Aircraft Design Thesis Report

CONCEPTUAL DESIGN OF A BUSINESS JET
FINAL REPORT
BSc. Aeronautical Engineering
Spring 2019
Date of Submission: 23-05-2019

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ABSTRACT
The aim of the project is to design and develop a Business aircraft that is more efficient
and can travel greater miles. The design will be implemented based on four reference
aircrafts in the same respective field. The four aircrafts which are used to conduct this
design are as follows: Global 6000, Global 5000, Gulfstream G-600 and Gulfstream G-
550. Using the characteristics of the previous mentioned aircrafts, the modification
that was intended was to remove the horizontal tail from the aft and making it as a
canard, this causes the aircraft to be more stable as all the three engines are aft
mounted along with the empennage initially. To Justify our modification we have
performed several calculations throughout the course of the design project.

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NOMENCLATURE
𝑾 𝟎 Take off Gross Weight, lb
𝑾 𝒄𝒓𝒆𝒘 Crew Weight, lb
𝑾 𝒑𝒂𝒚𝒍𝒐𝒅 Payload Weight, lb
𝑾 𝒇𝒖𝒆𝒍 Fuel Weight, lb
𝑾 𝒇
𝑾 𝟎
Fuel Fraction
𝑾 𝒆
𝑾 𝟎
Empty Weight Fraction
𝑾 𝒙
𝑾 𝟎
Mission Profile
𝑾 𝟏 Weight after Takeoff, lb
𝑾 𝟐 Weight at the end of Climb, lb
𝑾 𝟑 Weight after Cruise, lb
𝑾 𝟒 Weight after Descend, lb
𝑾 𝟓 Weight after Loiter, lb
𝑾 𝟔 Weight after Landing, lb
R Range, ft
E Endurance, s
C Specify Fuel Consumption
V Velocity, ft/s2
L/D Lift to Drag ratio
SFC Specific Fuel Consumption
BSFC Brake Specific Fuel Consumption
ηp Propeller Efficiency
AR Aspect Ratio
Sa Obstacle clearance distance, ft

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e Oswald efficiency factor
CD0 Zero lift drag coefficient
𝜳 Turn rate
𝒏 Load factor
𝑮 Climb gradient
Ct Tip chord, ft
Cr Root chord, ft
Λ Taper ratio
M Mach Number
Ww Weight on Wheel, lb
L Lift, lb
D Drag, lb
Nt Number of fuel tanks
Nz Ultimate load factor; =1.5*limit load factor
q Dynamic pressure at cruise, lb/ft2
Sf Fuselage wetted area, ft2
Sht Horizontal tail area, ft2
Svt Vertical tail area, ft2
Sw Wing area, ft2
t/cwing Wing thickness to chord ratio
t/cht Horizontal tail thickness to chord ratio
t/cvt Vertical tail thickness to chord ratio
Wdg Design gross weight, lb
Wen Engine weight, lb
Wfw Weight of fuel in wing, lb
Wl Landing Design Gross Weight, lb

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Wpresss Weight penalty due to pressurization, lb
Wuav Uninstalled avionics weight, lb
Λ Wing Sweep at 25% MAC
Λht Sweep angle at Horizontal Tail
Λvt Sweep angle at Vertical Tail
λw Taper ratio for wing
λht Taper ratio for horizontal tail
λvt Taper ratio for vertical tail
b Span
A Aspect Ratio
𝑺 𝒓𝒆𝒇 Wing reference Area
𝑺 𝒘𝒆𝒕
𝑺 𝒓𝒆𝒇
Wettest Area Ratio
Wing sweep back angle
𝒕/𝒄 Thickness ratio
𝜹 Wing twist angle
i Incidence angle
MAC Mean aerodynamic chord
𝒄= Mean aerodynamic chord
𝑻
𝑾
Thrust to weight ratio
𝑷
𝑾
Power to weight ratio
𝑽 𝒔𝒕𝒂𝒍𝒍 Stall speed
𝑽 𝟏 Decision speed
𝑽 𝑹 Rotation speed
𝑽 𝟐 Safety speed
𝝈 Density ratio
𝑪 𝑳𝑻𝑶 Takeoff lift coefficient
𝑺𝒍𝒂𝒏𝒅𝒊𝒏𝒈 Landing distance
𝑪 𝑳 Lift coefficient
𝑪 𝑫 Drag coefficient
SM Static Margin

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𝑨 𝑾 Aspect ratio for wing
𝑨 𝒉𝒕 Aspect ratio for the horizontal tail
𝑨 𝒗𝒕 Aspect ratio for the vertical tail
𝑩 𝑾 Wing Span
𝑯 𝒗
𝑯𝒕
Horizontal tail height above fuselage to vertical tail height above fuselage ratio
L/D Lift to Drag ratio
L Fuselage structural length
𝑳 𝒎 Length of main landing gear
𝑳 𝒏 Nose gear length
𝑳𝒕 Tail length, MAC of wing to MAC of tail, ft
M Mach number
V Cruising speed
𝑵 𝒆𝒏 Number of engines
𝑵𝒍 Ultimate landing factor
𝑵 𝒑 Number of personnel onboard

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Table of Contents
1.0. INTRODUCTION ......................................................................................................................................11
2.0. OBJECTIVES ...........................................................................................................................................11
3.0. MARKET RESEARCH.............................................................................................................................12
4.0. MISSION PROFILE..................................................................................................................................14
5.0. PRELIMINARY SIZING............................................................................................................................16
5.1. TAKE OFF GROSS WEIGHT .....................................................................................................................16
5.1.1. Estimating Empty Weight Fuel Fraction 𝑾𝒆𝑾𝟎.......................................................................17
5.1.2. Estimating Fuel Fraction 𝑾𝒇𝑾𝟎.................................................................................................18
6.0. WING DESIGN .........................................................................................................................................22
6.1. AIRFOIL SELECTION ...............................................................................................................................22
6.2. AIRFOIL GEOMETRY ...............................................................................................................................22
6.3. AIRFOIL LIFT AND DRAG.........................................................................................................................23
6.4. AIRFOIL SELECTION AND DESIGN ...........................................................................................................24
6.5. STALL ....................................................................................................................................................28
6.6. AIRFOIL THICKNESS RATIO .....................................................................................................................28
6.7. ASPECT RATIO .......................................................................................................................................31
6.8. WING SWEEP .........................................................................................................................................32
6.9. TAPER RATIO .........................................................................................................................................33
6.10. WING TWIST............................................................................................................................................33
6.11. WING INCIDENCE ....................................................................................................................................33
6.12. WING VERTICAL LOCATION ....................................................................................................................34
6.13. WING TIPS ...............................................................................................................................................34
7.0. TAIL CONFIGURATION..........................................................................................................................35
8.0. THRUST TO WEIGHT RATIO.................................................................................................................37
8.1. THRUST TO WEIGHT RATIO AT CRUISE CONDITION ..................................................................................37
8.2. THRUST TO WEIGHT RATIO AT TAKEOFF CONDITION ...............................................................................38
8.3. THRUST TO WEIGHT RATIO AT CLIMB CONDITION....................................................................................38
8.4. THRUST MATCHING ................................................................................................................................39
9.0. WING LOADING ......................................................................................................................................40
9.1. WING LOADING AT STALL SPEED ...........................................................................................................40
9.2. WING LOADING AT TAKEOFF ..................................................................................................................41
9.3. WING LOADING AT LANDING...................................................................................................................41
9.4. WING LOADING AT CRUISE.....................................................................................................................42
9.5. WING LOADING AT LOITER .....................................................................................................................42
9.6. WING LOADING AT INSTANTANEOUS TURN.............................................................................................42
10.0. INITIAL SIZING ........................................................................................................................................44
10.1. RUBBER ENGINE SIZING.........................................................................................................................44
10.1.1. Empty Weight Fraction ............................................................................................................44
10.1.2. Fuel Weight................................................................................................................................45
10.2. GEOMETRY SIZING .................................................................................................................................48
10.2.1. Fuselage .....................................................................................................................................48
10.2.2. Wing ............................................................................................................................................49
10.2.3. Tail...............................................................................................................................................50
10.2.4. Horizontal Tail............................................................................................................................51
10.2.5. Vertical Tail ................................................................................................................................52
11.0. CREW STATION, PASSENGERS, AND PAYLOAD............................................................................53
11.1. CREW STATION ......................................................................................................................................53
11.2. PILOT SIZES ...........................................................................................................................................54
11.3. SEATBACK ANGLE..................................................................................................................................55
11.4. OVER-NOSE VISION.................................................................................................................................55

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11.5. VISION ANGLE LOOKING UPWARD ...........................................................................................................57
11.6. TRANSPARENCY GRAZING ANGLE..........................................................................................................57
11.7. PASSENGER COMPARTMENT ..................................................................................................................58
11.8. CARGO PROVISIONS...............................................................................................................................60
12.0. PROPULSION AND FUEL SYSTEM .....................................................................................................61
12.1. ENGINE...................................................................................................................................................61
12.1.1. Engine Type...............................................................................................................................61
12.1.2. Engine Location ........................................................................................................................62
12.1.3. Inlet and Nozzle.........................................................................................................................63
12.2. FUEL SYSTEM.........................................................................................................................................65
12.2.1. Fuel System Integration...........................................................................................................65
12.2.2. Fuel Pump System....................................................................................................................65
13.0. LANDING GEAR......................................................................................................................................66
13.1. LANDING GEAR ARRANGEMENT .............................................................................................................66
13.2. TIRE SIZING ............................................................................................................................................67
13.3. SHOCK ABSORBERS...............................................................................................................................70
13.4. OLEO STRUTS.........................................................................................................................................71
13.5. GEAR RETRACTION GEOMETRY .............................................................................................................72
13.6. AIRCRAFT SUBSYSTEMS ........................................................................................................................72
14.0. COMPONENT WEIGHTS........................................................................................................................73
14.1. WEIGHTS REPORTING AND CG ESTIMATION...........................................................................................73
14.2. APPROXIMATE WEIGHT METHODS..........................................................................................................73
14.3. STATISTICAL WEIGHT METHODS ............................................................................................................74
14.3.1. Cruise..........................................................................................................................................77
14.3.2. Landing.......................................................................................................................................77
15.0. STABILITY, CONTROL AND HAND QUALITIES ................................................................................79
15.1. COORDINATE SYSTEMS AND DEFINITIONS ...............................................................................................80
15.2. STABILITY AXIS SYSTEM.........................................................................................................................81
15.3. LONGITUDINAL STATIC STABILITY AND CONTROL....................................................................................81
15.4. STATIC MARGIN......................................................................................................................................82
16.0. FINANCIAL ANALYSIS...........................................................................................................................83
16.1. LIFE-CYCLE COST ELEMENTS.................................................................................................................83
16.3. OPERATIONS AND MAINTENANCE COSTS ...............................................................................................84
16.4. FUEL AND OIL COST...............................................................................................................................84
18.0. CONSIDERING THE HEALTH, SAFETY, ECONOMIC, AND ENVIRONMENTAL IMPACT OF THE
SPECIFIC AIRCRAFT DESIGN ;.........................................................................................................................86
19.0. CONCLUSION..........................................................................................................................................86
20.0. REFERENCES ........................................................................................................................................87

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LIST OF FIGURES
Figure 1. Mission Profile (Business Jet)______________________________________________________________ 15
Figure 2. Empty weight fraction trends ___________________________________________________________ 17
Figure 3. Wetted Area Ratios___________________________________________________________________ 19
Figure 4. Maximum lift to drag ratio trends________________________________________________________ 20
Figure 5. Airfoil geometry ______________________________________________________________________ 22
Figure 6. Pressure distribution on airfoil__________________________________________________________ 23
Figure 7. Airfoil flow field and circulation _________________________________________________________ 23
Figure 8. Flow separation ______________________________________________________________________ 24
Figure 9. Airfoil cross section _____________________________________________________________________ 24
Figure 10. Thickness Analysis of GIII BL45 Airfoil ______________________________________________________ 25
Figure 11. Lift Characteristic slope _________________________________________________________________ 25
Figure 12. Drag characteristic curve________________________________________________________________ 26
Figure 13. Moment Characteristic Curve ____________________________________________________________ 26
Figure 14: Type of airfoils ______________________________________________________________________ 27
Figure 15: Types of stall _______________________________________________________________________ 28
Figure 16. Effect of t/c on drag__________________________________________________________________ 29
Figure 17. Effect of t/c on critical Mach number ___________________________________________________ 29
Figure 18. Effect of t/c on maximum lift __________________________________________________________ 30
Figure 19. Thickness ratio historical trend ________________________________________________________ 30
Figure 20: Range of Aspect Ratios ______________________________________________________________ 31
Figure 21. Effect of taper on lift distribution _______________________________________________________ 33
Figure 22. Wing incidence angle ________________________________________________________________ 33
Figure 23. Thrust lapse at cruise________________________________________________________________ 38
Figure 24. Maximum lift coefficient ______________________________________________________________ 40
Figure 25. Takeoff distance estimation___________________________________________________________ 41
Figure 26. Statistically improved Empty weight fraction equation ____________________________________ 44
Figure 27. Falcon 7x main fuselage layout _______________________________________________________ 53
Figure 28. Falcon 7x cockpit interiors____________________________________________________________ 54
Figure 29. Pilot RH and LH seat (Anon., n.d.)_____________________________________________________ 55
Figure 30: Aircraft over-nose angle (Anon., n.d.) __________________________________________________ 55
Figure 31. Falcon 7x nose-over angle (Anon., n.d.) ________________________________________________ 56
Figure 32. Vision angle looking upwards _________________________________________________________ 57
Figure 33. Falcon 7x grazing angle______________________________________________________________ 57
Figure 34. Cabin width and height_______________________________________________________________ 59
Figure 35. Cabin passenger compartment segments and cabin length (Anon., 2019)___________________ 59
Figure 36. Aircraft full layout____________________________________________________________________ 60
Figure 37. Baggage flat pallet (Anon., 2019)______________________________________________________ 60
Figure 38. high by-pass turbofan________________________________________________________________ 62
Figure 39. Falcon 7X__________________________________________________________________________ 63
Figure 40. Pitot/Normal Shock Inlet______________________________________________________________ 64
Figure 41. Fuel Pump System __________________________________________________________________ 65
Figure 42. Landing gear arrangement ___________________________________________________________ 66
Figure 43. Tricycle Geometry___________________________________________________________________ 67
Figure 44. Tire size table ______________________________________________________________________ 68
Figure 45, Tire Data___________________________________________________________________________ 69
Figure 46. Tire Pressure _______________________________________________________________________ 69
Figure 47. Shock Absorber_____________________________________________________________________ 70
Figure 48. Oleo Struts_________________________________________________________________________ 71
Figure 49. Oleo shock absorber machine ________________________________________________________ 71
Figure 50. Gear place _________________________________________________________________________ 72
Figure 51. Static and Dynamic Stability _____________________________________________________________ 79
Figure 52. Aircraft Coordinate System ______________________________________________________________ 80
Figure 53. Longitudinal Moments__________________________________________________________________ 81
Figure 54. Life-cycle Cost Elements. ________________________________________________________________ 83

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LIST OF TABLES
Table 1. Our Aircraft Specifications............................................................................................................................12
Table 2. Competitor Comparison Statistics.................................................................................................................13
Table 3. Defined Parameters for Mission Profile.........................................................................................................16
Table 4. Empty weight fraction vs Wo ..................................................................................................................17
Table 5. Historical mission segment weight fractions..........................................................................................18
Table 6. Specific fuel consumption .......................................................................................................................19
Table 7 L/D relation to (L/D)max...........................................................................................................................20
Table 8. Types of airfoil thickness and its effects ................................................................................................28
Table 9. Statistical equation for fuselage length ..................................................................................................48
Table 10. Cabin dimensions of Falcon 7x ............................................................................................................58
Table 11. Approximate Empty Weight Buildup...........................................................................................................73

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1.0. Introduction
In 1930’s Private jets commenced as a private affair and they used to accommodate only 2
passengers. Moreover, in the 1950’s Lockheed introduced the first Business jet in the
industry which accommodated about 10 passengers and 2 crew members. As the private
jets became a norm of air travel and due to increased demand, more manufacturing
companies such as Cessna, Embraer, Gulfstream, Beechcraft corporate jets, Bombardier
etc. joined in to create their own versions of Private jets which are classified into various
categories, Light jets, Mid-sized jets, Super mid-sized jets Large jets, Long range jets and
VIP airliners. At present these companies are competing in terms of making their private
jets more efficient, safe and more luxurious. Furthermore, the most notable advancement
by private jet manufacturers is supersonic jets using biofuel and long-range business jets.
Aircraft design is the engineering process to design an aircraft that meet the customers,
manufacturers and safety demands and should be cost efficient at the same time. Aircraft
design has 3 stages which are Conceptual design, which involves sketching the 3 views of
the aircraft and setting up the design layout and configurations. Then comes Preliminary
design phase which is testing of the aircraft design using wind tunnels and computational
fluid dynamics and all the structure and aerodynamics calculations are done then the
manufacturer decides whether to go ahead with the design or not. Lastly, it’s Detail design
phase which decides the number of structural elements in the aircraft and finalization of the
aircraft elements.
For this project we will be focusing on the ‘Conceptual design’ of a Business jet aircraft
“King Air B-100” and modifying it. “Aircraft design – A Conceptual Approach” written by
Daniel. P. Raymer was the main source of reference of this project.
2.0. Objectives
1. Design a Business jet in its conceptual phase based on market research and
historical trends.
2. To estimate and calculate initial aircraft Takeoff weight, thrust to weight ratio, wing
loading and other dimensions of our chosen aircraft.
3. Estimate dimensions, weight of all parts of the aircraft and the centre of gravity or the
whole aircraft
4. To sketch all 3 top, front and side views of our designed aircraft including its
dimensions.

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3.0. Market Research
Market research is the initial process of designing an aircraft. Getting to understand the
market situation, trends and patterns along with historical trends is very crucial in the
designing process of an aircraft as it saves a great deal of time and hence process of
iteration reduces. Below is our aircraft’s specifications:
Table 1. Our Aircraft Specifications
Manufacturer Dassault
Variant Falcon 7X
Crew 2 pilots + 1 crew
Capacity 12 to 16 passengers
Headroom 1.88 m
Cabin width 2.34 m
Cabin length 11.91 m
Height 23.38 m
Wingspan 7.83 m
Wing area 70.7 𝑚R
Wing loading 449 kg/𝑚R
Max takeoff weight 31,751 kg
Max landing weight 28,304 kg
Max zero fuel weight 18,598 kg
Fuel capacity 14,488 kg0
Basic operating weight 16,600 kg
Turbofan Pratt & Whitney Canada
PW307A
Thrust 28.48 KN
Range 11,019 km
ceiling 15,545 m
Max speed 956 km/h Mach(0.9)
Cruise speed 850 km/h Mach(0.8)
Approach 193 km/h
Landing 631 m
Takeoff (BFL,MTOW) 1,750 m
Avionics Falcon EASy flight Deck

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Below is the table of parameters of various manufacturers who are competitors to our
aircraft.
Table 2. Competitor Comparison Statistics
Manufacturer Bombardier Bombardier Gulfstream Gulfstream
Aircraft model Global 6000 Global 5000 G-600 G-550
Date 1993 1993 2002 2003
Range 39495000ft 31596000ft 39501312ft 41013780ft
Capacity Up to 19 Up to 13 Up to 19 14- 19
Top speed 865 ft/s 865 ft/s 894 ft/s 823 ft/s
Cruise speed 850 ft/s 850 ft/s 870 ft/s 774 ft/s
Takeoff
distance
5900 ft 5000 ft 5900 ft 5910 ft
Landing
distance
3100 ft 2207 ft 3100 ft 2770 ft
Maximum
operating
altitude
51000 ft 43000 ft 51000 ft 51000 ft
Initial cruise
altitude
41000 ft 51000 ft 41000 ft 41000
Engine Two Pratt &
Whitney
Canada
PW815GA
Rolls Royce
Deutschland
BR710A2-20
Prat&
Whitney
Canada
PW800
Rolls Royce
BR700
Thrust 14750 lb 14750 lb 15144 lb 15000 lb
Mach Number 0.925 0.85 0.925 0.885
Length 99 ft 5 in 96 ft 10 in 91 ft 2 in 96 ft 5 in
Passengers 13 13 19 16
Crew 4 3 3 3
The design process begins after thorough comparison between the competitor planes listed
in Table 2. above.

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4.0. Mission Profile
The journey of an aircraft from takeoff till landing is called mission profile. It is an integral
part of Aircraft design as gives direction and stability to our designing process. Mission
profile of any aircraft consists of takeoff, climb, cruise, descent, loiter and landing. In our
case the mission profile of our aircraft is very similar to a conventional aircraft.
In our case the mission is to takeoff smoothly, then it climbs. Moreover, it stays in cruise
condition for a while until it reaches its destination with stability and then it descends
gradually. If there are any environmental or air traffic concerns and the aircraft has to wait,
it can loiter for around 30 minutes. After all that it lands safely. In this case the passengers
would be VIP’s so the comfort of all the passengers is paramount for the pilot and the
airlines.

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The mission profile shown above is divided into various segments from 1 to
so below is the detail of each segment.
SEGMENT NUMBER AIRCRAFT POSITION
0-1 Take off
1-2 Climb
2-3 Cruise
3-4 Descend
4-5 Loiter
5-6 Landing
Lastly, the mission profile assists us in finding the takeoff weight and some
other values of the aircraft by letting us calculate Fuel weight in fractions of
each of its segments.
TAKEOFF
CRUISE
LOITER
LANDING
Figure 1. Mission Profile (Business Jet)
CLIMB
0
1
2 3
4
5
6

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5.0. Preliminary Sizing
Some parameters needed to be defined for the mission profile before starting the
preliminary sizing, as listed in table 3 below.
Table 3. Defined Parameters for Mission Profile
PARAMETER VALUE
Range 36154856 ft.
Cruise Speed 774 ft/s.
Aspect Ratio 6.3
Endurance 1800s
5.1. Take off Gross Weight
𝑊T =
𝑊VWXY + 𝑊[]^_`
1 − c
𝑊d
𝑊T
e − f
𝑊X
𝑊T
g
Where
f
hi
hj
g = 1.06 f1 −
hnop
hj
g
f
hn
hj
g = 𝐴𝑊T
r
𝐾tu
• The Falcon 7X carries 2 crew members, so according to FAR the average person
weighs 175lb; therefore 𝑊VWXY = 2 ∗ 175 = 350𝑙𝑏 for our aircraft.
• The Falcon 7X has the capacity to carry a total number of 12 passengers and
assuming each passenger carries 30lb baggage, and according to FAR the average
person weighs 175lb; therefore 𝑊[]^_` = 12 ∗ (175𝑙𝑏 + 30𝑙𝑏) = 2460𝑙𝑏 for our
aircraft.

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5.1.1. Estimating Empty Weight Fuel Fraction f
𝑾 𝒆
𝑾 𝟎
g
Figure 2. Empty weight fraction trends
According to historical trends graph shown in figure 2, we assume 𝑊T to be 60,000lb for our
jet transport aircraft.
Table 4. Empty weight fraction vs Wo
By using the formula given in Table 4 and since King Air B100 is a jet transport aircraft, we
get the values:
𝐴 = 1.02
𝐶 = −0.06
𝐾tu = 1.00, since the aircraft wings are fixed sweep
c
𝑊X
𝑊T
e = 𝐴𝑊T
r
𝐾tu = 1.02 ∗ (𝑊T)‚T.Tƒ
∗ 1.00 = 𝑊T
‚T.Tƒ

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5.1.2. Estimating Fuel Fraction f
𝑾 𝒇
𝑾 𝟎
g
According to our mission profile shown in figure 1 we have fuel fraction from point 0-6, which
includes take-off, climb, cruise, loiter and landing.
𝑊ƒ
𝑊T
= c
𝑊ƒ
𝑊„
e
^…`†…‡
∗ c
𝑊„
𝑊ˆ
e
^_†‰XW
∗ c
𝑊ˆ
𝑊Š
e
`XuVX…‰
∗ c
𝑊Š
𝑊R
e
VW‹†uX
∗ c
𝑊R
𝑊Œ
e
V^†•Ž
∗ c
𝑊Œ
𝑊T
e
‰•X_dd
Table 5. Historical mission segment weight fractions
By using Table 5, we get the values:
f
h•
h‘
g
^…`†…‡
= 0.995
f
h“
h”
g
V^†•Ž
= 0.985
f
h”
hj
g
‰•X_dd
= 0.970

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5.1.2.1. Cruise and Loiter Fuel Fraction
Table 6. Specific fuel consumption
• From table 6 we get the values for the specific fuel consumption during cruise and
loiter, since our engine is a high bypass turbofan then SFC of cruise is 0.5/hr and
SFC of loiter is 0.4/hr.
𝐶VW‹†uX =
0.5
3600
= 1.38 ∗ 10‚ˆ
/𝑠𝑒𝑐
𝐶^_†‰XW =
0.4
3600
= 1.11 ∗ 10‚ˆ
/𝑠𝑒𝑐
Figure 3. Wetted Area Ratios
• Using figure 3, our
™šn›
™œni
will be;
𝑆YX‰
𝑆WXd
= 5.5

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Figure 4. Maximum lift to drag ratio trends
• To find (L/D)max we need to calculate the wetted aspect ratio:
𝐴𝑅/(𝑆𝑤𝑒𝑡/𝑆𝑟𝑒𝑓) = 6.3/5.5 = 1.145
• Using figure 4, (L/D)max= 15.9
Table 7 L/D relation to (L/D)max
c
𝐿
𝐷
e
VW‹†uX
= 0.866 ∗ 15.9 = 13.7694
c
𝐿
𝐷
e
^_†‰XW
= 15.9
• To get the fuel fraction value of the cruise segment the Breguet range equation will
be used:
𝑊Š
𝑊R
= 𝑒
‚
¥r
¦f
§
¨
g
= 𝑒
‚
©T„ˆˆƒR∗Œ.Š©∗ŒT^‚ˆ
««ˆ(ŒŠ.«ƒ¬ˆ) = 0.6261
15.9
1.145

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• To get the fuel fraction value of the loiter segment the Breguet endurance equation
will be used:
(Assuming 30 minutes loiter, the endurance value will be 1800 seconds.)
𝑊„
𝑊ˆ
= 𝑒
‚
-r
f
§
¨
g
= 𝑒
‚
Œ©TT∗Œ.ŒŒ∗ŒT^‚ˆ
(Œ„.¬) = 0.9875
• Therefore, the fuel fraction weight is:
c
𝑊d
𝑊T
e = 1.06 c1 −
𝑊X…`
𝑊T
e
Where,
hnop
hj
= f
h•
hj
g = 0.995 ∗ 0.9875 ∗ 0.6261 ∗ 0.985 ∗ 0.970 = 0.5877
c
𝑊d
𝑊T
e = 1.06(1 − 0.5877) = 0.437
• Substituting all the values to calculate W0:
𝑊T =
𝑊VWXY + 𝑊[]^_`
1 − c
𝑊d
𝑊T
e − f
𝑊X
𝑊T
g
=
350 + 2460
1 − 0.437 − 𝑊T
‚T.Tƒ = 69427𝑙𝑏
We obtain the value of 69427lb after several iterations by assuming different values of W0.
• The Aspect ratio we used was 6.3 which we obtained after keeping all values of
Number of passengers, 𝑊T,
™šn›
™œni
the same and changing the values of Aspect Ratio
until we obtain the value of Maximum Takeoff 𝑊T close to 70000lb. Below is the table
which illustrates the various values we obtained from using different Aspect ratios.
NUMBER OF PASSENGERS 𝑺 𝒘𝒆𝒕
𝑺 𝒓𝒆𝒇
ASPECT RATIO 𝑾 𝟎
12 5.5 7.5 43495 lb.
12 5.5 6.3 69427 lb.
12 5.5 6.0 62575 lb.
12 5.5 5.8 75305 lb.

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6.0. Wing Design
The aircraft wing is a major contributor of lift for the aircraft, it’s the main control surface of
the aircraft. Therefore, the selection of wing design of an aircraft is very critical. Wing design
can be performed by taking certain parameters such as thickness ratio of airfoil, aspect
ratio, sweep angle, taper ratio, wing twist, wing tips and wing vertical location on fuselage.
6.1. Airfoil Selection
The airfoil is the heart of the aircraft and hence its selection affects the takeoff and landing
distances, cruise speed, stall speed, handling qualities and overall aerodynamic efficiency.
6.2. Airfoil Geometry
Airfoil geometry is the labelled parts on the airfoil, shown in figure below:
Figure 5. Airfoil geometry
Terminology:
• Leading edge: The front section of the airfoil
• Trailing edge: The rear section of the airfoil.
• Chord length: The distance from leading edge to trailing edge of the airfoil.
• Camber: Curvature characteristic of airfoil.
• Mean camber line: The line at equidistant from the upper and lower surfaces of airfoil.
• Total airfoil camber: Maximum distance of the mean camber line from the chord line.
• Thickness: distance from upper surface to lower surface, measured perpendicular to
mean camber line.
• Thickness ratio: Maximum thickness of the airfoil divided by its chord.

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6.3. Airfoil Lift and Drag
For an airfoil to generate lift it needs air at different velocities and pressure at its upper and
lower surface. Bernoulli’s law illustrates that air at high velocity has low pressure and vice
versa. Hence airfoil is designed in such a way that the bottom surface is flat and upper
surface is curved so that high pressure air pushes the low-pressure air at the top to generate
lift. Moreover, the airfoil angle of attack or camber generates air at the top of the wing to
travel quicker than the bottom of the wing. Figure below shows the pressure distribution on
upper and lower surfaces of a lifting airfoil at subsonic speeds.
Figure 6. Pressure distribution on airfoil
Flow field around an airfoil is a number of airflow velocity vectors. The vector length
illustrates the magnitude of the local velocity vector. The figure below shows the cause of
airfoil effect which alters the airflow which starts to circulate around the airfoil. These
calculations are crucial for classical calculations of lift and induced drag.
Figure 7. Airfoil flow field and circulation

23-05-2019 24 DR. ELHAM
Flow separation occurs on an airfoil when angle of attack is increased at high speeds thus
disturbing the flow and eventually reducing lift, to prevent that we curve the airfoils so flow
remains attached and produces lift.
Figure 8. Flow separation
6.4. Airfoil Selection and Design
Airfoils were developed in 1930’s by NACA. They were first called four-digit airfoils whereas,
the first digit defined as percent camber, second digit as location of maximum camber and
last two digits as maximum thickness of percent of chord in airfoil. Then five digits and six
digits airfoils were built for maximum lift and maximum camber forward. Later more modern
airfoil models were developed to comply with modern applications.
The airfoil chosen is the GIII BL45 airfoil, which is generally used on the Gulfstream aircrafts.
This airfoil is useful in terms of Mach number for business jets at transonic speeds.
Furthermore, this airfoil reduces the transonic drag and wave that a business jet experience.
Figure 9. Airfoil cross section
Figure below illustrates the thickness analysis of the airfoil mentioned below:

23-05-2019 25 DR. ELHAM
Figure 10. Thickness Analysis of GIII BL45 Airfoil
Figures below illustrate the Lift, Drag and Moment characteristics curves of the
BL45 airfoil which is chosen to be the airfoil for the aircraft wing.
Figure 11. Lift Characteristic slope

23-05-2019 26 DR. ELHAM
Figure 12. Drag characteristic curve
Figure 13. Moment Characteristic Curve

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Figure 14: Type of airfoils

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6.5. Stall
Stall characteristics play a crucial role in airfoil selection, some airfoils stall gradually while
others stall show a rough loss in lift during stall. There are 3 different types of stall:
1. Thick airfoils: Stall starts from trailing edge as angle of attack is increased, separation
starts from trailing edge and then gradually goes to leading edge.
2. Thin airfoils: Abrupt flow separation at high angle of attack at leading edge of wing
and flow fails to re-attach.
3. Very thing airfoil: At increased angle of attack, prior to flow separation at the leading
edge with flow reattaching back.
Below are the types of stall on an airfoil:
Figure 15: Types of stall
6.6. Airfoil thickness ratio
Airfoil thickness has a direct effect on stall characteristics, maximum lift and structural
weight, the table below shows the effects of the 2 types of airfoils:
THICK WINGS THIN WINGS
1. Huge drag due to increased
separation and bigger boundary
layer.
1. Less drag due to less airflow
separation.
2. Better stall properties as stall
begins from trailing edge of wing.
2. Stall begins from leading edge.
3. Thick airfoils add less weight to
aircraft structure according to
formula.
3. Thin airfoils add less weight to
structure of aircraft.
Table 8. Types of airfoil thickness and its effects

23-05-2019 29 DR. ELHAM
Figure 16 below shows the effect of thickness ratio on drag:
Figure 16. Effect of t/c on drag
Figure 17 below shows the effect of thickness ratio on critical Mach number:
Figure 17. Effect of t/c on critical Mach number

23-05-2019 30 DR. ELHAM
Figure 18 below shows the effect of t/c on maximum lift:
Figure 18. Effect of t/c on maximum lift
Figure. 19 below shows the effect of t/c on Design Mach number:
Figure 19. Thickness ratio historical trend
Proposed Reynolds number is another significant aspect in airfoil selection, as each airfoil
is designed for a specific Reynolds number, and for example the laminar airfoils demand
extremely smooth skins.

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6.7. Aspect Ratio
Aspect ratio of the wing is the ratio of the span of the wing to the mean chord of the wing.
The formula of the aspect ratio is given below:
𝐴𝑅 =
𝑏R
𝑆
There are 2 types of aspect ratios, a high aspect ratio wing is long and narrow, and a low
aspect ratio wing is short of length and wide. However, there are pros and cons of both
shown in table below:
HIGH ASPECT RATIO LOW ASPECT RATIO
1. Less wing tip vortices as tips are
far from Center of gravity and root
of wing so less strength of
vortices.
1. High strength of wing tip vortices
as they are closer to wing root and
center of gravity.
2. More lift produced as bigger wing
area.
2. Less lift produced as less wing
area.
3. Increased drag as wing tip loads
up due to increased lift.
3. Less drag due to less lift on wing.
4. Less maneuverability as wing is
too long.
4. More maneuverability due to short
wings span.
Figure below shows aircrafts with different aspect ratios:
Figure 20: Range of Aspect Ratios

23-05-2019 32 DR. ELHAM
6.8. Wing Sweep
Wings are swept back so wave drag across wing is reduced as incoming airflow is
distributed in two directions, Span-wise and Chord-wise flow. Other reasons of sweeping
the wing is to increase lateral stability and to delay critical Mach number. Table below shows
the properties of swept and un-swept wing:
SWEPT WINGS UNSWEPT WINGS
1. Reduces wave drag and flow
separation at high speeds and
high angle of attack by dividing
incoming airflow in two flows.
1. More drag as incoming airflow just
flows in one direction.
2. Creates dihedral effect by adding
stability when there is sideslip, it
provides extra lift on lower wing to
make aircraft stable.
2. Less stability of wing when in Side
or Forward slip.
3. Not great for stall as stall begins
from wing tips.
3. Takeoff and stall better as flow
separation begins from wing root.
4. Critical Mach number is delayed
as flow separation is delayed.
4. Critical Mach Number achieved
early and due to that there is wave
drag when flying at high speeds.

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6.9. Taper Ratio
Taper ratio of a wing is the ratio between the chords of the root to the tip of the wing.
Moreover, taper ratio can adopt various forms such as planform taper, platform taper,
thickness taper and inverted thickness. Commercial airlines use taper ratios in between 0.2-
0.5. The reason we taper the wing is, when the wing is un-tapered the lift distribution across
the wing is the same hence at the wing tips extra lift loads up which eventually increases
induced drag by 7%, therefore hampering the performance of the aircraft. The easiest form
of wing to make is a rectangular wing as it doesn’t require very complex designing, but an
ideal type of wing is an elliptical wing which isn’t practical to build as it’s very expensive.
Figure 21 below shows the effect of taper on lift distribution:
Figure 21. Effect of taper on lift distribution
6.10. Wing twist
A wing is twisted to improve its stall characteristics. The main aim of twisting the wing is to
reverse the lift distribution so that the root of the wing stalls before the wing tip for a smooth
and gradual stall. There are two types of twists, Geometric twist where a different airfoil
which is twisted is used on the wingtip and a less twisted airfoil is used on the wing root so
the wing root stalls first. Secondly, Aerodynamic twist is using various angle of incidences
of airfoils on the wing.
6.11. Wing incidence
It is the angle between the fuselage axis and the chord of the wing airfoil.
Figure 22. Wing incidence angle

23-05-2019 34 DR. ELHAM
6.12. Wing Vertical Location
Wing vertical position have three main configurations High wing, medium wing and low wing.
Table shows the trade-offs of having each wing configuration:
ADVANTAGES DISADVANTAGES
HIGH WING
1. Allows placing the fuselage closer to
ground so military aircrafts which carry
heavy payloads can easily load and
un-load.
1. Fuselage weight increased as it needs
to be strengthened to support landing
gear loads. This weight adds extra
drag.
2. Jet engines or propellers will have
sufficient ground clearance without
excessive landing gear length.
2. Blocks upward visibility of pilot when
in a climb.
MID WING
1. Allows good visibility for the pilot when
carrying ammunition under the wing.
1. Not feasible for passenger and cargo
aircrafts as mid wing’s box is attached
at the middle of the fuselage which
minimizes the area inside the fuselage
for passengers or baggage.
2. Best for aerobatic maneuver-ability. 2. Wing box connected by huge ring
frames into fuselage adding more weight.
LOW WING
1. More space for passengers and cargo
inside fuselage.
2. Difficult ground clearance for aircraft
with low wings.
3. Landing gear directly attached to wing
box which doesn’t need any extra
strengthening.
3. Low wing tips mean they can make
landing and takeoff difficult in order for
tips to touch the ground.
6.13. Wing Tips
Wing tips depend on the mission of the aircraft. The various configurations are rounder,
sharp, un-swept, cut-off, drooped, endplate, winglet etc. Their main function is to reduce
aerodynamic drag due to reducing the strength of wing tip vortexes. Their significant
reduction in Induced drag eventually leads to improved aircraft lift performance, increase in
cruise speed and saving fuel.

23-05-2019 35 DR. ELHAM
7.0. Tail Configuration
In an aircraft a tail is added as an additional control surface to increase stability across the
aircraft. The function of the tail is to create an opposite moment to the wing which levels the
aircraft to neutral position. However, choosing the right tail configuration is very crucial as
each type of aircraft has a different mission and requires different performance parameters.
Below is the characteristics of various tail configurations:
TAIL
CONFIGURATION
CHARACTERISTICS IMAGE
Conventional
• Easy to design and modify
• Provides adequate control and stability
at the lightest weight.
• Easy to mechanize control linkages
due to adequate structure in fuselage
where HT is attached.
T-Tail
• Has plenty of ground clearance.
• Provides end-plate effect to Vertical tail
which reduces size of vertical tail.
• Suitable configuration for rear fuselage
mounted engine.
• Vertical tail must be strengthened
which adds more mass to the overall
aircraft.
Cruciform Tail
• Compromise between conventional
and T-tail.
• Prevents the lower part of the rudder to
be exposed to undistributed flow.
• Less weight penalties as no
strengthening is needed.
• Doesn’t provide end plate effect.
H-Tail
• Positions the vertical tail in undisturbed
air during high airspeeds and angle of
attacks.
• Provides end-plate effect to horizontal
tail.
• Heavier than conventional tail.
Triple-Tail
• Similar to H-tail, used on Lockheed
constellation.

23-05-2019 36 DR. ELHAM
V-Tail
• Reduces wetted area when compared
to separate VT and HT.
• Offers reduced interference drag.
• Complex model as rudders and
elevators work together.
• Adverse roll yaw coupling effect.
Inverted V-Tail
• Prevents the issue of adverse roll yaw
coupling and provides pilot with
Proverse roll yaw coupling.
• Doesn’t provide adequate ground
clearance.
Y-Tail
• Similar to V-tail but rudders and
elevators are separated to reduce
complexity in system.
• Y-tail acts as a skid stopping the
propellers from hitting the ground.
• Adding a third tail adds more weight to
the aircraft.
Twin-Tail
• Reduces height of a single vertical tail.
• More effective in high alpha maneuver.
• Heavier than a single vertical tail.

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8.0. Thrust to Weight ratio
The Thrust to weight ratio is the ratio between engine of the aircraft to its weight. This ratio
varies in each aircraft. The T/W affects the performance of the aircraft directly, the higher
its value the more quickly it’ll accelerate, climb more rapidly, reach a higher maximum
speed and climb more quickly.
For the business jet we’ll first calculate the T/W values using takeoff weight value we
calculated before, then we’ll calculate the wing loading.
There are three equations to calculate T/W ratio at different stages of our aircraft flight, out
of the three values the largest value should be considered.
8.1. Thrust to Weight ratio at cruise condition
The value of (L/D)cruise was taken from our (L/D)max value which we obtained while
calculating the takeoff weight, W0.
c
𝑇
𝑊
e
VW‹†uX
=
1
f
𝐿
𝐷
g
VW‹†uX
=
1
13.7694
= 0.07262

23-05-2019 38 DR. ELHAM
8.2. Thrust to Weight ratio at takeoff condition
The value of (W cruise /W takeoff) is calculated by multiplying the takeoff and climb weight
fractions shown in the table above.
Figure 23. Thrust lapse at cruise
The value of (T takeoff /T cruise) is obtained from the graph above by assuming the altitude at
50,000ft and the engine as high BPR turbofan. The value taken from the graph should be
reciprocated as we require (T takeoff /T cruise).
c
𝑇
𝑊
e
‰•X_dd
= c
𝑇
𝑊
e
VW‹†uX
¯
𝑊VW‹†uX
𝑊‰•X_dd
°c
𝑇‰•X_dd
𝑇VW‹†uX
e = (0.07262)(0.955)c
1
0.23
e = 0.3015
8.3. Thrust to Weight ratio at climb condition
The value of Vvertical and V was taken from the aircraft specification sheet, a value of
2055fpm. The value of (L/D)climb is equal to 90% of (L/D)max value. The value of Velocity is
converted to fpm, it becomes 774 ∗ 60 = 46440𝑓𝑝𝑚
c
𝑇
𝑊
e
V^†•Ž
=
1
f
𝐿
𝐷g
V^†•Ž
+
𝑉tXW‰†V^
𝑉
=
1
14.31
+
2055
46440
= 0.1141

23-05-2019 39 DR. ELHAM
8.4. Thrust Matching
1. f
³
h
g
VW‹†uX
= (0.07262) ∗ f
h“
h”
g
V^†•Ž
∗ f
h”
hj
g
‰•X_dd
= (0.07262)(0.970)(0.985) = 0.069
𝑇VW‹†uX = f
³
h
g
VW‹†uX
∗ 𝑊T = (0.069)(69427) = 4811𝑙𝑏 for both engines
2. 𝑇‰•X_dd = (0.3015)(69427) = 20932𝑙𝑏 for both engines
3. f
³
h
g
V^†•Ž
= (0.1141) ∗ f
h”
hj
g
‰•X_dd
= (0.1141)(0.970) = 0.1106
𝑇V^†•Ž = (0.1106)(69427) = 7683𝑙𝑏 for both engines
T/W RATIOS VALUES
c
𝑻
𝑾
e
𝒄𝒓𝒖𝒊𝒔𝒆
0.07262
c
𝑻
𝑾
e
𝒕𝒂𝒌𝒆𝒐𝒇𝒇 0.3015
c
𝑻
𝑾
e
𝒄𝒍𝒊𝒎𝒃
0.1141
• We will choose the highest value from the table as our (
³
h
), which is 0.3015.

23-05-2019 40 DR. ELHAM
9.0. Wing Loading
9.1. Wing Loading at Stall Speed
We assume the density as sea level condition, 𝜌=0.00237slug/ft3
. Vstall is assumed to be
70knots. The value of 𝐶§·¸¹
is obtained from the graph below for a double slotted fowler flap
at a wing sweep angle of 30°.
Figure 24. Maximum lift coefficient
c
𝑊
𝑆
e
u‰^^ u[XX`
=
1
2
𝜌𝑉u‰^^
R
𝐶§·¸¹
=
1
2
(0.00237)(70)R(2.5) = 14.5
When we compare our f
h
™
g
u‰^^ u[XX`
value to the table below, our value is far away from Jet
transport value (120) therefore, we neglect this value.
2.5

23-05-2019 41 DR. ELHAM
9.2. Wing Loading at Takeoff
The takeoff parameter is obtained from the figure below, at a takeoff distance of 5710ft.
Figure 25. Takeoff distance estimation
Assume 𝜎 = 1
𝑪 𝑳 𝑻𝑶
:
𝐶§·¸¹
= 𝐶§¼½¾ ¿¸opÀoÁ
= 2.5
𝐶§¼½¾›¸ÂnÃii
= 𝐶§¼½¾ ¿¸opÀoÁ
∗ 0.8 = (2.5)0.8) = 2
𝐶§ÄÅ
=
𝐶§¼½¾›¸ÂnÃii
1.21
=
2
1.21
= 1.65
c
𝑊
𝑆
e
‰•X_dd
= ( 𝑇𝑂𝑃) 𝜎𝐶§ÄÅ
c
𝑇
𝑊
e = (290)(0.1)(1.65)(0.3015) = 144.26
9.3. Wing Loading at Landing
𝑆^…`†…‡ = 80c
𝑊
𝑆
e ¯
1
𝜎𝐶§·¸¹
° + 𝑆
2070 = 80 c
𝑊
𝑆
e c
1
(0.1)(2.5)
e + 600
c
𝑊
𝑆
e
^…`†…‡
= 19.985

23-05-2019 42 DR. ELHAM
9.4. Wing Loading at Cruise
𝑞 =
Œ
R
𝜌𝑉R
, We will take cruising altitude as 50000ft and hence 𝜌 = 0.000364 𝑠𝑙𝑢𝑔/𝑓𝑡Š
Hence, 𝑞 =
Œ
R
(0.000364)(774)R
= 109𝑙𝑏/𝑓𝑡R
c
𝑊
𝑆
e
ËÌÍ ÎX‰ W…‡X
= 𝑞Ï
𝜋𝑒𝐴𝑅𝐶¨j
3
= (109)Ï
𝜋(0.8)(6.3)(0.015)
3
= 30.669
9.5. Wing Loading at Loiter
c
𝑊
𝑆
e
ËÌÍ ÎX‰ ^_†‰XW
= 𝑞Ñ 𝜋𝑒𝐴𝑅𝐶¨j
= (109)Ò𝜋(0.8)(6.3)(0.015) = 53.12
9.6. Wing Loading at Instantaneous Turn
To calculate the load factor n, a value of 1.5 was assumed for 𝜓 based on historical trends.
𝒏:
𝜓 =
𝑔√𝑛R − 1
𝑉
∗ 57.3
1.5 =
32.2√𝑛R − 1
774
∗ 57.3
𝑛 = 1.18
c
𝑊
𝑆
e
†…u‰…‰…X_‹u ‰‹W…
=
𝑞𝐶§·¸¹
𝑛
=
(109)(2.5)
1.18
= 230.93

23-05-2019 43 DR. ELHAM
WING LOADING VALUE
Wing Loading for Stall Speed 14.5
Wing Loading for Takeoff Distance 144.26
Wing Loading for Landing distance 19.985
Wing Loading for cruise 30.669
Wing Loading for Loiter Endurance 53.12
Wing loading for Instantaneous turn 230.93
Comparing all the values calculated, the lowest should be selected as our Wing Loading
value. However, the range of wing loading that is applicable for business jets in regard to
historical trends is between 20 and 200. Hence regarding the other values of W/S, the least
value between 20 and 200 is chosen. Thus:
h
™
= 30.669

23-05-2019 44 DR. ELHAM
10.0. Initial Sizing
Initial sizing is a more refined estimation of the aircraft takeoff weight and the fuel weight
when compared to the preliminary sizing. An aircraft can be sized using existing engine or
a new design engine. For the existing engine it is fixed in size and thrust, but for the new
design it can be built in any size and thrust required, it called a ‘’ rubber engine’ ’since the
empty weight was calculated using a guess of the takeoff weight, it was necessary to iterate
towards a solution.
10.1. Rubber Engine Sizing
𝑊𝑜 = 𝑊𝑐𝑟𝑒𝑤+ 𝑊𝑓𝑖𝑥𝑒𝑑𝑝𝑎𝑦𝑙𝑜𝑎𝑑 + 𝑊𝑑𝑟𝑜𝑝𝑝𝑒𝑑𝑝𝑎𝑦𝑙𝑜𝑎𝑑 + 𝑊𝑒𝑚𝑝𝑡𝑦 + 𝑊𝑓𝑢𝑒𝑙
10.1.1. Empty Weight Fraction
The Empty weight fraction in this section is more precise as compared to the value
estimated in preliminary sizing. It’s found from figure below:
Figure 26. Statistically improved Empty weight fraction equation
𝑊𝑒
𝑊𝑜
= ¯ 𝑎 + 𝑏 𝑊𝑜VŒ
𝐴VR
c
𝑇
𝑊
e
VŠ
c
𝑊𝑜
𝑆
e
Vˆ
𝑀𝑚𝑎𝑥V„
° 𝐾𝑣𝑠
𝑊𝑒
𝑊𝑜
= [0.32 + 0.66( 𝑊T)‚T.ŒŠ(6.3)T.ŠT(0.3015)T.Tƒ(19.985)‚T.T„(0.9)T.T„](1)
= 0.32 + 0.914𝑊T
‚T.ŒŠ
𝑊X = (0.32 + 0.914𝑊T
‚T.ŒŠ
)𝑊T

23-05-2019 45 DR. ELHAM
10.1.2. Fuel Weight
The fuel weight can be calculated by estimating the fuel weight fraction at every segment
and then adding all of them.
𝑊𝑓=1.06[∑ f1 −
hÀ
hÀá”
g 𝑊†‚Œ]
• Takeoff (Based on historical estimation)
𝑊1
𝑊𝑜
= 0.9800
• Climb and accelerate (Subsonic Flow)
𝑊2
𝑊1
= 1.0065 − 0.325𝑀
𝑊2
𝑊1
= 1.0065 − 0.325(0.9) = 0.977
• Cruise
§
¨
=
Œ
â
ã äDo
ç
è
é ê
ç
è
ê
”
ã ë ìn
=
Œ
f
(”jí)(j.j”‘)
(îj.••í)
g ê (ŠT.ƒƒ¬)ê
”
ë(”jí) (j.ï)(•.î)
𝐿
𝐷
= 14.28
𝑊3
𝑊2
= 𝒆
ð
‚𝑹 𝑪
𝑽 f
𝑳
𝑫 g
ñ
= 𝒆
c
(𝟑𝟔𝟏𝟓𝟒𝟖𝟓𝟔)(𝟏.𝟑𝟖∗𝟏𝟎−𝟒
)
(𝟕𝟕𝟒)( 𝟏𝟒.𝟐𝟖 )
e
= 0.636
• Loiter
𝑊4
𝑊3
= 𝑒
‚
- r
§
¨ = 𝑒
‚
(Œ©TT) (Œ.ŒŒ∗ŒTáô)
(Œˆ.R©) = 0.986

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• Descent (Based on Historical trends)
W„
Wˆ
= 0.9925
• Landing (Based on Historical trends estimation)
Wƒ
W„
= 0.995
Wƒ
WT
= (0.98)(0.977)(0.636)(0.9766)(0.995)
𝑊d,Œ = c1 −
𝑊Œ
𝑊T
e WT = (1 − 0.98)(69427) = 1388.54
𝑊Œ = 69427 − 1388.54 = 68038𝑙𝑏

𝑊d,R = c1 −
𝑊R
𝑊Œ
e WŒ = (1 − 0.977)(68038) = 1564.87
𝑊R = 68038 − 1564.87 = 66473.13𝑙𝑏
𝑊d,Š = c1 −
𝑊Š
𝑊R
e WR = (1 − 0.636)(66473.13) = 24196
𝑊Š = 66473.13 − 24196 = 42344𝑙𝑏
𝑊d,ˆ = c1 −
𝑊ˆ
𝑊Š
e WŠ = (1 − 0.9925)(42344) = 317.6
𝑊ˆ = 42344 − 317.6 = 42026𝑙𝑏
𝑊d,„ = c1 −
𝑊„
𝑊ˆ
e Wˆ = (1 − 0.986)(42026) = 588.36

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𝑊„ = 42026 − 588.36 = 41437.64𝑙𝑏
𝑊d,ƒ = c1 −
𝑊ƒ
𝑊„
e W„ = (1 − 0.995)(41437.64) = 207.185
𝑊Œ = 41437.64 − 207.185 = 41229𝑙𝑏
÷ 𝑊d = 1388.54 + 1564.87 + 24196 + 317.6 + 588.36 + 207.185 = 28262.55𝑙𝑏
𝑊d = 1.06 ÷ 𝑊d = 1.06(28262.55) = 29958.3𝑙𝑏
Eventually:
𝑊T = 𝑊VWXY + 𝑊d†øX` []^_` + 𝑊`W_[[X` []^_` + 𝑊X•[‰] + 𝑊d‹X^
𝑊T = (350) + (2460) + (0) + ù0.32 + 0.914𝑊T
‚T.ŒŠ
ú𝑊T + (29958.3)
𝑊T = 32768.3 + ù0.32 + 0.914𝑊T
‚T.ŒŠ
ú𝑊T
𝑊T = 70350𝑙𝑏
The weight found with more precise equations from the preliminary sizing is 81549lb which
is close to the value found earlier in initial sizing as 69427lb. This illustrates that all the
calculations done are based on a valid estimate of the initial weight.

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10.2. Geometry Sizing
10.2.1. Fuselage
Table 9. Statistical equation for fuselage length
Fuselage length
According to table 9 the fuselage length can be calculated.
𝜄d = 𝑎𝑊T
V
𝜄d = 𝑎𝑊T
V
= (0.67)(70350)T.ˆŠ
= 81.35𝑓𝑡
Fuselage fineness ratio
Fineness ratio =
𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒 𝑙𝑒𝑛𝑔𝑡ℎ
𝑚𝑎𝑥𝑖𝑚𝑢𝑚 𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒 𝑑𝑖𝑎𝑚𝑡𝑒𝑟
=
𝜄d
𝑑d
=
81.35
9.84
= 8.27
The typical range for subsonic flight is between 6 and 8 hence this value is close to it and
is accepted.

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10.2.2. Wing
𝑊𝑖𝑛𝑔 𝑟𝑒𝑓𝑒𝑟𝑒𝑛𝑐𝑒 𝑎𝑟𝑒𝑎 =
𝑇𝑎𝑘𝑒𝑜𝑓𝑓 𝑔𝑟𝑜𝑠𝑠 𝑤𝑒𝑖𝑔ℎ𝑡
𝑇𝑎𝑘𝑒𝑜𝑓𝑓 𝑤𝑖𝑛𝑔 𝑙𝑜𝑎𝑑𝑖𝑛𝑔
𝑆 =
𝑊T
𝑊T
𝑆
=
(70350)
(30.669)
= 2294.95 𝑓𝑡R
𝑏 = Ò(𝑠 ∗ 𝐴𝑅) = √(2294.95)(6.3) = 120.23𝑓𝑡
𝜆 = 0.4
𝐶W__‰ =
2𝑆
𝑏(1 + 𝜆)
=
2(2294.95)
(120.23)(1 + 0.4)
= 27.244𝑓𝑡
𝐶‰†[ = 𝜆𝐶W__‰
𝐶‰†[ = 𝜆𝐶W__‰ = (0.4)(27.244) = 10.9𝑓𝑡
𝑀𝐴𝐶 = 𝐶̅ = c
2
3
e 𝐶W__‰ ∗
1 + 𝜆 + 𝜆R
1 + 𝜆
= c
2
3
e (27.244) ¯
1 + 0.4 + 0.4R
1 + 0.4
° = 20.24𝑓𝑡
𝑌* = c
𝑏
6
e c
1 + 2𝜆
1 + 𝜆
e = c
120.23
6
e¯
1 + 2(0.4)
1 + 0.4
° = 31.49𝑓𝑡

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10.2.3. Tail
𝐶+³ = 1, 𝐶¦³ = 0.09
Considering our modified aircraft is a controlled canard, hence our value for 𝐶+³ = 0.1

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10.2.4. Horizontal Tail
𝑆+³ =
𝑐+³ 𝐶h
**** 𝑆h
𝐿+³
=
(0.1)(20.24)(2294.95)
0.35(81.35)
= 163𝑓𝑡R
𝑏 = Ò(𝑠 ∗ 𝐴𝑅) = √(163)(3) = 22𝑓𝑡
𝜆 = 0.4
𝐶W__‰ =
2𝑆
𝑏(1 + 𝜆)
=
2(163)
(22)(1 + 0.4)
= 10.6𝑓𝑡
𝐶‰†[ = 𝜆𝐶W__‰
𝐶‰†[ = 𝜆𝐶W__‰ = (0.4)(10.6) = 4.24𝑓𝑡
2
3
e 𝐶W__‰ ∗
1 + 𝜆 + 𝜆R
1 + 𝜆
= c
2
3
e (10.6) ¯
1 + 0.4 + 0.4R
1 + 0.4
° = 7.87𝑓𝑡
𝑌* = c
𝑏
6
e c
1 + 2𝜆
1 + 𝜆
e = c
22
6
e¯
1 + 2(0.4)
1 + 0.4
° = 4.71𝑓𝑡

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10.2.5. Vertical Tail
𝑆¦³ =
𝑐¦³ 𝑏h 𝑆h
𝐿¦³
=
(0.09)(120.23)(2294.95)
0.35(81.35)
= 872𝑓𝑡R
𝑏 = Ò(𝑠 ∗ 𝐴𝑅) = √(872)(5) = 66𝑓𝑡
𝜆 = 0.4
𝐶W__‰ =
2𝑆
𝑏(1 + 𝜆)
=
2(872)
(66)(1 + 0.4)
= 18.8𝑓𝑡
𝐶‰†[ = 𝜆𝐶W__‰
𝐶‰†[ = 𝜆𝐶W__‰ = (0.4)(18.8) = 7.52𝑓𝑡
2
3
e 𝐶W__‰ ∗
1 + 𝜆 + 𝜆R
1 + 𝜆
= c
2
3
e (18.8) ¯
1 + 0.4 + 0.4R
1 + 0.4
° = 13.96𝑓𝑡
𝑌* = c
𝑏
3
e c
1 + 2𝜆
1 + 𝜆
e = c
66
3
e¯
1 + 2(0.4)
1 + 0.4
° = 28.3𝑓𝑡

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11.0. Crew station, Passengers, and Payload
The design of the interior of the aircraft is as important as the design of the exteriors. The
placement of the cockpit can affect the pilot’s performance and ease of handling the cockpit
features.
11.1. Crew Station
The design and placement of the crew station and the cockpit is essential, and it is a crucial
step during the aircraft conceptual design process. The pilot’s outside vision must be
perfectly clear with no obstruction coming in the way. Apart from the importance of the
location of the cockpit in the vicinity of the aircraft, the fuselage shape surrounding the
cockpit is of great importance as well.
A generalized crew station design consists of the placement and sizing of the following
segments of which are:
• Flight deck
• Crew rest compartment
• Upper avionics bay
• Main deck
• Main avionics bay
• FWD cargo compartment
The Falcon 7x aircraft is a 12-passenger aircraft, having only the main deck. The following
figure 27. contains the main layout of the Falcon 7x fuselage showing all of the main
compartments.
Figure 27. Falcon 7x main fuselage layout

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The following figure shows the interiors of the falcon 7x cockpit as well as the placement of
each of the interior features and the pilot’s seat.
Figure 28. Falcon 7x cockpit interiors
11.2. Pilot Sizes
A typical commercial aircraft has a certain set of height requirements for the pilot. The height
of the pilot flying a commercial aircraft is similar to that of the military aircraft. The minimum
height requirement if 1.66 m (65.2 in), and the maximum height requirement is 1.86 m (73.1
in). Any person with a height outside the provided height requirement would find great
discomfort in flying the aircraft and in the pilot seating.

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11.3. Seatback angle
The Falcon 7x pilot seat is shown and its parts are labelled in the figure to follow.
Figure 29. Pilot RH and LH seat (Anon., n.d.)
The seatback angle is the angle from the pilot’s head to the back of the pilot seat taken from
the seat reference point. The importance of the seatback angle depends on the aircraft
being designed, for example, a military aircraft pilot experiences a greater G-force than a
pilot flying a commercial aircraft. In that case, the seatback angle adjustment must be
accounted for. A greater seatback angle is beneficial for a military aircraft, decreasing the
effects of the G-force on the pilot. The seatback angle of a typical aircraft is an angle of 13°.
For the Falcon 7x the seatback angle is ranged from 8° to 48°, and is adjustable according
to (Anon., n.d.).
11.4. Over-nose vision
The over-nose angle is measured from the pilot’s sight line to the nose of the aircraft. The
following figure demonstrates how the over-nose angle may be measured/taken.
Figure 30: Aircraft over-nose angle (Anon., n.d.)

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The typical commercial aircraft has an allowable average over-nose angle of 11-20 degrees.
Figure 31. Falcon 7x nose-over angle (Anon., n.d.)
The over-nose angle of the Falcon 7x is measured to being 14 degrees. The over-nose
angle is a very important factor taken into consideration which is crucial for the aircraft
landing. In the case of a military aircraft, it is important for air-to-air combat. Having a longer
aircraft nose would make it more difficult for the pilot to see the runway while landing. The
nose’s aerodynamic shape must be streamline such that it does not increase the drag.
The over-nose angle (𝛼_tXW…_uX) can be calculated using the following equation:
𝛼_tXW…_uX = 𝛼[[W_V- + 0.07𝑉[[W_V-
Such that;
𝑉[[W_V- (Knots) = Approach velocity
𝛼[[W_V- = Approach angle
The approach velocity obtained for the Falcon 7x is 104 knots (Anon., n.d.), and the
approach angle is obtained as 3 degrees as per (Agency, n.d.). Thus, the over-nose angle
is calculated as:
𝛼_tXW…_uX = 10.28°
14°

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11.5. Vision angle looking upward
According to the vision angle looking upwards, the falcon 7x does not have an eyebrow
window that enables the pilot to have a greater vision upwards angle, thus the vision upward
angle is limited. The following image represents the vision angle looking upwards.
Figure 32. Vision angle looking upwards
Figure 32 represents a shape similar to that of the Falcon 7x, having a low wing
configuration.
11.6. Transparency Grazing Angle
The transparency grazing angle of an aircraft is the angle that lies between the pilot’s line
of vision and the edge of the windscreen of the cockpit. The angle must have a minimum
value of 30°. The grazing angle of the Falcon 7x is as seen in the following figure:
Figure 33. Falcon 7x grazing angle
The measured grazing angle of the falcon 7x as seen in figure 33 is 30 degrees.
30°

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11.7. Passenger compartment
Through the design of a commercial aircraft, the arrangement of the cabin seating is
determined by a set of dimensions which take into consideration the number of passengers
on the aircraft, and the number of seats which are placed in each row. The maximum seats
that may be accessed through each aisle are three seats, and there must be a distance of
70 ft between each door. The following are the dimensions which must be calculated in
order to design the appropriate cabin seating plan:
• The pitch: Distance from the seat backrest of a seat to the backrest of the seat to
follow.
• Headroom: The distance measured from the roof above the seats to the floor.
• Seat width
• Aisle width
• Aisle height
The design if the cabin interiors of the business jet is not a conventional aircraft cabin design
with rows of seats, but the seats are placed in a very specious setting and there aren’t any
seats placed one after the other, and thus the seat pitch is not measured. But typically, for
a small aircraft the seat pitch must be a value of 0.78 m.
The previous cabin dimensions mentioned have been calculated for the aircraft to be
designed as are as shown in the following table:
Table 10. Cabin dimensions of Falcon 7x
CABIN SECTION MEASUREMENT (m)
Seat pitch -
Seat width 0.752
Head height 1.480
Aisle width 0.705
Aisle height 1.927

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The following figures show the different cabin segments in the aircraft being designed, as
well as the cabin length, height and width.
Figure 34. Cabin width and height
Figure 35. Cabin passenger compartment segments and cabin length (Anon., 2019)

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11.8. Cargo Provisions
Since the business jet transporter (Falcon 7x) does not carry many passengers whilst
having a spacious cabin, the cargo compartment does not need to be too spacious. As seen
in figure 36, the cargo compartment is quite small. The cargo or baggage compartment is
placed in the rear end of the fuselage, following the cabin. The baggage compartment is
small and would fit 6 large passenger luggage bags.
Figure 36. Aircraft full layout
Since the cargo compartment is quite small, and the business jet aircraft is smaller in size
than the conventional aircraft (Boeing 747 for example), the cargo is not loaded in
containers, but instead they are loaded onto the aircraft on flat pallets. The following figure
shows a pallet on which the passenger baggage is placed and placed into the aircraft with.
Figure 37. Baggage flat pallet (Anon., 2019)
The cargo is placed underneath the cabin, and this is hown in the figure above of the cabin
cross-section.

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12.0. Propulsion and Fuel System
12.1. Engine
The propulsion system is an important part of the conceptual design of the aircraft.
Regardless of the type and size of the engine, it is one of the heaviest items in the aircraft.
The exact size values and specifications should be known in order to continue to develop
the propulsion system procedure followed by the fuel system because the fuel tank has a
considerable aircraft weight. As a basic operation of an aircraft engine, it compresses the
outside air, mixes it with fuel, and then burns and extracts the energy obtained from the
generated high-pressure gas.
When taking into account the available propulsion system options, the following could be
considered:
• Centrifugal turbojet
• Turbofan
• Turboprop
• Afterburner
• Piston prop
The high bypass turbofan engine is chosen for the business jet design of this project.
12.1.1. Engine Type
Turbofan engine
The selected propulsion system is turbofan, which is a commonly used jet engine. There
are two types of turbofan engines, low bypass, which generate a larger jet thrust than the
fan thrust, and vice versa high bypass turbofan engines. The ratio of the air mass flow
bypassing the engine core to the air mass flow through the core is referred to as the bypass
ratio. In the latter case, high bypass turbofans are typically used for commercial flights, while
low bypasses are commonly used for fighter aircraft.

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High by-pass Turbofan
Figure 38. high by-pass turbofan
Turbofan engines are made more fuel efficient by simultaneously increasing the inlet
temperature of the overall pressure ratio of the turbine rotor. The previously mentioned low
thrust is achieved by using a single stage instead of a multi-stage fan. Therefore, turbofan
high bypass engines can also be called low specific thrust engines. The core of the engine
must generate sufficient power to drive the fan in its design, Corresponding improvements
are made in turbine cooling to enhance the core thermal efficiency, which can in turn tolerate
higher inlet temperatures for the high-pressure turbine. So, decreasing the core flow can be
said to increase the bypass ratio. However, with reduced core mass flow, it affects efficiency
and increases the load on low-pressure turbine, in order to address this issue, further steps
are needed to maintain and reduce efficiency and stage loading on low-pressure turbines.
12.1.2. Engine Location
The location of the engine on an aircraft has numerous aerodynamic and structural effects
in all aspects. When it comes to engine location, there are several options, such as:
• In / on the wings
• Above the wings
• Below the wings
• After fuselage
• Top of fuselage
• Side of fuselage
The location for engine placement is selected at the aft of the fuselage for this design
project. The general idea is that as the aircraft becomes smaller, the positioning of engines
under a wing like a typical aircraft becomes problematic with regard to the ground clearance
of the nacelle and is therefore chosen to arrange the aft engine.

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Figure 39. Falcon 7X
The following points are considered to examine the advantages of the arrangement chosen:
• Greater CLMAX due to wing pylon removal and interference of exhaust flaps
• Less drag due to wing pylon interference elimination
• Avoid debris from getting into the engine
• Lower height of fuselage permitting shorter landing gears
The following may be noted as to the disadvantages:
• The center of gravity is moved backwards
• Bigger Tail
• Problem of balancing aircraft
• Maintenance problem
12.1.3. Inlet and Nozzle
The basic function of the entrance is to decelerate the incoming air to approximately half of
the speed of sound before entering the engine. Entrance geometry affects engine
performance.
The four type of inlet are:
• Conical, Round, Spike Inlet
• NACA Flush Inlet
• 2D Ramp Inlet
• Pitot or Normal Shock Inlet

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The pitot / normal shock inlet is selected here for the design project:
Figure 40. Pitot/Normal Shock Inlet
In subsonic flights, the Pitot / Normal shock inlet is commonly used. It is a simple hole facing
forward. At subsonic and low supersonic speeds, it is very effective. Cowl lip radius has a
significant impact on engine performance and drag. A blunt shape cowl with a radius of 7%
of the inlet front face radius for this design project.
There are different types of nozzle:
• Fixed convergent
• Converging iris
• Ejector
• Variable convergent
• 2-D vectoring
• Single expansion ramp
• Translating plug
• Converging-diverging ejector
The Fixed Convergent is selected for this project's business jet design. It is the type most
commonly used for turbojet and turbofan subsonic engines. With a focus on cruise
efficiency, the nozzle exit area is optimized. Its structure has an advantage in terms of
weight and simplicity.

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12.2. Fuel System
12.2.1. Fuel System Integration
Fuel system contains of fuel pump, tanks and pipes, fuel controller fuel system in the aircraft
manage the amount of fuel in the aircraft and enables the crew to transfer and pump the
fuel into the propulsion system and the APU. It depends on the performance of the aircraft
and the mission. The fuel system helps the pilot to control which tank will supply the engines
with fuel and the valves that can be switched off when an engine fire occurs. Fuel tanks are
a major component of the fuel system. There are different types of fuel tanks like bladder
tanks, integral tanks, tip tanks, etc.
The fuel systems for the smaller aircraft are divided into two:
• Gravity feed system
• Fuel pump system
12.2.2. Fuel Pump System
Basically, it consists of two fuel pumps. The main pump system is engine driven with an
auxiliary pump that is electrically driven for use in the engine start process. The auxiliary
pump, also known as a boost pump, makes the fuel system more reliable, as it can also be
used in the event that the engine pump fails.
Figure 41. Fuel Pump System
• Carburetor: a device mixing air and fuel in an appropriate ratio for engine combustion.
• Primer: it pump fuel directly into the cylinders from the tanks
• Strains: Eliminates any moisture in the system and other sediments before entering
the carburetor
• Selector valve: Allow the selection of fuel from tanks

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13.0. Landing Gear
The aircraft is supported by the landing gear, allowing it to take off and land and usually run
until it stops (taxi). The gearbox location is very important when it comes to ground because
of stability and controllability. landing gear positioning provides a higher handling quality
and must eliminate over-equilibrium during take-off and landing.
13.1. Landing Gear Arrangement
Figure 42. shows the most common options for the landing gear arrangement. The most
common is the tricycle gear with one wheel (auxiliary wheel) before the center of gravity
and two main wheels after the center of gravity. If center of gravity is before the main wheels
offers stability on the ground. Tricycle landing gear can also improve forward visibility on
the ground and can provide passengers with a flat cabin floor and cargo landing, thus
selecting the tricycle landing gear arrangement due to its vast number of advantages.
Figure 42. Landing gear arrangement

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The tricycle landing gear design as shown is very complex, the landing gear length has to
be designed so that the tail does not hit the ground in landing, in order to avoid such a
problem the design measurement start from the wheel in the static position assuming that
the aircraft angle of attack for landing is 90% of the maximum lift, for must aircraft range is
from 10-15 degrees, to avoid from tipping back its tail the vertical angle from the front wheel
to center of gravity must be greater than tip back angle. If the nose wheel carries less than
5% of the weight of the aircraft, the traction of the nose wheel will not be sufficient to steer
the aircraft. The angle of turnover is a measure of the tendency of the aircraft to change
when taxiing around the sharp corner, fig. 43 shows that the desired travel angle of the strut
is about (7) degrees, allowing the tire to go up or back when a large bump is in front.
Figure 43. Tricycle Geometry
13.2. Tire Sizing
The term wheel is sometimes used to describe the whole assembly of the wheel, brake, tire.
But the wheel is generally a circular metal object mounted on a rubber tire, the break
function is to slow down the aircraft by increasing rolling friction also are designed to carry
aircraft weight, the main tires carry 90% of the overall weight while the nose tire carries 10%
of the aircraft's overall weight, figure 44 provides equations for estimating the size of the
main tire. As in the final design layout, the catalog provided by the manufacturers must
select the actual tires to be used:

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Figure 44. Tire size table
Figure 45 shows tire data for various aircraft types in accordance with their specifications.
The section "Three-part name" is the newest and highest-pressure tires that are designed
to meet specific requirements. They are classified by outer diameter, width and diameter of
the rim. These tires are chosen based on the smallest tire that carries the maximum loads.
Ww = Maximum Loads
Ap= Contact area with the pavement. Footprint area
𝑊Y = 𝐴/ ∗ 𝑃
𝑊Y = 90%( 𝑊T) = 0.9(70350) = 63315𝑙𝑏
Main wheels diameter 𝐷 = 𝐴𝑊Y
1
= 2.69(63315)T.R„Œ
= 43.14𝑓𝑡
Main wheels width 𝑊 = 𝐴𝑊Y
1
= 1.17(63315)T.RŒƒ
= 12.744𝑓𝑡
𝑃 = 120𝑝𝑠𝑖 = 17280𝑙𝑏/𝑓𝑡R
for major civil airfield
𝑊Y = 𝐴[. 𝑃
𝐴[ =
63315
17280
= 3.66
𝐴[ = 2.3 ∗ √𝑤 ∗ 𝑑 ∗ (
𝑑
2
∗ 𝑅W)
𝑅W =
𝐴[
2.3 ∗ √𝑤 ∗ 𝑑 ∗
𝑑
2
= 3.15 ∗ 10‚Š

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Figure 45, Tire Data
The best tire size for the designed aircraft should be around (21*7.25-10) from historical
data. A rough estimation for tire, internal pressure should be kept below the values shown
in figure 46 and the maximum pressure should be more than (200psi).
Figure 46. Tire Pressure

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13.3. Shock Absorbers
The main objective of the landing gear is to absorb the landing shock and smooth the ride
while taxiing, as shown in figure 47.
Figure 47. Shock Absorber

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13.4. Oleo struts
The shock absorber selected for the designed aircraft is the most common type of shock
absorber used today. Oleo Shock struts is called oleo or air & oil struts because of the
combination of compressed air nitrogen and the hydraulic fluid to absorb shock loads while
landing. Oleo shock struts work at the external ends with double telescopic cylinders. The
aircraft is fitted with the top cylinder and the bottom cylinder is attached to the landing gear.
Also, the bottom cylinder can slide in and out of the top cylinder easily.
Figure 48. Oleo Struts
Hydraulic fluid fills the bottom cylinder while the top cylinder is filled with nitrogen or
compressed air. Small hole called a orifice that connects the two cylinders
Figure 49. Oleo shock absorber machine
As the aircraft lands, the pressure from the ground-touching wheels forces the hydraulic
fluid to the orifice and to the nitrogen or compressed air chamber in the top. As the fluid
moves up the orifice, heat in the form of kinetic energy begins, the shock of landing is
absorbed by transferring this energy into thermal energy.

23-05-2019 72 DR. ELHAM
13.5. Gear Retraction Geometry
Everything is known in gear retraction geometry design process, the size of wheels, tires
and shock absorbers the last task is to find a home for the landing gears. Choosing a poor
design can result in weight gain, decrease the volume of internal fuel, and create extra drag.
Figure 50 shows the retracted positions of the main landing gear.
Figure 50. Gear place
The geometry of the gear retraction for the gear is selected using the main gear to retract
or extend the gear into or out of the intersection junction of the fuselage wing and the nose
gear simply folds forward into the fuselage below the cockpit. This choice is due to its ease
of production, high reliability and low operating costs.
13.6. Aircraft Subsystems
Hydraulic System: At specified pressure, a light oil or liquid is pumped and then stored in a
tank.
Electrical system: Offers avionics, hydraulics, lighting and other subsystems with electrical
power.
Pneumatic system: Primarily supplies compressed air for pressure, environmental control,
anti-icing and in some cases, engine starting
Auxiliary/Emergency power system: is a hydraulic and electrical system for emergencies if
it is not backed up by the original system
Avionics: Radio, flight control, radar, sensor, instrument …etc.

23-05-2019 73 DR. ELHAM
14.0. Component Weights
Sections 14.1. and 14.2. of this report demonstrate quick methods for estimating the empty
weight of the aircraft. To acquire and estimate results more accurately, weights of separate
components are estimated individually. During the conceptual design phase, traditional
historic statistics are used, whereas during the preliminary and detailed design phase
component selection and structural analysis methods are used.
14.1. Weights Reporting and CG Estimation
The maximum takeoff weight which has been estimated during preliminary sizing should
not be altered. Based on historical trends where in a new component is taken to be similar
to a component of an aircraft in existence.
14.2. Approximate Weight Methods
The approximate weights method is the estimation of the aircraft components weight
through very simple weight ratios. While these estimations are not extremely accurate and,
it is a very quick and easy estimation. Such historical trends are summarized in table 26,
along with their location estimates.
Table 11. Approximate Empty Weight Buildup

23-05-2019 74 DR. ELHAM
14.3. Statistical Weight Methods
The process of weight estimation is an iterative process and there are no right or wrong
answers during this process. However the values should lie within sensible ranges for the
results to be accurate. The statistical weight method is a much more refined approach
towards the weight estimation. Various manufacturers create their own version of these
equations in order to estimate the weight and thus the centre of gravity in the most accurate
manner. Some of the equations have been released by the respective manufacturers. The
equation for general aviation are summarized as follows.
PARAMETER VALUE
Sw 1985ft2
Wfw 29958lb
A 6.3
λ 0.37
Λ 30º
q 109lb/ft2
t/c 0.15
Nz 2.1
Wdg 70350lb
Sht 164ft2
Ht/Hv 0
Sf 2386ft2
Lt 10ft
Lm 2.5ft
Ln 4ft
Wpress 445.77lb
Pdelta 8psi=1152lb/ft2
Vpr 5567ft3
Wl 70350lb
Nl 4.5
Wen 990lb
Vt 596ft3
Vi 750ft3
L 81.35ft
Bw 120ft
Kn 0.12
W 9.84ft
M 0.9
Wuav 1200lb
Np 15

23-05-2019 75 DR. ELHAM
𝑊Y†…‡ = 0.036𝑆Y
T.«„©
𝑊dY
T.TTŠ„
c
𝐴
𝑐𝑜𝑠RΛ
e
T.ƒ
𝑞T.TTƒ
𝜆T.Tˆ ð
100
𝑡
𝑐
𝑐𝑜𝑠Λ
ñ ‚T.Š
(𝑁4 𝑊`‡)T.ˆ¬

𝑊Y†…‡ = 0.036𝑆Y
T.«„©
𝑊dY
T.TTŠ„
c
𝐴
𝑐𝑜𝑠RΛ
e
T.ƒ
𝑞T.TTƒ
𝜆T.Tˆ ð
100
𝑡
𝑐
𝑐𝑜𝑠Λ
ñ ‚T.Š
(𝑁4 𝑊`‡)T.ˆ¬
𝑊-_W†4_…‰^ ‰†^ = 0.016(𝑁4 𝑊`‡)T.ˆŒˆ
𝑞T.Œƒ©
𝑆-‰
T.©¬ƒ
ð
100
𝑡
𝑐
𝑐𝑜𝑠Λ
ñ ‚T.ŒR c
𝐴
𝑐𝑜𝑠RΛ-‰
e
T.TˆŠ
𝜆-
‚T.TR
𝑊tXW‰†V^ ‰†^ = 0.073 c1 + 0.2
𝐻‰
𝐻t
e(𝑁4 𝑊`‡)T.Š«ƒ
𝑞T.ŒRR
𝑆t‰
T.©«Š ð
100
𝑡
𝑐
𝑐𝑜𝑠Λt‰
ñ ‚T.ˆ¬ c
𝐴
𝑐𝑜𝑠RΛt‰
e
T.Š„«
𝜆t‰
T.TŠ¬
𝑊d‹uX^‡X = 0.052𝑆d
Œ.T©ƒ
(𝑁4 𝑊`‡)T.Œ««
𝐿‰
‚T.T„Œ
c
𝐿
𝐷
e
‚T.T«R
(𝑞)T.RˆŒ
+ 𝑊[WXuu
𝑊•†… ^…`†…‡ ‡XW = 0.095(𝑁^ 𝑊^)T.«ƒ© c
𝐿•
12
e
T.ˆT¬
𝑊…_uX ^…`†…‡ ‡XW = 0.125(𝑁^ 𝑊^)T.„ƒƒ c
𝐿…
12
e
T.©ˆ„
𝑊†…u‰^^X` X…‡†…X = 2.575𝑊X…
T.¬RR
𝑁X…
𝑊d‹X^ u]u‰X• = 2.49𝑉‰
T.«Rƒ
6
1
1 +
𝑉†
𝑉‰
7
T.ŠƒŠ
𝑁‰
T.RˆR
𝑁X…
T.Œ„«
𝑊d^†‡-‰ V_…‰W_^u = 0.053𝐿Œ.„Šƒ
𝐵Y
T.Š«Œ
(𝑁4 𝑊`‡ ∗ 10‚ˆ
)T.©
𝑊-]`W‹^†Vu = 𝐾- 𝑊T.©
𝑀T.„
𝑊t†_…†Vu = 2.117𝑊‹t
T.¬ŠŠ
𝑊X^XV‰W†V^ = 12.57ù𝑊d‹X^ u]u‰X• + 𝑊t†_…†Vuú
T.„Œ
𝑊†W V_…`†‰†_…†…‡ …` …‰† †VX = 0.265𝑊`‡
T.„R
𝑁[
T.ƒ©
𝑊t†_…†Vu
T.Œ«
𝑀T.T©
𝑊d‹W…†u-†…‡ = 0.0582𝑊`‡ − 65

23-05-2019 76 DR. ELHAM
COMPONENT WEIGHT
(lb)
LOCATION
FROM
NOSE(ft)
MOMENT
(lb.ft)
LOCATION
FROM CL(ft)
MOMENT
(lb.ft)
Wing 14124 35.06 495187 0 0
Horizontal Tail 859 22.95 19714 0 0
Vertical Tail 4919 74.47 366317 14.35 70587
Fuselage 12442 40.67 506016 0 0
Main Landing
Gear
1430 43.5 62205 6.17 8823
Nose Landing
Gear
149 10.5 1564 6.92 1031
Installed Engine 10403 74.47 774711 6.52 67827
Fuel System 952 35.06 33377 0 0
Flight Controls 1407 38.76 1445 0 0
Hydraulics 4475 46 205850 2 8950
Electrical 9529 40 381160 3 2857
Avionics 3679 75 255925 3 11037
Air Conditioning 5407 25 135175 3 16221
Furnishing 163 40.5 6601 2 326
TOTAL 69938 3245247 187659
𝑥V‡ =
∑ 𝑊† 𝑥†
†
Œ
𝑊T
=
3245247
69938
= 46.4𝑓𝑡
𝑧V‡ =
∑ 𝑊† 𝑧†
†
Œ
𝑊T
=
187659
69938
= 2.68𝑓𝑡

23-05-2019 77 DR. ELHAM
14.3.1. Cruise
During cruise the only variation compared to take off condition is that half the fuel is
estimated to be burnt and weight of the landing gear is added to the fuselage as it is
retracted inwards.
𝑥V‡ =
∑ 𝑊† 𝑥†
†
Œ
𝑊T
=
2887061
66187
= 43.6𝑓𝑡
𝑧V‡ =
∑ 𝑊† 𝑧†
†
Œ
𝑊T
=
158850
66187
= 2.40𝑓𝑡
14.3.2. Landing
During landing, it is estimated that 1/5th
of the initial weight of fuel remains which accounts
for the reserve fuel. In addition the landing gear is also redeployed.
𝑥V‡ =
∑ 𝑊† 𝑥†
†
Œ
𝑊T
=
3065247
63438
= 48.3𝑓𝑡
𝑧V‡ =
∑ 𝑊† 𝑧†
†
Œ
𝑊T
=
122659
63438
= 1.93𝑓𝑡

23-05-2019 78 DR. ELHAM
14.3.3. Centre of Gravity Envelope Diagram
The plot of the changing centre of gravity with variation in flight condition along with change
in total gross weight in shown in figure (X). The limits of c.g. are fixed between the main and
the nose landing gears such that the aircraft remains stable during all the flight conditions.
It is observed that the aircraft so designed does stay well within the forward and aft c.g. limit
and thus the aircraft would remain stable which is extremely important in this case as
passenger safety and comfort is the first priority for a business jet.
c.g. different
flight condition
Forward c.g.
limit
Aft c.g. limit

23-05-2019 79 DR. ELHAM
15.0.Stability, control and hand qualities
The basic concept of stability is simply that a stable aircraft, when disturbed, tends to
return by itself to its original state (pitch, yaw, velocity). “static stability” is present if the
forces created by the disturbed state (such as a pitching moment due to an increased
angle of attack) push in the correct direction to return the aircraft to its original state.
Dynamic stability is present if the dynamic motions of the aircraft will eventually return the
aircraft to its original state. The way the aircraft returns to its original state depends upon
the restoring forces, mass distribution, and "damping forces." Damping forces slow the
restoring rates. For example, a pendulum swinging in air is lightly damped and will
oscillate back and forth for many minutes. The same pendulum immersed in water is
highly damped and will slowly return to vertical with little or no oscillation.
The figures below illustrate these concepts for an aircraft disturbed in pitch. In figure (a)
the aircraft has perfectly neutral stability and simply remains at whatever pitch angle the
disturbance produces. In figure (b) it shows static instability. The forces produced by the
greater pitch angle cause the pitch angle to further increase. Pitchup is an example of this.
In figure (c) the aircraft shows static stability with very high damping. The aircraft slowly
returns to the original pitch angle without any overshoot. In figure (d) shows a more typical
aircraft response; the aircraft returns to its original state but experiences some converging
oscillation. In figure (e) it shows the restoring forces are in the right direction, so the
aircraft is statically stable.
Figure 51. Static and Dynamic Stability

23-05-2019 80 DR. ELHAM
15.1. Coordinate systems and definitions
Figure 52 defines the two axis systems commonly used in aircraft analysis. The "body-axis
system" is rigidly fixed to the aircraft, with the X axis aligned with the fuselage and the Z
axis upward. The origin is at an arbitrary location, usually the nose. The body-axis system
is more "natural" for most people but suffers from the variation of the direction of lift and
drag with angle of attack.
The "stability" axis system, commonly used in stability and control analysis, is a
compromise between these two. The X-axis is aligned at the aircraft angle of attack, as in
the wind axis system, but is not offset to the yaw angle. Directions of X, Y, and Z are as in
the wind axis system.
Wing and tail incidence angles are denoted by i, which is relative to the body-fixed
reference axis. The aircraft angle of attack a is also with respect to this reference axis, so
the wing angle of attack is the aircraft angle of attack plus the wing angle of incidence.
Nondimensional coefficients for lift and drag have been previously defined by dividing by
dynamic pressure and wing area. For stability calculations, the moments about the three
axes (M, N, and L) must also be expressed as nondimensional coefficients.
Since the moments include a length (the moment arm) they must be divided by a quantity
with dimension of length as well as by the dynamic pressure and wing area. This length
quantity is the wing MAC chord for pitching moment and the wing span for yawing and
rolling moments, as shown in Eqs. (16.1-16.3). Positive moment is nose up or to the right.
Cm =M / qSc (16.1)
Cn = N / qSb (16.2)
Cf= L / qSb (16.3)
Figure 52. Aircraft Coordinate System

23-05-2019 81 DR. ELHAM
15.2. Stability Axis system
X axis: is aligned at the aircraft angle of attack (wind axis system)
Y axis: towards the right wing
Z axis: downwards
Pitching moment (M): nose-up (+)
Rolling moment (L): right wing down (+) nose to the right
Yawing moment (N): right wing backward (+) nose to the right
Derivative of moment coefficients with respect to angle of attack or
sideslip: 𝐶𝑚𝛼 , 𝐶𝑛𝛽 , 𝐶𝑙𝛽 , 𝐶𝑚𝛿
15.3. Longitudinal static stability and control
Pitching-Moment Equation and Trim:
Major contributors to aircraft pitching moment about the c.g. wing, tail, fuselage and
engine contributions.
Most aircraft being symmetrical about the centre line, moderate changes in angle of attack
will have little or no influence upon the yaw or roll. This permits the stability and control
analysis to be divided into longitudinal (pitch only) and lateral-directional (roll and yaw)
analysis. Figure 16.3 shows the major contributors to aircraft pitching moment about the
e.g., including the wing, tail, fuselage, and engine contributions. The wing pitching-
moment contribution includes the lift through the wing aerodynamic center and the wing
moment about the aerodynamic center. Remember that the aerodynamic center is defined
as the point about which pitching moment is constant with respect to angle of attack. This
constant moment about the aerodynamic center is zero only if the wing is uncambered
and untwisted.
Figure 53. Longitudinal Moments

23-05-2019 82 DR. ELHAM
Total moment about the c.g.:
𝑀𝑐𝑔 = 𝐿 (𝑋𝑐𝑔 − 𝑋𝑎𝑐,𝑤 )+ 𝑀𝑤 + 𝑀𝑤𝛿𝑓𝛿𝑓 + 𝑀𝑓𝑢𝑠 −𝐿ℎ (𝑋𝑎𝑐,ℎ − 𝑋𝑐𝑔) − 𝑇𝑧𝑡 + 𝐹𝑝(𝑋𝑐𝑔 − 𝑋𝑝)
(16.4)
Total coefficient of moment about the c.g.:
𝐶𝑚𝑐𝑔 = 𝐶𝐿(𝑋ത 𝑐𝑔 − 𝑋ത 𝑎𝑐,𝑤) + 𝐶𝑚𝑤 + 𝐶𝑚𝑤𝛿𝑓𝛿𝑓 + 𝐶𝑚𝑓𝑢𝑠 −𝜂ℎ 𝑆ℎ 𝑆𝑤 𝐶𝐿ℎ(𝑋ത 𝑎𝑐,ℎ −
𝑋ത 𝑐𝑔) − 𝑇 𝑞𝑆𝑤 𝑍ҧ 𝑡 + 𝐹𝑝 𝑞𝑆𝑤 (𝑋ത 𝑐𝑔 − 𝑋ത 𝑝) (16.7)
For a static "trim" condition, the total pitching moment must equal zero. For static trim, the
main flight conditions of concern are during the take-off and landing with flaps and landing
gear down and during flight at high transonic speeds. Usually the most forward e.g.
position is critical for trim. Aft-c.g. position is most critical for stability. Equation (16.7) can
be set to zero and solved for trim by varying some parameter, typically tail area, tail lift
coefficient (i.e., tail incidence or elevator deflection), or sometimes e.g. position. The wing
drag and tail trim drag can then be evaluated.
For longitudinal static stability: 𝐶𝑚𝛼 < 0
Derivative of pitching moment coefficient with respect to angle of attack:
𝑑𝐶𝑚 /𝑑𝛼 = 𝐶𝑚𝛼 = 𝐶𝐿𝛼(𝑋ത 𝑐𝑔 – 𝑋ത)𝑎𝑐,𝑤 + 𝐶𝑚𝛼𝑓𝑢𝑠 −𝜂ℎ(𝑆ℎ/𝑆𝑤) 𝐶𝐿𝛼ℎ(𝜕𝛼ℎ/ 𝜕𝛼)(𝑋ത 𝑎𝑐,ℎ −
𝑋ത 𝑐𝑔) + (𝐹𝑃𝛼/𝑞𝑆𝑤) (𝜕𝛼𝑝/𝜕𝛼) (𝑋ത 𝑐𝑔 − 𝑋ത 𝑝)
The pitching moment derivative equation changes with c.g. location. There is a c.g.
position, where pitching moment is constant with angle of attack, which is called airplane
aerodynamic centre or neutral point.
15.4. Static Margin
𝑋*…[ =
𝐶§A
𝑋*V,Y − 𝐶•A,d‹u
+ 𝑛-
𝑆-
𝑆Y
𝐶§A-
𝜕𝛼-
𝜕𝛼 𝑋*V,- +
𝐹[,A
𝑞𝑆Y
𝜕𝛼[
𝜕𝛼 𝑋*[
𝐶§A
+ 𝑛-
𝑆-
𝑆Y
𝐶§A-
𝜕𝛼‹
𝜕𝛼
+
𝐹[,A
𝑞𝑆Y
𝐶•A
= −𝐶§A
ù𝑋*…[ − 𝑋*V‡ú
𝑆𝑀 = −
𝑐•,A
𝑐^,
𝑋*V,Y = 7.26, 𝐶§A
= 6.19,
𝜕𝛼‹
𝜕𝛼
= 1.5,
𝜕𝛼-
𝜕𝛼
= 0

23-05-2019 83 DR. ELHAM
16.0.Financial Analysis
Before an aircraft can be proposed to a potential customer, the cost analysis is mandatory
to be included, as in the aviation industry, it is crucial to consider that any development
requires to be cost efficient. When designing an aircraft the cost estimation takes into
account of all the research, funding, the production costs, and the operational costs till the
aircraft is disposed. The life-cycle costs of the business jet and how these costs are
estimated will be used to discuss the different factors.
16.1. Life-cycle Cost Elements
The life-cycle cost elements suggests that it is important to study the production and
operational costs as well as the development costs; as can be seen in figure blow.
Figure 54. Life-cycle Cost Elements.
• RDT&E: Research, Development, Tests, and Evaluation of the aircraft throughout the
designing and developing process, includes the expenses of technological research and
tests which were carried out to develop the design on the aircraft, including the certifications
for civil aircrafts, airworthiness, and the mission capabilities of the designed airplane.
• Flyaway Cost: it is the cost production which includes labor and material costs, hence are
decided by manufacturers and are usually negotiated to be reduced as much possible to
maximize profit out of the project while staying within the limits of safety and good quality
regulations. The production costs are usually very expensive and considered of the highest
part of a life-cycle costs of an aircraft.
• Ground support equipment and special constructions cost: mainly for military aircrafts,
where special ground equipment is required in order to perform maintenance tasks and
ground operations. For civil aircrafts, the ground support equipment is common and well
known, therefore, their cost will be very little and sometimes not mentioned.

Aircraft Design Thesis Report

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