Design and Fabrication of Ultra Light Aircraft

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DESIGN AND FABRICATION OF ULTRALIGHT AIRCRAFT USING

INCOUNTRY RESOURCES

By

NUST CADET SYED HASSAN MAHMOOD WASTI (060901) NUST CADET MOHAMMAD USMAN USMANI (060906)

NUST CADET HUMAYUN YOUSAF (050803) NUST CADET MUHAMMAD ALI (060904)

NUST CADET BILAL (060907)

COLLEGE OF AERONAUTICAL ENGINEERING

PAF ACADEMY, RISALPUR

03 SEPTEMBER 2010

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A Project Report on


INCOUNTRY RESOURCES

By



NUST CADET BILAL (060907) 69EC

Submitted to the faculty of Department of Aerospace Engineering

In partial fulfillment of the requirements for the degree of

Bachelors of Aerospace Engineering

Major: Aerospace Engineering

Department of Aerospace Engineering

College of Aeronautical Engineering

PAF Academy, Risalpur

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COLLEGE OF AERONAUTICAL ENGINEERING

PAF ACADEMY, RSIALPUR


INCOUNTRY RESOURCES

By



NUST CADET BILAL (060907) 69 EC

A report submitted to the College of Aeronautical Engineering

In partial fulfillment of the requirements for the degree of B.E

APPROVED

(**************) (**************)

Wing Commander Messam Abbas Group Captain ABDUL MUNEM KHAN

Project Advisor Head of Aerospace Deptt.

College of Aeronautical engineering College of Aeronautical Engineering

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Contents List of Tables ....................................................................................................................................... 7

List of Figures ..................................................................................................................................... 7

CHAPTER I ......................................................................................................................................... 8

INTRODUCTION .................................................................................................................................... 8

Definition of Ultra Light Aircraft ........................................................................................................ 8

Objective .............................................................................................................................................. 8

Methodology followed ........................................................................................................................ 8

Market Surveys and Research ......................................................................................................... 8

Conceptual Design ............................................................................................................................. 8

Aerodynamic Evaluation ................................................................................................................... 9

Detailed design & CAD modeling .................................................................................................... 9

Full Scale Fabrication ........................................................................................................................ 9

CHAPTER II ...................................................................................................................................... 10

CONCEPTUAL DESIGN ..................................................................................................................... 10

Introduction ........................................................................................................................................ 10

Phases of Aircraft Design................................................................................................................ 10

Conceptual Design Process ........................................................................................................... 11

Selection of Specification ................................................................................................................ 12

Mission Profile ................................................................................................................................... 13

Base Structure .................................................................................................................................. 14

Wing Geometry Selection ............................................................................................................... 14

Airfoil Selection ................................................................................................................................. 17

Engine Location ................................................................................................................................ 19

Landing Gear .................................................................................................................................... 20

Propeller Selection ........................................................................................................................... 20

Engine Cooling System ................................................................................................................... 20

Horsepower to Weight Ratio and Wing Loading ......................................................................... 21

Aircraft Final Specifications ............................................................................................................ 22

Three Dimensional Conceptual Model .......................................................................................... 23

CHAPTER III ......................................................................................................................................... 24

AERODYNAMICS ................................................................................................................................ 24

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Lift Slope Curve ................................................................................................................................ 24

Profile Drag ....................................................................................................................................... 24

Induced Drag Factor ........................................................................................................................ 25

Thrust Available and Thrust Required .......................................................................................... 26

CHAPTER IV ..................................................................................................................................... 27

DETAILED DESIGN ............................................................................................................................. 27

Fuselage ............................................................................................................................................ 27

Cockpit ............................................................................................................................................... 29

Flight Instruments ............................................................................................................................. 30

LANDING GEAR ANALYSIS .......................................................................................................... 43

Firewall ............................................................................................................................................... 48

Safety ................................................................................................................................................. 48

Control Surfaces ............................................................................................................................... 48

Buckling Analysis ............................................................................................................................. 54

Bolt Size Calculation ........................................................................................................................ 56

POWER PLANT SELECTION ........................................................................................................ 58

Modifications: .................................................................................................................................... 59

Alternative engine suggestions: ..................................................................................................... 61

PROPELLER DESIGN .................................................................................................................... 62

Results ............................................................................................................................................... 63

Conclusion: ........................................................................................................................................ 67

Solid edge view of the designed propeller: .................................................................................. 67

Fuel system ....................................................................................................................................... 68

CHAPTER V ...................................................................................................................................... 70

FABRICATION ...................................................................................................................................... 70

Material Selection ............................................................................................................................. 70

Fabrication of Connectors ............................................................................................................... 73

Fabrication of Custom Designed Joints and Hinges ................................................................... 74

Fuselage Construction ..................................................................................................................... 76

2d Truss Construction ..................................................................................................................... 84

Wing Construction ............................................................................................................................ 87

Empennage Assembly ..................................................................................................................... 90

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Control Surfaces ............................................................................................................................... 94

Fabrication of Landing Gear ......................................................................................................... 101

Fabrication of Aircraft Skin Covering: .......................................................................................... 105

Solid Edge Figures ......................................................................................................................... 106

CONCLUSION ................................................................................................................................ 111

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List of Tables

Table 1: Aircraft Specifications ................................................................................................................... 12

Table 2: Airfoil Comparison ......................................................................................................................... 17

Table 3: Wing Loading and HP/W ............................................................................................................... 21

Table 4: Aircraft Final Dimensions .............................................................................................................. 22

Table 5: Loads Estimation ........................................................................................................................... 31

Table 6: Top View showing wing spars and major ribs of the wing ............................................................ 31

Table 7: Side View showing wing spars and struts ..................................................................................... 31

Table 8: Fuselage Truss members ............................................................................................................... 85

Table 9: Wing members .............................................................................................................................. 88

Table 10: Horizontal Tail members ............................................................................................................. 91

Table 11: Vertical Tail members ................................................................................................................. 93

Table 12: Landing Gear members ............................................................................................................. 101

List of Figures

Figure 1: Iterative Procedure of Concept Design ........................................................................................ 11

Figure 2: Mission Profile ............................................................................................................................. 13

Figure 3: Base Structure .............................................................................................................................. 14

Figure 4: Clark Y Airfoil Coordinates ........................................................................................................... 18

Figure 5: Clark Y Characteristics .................................................................................................................. 18

Figure 6: Constraint Diagram ...................................................................................................................... 21

Figure 7: Three Dimensional Views of Conceptual Design ......................................................................... 23

Figure 8: Lift Slope Curve ............................................................................................................................ 24

Figure 9: Profile Drag .................................................................................................................................. 24

Figure 10: Profile Drag at Different Altitude ............................................................................................... 25

Figure 11: Induced Drag Factor ................................................................................................................... 25

Figure 12: Drag Polar................................................................................................................................... 26

Figure 13: Thrust Available and Thrust Required ....................................................................................... 26

Figure 14: Effect of Fineness Ratio on Fuselage Drag ................................................................................. 28

Figure 15: Definition of Upsweep and its effect on Drag............................................................................ 28

Figure 16: FEM Model of the Wing ............................................................................................................. 32

Figure 17: Application of Loads on the Wing .............................................................................................. 33

Figure 18: von Misses Stress distribution in wing members ...................................................................... 33

Figure 19: Horizontal Tail Geometry ........................................................................................................... 35

Figure 20: von Misses Stress distribution in Horizontal Tail members ....................................................... 36

Figure 21: Vertical Tail Geometry ............................................................................................................... 37

Figure 22: von Misses Stress distribution in Vertical Tail members ........................................................... 38

Figure 23: Fuselage Truss dimensions ........................................................................................................ 39

Figure 24: von Misses Stress distribution in modified Fuselage Truss ....................................................... 42

Figure 25: Fuselage Cabin ........................................................................................................................... 43

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CHAPTER I

INTRODUCTION

Definition of Ultra Light Aircraft

1. The definition of ultra light aircraft varies from country to country the aviation

regulatory body of the country decides the weight of an aircraft to be classifies as ultra

light aircraft. However, the civil aviation authority of Pakistan has no such definition for

ultra light aircraft. As a result we selected the empty weight of our aircraft closer to that

used in India and set it to 800lb.

Objective

2. The aim of the project is to design and fabricate a full scale ultra light aircraft using

in-country resources.

Methodology followed

a) Market Survey’s and Research

b) Conceptual Design

c) Aerodynamic evaluation

d) Detailed Design, CAD modeling

e) Full scale fabrication

Market Surveys and Research

3. First step was initiated by carrying out a thorough web research followed by many

surveys done at Peshawar, Lahore and Karachi flying and ultra light hobby clubs. Raw

data of different ultra light aircraft made worldwide and in Pakistan was collected and

arranged for comparison purposes. This step helped in setting the initial specification of

the aircraft.

Conceptual Design

4. Conceptual design was generated following Design books by Daniel P Raymer and

Roskam. An iterative process resulted in optimized conceptual design.

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Aerodynamic Evaluation

5. Aerodynamic characteristic of the aircraft were evaluated using theoretical

aerodynamics and final configuration was set.

Detailed design & CAD modeling

6. Structural analysis was carried out using ANSYS software and a detailed sketch of

the aircraft was created using Solid Edge software. Geometric drawings were obtained

to use for fabrication process.

Full Scale Fabrication

7. After creating detailed geometry, full scale templates of different parts were

created. Fabrication was done using tools and technology available locally. Pipe

marking, cutting, drilling and welding are few of main processes done in fabrication

phase. All the structural parts including cabin, truss, wings, control surfaces, landing

gear and empennage assembly were fabricated separately and assembled together to

complete the aircraft structure.

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CHAPTER II

CONCEPTUAL DESIGN

Introduction

1. Airplane is the intellectual engineering process of creating. Aircraft design is a separate discipline of aeronautical engineering – different from aerodynamics, structures, propulsion and controls. The job of a designer involves a lot of capabilities including knowledge of the above disciplines, his experience, talent, good approach, hard work and utilization of available resources and tools.

2. A good aircraft design seems to miraculously glide through subsequent evaluations by specialists without major changes being required. Somehow the landing gear fits, the fuel tanks are near the center of gravity, the structural members are simple and light weight, overall arrangement provides good aerodynamics, the engines installed in simple and clean fashion and a host of similar detail seems to fall in space.

Phases of Aircraft Design

3. There are three phases of aircraft design process.

I. Conceptual Design: Conceptual design is the primary phase. It involves configuration arrangement, size, weight and performance parameters. An affordable aircraft will be the one which meets all these requirements.

II. Preliminary Design: A preliminary design begins when major changing is over. During this phase the areas of interest are structures, landing gears and control system. Testing is initiated in areas such as aerodynamics, propulsion and stability and control parameters. The ultimate objective during this phase is to get full-scale development

III. Detail Design: Assuming a favorable decision for entering full-scale development, the detail design phase begins in which the parts of the aircraft to be fabricated are redesigned. For example, individual ribs, bolts etc are designed and analyzed. Detail design ends with the fabrication of the aircraft.

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Conceptual Design Process

4. Conceptual design is the very first step of aircraft designing where the basic configuration, size, layout, weight and performance are set. It is done by going over the set or desired requirements repeatedly while validating their feasibility. This process answers the questions whether an aircraft with the requirements can be built to fly or not. As it has been made clear that there is a set of certain characteristics that the design has to meet. These specifications can be the requirement of the user or the designer. Conceptual design process starts with the rough sketch of the aircraft, which is being designed. This gives us a very crude idea of what we are going to design. This sketch may include approximate wing geometry, location of engines, payload, passengers, cockpit and landing gears etc.

5. After this initial sketch rough weight estimation will be done, it will be followed by wing geometry selection and the calculation of other important parameters of the aircraft such as wing loading, thrust to weight ratio etc. After all this has been done, initial sizing will be carried out. In the last an iteration process will be carried out which will result in the final values of all the parameters of the aircraft.

Figure 1: Iterative Procedure of Concept Design

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Selection of Specification

6. Ultra light aircrafts are usually used for recreation and therefore loiter time rather

than the range is more important. As no design requirements were given therefore a

survey was conducted and the specifications were selected accordingly.

7. The specifications that were selected were:

Range 10 mile

Endurance 20 min

Stall speed 25 mph

Max speed 60 mph

Cruise speed 52 mph

TO distance 400 ft

Landing distance 300 ft

Ceiling 5000 ft

Climb rate 550 fpm

Table 1: Aircraft Specifications

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Mission Profile

Figure 2: Mission Profile

TO

RANGE 10 MILES

LOITER 20 MIN

LANDING

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Base Structure

8. Afterwards the base structural design was selected:

Figure 3: Base Structure

Wing Geometry Selection

9. The wing geometry includes taper ratio, aspect ratio, dihedral, sweep, planform,

twist, wing location, thickness and incidence.

I. Taper Ratio

High Taper Ratio Low Taper Ratio

Weight High Low

Tip stall Good Poor

Manufacturing Easy Difficult

Hence, it was decided to have a straight wing as flow separation downstream from

the root region causes buffeting as it flows over the horizontal tail, thus providing

stall warning to the pilot. Moreover, as the wing tip still has attached flow control

surfaces would still be operatable. Besides, a straight wing is cheaper and easy to

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manufacture. However, special efforts would be made in manufacturing as Lift

distribution is such that the centroid is away from the root causing greater bending

moment.

II. Wing Location

High Mid Low

Interference drag Poor Good Poor

Dihedral effect Negative Neutral positive

Visibility Good Good Poor

High wing configuration was selected as:

a. It adds to the lateral stability of the aircraft.

b. The wings will not strike the ground on landing

c. Safe from FOD

d. Fuel system can be incorporated in it. (gravity fed)

e. Wing box straight through the fuselage.

f. Easy to manufacture

III. Monoplane/Biplane

Monoplane Bi - plane

Cantilever Braced

Weight High Low Very low

Profile Drag Low High Higher

Interference Drag Low High Higher

For less weight and greater structure strength, braced wings would be used and to

make sure that drag is not very high, monoplane was preferred over biplane. More

importantly from manufacturing point of view monoplane construction is much easier

than biplane.

IV. Wing Sweep

None Sweep

Lift curve slope High Low

Pitch attitude in low speed level flight Low High

Ride through turbulence Poor Good

Stall Good Poor

Lateral control at stall Good Poor

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Wing weight Low High

No sweep was selected as it gives lower weight, better performance at low speed

(our regime) and gives better stall characteristics(assuring safety).

V. Wing Dihedral

Effect of Dihedral angle

Positive Negative

Spiral stability Increased Decreased

Dutch Roll Stability Decreased Increased

Ground clearance of wing Increased Decreased

As Ultra light aircrafts are supposed to be stable therefore a small positive dihedral

angle would be used.

VI. Wing Incidence

Large Small

Cruise drag High Low

Cockpit visibility Good Average

As incidence angle would increase drag therefore it was decided to use zero

incidence angle.

VII. Wing Thickness

Low t/c High t/c

Wing weight High Low

Subsonic wing drag Low High

Wing Fuel volume Poor Good

Maximum Lift Poor Good

As greater thickness ratio increases lift as well as decrease weight, we will use

higher t/c but greater than 12 Cl max starts to deteriorate, thus , we will use t/c less

than 12.

VIII. Aspect Ratio

HIGH LOW

Induced Drag Low High

Lift-curve Slope High Low

Pitch Attitude (approach) Low High

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Ride in Turbulence Poor Good

Wing Weight High Low

Wing Span Large Small

Due to higher L/D, higher aspect ratio is selected. But higher value is limited by

increasing weight with aspect ratio. Thus, a trade off is carried out and aspect ratio

of 6 is finalized.

IX. Twist

Large Small

Induced drag High Small

Tipstall Good Poor

Wing Weight Mildly lower Mildly higher

To decrease complexity and to improve induce drag as wing span is already very

large, we will use no twist.

Airfoil Selection

Airfoil CLmax

CLARK Y 1.65

NASA GA(W)-1 1.7

NASA GA(W)-2 1.8

NACA 2412,43012

1.65

Table 2: Airfoil Comparison

Clark Y was used as an airfoil as it is easy to build and is most commonly used is Ultra

light aircrafts worldwide.

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Figure 4: Clark Y Airfoil Coordinates

Figure 5: Clark Y Characteristics

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Engine Location

Tractor configuration

Advantages

1. CG moves forward. Improving stability and allowing a shorter and smaller tail

2. Propeller works in an undisturbed free stream

3. More effective flow of cooling engine

Disadvantages

1. The propeller slipstream disturbs the quality of air over the wing

2. Skin friction increases over the wing

Pusher configuration

Advantages

1. Undisturbed flow over the wing and fuselage

2. Favorable pressure gradient at the rear of the fuselage prevents flow separation

3. Engine noise reduced

4. Pilot’s view improved

Disadvantages

1. CG shifts back which causes stability problem

2. Propeller damaged by FODs

3. Engine cooling problem more severe

After considering the pros and cons of both the configurations it was decided to use the

more conventional Tractor configuration. The major factor in making this decision was

the use of an automobile engine. As we have modified the engine therefore, it would be

a huge risk if we are using the pusher configuration as any cooling problem could result

in a crash.

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Landing Gear

Tricycle Tail dragger

Groundloop behavior Stable Unstable

Visibility over the nose Good Poor

Floor attitude on the ground Level Not level

Weight Medium Low

Steering after touchdown Good Poor

Takeoff rotation Good Good

Although the Tricycle landing gear has more advantages but the tail dragger

configuration was selected as:

Provides clearance for the propeller.

Less drag and weight

Wing creates more lift as it is already at an angle of attack

Easier to fabricate

However, it is inherently an unstable configuration during ground roll. If the

airplane starts to turn during ground roll CG tends to swing around causing the

turn to get tighter and tighter. So the pilot must keep the airplane always aligned

with the runway.

Propeller Selection

Three practical constraints would be kept in mind while buying the propeller:

1. Propeller tip must clear the ground

2. Propeller tip should not reach supersonic speeds. As compressibility effects

would ruin the propeller performance.

3. Propeller must be large enough to absorb engine power. The power absorption

of propeller is increased by increasing the diameter.

The final choice of propeller used would solely depend on the availability of propeller as

they are not easily available.

Engine Cooling System

Updraft cooling system is used as it, unlike downdraft cooling system, flows the

cooling air upward through the cylinders and exits it into low pressure air above the

fuselage, creating more efficient cooling flow due to a suction effect.

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Horsepower to Weight Ratio and Wing Loading

The thrust to weight ratio (HP/W) and the Wing Loading (W/S) are the two most

important parameters affecting the aircraft performance. Wing loading and horsepower

to weight ratio are interconnected for a number of performance. The major part of the

analytical design is the optimization of these factors as these are more interconnected

in the segments of take off, landing, turn and glide etc. To achieve their values,

constraint diagram was formulated, using the most critical segment of our profile:

Figure 6: Constraint Diagram

W/S 4.7941

HP/W 0.08

Power Loading 12.5

Table 3: Wing Loading and HP/W

0

0.05

0.1

0.15

0.2

0.25

0 5 10 15

HP

/W

W/S

Título del gráfico

STALL LIMIT

TO

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Aircraft Final Specifications

After going through different phases of concept design, final aircraft specifications come

out as follow:

All dimensions in feet.

Fuselage Control Surfaces

Length 17 Aileron

Width 2.5 Chord 0.25 of wing

Height 4.25 Span 0.45 of wing

Wing Elevator

Span 28 Chord 0.9 of stabilizer

Chord 4.66 Span 0.45 of stabilizer

Span area 130.33 Rudder

Sweep 0 Chord 0.4 of fin

Aspect Ratio 6 Span 0.9 of fin

Taper ratio 1 Flaps

Dihedral 0 Chord 0.25 of wing

Span 0.55 of wing

Horizontal Tail Cl design 0.75136

Chord 3.2

Span 9.5 Weights

Span area 29.7 payload 200

Lht 10.2 Structural weight 302

Aspect Ratio 3 Engine weight 150

Taper Ratio 0.6 Maximum weight (fully loaded)

652

Vertical Tail Clean Configuration

Chord 3.5 K 0.066306

Span 4.1 Cdo 0.033043

Span area 14.2 Take Off

Lvt 10.2 Cdo 0.048043

Aspect Ratio 1.2 K 0.070726

Taper Ratio 0.6 Landing

Power plant Cdo 0.098043

Power required 60 hp K 0.075778

Fuel required 32 lbs.

Fuel volume required 20 litres Take off and Landing

Air Cooling System ST/O 286.1

Air Intake Area 0.278 sq. ft. Sldg 390

Propeller Diameter 4.3 ft.

Tires

Diameter 12 inch.

Width 4.2 inch.

Table 4: Aircraft Final Dimensions

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Three Dimensional Conceptual Model

Isometric View Front View

Side View Top View

Figure 7: Three Dimensional Views of Conceptual Design

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CHAPTER III

AERODYNAMICS

Lift Slope Curve

Cl alpha versus Mach No:

Figure 8: Lift Slope Curve

Profile Drag

Profile Drag versus Mach No:

Figure 9: Profile Drag

4.282

4.284

4.286

4.288

4.29

4.292

4.294

4.296

0 0.02 0.04 0.06 0.08 0.1

CL

alp

ha

M

CL alpha vs M

0

0.005

0.01

0.015

0.02

0.025

0 0.02 0.04 0.06 0.08 0.1

CD

M

CD vs Mach No.

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Drag at different altitude:

Figure 10: Profile Drag at Different Altitude

Induced Drag Factor

Induced Drag Factor versus Mach No:

Figure 11: Induced Drag Factor

0

0.005

0.01

0.015

0.02

0.025

0 0.02 0.04 0.06 0.08 0.1

CD

M

CD vs M

SL

1000

2000

3000

4000

5000

0

0.02

0.04

0.06

0.08

0.1

0.12

0 0.02 0.04 0.06 0.08 0.1

K

M

K vs M

0-0.8

above 0.8

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Drag Polar

Figure 12: Drag Polar

Thrust Available and Thrust Required

Thrust Available versus Thrust Required:

Figure 13: Thrust Available and Thrust Required

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 0.1 0.2 0.3 0.4

Cl

CD

Drag Polar

Cruise

T/O

Landing

0

50

100

150

200

250

300

350

400

450

500

0 0.05 0.1

thru

st (

lb)

Mach

Thrust available vs Thrust Required

thrust available

thrust required

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CHAPTER IV

DETAILED DESIGN

Components of the Aircraft

1. Wing

2. Fuselage Cabin

3. Fuselage Truss

4. Horizontal Tail

5. Vertical Tail

6. Flaps

7. Control Surfaces (Ailerons, Elevators, Rudder)

8. Engine Mount

9. Wing Mount

10. Connectors/Joints

11. Landing Gear

12. Control System

Fuselage

Fuselage is responsible for the largest portion of overall drag for most of the airplanes.

Thus it should be sized and shaped accordingly for minimum drag. It contributes to

various drags:

Friction drag.

Profile drag.

Base drag.

Compressibility drag.

Induced drag.

To reduce friction drags, two options are available:

Shape the fuselage so that laminar flow is possible.

Reduce the length and perimeter as much as possible.

If fuselage length is decreased, for the same level of static stability, tail size can be

decreased, thus, decreasing the friction drag. Thus, the optimum fineness ratio is 4 to 8.

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Figure 14: Effect of Fineness Ratio on Fuselage Drag

Upsweep of fuselage:

Upsweep is needed for clearance during take-off. It also gives clearance during taxi in

tail dragger arrangement.

Figure 15: Definition of Upsweep and its effect on Drag

As we can see in figure, drag does not increase much up to the upsweep of 15

degrees.

Our proposed design:

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Upsweep 12.5 degrees

Fineness ratio 4

Cockpit

The basic idea behind the cockpit design is to provide comfort to the pilot besides easy

excess to the all vital controls and accessories without diverging the pilot’s attention.

Cockpit design especially for homebuilders:


Sidewise motion of stick: 15 degrees.

Distance between rudder pedals: 50cm.

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Flight Instruments

Following are the most common instruments used in most of the homebuilders:

1. airspeed indicator

2. altimeter

3. magnetic compass

4. tachometer

5. oil pressure gauge

6. oil temperature gauge

7. fuel quantity indicator

According to FAR 23, depth of instrument panel should be around 1 foot.

Load Estimation

Before going for further structural design, it is necessary to know the loads the structure

is going to sustain. The total takeoff weight of the aircraft is 568 lbs and it will take a

Load Factor of 2.0 during its flight. A Factor of Safety of 1.5 will be used for the

structural design of the aircraft.

Assuming that the wing generates all the lift, the wing must produce lift equal to the

weight of the aircraft. Thus, total vertical force experienced by the wing will be the

weight multiplied by both the Load Factor and the Factor of Safety which gives a value

of 1700 lbs. Wing drag is estimated to be 10% of its lift which equals 170 lbs.

The lift and drag of horizontal tail is estimated to be 20% of that of the wing. So, lift and

drag of horizontal tail becomes 340 lbs and 34 lbs respectively.

Since the vertical tail has similar configuration as half of the horizontal tail, so we can

take the side force and drag of the vertical tail to be half of the lift and drag of the

horizontal tail respectively. Thus, the side force becomes 170 lbs with a drag force of

17 lbs.

The following table summarizes the load estimation. These loads will be used to

analyze the design.

Member Vertical Loads (Lift, lbs)

Horizontal Loads (Drag, lbs)

Lateral Loads (Side Force, lbs)

Wing 1700 170 --

Horizontal Tail 340 34 --

Vertical Tail -- 17 170

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Table 5: Loads Estimation

Wing Design

Dimensions

As per Conceptual Design, the wing dimensions are:

Span 28 ft

Chord Length 3.5 ft

The wing is straight rectangular without any twist.

Components

The wing comprises of:

1. Leading edge Spar

2. Trailing Edge Spar

3. Major and Minor Ribs

4. Struts

Construction:

The wing will be constructed in two symmetrical parts which will be later joined at the

wing mount. Each of the part will be 13 ft long and a 2 ft connector at the wing mount

will make the total span to 28 ft. The following design is proposed for the wing part.

Table 6: Top View showing wing spars and

major ribs of the wing

Table 7: Side View showing wing spars

and struts

All the members consist of Al pipes. An iterative trial and error structural analysis will

determine the dimensions of these pipes.

Determination of Dimensions of Pipes

In order to determine the outer diameter and thickness of the pipes, the proposed

design was analyzed in ANSYS using different combinations of available diameters and

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thicknesses. The best combination was then selected as the final dimensions of the

pipes.

Procedure of Analysis

The basic geometry was predetermined and the variable parameters were outer

diameter and thickness of the spars, ribs and struts as well as the span wise location of

the struts. So, the basic geometry was generated in ANSYS and arbitrary values of

outer diameter and thickness were given.

Figure 16: FEM Model of the Wing

The geometry was then meshed with element type PIPE16. Loads were applied as

estimated before at 14 Hard Points. These Hard Points are the points where ribs will be

joined. So, applying loads at these points will be quit suitable. It is assumed that the

loads are equally distributed among these points. The Load Step was solved and von

Misses Stress was plotted and maximum stress was noted. These steps were repeated

until a satisfactory result was obtained.

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Figure 17: Application of Loads on the Wing

Result of the Analysis

The analysis showed the following combination to be best suited which was selected as

the final dimensions.

Dimension Value

Spar Outer Diameter 2.00 in

Spar Thickness 0.78 in

Strut Outer Diameter 1.50 in

Strut Thickness 0.75 in

Major Rib Outer Diameter 2.00 in

Major Rib Thickness 0.78

Strut Location Mid Span

For the above combination, plot of von Misses Stress is shown below:

Figure 18: von Misses Stress distribution in wing members

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The factor of Safety can be calculated as:

Maximum Stress = 29,510 psi

Tensile Strength of Material = 30,000 psi

𝐹𝑂𝑆 =𝑆𝑡𝑟𝑒𝑛𝑔𝑡ℎ 𝑜𝑓 𝑀𝑎𝑡𝑒𝑟𝑖𝑎𝑙

𝑀𝑎𝑥𝑖𝑚𝑢𝑚 𝑆𝑡𝑟𝑒𝑠𝑠=

30,000

29150= 1.03

Since, we have already multiplied the applied loads with 1.5 as a FOS, so the actual

FOS will be 1.5 X 1.03 = 1.54.

Minor Rib Construction:

One of the most important parts of wing design is to give its cross section the shape of

desired airfoil. In order to do so, seven ribs are placed on each side of the wing at a

distance of 2.25 ft apart. The ribs are constructed from two pipes. A straight pipe is used

for the lower surface to get a flat bottom, as designed. The upper surface will be given

the required curvature by using a bended pipe.

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Horizontal Tail

Dimensions

As per Conceptual Design, the horizontal tail dimensions are:

Span 9.50 ft (114 in)

Root Chord Length 3.25 ft (39 in)

Taper Ratio 0.60

Sweep Angle 15 Deg

Keeping in mind the above dimensions, the following geometry is proposed for the

horizontal tail.

Figure 19: Horizontal Tail Geometry


In order to determine the outer diameter and thickness of the pipes, the proposed

design was analyzed in ANSYS using different combinations of available diameters and

thicknesses. The best combination was then selected as the final dimensions of the

pipes.


The basic geometry was predetermined and the variable parameters were outer

diameter and thickness of the spars and ribs. So, the basic geometry was generated in

ANSYS and arbitrary values of outer diameter and thickness were given. The geometry

was then meshed with element type of PIPE16. Loads were applied as estimated before

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at 06 Hard Points. It is assumed that the loads are equally distributed among these

points. The Load Step was solved and von Misses Stress was plotted and maximum

stress was noted. These steps were repeated until a satisfactory result was obtained.


The analysis showed the following combination to be best suited which was selected as

the final dimensions.

Dimension Value

Outer Diameter 1.50 in

Thickness 0.75 in

For the above combination, plot of von Misses Stress is shown below:

Figure 20: von Misses Stress distribution in Horizontal Tail members

The factor of Safety can be calculated as:





30,000

28,496= 1.05


FOS will be 1.5 X 1.05 = 1.58.

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Vertical Tail

Dimensions

As per Conceptual Design, the horizontal tail dimensions are:

Span (Height) 04 ft (48 in)

Root Chord Length 2.5 ft (30 in)

Taper Ratio 0.60

Sweep Angle 15 Deg

Keeping in mind the above dimensions, the following geometry is proposed for the

horizontal tail.

Figure 21: Vertical Tail Geometry

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Since, dimensions of pipes for horizontal tail are selected, we will check, whether these

dimensions can be used for vertical tail or not. If the Factor of Safety is acceptable, we

will use the same pipes for vertical tail construction.


The basic geometry was generated in ANSYS and the values of outer diameter and

thickness were given as 1.5 in and 0.75 in respectively. The geometry was then meshed

with element type of PIPE16. Loads were applied as before at 06 Hard Points. It is

assumed that the loads are equally distributed among these points. The Load Step was

solved and von Misses Stress was plotted and maximum stress was noted.


The von Misses Stress distribution is shown below. The maximum stress comes out to

be 26,679 psi.

Figure 22: von Misses Stress distribution in Vertical Tail members

The Factor of Safety can be calculated as:





30,000

26,679= 1.12

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FOS will be 1.5 X 1.12 = 1.69.

The Factor of Safety came out to be 1.69 which is quiet acceptable. So, the same pipes

will be used for vertical tail.

Dimension Value


Thickness 0.75 in

Fuselage Truss

Dimensions

The fuselage truss must be 9.5 ft (114 in) long. The other dimensions are as shown

below.

Figure 23: Fuselage Truss dimensions


The Fuselage Truss will be constructed from the same pipes used for the tails i.e.

Dimension Value


Thickness 0.75 in

Determination of Truss Members

A 2 dimensional truss was proposed as the initial design. This design was analyzed

using ANSYS. By noting the results, necessary modifications were made. The proposed

design is shown below.

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The basic geometry was generated in ANSYS and the values of outer diameter and

thickness were given as 1.5 in and 0.75 in respectively. The geometry was then meshed

with element type of PIPE16. Loads were applied at the attachment points of the

horizontal and vertical tail as well as the rear landing gear. The loads at these points

were obtained from previous analysis of horizontal and vertical tails. An estimated

vertical load of 140 lbs was applied at rear landing gear attachment point. The loads

are summarized below:

Loads HT Point 1 HT Point 2 VT Point 1 VT Point 2 Rea LG

FX (lbs) 21 15 15 21 --

FY (lbs) 75 96 -28 28 140

FZ (lbs) 26 -26 97 74 --

The Load Step was solved and von Misses Stress was plotted and maximum stress

was noted.


The analysis showed that the lateral displace of the truss structure is very large. The

displacement came out to be 20 in which is not acceptable. So, design must be

modified to restrict the lateral diplacement of the tail. Plot of lateral displacement is

shown below.

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Fuselage Truss Modification

In order to restrict the lateral displacement, members must be added which can take

lateral loads. So, four members are added as shown below.

Analysis of Modified Truss

The modified truss was analyzed as done before.


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Lateral Displacement: The lateral displacement of the tail of the modified truss

structure came out to be about 4 in. this value is acceptable if the stress does not

exceeds the strength of the material used. Plot of lateral displacement is shown below.

von Misses Stress: The maximum von Misses Stress came out to be 26, 385 psi which

gives a FOS of about 1.70. So, the design is safe and can be accepted. Distribution of

von Misses Stress is shown as under:

Figure 24: von Misses Stress distribution in modified Fuselage Truss

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Fuselage Cabin

Fuselage cabin is the part of aircraft where most of the components get jointed. Wing,

landing gears, engine mounts and fuselage truss are connected to the cabin. It also has

the provision for installment of flight instruments and control system. It will also serve as

the cockpit to house the pilot. The following figures show the sketch of fuselage cabin.

Front View Side View

Figure 25: Fuselage Cabin

LANDING GEAR ANALYSIS

For the landing gear analysis of ultra light aircraft a standard drop test is carried out. In

this test the aircraft is dropped from a height of 3 ft and the impact taken by landing gear

is closely observed if the landing gear is able to fully survive the impact it is declared

safe. However, we have used ANSYS to simulate all the critical landing loads.

LANDING GEAR CRITICAL LOADS

Critical loads as stated by FAR 23 and those suggested on discussion forums of

ultralight aircraft were considered.

According to FAR 23 regulation the most severe conditions for landing were

• Vertical load equal to 75% of max ground reaction

– 10 fps for design landing weight

– 6 fps for design TO weight

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• Drag 40% of the max weight

• Side force 25% of vertical force

• This combination acts at the wheel axle centerline.

However, most discussion forums on the internet suggested that forces up to 3g’s could

be experienced during poor landing and hence vertical loads were selected accordingly.

The corresponding loads used were:

◦ Drag/Friction : 280 lbs

◦ Lateral loads due to bad landing : 350 lbs

◦ The vertical reaction on touchdown calculated on suggestions from discussion

forums describing landings at 3gs : 1400 lb

Dimensions Selection

Dimensions were selected on the following requirement:

The final landing gear dimensions are:

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SOLID EDGE VIEW

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ANSYS ANALYSIS

The greatest challenge we faced during the analysis was the non availability of damping

coefficient and spring constant value of dampers. Although spring constant could be

found. It was impossible to calculate the damping coefficient without proper

experimental setup. Therefore, it was decided to find the loads in the rods connecting

the dampers and hope that a damper could be found with the required load rating

inscribed on it.

The element used for landing gear analysis was Pipe 16:

Uniaxial element with tension-compression, torsion, and bending capabilities

The element has six degrees of freedom at two nodes: translations in the nodal

x, y, and z directions and rotations about the nodal x, y, and z axes.

Material Used

Steel 4130 N is used. Use of steel also means that the structure can be easily welded

therefore analysis of welded structure was done.

Density ρ [kg/dm3] 7.9

Elastic modulus E [GPa] 200

Poisson’s ratio ν 0.28

Thermal conductivity κ [W/mK] 31.4

Thermal expansion α (1/K) 20 e-6

Ultimate Tensile strength (Mpa) 730

PROCEDURE

As we couldn’t get the values of “k” and “c” for the shock absorber and the design

papers were only specified for oleo shock absorbers therefore they couldn’t be

used

Hence it was decided that if individual load in the member where shock absorber

is to be installed could be found then the shock absorber can be selected

The maximum stresses in each member were found.

The load was then calculated by multiplying with area

These loads were then used to find a suitable shock absorber

Usually FOS of 1.5 is used but I have used 2 as the local material might not have

perfect strength

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TWO POINT LANDING CONDITION

The results as calculated by ANSYS were max. stress of 3269 psi. which corresponded

to a force of 845 lbs. The maximum stress was also less than the failure stress.of 105

ksi

Fig. 2 point landing analysis

The deformation calculated was also acceptable

Fig. deformation diagram

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Firewall

Stainless steel or galvanized iron sheet of minimum thickness .015 inch (according to

FAR 23) is installed between cockpit and engine to prevent pilot from engine heat.

Safety

For the safety of pilot, cabin should consist of :

Ballistic Parachute

Fire Extinguisher

Control Surfaces

Flight control systems can be divided into primary and secondary flight control systems.

Primary control systems:

Lateral Controls: ailerons, spoilers and differential stabilizer.

Longitudinal controls: elevator, stabilizer and canard.

Lateral controls: rudder.

Secondary control systems:

Trim controls: lateral, longitudinal and directional controls.

High lift controls: trailing and leading edge flaps and slats.

Thrust controls: engine fuel controls (throttle), manifold gates and propeller blade

incidence.

Based on their design, flight control systems are further divided into:

Reversible flight control systems.

Irreversible flight control systems.

Reversible flight control systems:

In a reversible flight control system, when the cockpit controls are moved, the

aerodynamic surface controls moves and vice-versa. They are typically mechanized

with cables, push rods or a combination of both.

Irreversible flight control systems:

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In an Irreversible flight control systems (hydraulic or electrical), when the cockpit

controls are moved, the aerodynamic surface controls moves and not vice-versa.

Another way of stating this is to say that in an irreversible flight control system an

actuator moves the aerodynamic surface controls. The pilot merely signals the actuator

to move. This signaling process is usually an irreversible process.

Major advantages associated with reversible flight control system are:

Simplicity.

Low cost.

Reliability.

Relatively maintenance free.

In laying out a reversible flight control systems, the following important design aspects

need to keep in mind:

Mechanical design requirements for cable and push-rod systems.

Efficiency consideration.

Cable and push rod control force levels.

Control surface types and hinge moments.

Aerodynamic balance requirements.

Mass balance requirements.

Major design problems associated with reversible flight control system are:

Cable stretch.

Cable slack.

Friction.

Weight.

Handling qualities.

Elastic control system deformation.

Flutter.

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Cable stretch:

To prevent cable stretch l/s >6.

Cable slack:

Turnbuckles are used to prevent cable slack and to prevent cables from leaving pulley,

cable guards are used.

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Friction:

To prevent too much friction, following rules to be considered:

Keep cable runs as straight as possible.

Keep the number of guides and pulleys as small as possible.

For every cable turn, additional pulley is needed which introduces extra weight and

extra friction.

Elastic control system deformation:

To prevent elastic control system deformation, the following rules should be

observed:

Use oversized cables (this will cost weight).

Make sure pulleys are attached to stiff structural components. Do not attach them to flat

plates, they deform easily.

Flutter The solution to the control surface flutter is :

To make sure that centre of mass not behind hinge line

try to have chordwise cg at hinge line

prevent any play in the linkages

Design consideration:

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Primary flight control system cables should have a diameter greater than .11inch

(2.8mm) according to FAR 23.

Rotation efficiency is highest if the angle between the cable and the driving

sector is 90 degrees with the system in its neutral position.

Kinematic feasibility.

Mechanisms A and B are kinematically sound, the quadrangles ABFE and CDFE

remain parallelograms when the system is used. Mechanism A is better than B

because of its higher rotation efficiency. Mechanism C and D are unworkable

because the cable lengths AB and CD do not remain constant after some rotation.

Thus mechanism C and D are undesirable.


Reversible flight control system is selected, which consists of only cable pulley.

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Three Dimensional Sketches

Front View

Side View

Top View

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Buckling Analysis

Our aircraft cabin contained vertical members that would take compressive stresses at

landing as indicated by ANSYS Analysis therefore buckling analysis was necessary. As

we had chosen pipes for constructing cabin therefore we had the added advantage in

Buckling as a circular pipe is the most efficient column section to resist buckling. This is

because it has an equal radius of gyration in all directions and it has its area distributed

as far away as possible from the centroid.

For the purpose of Analysis a long column with central loading case was considered.

Euler column equation was used and the equation was formed by using boundary

condition for the case where both ends of the column are fixed. For obtaining the Euler

column formula we began by assuming a pipe of length l loaded by a force P acting

along the centroid axis as the bar is bent a negative moment is required, and hence

Fig. showing pipe bending under compressive loads

M = -Py

Comparing with the beam deflection formula 𝑑2𝑦

𝑑𝑥2=

𝑀

𝐸𝐼 we get the equation

2

20

d y Py

dx EI

Solving the differential equation for initial conditions gives us the final buckling analysis

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2

2

4cr

EIP

l

Since the maximum compressive force to be experienced by the pipes is known we

need to keep all variables constant while finding out one variable. In our case P, E, l is

known. We decided to keep pipe thickness and internal diameter constant and find out

the minimum required outer diameter to keep the pipe from buckling. The Moment of

Inertia of pipe is given by

4 4( )

64

o id dI

Substituting in Euler Equation and solving we got the final equation:

24 4

3

16 cro i

P ld d

E

All the variables were substituted in the equation the result we got for minimum outer

diameter requirement was 1.42 inch. Hence it was verified that a pipe of outer diameter

1.5 inch, thickness 2mm would not buckle under the given loads. As with aviation rules

a Factor of Safety of 1.5 was also kept in mind while doing calculations.

Fig. showing condition for Euler equation use

To ensure that it’s an Euler column one last check is done the slenderness ratio of this

size pipe was found out to be l/k=105.3 which is greater than the limiting slenderness

ratio of 82.3l

lk

hence the use of Euler Equation is justified

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Bolt Size Calculation

In the aircraft bolts have been used to join different members. In order to prevent

rotation of members, two bolts were used on each connection point. Before bolt

selection the bolt size was determined. The three basic types of failures associated with

bolts are:

1. Crushing/bearing failure

2. Tearing failure

3. Shearing failure

Fig. (a) Crushing failure (b) tearing failure (c) shearing failure

In our case the Crushing or bearing failure is the bolt size determining factor. In order to

prevent pipes from crushing during tightening a small 3 in piece of wood was inserted in

the pipe before drilling. Hence, the bolt passed through the wood sandwiched between

aluminum pipe of thickness 2mm.

To calculate the bolt size software, MITCALC was used however two calculations were

also done manually and the answers matched with the software to validate MITCALC.

𝑃𝑐 = 𝑡. 𝑑. 𝜎𝑐

Where,

t= pipe thickness

d= diameter of the bolt

𝜎𝑐= safe permissible crushing stress

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The calculated bolt sizes were as follows:

S No Member/ connections Bolt size required

1 Cabin connections 6 mm

2 Wing - cabin attachments 8 mm

3 Control surfaces / horizontal/vertical stabilizers 4 mm

4 Truss-empennage connections 6 mm

5 Landing gear attachments 8 mm

6 Engine mount 8 mm

Lock nuts were used with bolts. These nuts are not manufactured locally and are

imported. However, at the moment lock nuts for size smaller than 6 mm were not

available. As a result we were forced to use only 6mm and 8mm bolts.

The carbon steel bolts selected were of highest available grade of 12.9. Allen Key bolts

were selected as Hex Bolts of the specified grades were not available. Bolts of length

2.0 and 2.5 inch were used.

Fig. Allen key bolts with wrenches

http://www.google.com.pk/imgres?imgurl=http://img.tradeindia.com/fp/thumb/1/478/668.jpg&imgrefurl=http://www.tradeindia.com/suppliers/allen-bolt.html&usg=__RkVg3ev4A44qXb0qxI-4fXRJveM=&h=100&w=100&sz=3&hl=en&start=42&zoom=0&tbnid=fmwzAkf9qpmhCM:&tbnh=82&tbnw=82&prev=/images?q=allen+key+bolt+with+wrench&start=40&um=1&hl=en&sa=N&tbs=isch:1&um=1&itbs=1

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POWER PLANT SELECTION

Power requirement for the aircraft was first found out from the initial design which was

coming out to be 60 hp. Although aviation engines which fulfilled our requirement were

available but financial constraints prevented us from purchasing them. Therefore a

market survey was performed to find a suitable car engine which would fulfill our needs.

However, all these engines were water-cooled and added weight. Therefore it was

decided to use air-cooled engines and hence it was decided to use VW engines another

advantage of an original VW engine is that VW claims that its engines can automatically

adjust fuel requirement with altitude. An extensive search for VW engines was carried

out and a list of VW engines available in Pakistan was made:

Production Date Displacement Engine Number

Horsepower

Dec. 1953 through July 1960 1200cc 1, 2, 3 36

August 1965 through July 1966 1300cc FO 50

August 1965 through July 1970 1300cc E low comp 37

August 1966 through July 1967 1600cc HO 53

August 1967 through July 1969 1600cc H5 low comp 50

August 1967 through July 1970 1600cc L 50

August 1969 through July 1970 1600cc H1 53

August 1969 through July 1970 1800cc B 57

August 1970 through July 1973 1300cc AC 44

August 1970 through Sept. 1971 1800cc AE 60

August 1971 through July 1973 1800cc AD 65

August 1973 through July 1975 1300cc AR 48

August 1973 through Dec 1980 1800cc AS 60

From Dec. 1974 (fuel-injected) 1800cc AJ 60

After the survey it was decided to use 1800cc VW engine or 1600cc engine produced in

1966 or 1969. For this purpose the five major VW mechanics and engine suppliers were

contacted in Karachi in May and they provided us with surety that the engine would be

provided. The estimate engine cost was Rs 20,000/-.

Car Engine name Rpm Hp Torque (Nm) Weight (lb)

Cultus G-10 5100-5500 50 65 139

Cuore ED-10 5500 48 83 N/A

Alto SS80 5500 62 87 N/A

City 1.3 L (~79 cu in) I4 i-VTEC 5500-6000 100 128 N/A

Corolla 2NZ-FE 6000 80 119 N/A

http://en.wikipedia.org/wiki/I4_engine

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Modifications:

For engine modifications local mechanics and ultra light aircraft owners were consulted.

Although help was also taken from internet but local mechanics were unfamiliar with

most of these foreign techniques as a result following modifications were finalized:

Removal of outer cowl: This would assist in more efficient cooling

Removal of gear box: As our aircraft would operate at a constant rpm therefore

gearbox use was judged irrelevant.

Fig. the removal of bulky gear box would significantly reduce weight

Double carburetor: This would allow the pistons to work differentially allowing

for a much better and efficient compression

Engine Overhaul: As these engines are about 30 years old therefore engine

overhaul is necessary to ensure optimum performance.

Replacing flywheel: The standard and heavier steel flywheel would be replaced

with Al. flywheel. This would cause significant weight reduction.

Fig: A standard flywheel and a custom made flywheel

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Reduction drive installation: In order to change the RPM of engine a reduction

drive was necessary. This modification is held till last so as to achieve the most

favorable RPM. It is suggested to use engine piston material for making pulleys

and thrust bearings should also be used as car engines are not designed to take

thrust loads and therefore in order to increase their life thrust bearings should

also be used

Fig. a reduction drive popularly used in auto conversion engines

Change Generator: To reduce weight the heavier original generator could be

replaced with a lighter generator or the generator could be fully removed and

battery can be charged on ground.

Fig. as can be seen the VW generator is quite bulky

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Alternative engine suggestions:

Despite promises made by engine suppliers an 1800cc VW engine as required was not

made available. As a result a survey was done on the internet for alternate powerplants:

Rotax engine: A second hand dual carburetor Rotax 503 engine was available in

Peshawar at a cost of Rs 2 lakh. The engine had flown for 40 hours. It provided

47 hp but at the same time its weight was 85 lbs which is almost half the

estimated weight of VW engine. Propeller specifications for Rotax engines are

directly available on the internet. For Rotax 503 carrying a weight in the range of

700-850 lbs the propeller specs are: 56x32 where 56 is the diameter and 32

degrees is the pitch.

Fig. Rotax 503

Water-cooled engines: Using water-cooled engine would increase our aircraft

weight as the whole cooling mechanism had to be added. However, as a last

resort these engines too can be considered. Increasing weight would also

increase the Hp requirement as a result cultus and Coure engine were

considered inadequate. City and Corolla engines would exceed our requirement.

This leaves Alto SS80 engine as the best available option. However this engine

was introduced in 2008 and a brand new engine costs a staggering Rs 1.8 lakh.

While a used engine costs Rs 50,000/-. For the given Rpm and new weight the

propeller was designed using javaprop® software which gave us propeller

specification of 58x26.

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PROPELLER DESIGN

Initially it was decided that a propeller would be bought directly from the market for the

ultra light aircraft. However, no ready made wood propellers were available for sale in

the market leaving us with no choice but to fabricate the propeller. A manufacturer was

contacted and he asked for the diameter and pitch of the propeller.

For propeller design a free software javaprop® was used. The software is based on

Adkins, C. N.: Design of Optimum Propellers, AIAA-83-0190. The main improvements

were:

Elimination of the small angle approximation

Elimination of Light loading approximations prevalent in the classical design

theory

An iterative scheme is introduced for accurate calculation of the vortex

displacement velocity and the flow angle distribution.

Momentum losses due to radial flow can be estimated by either the Prandtl or

Goldstein momentum loss function.

In the analysis portion "Blade Element Method" is used and it uses the same

airfoil polars as the design procedure

The influence of blade number and tip loss are taken into account by the "Prandtl

Tip-Loss Factor

Four major design criterions were kept in mind while carrying out propeller design:

THE NUMBER OF BLADES

The number of blades has a small effect on the efficiency only. Usually a propeller with

more blades will perform slightly better, as it distributes its power and thrust more

evenly in its wake. But for a given power or thrust, more blades also mean more narrow

blades with reduced chord length, so it’s difficult to manufacture.

DENSITY:

The density of the fluid has no influence on the efficiency of a propeller, but strongly

affects its size and shape. As the forces and the power are directly proportional to the

fluid density. Aircraft operating at high alt. usually need a smaller diameter then those

operating at low alt.

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PITCH

Large pitch propellers may have a good efficiency in their design point, but may run into

trouble when the have to operate at axial velocity. In this case, the blades tend to stall.

Usually the best overall propellers will have a pitch to diameter ratio in the order of 1.

DIAMETER

The propeller diameter has a big impact on performance. Usually a larger propeller will

have a higher efficiency, as it catches more incoming fluid and distributes its power and

thrust on a larger fluid volume.

Results

The software javaprop® is a very user friendly software a six step procedure was

followed to come to final result.

Number of Blades 2 blades

Diameter 4.6ft

Rpm 2700

Thrust at design condition 220 lb

Efficiency at design condition 65%

Blade angle at 75% R 25

Geometric Pitch 3.8ft

Airfoils were selected according to the most popularly used airfoils in ultralight

propellers

At r/R=0 MH 126

Fig. MH 126

MH 126, this airfoil was designed for the root section of a full size propeller. It covers a

wide angle of attack range without separation and has the required thickness for this

region.

At r/R=0.33 MH 112

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Fig. MH 112

MH 112 16.2%, this airfoil was designed to follow the root section. It covers the typically

needed range of lift coefficients for the inboard region.

At r/R=0.67 MH 114

Fig. MH 114

MH 114 13%, this airfoil is well suited for the middle part of the propeller halfway

between root and tip.

At r/R=1.0 MH 116

Fig. MH 116

MH 116 9.8%, this airfoil can be used for the tip of propellers operating at tip Mach

numbers of 0.6 and below.

Then the front and side view of propeller was found

Fig. Front and side view of the propeller

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Next the thrust graph with respect to velocity were found

Fig. Thrust vs Velocity graph

The thrust graph was then imported to MS Excel and required thrust vs available thrust

graph was made

Fig. Thrust comparison

As can be seen there is excessive thrust available. Especially at cruise Mach which is

about 0.067.

0

50

100

150

200

250

300

350

400

450

0 0.05 0.1

thru

st (

lb)

Mach

Thrust comparison

thrust available

thrust required

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Next Cl, Cd plots were made against r/R. As desired the Cl value is greater than Cd

throughout the blade diameter

Fig. Cl, Cd vs r/R

A Mach number vs r/R curve was also plotted. The Mach number throughout the

diameter remained below 0.7 Mach as desired.

Fig. Mach vs r/R

A chart showing the velocity of the slipstream of propeller was also made. Vx is the axial

velocity, V is the inflow velocity. The graph shows that the flow has accelerated by

about 80%.

Fig. Velocity variation graph

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Conclusion:

The propeller design was satisfactory as it was close to common propeller results for

ultra light aircrafts with 50hp engine as shown in the table below

Commonly used

propeller Propeller designed using javaprop®

Prop Dia. (") 52 56 59 55.2

Pitch (degrees) 28 25

Thrust (lb.) 158 230

The propeller efficiency is 65% which is less than the ideal 80%. However the efficiency

is dependent on a number of factors like HP, RPM, Diameter, Velocity etc.In our case

the HP is fixed. The rpm and diameter are also restricted with the thrust requirement

and the need to keep the tip velocity below 0.75 Mach. All these restraints have resulted

in a lower efficiency. An iterative method was used to get the best combination of

Diameter and Rpm.

Solid edge view of the designed propeller:

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Fuel system

Two types of fuel systems were considered:

Gravity fed fuel system

Pressure fed fuel system

Gravity-Feed Systems use only the force of gravity to push fuel to the engine fuel-

control mechanism. The bottom of the fuel tank must be high enough to provide

adequate pressure to the fuel-control component. This type of system is often used in

high-wing light aircraft. We have preferred gravity fed system as it is simple, does not

need any installation of pump and suits our aircraft configuration i.e. high-wing light

aircraft. However, we need to install two valves to ensure fuel supply at all time

Thus, our fuel tank has two valves to counter for every phase of flight. Fuel from these

valves merged together before entering into the engine.

Fuel System Components:

Tanks

Lines

Valves

Fuel Flow-meters

Filters and Strainers

Fuel line routing:

Fuel lines and hoses should be routed free of conflict with moving parts and should be

secured so they will not vibrate against the airframe. It should also be ensured that fuel

lines do not hinder pilot’s movement.

The fuel line from the tank to the engine must be routed continuously down hill. If this is

not maintained, air pockets or bubbles can get trapped in any high spots and severely

restrict the flow of fuel.

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Solid edge view:

Towards engine

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CHAPTER V

FABRICATION

Material Selection

Before starting fabrication it was imperative to select material which had a high strength

to weight ratio. Al 6061-T6 is the most commonly used material in ultralight aircraft

industry. However, this material is not available locally and hence an alternate Al 6063-

T6 was used in aircraft structure. A comparison of AL6061-T6 and AL 6063-T6 is given

below:

T 6063 T 6061

Density (lb / cu. in.) 0.098 0.097

Specific Gravity 2.7 2.7

Melting Point (Deg F) 1090 1150

Modulus of Elasticity Tension 10 10

Modulus of Elasticity Torsion 3.8 3

Tensile strength 42,000 psi (290 MPa) 30,000 psi (207 MPa)

Yield strength 35,000 psi (241 MPa) 25,000 psi (172 MPa)

Although AL 6061-T6 was a better option, AL 6063-T6 is used as it fulfills our

requirement.

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Tools Used

Fabrication process started with acquisition of tools. Due to unavailability of resources

expensive tooling equipment could not be purchased that’s why stress was given upon

using simple and cheap yet effective tools. A few parts needed welding or machining,

and those were farmed out to local professionals. Following tools were used during the

fabrication process.

Tools Used: Had Access to:

Flexible tape measure 5 m long Sheet Metal Shear

Carpenters square Sheet Metal Break

Water level 2 foot Drill press

Rivet gun Metal Lathe

1/4-inch Socket wrench set Metal Mill

Open end wrench set Welding torch

4-inch bench vise

Hand Drill

Center punch

Hacksaw

High Speed Electric Cutter

Electric Baby Grinder

Assorted metal files

Handheld Pipe cutter

Pop Rivet Gun

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Fabrication Process Steps

Following steps were taken in building and assembling of the designed ultra light

aircraft. Some steps like pipe marking and cutting are in series while others were

completed in parallel. it should be noted that all the fabrication work was done in

backyard of a house and therefore this project is a true home built product.

Pipe Marking and Cutting

As the standard size of aluminum pipes available from local market was 12 ft..So they

had to be cut into smaller pieces so that they can be used on different locations as

required by structural design of the aircraft. This Task was completed in two steps

1) Marking:

Marking scheme was such that a single 12 ft piece could be fully utilized in

making different size pieces and minimum material was wasted. Also the defined

nomenclature for different members was used to write Name, Length and

Diameter of each piece

2) Cutting:

Pipe Cutter was used to cut the larger pipes into smaller precise pieces which

were to be used in different parts of aircraft. Procedure followed for cutting

purpose was to hold the marked pipe in bench vise and provide a support to

keep balance then a straight cut was made which ensured dimensions to be

exact as marked on pipes.

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Fabrication of Connectors

Following connectors were obtained from market directly.

I. 3d Elbow connector

These are 3d joints which were used to join cabin pipes meeting at 90 degree

angles. Mild steel material joints were used due to availability and cost

constraints.

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II. 3d TEE connector

These connectors are used to join 04 pipes meeting at 90 degree angles

simultaneously. They are also made of Mild steel. They provide ease of

assembling as strength to highly stressed portions of cabin.

Fabrication of Custom Designed Joints and Hinges

Due to unavailability of required connectors and joint members from local market, a

number of connectors were designed, analyzed and fabricated with the help of local

professionals, given below are the details of design and fabrication of different type of

joining members

I. 90’ Clip Connectors

These types of connectors were most commonly used in fabrication of wing,

empennage and control surfaces; they were used to join any 02 pipe members

which were at right (90deg) angle to each other.

Their fabrication started with cutting a piece of aluminum plate of 6inx4inx04mm

and shaping it into desired joint with the help of soft hammering???? and sheet

bending techniques.

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II. Angled Clip Connectors

This type of connectors were used to join the main members which were inclined

to each other at some specified angle, different angle connectors were used in

different parts of cabin and empennage assembly, they were fabricated in the

same fashion as 90 deg connectors and special care was taken in making the

correct angle of joint as well as to prevent the shear of soft aluminum sheet .

III. Plate Connectors

This special type of joint was used on rear portion of the cabin; their main

purpose was to support the extruded cabin members which support rear struts

coming from wing. As this portion is highly stressed and requires stronger joints

than simple angled joint, plate joints were designed and fabricated using 04mm

thick aluminum plate which connected angled members of rear cabin to extruded

strut attachment members.

IV. Ring Connectors

Connecting Wing airfoils with the main spar was a difficult task. As these types of

connectors are unavailable in the market therefore an innovative new ring

connector was designed. For their fabrication purpose, Mild steel pipes were

found which would snugly fit on the Spar and would be having thickness

only1mm as these connectors won’t be taking much stress. Then a jig was made

that copied the exact wing dimensions. Once the jig was made exact welding

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location of the smaller pipe onto the 2in pipe was confirmed and the ring

connectors were made accordingly.

Control Surface connectors/Hinges

For deflection of flaps and control surfaces some rotating mechanism was

necessary, market survey showed that desired deflection could not be achieved

using ordinary hinges and so male female couple type hinges of different desired

sizes were designed and fabricate, for this purpose MS bars were machined and

welded over MS pipes of different required diameters, then these completed

hinges were clipped to control surfaces mounted on Horizontal and vertical tail as

well as on wing spars.

Fuselage Construction

Fuselage design consisted of two main components

1) Cabin compartment

2) 2d Truss Structure

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Both the components were fabricated separately and then joined to complete the

fuselage structure. Details of both the processes are given below.

Cabin Construction:

Cabin is a critical portion of aircraft as it gives places for attachment of wing, rear truss,

engine mounts, and landing gear ,it also houses the pilot and works as a base for

placement of control components. List of members used is given below followed by

steps taken to complete the cabin structure.

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Material used in Cabin Fabrication

Vertical Cabin Members

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SN. Members Length(in) Diameter(in) Thickness(mm) Material

1 FC 001 30 1.5 2

Al 6063

2 FC 002 30 1.5 2

3 FC 003 54 1.5 2

4 FC 004 54 1.5 2

5 FC 005 54 1.5 2

6 FC 006 54 1.5 2

Lateral Cabin Members:

SN. Members Length Diameter/Width Thickness Material

1 FCL 015 18 1.5 2

AL6063

2 FCL 016 18 1.5 2

3 FCL 017 58 1.5 2

4 FCL 018 58 1.5 2

5 FCL 019 50 1.5 2

6 FCL 020 50 1.5 2

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Horizontal and Angled Cabin members:

No. Members Length(in) Diameter(in) Thickness(mm) Material

1 FCH 007 30 1.5 2

AL6063

2 FCH 008 30 1.5 2

3 FCH 009 30 1.5 2

4 FCH 010 30 1.5 2

5 FCA 021 30 1.5 2

6 FCA 022 30 1.5 2

7 FCA 023 35 1.5 2

8 FCA 024 35 1.5 2

Landing Gear and Wing Mounts:

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S No. L/g & Wing Mounts Length(in) Width(in) Thickness(mm) Material

1 FLB01 X 02 50 2 5 AL6063

2 FLB02 X 02 50 2 5

3 FLB03 X 02 50 2 5

Steel 4 FLB04 X 02 50 2 5

5 FLB05 X 02 32 2 5

Connectors Used

S No. Connector Type Quantity Material

1 3d Elbow 4 Mild steel

2 3d Tee 2 Mild steel

3 90’ Clip 4 Al 6063

4 Angled Clip 4 Al 6063

5 Plate Joint 2 Al 6063

Other Materials Used:

S No. Part Material Length(in) Width/Dia(in) Thickness(mm)

1 Flooring Hard Plywood 50 30 20

2 Windshield Plastic Sheet,Al pipes 32,32 24,15mm 7

3 Firewall Galvanized Iron Sheet 66 54 2

Construction process:

I. Pipe Structure and Connections

The pipes were connected using Elbow, TEE and Clip joints????Reason of using

MS (Mild steel) for Tee and elbow joints was unavailability of Aluminum joints.

The joints were strengthened using wooden pieces of 3inch length and were

snugly fit inside the Aluminum pipes .This procedure has been proved effective

by previous designs fabricated worldwide.

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Rear portion joints of the cabin were specially designed and custom fabricated to

ensure strength and sleekness of this portion. All the joints were Bolted using

6mm diameter,2 and 2.5 inch length Allen Bolts. Lock nuts were used to ensure

safety and reliability.

II. Wing Mount and Landing Gear attachments:

As wing and landing gear are two main loads bearing components of the aircraft,

their attachments to the cabin needed to be strong enough to bear loads

transferred from this component.

To fulfill this requirement L beam of Aluminum and Stain less steel were used for

Wing and landing gear attachment respectively. These beams were joined to

base pipe structure using nuts and bolts.

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III. Cabin Floor , Wind shield and Cabin Firewall

Cabin floor is required for mounting of pilot seat, control stick and pedals;

it also provides a base for installment of various subsidiary components.

For construction purpose hard plywood sheet was used. After cutting the

sheet in precise dimensions it was attached to base members (pipes and

L beams) with the help of nuts bolts and washers.

Wind Shield serves the purpose of protecting the pilot from prop wash,

bird strike and foreign objects .It also gives aerodynamic shape to cabin

front portion thus reducing drag and increasing performance.

It was constructed using plastic sheet which were cut to size. Front portion

of wind shield was bended to give curvature, for this purpose aluminum

pipes of 15 mm diameter were used. Finally sheet was riveted to Cabin

Pipes.

Cabin Firewall is used for thermal insulation or the cabin thus providing a

safety wall between engine and cabin so that engine heat might not injure

the pilot or damage instruments inside cabin.

Galvanized iron sheet was used for this purpose as suggested by Daniel P

Raymer.

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2d Truss Construction

Second main portion of fuselage fabrication was construction of rear truss, it provides

connections for Empennage assembly and is structural backbone of the aircraft,

structural analysis showed adequate strength of simple 2d truss and its ease of

construction made it an attractive choice for fabrication. Given below is the list of

members used in fabricating this part followed by the procedure of construction

Material Used:

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No. MEMBERS LENGTH (in) Diameter(in) Thickness(mm)

1 FT 001 114 1.5 2

2 FT 002 116.5 1.5 2

3 FT 003 23.1 1.5 2

4 FT 004 27.6 1.5 2

5 FT 005 18.8 1.5 2

6 FT 006 28.9 1.5 2

7 FT 007 15.6 1.5 2

8 FT 008 23 1.5 2

9 FT009 10.6 1.5 2

Table 8: Fuselage Truss members

Construction Process:

Designed 2d truss was constructed using aluminum members which were joined using

4 ft long,6 in wide and 5mm thick al plates. Members were fitted in two such aluminum

plated and bolted to complete the truss ,all the joints which had drilling holes were

strengthened by using 3in long snugly fit wooden pieces. Truss also contained the

connections for empennage assembly, for tail connection smaller pieces of aluminum

plates and L beams were used.

Assembling of Fuselage:

After completion of both cabin and truss structures, they were joined, following joints

and strengthening members were used to make the fuselage rigid and to constraint it in

all degrees of freedom.

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Members Used:

No. MEMBERS LENGTH (in) Diameter(in) Thickness(mm)

1 FA001 48 0.7 2

2 FA002 48 0.7 2

3 FA 003 36 0.7 2

4 FA 004 36 0.7 2

Connector Used:

Connector Quantity Material

90’ Clip 02 Al 6063

Assembling Procedure: Assembly was joined first with the help of 90 deg clip

connectors but they were unable to take all the loads acting on truss so to strengthen

the structure members FA001 and FA002 were used which joined the cabin roof L

beams to truss while Elements FA003 and FA004 supported the truss by connecting it

to cabin base members.

Fuel Tank Installation:

Fuel tank provides the engine with required fuel, gravity feed fuel system was used as it

is simple and easy to use there is no need of special fuel pump.

Following component complete the fuel system

I. Fuel Tank

Locally available Fuel tank of 5KVA electric generator which has 25 liter fuel

capacity was used. It is fitted with fuel level indicator.

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II. Fuel Tank Supports

Two (02) MS pipes of 0.7in diameter and 1mm thickness were used to support

fuel tank, they also served the purpose of giving aerodynamic shape to rear

fuselage portion.

Wing Construction

Wing is the lift producing surface. It bears aerodynamic loads like lift and drag and also

provides resistance towards flutter and fatigue phenomenon.

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Material Required

S No. member Length Diameter Thickness material

1 W 001 13 ft 2 in 2 mm

AL 6063

2 W 002 13 ft 2 in 2 mm

3 W 003 13 ft 2 in 2 mm

4 W 004 13 ft 2 in 2 mm

5 Rib 001 40 in 2 in 2 mm

6 Rib 002 40 in 2 in 2 mm

7 Rib 003 40 in 2 in 2 mm

8 Rib 004 40 in 2 in 2 mm

9 Rib 005 40 in 2 in 2 mm

10 Rib 006 40 in 2 in 2 mm

11 Rib 007 (Qty 14) 40 in 15 mm 2 mm

12 Airfoil 001 (Qty 14) 46.2 in 15 mm 2 mm

Table 9: Wing members

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Connectors Used:

S No. Connectors Used Quantity Material

1 MS pipes 5 ft long 2 Mild Steel

2 90⁰ clip 12 AL 6063

3 Airfoil ring connectors 28 Mild Steel

4 U bolts (dia 8mm grade 12.9) 8 Carbon Steel

5 Allen bolts(dia 6mm length 2.5in grade 12.9) 36 Carbon Steel

The aircraft’s wing span was calculated to be 28ft. however; 14ft aluminum pipes were

not available in the market. As a result 13ft pipes were bought and to join then a 5ft Mild

steel pipe was snugly inserted into the two pipes. This meant that a 1.5ft MS pipe was

inserted into each Spar.

The two main spars were connected with 6 ribs of AL diameter 2 inch while 14 15mm

ribs were also attached. The positioning of the 2 in Ribs was done to accommodate the

control system pulley besides providing strength to the wing. The 2 in ribs were

connected with clip joints and were bolted. Riveting was also an option but since the

aircraft had to be transported to Risalpur therefore Bolts were given preference.

Next Step was the bending of pipe into airfoil shape. For this purpose a full scale airfoil

shape was taken from plotter. Next this shape was cut onto a card board. This template

was given to the pipe bender who accurately bent 14 pipes according to the template.

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Airfoils were connected to main wing spars using specially designed ring connectors

which gave not only ease of installation of rib but also came up as strong and light

weight solution to the problem

Pictures:

Full wing

MS pipe

Clip connector

Welded connector

U bolts

Jig

Empennage Assembly

Empennage

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Horizontal Tail

Horizontal stabilizer is an integral part of any airplane it provides stability to the aircraft

and prevents it from going into spin. As before the material required is tabulated

followed by the joints required and then the fabrication process is mentioned

MATERIAL REQUIRED

S.no member Length (in)

Diameter (in)

Thickness (mm)

material

1 HT 001 (i) 24 1.5 2

AL 6063

2 HT 001 (ii) 24 1.5 2

3 HT 002 (i) 29 1.5 2

4 HT 002 (ii) 29 1.5 2

5 HT 003 (i) 34 1.5 2

6 HT 003 (ii) 34 1.5 2

7 HT 004 114 1.5 2

8 HT 005 (i) 59 1.5 2

9 HT 005 (ii) 59 1.5 2

Table 10: Horizontal Tail members

JOINTS USED

Joints Used Qty Material

L beams ( length 6 inch ) 4 AL 6063

Plate ( thickness = 5mm) 1 AL 6063

90⁰ clip 6 AL 6063

Angle clip 6 AL 6063

Allen bolts (dia 6mm length 2 in grade 12.9) 40 Carbon Steel

MS Struts 2 Mild Steel

During Horizontal tail construction it was of utmost importance that the horizontal tail be

kept symmetric. So that during flight the aircraft won’t tilt to one side.

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At first the 90⁰ clips were attached to HT 004. Afterwards members HT001, HT002,

HT003 were cut accurately and joined with the angle clip to HT 004. During this entire

process it was necessary that the members should be kept perpendicular to HT 004 for

this reason a triangle was kept with the connecting members.

After assembling the horizontal tail a pendulum was used to confirm the symmetry of

the Horizontal tail. The pendulum pointed at the midpoint of HT 004 which confirmed

that the HT is symmetric. For final test of symmetry diagonal length of the HT was

measured from both sides and was found equal.

The horizontal tail was connected by sandwiching the horizontal stabilizer between the

L beam and plate. It was ensured that the centre of horizontal stabilizer matched the

centre of fuselage truss structure.

As these connections were not sufficient therefore they were enforced by attaching

struts to the horizontal stabilizer.

Pics

Uncut pipes at 90

Full HT

Connectors

Struts

Pendulum hanging

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VERTICAL STABILIZER

Vertical stabilizer is an integral part of any airplane it provides stability to the aircraft and

prevents it from going into spin.

MATERIAL REQUIRED

S.no member Length (in) Diameter (in) Thickness (mm) material

1 VT 001 18 1.5 2

AL 6063

2 VT 002 22 1.5 2

3 VT 003 26 1.5 2

4 VT 004 64 1.5 2

5 VT 005 49.5 1.5 2

Table 11: Vertical Tail members

JOINTS USED


L beams ( length 8 inch ) 2 AL 6063

Plate ( thickness = 5mm) 2 AL 6063

90⁰ clip 4 AL 6063



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Horizontal tail fabrication was relatively easy first we joined members with 90 clips at a

distance of 18 in from centre to centre. Afterwards the pipes were cut to size and joined

using angle connectors. It should be remembered that the members were cut at an

angle to fit them in the vertical stabilizer.

The vertical tail was attached to the truss structure by using two plates of dimension 6x6

in front and by using a 90 connector and 2 L beams at the rear.

Control Surfaces

Four control surfaces were to be fabricated: elevators, ailerons, flaps and rudder. The

construction of control surfaces was relatively easy.

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Ailerons

MATERIAL REQUIRED


Diameter (in)

Thickness (mm)

material

1 AL 001 (i) 60 1.5 2

AL 6063

2 AL 001 (ii) 60 1.5 2

3 AL 002 (i) 60 1.5 2

4 AL 002 (ii) 60 1.5 2

5 AL 003 (Qty 6) 18 1.5 2

JOINTS USED


90⁰ clip 12 AL 6063


The construction of Ailerons was straight forward the Connector’s centre was placed at

a distance of 30in from each other. While joining the AL 003 members to AL 001

members it is suggested that first the connector should be placed at the marked position

and checked for its perpendicularity with member AL 001. Even a slight mishap in this

case would render the whole Aileron useless. Once the AL 003 members have been

joined with AL 001 the members should then be joined with AL 002 members if the

procedure is correctly followed the AL 003 members should automatically come

perpendicular to AL 002.

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Flaps

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MATERIAL REQUIRED

S.no Member Length (in)

Diameter (in)

Thickness (mm)

material

1 FL 001 90 1.5 2

AL 6063

2 FL 002 90 1.5 2

3 FL 003 90 1.5 2

4 FL 004 90 1.5 2

5 FL 005 (Qty 8) 18 1.5 2

JOINTS USED


90⁰ clip 16 AL 6063

AL connecting pipe (ID 1.5in Length 5ft) 3 AL 6063


For flaps it was necessary that the flap leading edge be a single piece as it was to be

connected with control system. This would ensure equal and timed movement of the

entire flap. To achieve this Aluminum pipe was bought whose internal diameter was 1.5

inch and hence the two FL 001 members would snugly fit into the pipe both the

members were inserted 1 ft into the Aluminum pipe.

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Once through with the leading edge of the flap rest of the procedure is same with the

connectors joined at a distance of 30 inch from each other and perpendicular to the FL

001 and FL 002.

Rudder

MATERIAL REQUIRED


Diameter (in)

Thickness (mm)

material

1 RR 001 9 1.5 2

AL 6063

2 RR 002 15 1.5 2

3 RR 003 21 1.5 2

4 RR 004 60 1.5 2

5 VT 005 61.34 1.5 2

JOINTS USED


90⁰ clip 3 AL 6063



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Rudder fabrication is similar to Vertical stabilizer construction the 90⁰ clips are placed at

a distance of 29 inch from centre to centre. Rest it should be kept in mind that a triangle

should be used while fixing the 90⁰ clip.

Elevator

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MATERIAL REQUIRED


Diameter (in)

Thickness (mm)

material

1 EL 001 (Qty 2) 11 1.5 2

AL 6063

2 EL 002 (Qty 2) 15.25 1.5 2

3 EL 003 (Qty 2) 32 1.5 2

4 EL 004 (Qty 2) 21 1.5 2

5 EL 005 (Qty 2) 50 1.5 2

JOINTS USED


90⁰ clip 4 AL 6063

Angle clip (12⁰) 4 AL 6063

Angle clip (30⁰) 4 AL 6063


Since the hinges designed for the movement of control surfaces did not allow for a

single elevator therefore the two elevators had to be fabricated. The first 90⁰ clip was

positioned at the end while the second was placed at a distance of 20 inch from the

centre of the first.

Although the lengths stated above are accurate but it is highly suggested that the angle

pipes be kept a bit longer and later cut down after fixing the members. Again it is

imperative that both the elevators be kept symmetric.

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Fabrication of Landing Gear

Landing gear plays a vital role during taxing, taking off and most important at landing, it

bears the ground loads and impact loads; landing gear was designed following 3g

impact regulation of Federal Aviation Authority (FAA).So it had to be one of strongest

components of the aircraft. The design included a shock absorbing system which

reduces the impact loads on aircraft.

Coming to the fabrication of the finally selected landing gear design following members

were used followed by the steps taken in fabrication.

Members of Landing Gear:

No. Component LENGTH (in) Diameter(in) Thickness(mm) Material

1 LR001 45 Mild Steel 2 LR002 45

3 LR003 50

4 LR004 50

Table 12: Landing Gear members

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Tires Used:

No. Tire Diameter(in) Thickness(In) Material Feature

1 TR01 14 4 Rubber tire and tube

Inbuilt brake system

2 TR02 14 4

3 TR03 6 3 All Hard Rubber Lock mechanism

Shock absorber:

No. Tire Spring Constant Damping Coefficient Rated load(lb)

1 SA01 NA NA 800

2 SA02

Fabrication Process:

I. Main landing Gear

a) Connecting Members: In main landing gear design four (04) connecting

members (LR001,LR002,LR003,LR004) were used which bear and

transferred the impact load to base cabin members(FLB05,06,07,08),they

were cut into size and two (LR003,LR004) of them were bent to take the

desired shape as per design. These rods were connecting to cabin base

by pin joints so that they can move in 02 degree of freedom, the angle

between the aircraft centerlines and front members was set as 27degrees

which is in conjunction with FAR regulation of minimum of 25degree of

landing gear angle.

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b) Tires: Two tires were selected for main landing gear design, for this

purpose local market survey was done to find the closest match to our

designed landing gear tires. Finally tires were obtained which met the

design specifications and also incorporated the feature of in build shoe

brake system for ground brake purposes These tires were welded to

landing gear rods to ensure strong joints.

c) Shock Absorbers: As mentioned previously, two shock absorbers were

selected using standard testing procedure. They were attached to

members (LR001, LR002) with the help of custom made welded holder

which acted as pin joints for shock absorbers.

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II. Rear Landing Gear

As the aircraft has a tail dragger configuration third landing was placed at rear tail

section of aircraft. Rear landing gear is not designed to bear heavy impact loads

rather its main purpose is to help in movement and directional control while at

ground. An all rubber tire has been used for rear landing gear which is fixed into

lower end of vertical tail rear spar. Rear landing gear also features a locking

mechanism which enables us to lock its rotational motion as well as its

translational motion or both.

Final Assembling of Aircraft:

After completion of each component separately they were attached as a final assembly

consisting of cabin fuselage truss, wing and tail assemblies and in the end control

surfaces and landing gears were attached to give the aircraft final shape. All the

components fit perfectly and were bolted, hinged and welded to make the whole

structure as one piece.

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Fabrication of Aircraft Skin Covering:

Aircraft covering is necessary to complete the structure such that it could be installed

with propulsion system, covering of wing, empennage and control surface is necessary

for production of aerodynamic forces on these parts while fuselage covering gives the

aircraft aerodynamic shape hence reducing drag and increasing efficiency.

Material Used:

Parachute material having dimensions 40 x 05 yards was used for fabricating

skin of the aircraft.

Rivets of 3mm diameter and 12 mm thickness were used to fix the stitched skin

cloth to the aircraft

Fabrication Process:

First of all clothing material was cut to dimension of wing, empennage and control

surfaces then it was stitched to give a tight fitting so that there won’t be any slack or

loose surface left. Last step was to fix the skin with aircraft components with the help of

riveting process. Thus giving the aircraft a complete aerodynamic shape.

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Solid Edge Figures

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CONCLUSION

Full detail design of the aircraft was carried out which would ensure the integrity of the

aircraft. Every possible aspect of design was taken care off from the selection of Bolt

size to installation of aircraft engine. Due to shortage of time the fabrication could not be

completed. However, the most important part of fabrication, aircraft structure, was fully

completed. Material for aircraft covering and controls has also been acquired while the

propeller has also been fabricated. The entire fabrication process was carried out using

the most economical resources and it was proven that an aircraft can be fabricated in

Pakistan using local resources and that too at an extremely low cost.

It is suggested that this project be continued. Inshallah by next semester this project

would be fully completed and CAE would be able to claim that its students have

successfully designed and flown their own aircraft.

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Design and Fabrication of Ultra Light Aircraft

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