Conceptual Design Study of an Unmanned Air Vehicle · 2012. 7. 29. · During design process,...

Istanbul Technical University

Faculty of Aeronautics and Astronautics

Aeronautical Engineering Department

Principles of Aircraft Design

UCK451E

Project Report

Conceptual Design Study of an Unmanned Air Vehicle

Instructor : Prof. Mehmet Şerif Kavsaoğlu

Course Assistant : Uğur Özdemir

Date : January 2010

Prepared by : GROUP 27

Halil ALTUN 110050015 [email protected]

Ulaş AKGÜN 110060024 [email protected]

mailto:[email protected]�mailto:[email protected]�

Abstract

ABSTRACT

The main aim of the aeronautical design project was to design a radio controlled mini Unmanned Air Vehicle with electric propeller motor type. This design will be processed step by step from initial sizing study than landing gear system. In all these progress, it will be focused on different aerodynamic and design parameters. Also in each study the formulas will be given and the necessary calculations will be made for the aircraft. All the studies performed throughout the semester was included in design report; Competitor Study, First Guess Sizing, Airfoil and Geometry Selection, Horsepower to Weight Ratio and Wing Loading, Initial Sizing, configuration layout and interior design, Propulsion and Fuel System Integration, Landing Gear and Subsystems, Aerodynamics.

The design will be made according to the needs by the mission, and the mission requirements have defined by the instructor of the course. After the calculations and researchs on which motor will be used then with the brand of “NEU 1506/3Y” was chosen to use for our aircraft. It has 1342 watt power and 0,6056kg weight as total. And battery type is “ELITE 1500 26-Cell Pack”. Propeller type is selected as “APC 18x12 E”.

The entire design requirements and mission were nearly met with a high accuracy with the competitor aircrafts at the end of the design processing. As a first step, starting from using the data of the competitor aircrafts was a very useful tool in the conceptual design.

Table of Contents

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TABLE OF CONTENTS

CHAPTER 1. INTRODUCTION 1

1.1. Purpose 1

1.2 Requirements and Mission Profile 1

1.3. About the Report 2

CHAPTER 2. COMPETITOR STUDY 3

2.1. Introduction 3

2.2. Progress 3

2.3.Specifications and Empirical Data 3

2.4.Aircraft Data Sheets 4

2.5. Plots from Aircraft Data Sheets 7

2.6. Histogram from Aircraft Data Sheets 10

2.7. Conclusion 10

CHAPTER 3. INITIAL CONFIGURATION SELECTION AND FIRST GUESS SIZING

11

3.1. Introduction 11

3.2 Aircraft Configuration and Figure of Merit 11

3.3. Internal and External Payload Compartment 14

3.4. First Guess Sizing and Wing Area Sizing for Specific Mission 16

3.5. Graphics 20

3.6. Conclusion 25

Table of Contents

ii

CHAPTER 4. AIRFOIL and GEOMETRY SELECTION 26


4.2. Progress and Airfoil Selection 26

4.3. Tail Geometry and Arrangements 31

4.4. Airfoil Geometry 32

4.5. Conclusion 32

CHAPTER 5. POWER TO WEIGHT RATIO AND WING LOADING 33


5.2. Power to Weight Ratio Estimation 33

5.3. Wing Loading Selection 35

5.4. Analysis 38

5.5. Summary of the Results 42

5.6. Conclusion 42

CHAPTER 6. INITIAL SIZING 43


6.2. Configuration Description 43

6.3. Mission Description 44

6.4. Rubber Engine Sizing Study 44

6.5. Motor Selection 45

6.6. Fixed Engine Sizing Study 46

Table of Contents

iii

6.7. Geomerty Sizing Study 47

6.8. Summary of Results 48

6.9. Conclusion 49

CHAPTER 7. CONFIGURATION LAYOUT AND INTERIOR DESIGN 50


7.2. Wing and Tail Surface 50

7.3. Locating Wing and Tail Surfaces 50

7.4. Wetted Area and Volume Determination

7.5. Conclusion

54

55

CHAPTER 8. PROPULSION AND FUEL SYSTEM INTEGRATION 56


8.2. Propulsion System Selection 56

8.3. Conclusion 63

CHAPTER 9. LANDING GEAR and SUBSYSTEMS 64


9.2. Landing Gear Arrangement 64

9.3. Tire Selection 65

9.4. Shock Absorber 67

9.5. Castoring Nose Wheel Geometry 67

9.6. Non Retractable Landing Gear 68

Table of Contents

iv

9.7. Subsystems 68

9.8. Conclusion 69

CHAPTER 10. AERODYNAMICS 70


10.2. Lift 70

10.3. Parasite(Zero Lift) Drag 72

10.4. Estimation Of Induced Drag (Drag due to Lift) 75

10.5. CL - α and Cl – Cd Curves 76

10.6. Tabulated Results 77

10.7. Conclusion 77

CHAPTER 12. CONCLUSION 78

REFERENCES

APPENDICES

79

80

Nomenclature

v

NOMENCLATURE Symbols A Aspect ratio, area

Definitions

a Speed of sound a/c Aircraft aoa Angle of attack Ap Footprint area B Distance between main and nose landing gears b Span C Specific fuel consumption (SFC) c Chord line Cbhp Propeller specific fuel consumption Cf Equivalent skin friction coefficient CD Drag coefficient CDo Parasite drag CL Lift coefficient Cm Pitching moment coefficient c/4 Quarter chord line D Drag, diameter E Endurance e Oswald span efficiency factor Eq. Equation F Fuselage lift factor FF Form factor Fig. Figure g Gravitational acceleration G Climb gradient H Height of c.g. h Altitude, height hp/W Horse power to weight ratio i Incidence angle K Drag-due-to-lift factor k Skin roughness value Kvs Variable swept constant l Characteristic length L Length, lift, yawing moment L/D Lift to drag ratio LHT Horizontal tail moment arm LVT Vertical tail moment arm M Mach number, mass, pitching moment Ma Distance from main gear to aft C.G. Mf Distance from main gear to forward C.G. n Load factor N Number of engines, Rolling moment Na Distance from nose gear to aft C.G.

Nomenclature

vi

Nf Distance from nose gear to forward C.G. nm Nautical mile P Pressure, power Q Component Interference factor, production quantity q Dynamic pressure R Range R/C, Vv Rate of climb Re Reynolds number Ref. Reference Rcutoff Cutoff Reynolds number Rr Rolling radius S Area, stroke Sexp Exposed area SG Ground roll distance ST Stroke of tire, transition distance Swet Wetted area Swf Flapped wing area T Thrust t Thickness, time t/c Maximum thickness of the section T/W Thrust to weight ratio u Upsweep angle V Velocity, volume W Weight w Width W/S Wing loading Wi Weight at i’th step Ww Weight on wheels X X location, nondimensionalized wrt. c Y Y location, nondimensionalized wrt. c y Location along the wing (0 at the centerline) Yp Engine-out moment arm Z Z location nondimensionalized wrt. c

Nomenclature

vii

Subscripts

Definitions

0L At the zero lift angle of attack ach Aerodynamic center of horizontal tail acv Aerodynamic center of vertical tail acw Aerodynamic center of wing alt Altitude c Component c/4 Quarter chord cg Center of gravity cr Critical e Elevator, empty exp Exposed F Flare FR Free roll f Flaps, fuselage, fuel fus Fuselage h Horizontal tail H.L. Hinge line HT Horizontal tail L Lift L&P Leakage and protuberance LE Leading edge m Maximum, pitching moment max Maximum min Minimum misc Miscellaneous ml Main lending gear np Neutral point o At takeoff, at zero lift p Payload r Rudder, root, required ref Reference SL Sea level SLS Sea level standard t Tip TO Takeoff TR Transition v Vertical tail VT Vertical tail w Wing, wheel wet Wetted Greek Symbols Definitions

Nomenclature

viii

α Angle of attack β Sideslip angle, yaw angle, Mach number correction factor δ Control surface deflection Λ Sweep angle σ Air density ratio ρ Air density ε Downwash angle γ Specific heat ratio of the air, climb angle ηp Propeller efficiency η Stroke efficiency ηt Tire efficiency λ Taper ratio ηh Ratio of the dyn pressures on hor.tail and the free stream ηv Ratio of the dyn pressures on ver.tail and the free stream µ Airplane relative density parameter, ground rolling resistance ψ Instantaneous turn rate

Acronyms

Definitions

APU Auxiliary power unit B.F.L. Balanced field length C.G. Center of gravity CUT Commuter utility transport DAPCA Development and procurement costs of aircraft EPU Emergency power unit ESHP Equivalent shaft horse power FFR Fuselage fineness ratio FTA Number of flight test aircraft KE Kinetic energy l.e. Leading edge L.F.L. Landing field length MAC Mean aerodynamic chord OEI One engine out RAT Ram air turbine SFC Specific fuel consumption SHP Shaft horse power t.e. Trailing edge TAI Turkish Aerospace Industries T.O.D. Takeoff distance T.O.F.L. Takeoff field length TOGW Takeoff gross weight TOP Takeoff parameter

Chapter 1 Introduction

1

1. INTRODUCTION Electric propeller radio controlled mini Unmanned Air Vehicles are mostly used for the specific missions. Mainly these aircrafts are being chosen because of the fact of their size can be minimized according to aerodynamic limits and during mission pilots’ confidence can be neglected by the reason of radio controlled. Mini air vehicles are also use for military tasks for discovering some dangerous and hazardous zones. Also there are several kinds of competitions which are organized for the aim of improving the young engineers’ abilities and contributing of their imaginations.

In this project, a conceptual design of a radio controlled mini unmanned air vehicle with electric propeller has been performed.

1.1 Purpose The aircraft will be designed as radio controlled mini Unmanned Air Vehicle. An electric propeller motor will be used for creating thrust power on this air vehicle.

1.2 Requirements and Mission Profile For a radio controlled mini unmanned aircraft with the following objectives perform a first guess sizing;

Range : 3657,6 m

Vcruise : 20 m/s

Wpayload : 2.835kg (5bats)

Design mission:

Figure 1.1 Required Mission Segments

1

4

0

3

2

5

Chapter 1 Introduction

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0–1: Takeoff 1–2: Climb and Turn 2–3: 500m Cuise 3–4: 360 deg Loiter 4–5: Landing

1.3 About the Report During design process, mainly the textbook, AIRCRAFT DESIGN: A CONCEPTUAL APPROACH, Raymer D.P, AIAA Education Series, Washington, 1992, is used. The ones which are used also will be mentioned in the part of references as a final step for this project. The tables, equations, and figures are mainly used from the text book of Raymer, 1992. In the related chapters the detailed drawings of the design aircraft will be given with all its necessary dimensions.

Chapter 2 Competitor Study

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2. COMPETITOR STUDY 2.1. Introduction

As a beginning it is needed to find the almost entire information about existing aircrafts which are in the category of electric propeller and radio controlled mini Unmanned Air Vehicle. And also these aircrafts could have the mass as maximum 10 kg and as minimum 3kg. The information needed has to include the basic features of unmanned air vehicles’ mainly about weight, structure, material, speed, motor and take off distance data.

Additionally, if they are published more, battery type, country made of, airfoil tip, endurance, take off distance and some more ratios by using some parameters. After collecting all values which are needed, they will be compared each other. Because it is aimed to find out the similarities and differences for each air vehicle by plotting the graphs and diagrams with some more revealing comments. 2.2 Progress

In this part, the differences and the similarities of 10 different air vehicles will be being worked on by using the values found and calculated. According to the research work restriction, they will have some same features, as they all have to have the mass in between 3kg to 10 kg and they all have to be electric propelled and radio controlled mini unmanned air vehicle. However it can not be expected that entire features will be the same for all because they had been built for different missions and these facts create the different features for each.

The aim of the design has to be considered in the beginning because while working on, if it is obligatory to make an optimization, everything must be clear and creator must know about the target. This is why each of these 10 air vehicles have some mission as Assembly mission, Ferry flight mission, Payload mission, Delivery mission, Surveillance flight mission, Store release flight mission, Asymmetric loads mission etc.

The project aircraft will be designed according to the missions that have to be done during competition.

2.3 Specifications and Empirical Data

The data needed mostly is found from JANE’s All the World’s Aircraft

, past years of DBF Reports and some other sources which are given in the references’ section. While making selection for the existing aircrafts, some of the criteria’s are:

-Electric Propeller -Radio Controlled mini Unmanned Air Vehicle -Take Off Mass:3-10 kg


4

2.4 Aircraft Data Sheets

Table 2.1: Data Sheets of Selected Aircrafts

Aircraft ATA 5 ATA 7 ATA -8 Kaan S.U.I.T Shock and Awe Reference Number(s) 1 2 3 4 5 Country Turkey Turkey Turkey Portugal America Configuration Conventional Conventional Canard V-Tail V-Tail Payload Weight, WP (kg) 2,85 2,94 2,04 2,28 4,24 Empty Weight, We (kg) 10,43 2,35 2,33 3,13 3,36 Battery Weight, Wf (kg) 0,78 0,6056 0,63 0,62 0,635

Battery Type ELITE 1500 26-

Cell Pack 15 GP 2200 Maximum Take Off Weight, W0 (kg) 14,06 5,9 4,99 6,03 6,08

Powerplant (Motor) NEU 1506/3Y Neu 1506/3Y NEU 1110/1.5Y Power (watt) 1000-1500 1342 Propeller APC 18x12 E 17x12 APC 12x6 APC Wing Span (m) 2,17 1,86 1,17 0,04 0,58 Wing Area (m2) 0,8 0,49 0,33 0,09 0,27 Wing Aspect Ratio 5,9 7 4,23 6,65 4 Wing Taper Ratio 0,05 0,65 Sweep Angle (c/4) 0 Wing thickness ratio (root) Wing thickness ratio (tip) Airfoil (root) DAE 21 SD 7043 SD 7062 NACA-64 LA203KB Airfoil (tip) DAE 21 SD 7043 SD 7062 NACA-64 LA203KB Flaps Horizontal Tail Volume Coeff., CHT 0,35 Vertical Tail Volume Coefficient, CVT 0,031 Horizontal Tail Airfoil NACA 0009 NACA 2410 CLARK-Y LA203KB SYM Vertical Tail Airfoil NACA 0009 NACA 0009 SD 7062 CLARK-Y LA203KB SYM Max. Wing Loading W0/S, (lb/ft^2) 3,6 2,66 2,05 1,28 4,62 Horsepower to Weight Ratio (h.p./lb) 0,03 0,13 0,117 0,03 0,03 Cruising Speed (m/s) 19,81 13,59 16,49 10,66 30,82 Stalling Speed, flaps up (m/s) 13,72 7,53 9,48 9,27 18,01 Staliing Speed, flaps


5

down (m/s) Take Off Distance (m) 45,72 7,25 30,48 60 30,48 Landing Distance (m) Endurance (minutes) Wing Structure and Materials Composite Composite Composite Composite Carbon fiber Fuselage Structure and Materials Composite Composite Composite Composite Composite Landing Gear Arrangement Tricycle Tricycle Tricycle Tricycle Tricycle

Landing Gear Structure Aluminum and

Composite Carbon Fiber

and Aluminum Aluminum and

Plastic Carbon Fiber Carbon Fiber

Main Wheel Diameter Steering Wheel Diameter

Aircraft ATA-9 Turbo

Encabulator Aeroshock Fast Back Team Shadow

Drag Reference Number(s) 6 7 8 9 10 Country Turkey U.S.A U.S.A U.S.A U.S.A Configuration Conventional V-tail conventional conventional conventional Payload Weight, WP (kg) 3,13 4,08 4,08 4,54 1,65 Empty Weight, We (kg) 2,27 1,25 1,36 12,13 3,26 Battery Weight, Wf (kg) 0,6 0,38 0,34 2,69 0,94 Battery Type ELITE 1500 15GP 2200 N GP 4600s Maximum Take Off Weight, W0 (kg) 6 5,29 5,44 22,79 7,45

Powerplant (Motor) NEU 1506/3Y Neu 1506/2Y NEU

1506/3Y Neu 1905/1.5Y Neu 1107 Power (watt) 480 watt 450 watt 630 watt 600 watt Propeller 16x10 APC 16x10 18x10E 18x10 APC 18x10 E Wing Span (m) 0,13 2,04 1,47 1,52 1,52 Wing Area (m2) 0,6 0,46 0,89 0,56 0,57 Wing Aspect Ratio 3,75 6,5 4,082 Wing Taper Ratio 0,45

Sweep Angle (c/4)

Wing thickness ratio Wing thickness ratio (tip)


6

Airfoil (root) Custom SD 7062 NACA 0009

Airfoil (tip) Custom SD 7062 NACA 0009

Flaps Horizontal Tail Volume Coeff., CHT Vertical Tail Volume Coefficient, CVT

Horizontal Tail Airfoil NACA 2412 Clark-y NACA 2408

Vertical Tail Airfoil NACA 0015 Inverted Clark-y NACA 2408 Max. Wing Loading W0/S, (lb/ft^2) 2,165 11,5 6,11 40,49 13,07 Horsepower to Weight Ratio (h.p./lb) 0,17 0,21 0,24 0,25 0,84

Cruising Speed (m/s) 13,86 14,34 15,24 16,52 22,4 Stalling Speed, flaps up (m/s) 19,6 12,26 10,67 11,49 16,4 Staliing Speed, flaps down (m/s)

Take Off Distance (m) 21,55 27,2 22,86 26,21 30,48

Landing Distance (m)

Endurance (minutes) 5 minutes Wing Structure and Materials Composite Carbon fiber Carbon fiber

Composite/Balsa Build Up Carbon fiber

Fuselage Structure and Materials Composite Carbon fiber Carbon fiber Composite Carbon fiber Landing Gear Arrangement Tricycle

Landing Gear Structure Carbon Fiber Aluminum and

Composite Carbon Fiber Carbon Fiber

and Aluminum Fiber glass

Main Wheel Diameter Steering Wheel Diameter


7

2.5 Plots from Aircraft Data Sheets

Figure 2.1: Empty Weight/Total Weight to Total Weight

Figure 2.2: Battery Weight/Total Weight to Total Weight

Figure 2.3: Wing span to Total Weight

0

5

10

15

20

25

0 0.5 1 1.5 2 2.5

b - Wo

b - Wo


8

Figure 2.4: Wing area to Total Weight

Figure 2.5: Total Weight for each Air vehicle

Figure 2.6: Payload Weight for each Air vehicle

0

5

10

15

20

25

0 0.2 0.4 0.6 0.8 1

S - Wo

S - Wo


9

Figure 2.7: Take off distance for each Air vehicle

Figure 2.8: Total weight/Wing area for each Air vehicle

Figure 2.9: Motor power(horse power)/Total weight for each Air vehicle


10

2.6 Histogram from Aircraft Data Sheets

Figure 2.10: Differences of Vcruise values for each Air vehicle

2.7 Conclusion

According to the parameters, the calculation, the graphs and the diagrams could

be plotted. This makes to realize the differences and similarities easily in one chart. The

air vehicles which has built for the same or similar aim they had the approximately same

values. It is obvious that all the parameters of an aircraft are dependent on each other. So

that means if one single feature of the aircraft will be changed, almost entire design

parameters has to be modified according to the airworthiness parameters. Also these facts

can be integrated to the heavy aircrafts. In real life, with this, it can be seen the mini air

vehicles performance and then decided to make the real heavy air craft.

Chapter 3 First Guess Sizing

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3. INITIAL CONFIGURATION SELECTION AND FIRST GUESS SIZING

3.1 Introduction In this part of the design process, we will decide which configurations, training and specifications to choose and calculate. This will include the aircraft’s weight, drag, aircraft configuration, speed, manufacturability, aerodynamic efficiency, internal and external payload compartments and also pilot and ground crew training.

3.2 Aircraft Configuration and Figure of Merit 3.2.1 Aircraft Configuration The most used aircraft configurations are conventional, canard, bi-plane, tri-plane and flying wing.

3.2.1.1 Conventional This is the most used configuration among all the configuration types. There is so much data that can be obtained about conventional aircrafts throughout the competitor study.

Figure 3.1: Conventional wing type

3.2.1.2 Canard The main advantage of the canard configuration is that it's unstable and in theory

could be safer by preventing stall / spin departures from coordinated flight. This works because as the angle of attack increases and approaches a "stall" condition, the canard stalls before the main wing. If the canard stalls it loses lift and thus the nose drops thus decreasing angle of attack and breaking the stall. To talk about disadvantages, fuel center of gravity lies farther behind aircraft center of gravity than in conventional designs. This means that a large center of gravity range is produced or that the fuel must be held elsewhere. The distance from the aircraft center of gravity to the most aft part of the airplane is usually smaller on canard aircraft. This poses a problem for locating a vertical stabilizer and may result in very large vertical surfaces. Most importantly, canard sizing is much more critical than aft tail sizing. By choosing a canard which is somewhat too big or too small the aircraft performance can be severely affected. It is easy to make a very bad canard design.


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Figure 3.2: Canard wing type

3.2.1.3 Bi-plane Aircraft built with two main wings can usually lift up to 20% more than can a

similarly sized monoplane of similar wingspan, which tends to afford greater maneuverability. The struts and wire bracing of a typical biplane form a box girder that permits a light but very strong wing structure. On the other hand, each wing negatively interferes with the aerodynamics of the other. For a given wing area the biplane produces more drag and less lift than a monoplane, but this effect can be slightly reduced by placing one wing forward of the other though NACA research into this suggests that the reduction of drag is minimal.

Figure 3.3: Bi-plane wing type

3.2.1.4 Tri-plane The advantages of a tri-plane are, the high aspect ratio of the wings allows an

excellent rate of climb and a good glide ratio if the wing interference drag can be reduced. A tri-plane can also have a high roll rate from the short span wings. Additionally it has a good longitudinal stability due to the extra lift centres. The disadvantages of a tri-plane are, it has high drag from wing interference and lift loss from six wing tips. Moreover, extra work is required to build the wings.

Figure 3.4: Tri-plane wing type


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3.2.1.5 Flying wing

A clean flying wing is theoretically the most aerodynamically efficient (lowest drag) design configuration for a fixed wing aircraft. It also offers high structural efficiency for a given wing depth, leading to light weight and high fuel efficiency. Because it lacks conventional stabilising surfaces or the associated control surfaces, in its purest form the flying wing suffers from the inherent disadvantages of being unstable and difficult to control. These compromises are difficult to reconcile, and efforts to do so can reduce or even negate the expected advantages of the flying wing design, such as reductions in weight and drag. Moreover, solutions may produce a final design that is still too unsafe for certain uses, such as commercial aviation.

Figure 3.5: Flying wing type

3.2.2 Figures of Merit 3.2.2.1 Rated Aircraft Cost

The aircraft is needed to be as cheap as possible without compromising the important specifications of the aircraft. There must be a balance between the other specifications and the cost.

3.2.2.2 Speed The higher the speed is, the faster and in the less time the mission can be completed.

3.2.2.3 Handling Qualities Handling of the aircraft must be easy so that the pilot can fly, take off and land on the

airplane as efficiently as possible.

3.2.2.4 Manufacturability Manufacturability is an important factor on the design process because it critically

states the rated aircraft cost.

3.2.2.5 Aerodynamic Efficiency Aerodynamic efficiency is one of the key parameters that determines the weight and

cost of an aircraft. Roughly speaking, an aircraft's range is directly proportional to its aerodynamic efficiency without any increase in fuel usage.


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3.3 Internal and External Payload Compartment In the design process, the payload compartments must be diligently designed so that

the loading and unloading can be done easily and quickly, also the compartment shouldn’t have a big effect on the aircraft’s weight. Moreover, the compartments must be designed so well that the center of gravity of the aircraft should be stable when it is loaded.

In the figure-3.6, the drawing of the aircraft’s fuselage with it’s internal compartment is shown.

Figure 3.6: Aircraft’s fuselage with internal compartment

The aircraft must be able to carry 6 to 10 softballs without the center of gravity being changed. So here are some top views that show how the balls can be applied to the compartment.

Figure 3.7: Loading possibilites for 6 balls


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Figure 3.8: Loading possibility for 10 balls

To stabilize the balls in the y-axis (the axis on the direction of ground to sky), a cover is designed as shown in figure – 3.9:

Figure 3.9: Internal compartment cover design

Also external payload compartment is needed. This compartment can cause an increase in weight, so we can make something that can be installed over the cover of the payload compartment in case of external loading and the design must be as light as possible. As shown in figure –3.10, in the middle of the design, there is an additional support for fixing the bats. The cuffs are on a rail system as it is seen on the other figure, so the cuffs can be adjusted to the height of the bats.

Figure 3.10: External compartment (fixing supports open on left, closed on right)


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The installed external compartment on the cover is shown in figure – 3.11.

Figure 3.11: Installed external compartment on the cover

3.4 First Guess Sizing and Wing Area Sizing for Specified Mission

3.4.1 Design missions and results

3.4.1.1 Mission 1a

Empty take off Range = 2x4000=8000 ft (2Laps along the flight course) Vcruise = 15m/s First guess take off Weight : 8.4 kg Cl/Cd=10

By using Excell formulations, take off weight for Mission 1a was calculated.

Table-3.1 Mission 1a W0 Estimated (kg) 8.4 W0 Calculated (kg) 0.04

t (sec) 162.56 We (kg) 0.02 P0 (kg) 1.12

PCR (Watt) 0.84 V (Volt) 0.03 E (Volt) 0.03

Energy (Joule) 156.74 Q (Coloumb) 1408.28

S (m2) 0.3

Wbattery (kg) 0.02 n 1


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3.4.1.2 Mission 1b

In this mission everything is the same as Mission 1-a except Vcr = 20m/s.

Table-3.2 Mission 1b

W0 Estimated (kg) 8.4 W0 Calculated (kg) 0.04

t (sec) 121.92 We (kg) 0.02 P0 (kg) 1.5



S (m2) 0.3


3.4.1.3 Mission 2a

Take off with 10 softballs Wpayload = 10*200gr = 2000gr Range = 3x4000=12000ft (3Laps along the flight course) Vcr = 15 m/s Cl/Cd=10

Table-3.3 Mission 2a


t (sec) 243.84 We (kg) 1.49 P0 (kg) 99.72



S (m2) 0.41



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3.4.1.4 Mission 2b




t (sec) 182.88 We (kg) 1.51 P0 (kg) 134.54



S (m2) 0.41


3.4.1.5 Mission 3a

Take off with 3 bats Wpayload: 3*567 g = 1701 g. Range = 3x4000=12000ft (3Laps along the flight course) Vcr = 15 m/s Cl/Cd=10

Table-3.5 Mission 3a


t (sec) 243.84 We (kg) 1.43 P0 (kg) 481.22



S (m2) 0.4



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3.4.1.6 Mission 3b




t (sec) 182.88 We (kg) 1.24 P0 (kg) 112.66



S (m2) 0.07


3.4.1.7 Mission 3c

In this mission everything is the same as Mission 3-a except Wpayload=5x567=2835gr.

Table-3.7 Mission 3c


t (sec) 243.84 We (kg) 2.26 P0 (kg) 145.51

PCR (Watt) 109.13 V (Volt) 3.64 E (Volt) 4


S (m2) 0.46



20

3.4.1.8 Mission 3d

In this mission everything is the same as Mission 3-b except Wpayload=5x567=2835gr

Table-3.8 Mission 3d


t (sec) 182.88 We (kg) 2.29 P0 (kg) 195.85



S (m2) 0.46


3.5 Graphics After making the calculations through a series of iterations, we have plotted the

needed graphs such as W0 – Wpayload , We – Wpayload , Wbattery – Wpayload , Swing – Wpayload , P0 – Wpayload , W0 – Vcr , We – Vcr , Wbattery – Vcr , Swing – Vcr , P0 – Vcr.

3.5.1 For Vcruise = 15 m/sec; Wo-Wpayload, We-Wpayload, Wbattery-Wpayload, Swing-Wpayload Po-Wpayload

Figure 3.12 Wo - Wp

0.00

1.00

2.00

3.00

4.00

5.00

6.00

0.000 0.500 1.000 1.500 2.000 2.500 3.000

W0-Wp

W0-Wp V=15m/s


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Figure 3.13 We-Wp

Figure 3.14 Wf-Wp

Figure 3.15 Swing-Wp

0.00

0.50

1.00

1.50

2.00

2.50

0.000 0.500 1.000 1.500 2.000 2.500 3.000

We-Wp

We-Wp V=15m/s

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.000 0.500 1.000 1.500 2.000 2.500 3.000

Wf-Wp

Wf-Wp V=15m/s

0.00

0.10

0.20

0.30

0.40

0.50

0.000 0.500 1.000 1.500 2.000 2.500 3.000

Swing-Wp

S_wing-Wp V=15m/s


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Figure 3.16 Po-Wp

Figure 3.17 Wo-Wp

Figure 3.18 We-Wp

0.00

100.00

200.00

300.00

400.00

500.00

600.00

0.000 0.500 1.000 1.500 2.000 2.500 3.000

Po-Wp

Po-Wp V=15m/s

-1.000.001.002.003.004.005.006.00

0.000 1.000 2.000 3.000

W0-WpDoğrusal (W0-Wp)

W0-Wp V=20m/s

-0.50

0.00

0.50

1.00

1.50

2.00

2.50

0.000 0.500 1.000 1.500 2.000 2.500 3.000

We-WpDoğrusal (We-Wp)

We-Wp V=20m/s


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Figure 3.19 Wf-Wp

Figure 3.20 Swing-Wp

Figure 3.21 Po-Wp

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.000 0.500 1.000 1.500 2.000 2.500 3.000

Wf-WpDoğrusal (Wf-Wp)

Wf-Wp V=20m/s

0.00

0.10

0.20

0.30

0.40

0.50

0.000 0.500 1.000 1.500 2.000 2.500 3.000

Swing-Wp

S_wing-Wp V=20m/s

-50.00

0.00

50.00

100.00

150.00

200.00

250.00

0.000 0.500 1.000 1.500 2.000 2.500 3.000

Po-WpDoğrusal (Po-Wp)

Po-Wp V=20m/s


24

Figure 3.22 Wo-Vcr

Figure 3.23 We-Vcr

Figure 3.24 Wf-Vcr

3.563.563.573.573.583.583.593.593.603.603.61

0 5 10 15 20 25

W0-VcrDoğrusal (W0-Vcr)

W0-Vcr Wp=2000gr

1.49

1.49

1.50

1.50

1.51

1.51

1.52

0 5 10 15 20 25

We-VcrDoğrusal (We-Vcr)

We-Vcr Wp=2000gr

0.00

0.02

0.04

0.06

0.08

0.10

0 5 10 15 20 25

Wf-VcrDoğrusal (Wf-Vcr)

Wf-Vcr Wp=2000gr


25

Figure 3.25 Swing-Vcr

Figure 3.26 P0-Vcr

3.6 Conclusion In this section of the design, we started making initial guesses for our aircraft specifications. Through the calculations made, we tried to have a zero error between our estimation and calculation so that we could find the right value for the specification. These calculations have been made for 8 missions given. For each mission, we calculated the take off weight, empty weight and battery weight. Also we estimated and calculated the take off power, P0 and wing area.

0.41

0.41

0 5 10 15 20 25

Swing-Vcr

Swing-Vcr Wp=2000gr

0.0020.0040.0060.0080.00

100.00120.00140.00160.00

0 5 10 15 20 25

P0-VcrDoğrusal (P0-Vcr)

P0-Vcr Wp=2000gr

Chapter 4 Airfoil and Geometry Selection

26

4. AIRFOIL & GEOMETRY SELECTION

4.1 Introduction

As a beginning it is needed to calculate the value of Cl (lift coefficient) and Reynold number by using the previous studies which has been done. Because according to the researches and calculations made the value of W (weight), S (wing area), ρ(density), c(chord line), µ(viscosity), V(velocity) are already known. Then it will be dicussed about the airfoil characteristics of the wing in terms of Reynolds Number, Cl,design, Cl,max, L/D)max Cd0, Cm0, stall characteristics, thickness ratio, weight, ease of manufacturing. These all are important to have the nearly best design. While making this process, 5 different type of airfoils will be taken as examples, then they all will be compared each other with their variety of aerodynamic characteristic.

After deciding on the best airfoil for the aircraft it will be discussed about aspect ratio, sweep angle, twist, wing incidence, dihedral, wing vertical location and wing tip. These same process will applied for finding the tail characteristics tail geometry and arrangements as well. Additionally with plotting some graphics, it will try to be seen much more obvious that why the one has chosen as the best airfoil.

4.2 Progress and Airfoil Selection According to the values it was mentioned in introduction part, we have to find the

lowest value of Reynolds number and Cl,design. The situations of the air as we know about the mission, all the unknowns as W (weight), S (wing area), ρ(density), c(chord line), µ(viscosity), V(velocity) are known as follows;

W= 5.19kg c= 0,35 m S= 0.46 m2 µ=18,97x10^-6 Pa.s ρ= 1.1 kg/m3 V= 15 m/s

Then with the formulas of Re = ρVdµ

and Cldesign =W

12ρV

2S

they can be calculated as:

Re = ρVdµ

= (1.1)∗(15)∗(0.35)18,97∗10−6

= 3,04*105 Cldesign =

W12ρV

2S= 5,191

2(1,1)(152)0,46

= 0,0912

According to these calculations and the airfoils selected we will try to find out the best airfoil for our aircraft. Here are some graphics to consider easily.


27

4.2.1 Graphics for NACA 64

Figure 4.1 Cl-Alpha and Cd-Alpha changes

Figure 4.2 Cl-Cd

Figure 4.3 ClCd-Alpha_Cm-Alpha


28

4.2.2 Graphics for NACA 2412


Figure 4.5 Cl-Cd



29

4.2.3 Graphics for DAE 21(chosen)


Figure 4.8 Cl-Cd



30

4.2.4 Graphics for CLARK-Y


Figure 4.11 Cl-Cd



31

Table 4.1 Comparison of the aerodynamic values for airfoils

It was chosen DAE 21 as the airfoil type for our aircraft during the mission. Because we want to have the high value of the Clmax and low value of Cmα in between some criterias. Of course there are some limits for choosing these values, in the meaning of, if it is going to be chosen the highest value of Clmax it will be difficult to control the aircraft and also will be hard to control it. On the other hand if the value Cmα will be the lowest value of itself the pith down effects will be occurred and the aircraft will have the tendency of loosing altitude all the time. In case of these all to have the desired stall characteristics the relatively fat airfoils are chosen because of the fact of stall phenomena starts from the trailing edge across to the way of wing upper surface. And also Fat airfoils are described to contain more than %15 thickness ratio. Deciding about the best airfoil for the missions has to be done, even if the values of Cd was a bit higher than the others, DAE21 is the best of these airfoils worked on.

While working on aeronautical effects also it has to be considered about ease of

manufacturing. Since DAE profiles are the comman-used profile types, producing it is not going to cost more than the other with the advantageous of we can ask more related with our misson.


4.3 Tail Geometry and Arrangements As it is explored on the second phase of our design process, the conventional tail was

chosen. According to Daniel Raymer’s book, conventional tail provides adequate stability and control at the lightest weight. It also has a good pitch up recovery and provides a good spin maneuverability. It also has a good stall recovery at high angle of attacks, because the downwash effect from the wings don’t effect the rudders in the tail much in these angles. Airfoil of the tail is chosen the same as the wing airfoil.


32

Tail aspect ratio is chosen 3 just below the wing aspect ratio. Because it would weight

a lot and isn’t necessary for our kind of aircraft, no sweep was given to the wing. A little twist is given to have a gradual stall starting from root to the tip of the wing. A dihedral effect is not necessary for the aircraft due to the missions expected.

Vertical Tail Horizontal Tail (Conventional)

Airfoil Type NACA009 NACA 0009 A 1,4 3,755 λ 0.3 0.3

4.4 Airfoil Geometry

Figure 4.14 Airfoil Geometry

4.5 Conclusion I n this chapter selection of the airfoil, the wing and tail geometry were performed. These results were important already to be used in the following chapters. Also in later design processes and their manufacturability, validity will be checked by mostly focusing on other factor and in later studies. On the other hand, the results which were tabulated on the tables are also important while discussing and making decision for the airfoil geometry and for the other parameters related with our aircraft.

Chapter 5 Power to Weight Ratio and Wing Loading

33

5. POWER TO WEIGHT RATIO AND WING LOADING

5.1 Introduction With the performance requirements, atmospheric conditions and some other parameters given, we are supposed to estimate and tabulate some parameters with the help of our analysis. We are asked to find stall speed in four conditions: empty-flaps up, empty-flaps down, fully loaded-flaps up, fully loaded-flaps down. Then we need to find the take off ground roll, best range speed and maximum speed for both empty and fully loaded conditions. Finally, we have to find climb gradient and rate of climb at the beginning of climb.

5.2 Power To Weight Ratio Estimation In the process of power to weight ratio selection, firstly we shall check our competitor study. According to the data found on the competitor study, power to weight ratio can be estimated as the average of the values, which is 0,2047 h.p/lb. Another selection must be made based on the thrust matching calculations. From thrust matching calculations we will have different values of W/S (wing loading). Then at the end one of them will be decided to use finding all the answers for the questions asked. • From Competitor Study: Power to Weight Ratio (h.p./kg)

0,03 0,13 0,117 0,03 0,03 0,17 0,21 0,24 0,25 0,84

𝑃𝑊

=0,03 + 0,13 + 0,117 + 0,03 + 0,03 + 0,17 + 0,21 + 0,24 + 0,25 + 0,84

10

𝑃𝑊

= 0,1828 ℎ. 𝑝/𝑘𝑔

• To find the values of Power to weight ratio, some parameters are given and

calculated as follows;

𝜂𝑝 = 0,6 ρ=1,1kg/m3 µ𝑇𝑂 ≌ 1,0 𝐶𝐿 ≌ 0,5 𝐶𝐷0 = 0,025 AR = 3,755 e = 0,8


34

𝑇

𝑊)𝑐𝑟𝑢𝑖𝑠𝑒 =

1𝐿𝐷)𝑐𝑟𝑢𝑖𝑠𝑒

= 1𝐿𝐷)𝑚𝑎𝑥

𝑇.𝑉 = 𝜂𝑃 .𝑃 𝑇 =𝜂𝑃𝑉

.𝑃

𝜂𝑝.𝑃𝑉.𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 =1

𝐿𝐷)𝑚𝑎𝑥

𝑃𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 =𝑉𝜂𝑝

. 1𝐿𝐷)𝑚𝑎𝑥

𝐾 = 1

п.𝐴𝑅.𝑒

𝐾 = 1п.(3,755).(0,8)

𝐾 = 0,10596

𝐿𝐷

)𝑚𝑎𝑥 =𝐶𝑙

𝐶𝑑𝑜+𝐾𝐶𝑙^2 𝐿

𝐷)𝑚𝑎𝑥 = 9,7146

𝑉𝑇𝑂 = (1,1)𝑉𝑠𝑡𝑎𝑙𝑙 𝑉 = 0,7𝑉𝑇𝑂

𝑉𝑇𝑂 = (1,1). 10 𝑉 = 0,7.11

𝑉𝑇𝑂 = 11 𝑚/𝑠 𝑉 = 7,7 𝑚/𝑠 𝑃

𝑊)𝑐𝑟𝑢𝑖𝑠𝑒 =

𝑉𝜂𝑝

. 1𝐿𝐷)𝑚𝑎𝑥

𝑃𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 =7,70,6

. 19,7146

𝑃𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 = 1,321 ℎ. 𝑝/𝑙𝑏

𝑃𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 = 0,5941 ℎ. 𝑝/𝑘𝑔 Since we would like to have more horse power according to the weight, the bigger power to weight ratio will be chosen. So the value of power to weight ratio is 0,59 h.p/kg.

𝑃𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 = 0,59 ℎ.𝑝/𝑘𝑔

Table 5.1 Selection of P/W

P/W [hp/kg]

Competitor Study 0.1828

Thrust Matching 0.5941


35

5.3 Wing Loading Selection In the process of wing loading selection, firstly we shall check our competitor

study. According to the data found on the competitor study, wing loading can be estimated as the average of the values, which is 12,7545 lb/ft2.

𝑾𝑺

= 12,7545 [kg/m2]

5.3.1 Stall Speed Requirements (𝐶𝐿)max𝐿/𝐷 = �𝐶𝐷0 .𝜋.𝐴𝑅. 𝑒 => (𝐶𝐿)max𝐿/𝐷 = �0,025𝑥𝜋𝑥3,755𝑥0,8 =0,2359

𝑊𝑆

=12𝜌𝑉𝑠𝑡𝑎𝑙𝑙2 𝑐𝐿𝑚𝑎𝑥

𝑊𝑆

= 12

(1,1). 102. 0,2359

𝑊𝑆

= 12,97 𝑁/𝑚2

5.3.2 Take off Ground Roll Requirements

𝑆𝐺 =1

2𝑔𝐾𝐴ln (𝐾𝑇+𝐾𝐴𝑉𝐹

2

𝐾𝑇+𝐾𝐴𝑉𝑖2)

𝐾𝑇 =

𝜂𝑝𝑉

𝑃𝑊− 𝜇

𝐾𝐴 =𝜌

2.(𝑊/𝑆) . (𝜇𝑐𝐿 − 𝑐𝐷0 − 𝐾. 𝑐𝐿

2)

To find the values of KT and KA, some parameters are given and

calculated as follows;

𝜂𝑝 = 0,6 ρ=1,1kg/m3 µ𝑇𝑂 ≌ 1,0 𝐶𝐿 ≌ 0,5 (0,7 with flaps) 𝐶𝐷0 = 0,025 AR = 3,755 e = 0,8

𝑃

𝑊= 1,321

SG=90ft=27,4mt g=9,81


36

𝐾𝑇 =𝜂𝑝𝑉𝑃𝑊 − 𝜇

𝐾𝑇 =0,67,7

.1,321 − 0,1 𝐾𝑇 = 2,9351 𝑥 10−3

𝐾𝐴 = 12𝑔 𝑆𝐺 𝑙𝑛 �𝑉𝑇𝑂2

2𝑔 𝑆𝐺𝐾𝑇�= 1

2𝑥9.81𝑥27.432ln � 7.7

2

2𝑥9.81𝑥27.432𝑥0.002935�

𝐾𝐴 == 6,7356𝑥10−3

• Flaps up (CL=0.5) :

𝑊𝑆

= 12𝐾𝐴.

�𝜇𝐶𝐿 − 𝐶𝐷0 − 𝐾𝐶𝐿2� = 1

2𝑥0,006736. (1𝑥0,5 − 0,025 − 0,10596𝑥0,52)= 33,29 𝑘𝑔

𝑚2

• Flaps down (CL=0.7) :

𝑊𝑆

= 12𝐾𝐴.

�𝜇𝐶𝐿 − 𝐶𝐷0 − 𝐾𝐶𝐿2� = 1

2𝑥0,006736. (1𝑥0,7 − 0,025 − 0,10596𝑥0,72)= 46,25 𝑘𝑔

𝑚2

5.3.3 Cruising speed requirement

𝑉𝑐𝑟𝑢𝑖𝑠𝑒 =𝜂𝑝. (𝑃/𝑊)

�𝑞.𝐶𝐷𝑂𝑊/𝑆 � + �𝑊/𝑆

𝑞.𝜋.𝐴𝑅. 𝑒�

𝑞 = 1

2𝜌𝑉𝑐𝑟𝑢𝑖𝑠𝑒2

𝑞 = 12

(1,1)(25)2 𝑞 =343,75 kg/m.s^2

𝑉𝑐𝑟𝑢𝑖𝑠𝑒 =𝜂𝑝.(𝑃/𝑊)

�𝑞.𝐶𝐷𝑂𝑊/𝑆 �+�𝑊/𝑆

𝑞.𝜋.𝐴𝑅.𝑒�

25 =(0,6). (1,321)

�(343,75). (0,025)𝑊𝑆

� + � 𝑊/𝑆(343,75).𝜋. (3,775). (0,8)�


37

*If (W/S) a , the equation turns into;

a2 - 103,4453xa 28027,363 = 0 a1 = - 108,2 (can not be negative) a2= 210,2 (selected) √ *Converting from British System to SI System 𝑊𝑆

= 210,2𝑥2,205

10,764= 43,05 𝑘𝑔/𝑚2

Since the value of wing loading must be the most suitable one for all

conditions, one of these wing loading results has to be chosen. The lowest value of wing loading is calculated with the method of Stall speed, and the highest value of the wing loading is calculated with the method of T.O Ground roll (flaps down). Here it has to be chosen the minimum result of W/S but because of the fact that producing the lowest value of W/S is going to be a more expensive than the others.

On the other hand choosing the highest value of W/S is not going to be good

idea because the result will not fit in all the equations. So this is why it has to be estimated all situations and has to be chosen the optimal value of W/S. According to the reasons mentioned about the value of 33,29 was chosen as W/S.

Table 5.2 Selection of W/S

Method W/S (kg/m^2) Competitor Study 12,75 Stall Speed (10 m/s) 12,97 T.O. Ground Roll (Flaps up) 33,29 T.O. Ground Roll (Flaps down) 46,25

CruiseSpeed 43,05


38

5.4 Anlysis After calculating the values of power to weight ratio and wing loading now all

the speed requirements can be found by putting the unknowns of P/W and W/S in the formulas.

5.4.1 Stall Speed: Empty and flaps up

𝑉𝑠𝑡𝑎𝑙𝑙 = �2(𝑊𝑆 )𝑒𝜌𝐶𝐿𝑚𝑎𝑥

= �2𝑥(33.29)

1.1𝑥0.5= 11,00

𝑚𝑠𝑛

5.4.2 Stall Speed: Empty and flaps down

𝑉𝑠𝑡𝑎𝑙𝑙 = �2(𝑊𝑆 )𝑒𝜌𝐶𝐿𝑚𝑎𝑥

= �2𝑥(33.29)

1.1𝑥0.7= 9,29

𝑚𝑠𝑛

5.4.3 Stall Speed: Fully Loaded and flaps up

𝑉𝑠𝑡𝑎𝑙𝑙 = �2(2𝑥𝑊𝑆 )𝑒𝜌𝐶𝐿𝑚𝑎𝑥

= �2𝑥(66.58)

1.1𝑥0.5= 15,56

𝑚𝑠𝑛

5.4.4 Stall Speed: Fully Loaded and flaps down

𝑉𝑠𝑡𝑎𝑙𝑙 = �2(2𝑥𝑊𝑆 )𝑒𝜌𝐶𝐿𝑚𝑎𝑥

= �2𝑥(66.58)

1.1𝑥0.7= 13,15

𝑚𝑠𝑛

5.4.5 Take Off Ground Roll: Empty

𝑆𝐺 =1

2.𝑔.𝐾𝐴 𝑙𝑛 �

𝐾𝑇+𝐾𝐴𝑉𝑓2

𝐾𝑇+𝐾𝐴𝑉𝑖2�

Vf = VTO = 1,1x Vstall = 11m/s

Vi = 0 m/s

𝐾𝐴 =𝜌

2.(𝑊/𝑆) . (𝜇𝑐𝐿 − 𝑐𝐷0 − 𝐾. 𝑐𝐿

2) = 0,007410

𝐾𝑇 =𝜂𝑝𝑉

𝑃𝑊− 𝜇 = 0,0029351

𝑆𝐺 =1

2𝑥(9,81)𝑥(0,007410) 𝑙𝑛 �0,0029351+(0,007410)𝑥(11)

2

0,0029351� = 39,379 m


39

5.4.6 Take Off Ground Roll: Fully Loaded

𝑆𝐺 =1

2.𝑔.𝐾𝐴 𝑙𝑛 �

𝐾𝑇+𝐾𝐴𝑉𝑓2

𝐾𝑇+𝐾𝐴𝑉𝑖2�

Vf = VTO = 1,1x Vstall = 11m/s Vi = 0 m/s 𝑃𝑊

)𝑒𝑚𝑝𝑡𝑦 = (2,0)𝑃𝑊

)𝑙𝑜𝑎𝑑𝑒𝑑 𝑊𝑆)𝑒𝑚𝑝𝑡𝑦 = (0,5)

𝑊𝑆)𝑙𝑜𝑎𝑑𝑒𝑑

𝐾𝐴 =𝜌

2.(𝑊/𝑆) . (𝜇𝑐𝐿 − 𝑐𝐷0 − 𝐾. 𝑐𝐿

2) = 0,003705

𝐾𝑇 =𝜂𝑝𝑉

𝑃𝑊− 𝜇 = 0,048532

𝑆𝐺 =1

2𝑥(9,81)𝑥(0,003705) 𝑙𝑛 �0,048532+(0,003705)𝑥(11)

2

0,048532� = 32,998 m

5.4.7 Best Range Speed: Empty

𝑉𝑏.𝑟 = 𝑉𝑚𝑎𝑥𝐿𝐷

= �2𝜌

.𝑊𝑆

.�𝐾𝑐𝐷0

𝐾 = 0,10596 𝑐𝐷0 = 0,025 𝑊𝑆

= 33,29

𝑉𝑚𝑎𝑥𝐿𝐷

= 11,16 m/s

5.4.8 Best Range Speed: Fully Loaded

𝑉𝑏.𝑟 = 𝑉𝑚𝑎𝑥𝐿𝐷

= �2𝜌

.𝑊𝑆

.�𝐾𝑐𝐷0

𝐾 = 0,10596 𝑐𝐷0 = 0,025 𝑊𝑆

= 66,58

𝑉𝑚𝑎𝑥𝐿𝐷

= 15,79 m/s


40

5.4.9 Maximum Speed: Empty

𝑇𝑊



𝐿𝐷

)𝑚𝑎𝑥 =12�

п𝐴𝑒𝑐𝐷0

= 9,7146

𝑇𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 = 0,10294

𝑇𝑊

=𝑞.𝐶𝐷𝑜𝑊𝑆

+ 𝑊𝑆

𝑞. π.𝐴. 𝑒 (1)

*From (1) formula when q unknown found as q1= 58,53 , √ q2= 78,53 (selected) q=( ½).ρ.Vmax2 Vmax=11,95 m/s

5.4.10 Maximum Speed: Fully Loaded

𝑇𝑊



𝐿𝐷

)𝑚𝑎𝑥 =12�

п𝐴𝑒𝑐𝐷0

= 9,7146

𝑇𝑊

)𝑐𝑟𝑢𝑖𝑠𝑒 = 0,10294

𝑇𝑊

=𝑞.𝐶𝐷𝑜𝑊𝑆

+ 𝑊𝑆

𝑞. π.𝐴. 𝑒 (1)

*From (1) formula when q unknown found as q1= 37,122 , √ q2= 237,022 (selected) q=( ½).ρ.Vmax2 Vmax=20,76 m/s

5.4.11 Climb gradient and rate of climb at the beginning of climb

Before calculating the climb gradient, first we need to find Vclimb and q values :

𝑉𝑐𝑙𝑖𝑚𝑏 = 1,2 .𝑉𝑠𝑡𝑎𝑙𝑙 𝑉𝑐𝑙𝑖𝑚𝑏 = 1,2 .10 𝑉𝑐𝑙𝑖𝑚𝑏 = 12

𝑚𝑠


41

𝑞 =12𝜌𝑉𝑐𝑙𝑖𝑚𝑏2

𝑞 =12

. 1,1 . 122

𝑞 = 79,2 kg/ms2

After finding Vclimb and q, we can calculate T/W and D/W to find climb gradient (G) :

𝑇𝑊

)𝑇𝑂 =𝜂𝑝

𝑉𝑐𝑙𝑖𝑚𝑏𝑃𝑊

𝑇𝑊

)𝑇𝑂 =0,612

.0,1828 .550

𝑇𝑊

)𝑇𝑂 = 5,027

𝐷𝑊

=𝑞𝑆𝑐𝐷0 + 𝑞𝑆 �

𝑐𝐿2п𝐴𝑒�

𝑊=𝑆𝑊

. ( 𝑞𝑐𝐷0 + 𝑞 �𝑐𝐿2

п𝐴𝑒�)

𝐷𝑊

=1

66,58 . (79,2 . 0,025 + 79,2 .�

0,72

п . 3,755 .0,8�)

𝐷𝑊

= 0,0915

Now we can calculate the climb gradient :

𝐺 =𝑇𝑊−𝐷𝑊

𝐺 = 5,027 − 0,0915 = 4,9355 Then to find the rate of climb (Vy):

𝐺 =𝑉𝑦

𝑉𝑐𝑙𝑖𝑚𝑏

𝑉𝑦 = 𝐺𝑉𝑐𝑙𝑖𝑚𝑏 𝑉𝑦 = 4,9355 .12 = 59,226 𝑚/𝑠


42

5.5 Summary of Results

Table 5.3 The results in SI and British Units

5.6 Conclusion In this section of the design process, we have decided our horsepower to weight ratio and wing loading. In the decision process of power to weight ratio selection, we have used competitor study’s data and thrust matching, in the decision process of wing loading selection we have used competitor study’s data, stall speed requirement, take off ground roll requirement and cruising speed requirement. After making the calculations, we have made the final decision based on the results. In those calculations, we have considered cL values when flaps are down, so that leaded us to a decision that we can use a plane flap on our airplane.

Chapter 6 Initial Sizing

43

6. INITIAL SIZING

6.1 Introduction In this section of the design, firstly, we are asked to perform a rubber engine sizing

according to a similar mission in one of the missions in the second part of our design process. The airplane has to carry 5 bats in a flight course with 3 laps along which has 12000 ft. cruise. After the sizing, we shall select a motor according to the competitor study wmade for first study. We collected the information about the past aircrafts which was made before then we have already had some pr-study about this motor selection. Motor selection will be made by using the aerodynamic calculations, some engineering estimations and the other parameters of our aircraft. And then we will perform a fixed engine sizing by keeping and using the old value of 𝑃

𝑊)o and we will recalculate the range. Finally we need to perform a geometry

sizing study and show the results in a table format.

6.2 Configuration Description

From previous of our studies we have already know some idea about the changes in between Horsepower to Weight Ratio (watt/kg) and Maximum Take Off Weight, W0 (kg). According to this change with the parameters in our competitor studies when we put the inputs to the graphic which Wo is in “x” axis and P

W is in “y” axis, and then when we plot a

line which passes approximately around these points, we would have the lineer equation of:

y= (-33,135).x + 336,17

Here this equation will be using to estimate the Take off weight, Empty Weight, and Battery Weight.

Po𝑊𝑜

= (−33,135). Wo + 336,17

This equation is formed by rubber engine sizing study, which is shown in the 4th part of the report.

Figure 6.1 Conceptual Sketches (from top and side view)


44

6.3 Mission Description

Design mission is similar to a mission in one of the mission in the second part of our design process. That was our last mission which called as 3-d. Our plane has to take off, travel a 4000 feet circuit for 3 times with a 20 m/sec cruise speed and land on. It has to carry 5 bats, which corresponds to a 2835 gram payload weight.

6.4 Rubber Engine Sizing Study In the second part of our design process, we have used a quick method of sizing an aircraft. In this part, we have to make some refining to the sizing equation so that we can get better and more realistic results. As a first step of our calculations we have to know about the

changes between Po𝑊𝑜

and Wo. So here it is the relations in between mentioned again as plotted

in the graphic.

Po𝑊𝑜

= (−33,135). Wo + 336,17

Figure 6.2 Relation between Wo - Po/Wo

So then we will start our calculation with guessing the value of Wo as 8,40. When this parameter is put in the equation we will be finding the value of Po/Wo as 138,3078 watt/k.g. And Wo is found as 5,97kg. However to find the Wo_calculated as 5.97 it has to be made 46 steps of iterations. This is why after that as the value of Wo was taken 6,40. After 14 steps of iterations approximately the same value of Wo=5,97.

• Calculation of Po (power,watt)

Po =Po𝑊𝑜

. Wo = (138,3078)x(5,97) ----------> Po = 825,89 watt

y = -33,135x + 336,17

050

100150200250300350400

0 5 10 15

Wo - P_o/W_o

wo-p/wDoğrusal (wo-p/w)


45

• Calculation of S (wing area)

Again while using the values and parameter from competitor study we have the equation which shows the relations in between S and Wo. The equation plotted and found is:

S=(0,0305)x(Wo)+0,2902 And with using this formula;

𝑆 = (0,0305)x(Wo) + 0,2902 = (0,0305)x(5,97) + 0,2902 𝑆 = 0,4723 𝑚2

• Calculation of We (empty weight)

We =WeW0

xW0 = 0,44 x 5,97 = 2,6268 kg

6.5 Motor Selection In this section we will decide for a motor as which one is going to be the best for our aircraft and the missions. To make this decision of course we have some parameters and some values which are calculated as P0 (power needed), and how many batteries we need to use, battery weight allowed etc. However while deciding the motor, the most important criteria was “how much trust power needed”.

Of course there is variety kind of motors but we have to choose the one it has to be matched with our mission. Our motor performance is the basic factor which automatically and directly affects the flight in any cases, while taking off round, during cruise, carrying the bats and landing etc. That was why the power calculated 825,89 watt will be based on as a first motor selection criteria.

And the other topic of “how many batteries will be used” was the one of the important criteria. During calculations each battery has 1.1 V was chosen and with the basic formula of P=I.V how much volt energy we needed was calculated. On the other hand for the joints and cable there were some voltages more than our result which need to be added more. That was why %10 percent of the calculation voltage was added more into the previous one.

This was important, coz according to our mission and how much power the aircraft needs, gave an idea about choosing the motor to us. It has to have enough power but at the same time we have some limit for Battery weight. If we think about each battery has 22,95gr weight, here we have to make engineering decision with our optimizations.

So finally according to the result of we need to have a motor which has 825,89 watt energy and as we have already calculated the battery weight.

𝑊𝑏𝑎𝑡𝑡𝑒𝑟𝑦𝑊0

= (−0,003)x(Wo) + 0,1228 = (−0,003)x(5,97) + 0,1228

𝑊𝑏𝑎𝑡𝑡𝑒𝑟𝑦 = 0,6263 kg


46

We can say about this battery weight as allowable maximum battery weight. According to our calculation 21 batteries will be used for the motor. This calculation was made for all of these batteries’ weight.

After calculation and researching for which motor will be used then with the brand of “NEU 1506/3Y” was chosen to use for our aircraft. It has 1342 watt power and 0,6056kg weight as total. And battery type is “ELITE 1500 26-Cell Pack”. Propeller type is selected as “APC 18x12 E”.

Table 6.1 Motor and Propeller Features

Powerplant (Motor) NEU 1506/3Y Power (h.p.) 1342 watt Propeller APC 18x12 E

6.6 Fixed Engine Sizing Study The sizing procedure for the fixed-size engine is similar to the rubber engine sizing, but there are some exceptions. A parameter which is allowed to vary as the aircraft is sized is needed. By adding this parameter to the takeoff weight equation, fixed engine sizing can be made.

• Estimate Take off weight (Newton)

WO =NTPerEngine

(T/W)=

1.1342217,0632

= 6,18 kg

WO = (6,18)x(9,81) = 60,65 N

• Empty Weight (Newton)---(for W0=6,18kg)

We =WeW0

x W0 = (0,44) x (6,18) =2,7192kg

We = (2,7192)x(9,81) = 26,68 N

• Battery Weight (kg)

Wf = 21 x22,951000

= 0,48kg

Wf = (0,48)x(9,81) = 4,709 N

• Wing area ( m2 )

𝑆 = (0,0315). Wo + 0,2962

𝑆 = (0,0315). (6,18) + 0,2962 ---------------> 𝑆 = 0,49087 m2


47

• Payload Weight ( Wpayload )

Wpayload = 2,835 kg = 27,811 N

6.7 Geometry Sizing Study In the geometry sizing study part, firstly we should find our fuselage length. To find it,

we will use the higher W0 value which we have obtained in the engine sizing study. The formula for the fuselage length is :

𝐿𝑒𝑛𝑔𝑡ℎ = 𝑎.𝑊0𝐶

For W0 = 6,18 kg and for a homebuilt-composite airplane, a = 1,28 m and C = 0,23 . If we use those values in the formula :

𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒 = 𝑎.𝑊0𝐶 = 1,28 . 6,180,23 = 1,94597 𝑚

Then, we need to find the wing size. To find it, we will simply divide the weight by the takeoff wing loading.

𝑆𝑤𝑖𝑛𝑔 =𝑊0𝑊0𝑆

𝑆𝑤𝑖𝑛𝑔 =6,18

12,7545

𝑆𝑤𝑖𝑛𝑔 = 0,4845 𝑚2

To calculate tail volume coefficient, we will use the following formulas.

𝑐𝑉𝑇 =𝐿𝑉𝑇. 𝑆𝑉𝑇𝑏𝑤. 𝑆𝑤

𝑆𝑉𝑇 = 𝑐𝑉𝑇𝑏𝑤𝑆𝑤/𝐿𝑉𝑇

𝑐𝐻𝑇 =𝐿𝐻𝑇. 𝑆𝐻𝑇�̅�𝑤. 𝑆𝑤

𝑆𝐻𝑇 = 𝑐𝐻𝑇�̅�𝑤𝑆𝑤/𝐿𝐻𝑇

For a homebuilt airplane, cHT=0,5 and cVT=0,04 . �̅�𝑤 = 0,35 𝑚 , Swing=0,4845 m2 , and to find bw :

𝑏𝑊 =𝑆𝑤𝑖𝑛𝑔�̅�𝑤

=0,4845

0,35= 1,3843 𝑚


48

Our airplane has a front-mounted propeller engine, so LHT is 60% and LVT is 55% of the fuselage. So :

𝐿𝐻𝑇 = 𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒𝑥0,6 = 1,94597𝑥0,6 = 1,167582 𝑚

𝐿𝑉𝑇 = 𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒𝑥0,55 = 1,94597𝑥0,55 = 1,070284 𝑚

𝑆𝐻𝑇 =0,5𝑥0,35𝑥0,4845

1,167582= 0,07262 𝑚2

𝑆𝑉𝑇 =0,04𝑥1,3843𝑥0,4845

1,070284= 0,025066 𝑚2

6.8 Summary of Results

Table 6.2 : Results in SI unit system

Rubber Engine Sizing

Fixed Engine Sizing

Wo (N) 58,5657 60,6258

We (N) 25,768908 26,675352

Wp (N) 27,81135 27,81135

Wf (N) 6,144003 4,7088

Po (watt) 825,89 -

S (m2) 0,4723 0,49087

Table 6.3 : Results in British Engineering systems

Rubber Engine Sizing

Fixed Engine Sizing

Wo (lbf) 13,16609316 13,62922207

We (lbf) 5,793080992 5,996857711

Wp (lbf) 6,252240221 6,252240221

Wf (lbf) 1,381226826 1,058580355

Po (h.p) 1,107536736 -

S (ft2) 5,08379489 5,283680706


49

Table 6.4 : Geometry results Table 6.54 : Geometry results British

SI unit system Engineering systems

Fuselage_Length(m) 1,94597

Fuselage_Length(ft) 6,384416

Swing(m2) 0,4845

Swing(ft2) 5,215115

cHT 0,5

cHT 0,5

cVT 0,04

cVT 0,04

cw (m) 0,35

cw (ft) 1,148294

bw (m) 1,3843

bw (ft) 4,541667

LHT (m) 1,167582

LHT (ft) 3,83065

LVT (m) 1,070284

LVT (ft) 3,51143

SHT (m2) 0,07262

SHT (ft2) 0,781675

SVT (m2) 0,025066

SVT (ft2) 0,269808

6.9 Conclusion In this section, we performed a rubber engine sizing study. Then according to the data

we obtained, we selected a motor for our aircraft. We continued our process with a performing of fixed engine sizing study. If we look at the results in these to studies in table 1 and 2, we can see that the values are not absurdly different than each other, and according to the competitor study, we can see that the results are acceptable.

Finally we have performed a geometry sizing study so that we could see that dimensions of our airplanes components such as fuselage, wing and tail.

Now we have our motor model and dimensions of our airplane. As we can see, our airplane is started to have a numeric shape which is calculated in this study.

Chapter 7 Configuration Layout and Interior Design

50

7. CONFIGURATION LAYOUT AND INTERIOR DESIGN

7.1 Introduction The configuration layout and interior design of the aircraft are performed in this chapter. So many values of the parameters which are belonging to the aircraft have already been calculating until now. It will be needed to know some values of aircrafts to calculate the other unknowns related with wing, tail, fuselage. First of all it will be rearticulated the mean aerodynamic chord. Then it will be studied on the airfoils and twist distributions which were mentioned in the last studies. As a second step, the location of the wing and the tail will be decided on fuselage. This decision will be made by using the equations and relations in between the wing and tail parameters. As a third step 3 dimension drawings will be added to be able to consider more detailed way on the aircraft and to be able to have much more realistic results.

7.2 Wing and Tail Surfaces The wing area, the horizontal tail area and the vertical tail area are calculated on previous studies. Also, according to the calculations before which were made, the values of aspect ratio of the wing, the horizontal tail and vertical tail were found already. With the values of their own taper ratio the known parameters are tabulated below as:

Table 7.1: The calculated parameter before

λ AR S (m2) Wing 0,5 3,755 0,4845 Horizontal tail 0,7 3 0,07262 Vertical tail 0,7 3 0,025066

According to these values we will be calculating the span distance as first. Then with using the result of the previous step we will try to find the root chord and tip chord. Meanwhile we will have the all unknown parameters to calculate the value of the mean aerodynamic chord (m.a.c), location of mean aerodynamic chord. Respectively the following equations will be used to find the dimensions for the wing, the horizontal and vertical tail.

7.3 Locating Wing and Tail Surfaces Span 𝑏𝑊 = �𝐴𝑅𝑥𝑆𝑤𝑖𝑛𝑔

Root Chord 𝐶𝑟𝑜𝑜𝑡 =2𝑆𝑤𝑖𝑛𝑔

𝑏𝑤𝑖𝑛𝑔𝑥(1+𝜆)

Tip Chord 𝐶𝑡𝑖𝑝 = 𝜆𝑥𝐶𝑟𝑜𝑜𝑡


51

𝐶𝑤 =�23�𝑥𝐶𝑟𝑜𝑜𝑡𝑥(1+𝜆+𝜆

2)

(1+𝜆)

Location of m.a.c Y= 𝑏

6 1+2𝜆1+𝜆

• Wing Calculations

𝑆𝑤𝑖𝑛𝑔 =𝑊0𝑊0𝑆

=6,18

12,7545= 0,4845 𝑚2

𝑏𝑊 = �𝐴𝑅𝑥𝑆𝑤𝑖𝑛𝑔 = �3,755𝑥0,4845 = 1,3488 𝑚

𝐶𝑟𝑜𝑜𝑡 =2𝑆𝑤𝑖𝑛𝑔

𝑏𝑤𝑖𝑛𝑔𝑥(1 + 𝜆)=

2𝑥0,48451,3488𝑥(1 + 0,5)

= 0,4789 𝑚

𝐶𝑡𝑖𝑝 = 𝜆𝑥𝐶𝑟𝑜𝑜𝑡 = 0,5𝑥0,4789 = 0,2395 𝑚 = 𝟐𝟑,𝟗𝟓𝒄𝒎

• Tail Calculations

𝐿𝐻𝑇 = 𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒𝑥0,6 = 1,94597𝑥0,6 = 1,167582 𝑚

𝑆𝐻𝑇 = 0,07262 𝑚2

𝐿𝑉𝑇 = 𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒𝑥0,55 = 1,94597𝑥0,55 = 1,070284 𝑚

𝑆𝑉𝑇 = 0,025066 𝑚2

𝑏𝑉𝑇 = (𝐴𝑅𝑉𝑇𝑥𝑆𝑉𝑇)1/2 = (3𝑥0,024423)1/2 = 0,2707 𝑚

𝑏𝐻𝑇 = (𝐴𝑅𝐻𝑇𝑥𝑆𝐻𝑇)1/2 = (3,0𝑥0,077286)1/2 = 0,4815 𝑚

Figure 7.1 Location of the mean aerodynamic chord for main wing

Mean Aerodynamic center(m.a.c)

Mean Aerodynamic Chord = 299,375mm

674 mm


52

Figure 7.2 Location of the mean aerodynamic chord for vertical tail

Figure 7.3 Top view of the aircraft with all components

Figure 7.4 Side view of the aircraft with all components

145,9 mm

237,5 mm

M. A. C= 191,7mm

286 mm


53

Figure 7.5 Front View

Figure 7.6 Isometric View (Before loading the balls and bats )

Figure 7.7 Isometric View (After loading the balls and bats )


54

Figure 7.8 Transparent view (motor component)

7.4 Wetted Area and Volume Determination In this step we will try to find the wetted area first and then with using this result we will try to calculate the volume of our aircraft. First it will be estimated a value of exposed area for our aircraft and then find the approximate value of the wetted area. The wetted area directly depends on the exposed area and the equation which shows the relationship in between is;

Swet = Sexposed [ 1,977 + 0,52 𝑡𝑐

]

• Wing Wetted Area Sexposed will be guessed as the exact value of our area of the wing as 0,4845 m2. So;


] Swet = 0,4845 [ 1,977 + (0,52 )𝑥(0,13)]

Swet = 0,9906 m2

• Horizontal Wetted Area Sexposed will be guessed as the exact value of our area of the wing as 0,07262m2. So;


] Swet = 0,07262 [ 1,977 + (0,52 )𝑥(0,11)]

Swet = 0,1477 m2


55

• Vertical Wetted Area Sexposed will be guessed as the exact value of our area of the wing as 0,025066m2. So;


] Swet = 0,025066 [ 1,977 + (0,52 )𝑥(0,07)]

Swet = 0,05047 m2

7.5 Conclusion After calculating, this is very obvious that initial layout of the aircraft is extremely heuristic effect of the aircraft. Especially with the locations of the wing, the vertical tail and the horizontal tail are one of the most effective design parameters while designing an aircraft. And also we have seen that the locations of the vertical and the horizontal tail depend on their moment arm directly.

Chapter 8 Propulsion and Fuel System Integration

56

8. PROPULSION AND FUEL SYSTEM INTEGRATION

8.1 Introduction In this section of our aircraft design, we will be discussing about the propulsion system criteria. And we will be describing in detail the motor which is going to be selected. After that we will size the propeller and describe its properties. Then the propeller location will be selected and it will be discussed with our selection criteria. And then the installation of the motor will be decided with the reasons why it was chosen as like that. As a second step the battery compartment will be designed and the criterias for the selection and decision will be discussed. While making all these decisions and discussions it will be checked already the volume provided is sufficient or not. At the end the drawing figures will be added to the report and our last design shape for the aircraft will be showed with the configuration layout drawings.

8.2 Propulsion System Selection For the real aircrafts there are several choices to make the propulsion. Because there are several choices of motors according to the design criterias for the aircraft and the mission it has to be designed for. But our aircraft, being designed now, will be mini unmanned air vehicle, so this is why we will definitely have to choose the most effective mini motor to have the highest trust from it. And we already made a chose for the motor in our study 5. And according to our selections our motor was chosen as “NEU 1506/3Y”, our battery type was chosen as “ELITE 1500 26-Cell Pack” and the propeller type as “APC 18x12 E”.

• Battery type Battery type was chosen as “ELITE 1500 26-Cell Pack”. According to our calculations, to get the maximum trust with the minimum weight, this type of the batteries were the optimum ones. And in our aircraft 21 batteries will be used for creating the energy for our motor.

Figure 8.1: The picture of the battery “ELITE 1500 26-Cell Pack”


57

Table 8.1. The dimensions for the battery type

Diameter (mm) Height (mm)

ELITE 1500-26 Cell Pack 17 28,5

• Battery box As it was calculated in study 5, how many batteries will be used, to get the enough voltage for our motor, at the end 21 batteries will be used. However to hold the all batteries together, obviously a battery box has to be designed. And according to our aircrafts size and the minimum weight criteria considered, finally it has the dimensions as follows:

Table 8.2. The dimensions for the battery box designed

Height (mm) Width (mm) Depth (mm)

Battery Box 12 17,5 2,5

Basicly the design criterias for the battery box, were minimum weight, minimum volume inside of the aircraft, enough space between the batteries and the battery box’s walls, easy way to plug the batteries in and out. Actually to leave enough space in between the batteries and the battery box’s wall was the one of the important optimized criteria, because while the aircraft is working the batteries’ heat will be increasing gradually. Additionally for this situation, if the enough space will not be left it may effect the batteries performance in a bad way or damages the aircraft directly. Even it may cause a fire during flight. So with leaving some space it will be cooled by its own.


58

Figure 8.2: The batteries loaded in the battery box

The battery box will get the energy from the batteries and according to our design the energy will be carried with 2 cables, as anode and cathode, then it can send the energy to the motor.

Figure 8.3: After connecting the battery box and the motor

Then the battery box will be sticked to the payload compartment. While we were analyzing and searching for which part would be best to connect the battery type it was decided to stick the battery type into the payload compartment. As a first reason for this decision, it can be said that according to our design, the payload compartment is so near to the nose of the aircraft. And because of the fact of our motor has to be behind of the motor, the payload compartment is the best place that we stick the battery box on it. Otherwise we would stick this into the fuselage of our aircraft and that may affect the aerodynamical forces and create some drug also. And the other reason for choosing the payload compartment is, our payload compartment has holes to be able to put the balls inside. And there is a cover for that which we will put the bats on. That means payload compartment directly connects to the air. And the air passing around the payload compartment will cool the air inside as well. That means if we stick the battery box into the payload compartment, this cooled air inside of it, will also cool the air around the batteries, which are increasing the heat during the flight.


59

Figure 8.4: Sticked battery box into the payload compartment

• Cuff system To connect the motor to the fuselage, there will be need to design a new system which makes the motor tight to the aircraft. So a cuff system was chosen for this problem as the one in the front of the motor and the one in the back of the motor. It will have 6 clips as total. 3 clips from the front of the motor and the other 3 clips will be back of the motor. Because of the fact of two middle clips will be carrying the motor mostly, these have to be designed as much stronger which means much wider than the other 4 clips. And also according to the other fact of the weight of the motor will be heavier on back side of the motor, the clips on the back are also much wider than the one from front. The right and left clips will keep the motor stable especially while rolling during the flight.

Figure 8.5: The Cuff system on the motor zoomed view


60

Figure 8.6: The Cuff system on the motor from back

Figure 8.7: The Cuff system on the motor from front

Figure 8.8: The final model after making all connections from left view


61

Figure 8.9: The final model after making all connections transparent

• Propeller Our propeller was chosen as “APC 18x6”. Each propeller has 457,2mm length. And the distance in between its center and the top of the fuselage is 85mm. The width and height of the fuselage are 280mm, 250mm respectively. According to these dimensions which has the information about the location of the propeller tabulated as follows:

Table 8.3: Length for some size of the propeller

Length (mm)

Propeller 457,2

Width of the fuselage 280

Height of the fuselage 250

Center of the propeller to top of fuselage 85


62

Figure 8.10: Front view of propeller and size of it

Figure 8.11: Another view for a close look at the propeller and motor location

Figure 8.12: Design on the configuration layout drawing


63

8.3 Conclusion The location of the propeller and the design of it were decided according to the calculations and the choices from the previous studies. It was given information about the battery type and then designed a box for all batteries. Also it was mentioned why we need a battery box and why we sticked it to the place of its decided. Then the need of the cuff system was explained and the details about its own with strengths and width etc. As last all the compartments were assemblied all together and showed the details with the locations of each others.

Chapter 9 Landing Gear and Subsystems

64

9. LANDING GEAR AND SUBSYSTEMS

9.1 Introduction In this section of our aircraft design, we will be designing the landing gear. Also the criteria for selecting the landing gear arrangement and landing gear design will be discussed. In the same topic we will be estimating the maximum loads to be carried by the main and nose wheels then we will show the details of the landing gear system which was drawn by using 3-d computer aided drawing program. Then about the tire and its pressure the required tire size will be estimated based on load carrying and braking requirements. As a next step stroke will be determined and we will decide whether we will need an oleo and/or solid spring study depending on our system. Then the castoring wheel geometry will be designed and showed its details. We will also design and discuss about using retractable or non retractable landing gear with its reasons. And as last, we will discuss and make a decision on the aircraft subsystems which will be installed.

9.2 Landing Gear Arrangement In this section we will choose the type of our landing gear. There are several types of landing gear systems for real aircrafts, with considering its stability and payload. But because of the fact of we are designing and mini unmanned air vehicle we will not focusing of payload, duration, cruise in a detailed way. However there will be some other criteria appears at this point as the landing gear will not be fold up to the fuselage. Because if we try to design a system which will pull the landing gear system up and hide the landing gear inside of the fuselage, the take off weight of our aircraft will increase and also the manufacturing cost will get higher as well. So this is why as a first step this system which holds the landing gear will not be designed as explained above. After all these calculations and estimations for our aircraft, it is nearly done for the entire design of our aircraft. It will try to choose about which style of landing gear arran

Conceptual Design Study of an Unmanned Air Vehicle · 2012. 7. 29. · During design process,...

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Transcript of Conceptual Design Study of an Unmanned Air Vehicle · 2012. 7. 29. · During design process,...