aerochopper(vtol concept)

1

Aero-chopper

(VTOL) Supervisor: Mr. Nasser Chakra

by,

Shashank Dathatreya

2

Table of Content INTRODUCTION ............................................................................................................................... 4

UAV Types ................................................................................................................................... 6

Understanding the Project .............................................................................................................. 8

AIM .............................................................................................................................................. 8

ABSTRACT .................................................................................................................................... 8

SCOPE .......................................................................................................................................... 8

Parametric Study ........................................................................................................................... 10

Specifications and details .......................................................................................................... 12

Specifications and details .......................................................................................................... 14

GANTT CHART ................................................................................................................................ 17

DESIGN CONCEPT .......................................................................................................................... 19

COST ANALYSIS .............................................................................................................................. 21

MAN POWER ................................................................................................................................. 23

Cost analysis .................................................................................................................................. 24

Cost of materials and electricals ............................................................................................... 26

Materials ....................................................................................................................................... 27

Tools .......................................................................................................................................... 30

ELECTRICALS .............................................................................................................................. 32

Overview ................................................................................................................................... 33

Overview ................................................................................................................................... 35

Airfoil Selection ............................................................................................................................. 42

AIRCRAFT DESIGN .......................................................................................................................... 47

Structure designing :(PROFILI) ....................................................................................................... 49

3D Drawing .................................................................................................................................... 54

3

CONSTRUCTION (ASSEMBLY) ........................................................................................................ 58

GRAPHS ......................................................................................................................................... 74

Area Calculation ............................................................................................................................ 78

PERFORMANCE ANALYSIS ............................................................................................................. 97

CENTRE OF GRAVITY .................................................................................................................... 128

PERFORMANCE ANALYSIS ........................................................................................................... 132

Troubleshooting .......................................................................................................................... 135

Safety and Risk Assessment ........................................................................................................ 138

CONCLUSION ............................................................................................................................... 139

4

Acknowledgement

Of the many people who have been enormously helpful in the preparation of this

project, we are especially thankful to, Mr. Nasser Chakra for his help and

support in guiding us to through to its successful completion.

We would also like to extend our since gratitude to Emirates Aviation College for

the use of their resources, such as online databases and library, without which

the completion of this project would have been extremely difficult.

A very special recognition needs to be given to Ms. Kavita, our librarian, for her

extensive help and support during research and in dealing with online resources.

In addition, a special thanks to our friends Cibin, Suraj and Yogesh for their help,

consideration and guidance.

Last but not least, we would like to say a special thank you to our parents and

family members for their moral and financial support this semester.

5

INTRODUCTION1

Figure 1

UAV is an acronym for Unmanned Aerial Vehicle, which is an aircraft with no pilot on board.

UAVs can be remote controlled aircraft, for example, flown by a pilot at a ground control

station, or can fly autonomously based on pre-‐programmed flight plans or more complex

dynamic automation systems. UAVs are currently used for a number of missions, including

reconnaissance and attack roles. To distinguish UAVs from missiles, a UAV is defined as being

capable of controlled, sustained level flight and powered by a jet or reciprocating engine. In

addition, a cruise missile can be considered to be a UAV, but is treated separately on the basis

that the vehicle is the weapon. The acronym UAV has been expanded in some cases to UAVS

(Unmanned Aircraft Vehicle System). The FAA has adopted the acronym UAS (Unmanned

1. 1 http://www.theuav.com/

6

Aircraft System) to reflect the fact that these complex systems include ground stations and

other elements besides the actual air vehicles.

Officially, the term 'Unmanned Aerial Vehicle' was changed to 'Unmanned Aircraft System' to

reflect the fact that these complex systems include ground stations and other elements besides

the actual air vehicles. The term UAS, however, is not widely used as the term UAV has become

part of the modern lexicon.

UAV Types

• Target and decoy -‐ providing ground and aerial gunnery a target that simulates

an enemy aircraft or missile

• Reconnaissance -‐ providing battlefield intelligence

• Combat -‐ providing attack capability for high-‐risk missions (see Unmanned

Combat Air Vehicle)

• Research and development -‐ used to further develop UAV technologies to be

integrated into field deployed UAV aircraft

• Civil and Commercial UAVs -‐ UAVs specifically designed for civil and commercial

applications.

Degree of Autonomy

Some early UAVs are called drones because they are no more sophisticated than a simple radio

controlled aircraft being controlled by a human pilot (sometimes called the operator) at all

times. More sophisticated versions may have built-‐in control and/or guidance systems to

perform low level human pilot duties such as speed and flight path stabilization, and simple

prescript navigation functions such as waypoint following.

From this perspective, most early UAVs are not autonomous at all. In fact, the field of air vehicle

autonomy is a recently emerging field, whose economics is largely driven by the military to

develop battle ready technology for the war fighter. Compared to the manufacturing of UAV

flight hardware, the market for autonomy technology is fairly immature and undeveloped.

Because of this, autonomy has been and may continue to be the bottleneck for future UAV

7

developments, and the overall value and rate of expansion of the future UAV market could be

largely driven by advances to be made in the field of autonomy.

Autonomy technology that will become important to future UAV development falls under the

following categories:

• Sensor fusion: Combining information from different sensors for use on board

the vehicle

• Communications: Handling communication and coordination between multiple

agents in the presence of incomplete and imperfect information

• Motion planning (also called Path planning): Determining an optimal path for

vehicle to go while meeting certain objectives and constraints, such as obstacles

• Trajectory Generation: Determining an optimal control maneuver to take to

follow a given path or to go from one location to another

• Task Allocation and Scheduling: Determining the optimal distribution of tasks

amongst a group of agents, with time and equipment constraints

• Cooperative Tactics: Formulating an optimal sequence and spatial distribution of

activities between agents in order to maximize chance of success in any given

mission scenario

8

Understanding the Project

AIM

The Aim of this project is to design and construct an Unmanned Aerial Vehicle which will be a

hybrid between a helicopter and an airplane, so that we can achieve advantages of both

helicopter and airplane.

ABSTRACT

The purpose of this project is to design and construct a tilt-‐rotor aircraft with both a vertical

takeoff and landing. The aircraft being a hybrid of airplane and helicopter, which gives the

structure a superior performance and enhanced abilities having both the functions of a

helicopter and the aircraft, which include vertical take-‐off/landing and required forward speed.

The model aircraft can be constructed with balsa wood or any composite materials. The

airframe consists of the fuselage, which is the main component of the airplane, the wings(large

section of the aircraft), and the empennage (tail section, or tail feathers). The components of

the wings and tail sections are also known as the control surfaces since they are of course

important in controlling the airplane. The attached to the wings are flaps and ailerons. The

empennage is the tail assembly consisting of the horizontal stabilizers, the elevators, the vertical

stabilizer, and the rudder.

SCOPE

Scope of the project of constructing a UAV which will possess the capabilities of both helicopter

and airplane. Many reasons to this purpose , most important being because this branch of

aerospace industry has not fully been succeeded. Their success is limited to jet aircraft with

VTOL which use thrust vectoring and helicopters which use cyclic pitch and collective pitch to

hover.

9

The success of this model could be a breakthrough for larger scale models and eventually there

could be a new era of transportation where the private, military aircrafts could also implement

this concept an use shorter runway for take-‐off and cruise at a higher speed.

One of the main advantages of this type of aircraft is that if in case the engine fails, the aircraft

can glide and land as a normal aircraft since it has wings to create lift unlike helicopter, similarly

vice versa.

10

Parametric Study The aircrafts made with the similar concept is taken into this parametric study

RC TWIN VTOL PROTOTYPE

Figure 2

Specifications and details

Dimensions

Length -‐ 43 inches

Wingspan -‐ 48 inches

Center wing -‐ 29 inches

Motor spacing -‐ 19 inches

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Specification

Motors -‐ AXI 2212/26

Propellers -‐ MPI MAXX PRODUCTS counter rotating pair 10 x 4.5 slow flyer

ESC controller -‐ Castle Creations newer phoenix 25 Amp with 3 amp BEC

Batteries tested -‐ Polypus PQ-‐2100XP-‐3S 2100ma 20 C rated 167 grams or PD-‐B2600N-‐SP 3S

2600ma 12C rates 192 grams

External mixers -‐ 2 VEE-‐TAIL OMNI mixers

Aircraft Structure -‐ 1/4 inch balsa tail and fuselage, 2 @ 8mm diameter carbon fiber tubes

C of G -‐ on the tilt spar tube of maximum 1/4 inch front of the C of G -‐30% of wing chord

position WING

Static Thrust -‐ max 1300 grams

Weight -‐ 920-‐950 grams

12

V-‐22 OSPREY MODEL

Figure 32

Specifications and details

Dimension

- Length – 38.5 inches

- Span – 36 inches

- Center wing – 22 inches

- Weight – 1500 g

Power system – 2 Scorpion HK 2221-‐10 motors

Propeller – APC 12 x 3.8 slow flyer

Servos used

- 2 HITEC HS-‐5085MG (for tilting motors)

- 2 micro servos (for controlling movable surfaces )

2. 2 2 http://www.theuav.com/

13

Structure

- Primary – Balsa wood

- Secondary – Carbon rods and aluminum pipes

Electricals

- Receiver – Futaba R617FS

- Battery – Two EM2200 4S

- ESC – Two Phoenix ICE Lite 50SB

- Gyro – Three Futaba GY401

- Receiver power – CC regulator 20A Pro

Airfoil used – NACA 2413

Wing used – Straight wing

Empennage:

Horizontal and vertical stabilizer – conventional

Adhesion

- E-‐poxy 30 minutes

- E-‐poxy 5 minutes

- Hot glue

14

Specifications and details (AERO-CHOPPER)

Dimensions

- Length – 39 inches

- Span – 34 inches

- Center wing – 22 inches

- Approximate weight estimation – 2.5 to 3 kg

Power system – two Power electric motors

Propeller – APC 12 x 3.8 slow flyer

Servos used

- 2 high torque and high speed servo (for tilting the motors)

- 4 Micro servos ( for controlling movable surfaces )

Structure

- Primarily : Balsa wood

- Secondary : Carbon rods and aluminum pipes

Electricals

- Minimum 9 channel receiver and transmitter

- Minimum 3 gyros

- 2 external V-‐mixer

- Wire extensions

- Y-‐splitters

- Two 4cell battery packs

- 1 BEC

- 2 Electronic Speed Controllers – Minimum 60amps

Airfoil used – NACA 2414

Wing used – straight single high wing with uniform chord

15

Empennage:

Horizontal and vertical stabilizer – conventional

Engine mount – is tilted inwards by 2.3degrees

Adhesion

- Z-‐poxy 30 minutes

- Z-‐poxy 5 minutes

16

Mission Objectives

• Design and construct a hybrid aircraft of a helicopter and an airplane.

• Ensuring stable takeoff, land and transition from hover mode to forward mode.

• Ensure that the aircraft has an average endurance of a minimum 15 minutes in

hover mode or normal mode.

Outcomes

• Gathering information about How VTOL mechanism works.

• The type of wings and body constructed suitable to the VTOL concept

• Defining a set of parameters that we want the plane to conform to.

• Identify the materials and the budget required.

• Mathematical and aerodynamic calculations and maneuver calculation.

• Design the aircraft in a 2D & 3D sketch on AUTOCAD.

• Create an effective launch system in hover mode.

• Experiment the prototype model & troubleshoot safety & related issues.

• A Presentation of the aircraft.

Table 1

SPECIFICATIONS

Wing span (A) Span < 1m

Type of Wing Straight wing

Weight Weight < 2kg

Fuselage Length(a) Length < 1m

17

GANTT CHART

19

DESIGN CONCEPT The concept of Aero-‐chopper is very simple but involves sophisticated electrical and mechanics

for it to work.

The aircraft will be a twin engine and the engines will be on both the ends of the wing and will

be placed exactly on the C.G of the aircraft so that when the thrust is given, and if the aircraft is

balanced exactly on the C.G(motors), Aero-‐chopper should lift vertically.

Figure 4

The control of Aero-‐chopper on the different axis will be done by moving the engines and also

by powering up and down of the motors.

For the control of the pitch, Aero-‐chopper will tilt its wing anti-‐clockwise which would move the

direction of the propellers too. This will cause a change in the pitch of the aircraft.

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Figure 5

The Aero-‐chopper should also have the capability to tilt its engine forward about 45o to help it

transit from Hover mode to normal aircraft mode where its engine will be 0o (parallel to the

direction of flight)

Figure 6

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COST ANALYSIS Man hours analysis

Table 2

WBS Sheet: 1 Analysis in hours

Activity description Est to

complete

Est @

complete

Variance

GANTT CHART 2 3 1

Research 15 20 5

Aircraft Design 25 40 15

Design Approval 3 5 2

Parametric Design 4 4 0

Airfoil selection 2 2 0

3D design 15 20 5

Mission 1 1 0

Selection of aircraft parts 3 5 2

Cost Analysis 2 2 0

Tools and Materials(separate sheet) 10 16 6

Construction plan printing/Tracing 20 23 3

Cutting of material parts and organizing 4 4 0

Construction of aircraft structure accordingly 45 55 10

Assembly made rigid and Shaping 3 4 1

22

Electricals and Servos purchase(separate sheet) 5 5 0

Custom circuit made and tested 4 5 1

Fixing of Electricals and Servos 10 13 3

Aircraft performance test 3 3 0

Performance calculations 15 20 5

Centre of Gravity placement/calculations 2 2 0

Flight Test -‐ 1 3 3 0

Painting and finishing of aircraft structure 2 2 0


Final Calculations 2 4 2


Project report writing 10 15 5

Finalization of Aircraft 1 1 0

Deliverable 2 2 0

TOTAL 219 286 67

23

MAN POWER

Table 3

Days for the Project 90 days

Days devoted to the project

70 days

Average hours worked per day

4hours/day

Total hours for the days worked

45 x 4 = 180 hours

Average Man power = no. of persons/ hours

1/180

So a person has to work for 280 hours on this project.

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Cost analysis Table 4

WBS Sheet: 2 Analysis in costs

Activity description Est to

complete

Est @

complete

Variance

GANTT CHART 0 0 0

Research 0 30 30

Aircraft Design 50 70 20

Design Approval 50 65 15

Parametric Design 0 0 0

Airfoil selection 0 0 0

3D design 0 0 0

Mission 0 0 0

Selection of aircraft parts 0 0 0

Cost Analysis 0 0 0

Tools and Materials(separate sheet) 600 820 220

Construction plan printing/Tracing 50 85 35

Cutting of material parts and organizing 30 35 5

Construction of aircraft structure accordingly 50 60 10

Assembly made rigid and Shaping 30 30 0

25

Electricals and Servos purchase(separate sheet) 6000 7740 1740

Custom circuit made and tested 0 0 0

Fixing of Electricals and Servos 0 30 30

Aircraft performance test 0 0 0

Performance calculations 0 0 0

Centre of Gravity placement/calculations 0 0 0

Flight Test -‐ 1 200 200 0

Painting and finishing of aircraft structure 30 40 10

Flight Test -‐ 2 200 200 0

Final Calculations 0 0 0

Flight Test -‐ 3 200 200 0

Project report writing 0 0 0

Finalization of Aircraft 0 0 0

Deliverable 50 50 0

TOAL 7540 9655 2115

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Cost of materials and electricals Table 5

Items Quantity Cost per piece(aed) Total Amount (aed)

Materials

Balsa wood pack 1 500 500

Glues 5 35 175

Sand Paper 10 5 50

Cutter 3 15 45

Monocot 1 50 50

Electricals

Propellers 3 60 180

Electric Speed Control 3 480 1440

Battery 3 640 1920

Engine(Motor) 3 460 1380

Radio unit 1 1000 1000

Landing Gear unit 1 500 500

Servo pack 4 170 680

Servo(Tilt rotor) 2 290 580

Hinges pack 3 20 60

Total 43 4225 8560

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Materials3 Balsa wood

Figure 7

Balsa wood is the main material that we have used to construct the aircraft. Balsa wood is

lightweight, inexpensive and relatively strong. We have used it to construct the fuselage, wing

and tail-‐plane as well as in the sheeting of the plane.

Ply wood

Figure 8 We used ply wood on our model on the places where we need more strength like the root rips

of the wing, the front side cover of the fuselage, servo plates etc.

The materials that were mainly used were Balsa and Plywood

3. 3 http://www.moneysmith.net/Soaring/soaring4.html

28

Table 6

Component Material Thickness

flat fuselage sides, wing ribs, wing

spruces, main frame of fuselage,

servo holder, battery pack holder ,

frame and landing gear support area

etc.

B-‐Grain balsa wood 4 mm

Elevator ,horizontal stabilizer,

vertical stabilizer, aileron and rudder

C-‐grain balsa wood 3.2 mm

Covering rounded the fuselage,

planking fuselage and nose and wing

surface

A-‐grain 1.5 mm

To support some particular area like

inside the fuselage, Tilt roll of the

wing, and landing gear hold and

support area etc.

we used very small amount of ply

wood to make structure strong.

Plywood 4 mm

Thickness

▬ Mostly we used 4mm balsa for our main construction like wings , flat fuselage sides,

wing ribs, formers, trailing edges where more strength are required.

▬ We used 3.2 mm for body where it is not required to be very strong and it’s because to

reduce weight. we also used it for rudder, elevator, stabilizer, other attachments etc.

▬ In our project used 1.5mm where it is required for covering.

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E-poxy Glue

Figure 9

Epoxy is a strong, important modeling glue but one which must be used sparingly because of its

heavy weight.

Epoxy is classified by its strength and working time. Quick cure, or five minute epoxy, is strong

enough for most modeling applications, and is very handy for quick repairs. Slow cure (30

minute or more) epoxy is used when extra strength is required.

We have used epoxy to join the major parts of the airplane. This includes joining the wing

mounts to the fuselage, and attaching the tail to the fuselage. We have also used slow cure

epoxy for bonding the wood skins to the foam wing and stabilizer core.

Masking Tape

Figure 10

We used masking tape for minor repairs in the airplane. Masking tape was chosen due to its

convenient size, shape and ease of removal. It was mainly used for fixing small cracks in the

balsa wood.

30

Tools

Drill tools Figure 11

We used a small hand drill to drill holes in the balsa wood. A drill press was also used to make

sure that the holes were straight. Our hand drill was able to make holes of 2mm thickness.

Protractor Figure 12

We used a protractor to measure various angles in the model aircraft, which were needed in the

calculations. For example, we used it to measure the sweptback angle and the angle of the tail

planes.

Cutter

Figure 13

31

We used a normal cutter as it was very useful to cut the balsa wood, it easily cut through the

wood and was simple to handle. We sometimes used it to file the surface of the wood to make it

smooth and even.

Rulers

Figure 14

We used rulers for measuring the dimensions of the aircraft like wingspan, length of the

fuselage etc.

Sand paper Figure 15

Sandpaper is used to remove small quantities of material at a time from the surface of an

object. Sandpaper can be used to remove a specific material from an object (such as a layer of

paint) or to level and/or smooth the surface of the object. Sandpaper comes in many numbered

"grades," with smaller numbers being coarser and removing more surface material with each

pass. Higher numbers are finer and remove less material.

We have mostly used ‘low grade’ sandpaper for polishing and smoothing the aircraft. We have

also used it to shape the ribs and spars of the model aircraft.

32

ELECTRICALS4

MOTORS

Main wing motors(2 on the either sides of the wing)

Figure 16

Power 10 Brushless Out-‐runner Motor, 1100Kv

Key Features

• Equivalent to a 10-‐size glow engine for 32–48 ounce (910–1360 g) airplanes

• Ideal for 3D airplanes weighing 28–36 ounces (790–1020 g)

• Ideal for models requiring up to 450 watts of power

• High-‐torque, direct-‐drive alternative to in-‐runner brushless motors

• Includes mount, prop adapters and mounting hardware

• External rotor design—5mm shaft can easily be reversed for alternative motor

installations

• Slotted 14-‐pole out-‐runner design

4. www.e-‐fliterc.com/Products

33

• High-‐quality construction with ball bearings and hardened steel shaft

• Quiet, lightweight operation

Overview

The Power 10 is designed to deliver clean and quiet power for 10-‐size sport and scale airplanes

weighing 32 to 48 ounces (910 to 1360 grams), 3D airplanes weighing 28 to 36 ounces (790 to

1020 grams), or models requiring up to 375 watts of power. It’s an especially good match for the

E-‐flite Brio 10 for high speed F3A precision or artistic aerobatics.

Product Specifications

Type: Brushless out-‐runner motor

Size: 10-‐size

Bearings or Bushings: One 5 x 14 x 5mm Bearing, and One 5 x 11 x 5mm Bearing

Wire Gauge: 16

Recommended Prop Range: 10x5–12x6

Voltage: 7.2–12

RPM/Volt (Kv): 1100

Resistance (Ri): .043 ohms

Idle Current (Io): 2.10A @10V

Continuous Current: 32A

Maximum Burst Current: 42A (15 sec)

Cells: 6–10 Ni-‐MH/Ni-‐Cd or 2–3S Li-‐Po

Speed Control: 35–40A brushless

Weight: 122 g (4.3 oz)

34

Overall Diameter: 35mm (1.40 in)

Shaft Diameter: 5mm (.20 in)

Overall Length: 43mm (1.60 in)

Needed to Complete

E-‐flite Brio 10

40A ESC

6-‐ to 10-‐cell Ni-‐MH/Ni-‐Cd or 2–3S Li-‐Po

10x5 to 12x6 electric props

Tail Wing motor(1 at the rear)

Figure 17

Park 370 BL Outrunner,1200Kv with 4mm Hollow Shaft

Key Features

• Ideal for models requiring up to 120 watts of power

• Optimized windings for 3D performance

35

• High-‐torque, direct-‐drive alternative to in-‐runner brushless motors

• Includes mount, prop adapters and mounting hardware

• 4mm hollow shaft is easily reversed for alternative motor installations

• Excellent motor for small 3D airplanes 7–14 oz (200–400 g)

• Extremely lightweight—just 1.6 ounces

• Ideal for variable pitch props such as the E-‐flite® Showstopper Variable Pitch

Prop System

• External rotor design for better cooling

• High-‐quality construction with ball bearings

Overview

E-‐flite’s latest Park 370 is a brushless out-‐runner motor that features a 4mm hollow shaft, ideal

for use with variable pitch propellers. It’s perfectly designed for electric models equipped with

variable-‐pitch propeller systems, such as the E-‐flite® ShowStopper VPP system. However, you

don’t need a VPP to use this motor—it’s an excellent motor for small 3D airplanes that weigh 7–

14 ounces. A motor mount, prop adapter and all hardware are included.

Product Specifications

Type: Brushless out-‐runner

Size: Park 370

Bearings or Bushings: One 4 x 8 x 4mm Bearing, and One 4 x 9 x 4mm Bearing

Recommended Prop Range: 8x3.8–10x4.7 or Variable Pitch systems

Voltage: 7.2–12V

RPM/Volt (Kv): 1200

Resistance (Ri): .18 ohms

Idle Current (Io): .60A

36

Continuous Current: 10A

Maximum Burst Current: 12A (15 sec)

Cells: 6–10 Ni-‐MH/Ni-‐Cd or 2–3S Li-‐Po

Speed Control: 12–20A Brushless

Weight: 45 g (1.6 oz)

Overall Diameter: 28mm (1.10 in)

Shaft Diameter: 4mm (.16 in) hollow

Overall Length: 25mm (1.00 in)

Needed to Complete

12–20A brushless ESC

2–3S Li-‐Po or 6–10 Ni-‐Cd/Ni-‐MH

8x3.8–10x4.7 Slow Flyer Prop

Variable Pitch Prop option

37

ELECTRONIC SPEED CONTROLLER

ESC Eletronic Speed Control Detrum 30-‐40a 2-‐6s LIXX / 5-‐18s NC -‐ 0km

Model: E

Figure 18

• Model: ESC-‐40A

• Size(mm): 50 X 25 X 13

• Weight: 36g

• Current:40A

• NiCd/NiMh] /servos: 6/5 8/5 10/4 12/3

• [Li-‐xx]/servos: 2/5 3/4

38

BATTERY

Esky EK1-‐0186 20C 11.1v 1800mah Li-‐Polymer battery

Figure 19

Product Description

Table 7

Item NO. EK1-‐0186

Size 100*34*25mm

Weight(g) 47.0ï؟½ï 5.0½؟

(single electric core)

discharge magnification 20C

compages form connection in series

charging port XH2.5-‐4P reversal

(equilibrium charge)

Inner resistance 20mï؟½ï؟½ max


discharging cut-‐off voltage 2.75V


charging cut-‐off voltage 4.20ï؟½ï 0.05½؟ V


long-‐time load voltage 3.6V~4.1V


39

Radio

JR Propo DSX7 7-‐Channel 2.4GHz Computer Radio Control System (DSMJ), Package includes

Transmitter 2.4GHz DSMJ, RD731 7Ch 2.4G DSMJ Receiver w/EA131 Remote Receiver, ES539

Standard Servo x3, TX 8N 1500mah Ni-‐MH battery, Switch and 220V charger. English manual

included. | Mode 1, Mode 2 inter-‐changeable.

Figure 20

Product Code : [DSX7 2.4G DSMJ w/ES539 [DSX7JES539]]

Quality product from JR Propo.

JR Propo -‐ 2.4GHz Spread Spectrum Technology (DSMJ)

JR Propo DSX7 2.4GHz Computer Radio Control System (DSMJ) is suitable for Beginner to

Intermediate flyers and also the only model for even the advanced. It is reliable and stable with

2.4GHz with built-‐in system, promising an exciting flight in the comfort of all flyers.

It comes with Transmitter 2.4GHz, RD731 2.4GHz DSMJ 7 Channel Receiver w/EA131 remote

receiver, 3pcs x ES539 standard servos for electric model or glow model use.

The system comes with Mode 1 which can be changed to Mode 2 by editing system software

with stick spring.

The Flight Mode is at the right hand side.

40

Content

JR Propo DSX7 Transmitter 2.4GHz DSMJ

RD731 7Ch 2.4GHz DSMJ Receiver w/EA131 Remote Receiver

JR04884 2.4GHz Remote wire extension (150mm/6")

JR ES539 Standard Analog Servo x 3pcs (Servo Horns & mounting accessories included)

TX 8N 1500mah Ni-‐MH battery

NEC-‐322 220V Tx & Rx charger

Bind Plug Set

Switch

2mm Allen Wrench

English manual included | Mode 1 or Mode 2 inter-‐changeable.

Spec

-‐Method: DSMJ / Computer Mixing

-‐Number of Channels: 7ch

-‐Transmitter Weight: 640g (excluding battery)

-‐Battery fit: 8N1500

-‐For Helicopters or Airplane

Features

Band: 2.4 GHz

Servos: ES539 X 3

Receiver: RD731 (DSMJ)

Transmitter (Tx) Battery Type: 1500mah Ni-‐MH

AC: 220V

20-‐model memory

Airplane and Heli software

Switch assignment

P-‐mixes

41

3-‐axis dual rate and expo

3-‐position flap (Airplane)

5-‐point throttle & pitch curve (Heli)

3 flight modes plus hold (Heli)

Gyro programming (Heli)

CCPM swash mixing 90/120/180 degree (CCPM: Cyclic Collective Pitch Mixing System)

English manual

ES539 Standard Analog Servo Specification

Torque: 4.8kg.cm (66.67oz.in)

Speed: 0.23S/60°

Size: 32.5 x 19 x 38.5mm (1.28x0.75x1.52in)

Weight: 38g (1.34oz)

42

Airfoil Selection5 Airfoil used – NACA 2414

As this airfoil seems to be the most suited for this application according to the study shown

below

Comparing the airfoils; NACA 2412, NACA2414, NACA 2414

Naca-‐2412

Thickness: 12.0% Max CL angle: 15.0

Camber: 2.0% Max L/D: 50.702

Trailing edge

angle: 14.5o Max L/D angle:

5.5

Lower flatness: 45.2% Max L/D CL: 0.927

Leading edge

radius: 1.7% Stall angle:

7.0

Max CL: 1.204 Zero-‐lift angle: -‐2.0

5. 5 http://www.worldofkrauss.com/foils

Figure 21

43

Figure 22

NACA 2414

Figure 23

Thickness: 14.0%

Camber: 2.0%

Trailing edge

angle: 17.8o

Lower flatness: 50.5%

Leading edge

radius: 3.0%

Max CL: 1.245

Max CL angle: 10.5

Max L/D: 41.542

Max L/D angle: 6.0

Max L/D CL: 0.943

Stall angle: 10.5

Zero-‐lift angle: -‐2.0

44

NACA 2415

Figure 25

Thickness: 15.0%

Camber: 2.0%

Trailing edge angle: 19.1o

Lower flatness: 43.6%

Leading edge radius: 3.3%

Max CL: 1.281

Max CL angle: 11.5

Max L/D: 40.672

Max L/D angle: 6.5

Max L/D CL: 0.991

Stall angle: 11.5

Zero-‐lift angle: -‐2.0

45

Figure 26

From the above figures NACA 2414 is the most suitable.

NACA 2414 is selected since it has good enough thickness to accommodate 1 cm rod for the

engine tilting mechanism.

46

Ribs shapes generated with the help of the software "profili"

Figure 27

47

AIRCRAFT DESIGN The 2-‐D drawing that guided through the dimensions and construction process showing the 3-‐

isometric views of the aircraft.

Top view

Figure 28

48

Side view

Figure 29

Front view

Figure 30

49

Structure designing :(PROFILI) Wing structure with ribs placements designed with the help of Profilli

Figure 31

50

Figure 32

Figure 33

51

Rib structure Design on AutoCAD

Wing with ribs placement

52

Fuselage ribs and wing ribs placement

Spars supporting the ribs

Total structure of the wing

Fuselage ribs for support of the structure

54

3D Drawing Side view

Figure 34

Bottom view

Figure 35

55

Top view

Figure 36

56

Circuits

Tilt rotor circuit

Figure 37 Figure 38

Tilt rotor mechanism

Figure 39 Figure 40

57

Servo circuit

Figure 41

58

CONSTRUCTION (ASSEMBLY)

Figure 42

The Design(plan) gave us a green signal to finally start with the construction of the aircraft. The

component parts that were needed to form an assembled aircraft were each traced and draw

on the balsa wood with the respective dimensions using the carbon paper. These designs of the

parts were traced with the help of a transparent paper.

59

Figure 43

And then all the shapes were cut with the help of a normal metal cutter, and then placed

separately.

Figure 44

60

Starting with the wing, which had the following units:

• 8 airfoil shaped ribs each wing.

• 3 spars

• Tilt rotor holder ribs

Figure 45 Figure 46

Figure 47

61

Holes made with the help of a small drilling machine done by a professional. Holes made for the

space provision of the spars going through the airfoil parts.

Figure 48

Wing placed according to the design with the spars going through them making the entire inner

structure of the wing.

62

Fuselage

Figure 49

The Fuselage ribs cut accordingly and shaped as per the design.

Figure 50

Figure 51 Figure 52

63

Figure 53

Spaces at the sides of the fuselage ribs provided for placement of the support balsa sticks

making the fuselage structure rigid.

Figure 54

The fuselage ribs placed accordingly at correct distances as per the design.

Long balsa sticks glued to the spaces provided at the sides of the fuselage ribs.

64

Figure 55

Figure 56

The center part of the wing where it is placed on top of the fuselage structure, is constructed

accordingly for the holding of the tilt rotor mechanism parts.

65

Figure 57

Ply wood used for the support of the wing structure against the fuselage to give the area a

better rigidness and support, and a free movement in direction for the wing.

Figure 58

Landing gear support is constructed at the lower part of the fuselage on the either sides, so as to

give the landing gear a space away from the main fuselage structure.

66

Engine Mount

Figure 59

The engines placed on the either sides of the wing must be supported very strong as high stress

is faced in this area due to maximum throttle of the motor.

This area is mounted with balsa and ply wood together giving it a very good hold preventing

from breaking due to stress.

Figure 60

67

Tail wing

Figure 61

The tail-‐wing includes the horizontal stabilizer , the rudder and the tail motor mount.

Figure 62

Parts of the tail wing placed and fixed accordingly forming the internal structure of the

horizontal stabilizer and the rudder.

68

Figure 63

The total internal structure is constructed.

Figure 64

Figure 65

69

Tilt-Rotor mechanism structure

The tilt rotor section, constructed accordingly with the provision of the spar going through the

whole wing and the strong support for the tilt rotor mechanism structures.

Figure 66 Figure 67

Figure 68 Figure 69

70

Figure 70

Wing at the tilt position for the hovering part of flight

Electricals and servos fixed at the appropriate locations

Figure 71 Figure 72

71

Tail-‐motor fixed with the mount supporting it and giving the propeller blades a clearance

distance from the tail wing.

Figure 73

Figure 74 Figure 75

Landing gear attached, one on either sides and one at the tail-‐part of the fuselage

Figure 76

72

Battery holder is made by creating a space exactly measured for the battery to fit it.

Figure 77

Motors fixed to the mounts on the either side of the wings.

Figure 78

73

Tilt-‐Rotor Aircraft sheeting, shaped and painted.

Figure 79

Aero-‐chopper presented with the Tilt-‐rotor function.

Figure 80

74

GRAPHS6 The graphs that we are going to use are the following

The aerodynamic form factor graph

Figure 81

6. 6 Fundamentals of Flight by Richard S Shovel

75

Figure 82

Figure 837

76

Figure 84

Figure 85

77

Table 8

78

Area Calculation CALCULATIONS:

Wing

Drawing 1

Drawing 2

Rectangle

Area = l x b

= 35 x 18 = 648cm2

RIGHT + LEFT = 648 + 648 = 1296cm2

79

Drawing 3

Rectangle

Area = l x b

= 13.5 x 9.1 = 122.85cm2

Total wing area(TOP) = Left section + Right section + Center section

= 648 + 648 + 122.85

= 1418.85cm2

Total wing area(BOTTOM) = Total wing area(TOP)

Total Wing Area(TOP and BOTTOM) = TOP + BOTTOM

= 1418.85cm2 + 1418.85cm2

= 2837.7cm2

Airfoil shaped side section of wing

With the help of AutoCAD the exact area of the side section of the wing could be taken by

calculating the area of the airfoil.

80

Drawing 5

Area = 4.1455inch2 = 26.745cm2

2 sides = 26.745cm2 x 2 = 53.49cm2

Area of the sides view of the wing = 53.49cm2

Total surface Area of the Wing = Side-‐view Area + Top & Bottom view Area

= 53.49cm2 + 2837.7cm2

= 2891.19cm2

Total Surface Area of the Wing(MAIN) = 2891.19cm2

Drawing 4

81

Horizontal Stabilizer

Drawing 6

Rectangle

Area = l x b = 7.5 x 12.2 = 91.5cm2

Triangle

Area = 1/2 b h = 1/2 x 4.5 x 12.2 = 27.45cm2

Total = Rectangle + Triangle = 27.45 + 91.5 = 118.95cm2

2 sides = 118.95 + 118.95 = 237.9cm2

82

Vertical Stabilizer

Drawing 7

Rectangle

Area = l x b = 52 x 10 = 520cm2

Top and bottom = 520 x 2 = 1040cm2

Total Area = 1040cm2

Engine Mount

Drawing 8

l = 11.7cm

83

w = 5.1cm

h = 5.1cm

Area of cuboid = 2 ( lw + wh + hl )

= 2 ( 11.7x5.1 + 5.1x5.1 + 5.1x11.6 )

= 2 (145.35) = 290.7cm2

Total Area of Engine Mount(2 sides) = 290.7cm2 x 2 = 581.14cm2

Engine Mount(TAIL)

Drawing 9

l = 11.5cm , w = 3.8cm , h = 2cm

Area of cuboid = 2 ( lw + wh + hl )

= 2 ( 11.5x3.8 + 3.8x2 + 11.5x2 )

= 2( 74.3 ) = 148.6cm2

Total Area of Engine Mount(TAIL) 148.6cm2

84

Fins

Drawing 10

Triangle

Area = 1/2 b h = 1/2 x 4.9 x 6.8 = 1/2 x 33.32 = 16.66cm2

Rectangle(R1)

Area = l x b = 6.8 x 5 = 34cm2

Rectangle(R2)

Area = l x b = 9.9 x 2 =19.8cm2

Total Area(one side-‐one fin)=Triangle + R1 + R2 = 16.66 + 34 + 19.8 = 70.46cm2

2 Sides = 70.46 x 2 = 140.92cm2 ; 2 Fins = 140.92 x 2 = 281.84cm2

Total Area of the fins = 281.84cm2

85

Landing Gear Hold

Drawing 11

Airfoil area -‐ 2.7910cm2 = Side surface

53.49 -‐ 2.7910 = 50.699

Area = 50.699cm2

Forward section

Drawing 12

Area = l x b = 7.5 x 2 = 15cm2

86

Top and Bottom surface

Drawing 13

Area = l x b = 7.5 x 16 = 120cm2

Top and bottom = 120 x 2 = 240cm2

Total Landing gear hold Area = Side + Forward + Top and Bottom

= 50.699 + 240 +15

= 305.699cm2

87

Area of Fuselage

Drawing 14

Area of the fuselage side section

= Area A + Area B +Area C + Area D + Area E + Area F

Area A

(Trapezium)

Drawing 15

Area of Trapezium = (a + b)/2 x h

= (5 + 6.3)/2 x 2.9 = 16.385cm2

88

Area B

Drawing 16

Area of Trapezium = (a + b)/2 x h

= (6.3 + 8.5)/2 x 4 = 29.6cm2

Area C

Drawing 17

Rectangle

Area = l x b = 56.95cm2

Triangle

Area = 1/2 x b x h = 1/2 x 4.6 x 6.7 = 15.41cm2

Total Area = 56.95 + 15.41 = 72.36cm2

89

D

Drawing 19

Can be assumed as:

Area = l x b = 13.4 x 16.5 =

221.1cm2

E

Drawing 20

Drawing 18

90

Rectangle

Area = l x b = 7.8 x 23.2 = 180.96cm2

Triangle

Area = 1/2 x b x h = 1/2 x 5.9 x 23.2 = 249.4cm2

F

Drawing 21

Area = 1/2 x b x h = 1/2 x 7.8 x 24.3 = 94.77cm2

Total Area of Fuselage Side section = A + B + C + D + E + F

= 16.385 29.6 + 72.36 + 221.1 + 249.4 + 94.77 = 683.615cm2

Both sides = 683.615 + 683.615

= 1,367.23cm2

91

Fuselage(TOP AND BOTTOM)

Drawing 22

92

A'

Drawing 23

Area of trapezium

= ( a + b )/2 x h = ( 6.2 + 12.1 )/2 x 6

= 225.06cm2

B'

Drawing 24

Area of Trapezium

= ( a + b )/2 x h = ( 12.1 + 13 ) / 2 x 7.6

= 95.38cm2

93

C'

Drawing 25

Area = l x b = 13 x 10

= 130cm2

D'

Drawing 26

Area = l x b = 8.2 x 5.5 = 45.1cm2

94

E'

Drawing 27

Area = l x b = 13 x 29.4 = 382.2cm2

F'

Drawing 28

Area = 1/2 x b x h

= 1/2 x 13 x 24 = 156cm2

95

Total Area of Fuselage(TOP)

Area A' + B' + C' + D' + E' + F' = Total Area

225.06 95.38 + 130 + 45.1 + 382.2 + 156 = 1,033.74cm2

TOP and BOTTOM = 1033.74 x 2 2067.48cm2

Front surface

Drawing 29

Area = l x b = 6.2 x 5 = 31cm2

Total Area of Fuselage

SIDES + TOP/BOTTOM + FRONT

= 1367.23 + 2067.48

= 3465.71cm2

Total Area of Fuselage = 3465.71cm2

Total Area of Wing = 2891.19cm2

Total Area of Tailing = 1040cm2

Total Area of Horizontal Stabilizer = 237.9cm2

96

Total Area of Fins = 281.84cm2

Total Area of Engine Mounts(MAIN) = 581.14cm2

Total Area of Engine mount(TAIL) = 148.6cm2

Total Area of Landing Gear Hold = 305.699cm2

TOTAL SURFACE AREA OF THE AIRCRAFT =

Fuselage + Wing + Tailing + Horizontal Stabilizer + Fins + Engine Mounts(MAIN) + Engine

mount(TAIL) + Landing Gear Hold

= 3465.71cm2 + 2891.19cm2 + 1040cm2 + 237.9cm2 + 281.84cm2 + 581.14cm2 + 148.6cm2 +

305.699cm2

Total Surface Area of the Aircraft = 8952.079cm2

97

PERFORMANCE ANALYSIS LIFT

Airfoil used is NACA 2414

The software profili gives us the following values;

CLmax = 1.379 at 15o AOA

CL = 0.752 at 4o AOA

During cruise

Considering the angle of attack of wing, during cruising will be 40.

We know that L = W during cruise.

L = 2 KG

L = 20 N

Ρ(density) at sea level = 1.225 kg/ m3

S(wing area) = 0.05898m2

L = 1/2 P CL V2 S Eq (1)

V2 = 2L / ρ CL S = 2 * 20 / 1.225 * 0.752 * 0.05898

V2 = 737.2

V = 27.15m/s

98

During Landing

The Cl is at max = 1.379

V2 = 2L / ρ CL S = 2 * 20 / 1.225 * 1.379 * 0.0589

V2 = 409.50

V = 20.2 m/s

99

Vstall is the lowest speed at which steady controllable flight can be maintained any further

increase in AOA will cause flow separation on the wing upper surface, a drop in lift, a large

increase in drag. In a well-‐designed airplane, a strong pitch-‐down moment is experienced.

Vstall = Vat landing

Vstall = 20.2 m/s

LIFT during landing

L = 1/2 P CL V2 S Eq (1)

= 1/2 *1.225* 1.379*20.22 * 0.05898

L = 20.2 N

During take off

Velocity at take-‐off is 20% greater than Vstall.

VTO = 20.2 + 20.2 x 0.20

= 24.24 m/s

100

OR

Around 70% of CLMAX

i.e. 0.70x1.379 = 0.966

VTO2 = 2L / ρ CL S = 2 * 20 / 1.225 * 0.966 * 0.05898

i.e. At 60 AOA VTO = 23.9 m/s

LIFT during Take-off

LTO = 1/2 P CL V2 S Eq (1) [since CL = 0.966

= 1/2 *1.225* 0.966*20.22 * 0.05898 and VTO = 21.8]

LTO = 14.43N

DRAG

The total Drag of the Aircraft is calculated by summing the parasite and induced drag together.

101

CD = CDP + CDi

D = CDqS

Drag is calculated for three phase of flight i.e. Take off, Cruise and landing.

At Take Off

CDptotal is calculated by computing CDp for wing, fuselage , horizontal and vertical stabilizer

separately.

CDp of wing

∑ 𝐾𝐶fswet / Sref Eq (4)

Sref = 0.028912m3

Swet = 0.05898 m3

Cr = 0.18 m

CT = 0.18 m

σ = CT / Cr = 1

L = MAC Eq (5)

= 2/3 x Cr (1 + σ -‐ σ/(1 + σ) )

= 2/3 * 0.18 ( 1 + 1 -‐ 1/(1 + 1)) = 2/3 *0.18(2 -‐ 1/2)

= 0.18

L = 0.18

102

V = 1.4607 * 10-5

RN = V0 x L/v Eq (6)

= 24.24 * 0.18 / 1.4607 * 10-‐5

RN = 298,706.0998

Since RN > 200,000

The flow is turbulent

Cf for turbulent

Cf = 0.455 / [(log10RN)2.58 = 0.455 / [log10298,706.0998) 2.58

= 5.6 x 10-‐3 Eq (7)

103

From fig 9c.1 we get K = 1.265

Hence CDp = ∑ 𝐾𝐶fswet / Sref Eq (4)

= 1.265 x 5.6 x 10-‐3 x 0.05898 / 0.028912

CDpwing = 0.01443

104

CDp of fuselage

Fuselage Area = 0.3465 m2

Fuselage length = 1.016 m

RN = ( Vo * L)/v Eq (6)

= (24.24 * 1.016)/ 1.4607 x 10-‐5 = 1,686,029.986

RN = 1,686,029.986 > 200,000 hence the flow is turbulent

Cf = 0.455 / [(log10RN)2.58 Eq (7)

Cf = 4.062 x 10-‐3

CDpfuselage = ∑ 𝐾𝐶fswet / Sref

K from fig 9c.2

K = 1.16

105

CDpfuselage = 9.599 x 10-‐3

= 0.00959

CDp of Horizontal Stabilizer

H.S of Area = 0.02379m2

106

Cr = 0.12 m

Ct = 0.075 m

σ = CT / Cr = 0.625

t/c = =

Swet = 0.0485

Sref = 0.02379

L = MAC

= 2/3 x Cr (1 + σ -‐ σ/(1 + σ) )

= 2/3 x 0.12 ( 1 + 0.625 -‐ 0.625/(1+0.625) ) = 2/3 x 0.12 ( 1.625 -‐ 0.3846)

L = 0.0992

RN = V0 x L / v

= 24.24 x 0.0992 / 1.4607 x 10-‐5 =

as 164,620.25 > 200,000

Hence flow is laminar

Cf = 1.328 / √RN Eq (7)

= 3.27 x 10-‐3

K from fig 9c.1

K = 1.155

CDp = ∑ 𝐾𝐶fswet / Sref

= 1.155 x 3.27 x 10-‐3 x 0.0485/0.02379

107

CDp H.S = 0.00767

108

CDp of Vertical Stabilizer

Area of Vertical Stabilizer = 0.01040

Cr = 0.1 m

Ct = 0.1 m

σ = CT / Cr = 1 m

Swet = 0.021216

Sref = 0.0104

t/c = =

L = MAC

= 2/3 x Cr (1 + σ -‐ σ/(1 + σ) )

= 2/3 x 0.1 (1 + 1 -‐ 1/(1 + 1) = 2/3 x 0.1 (2 -‐ 1/2)

L = 0.1 m

RN = V0 x L / v Eq (6)

= 24.24x 0.1 / 1.460 x 10-‐5

= 166,027.397 < 200,000

Cf for Laminar Boundary layer

Cf = 1.328 / √RN Eq(8)

= 1.328 / √166,027.397

= 3.25 x 10-‐3

K from fig 9c.1

K = 1.185

109

CDp = ∑ 𝐾𝐶fswet / Sref Eq(4)

= 1.185 x 3.25 x 10-‐3 x 0.021216/0.0104

CDp V.S = 7.85 x 10-‐3

110

Total CDp for the Aero-‐chopper during Take-‐Off

111

CDptotal = CDpwing + CDpfuse + CDp H.S + CDp V.S

= 0.01443 + 0.0095 + 0.00767 + 0.00785

CDptotal = 0.03945

Induced Drag CDi

CDi = CL2/πARe

AR = b2 / s = 0.862 m / 0.028912 m2

= 2.5581

Cl at Take-‐Off = 0.966

(constant) U = 0.99

(constant) S = 0.975

K = 0.38 x CDptotal = 0.38 x 0.03945 = 0.014991

e = 1 / (πARK) + 1 / u + s

= 1 / (π x 2.558 x 0.01499 ) + 1/(0.99 + 0.975)

112

e = 8.8135

CDi = (0.966)2 / (π x 2.55 x 0.933)

= 0.124

The total efficient of drag is

CDptotal = CDp + CDi

= 0.03945 + 0.124

CDptotal = 0.1642

114

DRAG total

D = CDPV2S / 2

= (0.1642 x 1.225 x 24.242 x 0.0589) / 2

D = 3.48N

L/D Ratio during Take-‐off = 14.43 / 3.48 N

= 4.14 N

Wing loading = W / S = 2 / 0.028912 = 69.17 kg/m2

Thrust to Weight ratio = T / W = 4 / 2 = 2

116

AIRCRAFT MANEUVERS

Turning Performance

V = 27.1 m/s

θ = 60o

g = 10 m/s 2

n = 1/Cosθ = 1/Cos60 = 2

Level Turn

Radius of turn (r)

r = V2/g√n2 – 1) = 27.12/10√(22 -‐1) = 27.12/10√3 = 42.4m

Angular Velocity (ω)

ω= g√n 2 – 1/v = 10 * √3/27.1 = 0.63 m/s

117

Vertical Turn Pull up

Radius of turn (r)

r = r 2/g(n-‐1) = 27.1 2/10(2-‐1) = 73 m

Angular velocity

ω= g(n-‐1)/v = 10(2-‐1)/27.1 = 0.36 m/s

VERTICAL TURN PULL DOWN

Radius of turn (r)

r= v 2/g(n+1) = 27.12 /10(2+1) = 24m

Angular velocity

ω = g(n +1)/v = 10* 3/27.1 = 1.1 m/sθ

Vertical turn

Radius of turn r

r = v2/gn = 27.12/10*2 = 36 m

Angular velocity (ω)

ω= gn/v = 10 * 2/27.1 = 0.73 m/s

118

HELICOPTER PERFORMANCE

Hover

Figure 86

Disc Area = A

A = π * 0.1552

= π * 0.0240

A = 0.075 m2

Density at sea level ρ = 1.2550 kg/m3

Assuming V∞ = 7 m/s

Upstream or Downstream Velocity = V∞

Induced Velocity = Vi

∆P = 1* ρ * V∞* V∞ / 2

∆P = 1x1.2250x7x7 / 2

119

= 60/2

∆P = 30 kg/m3

Thrust required to Hover

∆P = Thrust / Area

Thrust = ∆P * Area

= 30 x 0.075

Thrust = 2.25 kg/s

Since Aero-‐chopper has two motors on the either side of the main wing the total thrust is 2.25

+ 2.25 = 4.5kg/s

Velocity required to Hover

Vi = √(T/2*ρ*A)

120

= √(W / 2 * ρ where, W( Disc Loading ) = T / A

= √ 30/ (2x1.225) = 4.5 /( 2 x 0.075 )

= 30kg/m3

Vi = 3.49 m/s

121

Power required to Hover

Pi = T * Vi

= 2.25 x 3.49

P = 7.87kgm/s2

The total power required by both the motors is P * 2 = 7.87 x 2

Ptotal = 15.74kgm/s2

122

Vertical Climb

Figure 87

Rate of climb = VC

Assuming the VC = 1m/s

Vh2 = (VC + Vi) Vi

( 3.49 )2 = (1 + 0.5V) x 0.5V

12.18 = 0.5V + 0.25V2

0.25V2 + 0.5V -‐ 12.18 = 0

V = 6.05 m/s

Vi = 0.50 *V

Vi = 0.5 x 6.05 = 3.025

= 3.025 m/s

Total Velocity for Climbing = ( VC + Vi )

= ( 1 + 3.025 )

Total Velocity for Climbing = 4.025m/s

123

Thrust required for vertical climb

T = 2 ρA(VC + Vi) Vi

= 2 x 1.2550 x 0.075 x 4.025 x 3.025

T = 2.29kg/s

Total thrust required is T*2 = 4.58kgm/s2

Power required for Vertical climb

Pi = T(VC + Vi)

Pi = 2.29 ( 4.025 )

Pi = 9.21kgm/s2

Total power required is

Pi = T(VC + Vi)

Pi = 4.58 ( 4.025 )

Pi = 18.43kgm/s2

As we know the relationship between induced velocity in hover and vertical climb, the induced

velocity decreases as the climbing speed and similarly induced power increases with the

increase in climb speed.

124

Comparing the Hover and the Vertical Climb, we get.

Vertical Descend

With a flow pattern as in fig simple momentum theory gives a reasonable approximation: thus

with VC negative and Vi positive the thrust is

T = -‐2 ρA(VC + Vi) Vi

Assuming the Descend rate to be VC = -‐1m/s

T/ -‐2 ρA = (VC + Vi) Vi We know, thrust = 4.5kg

4.5 / -‐2 x 1.255 x 2 x 0.075 = (VC + Vi) Vi

-‐11.9 = (-‐1 + Vi) Vi

-‐11.9 = -‐ Vi + Vi2

Vi2 -‐ Vi -‐ 11.9 = 0

Vi = -‐2.98 m/s

And hence V∞ = -‐2 x V1

V∞ = 5.96m/s

125

Maneuvers in Helicopter Mode8

Aero-‐chopper , has 4 basic Maneuver's or has to be controlled on all the 3 axis of stability and

also on the attitude or vertical ascend;

Figure 88

• Increasing and Decreasing altitude

This is achieved by the increasing or decreasing the throttle and thereby changing the RPM

and thrust produced by the two motors which are placed on the wing tips. For Vertical

takeoff the max thrust produced by both the engines are 4 kg i.e. N whose vector

component are Vertical.

• For controlling the pitch i.e. longitudinal axis.

Since the motors are placed on the C.G. line when both the motors are tilted front or back

this aircraft is effected on its pitch axis thereby causing the aircraft to pitch up or down.

7. 8 Basic Helicopter Aerodynamics by J.Seddon

126

Figure 89

Endurance calculation

For instance an RC LiPo battery that is rated at 2000mAh would be completely discharged in one

hour with a 2000 milliamp load placed on it.

Since the motor POWER10 has been chosen it requires Electronic speed controller at

"40"amps with power supply from a battery of 1800mAh

1800mAh/ 4000 ma x 60 = 2.7 mins

= 2.7 mins

When the motor is on full throttle for 27 min the battery will drain. But this calculation is not

for completely draining the battery. The battery actually drains out after 1.5 times the

calculated time. This margin is a safety margin for avoiding the aircraft to crash because of loss

of power.

So the actual Endurance for hover is 4.05 mins

Now for normal flight the endurance is

1800mAh/ 4000 ma x 60 x 1.5 = 4.05 mins

= 4.05 mins

2.7mins of safe flight and the actual endurance for normal flying is 4.05mins

127

Range Calculation

Range is calculated by using a very simple calculation.

Range of Aero-‐chopper when it is in helicopter mode

Range = Endurance * Cruise velocity

Range = 243 sec * 27.15 m/s

= 6002.1 m

Range of Aero-‐chopper when it is in aircraft mode

Range = Endurance * Cruise velocity

Range = 342 sec * 27.15 m/s

= 8447.4 m

128

CENTRE OF GRAVITY LONGITUDINAL AXIS -‐ DATUM point is the Tip of the nose

Components Weight (gm) Distance (cm) Moment

(gm*cm)

Right ESC 60 34 2040

Left ESC 60 34 2040

Wing tilt servo 64 28 2432

Left wing aileron servo 23 35 805

Right wing aileron servo 23 35 805

Left Tailing aileron servo 20 68 1360

Right Tailing aileron servo 20 68 1360

Receiver + Electricals 60 27 1620

Nose 31 7 217

Vertical Stabilizer 15 76 1140

Rudder 15 78 1170

Left wing aileron 10 45 450

Right wing aileron 10 45 450

Fuselage 342 45 15,390

Wing 288 32 9,216

Main Landing gear(Right) 30 25 750

Main Landing gear(Left) 30 25 750

Battery pack 480 14 6,720

Left motor with mount 122 + 40 33 5,346

Right motor with mount 122 + 40 33 5,346

Tail motor with mount 50 + 40 100 9000

Monocot 5 0.1 0.5

Total 2011 68,407.5

Longitudinal Axis C.G = Total Moments / Total weight

= 68407.5 / 2011 = 34cm from the Datum point

129

LATERAL AXIS -‐ DATUM point is the Tip of the Right Engine


(gm*cm)


Left ESC 60 51 3060



Right wing aileron servo 23 19.5 448.5




Nose 31 29 899


Rudder 15 74 1110



Fuselage 342 37 12,654

Wing 288 33.5 9,648

Main Landing gear(Right) 30 20 600

Main Landing gear(Left) 30 45 1350

Battery pack 480 28 13,440

Left motor with mount 122 + 40 57 9,234

Right motor with mount 122 + 40 12 1944


Monocot 5 0 0

Total 2011 73,832.5

Lateral Axis C.G = Total Moments / Total weight

= 73832.5 / 2011 = 36.7cm from the Datum point

130

DIRECTIONAL AXIS -‐ DATUM point is the mid-‐point of the Landing Gears separation


(gm*cm)


Left ESC 60 34 2040



Right wing aileron servo 23 35 805




Nose 31 21 651


Rudder 15 57 855



Fuselage 342 26 8892

Wing 288 14 4032

Main Landing gear(Right) 30 16.5 495

Main Landing gear(Left) 30 16.5 495

Battery pack 480 10 4800

Left motor with mount 122 + 40 29 4698

Right motor with mount 122 + 40 29 4698


Monocot 5 0 0

Total 2011 47,260

Lateral Axis C.G = Total Moments / Total weight

= 47260 / 2011 = 23.5cm from the Datum point

131

The Centre of Gravity of Aero-‐chopper falls on

34cm from the tip of the nose, 36.7 cm from the tip of the right motor

mount and 23.5 cm from the tip of the landing gear.

132

PERFORMANCE ANALYSIS

Formulae used throughout the report

Lift and drag calculation

• Lift (L) = ½ ρV2SCL

• D = Cads

• CD = CDP + CDi

• CDP = (Cf * k * Swet)/Sref

• MAC = 2/3 * CR(1 + σ -‐ σ/1+σ)

• RN = ( Vo * L)/vv

• CF = 0.455/(Log RN)2.54

• CF = 1.328/√RN

• CDi = CLcruise2/πARe

• e = 1/[πARK + 1/(u * s)]

Climb performance

• (R/C) = TV∞ -‐ DV∞ / W

Take-off Performance

• Sa = RsinθOB

• Sg = 1 / 2gKA ln(1 + KA/KT x V2LO ) + NVLO

• R = 6.9(Vstall)2 / g

• θOB = cos-‐1 (1-‐ hOB / R)

• SƟ = 1/2gK

• KT = (T/W) -‐ µT

• G = (16h/b)2 / 1 + (16h/b)2

133

• VLO = 1.1 * Vstall

• CD,O ~ ∆ CD,O = w/s * Kucm-‐0.215

• KA = ρ∞/2(w/s) [CD,O + ∆ CD,O + ( K1 + G/ πeAR) CL2 = µrCL ]

Landing Performance

• Vf = 1.15Vstall

• R = Vf2 / 0.2g

• hf = R(1 -‐ cosθo )

• So = S-‐ hf / tanθa

• Sa = 50 -‐ hf / tanθa

• Sf = R sin θa

• JT = Trev / W + µr

• JA = ρ∞ / 2(w/s) [CD,O + ∆ CD,O + ( K1 + G/ πeAR) CL2 = µrCL ]

• VTD = 1.15Vstall

Aircraft Maneuvers

• r = V2 / g * tan ϕ = V2 / g√(n2 -‐ 1)

• ω = V / r = V / V2 / g√(n2 - 1)

• ω = g√(n2 -‐ 1)

• r = V2 / g (n-‐1)

• ω = g (n-‐1) / V

• r = V2 / (n+1) g

• ω = g (n + 1) / V

• r = V2 / g * n

• ω = g * n / V

134

Helicopter performance

• ∆P = 1* ρ * V∞* V∞ / 2

• ∆P = Thrust / Area

• Vi = √(T/2*ρ*A)

• Pi = T * Vi

• Vh2 = (VC + Vi) Vi

• T = 2 * ρ * A * (VC + Vi) Vi

• Pi = T(VC + Vi)

• T = -‐2 ρA(VC + Vi) Vi

135

Troubleshooting After testing the plane problem with the following things occurred:

Landing Gear

Figure 90

The aircraft had crashed while taxiing. It was out of control when it had maximum throttle and

ended up having a bad crash resulting in a broken support section of the landing gear.

This was anyway fixed with epoxy glue 5min as the support material is plywood , making it

stronger and rigid.

136

Propeller Directional Balance

Figure 91

Before the VTOL aircraft was flown by itself , testing its balance of hovering was very important

or it could lead to a disastrous crash.

The balance was tested by holding the aircraft while full throttle was applied. Hovering was

achieved, but the balance of the both sides of the wing was not satisfying enough to let it fly on

its own.

The reason being, both the propellers on the either side of the wing were clockwise directional.

There was no counter support for the balance. The Counter-‐clockwise propeller could not be

found anywhere in the local country or nearby. So both clock-‐wise directional propellers were

tried anyway, resulting in a very unstable balance.

137

Throttle Balance

Figure 92

The self-‐hovering of the aircraft model was put to test after the previous troubleshooting was

taken care. The problem of Throttle balance did not allow the flying to be achieved.

The tail-‐motor that was fixed to upwards direction was restricting the airplane to take-‐off as the

power of the Tail-‐motor was greater than the ones on the wing. The wing-‐motors also faced a

difference in power produced, which got the aircraft pushing itself sideways and not taking off

vertically.

This problem was corrected by fixing one main gyro board to all the controls intending to trim

the power produced and help the controls to perform uniformly. But unfortunately the electrical

components did not respond to the gyro board even after proper fixing was done.

138

Safety and Risk Assessment

• When working with glue, accelerator or acetone, remember that they are toxic and

hazardous materials. Follow all guidelines and precautions accompanying these

materials. It is easy to become complacent, as the hazard is not immediately obvious.

• Always wash hands after working with glue materials. Keep glue, accelerator and

acetone away from the eyes. Safety glasses are recommended. Avoid rubbing the eyes,

and keep the hands away from the face when working with these materials

• If power tools are used, eye protection, instruction in the safe use of the tools and

proper supervision should all be considered prerequisites.

• If not flying at a club field, make sure the site you choose is adequate and appropriate

,not too small an area and not too close to people, animals, trees, power lines, buildings,

roads, etc. Also find out if there are any local ordinances that prohibit flying RC airplanes

in public spaces.

• When working with a hand cutter make sure not to apply big forces as it might lead to

hurting your hand.

• Unless your radio system is 2.4GHz, use a frequency checker or some other method of

frequency control before turning on your transmitter. Having two or more people flying

RC airplanes on the same frequency does not work; if you interfere with another pilot’s

frequency, you will cause an accident.

• Never ever keep your hands close to the engine propeller blade it can cut anything with

the speed of 17,000 rpm.

• Don’t try flying RC airplanes in “adverse” wind conditions. Depending on your model,

that could be anything over 10-‐15 mph. Know your plane’s limitations and if unsure

about wind speed, wait for another day.

139

CONCLUSION

Based on the theoretical calculations Aero-‐chopper was successfully designed and constructed.

Aero-‐chopper has been designed to accomplish VTOL and transition to forward flight.

By completing this project it has enlightened me on many topics such as the Helicopter and

Aircraft performance individually and together.

Since Aero-‐chopper is almost a successful project, and it explains the concept of vertical take-‐off

and landing, in future we hope to see the real aircraft perform vertical take-‐off and transition

and bring alive the legendary V-‐22 osprey.

I would like to conclude by saying that according to the study conducted, calculated results and

the test performed Aero-‐chopper can perform VTOL and transition to forward flight with no

hustle.

140

References

1. http://www.theuav.com/

2. espritmodel.com

3. helicopterpage.com

4. http://www.moneysmith.net/Soaring/soaring4.html

5. www.e-‐fliterc.com/Products

6. http://www.worldofkrauss.com/foils

7. Fundamentals of Flight by Richard S Shevell

8. Basic Helicopter Aerodynamics by J.Seddon

141

Index

3

3D Drawing ·∙ 54

A

ABSTRACT ·∙ 8

Acknowledgement ·∙ 4

AIRCRAFT DESIGN ·∙ 47

Airfoil Selection ·∙ 42

Area Calculation ·∙ 78

C

CENTRE OF GRAVITY ·∙ 128

CONCLUSION ·∙ 139

CONSTRUCTION (ASSEMBLY) ·∙ 58

COST ANALYSIS ·∙ 21

D

DESIGN CONCEPT ·∙ 19

E

ELECTRICALS ·∙ 32

G

GANTT CHART ·∙ 17

GRAPHS ·∙ 74

I

INTRODUCTION ·∙ 5

M

MAN POWER ·∙ 23

Materials ·∙ 27

Mission ·∙ 16

P

Parametric Study ·∙ 10

PERFORMANCE ANALYSIS ·∙ 97, 132

R

References ·∙ 140

S

Safety and Risk Assessment ·∙ 138

SCOPE ·∙ 8

Specifications and details ·∙ 10, 12

Structure designing :(PROFILI) ·∙ 49

T

Tools ·∙ 30

Troubleshooting ·∙ 135

aerochopper(vtol concept)

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