Aircraft 8 Passengers Design [RAVEN].

THIRD Year
Design and Manufacturing of Aircraft Parts AER(320)
F
Submitted to : Dr. Muhammad Nader
Date: 24th
DECEMBER,2019
Name section B.N.
Ahmed Samir Salah 1 1
Ahmed Tarek Mohamed 1 3
Bahaa Ibrahim Ibrahim 1 12
Peter Wageeh Roshdy 1 13
Abdelhaleem Hamada 1 20
AIRCRAFT DESIGN REPORT

1
AIRCRAFT DESIGN REPORT
Department of Aerospace Engineering
Cairo University, Faculty of Engineering
Giza
Technical Report
December 24, 2019
Abstract: The aim of this project is to design a light Twin Engine propeller aircraft
that can cater to small range about 450 miles and with a cruise speed 240 Knots. The
project involves the design of light Twin Engine propeller aircraft which could
accommodate about 6 passengers not including the pilot and co-pilot and considering
no attendants involved providing a medium level of comfort that a twin propeller
driven with small range is expected to provide. The aircraft design has been construct
on some certain step processes. The processes that are applied in this project is goes
like this; weight estimation, initial sizing, airfoil and geometry selection ,Thrust to
weight and wing loading analysis, configuration layout, propulsion and fuel system
analysis, aerodynamic analysis, stability and finally the cost analysis. Those processes
has been implemented in this project and each and every study is applied by comply
the methods and rules that are gathered from aircraft design books.
Keywords: Aircraft design, Twin-Engine, Propeller-Driven, Performance analysis,
Weight Estimation, Propulsion analysis, Sizing, Drag analysis.

Table of Contents
1. INTRODUCTION TO DESIGN ..................................................................................................14
2. RAVEN.........................................................................................................................................15
2.1 Description .......................................................................................................................... 15
2.2 Summary of Key Parameters............................................................................................... 15
2.3 Configuration Layout .......................................................................................................... 15
3. Requirements Analysis .................................................................................................................16
3.1 Requirements Summary ......................................................................................................16
3.2 Mission Profile .................................................................................................................... 16
3.3 Reference Design Concepts................................................................................................. 17
4. Weight estimation .........................................................................................................................18
4.1 Fuel-Fraction Method [Twin-Engine Propeller-Driven Airplane] ......................................19
4.2 Airplane Gross-Weight Estimation ..................................................................................... 20
5. Weight sensitivity .........................................................................................................................21
5.1 Sensitivity of TO .............................................................................................. 21
5.2 Sensitivity of TO ................................................................................................21
5.3 Sensitivity of TO , , , , ....................................................... 21
5.3.1 For the case involving the ratio + 1 dependent on range : ................................21
5.3.2 For the case involving the ratio + 1 dependent on endurance :........................21
6. Matching Plot................................................................................................................................23
6.1 Introduction ......................................................................................................................... 23
6.2 Stall Speed Sizing................................................................................................................ 23
6.3 Take-off Distance Sizing..................................................................................................... 24
6.4 Landing Distance Sizing......................................................................................................25
6.5 FAR 23 Climb Requirements Sizing...................................................................................27
6.5.1 Drag Polar Estimation...................................................................................................27
6.5.2 Rate of Climb Sizing.....................................................................................................28
6.6 Climb Gradient Sizing......................................................................................................... 29
6.7 Cruise Speed Sizing............................................................................................................. 30
6.8 Time-to-Climb Sizing.......................................................................................................... 31
6.9 Matching curves .................................................................................................................. 32
7. Wing design ..................................................................................................................................34
7.1 Number of wings ................................................................................................................. 35
7.2 Wing vertical location ......................................................................................................... 36
7.2.1 High wing......................................................................................................................36
7.2.2 Low wing ......................................................................................................................37
7.2.3 Mid-Wing......................................................................................................................38
7.2.4 The selection process ....................................................................................................38
7.3 Airfoil selection...................................................................................................................38
7.3.1 Airfoil Selection Criteria...............................................................................................38
7.3.2 Practical Steps for Wing Airfoil Section Selection.......................................................39
7.4 Wing Incidence.................................................................................................................... 42
7.4.1 The wing incidence must satisfy the following design requirements: ..........................43
7.5 Aspect Ratio ........................................................................................................................ 43
7.6 Taper Ratio.......................................................................................................................... 44
7.6.1 Calculating and selecting taper ratio for wing ..............................................................46
7.7 Sweep Angle........................................................................................................................ 46
7.7.1 Advantages of sweep ....................................................................................................47

7.8 Dihedral Angle .................................................................................................................... 48
7.9 High-Lift Device ................................................................................................................. 49
7.9.1 The Functions of a High-Lift Device............................................................................49
7.9.2 High-Lift Device Classification....................................................................................50
7.9.3 Design Technique .........................................................................................................51
7.10 Aileron................................................................................................................................. 52
8. Wing configuration.......................................................................................................................53
8.1 Wing Configuration Parameters:......................................................................................... 53
9. AIRPLANE CG ESTIMATION...................................................................................................55
9.1 Stability and Balance Control.............................................................................................. 55
9.2 Estimating CG Position ....................................................................................................... 56
10. Wing Body Balance ..................................................................................................................57
11. Trimming ..................................................................................................................................59
12. Tail Configuration.....................................................................................................................59
12.1 Basics Tail Configuration.................................................................................................... 59
12.2.3 Optimum Tail Arm........................................................................................................61
13. Horizontal Tail Design..............................................................................................................63
13.1 Horizontal Tail Location ..................................................................................................... 64
13.2 Select Horizontal Tail Volume Coefficient......................................................................... 64
13.3 Determine Optimum Tail Arm ............................................................................................ 64
13.4 Determine Planform Area.................................................................................................... 65
13.5 Horizontal Airfoil Selection ................................................................................................65
13.5.1 NACA 0006 Specifications...........................................................................................65
13.6 Determine Sweep angle and Dihedral angle........................................................................ 67
13.6.1 Sweep Angle .................................................................................................................67
13.6.2 Dihedral Angle..............................................................................................................67
13.7 Aspect Ratio ........................................................................................................................ 67
13.8 Taper Ratio.......................................................................................................................... 68
13.9 Setting Angle (incident angle).............................................................................................68
13.10 Calculation of MAC of Horizontal Tail, b, ......................................................... 69
14. Vertical Tail Design..................................................................................................................71
14.1 Location of Vertical Tail ..................................................................................................... 71
14.2 Selection of Vertical Tail Volume Coefficient.................................................................... 72
14.3 Determine Tail Arm............................................................................................................. 72
14.4 Planform Area ..................................................................................................................... 73
14.5 Airfoil Selection .................................................................................................................. 73
14.6 Sweep Angle........................................................................................................................ 73
14.7 Dihedral Angle .................................................................................................................... 74
14.8 Incident Angle ..................................................................................................................... 74
14.9 Aspect Ratio ........................................................................................................................ 74
14.10 Taper Ratio.......................................................................................................................... 75
14.11 Calculation MAC of VT, b of VT, .................................................................. 75
15. Spinning and Spinning Recovery..............................................................................................76
15.1 Spinning............................................................................................................................... 76
15.2 Spin Recovery ..................................................................................................................... 76
15.3 Check the ability of Rudder to prevent incipient spin......................................................... 77
15.3.1 Rudder power to prevent incipient spin ........................................................................77
15.3.2 Rudder Volume to recover from spin............................................................................77
16. VT Aileron Balance ..................................................................................................................78
16.1 Introduction ......................................................................................................................... 78

16.2 Types of Aileron.................................................................................................................. 78
16.3 Derivations .......................................................................................................................... 78
16.4 Calculation........................................................................................................................... 79
16.5 Flaps .................................................................................................................................... 80
16.6 Vertical Tail Engine Balance............................................................................................... 80
16.7 Rudder Deflection ............................................................................................................... 81
17. XFLR5 Design Performance.....................................................................................................82
18. Longitudinal Stability ...............................................................................................................83
18.1 Definition Of Longitudinal Stability ...................................................................................83
18.2 Contribution Of Aircraft Components................................................................................. 83
18.2.1 Wing Contribution ........................................................................................................84
18.2.2 Tail Contribution- Aft Tail............................................................................................85
18.2.3 Total Aircraft Components Contribution......................................................................86
19. Lateral Stability.........................................................................................................................87
19.1 Definition Of Lateral Stability.............................................................................................87
19.2 Contribution of Aircrafts Components................................................................................ 87
19.2.1 Wing-Fuselage Contribution.........................................................................................87
19.2.2 Vertical tail Contribution ..............................................................................................88
19.3 Rolling Stability...................................................................................................................89
19.3.1 Parameters Of Rolling Stability....................................................................................90
20. Propeller Configuration.............................................................................................................91
20.1 Propeller Definition............................................................................................................. 91
21. Undercarriage Design and Configuration .................................................................................94
21.1 Introduction ......................................................................................................................... 94
21.2 Landing Gear Arrangement................................................................................................. 94
21.2.1 Taildragger Undercarriage ............................................................................................95
21.2.2 Monowheel with Outriggers .........................................................................................95
21.2.3 Tricycle Landing Gear ..................................................................................................95
21.3 Tires, Wheels And Brakes Design....................................................................................... 95
21.3.1 Estimation.............................................................................................................96
21.3.2 Overturning Coefficient................................................................................................96
21.3.3 Wheel Track..................................................................................................................96
21.3.4 Turning Radius..............................................................................................................97
21.3.5 Tire Deflection and Shock Absorption .........................................................................97
21.3.6 Tire Footprint................................................................................................................98
21.3.7 Tire Selection................................................................................................................99
.22 Fuselage Configuration .................................................................................................................101
22.1 Introduction ....................................................................................................................... 101
22.2 Functional Analysis and Design Flowchart....................................................................... 101
22.3 Fuselage Configuration Design and Internal Arrangement............................................... 103
22.4 Ergonomics........................................................................................................................ 105
22.4.1 Definitions...................................................................................................................105
22.4.2 Human Dimensions and Limits...................................................................................105
22.5 Cockpit Design .................................................................................................................. 106
22.5.1 Number of Pilots and Crew Members.........................................................................107
22.5.2 Pilot/Crew Mission .....................................................................................................107

22.5.3 Pilot/Crew Comfort/Hardship Level...........................................................................107
22.5.4 Control Equipment......................................................................................................108
22.5.5 Measurement Equipment ............................................................................................108
22.5.6 Cockpit Integration .....................................................................................................109
22.6 Passenger Cabin Design .................................................................................................... 110
22.7 Cargo Section Design........................................................................................................ 112
22.8 Optimum Length-to-Diameter Ratio ................................................................................. 113
22.8.1 Optimum Slenderness Ratio for Lowest .............................................................113
22.9 Designed Fuselage Parameters .......................................................................................... 115
23. Fuel tanks................................................................................................................................115
24. Performance ............................................................................................................................119
24.1 Expression For Power Required For Level Flight............................................................. 119
24.2 Expression For Thrust Required For Level Flight............................................................. 120
24.3 Expression For Excess Power Required for Climb rate and Gradient and Horizontal
Acceleration from Level Flight....................................................................................................... 120
24.4 The Maximum Climb Rate and the Speed Which it Occurs.............................................. 121
24.5 The Climb Gradient at Max Climb Rate and the Max Climb Gradient............................. 121
24.5.1 Climb gradient at :..........................................................................................121
24.5.2 Max climb gradient :.........................................................................................121
24.6 Plot 32 , , 12 versus speed, evaluate their maximum........................................ 122
24.7 The Clean Aircraft Drag Polar........................................................................................... 123
24.8 Calculate and verify if it is satisfactory and how to adjust it if not satisfied?....... 123
24.9 Calculate at minimum power consumption ................................................................... 124
24.10 The maximum range for the aircraft.................................................................................. 124
24.11 Plot the aircraft envelop determined by available thrust ................................................... 124
24.12 Plot − curves at = 1,2,3,4..................................................................................... 125
24.13 = 0 plot – curves for at = 1,2,3,4......................................................... 125
24.14 Plot − for constant energy height ( 0000,40000,50000 .)...................... 126
24.15 Plot max turn − .................................................................................................... 126
24.16 Given = 4 , = −2 , = 1.3 , , = 392.2 ......................... 127
24.17 Calculate max ceiling ........................................................................................................ 128
24.18 Calculate Landing and Take-off Distance......................................................................... 129
24.18.1 Takeoff Distance.....................................................................................................129
24.18.2 Landing Distance ....................................................................................................129
25. Finance Issue...........................................................................................................................130
25.1 Introduction ....................................................................................................................... 130
25.2 Airplane Cost..................................................................................................................... 130
26. References...............................................................................................................................132
27. Appendix.................................................................................................................................133
27.1 Weight Estimation............................................................................................................. 133
27.2 Sensitivity.......................................................................................................................... 134
27.3 Matching Plot .................................................................................................................... 134
27.4 Longitudinal Stability........................................................................................................ 136
27.5 Lateral Stability ................................................................................................................. 137
27.6 Performance....................................................................................................................... 138
27.7 Absolute Ceiling................................................................................................................ 142

Table of Figures
FIGURE 1. CONFIGURATION LAYOUT....................................................................................................15
FIGURE 2 INITIAL MISSION PROFILE...........................................................................................16
FIGURE 3. DA 62 AND DA 42 MPP SPECS..............................................................................................17
FIGURE 4. STALL SPEED SIZING .............................................................................................................24
FIGURE 5. DEFINITION OF FAR 23 TAKE-OFF DISTANCES......................................................................24
FIGURE 6. TAKE-OFF DISTANCE SIZING.................................................................................................25
FIGURE 7. DEFINITION OF FAR 23 LANDING DISTANCES ......................................................................25
FIGURE 8. LANDING DISTANCE SIZING..................................................................................................26
FIGURE 9. RATE OF CLIMB SIZING.........................................................................................................28
FIGURE 10. CLIMB GRADIENT SIZING....................................................................................................29
FIGURE 11. CRUISE SPEED SIZING.........................................................................................................30
FIGURE 12. TIME TO CLIMB SIZING.......................................................................................................32
FIGURE 13. MATCHING PLOT................................................................................................................32
FIGURE 14.WING DESIGN PROCEDURE.................................................................................................35
FIGURE 15.THREE OPTIONS IN NUMBER OF WINGS: (A) MONOPLANE, (B) BIPLANE .....36
FIGURE 16.OPTIONS IN VERTICAL WING POSITIONS: (A) HIGH WING; (B) MID-WING; (C)
LOW WING; AND (D) PARASOL WING .................................................................................36
FIGURE 17 MAXIMUM LIFT COEFFICIENT VERSUS IDEAL LIFT COEFFICIENT FOR
SEVERAL NACA AIRFOIL SECTIONS. REPRODUCED FROM PERMISSION OF DOVER
PUBLICATIONS, INC.................................................................................................................40
FIGURE 18. VARIATION OF WING ZERO-LIFT AND WAVE DRAG COEFFICIENT VERSUS
MACH NUMBER FOR VARIOUS AIRFOIL THICKNESS RATIOS......................................40
FIGURE 19. AIRFOIL TESTING CONDITIONS...........................................................................................41
FIGURE 20. AIRFOIL TESTING DATA 1.........................................................................................41
FIGURE 23 WING SETTING (INCIDENCE) ANGLE......................................................................42
FIGURE 24 WING SETTING ANGLE CORRESPONDS WITH IDEAL LIFT COEFFICIENT .....42
FIGURE 25 SEVERAL RECTANGULAR WINGS WITH THE SAME PLANFORM AREA BUT
DIFFERENT ASPECT RATIO....................................................................................................43
FIGURE 26 THE TYPICAL EFFECT OF TAPER RATIO ON THE LIFT DISTRIBUTION ...................................44
FIGURE 27 WINGS WITH VARIOUS TAPER RATIOS: ..............................................................................44
FIGURE 28 STRAIGHT TAPERED AND SEMI‐STRAIGHT PLANFORM SHAPES....................44
FIGURE 29 SWEPT BACK PLANFORM ..........................................................................................45

FIGURE 30 LIFT DISTRIBUTION, ROOT-BENDING MOMENT AND SPAN EFFICIENCY FATOR FOR
DIFFERENT TAPER RAIOS...............................................................................................................45
FIGURE 31 MEAN AERODYNAMIC CHORD AND AERODYNAMIC CENTER IN A
STRAIGHT WING.......................................................................................................................45
FIGURE 32 FIVE WINGS WITH DIFFERENT SWEEP ANGLES..................................................................46
FIGURE 33 TYPICAL EFFECT OF SWEEP ANGLE ON LIFT DISTRIBUTION ..........................47
FIGURE 34. WING SHAPE.................................................................................................................47
FIGURE 35 (A) DIHEDRAL AND (B) ANHEDRAL (AIRCRAFT FRONT VIEW) .......................48
FIGURE 36. THE EFFECT OF DIHEDRAL ANGLE ON A DISTURBANCE IN ROLL
(AIRCRAFT FRONT VIEW): (A) BEFORE GUST; (B) AFTER GUST...................................48
FIGURE 37 TYPICAL VALUES OF DIHEDRAL ANGLE FOR VARIOUS WING
CONFIGURATIONS ...................................................................................................................49
FIGURE 38 EXAMPLE OF PRESSURE DISTRIBUTION WITH THE APPLICATION OF A
HIGH-LIFT DEVICE ...................................................................................................................50
FIGURE 39 TYPICAL EFFECTS OF A HIGH-LIFT DEVICE ON WING AIRFOIL SECTION
FEATURES ..................................................................................................................................50
FIGURE 40 MAXIMUM LIFT COEFFICIENT FOR SEVERAL AIRCRAFT .................................50
FIGURE 41 VARIOUS TYPES OF HIGH-LIFT DEVICE: (A) TRAILING EDGE HIGH-LIFT
DEVICE; (B) LEADING EDGE HIGH-LIFT DEVICE..............................................................51
FIGURE 42 LIFT COEFFICIENT INCREMENT BY VARIOUS TYPES OF HIGH-LIFT DEVICE
(WHEN DEFLECTED 60 DEG)..................................................................................................51
FIGURE 43. HIGH-LIFT DEVICE PARAMETERS: (A) TOP VIEW OF THE RIGHT WING; (B)
SIDE VIEW OF THE INBOARD WING (FLAP DEFLECTED) ...............................................52
FIGURE 44 TYPICAL LOCATION OF THE AILERON ON THE WING .......................................52
FIGURE 45. LONGITUDINAL FORCES ACTING ON AN AIRPLANE IN FLIGHT. .........................................55
FIGURE 46. CG IS TOO FAR AFT AT LOW STALL AIRSPEED....................................................................55
FIGURE 47. CG IS TOO FAR FORWARD..................................................................................................56
FIGURE 48. THE MAC IS THE CHORD DRAWN THROUGH THE GEOGRAPHIC CENTER OF THE PLAN
AREA OF THE WING.......................................................................................................................56
FIGURE 49. CMCG WITH Α...................................................................................................................58
FIGURE 50. C_L WITH Α.........................................................................................................................58
FIGURE 51. TAIL CONFIGURATION........................................................................................................60
FIGURE 52. CONVENTIONAL TAIL .........................................................................................................61
FIGURE 53. TOP VIEW OF AFT PORTION OF AIRCRAFT.........................................................................61
FIGURE 54. OPTIMUM TAIL ARM..........................................................................................................62

FIGURE 55. HORIZONTAL TAIL DESIGN PROCEDURE ............................................................................63
FIGURE 56. RECTANGULAR PLANFORM.....................................................................................65
FIGURE 57. NACA 0006 SPECIFICATIONS..............................................................................................65
FIGURE 58. LIFT COEFFICIENT VERSES DRAG COEFFICIENT AND ANGLE OF ATTACK...........................66
FIGURE 59. CL/CD VERSUS ALPHA AND CD VERSUS ALPHA .................................................................66
FIGURE 60. MOMENT COEFFICIENT VERSUS ALPHA.............................................................................66
FIGURE 61 (VT DESIGN PROCEDURE)..........................................................................................71
FIGURE 62 VERTICAL TAIL PARAMETERS.................................................................................72
FIGURE 63 VERTICAL IN WAKE REGION OF HORIZONTAL TAIL .........................................77
FIGURE 64 .TOP VIEW OF WING AILERON..................................................................................79
FIGURE 65. RUDDER DEFLECTION FOR ONE ENGINE CASE...................................................................80
FIGURE 66. DISTRIBUTION OF LIFT AT = 10°....................................................................................82
FIGURE 67.DISTRIBUTION OF VISCOUS DRAG AT = 10°...................................................................82
FIGURE 68. VELOCITY AT SURFACE AT = 10°....................................................................................82
FIGURE 69. PITCHING MOMENT COEFFICIENT VERSUS ANGLE OF ATTACK.........................................83
FIGURE 70. WING CONTRIBUTION TO PITCHING MOMENT.................................................................84
FIGURE 71.WING CONTRIBUTION IN PITCHING MOMENT...................................................................85
FIGURE 72.AFT TAIL CONTRIBUTION TO PITCHING MOMENT .............................................................85
FIGURE 73. AFT TAIL CONTRIBUTION ON PITCHING MOMENT............................................................86
FIGURE 74. AIRCRAFT TOTAL CONTRIBUTION ON PITCHING MOMENT...............................................86
FIGURE 75.STATIC DIRECTIONAL STABILITY..........................................................................................87
FIGURE 77.FUSELAGE GRAPHICAL PARAMETERS .................................................................................88
FIGURE 77. VERTICAL TAIL CONTRIBUTION TO LATERAL STABILITY.....................................................88
FIGURE 79.VERTICAL TAIL GEOMETRY..................................................................................................89
FIGURE 79. STATIC ROLLING STABILITY.................................................................................................89
FIGURE 80.FUSELAGE ROLLING STABILITY [LOW-WING]......................................................................90
FIGURE 81.DIHEDRAL ANGLE POSITIVE STABILIZE ROLLING.................................................................90
FIGURE 82. PROPELLER DIAMETER AGAINST RATED POWER...............................................................91
FIGURE 83. LANDING GEAR ARRANGEMENT........................................................................................94
FIGURE 84.GEOMETRIC DEFINITIONS FOR TAILDRAGGER ARRANGEMENT.........................................95
FIGURE 85.TAILWHEEL UNDERCARRIAGE.............................................................................................96
FIGURE 86. FOR OVERTURNING COEFFICIENT......................................................................................96
FIGURE 87. TURNING RADIUS PARAMETERS........................................................................................97
FIGURE 88.TIRE DEFLECTION ................................................................................................................98

FIGURE 89.TIRE FOOTPRINT..................................................................................................................98
FIGURE 90. TYPICAL CROSS-PLY DATA ..................................................................................................99
FIGURE 91. FUSELAGE DESIGN FLOWCHART......................................................................................103
FIGURE 92. FOUR GENERIC FUSELAGE CONFIGURATIONS. (A) LARGE TRANSPORT AIRCRAFT, (B)
FIGHTER AIRCRAFT......................................................................................................................104
FIGURE 93. INTERNAL ARRANGEMENT OF A CIVIL PASSENGER AND A FIGHTER AIRCRAFT. (A) LOW-
WING PASSENGER AIRCRAFT, AND (B) FIGHTER AIRCRAFT........................................................104
FIGURE 94.TWO TYPES FUSELAGE CONFIGURATIONS: (A) AIRBUS 321 (B) SUKHOI SU-27U.............105
FIGURE 95. EXAMPLES OF VARIATIONS IN HEIGHT BETWEEN MALES AND FEMALES AND DIFFERENT
ETHNIC GROUPS..........................................................................................................................105
FIGURE 96. BASIC INSTRUMENT PANEL..............................................................................................109
FIGURE 97. COCKPIT GEOMETRY FOR A LARGE TRANSPORT AIRCRAFT.............................................110
FIGURE 98. PASSENGER CABIN PARAMETERS ....................................................................................111
FIGURE 99. CABIN WIDTH AND CABIN LENGTH (TOP VIEW)..............................................................112
FIGURE 100. CARGO CONTAINER........................................................................................................113
FIGURE 101. DESIGNED FUSELAGE TOP VIEW....................................................................................114
FIGURE 102.FUSELAGE GRAPHICAL PARAMETERS .............................................................................115
FIGURE 103. SCHEMATIC OF AIRBUS A380 FUEL................................................................................117
FIGURE 104. SCHEMATIC OF FUEL TANK INSIDE THE WING...............................................................118
FIGURE 105. SCHEMATIC OF FUEL TANK DIMENSIONS ......................................................................118
FIGURE 106. CL/CD CURVES VERSUS VELOCITY..................................................................................122
FIGURE 107. DRAG POLAR ..................................................................................................................123
FIGURE 108. POWER VELOCITY CURVE...............................................................................................123
FIGURE 109. ALTITUDE VERSUS TRUE AIRSPEED................................................................................124
FIGURE 110. SPECIFIC POWER VERSUS VELOCITY AT DIFFERENT N ...................................................125
FIGURE 111. ALTITUDE VERSUS AIRSPEED [ENVELOP].......................................................................125
FIGURE 112. ALTITUDE VERSUS AIRSPEED FOR CONSTANT ENERGY .................................................126
FIGURE 113. MAX TURN RATE ............................................................................................................126
FIGURE 114. V_N DIAGRAM................................................................................................................127
FIGURE 115. POWER AVAILABLE & REQUIRED AT DIFF H...................................................................128
FIGURE 116. MAX CEILING DETERMINATION .....................................................................................128

List of Tables
TABLE 1.SUMMARY OF KEY PARAMETERS............................................................................................15
TABLE 2.DA62 MPP SPECIFICATIONS....................................................................................................17
TABLE 3. DA42 MPP SPECIFICATIONS ...................................................................................................17
TABLE 4. AIRCRAFT SPECIFICATIONS.....................................................................................................18
TABLE 5. RANGE AND ENDURANCE CASE RULES..................................................................................22
TABLE 6.RANGE AND ENDURANCE CASE RESULTS ...............................................................................22
TABLE 7. DRAG POLAR CORRELATIONS CONSTANTS............................................................................27
TABLE 8. DESIGN REQUIREMENTS........................................................................................................43
TABLE 9. WING CONFIGURATION PARAMETERS 1 ...............................................................................53
TABLE 10. WING CONFIGURATION PARAMETERS 2 .............................................................................54
TABLE 11. AVERAGE GROUP WEIGHT BREAKDOWN ............................................................................56
TABLE 12. STRUCTURAL WEIGHT BREAKDOWN...................................................................................57
TABLE 13. CG ESTIMATION ...................................................................................................................57
TABLE 14 TYPICAL VALUES FOR HORIZONTAL AND VERTICAL TAIL VOLUME COEFFICIENTS...............64
TABLE 15. HORIZONTAL TAIL PARAMETERS .........................................................................................70
TABLE 16 (VALUES FOR VERTICAL TAIL VOLUME COEFFICIENT)...........................................................72
TABLE 17. VERTICAL TAIL PARAMETERS ...............................................................................................76
TABLE18.PARAMATERS OF WING & AFT TAIL CONTRIBUTION ON PITCHING MOMENT.....................84
TABLE 19. FUSELAGE PARAMETERS......................................................................................................88
TABLE 20.PARAMETERS OF VERTICAL TAIL...........................................................................................89
TABLE 21 (TIP SPEED LIMITS) ................................................................................................................92
TABLE 22. SPECIFICATIONS OF OUR PROPELLER ..................................................................................93
TABLE 23. TIRE SPECIFICATIONS ...........................................................................................................99
TABLE 24. FUNCTIONAL ANALYSIS OF THE FUSELAGE........................................................................102
TABLE 25. COCKPIT DESIGN PARAMETERS .........................................................................................107
TABLE 26.AISLE WIDTH REQUIREMENTS FROM FAR 25 FOR TRANSPORT AIRCRAFT ........................111
TABLE 27. RECOMMENDED CABIN DATA (IN CENTIMETERS).............................................................111
TABLE 28. DESIGNED FUSELAGE PARAMETERS ..................................................................................115
TABLE 29. DENSITY OF VARIOUS FUELS AT 15◦C ................................................................................116
TABLE 30. COST ESTIMATES OF PERVIOUS AIRCRAFT ........................................................................130
TABLE 31. OUR DESIGNED AIRCRAFT EXPECTED COST.......................................................................130

List of Symbols and Abbreviations
a −
g −
s −
r −
G − ℎ
f −
q −
w −
t − ℎ
y −
L1/4 − ℎ
¶ /¶d −
¶ /¶d −
r −
( / ) − − −
( / ) − − −
−
−
−
−
−
−
−
− ℎ ℎ
ĉ − ℎ
. . −
−
−
0 − −
−
−

− ℎ ℎ ℎ ℎ
−
− ℎ
−
− ℎ
− ℎ
−
−
–
−
−
−
−
−
− ℎ
−
− ℎ
− ℎ
−
−
−
−
/ −
−
−
−
−
−
−
_ −
− ℎ
/ − ℎ

/ − ℎ
− ℎ ℎ ℎ
− /
−
−
−
− ℎ
/ −
0 − ℎ
− ℎ
− ℎ
− ℎ
– ℎ
− ℎ
−

1. INTRODUCTION TO DESIGN
Aircraft design is essentially a branch of engineering design. Design is primarily an analytical
process, which is usually accompanied by drawing/drafting. Design contains its own body of
knowledge, independent of the science-based analysis tools usually coupled with it. Design is a more
advanced version of a problem-solving technique that many people use routinely. Design is exciting,
challenging, satisfying, and rewarding. The general procedure for solving a mathematical problem is
straightforward. Design is much more subjective, there is rarely a single “correct” answer. The world
of design involves many challenges, uncertainties, ambiguities, and inconsistencies.
Air passengers demand more comfort and more environmentally friendly aircraft. Hence many
technical challenges need to be balanced for an aircraft to economically achieve its design
specification. Aircraft design is a complex and laborious undertaking with a number of factors and
details that are required to be checked to obtain optimum the final envisioned product. The design
process begins from scratch and involves a number of calculations, logistic planning, design and real
world considerations, and a level head to meet any hurdle head on.
Every airplane goes through many changes in design before it is finally built in a factory. These
steps between the first ideas for an airplane and the time when it is actually flown make up the design
process. Along the way, engineers think about four main areas of aeronautics: Aerodynamics,
Propulsion, Structures and Materials, and Stability and Control.
Aerodynamics is the study of how air flows around an airplane. In order for an airplane to fly at all,
air must flow over and under its wings. The more aerodynamic, or streamlined the airplane is, the less
resistance it has against the air. If air can move around the airplane easier, the airplane's engines have
less work to do. This means the engines do not have to be as big or eat up as much fuel which makes
the airplane more lightweight and easier to fly. Engineers have to think about what type of airplane
they are designing because certain airplanes need to be aerodynamic in certain ways. For example,
fighter jets maneuver and turn quickly and fly faster than sound (supersonic flight) over short
distances. Most passenger airplanes, on the other hand, fly below the speed of sound (subsonic flight)
for long periods of time.
Propulsion is the study of what kind of engine and power an airplane needs. An airplane needs to
have the right kind of engine for the kind of job that it has. A passenger jet carries many passengers
and a lot of heavy cargo over long distances so its engines need to use fuel very efficiently. Engineers
are also trying to make airplane engines quieter so they do not bother the passengers onboard or the
neighborhoods they are flying over. Another important concern is making the exhaust cleaner and
more environmentally friendly. Just like automobiles, airplane exhaust contains chemicals that can
damage the earth's environment.
Structures and Materials is the study of how strong the airplane is and what materials will be used
to build it. It is really important for an airplane to be as lightweight as possible. The less weight an
airplane has, the less work the engines have to do and the farther it can fly. It is tough designing an
airplane that is lightweight and strong at the same time. In the past, airplanes were usually made out
of lightweight metals like aluminum.
Stability and Control is the study of how an airplane handles and interacts to pilot input and feed.
Pilots in the cockpit have a lot of data to read from the airplane's computers or displays. Meanwhile,
the airplane should display information to the pilot in an easy-to-read and easy-to-understand way.
The controls in the cockpit should be within easy reach and just where the pilot expects them to be.
It is also important that the airplane responds quickly and accurately to the pilot's instructions and
maneuvers.
“A beautiful aircraft is the expression of the genius of a great engineer who is also a great artist.”

2. RAVEN
2.1 Description
Raven is considered to be new class aircraft; a light twin engine propeller-driven aircraft. If this
aircraft marketed, it ‘ll be for small business flight, schools and the government. It is 1700 HP propeller-
driven. The nominal cruising altitude is 10,000 feet PA and the aircraft is capable of carrying six
passengers in addition to the pilot, co-pilot with no attendants involved. Its state of the art avionics
package will attract many customers and make the pilot’s job much easier.
2.2 Summary of Key Parameters
Table 1.Summary of Key Parameters
2.3 Configuration Layout
As shown in Figure 1 the configuration layout of our airplane, which will have 6 seats for passengers
and two seats for pilots. The seats for the passengers will be 2 seats at the front, 2 seats at the middle
and 2 seats at the end of the airplane. In according to the cabin will be also 2 seats for pilot and co-
pilot.
Figure 1. Configuration Layout
Basic Performance
Max Airspeed 260 Knts
Cruise Speed 240 Knts
Service Ceiling 20,000 ft
Range 450 sm
Wing Geometry
Wing Span 54 ft
Wing Chord 9 ft
Aspect Ratio 6.1
Wing Surface 486
Wing Loading 15.94
Performance Parameters
Engine Type 2*R-1830-
64 850 HP
Static Thrust HP 1700 HP
SFC 0.49
lb./(hph)
MGWTO 7460 Lbs.

3. Requirements Analysis
3.1 Requirements Summary
 The design will be 14 CFR Part 23 compliant.
 The design team will utilize Part 21 Certification procedures.
 The aircraft will be capable of carrying 6 passengers in addition to the pilot and co-pilot.
 The aircraft will have a range of 450 miles.
 The R-1830-64 engine incorporates a FADEC system for reduced maintenance
costs as well as an electric starter for weight reduction.
 The aircraft will be capable of short take-offs and landings.
 The aircraft will capable fly at 240 Knots at Cruise.
3.2 Mission Profile
Figure 2 Initial mission profile
It was initially assumed that the presumed roadable aircraft should start-off with a ground cruise
phase, where the pilot will drive the vehicle in its roadable mode to the airport, from which he desires
to take-off. The weight ratios of such a phase would not be expected to be in aircraft books, but it was
decided that statistical data of ground vehicles can be used where one of the most frequently used
specifications in cars is mileage. Mileage means the number of miles the car can cross per gallon of
fuel used. This specification could be chosen from statistical data of cars similar to our passenger
number and general configuration. Then, the aerial mission would start with a warm-up phase, a taxi
phase and finally the take-off. Then the vehicle will climb to 10,000ft which was chosen to be the
cruising altitude of the vehicle and then, it would cruise over a range of 250 nautical mile. Afterwards,
it would be expected to loiter for about 33 minutes and then descend and perform the landing.
However, after consultation with our faculty advisor, Dr. Nader Aboulfotouh, it was decided to
eliminate the ground cruise phase, due to the fact that the aircraft can be refueled at the airport from
which it will departure from. What follows is the final mission profile for the roadable vehicle to be
designed:
As seen from figure 3, the ground cruise phase was removed due to its uselessness in case the
aircraft is refueled at the airport from which it will take-off. The targets of the previous configuration
remained the same in terms of loiter time and cruise range and altitude.

3.3 Reference Design Concepts
DA62 MPP
Power Plant Mass and loading Performance
Engine 2×Austro Engine AE
330 with 180 HP.
Empty
weight
3,505 lbs. Max. cruise speed
(14,000 ft, MCP)
190 kts TAS
Propeller 2×MT propeller 3 blade
constant speed propeller
MTOM 5,071 lbs. Min. operation
speed
76 kts TAS
Useful
load
1,565 lbs. Max. range at 50%
power
732nm-
1,264nm
Passengers 7 Max service ceiling 20,000 ft
Table 2.DA62 MPP Specifications
DA 42 MPP
Power Plant Mass and loading Performance
Engine 2×Austro Engine AE
300 with 168 HP.
Empty
weight
3,008 lbs. Max. cruise speed
(14,000 ft, MCP)
171 kts TAS
Propeller 2×MT propeller 3 blade
constant speed propeller
MTOM 5,071 lbs. Min. operation
speed
71 kts TAS
Useful
load
1,398 lbs. Max. range at 50%
power
698 nm /
1,065 nm
Passengers 4 Max service ceiling 18,000 ft
Table 3. DA42 MPP Specifications
Figure 3. DA 62 and DA 42 MPP Specs

4. Weight estimation
The purpose of this part of the design process is to produce an initial sizing of the vehicle take-of
weight, empty weight and fuel weight. This means that initial estimates of weight fractions for each
mission phase must be assumed. Since there is no similar existing class of vehicles in aircraft design
literature such as (Roskam, 1985) and (Sedrey). it was intuitive to first perform a statistical study on
the existing roadable aircrafts. It was surprising that there were more than 18 actual tested roadable
aircrafts designed since the 1940s and even more than that number in conceptual designs that never
reached the detailed design phases. Of those 18, we were able to manifest information about the
weights of 12 of them, which served as adequate raw materials for a statistical relation. However, all
of these vehicles had only 2 passengers, which meant that it was not guaranteed that the resulting
statistical relation would prove useful to our weight estimation problem.
We want to define some parameters . The fuel-fraction method will be used to determine
in which the airplane mission profile, shown in the subsequent Figure 2 , will be broken down into a
number of mission phases. The fuel used during each phase is found from a simple calculation or
estimated on the basis of experience. Each phase has a number and an associated fuel fraction.
Aircraft Specifications
Aircraft type Twin Engine Propeller Driven
Number of passengers 8
Weight of Each Passenger 175 Lbs.
Baggage for one passenger 30 Lbs.
Reserved fuel 20%
Weight of the payload 0
trapped fuel 5%
(L/D ) 9
0.6
0.82
(L/D) 10
0.6
0.72
240 Knots
0.8*
Loiter 15 minutes
Range of cruise 450 sm
Aspect ratio 6
Historical data A= .1048 , B=1.03505
Table 4. Aircraft Specifications

4.1 Fuel-Fraction Method [Twin-Engine Propeller-Driven Airplane]
A numerical value will be assigned to the fuel-fraction corresponding to each mission phase. This
can be done as follows.
Refer to (Roskam, 1985), table 2.1.
 Phase 1 : Engine start and warm-up
= .992
 Phase 2 : Taxi
= .996
 Phase 3 :Take off
= .996
 Phase 4: Climb to cruise altitude
= .99
 Phase 5: Cruise
The fuel-fraction associated with cruise can be calculated from Breguet’s range equation which is
expressed as follows for a propeller-driven airplane
= 375 ∗ ∗ ∗ ln
Where
This equation Yields
4
= .9142
 Phase 6: Loiter
The fuel-fraction associated with cruise can be calculated from Breguet’s endurance equation
which is expressed as follows for a propeller-driven airplane
=
175
∗ ∗ ∗ ln( )
Where
ℎ , ℎ
This equation Yields
= .9958

 Phase 7 : Descent
= .992
 Phase 8:Landing ,Taxi and shut down
= .992
Since
= = ( ) ( )
By substitution in the above expression Yields
= .8464
4.2 Airplane Gross-Weight Estimation
Refer to (Roskam, 1985), table 2.15. Historical data for this aircraft
A = .1048 , B = 1.03505
Since
= 1 − 1 − (1 + ) −
= +
Thus
= .8678
= 820
The take-off gross weight, ,can be calculated from
( ) = + ∗ ( ∗ − )
Hence
= 7460
= = 1 − ∗
ℎ = = 1 − (1 + )
ℎ = = −
= .
=
=

5. Weight sensitivity
After preliminary sizing of a new airplane, it is mandatory to conduct sensitivity studies to find out
which parameters drive the design and which area of technological change must be pursued.
5.1 Sensitivity of TO
= ∗ ( − ∗ (1 − ) ∗ )
Where = the air plane growth factor due to payload
This equation yield that
= 4.082
5.2 Sensitivity of TO
= ∗
−
This equation yield
= 1.7675
5.3 Sensitivity of TO , , , ,
The sensitivity of the take-off weight , to any parameter, , can be written as follows:
5.3.1 For the case involving the ratio dependent on range :
=
5.3.2 For the case involving the ratio dependent on endurance :
=
Where the factor is expressed as
= − ∗ [ (1 − ) − ] ∗ (1 + ) ∗
Tis equation Yields :
= 3.0947 ∗ 10
And the terms & are called .
For a propeller driven airplane, are given as follows ;

Parameter = Range case Endurance case
Range =R
= 375 ∗
Not applied
Endurance=E Not applied
= 375 ∗
Specific fuel consumption =
= 375 ∗ = 375 ∗
Propeller Efficiency =
= − 375 ∗ = − 375 ∗
Velocity Not applied
= 375 ∗
Lift TO Drag =
= − 375 ∗ = 375 ∗
Table 5. Range And Endurance Case Rules
Where R in sm & V in mph
= − 375 ∗ ∗
Parameter = Range case Endurance case
Range =R
= 6.7093
ibs
sm
---
Endurance=E ---
= 1519.8
ibs
hour
Specific fuel consumption =
= 5791.8 hp. hour = 633.24 hp . hour
Propeller Efficiency =
= −4237.9 ibs = −527.7 ibs
Velocity ---
= 1.72
ibs
mph
Lift TO Drag = = −386.1231 ibs = −38 ibs
Table 6.Range And Endurance Case Results

6. Matching Plot
6.1 Introduction
It’s necessary to get limit which design on it. To get this limit we do sizing for plane performance
according to following categories
 Stall Speed.
 Take-off Field Length.
 Landing Field Length.
 Climb Rate (AEO).
 Cruise Speed.
 Time to Climb.
The method, to be used, will allow a rapid estimation of airplane design parameters which are
 Wing Area
 Required Take-Off Power
 Clean, Take-Off and Landing Maximum Lift Coefficients , and .
6.2 Stall Speed Sizing
We design according to FAR 23 certified twin-engine airplanes, usually propeller-driven airplanes,
may not have a stall speed greater than 61 at . The power-off stall speed of an airplane may
be determined from
=
2
/
Where = 1.22
At sea level, the free stream density is given by
= 0. 002377 slug/ft
For a given value of and a specified maximum allowable stall speed at some altitude. The
maximum allowable wing loading can be identified.
In case of our airplane, according to FAR 23
≤ 61 = 102.907 /
Hence, we get
= 15.3732

Figure 4. Stall Speed Sizing
6.3 Take-off Distance Sizing
To make sizing in case of Take-off of airplanes we should get relation between W/S and W/p
= 1.66
Figure 5. Definition of FAR 23 Take-off Distances
In case of our airplane, take-off distance is assumed to be
= 1500
= 4.9 + 0.009( ) = 1500
When solving it we get
= 218.4626
ℎ
=
∗ ∗
/

At sea level
= 1
Assuming take-off maximum lift coefficient to be
= 1.7
=
.
/
Figure 6. Take-off Distance sizing
6.4 Landing Distance Sizing
Figure 7. Definition of FAR 23 Landing Distances

To make sizing in case of landing we should get .Frist we calculate from following equation
=
0.265
In case of our airplane, landing distance is assumed to be
= 1500
→ = 75.2355 = 126.997 /
=
1
2
Assuming landing maximum lift coefficient to be
= 2.05
→ = 39.286
For a single-engine propeller-driven airplane, the average landing to take-off weight is
approximately
= 0.99
So we can calculate from following equation
= = 38.893
Figure 8. Landing Distance Sizing

6.5 FAR 23 Climb Requirements Sizing
All airplanes must meet certain climb rate or climb gradient requirements. To size an airplane for
climb requirements, we should have an estimate for the airplane drag polar.
6.5.1 Drag Polar Estimation
The drag coefficient of an airplane express by following equation
= +
Where
→ =
log10
= + 10
Where ‘ ’ and ‘b’ are the correlation coefficients obtained, for a given value of airplane equivalent
skin friction ,from Table 3.4 (Roskam, 1985).
In case of our airplane, assuming equivalent skin friction coefficient
= 0.005
To get ‘f’ we need to calculate
log10
= + 10
Where ′ ′ and ′ ′ are the regression-line constants obtained from Table 3.5, (Roskam, 1985). We
have from Weight estimation section.
Thus
= 1108.22
Also
= 5.54148
Since
= = 404.203
Hence, the zero-lift drag coefficient is
= 0.014
Correlation Coefficients Regression-Line Constants
a = -2.301 C = 0.8635
b = 1 d = 0.5632
Table 7. Drag Polar Correlations Constants

6.5.2 Rate of Climb Sizing
The rate of climb is given by
= 33,000
= −
19
3
2
.
In case of our airplane, according to FAR 23.65 (AEO):
≥ 300
Configuration: gear up, take-off flaps, max.cont.power on all engine. To maximize , it is evident
to make as large as possible
3
2
=
1.345( ∗ )
3
4
1
4
Where
= 6.1 = 0.8 = 0.014 = 0.82
3
2
= 12.68
Thus;
=
0.82
+
.
240.92
Figure 9. Rate of Climb Sizing

6.6 Climb Gradient Sizing
The climb gradient parameter is given by
=
+
=
18.97
/
In case of our airplane, according to FAR 23.65 (AEO):
≥
1
12
Thus
=
Where
= 1.5
= + ∆ +
∗ ∗
Considering additional zero-lift drag coefficient due to take-off flaps
∆ = 0.015
= 0.1682
So
= 8.91795
=
+ 0.11213
1.2247
All of the above yields
=
15.56
∗
1
.
Figure 10. Climb Gradient Sizing

6.7 Cruise Speed Sizing
The cruise speed for propeller-driven airplanes is proportional to the following
∝
∗
/
Thus
∝
Where
=
In our case of our airplane cruise at = 240 can get from figure 3.28 in (Roskam, 1985).
So
= 1.6
( 10,000 ) = 0.7387
=
0.71015
Figure 11. Cruise Speed Sizing

6.8 Time-to-Climb Sizing
There are linear relationship between rate of climb and altitude ℎ
= 1 −
ℎ
ℎ
Where
≡
≡
For a given value of airplane ℎ ,from Table 3.7, (Roskam, 1985).
ℎ = 20,000
Can calculate from following equation
=
ℎ
ln 1 −
ℎ
ℎ
−1
In our plane we assume
= 10
ℎ = 10,000
= 2772.5887
For a propeller-driven airplane, since
= 33,000
Where
= −
19
3
2
.
= 693.147
= 0.021
=
0.82
0.021 +
.
122.562

Figure 12. Time to Climb Sizing
6.9 Matching curves
We matching curves of all sizing case to get the wing loading in [ ] and the power loading in
From design point .
Figure 13. Matching Plot

It is obvious that the design point is that at which stall and cruise curves intersect. At the design
point the corresponding
Where
→ = 15.3732
→ = 5.08
ℎ
From section [4.2] take-off gross weight was estimated to be
= 7460
Thus
= 485.26
= 1468.5 ℎ

7. Wing design
The wing may be considered as the most important component of an aircraft, since a fixed-wing
aircraft is not able to fly without it. Since the wing geometry and its features influence all other aircraft
components, we begin the detail design process with wing design. The primary function of the wing is
to generate sufficient lift force or simply lift (L). However, the wing has two other productions, namely
the drag force or drag (D) and nose-down pitching moment (M ). While a wing designer is looking to
maximize the lift, the other two (drag and pitching moment) must be minimized. In fact, a wing is
considered as a lifting surface where lift is produced due to the pressure difference between the lower
and upper surfaces. Aerodynamics textbooks are a good source to consult for information about
mathematical techniques to calculate the pressure distribution over the wing and for determining the
flow variables.
During the wing design process, 18 parameters must be determined. They are as follows:
 Wing reference area ( S or );
 Number of wings;
 Vertical position relative to the fuselage (high, mid-, or low wing);
 Horizontal position relative to the fuselage;
 Cross-section (or airfoil);
 Aspect ratio (AR);
 Taper ratio (λ);
 Tip chord ( );
 Root chord ( );
 Mean aerodynamic chord (MAC or C );
 Span (b);
 Twist angle (or washout) ( );
 Sweep angle (Λ);
 Dihedral angle (Γ);
 Incidence ( ) (or setting angle, );
 High-lifting devices such as flap;
 Aileron;
 Other wing accessories.
Of the above long list, only the first one (i.e., planform area) has been calculated so far (during the
preliminary design step).
= = 486

Figure 14.Wing Design Procedure
7.1 Number of wings
A number of wings higher than three is not practical Figure 15 illustrates a front view of three
aircraft with various configurations. Nowadays, modern aircraft almost all have a monoplane.
Currently, there are a few aircraft that employ a biplane, but no modern aircraft is found to have three
wings. In the past, the major reason to select more than one wing was manufacturing technology
limitations. A single wing usually has a longer wing span compared with two wings (with the same
total area). Old manufacturing technologies were not able to structurally support a long wing, to stay
level and rigid. With advances in manufacturing technologies and also new strong aerospace materials
(such as advanced light aluminum and composite materials), this reason is no longer valid. Another
reason was the limitations on the aircraft wing span. Hence a way to reduce the wing span is to
increase the number of wings. The most significant is the requirement for aircraft controllability. An
aircraft with a shorter wing span delivers higher roll control, since it has a smaller mass moment of
inertia about the x-axis. Therefore, if one is looking to roll faster, one option is to have more than one
wing leading to a shorter wing span.

Figure 15.Three options in number of wings: (a) Monoplane, (b) Biplane
And (c) Tri-wing
So from technical and commercial we will choose monoplane aircraft .
7.2 Wing vertical location
One of the wing parameters that could be determined at the early stages of the wing design
process is the wing vertical location relative to the fuselage center line. This wing parameter will
influence the design of other aircraft components directly, including aircraft tail design, landing gear
design, and center of gravity. In principle, there are four options for the vertical location of the wing. .
The primary criterion to select the wing location originates from operational requirements, while
other requirements such as stability and producibility are the influencing factors in some design cases.
Figure 16.Options in vertical wing positions: (a) High wing; (b) Mid-wing; (c) Low wing; and (d)
Parasol wing
7.2.1 High wing
The high-wing configuration Figure 15 (a) has several advantages and disadvantages that make it
suitable for some flight operations, but unsuitable for other flight missions.
7.2.1.1 Advantages
 Eases and facilitates the loading and unloading of loads and cargo into and out of cargo aircraft.
 Facilitates the installation of an engine on the wing, since the engine (and propeller) clearance is higher
(and safer).
 Facilitates the installation of a strut.
 The aircraft structure is lighter when struts are employed (as item 4 implies).
 Facilitates aircraft control for a hang glider pilot, since the aircraft center of gravity is lower than the
wing.

 Increases the dihedral effect (Clβ). It makes the aircraft laterally more stable. The reason lies in the higher
contribution of the fuselage to the wing dihedral effect (ClβW)
 The wing will produce more lift compared with a mid- and low wing, since two parts of the wing are
attached at least on the top part
 For an engine that is installed under the wing, there is less possibility of sand and debris entering the
engine and damaging the blades and propellers.
 The aerodynamic shape of the fuselage lower section can be smoother.
 There is more space inside the fuselage for cargo, luggage, or passengers.
 The wing drag produces a nose-up pitching moment, so it is longitudinally destabilizing. This is due to
the higher location of the wing drag line relative to the aircraft center of gravity ( > 0).
7.2.1.2 Disadvantages
 The aircraft tends to have more frontal area (compared with mid-wing). This will increase aircraft drag.
 The ground effect is lower, compared with low wing. During take-off and landing, a high-wing
configuration is not the right option for short take-off and landing (STOL) aircraft
 The landing gear is longer if connected to the wing. This makes the landing gear heavier and requires
more space inside the wing for the retraction system. This will further make the wing structure heavier.
 The wing produces more induced drag (Di) due to the higher lift coefficient.
 The horizontal tail area of an aircraft with a high wing is about 20% larger than the horizontal tail area
with a low wing. This is due to more downwash of a high wing on the tail.
 A high wing is structurally about 20% heavier than a low wing
 The aircraft lateral control is weaker compared with mid-wing and low wing, since the aircraft has more
laterally dynamic stability.
7.2.2 Low wing
In this section, the advantages and disadvantages of a low-wing configuration Figure 16 (c)
7.2.2.1 Advantages
 The aircraft take-off performance is better, compared with a high-wing configuration, due to the ground
effect.
 The retraction system inside the wing is an option, along with inside the fuselage.
 The landing gear is shorter if connected to the wing. This makes the landing gear lighter and requires
less space inside the wing
 The aircraft is lighter compared with a high-wing structure
 The aircraft frontal area is less.
 the aircraft drag is lower
 It is more attractive to the eyes of a regular viewer.
 The aircraft has higher lateral control compared with a high-wing configuration, since the aircraft has
less lateral static stability, due to the fuselage contribution to the wing dihedral effect (ClβW ).
 The wing has less downwash on the tail, so the tail is more effective
 The wing drag produces a nose-down pitching moment, so a low wing is longitudinally stabilizing. This
is due to the lower position of the wing drag line relative to the aircraft center of gravity ( < 0).
7.2.2.2 Disadvantages
 The wing generates less lift, compared with a high-wing configuration, since the wing has two separate
sections
 With the same token as item 1, the aircraft will have a higher stall speed compared with a high-wing
configuration, due to a lower .
 the take-off run is longer
 The wing makes a lower contribution to the aircraft dihedral effect, thus the aircraft is laterally
dynamically less stable.
 The aircraft has a lower landing performance, since it needs more landing run

7.2.3 Mid-Wing
The features of the mid-wing configuration Figure 16 (b) stand somewhere between the features
of a high-wing configuration and the features of a lowing configuration. The major difference lies in
the necessity to cut the wing spar in half in order to save space inside the fuselage. However, another
alternative is not to cut the wing spar and to let it pass through the fuselage, which leads to an
occupied space of the fuselage.
7.2.3.1 Advantages and Disadvantages
 The aircraft structure is heavier, due to the necessity of reinforcing the wing root at the intersection with
the fuselage
 The mid-wing is more expensive compared with high- and low-wing configurations
 The mid-wing is more attractive compared with the two other configurations.
 The mid-wing is aerodynamically streamlined compared with the two other configurations.
 A strut is usually not used to reinforce the wing structure.
7.2.4 The selection process
From technical Data and commercial Data we will use Low Wing configuration
7.3 Airfoil selection
This section is devoted to the process of determining the airfoil section for a wing. It is appropriate
to claim that the airfoil section is the second most important wing parameter, after the wing plan form
area. The airfoil section is responsible for the generation of the optimum pressure distribution on the
top and bottom surfaces of the wing such that the required lift is created with the lowest aerodynamic
cost (i.e., drag and pitching moment). If you are not ready to design your own airfoil, you are
recommended to select a proper airfoil from the previously designed and published airfoil sections.
7.3.1 Airfoil Selection Criteria
 The airfoil with the highest maximum lift coefficient ( ).
 The airfoil with the proper ideal or design lift coefficient ( ).
 The airfoil with the lowest minimum drag coefficient ( ).
 The airfoil with the highest lift-to-drag ratio
 The airfoil with the highest lift curve slope ( ).
 The airfoil with the lowest (closest to zero; negative or positive) pitching moment coefficient (Cm).
 The proper stall quality in the stall region (the variation must be gentle, not sharp)
 The airfoil must be structurally reinforceable. The airfoil should not be so thin that spars cannot be placed
inside.
 The airfoil must be such that the cross-section is manufacturable.
 The cost requirements must be considered.
 Other design requirements must be considered. For instance, if the fuel tank has been designated to be
placed inside the wing inboard section, the airfoil must allow sufficient space for this purpose.

7.3.2 Practical Steps for Wing Airfoil Section Selection
 Determine the average aircraft weight ( ) in cruising flight
=
+
2
∶ = = 7460
 Calculate the aircraft ideal cruise lift coefficient ( ). In a cruising flight, the aircraft weight is equal to
the lift force so:
=
2
→ =
2 ∗ 7460
. 001745 ∗ (240 ∗ 1.6878) ∗ 486
=. .10722
 Calculate the wing cruise lift coefficient (C ). Basically, the wing is solely responsible for the
generation of the lift. However, other aircraft components also contribute to
=
. 95
→ = .11286 ≅ .113
 Calculate the wing airfoil ideal lift coefficient ( ), The wing is a three-dimensional body, while an
airfoil is a two-dimensional section we have to resort to an approximate relationship. In reality, the span
is limited, and in most cases, the wing has a sweep angle and a non-constant chord, so the wing lift
coefficient will be slightly less than the airfoil lift coefficient. For this purpose, the following approximate
equation is recommended at this moment
=
. 9
→ ≅ .1254
 Calculate the aircraft maximum lift coefficient ( ):
=
2
→ ≅ 1.2334
 Calculate the wing maximum lift coefficient ( ). With the same logic that was described in step
3, the following relationship is recommended:
=
. 95
→ ≅ 1.2983 ≅ 1.3
 Calculate the wing airfoil gross maximum lift coefficient ( ):
=
. 9
→ ≅ 1.444
 Select/design the HLD (type, geometry, and maximum deflection)
 Calculate the wing airfoil net maximum lift coefficient ( ):
= − ∆ → ≅ 1: 1.5
 Identify airfoil section alternatives that deliver the desired (step 4) and (step 10), This is an
essential step Figure 17 shows a collection of and for several NACA airfoil sections in just
one graph.

Figure 17 Maximum lift coefficient versus ideal lift coefficient for several NACA airfoil sections.
Reproduced from permission of Dover Publications, Inc
 If the wing is designed for a high subsonic passenger aircraft, select the thinnest airfoil (the lowest
The reason is to reduce the critical Mach number ( ). Figure 5.24 shows the typical variation
of the wing zero-lift and wave-drag coefficient versus Mach number for four wings with airfoil thickness
ratio as a parameter.
ℎ =
240
667
≅ .36 → ℎ ℎ
Figure 18. Variation of wing zero-lift and wave drag coefficient versus Mach number for various
airfoil thickness ratios
from technical data , imperial data we will chose NACA 1408 airfoil

Figure NACA 1408 airfoil
Figure 19. Airfoil Testing Conditions
Figure 20. Airfoil Testing Data 1

7.4 Wing Incidence
The wing incidence ( ) is the angle between the fuselage center line and the wing chord line at its
root Figure 23 It is sometimes referred to as the wing setting angle ( ).
Figure 23 Wing setting (incidence) angle
Figure 24 Wing setting angle corresponds with ideal lift coefficient

7.4.1 The wing incidence must satisfy the following design requirements:
 The wing must be able to generate the desired lift coefficient during cruising flight.
 The wing must produce minimum drag during cruising flight
 The wing setting angle must be such that the wing angle of attack could be varied safely (in fact
increased) during take-off operation.
 The wing setting angle must be such that the fuselage generates minimum drag during cruising flight
(i.e., the fuselage angle of attack must be zero in cruise).
The typical wing incidence number for the majority of aircraft is between 0 and 4 deg. From
trimming-stability data and airfoil data the incident angle.
= .
7.5 Aspect Ratio
The aspect ratio (AR) is defined as the ratio between the wing span b (see Figure 25) and the
wing MAC or ̅:
=
̅
Figure 25 Several rectangular wings with the same planform area but different aspect ratio
Wing Area 486
Span length 54
Table 8. Design Requirements
The wing planform area with a rectangular or straight tapered shape is defined as the span times
the MAC:
= ̅
Thus, the aspect ratio shall be redefined as:
=
̅
= → =
54
486
= 6 → ̅ = 9

7.6 Taper Ratio
The taper ratio ( ) is defined as the ratio between the tip chord ( ) and the root chord ( ).6
This definition is applied to the wing, as well as the horizontal tail and the vertical tail. Root chord
and tip chord are illustrated in Figure 5.31:
=
Figure 28 Straight Tapered and Semi‐Straight Planform Shapes
Figure 27 Wings with various taper ratios:
(a) Rectangle (λ = 1); (b) Trapezoid 0 < λ < 1
(straight tapered); and (c) Triangle (delta) λ = 0
Figure 26 The typical effect of taper ratio on the lift
distribution

Figure 29 Swept Back Planform
Figure 30 LIFT DISTRIBUTION, ROOT-BENDING MOMENT AND SPAN EFFICIENCY FATOR FOR
DIFFERENT TAPER RAIOS
Figure 31 Mean aerodynamic chord and aerodynamic center in a straight wing

7.6.1 Calculating and selecting taper ratio for wing
 The wing taper will change the wing lift distribution. This is assumed to be an advantage of the taper,
since it is a technical tool to improve the lift distribution.
 The taper will reduce the wing weight, since the center of gravity of each wing section (left and right)
will move toward the fuselage center line. This results in a lower bending moment at the wing root. This
is an advantage of the taper
 Due to item 3, the wing mass moment of inertia about the x-axis (longitudinal axis) will be decreased.
Consequently, this will improve the aircraft lateral control
 The taper will influence the aircraft static lateral stability (Clβ ), since the taper usually generates a sweep
angle (either on the leading edge or on a quarter chord line).
From technical data. We will take our configurations
According to → Straight Taper
For suitable sweep angle :
= , = , = , Λ = tan
−
Taber ratio
= =.
7.7 Sweep Angle
The angle between a constant percentage chord line along the semi span of the wing and the
lateral axis perpendicular to the aircraft center line (y-axis) is called the leading edge sweep (LE). The
angle between the wing leading edge and the y-axis of the aircraft is called the leading edge sweep
(LE). Similarly, the angle between the wing trailing edge and the longitudinal axis (y-axis) of the aircraft
is called the trailing edge sweep (TE). In the same fashion, the angle between the wing quarter chord
line and the y-axis of the aircraft is called the quarter chord sweep (C/4). And finally,the angle between
the wing 50% chord line and the y-axis of the aircraft is the 50% chord sweep (C/2).
Figure 32 Five wings with different sweep angles

Figure 33 Typical effect of sweep angle on lift distribution
7.7.1 Advantages of sweep
 Improving the wing aerodynamic features (lift, drag, and pitching moment) at transonic, supersonic, and
hypersonic speeds by delaying the compressibility effects.
 Adjusting the aircraft center of gravity.
 Improving static lateral stability
 Impacting longitudinal and directional stability
 Increasing pilot view (especially for fighter pilots).
For our configuration :
Figure 34. Wing Shape
=
−
, = .

7.8 Dihedral Angle
When you look at the front view of an aircraft, the angle between the chord line plane of a wing
with the xy plane is referred to as the wing dihedral (). The chord line plane of the wing is an imaginary
plane that is generated by connecting all chord lines across the span. If the wing tip is higher than the
xy plane, the angle is called positive dihedral or simply dihedral, but when the wing tip is lower than
the xy plane, the angle is called negative dihedral or anhedral (see Figure 35).
Figure 35 (a) Dihedral and (b) anhedral (aircraft front view)
Figure 36. The effect of dihedral angle on a disturbance in roll (aircraft front view): (a) before
gust; (b) after gust
The necessary restoring rolling moment. The lateral static stability is primarily represented
by a stability derivative called the aircraft dihedral effect ( = ) that is the change in aircraft rolling
moment coefficient due to a change in aircraft sideslip angle (β). Observe a level-wing aircraft that has
experienced a disturbance (see Figure 36Figure 36) which has produced an undesired rolling moment
(e.g., a gust under one side of the wing). When the aircraft rolls, one side of the wing (say the left)
goes up, while the other side (say the right) goes down. This is called a positive roll. The right wing
section that has dropped has temporarily lost a small percentage of its lift. Consequently, the aircraft
will accelerate and slip down toward the right wing, which produces a sideslip angle (β). This is
equivalent to a wing approaching from the right of the aircraft; the sideslip angle is positive. In
response, a laterally statically stable aircraft must produce a negative rolling moment to return to the
original wing-level situation. This is technically translated into a negative dihedral effect ( < 0). The
role of the wing dihedral angle is to induce a positive increase in angle of attack (∆ ). This function of
the wing dihedral angle is carried out by producing a normal velocity ( = ):
∆ ≈
Γ
≈
Γ
≈ Γ

Figure 37 Typical values of dihedral angle for various wing configurations
From lateral stability ,technical and historical data. The dihedral angle for our low wing
configuration
=
7.9 High-Lift Device
7.9.1 The Functions of a High-Lift Device
One of the design goals in wing design is to maximize the capability of the wing in the generation
of the lift. This design objective is technically shown as the maximum lift coefficient ( ). In a
trimmed cruising flight, the lift is equal to the weight. When the aircraft generates its maximum lift
coefficient, the airspeed is referred to as stall speed:
= =
1
2
=
7.9.1.1 Two design objectives among the list of objectives are:
 maximizing the payload weigh.
 minimizing the stall speed (Vs).
The primary applications of HLDs are during take-off and landing operations. Since the airspeed is
very low compared with the cruising speed, the wing must produce a bigger lift coefficient. The aircraft
speed during take-off and landing is slightly greater than the stall speed. Airworthiness standards
specify the relationship between take-off speed and landing speed with stall speed. As a general rule,
we have:
=
 LEHLD tends to improve the boundary layer energy of the wing. Some type of HLD has been
used on almost every aircraft designed
 At the airfoil level, a HLD deflection tends to cause the following six changes in the airfoil features:

Figure 38 Example of pressure distribution with the application of a high-lift device
Figure 39 Typical effects of a high-lift device on wing airfoil section features
Figure 40 Maximum lift coefficient for several aircraft
7.9.2 High-Lift Device Classification
Two main groups of HLDs are:
 leading edge high-lift device (LEHLD)
 trailing edge high-lift device (TEHLD or flap)

Figure 41 Various types of high-lift device: (a) Trailing edge high-lift device; (b) Leading edge
high-lift device
7.9.3 Design Technique
In designing the HLD for a wing, the following items must be determined:
 HLD location along the span.
 The type of HLD (among the list in Figure 41)
 HLD chord (Cf).
 HLD span (bf).
 HLD maximum deflection (down) (δfmax)
Figure 42 Lift coefficient increment by various types of high-lift device (when deflected 60 deg)

Figure 43. High-lift device parameters: (a) Top view of the right wing; (b) Side view of the inboard
wing (flap deflected)
7.10 Aileron
An aileron is very similar to a trailing edge plain flap except it is deflected both up and down. An
aileron is located at the outboard portion of the left and right sections of a wing. Unlike a flap, ailerons
are deflected differentially, left up and right down or left down and right up. Lateral control is applied
on an aircraft through the differential motions of ailerons.
Figure 44 Typical location of the aileron on the wing

8. Wing configuration
From the data from section 6 of wing design we can summarize the wing configurations as follow
8.1 Wing Configuration Parameters:
Table 9. Wing Configuration Parameters 1
Wing parameters values
number of wings Monoplane
vertical position relative to the fuselage
low wing
cross-section (airfoil )
NACA 1408
aspect ratio ( ) 6.1
taper ratio ( ) . 5
tip chord ( ) 6
root chord ( ) 12
mean aerodynamic chord ̅ 9
Span ( ) 54.4
Twist angle ( ) 0
sweep angle (Λ) 6.34
dihedral angle (Γ) 5
incidence ( ) 0.9

Table 10. Wing Configuration Parameters 2
high-lifting devices and wing
geometry shape
Slats and flaps
Ailerons &flaps
straight taper wing

9. AIRPLANE CG ESTIMATION
9.1 Stability and Balance Control
Balance control refers to the location of the CG of an aircraft. The CG is the point at which the total
weight of the aircraft is assumed to be concentrated, and the CG must be located within specific limits
for safe flight. Both lateral and longitudinal balance are important, but the prime concern is
longitudinal balance; that is, the location of the CG along the longitudinal.
An airplane is designed to have stability that allows it to be trimmed so it will maintain straight and
level flight with hands off the controls. Longitudinal stability is maintained by ensuring the CG is slightly
ahead of the center of lift. This produces a fixed nose-down force independent of
The airspeed. This is balanced by a variable nose-up force, which is produced by a downward
aerodynamic force on the horizontal tail surfaces that varies directly with the airspeed.
Figure 45. Longitudinal forces acting on an airplane in flight.
If a rising air current should cause the nose to pitch up, the airplane will slow down and the
downward force on the tail will decrease. The weight concentrated at the CG will pull the nose back
down. If the nose should drop in flight, the airspeed will increase and the increased downward tail
load will bring the nose back up to level flight. As long as the CG is maintained within the allowable
limits for its weight, the airplane will have adequate longitudinal stability and control. If the CG is too
far aft, it will be too near the center of lift and the airplane will be unstable, and difficult to recover
from a stall. If the unstable airplane should ever enter a spin, the spin could become flat and recovery
would be difficult or impossible.
Figure 46. CG is too Far aft at Low Stall Airspeed.

If the CG is too far forward, the downward tail load will have to be increased to maintain level
flight. This increased tail load has the same effect as carrying additional weight.
A more serious problem caused by the CG being too far forward is the lack of sufficient elevator
authority. At slow takeoff speeds, the elevator might not produce enough nose-up force to rotate and
on landing there may not be enough elevator force to flare the airplane. Both takeoff and landing runs
will be lengthened if the CG is too far forward.
Figure 47. CG is too Far Forward
9.2 Estimating CG Position
Figure 48. The MAC is the chord drawn through the geographic center of the plan area of the wing.
Twin Engine
Aircraft.
Fuel Weight
(%)
Crew
(%)
Engine
(%)
Structure
(%)
Equipment
(%)
14 1 24 31 3
Table 11. Average group weight breakdown

Structure Wing
(%)
Fuselage
(%)
Tail
(%)
Landing gear
(%)
31 14 11 2 4
Table 12. Structural Weight Breakdown
Item Weight (lb.) Arm (ft.) Moment (lb. ft.)
Wing 323.764 12.42 4021.1489
Tail 46.252 40.76 1885.2315
Fuselage 254.38 19.8 5036.724
Landing gear 92.504 6 555.024
Engine 1790.4 10 1611.6
Fuel 1044.4 13.23 1381.412
Human while seating 1360 5 5440
payload 223.8 4 895.2
Table 13. CG Estimation
=
ℎ
∗
1
=
12895.2656
5135.51
∗
1
9
= 0.275
10.Wing Body Balance
During the mission the aircraft weight decreases as the fuel is consumed, resulting in a shift in the
center of gravity position, this shift must have a limit, to maintain the stability of the aircraft, this limit
is at the neutral point at which the aircraft is neutrally stable.
The neutral point definition: the slope of the moment coefficient about CG become zero
Calculation of .
− =
−
From following figures we can get and
= − .
= .

Figure 49. Cmcg with α
Figure 50. C_L with α
Also, From CG estimation
= .
= .
Now we can calculate static margin = − = 0.18007
 The range of static margins for general aircrafts is between 0.1 and 0.4, so our aircrafts static
margin of 0.18007 is a very good value as it balances between stability and controllability of
the aircraft.

11.Trimming
Airplane can be trimmed in two critical Condition
1. At stall, with , flaps down and CG on the most forward limit
2. At low and aft CG case
Check equation must be applied in each case
= (ℎ − ℎ ) + − = 0
Trimming cases will be checked in longitudinal stability part after design the tail and calculated the
CG of the airplane
12.Tail Configuration
12.1 Basics Tail Configuration
This section presents the design requirements and information for tail Configuration Horizontal tail
must satisfy the following requirements:
 Longitudinal Trim
 Directional Trim
 Lateral Trim
 Longitudinal Stability
 Directional Stability
 Lateral Stability
 Manufacturing and Controlling
 Handling quality
 Airworthiness
 Spin Recovery
 Cost
 Competitivity
 Size limits (e.g. limited height)
The following tail configurations are available that are satisfy the design requirements:
 Aft tail and one aft vertical tail
 Aft tail and twin aft vertical tail
 Delta wing and one aft vertical tail
 Canard and aft vertical tail
 Tailless

Figure 51. Tail Configuration
12.2 Tail Selection
For our design we will choose an Aft tail and one aft vertical tail
12.2.1 Aft Tail Configuration
Aft tail has several configurations that all can satisfy design configuration the following list shows
different types of Aft tail
 Conventional
 T-Shape
 H-Shape
 V-Tail
 Y-Tail
 Twin vertical tail
 Twin T
 U-Tail
 Triple Tail
12.2.2 Aft Tail Selection
For our design we will choose Conventional tail
12.2.2.1 Advantage of Conventional Tail
 Almost all textbooks examine its features
 Has light weight
 Efficient and perform at regular flight conditions
 Provide appropriate stability and control
 A large vertical tail plane height is more appropriate for conventional tail than T-Tail

12.2.2.2 Disadvantage of Conventional Tail
 Spin characteristic can be bad cause of the blanketing of vertical tail
 The downwash of the wing is relativity large in the area of horizontal tail plane
 Rare engine can’t be teamed with Conventional tails
Figure 52. Conventional Tail
12.2.3 Optimum Tail Arm
Tail Arm ( ) is the distance between tail aerodynamic center and the center of gravity. Tail Arm
service as the arm of pitching moment about CG to maintain the longitudinal trim. As tail arm increased
tail area is decreased, while as tail arm decreased the tail area increased. Short tail arm as in fighter’s
aircraft and long tail arm as in most transport aircraft.
Two very significant aircraft general design requirements are aircraft low weight and low drag. As
the horizontal tail arm is increased, the fuselage wetted area is increased, but horizontal tail wetted area
is decreased. Also, as the horizontal tail arm is decreased, the fuselage wetted area is decreased, but
horizontal tail wetted area is increased. Hence, we want the optimum tail arm to minimize drag; which
means to minimize the total wetted area of the aft portion of the aircraft. The following approach is
based on the fact that the aircraft zero-lift drag is essentially a function of the aircraft wetted area.
Figure 53. Top View of aft Portion of Aircraft

= +
=
1
2
+
2 ̅
= 0
By solving this equation, we can get optimum tail arm as following
=
4 ̅
Where
Optimum Tail Arm
̅ Main Aerodynamic chord
Wing Area
Tail Volume Coefficient
Fuselage diameter
Correction Factor and various between 1 to 1.4 depending on the aircraft configuration.
The = 1 is used when the aft portion of the fuselage has a conical shape. As the shape of the aft
portion of the fuselage goes further away from a conical shape, the factor is increased up to 1.4.
Figure 54. Optimum Tail Arm

13.Horizontal Tail Design
After the selection of tail configuration, the following procedure will used to design the horizontal
tail:
No
Yes
Figure 55. Horizontal Tail Design Procedure
Select Horizontal Tail Location
Select Horizontal Tail Volume Coefficient
Determine Optimum Tail Arm
Determine Planform Area
Determine Airfoil Selection
Determine sweep and Dihedral angle
Determine Aspect and Taper Ratio
Check Tail
stall
Analysis longitudinal Stability and Optimize
Calculate Setting angle
Calculate , , ,

13.1 Horizontal Tail Location
As we said in Configuration of tail part, we will choose an Aft Horizontal Tail
13.2 Select Horizontal Tail Volume Coefficient
The following table represent a reasonable Assumptions for Vertical tail coefficient for different
aircrafts types:
NO Aircraft Horizontal Tail Volume Coefficient
1 Glider and motor Glider 0.6
2 Home-built 0.5
3 GA-single prop driven engine 0.7
4 GA-twin prop-driven engine 0.8
5 GA with canard 0.6
6 Agricultural 0.5
7 Twin turboprop 0.9
8 Jet trainer 0.7
9 Fighter aircraft 0.4
10 Fighter (with canard) 0.1
11 Bomber/military transport 1
12 Jet Transport 1.1
Table 14 Typical values for horizontal and vertical tail volume coefficients
For our Airplane we choose Horizontal Tail Coefficient for GA-twin propeller-driven engine So,
= 0.8
13.3 Determine Optimum Tail Arm
The horizontal tail arm is less than three time the wing MAC (3 ̅ ), the aircraft is said to be short-
coupled. An aircraft with such tail configuration possesses the longitudinal trim penalty (e.g. fighters) .
We will calculate Optimum Tail Arm as we discuss in tail configuration part using the following
equation an Assuming that = 1 :
=
4 ̅
= 1 ∗
4 ∗ 9 ∗ 486 ∗ 0.8
∗ (1.43 ∗ 3.28)
= 30.82 = 9.39
Check
As we said should be more than MAC (3 ̅ )
> 3(9)
> 27

13.4 Determine Planform Area
Many areas of the aircraft design process rely on accurate lift estimation. Our selection will be
rectangular planform.
Figure 56. Rectangular Planform
13.5 Horizontal Airfoil Selection
Horizontal tail plane is a lifting surface (similar to the wing) and requires a special airfoil section.
the tail plane airfoil lift curve slope ( ) must be as large as possible along with a considerably wide
usable angles of attack. Since the aircraft center of gravity moves during the cruising flight, the airfoil
section must be able to create sometimes a positive lift (+ ) and sometimes a negative lift (− )
This requirement necessitates the tailplane to behave similar in both positive and negative angles of
attack. For this reason, a symmetric airfoil section is suitable candidate for horizontal tail.
A symmetric airfoil is that the second digit in a 4-digit and the third digit in a 5-digit and 6-series
NACA airfoil sections is zero. This denotes that the airfoil design lift coefficient and zero-lift angle of
attack are both zero.
In addition, another tail requirement is that horizontal tail must be clean of compressibility effect.
In order the tail to be out of the compressibility effect, the tail lift coefficient is determined to be less
than the wing lift coefficient. To ensure this requirement, the flow Mach number at the tail must be
less than the flow Mach number at the wing. This objective will be realized by selecting a horizontal
tail airfoil section to be thinner (say about 2 percent of MAC) than the wing airfoil section. For our
selection the wing airfoil section is NACA 1408 (( / ) = 8%), the horizontal tail airfoil section can
be chosen to be NACA 0006 to satisfy the above requirements.
13.5.1 NACA 0006 Specifications
Figure 57. NACA 0006 Specifications

Figure 58. Lift Coefficient verses Drag Coefficient and Angle of Attack
Figure 59. Cl/Cd versus Alpha and Cd versus Alpha
Figure 60. Moment Coefficient versus Alpha

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