Kisyan Light Jet THESIS ( University of Brighton)

K I S Y A N P E K A S A R O U F 1 | P a g e

School of Computing, Engineering and Mathematics

Division of Engineering and Product Design

DESIGN OF LIGHT JET

KISYAN PEKASAROUF

XE337 GROUP PROJECT

DR.NICHOLAS MICHE

DR.STEVEN BEGG

9th

May 2014

Bachelor of Science (Honours) in Mechanical and Manufacturing Engineering


I hereby certify that the attached is my own work except where otherwise indicated.

I have identified my sources of information; in particular I have put in quotation

marks any passages that have been quoted word-for-word, and identified their

origins.

DISCLAIMER

Signed

Date....


ABSTRACT

The project describes the major features of Skysonic jet designed with a

tadpole-like configuration that has performance requirements for light business jet purposes.

It is therefore desirable to be able to effectively investigate and analyse solutions from a

variety point of logical view, weighing together the results and conclusions. Solidworks

Computer Aided Drawing system and Foil SIM (III) software were used throughout the

analysis process to reflect a great extent of output considering the design parameters. The

experimental engine DGEN 390 specifications were used throughout the project to set the

limitations in designing Skysonic jet. The preliminary results which were adapted from the

Maverick Smartjet such as maximum cruising altitudes and cruising speeds were optimized

with a proper engineering approach in order to show the correct process sequence. Each

member was delegated with specific task and the output of the team effort on designing,

optimizing and finalising the process has been clearly shown with chronological order. The

project also discussed on pending further works to be done in future in order to improve the

quality of results outcome. Skysonic jet has produced a creative design where it has

successfully met all the conditions applied corresponding to vast resources and references.


TABLE OF CONTENTS

LIST OF FIGURES ................................................................................................................................ 6

LIST OF TABLES .................................................................................................................................. 7

NOMENCLATURE ............................................................................................................................... 8

INTRODUCTION .................................................................................................................................. 9

BACKGROUND .................................................................................................................................. 10

THEORY .............................................................................................................................................. 11

1.1 Aerofoil Decision Matrix ................................................................................................... 13

1.2 Liebeck LNV109A Aerofoil ............................................................................................. 14

1.3 Aspect ratio ........................................................................................................................ 15

1.4 Induced Drag ...................................................................................................................... 17

STRUCTURES ..................................................................................................................................... 17

2.1 Wing Component Mass Distribution ................................................................................. 18

2.2 Calculation of the mass using the Solidworks software .................................................... 18

TOOLS .................................................................................................................................................. 20

3.1 Foil SIM (III) version outlooks tested at 12,000 feet (3658m) with 250 knots (463kph) . 21

3.2 Foil SIM (III) version outlooks tested at 22,000 feet(6706m) and 300knots (555.6kph) .. 21

3.3 Solidworks Stress Analysis ................................................................................................ 23

3.4 Process flow ....................................................................................................................... 24

RESULTS AND DISCUSSION ........................................................................................................... 27

4.1 Distribution of data at 12,000 feet cruise with 250knots. .................................................. 27

4.2 Distribution of data at 22,000 feet cruise with 300knots. .................................................. 27

4.3 Accuracy of Foil Sim (III) software and manual calculations ........................................... 28

4.4 Induced drag corresponding to Aspect ratio ...................................................................... 29

4.5 Stress Analysis Results and Discussion ............................................................................. 30

CONCLUSION ..................................................................................................................................... 32

FUTURE TASKS ................................................................................................................................. 33

CONTEXT REFERENCES .................................................................................................................. 34

FIGURES REFERENCES .................................................................................................................... 36

APPENDIX A ....................................................................................................................................... 38

APPENDIX B ....................................................................................................................................... 39


APPENDIX C ....................................................................................................................................... 41

APPENDIX D ....................................................................................................................................... 42

APPENDIX E & F ................................................................................................................................ 44

APPENDIX G ....................................................................................................................................... 46

APPENDIX H ....................................................................................................................................... 47

APPENDIX I ........................................................................................................................................ 50

APPENDIX J ........................................................................................................................................ 51

APPENDIX K ....................................................................................................................................... 55

APPENDIX L ....................................................................................................................................... 56

APPENDIX M ...................................................................................................................................... 57


LIST OF FIGURES

Figure 1a & 1b : Angle of Attack ......................................................................................... 11

Figure 1c & 1d : Pressure distribution around aerofoil............................................................ 11

Figure 1e : Lift, Drag and Moment of Typical Aerofoil .......................................................... 12

Figure 1f : Chamber and Symmetrical Aerofoil ...................................................................... 12

Figure 1.1a : LNV109A Lift and Drag graphs ......................................................................... 14

Figure 1.3 : Vortex flow on the wing ....................................................................................... 15

Figure 1.3a : Type of wings ..................................................................................................... 15

Figure 1.3b : Typical aspect ratio graphs ................................................................................. 15

Figure 2a : Internal and External Wing structure..................................................................... 17

Figure 2b : Solidworks software drawing of wing ................................................................... 18

Figure 2c : 3D prototype of Skysonic wing ............................................................................. 18

Figure 2.2a : Drawing of Half wing ......................................................................................... 18

Figure 2.2b : Material properties ............................................................................................. 19

Figure 2.2c & 2.2d : Mass properties ...................................................................................... 19

Figure 3a & 3b & 3c : Foil SIM setup ..................................................................................... 20

Figure 3.1a & 3.1b & 3.1c & 3.1d : Foil SIM outlook at 12000 feet (3658m). ....................... 21

Figure 3.2a & 3.2b & 3.2c & 3.2d : Foil SIM outlook at 22000 feet (6706m) ....................... 21

Figure 3.4a : Whole wing drawing using Solidworks.............................................................. 24

Figure 3.4b & 3.4c : Fixture application .................................................................................. 25

Figure 3.4d :Load application. ................................................................................................. 25

Figure 3.4e: Material selection. ............................................................................................... 25

Figure 3.4f : Meshing selection ............................................................................................... 26

Figure 3.4g & 3.4h & 3.4i : Stress analysis diagrams ............................................................. 26

Figure 4.5a : Von Mises Stress versus Altitude graph ............................................................. 31


LIST OF TABLES

Table 1.1a : Aerofoil Selection decision matrix ...................................................................... 13

Table 1.3a : Aspect ratio of various aircraft ............................................................................ 16

Table 1.3b : Advantages and Disadvantages of different aspect ratio range ........................... 16

Table 2.1a : Weight distribution of Skysonic wing ................................................................. 18

Table 3.3a : Advantage and Disadvantage of Solidworks software ........................................ 23

Table 4.1a : 12000 feet (3658m) parameters breakdown ........................................................ 27

Table 4.2a : 22000 feet (6706m) parameters breakdown ........................................................ 27

Table 4.3a & 4.3b :Accuracy of Foil SIM ............................................................................... 28

Table 4.4a : Data distribution of Induced drag ........................................................................ 29

Table 4.5a : Mission of Stress Analysis ................................................................................... 30

Table 4.5b : Results of Aluminium Alloy 7075-T6 test .......................................................... 30

Table 4.5c : Result of Graphite test ......................................................................................... 30

Table 4.5d : Justification of Stress Analysis Done .................................................................. 31


NOMENCLATURE

3D

AOA

L

D

Re

Millimetre

Meter

Nautical Miles

Kilometre per hour

Kilometre

Kilogram

Square meter

Pi

Newton

Velocity

Mega Pascal

Density

Wingspan

Chord length

Kinematic viscosity

Wing area

Aspect ratio

Average aspect ratio

Three dimension

Angle of Attack

Lift

Drag

Lift Coefficient

Maximum Lift Coefficient

Drag Coefficient

Minimum Drag Coefficient

Lift and Drag ratio

Reynolds Number


INTRODUCTION

The Aero Group 2 which consist of myself, Kisyan Pekasarouf and members Konstantinos

Karvelis, Prevind Jagadesan, Rajendra Kumar and Ali Syed have designed a 1 pilot business jet,

capable of transporting 3 passengers and their baggage with the range of 800 nm (1481.6 km) ,

cruise speed of 250knots (463kph) and 300knots (555.6kph) for 12,000 feet(3658m) and 22,000

feet(6706m) altitude respectively. We have planned for specific situations by analysing the

performances of the certain parts of the aircraft relating to DGEN 390 specifications [1]. Since the

DGEN 390 is an experimental engine, never used in any aircraft before, the aim to design a

business jet with logical parameters and computer drawings were carefully handled. Each part and

dimensions of the Skysonic jet have been designed using the correct tools so that it fulfils the

objective of our project. Plus, our aircraft uses specific material for each individual part in an effort

both to keep the weight down and to help examine the feasibility of designing future aircraft in a

similar manner. The task which was delegated to me in this project is research on wing section.

Below table shows all my findings on Skysonic jet wing plus the general specification of our

aircraft.

GENERAL Aircraft purpose Light business jet

GENERAL Crew 1 Pilot

GENERAL Capacity 1 Pilot + 3 Passengers

GENERAL Empty Weight 850kg

GENERAL Useful Weight 934kg

GENERAL Maximum Take-off Weight 1784kg

GENERAL Power plant 2 x DGEN 390 Engine (330daN

max thrust each)

GENERAL Range 800 nm

WING Wingspan 8.24m

WING Aerofoil LNV109A

WING Wing tip and root 0.45m

WING Taper ratio 1 to 1

WING Wing Area 3.708

WING Wing aspect ratio 18.31

GENERAL Minimum cruising speed with

altitude 250 knots with 12,000 feet [1]

GENERAL Maximum cruising speed with

altitude 300 knots with 22,000 feet

GENERAL Fuel Capacity 650 litres

GENERAL Centre of Gravity position 3.66m from nose


BACKGROUND

MAVERICK SMART JET SPECIFICATIONS SKYSONIC JET

Light business jet Aircraft purpose Light business jet

2 Pilot Crew 1 Pilot

2 Pilot + 3 Passengers Capacity 1 Pilot + 3 Passengers

975kg Empty Weight 850kg

912kg Useful Load 934kg

1887kg Maximum Take-off

Weight 1784kg

2 X Pratt Witney JT-15 -5

( 6595N max thrust each) [4] Power plant

2 x DGEN 390 Engine

(330daN max thrust each)

1250nm Range 800 nm

10.30m Wingspan 8.24m

NACA 65(2) 215 [26] Aerofoil LNV109A

Wing tip = 0.9m , Wing root =

2.2m, Average = 0.65m Wing tip and root 0.45m

1 to 2 Taper ratio 1 to 1

6.695 Wing Area 3.708

15.85 Wing aspect ratio 18.31

290 knots at 22,000 feet [2] Maximum cruising speed

with altitude 300 knots at 22,000 feet

957 litres [2] Fuel Capacity 650 litres

We considered Maverick Smart Jet as a basis for our design .Below shown

the optimization of Skysonic jet from Maverick Smartjet [2][3] :

Both aircrafts are designed for light business jet and having maximum cruising altitude

of 22,000 feet (6706m).All the wing dimensions (wingspan, wing tip and root, wing area and

wing aspect ratio) matches each other however they are reduced by 20% since DGEN 390

engine could only provide half of the Pratt Witney thrust. At the initial stage of the design, all

the members were allocated with specific tasks. During the whole process of designing,

optimizing and finalizing, the tasks allocations slightly changes for each member as we have

re-evaluated the Gantt chart due to the time limits and lack of resources. Hence for further

detailed view of Gantt chart, refer Appendix M.


THEORY

Any section of the wing cut by a plane parallel to the aircraft xz plane is called an

airfoil[6].An aerofoil-shaped body moved through the air will vary the static pressure on top

surface and on bottom surface of the aerofoil [6]. If the mean chamber line in a straight, the

aerofoil is referred to as symmetric aerofoil, otherwise it is called cambered aerofoil[6].The

camber aerofoil is positive,upper surface static pressure is less than ambient pressure,while

the lower surface static pressure is higher than ambient pressure[6].This is due to higher

speed at upper surface and lower speed at lower surface of the aerofoil[6].Refer Figure 1.1

and 1.2 below.For this aerofoil terms definition, refer Appendix B.

As the aerofoil angle of attack increases, the pressure difference between upper and

lower surface will be higher.[Figure 1a]

As the aerofoil angle of attack decreases, the pressure difference between upper and

lower surface will be lower.[Figure 1b]

Figure 1c and 1d showing the pressure distribution around the aerofoil.Blue indicates the positive lift

on the upper surface of aerofoil and red indicates negative lift on the bottom surface of aerofoil.

The lift due to the angle of attack normally acts 25% of the chord line of an aerofoil.It is

called quarter chord point[6].Since the lift acting at this point, hence we will call this as

aerodynamic centre (ac)[6].The aerodynamic force from this point is divided into two

components :

Lift is equals to aerodynamic force perpendicular to relative wind[6].

Drag is equals to aerodynamic force parallel to relative wind[6].

Figure 1a Figure 1b

Figure 1c Figure 1d


Figure 1e showing the lift,drag moment of a typical aerofoil.

The aerofoil section lift,drag and pitching moment are defined in non-dimensional form in

below equations [5]:

Section Lift Coefficient :

[5]

Section Drag Coefficient :

[5]

Section Moment Coefficient :

[5]

Figure 1f showing the plot of lift coefficient and drag coefficient of

typical chamber and symmetrical aerofoil.

This section explains the relationship between positive camber, symmetric and negative

chamber aerofoils. Above theory diagram shows that the typical positive chamber aerofoils

having the highest lift coefficient corresponding to the angle of attack. However, positive

chamber aerofoil stalls early and produces more drag than symmetric aerofoils. Hence, the

drag coefficient of cambered aerofoil produced will be also high since its coefficient of lift is

much higher than symmetric.

Figure 1e

Figure 1f


1.1 Aerofoil Decision Matrix

After a simple analysis done for the aerofoil selection, above aerofoils parameters were

gathered and placed on Table 1.1a [7].On this table (1.1a) aerofoils are simply scored by

numbers 1,2, and 3.Then total is entered at the bottom of the table as shown above.There are

two NACA aerofoils of 6 digits and 4 digits,one aerofoil of Lockheed and another one from

Liebeck aerofoil.The (2)thickness ratio of NACA 65(2)-215 is leading the factor because it

has much great thickness to allow the smooth airflow around the curve while LNV109a and

Lockheed L-188 comes next to it since the it is not a symmetric aerofoil.The least score goes

to NACA 0006 and this aerofoil will experience less airflow to generate lift on top of it.The

best aerofoil for (3)lift coefficients (when angle of attack is zero) factor was rated based on

the highest value that an aerofoil can achieve.In this case , LNV109a stays the top, followed

by others.Since the value is higher, it produces more lift (broad curvature) along the both x

and y axis.Aerofoil with (4)maximum chamber contributes more in lift generation because

of the curvature shape,which has the maximum chamber at 25% of the chordline of the

aerofoil.However,aerofoil 1 and 3 could not score well despite of its symmetrical shape(less

curvature at 25% of the chordline of the aerofoil).NACA 0006 scores the least since the

thickness of the aerofoil is thinner.There will be no lift produced if the lift coefficient is zero

even though it has a specific angle of attack(5).Larger negative values have the advantage of

Number Parameters Aerofoil 1 Aerofoil 2 Aerofoil 3 Aerofoil 4 SCORE

1 Name of aerofoil Lockheed L-188 [8]

LNV109a [9] NACA 65(2)-

215 [10] NACA 0006

[11] Aerofoil

1 Aerofoil

2 Aerofoil

3 Aerofoil

4

2 Thickness ratio 12% 13% 15% 6% 2 2 3 1

3 Coefficient Lift when

AOA = 0 0.3 0.5 0.4 0.0 1 3 2 1

4 Maximum Chamber (%) 2.7 6.0 1.1 0.0 1 3 1 1

5 AOA when Coefficient

lift =0

negative

2.5 negative 4 negative 1 0 2 3 1 1

6 Maximum coefficient of

lift 1.262 1.505 1.098 0.616 2 3 2 1

7 Angle of Attack at Cl

max 9 15 15 7 1 3 3 2

8 Stall characteristics

3 1 2 3

9 Drag coefficient

minimum 0.015 0.030 0.017 0.015 3 1 2 3

10 CL/CD max 48.062 21.635 38.478 21.415 3 2 2 2

18 21 18 15 Table 1.1a


producing the lift in a very long range such as from -4 to 16 rather than 0 to 7.In this factor

,LNV109a scores the highest.The maximum coefficient of lift (6) is the final limit of an aerofoil to

produce lift and the stall will initiate at this point.Referring to this factor Aerofoil 2 scores the

maximum because it delays stall (at 15) unlike other aerofoils which stall early. Aerofoil 4

experiencing the most earliest stall (7) with maximum lift coefficient of 0.6 since there's no curvature

(flat) around the shape and this aerofoil mostly applicable in supersonic aircrafts.The stall

characteristics(8) usually splitted into three; gentle,abrupt or medium and based on the above table,

Aerofoil 2 having a abrupt flow since its large angle of attack ,compare to other aerofoils which have

a small angle of attack.Since Aerofoil 4 do not produce a larger tilt as Aerofoil 2, a smooth separation

flow will be generated there.Next on the drag coefficient(9), the least value will be preferable and in

this case, Aerofoil 1 and 4 scores the highest.However,in real situation we cannot expect the lift to be

inversely proportional to drag because when the lift increases the drag increases too.Since Aerofoil 1

and 4 obtained the least drag and it produces less lift, we still cannot conclude both these aerofoils are

the suitable one.The ratio of coefficient lift and coefficient of drag are called as lift drag ratio(10).In

this category , Aerofoil 1 score the highest and these aerofoils are often used in high lift aircraft with

large payloads and long range travelling distance .However I still considering the remaining aerofoils

since my target is for the subsonic speed aerofoils with small number of passengers and short range of

travelling distance.Since Aerofoil 2, LNV109A obtained the highest score, hence I will conclude this

will be my final aerofoil.

1.2 Liebeck LNV109A Aerofoil

The unconventional shaped aerofoil shown in diagram is named after Dr.Robert Liebeck,who was

the first to try and find out what sort aerofoil-shaped geometry would yield the highest possible

Cl max[12].By using the lift coefficent at zero degree angle of attack(cruising position)[13] ,I

have tested LNV109A aerofoil in the Foil SIM (III) student package of NASA[14] to obtain

lift,drag, and necessary graph parameters.Diagram 1 and 2 shows the graph curve of aerofoil at

6% chamber and 13% thickness. Plus, parameters of altitude,speed, and reynolds number are

obtained from the FoilSIM(III) Software .

Angle of Attack

Lift

Co

eff

icie

nt

(Cl)

Dra

g C

oe

ffic

ien

t (C

d)

Angle of Attack

Figure 1.1a

Figure 1.1a shows the airfoiltools generated lift coefficient and

drag coefficient graph with 6% chamber and 13% thickness


All the lift generating components on aircraft such as engine and wing will contribute their

forces to fly the aircraft.Hence,the mission of our wing is how much of lift is needed to carry

the 17840 N aircraft at least in crusing position (12000 feet(3658m) and 22000 feet( 6706m)

)where the angle of attack of the aerofoil stays 0[13].Since the DGEN 390 engine producing

1400N(Refer Appendix A) of thrust, hence our Skysonic wing should give lift at least 92.2%

of 17840 N ( Refer Appendix C) to balance the aircraft weight .This will be my mission to be

achieve.

1.3 Aspect ratio

The aspect ratio is defined as the ratio between wingspan and area of the Skysonic aircraft

wing.Below shown the equation to find the aspect ratio.

=

[15]

When a wing is generating lift, it has a reduced pressure on the upper surface and an

increased pressure on the lower surface[16].The air would like to escape from the bottom

of the wing,moving to the top as shown in the diagram below[16].Therefore , air will escape

around the wing tip[16].Air escaping around the wing tip lowers the pressure difference

between the upper and lower surfaces.This certainly reduces the lift near the tip.Plus, the air

flowing around the tip flows in a circular path when seen from the front, and in effect pushes

down on the wing[16].This circular or vortex flow pattern continues downstream behind

the wing[16].Hence, the amount of the wing affected by the tip vortex is less for a high

aspect ratio wing than for a low aspect ratio wing .

Figure 1.3a shows the the vortex flow when airflow passes the upper and lower surface of the wing

Upper surface flow

Lower surface flow Tip vortex

Figure 1.3

Figure 1.3a Figure 1.3b


Table 1.3b shows the advantages and disadvantages of aspect ratio with given ranges.

The aspect ratio of Skysonic jet is 18.31 as calculated (Refer Appendix C) and it falls on fourth

category as a glider wing. Next, I will relate the aspect ratio to the induced drag to justify my

selection.

Aspect ratio Type of aircraft Average value

3 to 4 Military fighters 3.5

5 to 12 GA Aircraft 8.5

7 to 10 Commercial Jetliners 8.5

10 to 51 Sailplane (Glider) 30.5

Aspect

Ratio Advantages Disadvantages

3 to 4

This range will have high stall angle of attack causing the

lift generation very low. This range aircrafts are very low

structural weight and produces high flutter speed. This is

very suitable for military aircrafts since they only target

for speed, not lift.

This is inefficient because of the high

induced drag produced because

aspect ratio is always inversely

proportional to induced drag.

Generation of shallow lift coefficient

and low maximum lift coefficient

will the other minuses.

5 to 12

This range has a good roll response because of slight long

shape and also produces high flutter speed .The GA

aircraft will have a limited yaw since this aircraft still

aiming for lift unlike the military, aiming for speed.

This range aircrafts also inefficient

for long range design because of low

range speeds .Besides, it relatively

produces a high induced drag since it

is aiming for good lift generation.

7 to 10

The commercial jetliners have a good balance between low

induce drag and roll response and a smooth glide

characteristics since it travels for long ranges and target

for high lift.

It has a slight steep maximum lift

coefficient (larger change in Cl with

small changes of angle of attack).

10 to 51

This category mostly applicable to sailplane or gliders

which owns a low induced drag values because of huge

wingspan and a great glide characteristics since these

aircrafts are very light in weight. Next, producing a steep

Cl and maximum coefficient lift will be the plus points

because of large change in Cl with small changes in angle

of attack.

Since it has a very steep Cl, these

range aircraft wings could not able to

adjust the angle of attack in a large

range. High roll damping and higher

adverse yaw will be also the minuses

of this category.

There are two typical different lengths of wings which are shown in Figure 1.3a above. The longer

wingspan (glider) produces a high maximum lift coefficient (steep) while the shortest one (normal

wing) produces low maximum lift coefficient (shallow) (shown in Figure 1.3b). The longer the

wing, the slower the air moves from bottom surface of the wing to upper surface of the wing,

hence induced drag produced will be lower. On the right graph shown, the stalling angle of glider

wing is much earlier than the normal wing.

Table 1.3a shows the aspect ratio and average values of different type of aircraft

Table 1.3a [17]

Table 1.3b [17]


1.4 Induced Drag

Since the wing tip vortices produce a swirling airflow behind the wing, this component also

produces some drag, also named as induced drag. Normally, the vortices are always starts

strongest at tip, then weaker to the root and this will cause the effective of angle of attack to

be reduced. Most importantly, it also called as drag due to lift because it only occurs on

finite, lifting wings and the magnitude of the drag depends on the lift of the wing [18]. By

relating the aspect ratio with induced drag, I can justify and conclude my selection. (Refer

Results and Discussion)

STRUCTURES

Rear

spar

Trailing edge

assembly

Rib

Stringers Middle spar

Wing root rib

Front spar

Leading edge

assembly

Lightening

holes

Figure 2a shows the internal and external components of wing structure

Figure 2a

This section explains in detail on the internal structures included in Skysonics wing. All internal

and external parts are constructed using the Solidworks software. The dimensions and functions

[19] of each part is explained briefly Appendix D and E respectively. When the position of rear

spar is being decided, I had conversation with Konstantinos to determine the percentage of wing

chord needed for the aileron system. We decided to give 40% of the wing chord to satisfy the

space required for aileron mechanism and aileron .Hence, we decided to provide 20% of the wing

chord for the placement of aileron mechanism and 20% for the aileron chord.


Figures 2a and 2b above show the Solidworks software drawing and 3D

prototype printing of Skysonic cross section wing respectively.

2.1 Wing Component Mass Distribution

Table 2.1a shows the weight distribution of Skysonic wing

Quantity Parts Mass (kg)

Single unit (kg) Total (kg)

1 Front I beam spar 6.69 6.69

1 Rear I beam spar 4.00 4.00

1 Middle I- beam spar 10.77 10.77

2 Wing root and tip rib 1.39 2.78

6 Wing ribs 1.48 8.88

1 Leading edge assembly 10.47 10.47

1 Trailing edge assembly 20.03 20.03

4 Stringers on top rib surface 1.29 5.16

4 Stringers on below rib surface 0.89 3.56

1 Wing skin 91.55 91.55

Subtotal = 163.89

The total weight of wing on port and starboard side is :

163.89kg X 2 = 327.80kg

2.2 Calculation of the mass using the Solidworks software

i) The part of half wing is drawn as shown in Figure 2.2a

Figure 2b Figure 2c

Table 2.1a

Figure 2.2a


ii) The Aluminium Alloy 7075-T6 is choosen and its material properties is shown in the

circle below.(Figure 2.2b)

iii) The mass properties tab at the top left of the Solidworks software is selected (Figure 2.2c)

iv) Next, mass properties box will appear and the required details will be shown here(Figure

2.2d).

In Solidworks we will not able to calculate all the masses in one flow. However, I have to calculate it

separately for each part and compare the masses with the whole model wing. For example, in whole

wing model it shows 327.80 kg while when calculated separately obtained 326.6kg.Hence, this will be

the error of mass calculation observed in Solidworks software.

Mass of wing shown:

163.30 kg

163.30 X 2 = 326.6 kg

Figure 2.2b

Figure 2.2c

Figure 2.2d


Figure 3c

TOOLS

Foil Sim is a simulation software which is available in NASA website.It give guidance for

students to determine the lift,drag, angle of attack and other parameters with the software

calculations.Since it is a software calculation based parameters, there will be some errors in

values obtained.In order to rectify and reconfirm these values, I have done the manual

calculation below, and compared the differences or accuracy between manual and software

calculation.Hence, I have shown few steps how to set up the simulation process.

There are few tabs available on the top right of the programme .However

tabs which is necessasry for this experiment to use includes shape of

aerofoil, flight,select data(which provides graph parameters),and size of

wing.

Setting up the software in Metric units (Figure 3a)

Setting the aerofoils angle of attack,camber and thickness ratio(Figure 3b)

Setting the chord length, wing span and area of the wing(Figure 3c)

Figure 3a

Figure 3b


3.1 Foil SIM (III) version outlooks tested at 12,000 feet (3658m) with 250 knots (463kph)

3.2 Foil SIM (III) version outlooks tested at 22,000 feet(6706m) and 300knots (555.6kph)

Velocity of 200kph (55.56 )

Velocity of 300kph (83.33 ) Velocity of 400kph (111.11 )

Velocity of 100kph (27.78 )


Figure 3.1a [14] Figure 3.1b [14]

Figure 3.1c [14] Figure 3.1d [14]

Figure 3.2a [14] Figure 3.2b [14]

0.73



Figure 3.2c [14] Figure 3.2d [14]


3.3 Solidworks Stress Analysis

Stress analysis is a process to test a computer drawing model with mechanical stresses,

strains and deformations. I have conducted the stress analysis process using Solidworks

Simulation based on the following flow chart below [20]. Plus I have also list down the

advantages and disadvantages of Solidworks software (Table 3.3a).

Advantages Disadvantages

It can give an accurate result of displacement,

Von Misses stress and strain on any

dimensions of shape and object constructed

In Solidworks we will not able to calculate all the

masses in one flow. However, I have to calculate it

separately for each part and compare the masses. For

example, for whole wing model it shows 327.80 kg

while when calculated separately we have got

326.6kg.This shows slight error of 1.2kg between

both weights.

It provides a proper step to do stress analysis

so that the sequence is not missed out starting

from, fixture, load, material, and run

simulation

It does not show any proper equations on how the

stress, displacements are calculated.

It does provide a good understanding of

maximum and minimum value with colour

coding. For example, red for maximum value

and blue for minimum value

Some materials does not show all the properties

which causes the stress analysis in failure such as

Carbon fibre materials

Start

Determine the lift force

Meshing

Apply fixture, force and specific material

Solidworks Simulation Analysis

Results and Discussion

Design the wing

Table 3.3a


3.4 Process flow

1) a) The lift force which was calculated at 12,000 feet(3658m) with 250knots(463kph)

is 18990N

b) The lift force which was calculated at 22,000 feet (6706m) with 300 knots (555.6kph)

is 19631N

2) The model is constructed in Solidworks software with below dimensions (Figure 3.4a).

Length of half wing span = 3.44m

Width of fuselage = 1.37m

3) The fixture was applied with following steps :

The Simulation Xpress Analysis wizard on the top right of the Solidworks programme is

been clicked and a sequence step feature will appear. The fixtures option which is bolded

will be first task to be completed. All the future steps will be bolded once each task has

been completed (Figure 3.4b)

Figure 3.4a

Figure 3.4b

[30]


4) Next the force is applied on both port (left) and starboard(right) of the wing.

5) The material selection is done as shown below.

The outlook of fixed geometry is shown in circle (Figure 3.4c)

The load amount was typed in the column provided on the next step is triggered as

shown in the circle (Figure 3.4d). Once the arrow is shown (circled) towards

upward direction from beneath the wing, we moved on to next step

The next step is triggered (materials) and Aluminium Alloy 7075 T-6 is

selected. The material properties is shown in circle (Figure 3.4e)

Figure 3.4c

Figure 3.4d

Figure 3.4e


6) The meshing is created.

7) The outlook of Solidworks Simulation Xpress is shown below.

The meshing level is selected in between the coarse and fine

as shown in the circle and the final task is assigned (Figure 3.4f)

The first simulation was done on stress analysis (Figure 3.4g). The colour coding is applicable to

displacement (Figure 3.4h) and factor of safety (Figure 3.4i) as well. The red colour shows the most

critical part, green shows the average part and the blue shows safest part. The yield stress (circled)

will be shown on bottom section of the colour coding. In case a material fails, there will be a red

colour indication (circled) will be shown since that will be the most critical part.

Figure 3.4f

Figure 6.7

Figure 6.8

Figure 6.9

Figure 3.4g Figure 3.4h

Figure 3.4i


RESULTS AND DISCUSSION

4.1 Distribution of data at 12,000 feet cruise with 250knots.

Table 4.1a and 4.2a shows the breakdown of parameters tested using Foil SIM (III) software.Bolded values

which tested in Foil SIM (III) is shown below.

Manual Calculation

At the 12,000feet (3658m) with 100kph is converted to 27.78 ( )

SIM (III) software result.

Reynolds Number is a dimensionless number used in fluid mechanics to indicate wheter fluid

past a body or in a duct is steady or turbulent[31].It is normally affected by ,kinematic

viscosity, chord length , ,fluid density, and fluid velocity, Reynolds number,Re always

directly proportional to aircraft speed.

)

4.2 Distribution of data at 22,000 feet cruise with 300knots.

Speed (kph) Lift (N) Drag (N) Lift/Drag

Ratio

Lift

Coefficient

(Cl)

Drag Coefficient

(Cd)

Reynolds

Number

100 (Fig 3.1a) 890.00 43 20.258 0.73 0.045 655296

200 (Fig 3.1b) 3563.30 166 21.349 0.73 0.045 1310592

300 (Fig 3.1c) 8018.00 364 22.001 0.73 0.045 1965888

400 (Fig 3.1d) 14255.00 634 22.470 0.73 0.045 2621184

500 22139.60 1364.78 16.220 0.73 0.045 3276480

600 31881.84 1965.32 16.220 0.73 0.045 3931776

Speed (kph) Lift (N) Drag

(N)

Lift/Drag

Ratio

Lift

Coefficient

(Cl)

Drag Coefficient

(Cd)

Reynolds

Number

100 (Fig 3.2a) 639 32 19.843 0.73 0.045 500986

200 (Fig 3.2b) 2559 122 20.922 0.73 0.045 1001972

300 (Fig 3.2c) 5759 267 21.568 0.73 0.045 1502958

400 (Fig 3.2d) 10238 464 22.06 0.73 0.045 2003946

500 15899.81 980.125 16.22 0.73 0.045 2504930

600 22896.27 1411.41 16.22 0.73 0.045 3005916

Table 4.1a

Table 4.2a


Manual Calculation

At the 22,000feet (6706m) with 100kph is converted to 27.78 ( )

Sim (III) software result.

Reynolds Number

Above calculations for the speed of 100kph are done in order to prove that the Foil SIM (III) software

calculations matches with manual calculations.However,the actual speed of aircraft at 12,000 feet

(3658m) will be 250knots (463kph) and at 22,000 feet (6706m) will be 300knots (555.6kph).Hence

the lift generated at 12,000 feet (3658m) and 22,000 feet (6706m) are 18990N and 19631N

respectively.Hence, this successfully fulfill my mission of reaching the wing lift above 16448.48 N

to carry the aircraft at both cruising altitudes.(All the manual calculations from the speed of 200kph

to 600kph are stated in Appendix F and G.Plus, related parameter graphs with explanations are shown

in Appendix H).

4.3 Accuracy of Foil Sim (III) software and manual calculations The results obtained through Foil SIM (III) and manual calculations are almost accurate except for a

few number of parameters which as shown below (Table 4.3a and Table 4.3b). The red indicates the

differences in values and the yellow indicates as not available since the software can only calculate

up to 400kph (Refer Appendix I)

Accuracy of Foil SIM (III) with manual calculation 12000 feet (3658m)

Speed

(kph)

Lift (N) Drag (N) Reynolds Number L/D ratio

FOILSIM

(III) Manual

FOILSIM

(III) Manual

FOILSIM

(III) Manual

FOILSIM

(III) Manual

100 890 890 43 43 655296 655296 20.258 20.258

200 3563 3563 166 166 1310592 1310592 21.349 21.349

300 8018 8018 364 364 1965888 1965888 22.001 22.001

400 14255 14255 634 634 2621184 2621184 22.470 22.470

500

22140

1365

3276480

16.220

600

31882

1965

3931776

16.220

[32]

Table 4.3a


Accuracy of Foil SIM (III) with manual calculation 22000 feet (6706m)

Speed (kph)

Lift (N) Drag (N) Reynolds Number L/D ratio

FOILSIM

(III) Manual

FOILSIM

(III) Manual

FOILSIM

(III) Manual

FOILSIM

(III) Manual

100 639 639 32 32 500986 500986 19.843 19.843

200 2559 2559 122 122 1001973 1001972 20.922 20.922

300 5759 5759 267 267 1502959 1502958 21.568 21.568

400 10238 10238 464 464 2003946 2003946 22.032 22.060

500

15900

980

2504930

16.220

600

22900

1411

3005916

16.220

4.4 Induced drag corresponding to Aspect ratio

Below calculation is used to find Induced drag coefficient using aspect ratio values. Average aspect

ratio is taken since each type of aircraft in above Table 1.3a shows different range of values. The drag

is calculated overall in 12,000 feet (3658m) and 22,000 feet (6706m) cruise.

Induced Drag Coefficient =

[22]

= Lift coefficient when Liebeck LNV109A at zero degree angle of attack, 0.73 (Refer 3.1a where

the ORANGE arrow is showed)

= Average aspect ratios

Then, the lift coefficient is replaced by induced drag coefficient in lift formula (as shown in box)

[22]

Type of aircraft Average aspect ratio (except

for SKYSONIC)

Induced drag

coefficient

Induced drag (N) at

12,000feet (3658m)

Induced drag (N) at

22,000feet (6706m)

Military fighters 3.5 0.048 1248.23 1290.33

GA Aircraft 8.5 0.020 520.10 537.64

Commercial

Jetliners 8.5 0.020 520.10 537.64

SKYSONIC 18.31 (real value, not

average)

0.009 234.04 241.94

Table 4.4a above shows the data tabulation of induced drag at cruising altitudes

Since , Skysonic is included in the category of sailplane , it is compared now with other types of aircrafts.GA

aircraft, Commercial Jetliners and Military fighters aspect ratios are taken as average since they have certain

range as shown in Table 1.3a. Hence, from above data, I can conclude that the induced drag produced by

Skysonic at 12,000 feet and 22,000 feet are lower compared to other aircrafts. Plus, the high aspect ratio of

Skysonic proves that it has a great capability to reduce tip vortices (causes induce drag) more efficiently.

Table 4.3b

Table 4.4a


4.5 Stress Analysis Results and Discussion

Mission of Stress Analysis for Skysonic wing Cruising Loads (Table 9.1)

Wing tested using Aluminium Alloy 7075-T6 with yield stress of 505,000,000 MPa

(Refer Appendix J)

Wing tested with Graphite with yield stress of 120,594,000 MPa

(Refer Appendix J)

Criteria High Low

Von Mises Stress (MPa) *

Displacement (mm) *

Factor of safety *

Study Von Mises stress (MPa) Resultant displacement (mm) Factor of

safety Maximum Minimum Maximum Minimum 12000feet(3658)

with the load of

18990 N

164,936,960 13,744,747 199.88

3.06

22000feet(6706m)

with the load of

19631 N

170,504,336 14,208,695 206.54

2.96

Study Von Mises stress (MPa) Resultant displacement (mm) Factor of

safety Maximum Minimum Maximum Minimum

12000 feet(3658m)

with the load of

18990 N

165,900,528 13,825,044 3010.71

0.73

22000 feet(6706m)

with the load of

19631 N

171,500,464 14,291,705 3112.33

0.70

Table 4.5b

Table 4.5a

The results of the simulation analysis is based on the process flow which was done in ( flow chart Pg.23).All the

steps were followed in sequence order with both wings were tested with lift forces of 18990N and 19631N

obtained at 12000 feet (3658m) and 22000 feet (6706m) cruise respectively. Aluminium Alloy 7075-T6 and

Graphite materials were selected for comparison based on the research done by the Mechanical Engineering

Department of Al-Nahrain University, Baghdad, Iraq on light weight aircraft wing structure in the year of

2008[23].Hence, considering the similar research concept on light aircraft wing, I have chosen these two materials

to identify the maximum and minimum Von Mises stress, maximum and minimum displacements and factor of

safety. Below results are obtained from the analysis:

Table 4.5c


Aspects Justification

Effect on Von Mises Stress

(MPa)

The stress range of Aluminium Alloy is lower compared to Graphite

with same load (18990N and 19631N) applied in both conditions.

Aluminium Alloy is still far more safer compared to Graphite since the

safety factor of Aluminium Alloy is almost three times greater than

graphite even though Graphite's maximum stress exceed Aluminium

Alloy.

Effect on Displacement

(mm)

The resultant displacement of Aluminium Alloy is approximately 15

times lower than Graphite. It clearly shows the wing stays more stiffer

and stronger when Aluminium Alloy is used because smaller number of

displacement range induces low fatigue characteristics [24]

Effect on Factor of Safety

This actually proves all the justifications above where Aluminium Alloy

material dominates the Graphite in both aspects, on Von Mises stress and

displacement. The safety factor of Aluminium Alloy exceeds Graphite

by 3 times and since Graphite exceeds its yield strength of 120, 594, 00

MPa, it will not able to experience and accept such heavy loads imposed

on wing at both cruising positions. Meanwhile, Aluminium Alloy will

able to withstand these loads easily with less displacements since the

range to achieve its yield strength is still far. Referring to British Civil

Airworthiness Requirements CAP 482 under Sub-section C

Structure,S303,minimum factor of safety required for small light

aeroplane is 1.5[25].Hence, the result obtained also successfully

complied to the regulation.

A B

The A and B dotted line (as shown in Figure 4.5a)having almost same length to each other .

Hence, the difference of maximum and minimum value at both A and B lines does not have much

variance unlike the displacement and the factor of safety (Refer Appendix J).

Table 4.5d

Table 4.5d shows the aspects and justification of the stress analysis done

Figure 4.5a

Figure 4.5a below shows the maximum Von Mises stresses at 12,000 feet (3658m)and 22,000 feet(6706m)

loads corresponding to Aluminium Alloy 7075-T6 and Graphite materials


CONCLUSION

The study has examined the characteristics of Skysonic jet wing under certain conditions

set by the DGEN 390 Engine specifications such as altitudes and speeds .Returning to the

mission posed at the beginning of the study to achieve 16448.48 N to carry the aircraft ,it

is now possible to state that, Skysonic wings have successfully overcome it at both 12,000

feet and 22,000 feet cruising altitudes.Plus, this study also proves that Skysonic wings

were able to withstand the induced drag with appropriate material selection.Hence, the

decision of reducing the 20% of wing dimensions from preliminary aircraft at the initial

stage of the project is absolutely correct since all the parameters at the final stage are

within the limits corresponding to DGEN 390 engine specifications.The findings of this

study suggest that the DGEN 390 engine has a lower thrust and in order to carry a business

light jet in 12,000 feet (3658m) and 22,000 feet (6706m) altitudes, the wings have to play a

very big role in which in this study the Skysonic wing cruising lifts are almost 15 times

higher than engine thrust. The study has gone someway towards enhancing my

understanding of wings especially on detailed frame structures and vast material

applications using the appropriate computer tools such as Solidworks Computer Aided

Drawing (CAD) and Foil SIM(III) software.During the process of research, the study was

limited in several ways. First, the time consumption for in depth research and second lack

of resources since the information available were not converged to our requirement.

Further investigations and experimentations such as wind tunnel test , cost analysis and

utilization of Computational Fluid Dynamics (CFD) are strongly recommended in our

Skysonic aircraft so that the analysis will give a full completion and in depth of

understanding.


FUTURE TASKS

i) Wind tunnel test

- Since all the results at 12,000 feet (3658m) and 22,000 feet (6706m) are

obtained from software and manual calculation, a realistic experiment has to

be done with a prototype design. The wind tunnel results might be same or

not with the software based calculation but for sure it will give a great

exposure in accuracy and precision comparing both outputs.

ii) Cost estimation

- The financial planning for Skysonic jet should be done especially in

materials and wing structures to fulfil the expectation and requirement of

customers. The cost of each wings of Skysonic jet might be higher or lower

than the preliminary aircraft wing , but cost estimation element is very

important since it will be covering the manufacturing cost, break-even

analysis and maintenance cost.

iii) Application of Computational Fluid Dynamics (CFD)

- The CFD programme should be used wisely for designing in depth

features and for the analysis methods since the software is more likely to be

used in most of the aircraft designing institutions and research centres.

Plus, I have to improve my designing skills in CFD since I only have

experience in Solidworks programme.


CONTEXT REFERENCES

1. Price-Induction. (2014). DGEN 380-390. Available: http://www.price-

induction.com/site_media/plaquettes/DGEN%20specifications%20sheet.pdf. Last accessed

4th May 2014.

2. Maverick Jets. (2006). Smart Jet. Available: http://www.maverickjets.com/jets/smartjet.php.

Last accessed 12th March 2014.

3. FindTheBest. (2014). Maverick Smart JET. Available:

http://planes.findthebest.co.uk/l/68/Maverick-SmartJET. Last accessed 12th March 2014.

4. XO Jet Incorporation. (2000). Hawker 400XP. Available:

http://www.xojet.com/Fleet/hawker-400xp/Hawker-400XP-Private-Jet.asp. Last accessed 01st

April 2014.

5. Daniel P.Raymer (1992). Aircraft Design : A Conceptual Approach. 2nd ed. California:

AIAA Education Series. p37.

6. Mohammad Sadraey (2013). Chapter 5 Wing Design. Daniel Webster College: Mohammad

Sadraey. p179-182.

7. Snorri Gudmundsson (2014). General Aviation Aircraft Design : Applied Methods and

Procedures. United States of America: Elsvier Incorporation. p293.

8. Airfoil Investigation Databese. (2013). LOCKHEED L-188 TIP AIRFOIL.Available:

http://www.airfoildb.com/foils/547. Last accessed 03rd March 2014.

9. Airfoil Investigation Databese. (2013). LNV109A. Available:

http://www.airfoildb.com/foils/541.

Last accessed 03rd March 2014.

10. Airfoil Investigation Databese. (2013). NACA 65(2)-215. Available:


11. Airfoil Investigation Databese. (2013). NACA 0006. Available:




13. John Dreese . (2007). Part 2: Basic Terms & Geometry . Available:

http://www.dreesecode.com/primer/airfoil2.html. Last accessed 21st Feb 2014.

14. National Aeronautics and Space Administration (NASA). (2014). FoilSim III Student Version

1.5a. Available: http://www.grc.nasa.gov/WWW/k-12/airplane/foil3.html. Last accessed

17th Apr 2014.


Sadraey. p207.


16. Daniel P.Raymer (1992). Aircraft Design : A Conceptual Approach. 2nd ed. California:

AIAA Education Series. p50.


Procedures. United States of America: Elsvier Incorporation. p309-310.

18. National Aeronautics and Space Administration (NASA). (2014). Induced Drag

Coefficient. Available: http://www.grc.nasa.gov/WWW/k-12/airplane/induced.html. Last

accessed 11th April 2014.

19. ANSYS. (2013). Wing modeling . Available:

http://www.structures.ethz.ch/education/master/analysis/FS2013PDFs/Ansys/ANSYS_ExII_-

_Problem_description.pdf. Last accessed 18th Apr 2014.

20. University of Science Malaysia. (2011). Sky Eye TM. Available:

http://aerospace.eng.usm.my/rcp/index.php/analysis/computational-fluid-dynamics-

cfd/skyeyetm. Last accessed 13th Mar 2014.

21. A.C.Kermode (2006). Mechanics of Flight. 11th ed. England: Pearson Education Limited.

p159-160.

22. A.C.Kermode (2006). Mechanics of Flight. 11th ed. England: Pearson Education Limited.

p104.

23. Professor Dr.Muhsin J.Jweeg (2008). Optimization of Light Weight Aircraft Wing Structure.

Iraq: Al-Nahrain University. p15.

24. AUTODESK. (2014). Small Deflection Stress analysis. Available:

http://help.autodesk.com/view/MFIWS/2014/ENU/?guid=GUID-1B128FF9-2E3A-4170-

8984-618E66FA4675. Last accessed 23rd Apr 2014.

25. Safety Regulation Group (1983). CAP 482 British Civil Airworthiness Requirements. 3rd ed.

United Kingdom: Civil Aviation Authority. Part 1 Sub-Section C Page 1.

26. X-Plane. (2012). Airfoils. Available: http://strategywiki.org/wiki/X-

Plane/Developing/PlaneMaker/Airfoils. Last accessed 18th Dec 2013.

27. Airfoil Tools. (2014). Airfoil Comparison. Available: http://airfoiltools.com/compare/index.

Last accessed 3rd Apr 2014.


Sadraey. p184-185.

29. Solidworks 2014 3D Mechanical Computer Aided Design (CAD) Programme

30. National Aeronautics and Space Administration. (2014). Lift to Drag ratio. Available:

https://www.grc.nasa.gov/www/k-12/airplane/ldrat.html. Last accessed 01st May 2014.

31. Oxford University Express. (2014). Reynolds Number. Available:

http://www.oxforddictionaries.com/definition/english/Reynolds-number. Last accessed 21st Apr 2014.

32. National Aeronautics and Space Administration. (2009). Reynolds Number. Available:

http://www.grc.nasa.gov/WWW/BGH/reynolds.html. Last accessed 19th Mar 2014.


FIGURES REFERENCES

1. Figure 1a and 1b

National Aeronautics and Space Administration (NASA). (2014). FoilSim III Student

Version 1.5a. Available: http://www.grc.nasa.gov/WWW/k-12/airplane/foil3.html. Last

accessed 17th Apr 2014.

2. Figure 1c and 1d

Aerofoil Engineering. (2014). AeroFoil version 3.2. Available:

http://aerofoilengineering.com/. Last accessed 17th Apr 2014.

3. Figure 1e

Mohammad Sadraey (2013). Chapter 5 Wing Design. Daniel Webster College:

Mohammad Sadraey. p182.

4. Figure 1f

Snorri Gudmundsson (2014). General Aviation Aircraft Design : Applied Methods and


5. Figure 1.1a

Airfoil Tools. (2014). LNV109A (lnv109a-il). Available:

http://airfoiltools.com/airfoil/details?airfoil=lnv109a-il. Last accessed 11th Jan 2014.

6. Figure 1.3

Glue-it.com. (1999). Model Aircraft Glossary. Available: http://www.glue-

it.com/aircraft/general-information/glossary/v_summ.htm#.U2egn_ldV8E. Last accessed

1st Apr 2014.

7. Figure 1.3a

X-Plane. (2011). What's With Winglets?. Available: http://forums.x-

plane.org/?showtopic=54184. Last accessed 17th Mar 2014.

8. Figure 1.3b

RC Groups. (2003). Aerodynamics Stall and Spin. Available:

http://adamone.rchomepage.com/index6.htm. Last accessed 17th Mar 2014.

9. Figure 2a , 2b

Solidworks 2014 3D Mechanical Computer Aided Design (CAD) Programme

10. Figure 2.2a , 2.2b , 2.2c , 2.2d


11. Figure 3a, 3b and 3c





12. Figure 3.1a, 3.1b, 3.1c, 3.1 d




13. Figure 3.2a , 3.2b, 3.2c, 3.2d




14. Figure 3.4a to Figure 3.4i



APPENDIX A

Altitude (m) Thrust (N)

3048 1500

3658 1400

5486 1150

6706 1100

Maximum Take - off Weight (N) 16500 N

Maximum speed at 12,000 feet (3658m) 250 knots (463kph)

Flight Envelope 25,000 feet (7620m)

Thrust at 10,000 feet (3048m) 1500 N

Thrust at 18,000 feet (5486m) 1150 N

Table above shows the DGEN-390 performance specifications [1]

0

200

400

600

800

1000

1200

1400

1600

0 1000 2000 3000 4000 5000 6000 7000 8000

Thru

st (

N)

Altitude (m)

Thrust versus Altitude


APPENDIX B

Aerofoil term definitions [19]

i) Chord

- The distance between leading edge and trailing edge of the aerofoil.

ii) Chord line

- The line which connect the leading edge and trailing edge along the aerofoil.

iii) Mean camber

- The line which is equidistant from upper and lower surfaces.

iv) Angle of attack

- The angle which forms between relative airflow and chord line.

v) Leading Edge

- The part of aerofoil where the air hits first.

vi) Trailing edge

- The part of aerofoil where the air hits last.

vii) Maximum thickness

- The maximum distance of the upper surface and lower surface.


Stall angle (28) is the angle of attack at which the aerofoil stalls

Maximum lift coefficient (28) is the maximum capability for the aerofoil in wing to lift

the aircraft weight.

Zero lift angle of attack (28) is the aerofoil angle of attack in which the lift coefficient is

zero.

The lift coefficient at zero angle of attack (28) is the lift coefficient when angle of

attack is zero. Mostly, during zero degree angle of attack, positive lift coefficient will be

produced.

AEROFOIL GRAPH CHARACTERISTICS [27]

Lift Coefficient vs Drag Coefficient

Pitch Coefficient vs Angle of Attack Drag Coefficient vs Angle of Attack


APPENDIX C

ESTIMATING THE LIFT MISSION BY PERCENTAGE [21]

The engine thrust will be 1400N where I have refer it

to the DGEN 390 specification [Refer Appendix A]

1400N + Wing force (N) = 17840 N

To calculate the percentage amount of force that produced by DGEN engine :

To calculate the percentage amount of force that need to produce by both straight wings

to lift the aicraft at cruising stage

100% - 7.8% = 92.2%

Hence, 92.2% of the lift is contributed by the straight wings at cruising stage

Conversion of percentage to lift force in N

92.2% = 16,448.48 N

Hence, 16,448.48 N is needed at least to lift the aircraft at 12,000 feet (3658m) at 250

knots (463kph and 22,000 feet (6706m) at 300knots (555.6kph) .So this will be my mission

to achieve in order to lift Skysonic at both altitude cruising stages.

ASPECT RATIO

= 8.24m

= 3.708

=

AR = 18.31

( Engine Thrust + Wing Force) at cruising position = Aircraft Gross weight or Maximum Take Off weight


APPENDIX D

Wing Parts Dimensions (mm) Mass (Kg)

Front Spar

Length : 18mm

Height : 51.07mm on right

44.88mm on left

Width : 3405mm

6.69

Rear Spar

Length : 18.44mm


28.57mm on left

Width : 3405 mm

4.00

Middle spar

Length : 28.78mm


57.90mm on left

Width : 3405 mm

10.77

Wing root and tip

rib (2 ribs in half wing)

Length : 450.0mm

Height : 58.17mm

Width : 35mm

2.78


Wing ribs (6 ribs in half wing) Length : 450.0mm

Height : 58.17mm

Width : 40mm

Hole diameter : 25mm

8.88

Leading edge assembly

Length : 30.03mm

Height : 31.95mm on left

44.97mm on right

Width : 3405 mm

10.47

Trailing edge assembly

Length : 181.52mm

Height : 24.34mm on left

1.63mm on right

Width : 3405 mm

20.03

Stringers on top rib surface (4 stringers)

Diameter : 12mm

Width : 3405 mm

5.16

Stringers on bottom rib surface (4 stringers)

Diameter : 10mm

Width : 3405 mm

3.56


Wing Skin

Length : 450.0mm

Height : 58.17mm

Width : 3405 mm

Thickness : 7.0mm

91.55

APPENDIX E

Parts Function

Spar Vertical structure running along the wingspan and it carries

shear loads

Ribs Elements are located in the cross-section plane. It has main

function to distribute the load and maintaining the shape.

Root wing Part of the wing which attached to the fuselage

Tip wing Part of the wing which attached away from the fuselage

Leading edge

assembly Front part of the wing which holds the support structure

Trailing edge

assembly Back part of the wing which holds the support structure

Stringers

Part of the wing which carries the axial force and balancing

the bending moment .It also stabilises the wing skin from

buckling

Wing Skin The outer layer which covers the wing structure


APPENDIX F

(i) At the 12,000 feet (3658m) = applicable for 100kph, 200kph, 300kph and 400kph

Calculating the Reynolds Number

)

Calculating the Lift Drag ratio

Ratio =

[30]

(ii) At the 22,000 feet (6706m) = applicable for 100kph, 200kph, 300kph and 400kph



Ratio =

[30]


APPENDIX G

Manual Calculation of parameters exactly at 250knots (463kph) and 300knots (555.6kph)

At the 12,000 feet (3658m)


)


Ratio =

At the 22,000 feet (6706m)



Ratio =

Parameters Values

Speed (kph) 463

Lift (N) 18990

Drag (N) 1170

Reynolds Number 3030602

L/D ratio 16.22

Parameters Values Speed (kph) 555.6

Lift (N) 19631

Drag (N) 1211

Reynolds Number 2779949

L/D ratio 16.22


0

5000

10000

15000

20000

25000

30000

35000

0 100 200 300 400 500 600 700

Lift

(N

)

Aircraft speed (kph)

Lift versus Aircraft speed

0

500

1000

1500

2000

2500

0 100 200 300 400 500 600 700

Dra

g (N

)

Aicraft speed (kph)

Drag versus Aircraft speed

0

5

10

15

20

25

0 100 200 300 400 500 600 700

Lift

Dra

g ra

tio


Lift Drag ratio versus Aircraft speed

APPENDIX H

Skysonic wing graph parameters at 12,000feet(3658m) and 22,000(6706m) feet cruise.

Note : Red indicates parameters at 12,000 feet

Blue indicates parameters at 22,000feet

Yellow indicates equal parameters


0

500000

1000000

1500000

2000000

2500000

3000000

3500000

4000000

4500000

0 100 200 300 400 500 600 700

Re

yno

lds

Nu

mb

er


Reynolds number versus Aircraft speed

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0 100 200 300 400 500 600 700

Lift

Co

eff

icie

nt

(CL)

Aircraft Speed (kph)

Lift Coefficient versus Aircraft speed


Lift or Drag versus Aircraft Speed

-The lift or drag and aircraft speed is directly proportional to each other

since the Cl and Cd at zero degree angle of attack remains 0.73 and 0.045

respectively throughout the cruising position. The curve of 12,000 feet

(3658m) is steeper than 22,000 feet (6706m) curve since the density at high

altitude is lower than low altitude for both lift and drag magnitudes.

Lift Drag Ratio versus Aircraft Speed

- The lift drag ratio at 12,000 feet(3658m) is much higher than 22,000

feet(6706m) .This show in 12,000 feet (3658m) the amount of drag

produced by the lift is much lower than 22,000feet (6706m).Thus this will

cause fuel consumption at 12,000 feet is much lower than 22,000 feet [30]

Reynolds number versus Aircraft Speed

- Reynolds number at 12,000 feet (3658m) is higher than 22,000 feet

(6706m) due to the velocity increment with different densities. Hence,

the ratio of inertial forces (resistant to change) to viscous forces

(heavy and gluey) is higher at 12,000feet(3658m) compared to

22,000 feet (6706m)[32]

Lift Coefficient versus Aircraft Speed

- The lift coefficients in both altitudes are remains equal due to the

zero angle of attack. The lift coefficient only varies if the angle of

attack changes.


APPENDIX I

ACCURACY PERCENTAGE

Based on the Table 4.3b,

Total number of values : 16 (values are taken up to 400kph since Foil SIMs ability

to calculate is only up to that speed)

Total number of errors : 3 (red boxes of Foil SIM (III) data, excluding manual )

Table 4.3a does not meet any errors however Table 4.3b meet 3 errors. Above calculation proves that,

the Foil SIM software calculation will differ with manual calculation up to 18.75 % in every

calculation from 100kph to 400kph.


APPENDIX J

12000 feet altitude (3658m) with 18990N of load using Aluminium Alloy 7075-T6

Elastic Modulus (MPa) 72000

Yield Strength (MPa) 505

Mass Density ( ) 2810

Tensile Strength (MPa) 570

Possion Ratio 0.33

Von Mises Stress (MPa) Displacement (mm)

Factor of Safety Standard Meshing

[29]


22000 feet (6706m) with 19631N of load using Aluminium Alloy 7075-T6




12000 feet altitude (3658m) with 18990 N of load using Graphite

Elastic Modulus (MPa) 4800

Yield Strength (MPa) 120.59

Mass Density ( ) 2240

Tensile Strength (MPa) 100.83

Possion Ratio 0.28


Standard Meshing Factor of Safety

[29]


22000 feet (6706m) with 19631 N of load using Graphite




APPENDIX K

Both A and B dotted line are not having same length to each other, hence they have different

maximum and minimum ranges compared to each other. Graphite is having almost 15 times

greater displacement range compared to Aluminium Alloy 7075-T6.Since the Graphite

material wing have a large displacement[24], it will not be stiff and strong enough to absorb

the cruising loads of Skysonic. However for the Aluminium Alloy 7075-T6 material, they

have very small gap of difference in displacement, which will able to withstand the loads

without any damage.

The A and B dotted lines are also not having the same length to each other in this case where

the wing safety factor range using Aluminium Alloy 7075-T6 exceeds the Graphite. Hence,

this concludes that Aluminium Alloy 7075-T6 is 3 times safer to use in Skysonic wing rather

than Graphite in order to carry the cruising loads.

A

B

A

B


APPENDIX L


APPENDIX M

Gan

tt C

hart

Vers

ion

On

e


Gan

tt C

hart

Vers

ion

Tw

o


Gan

tt C

hart

Vers

ion

Th

ree

(Fin

al)

Documents

Kisyan Light Jet THESIS ( University of Brighton)