This presentation explains the different design phases of aircraft and some detailed descriptions of each phase. Then, it explains the preliminary design procedure of the aircraft which consists of weight estimation, wing design, fuselage design, drag polar, maximum speed, thrust requirement and selection of engine, landing gear selection, tail surface design, and different performance calculations.

1. Seminar-II
ROSHANSAH
USN :- 17AE60R01
M.Tech (1st Year)
Dept.of Aerospace Engg.
IndianInstituteof Technology
Kharagpur(IIT KGP)
PRELIMINARY DESIGN OF A 100 SEATER
AIRCRAFT

2. PHASES OF AIRPLANE DESIGN
Conceptual Design
-Variety of possible aircraft configuration arrangement, size, weight
& performance that meet the design specifications are sketched.
-New ideas & problem emerge as design is investigated.
Preliminary Design
-Initial stages of design
-Specialist in area of structure, landing gear ,control system will
design and analyze their portion of aircraft.
-Testing in area of aerodynamics, propulsion & stability and control
is initiated.
-One of key activity occur in it is lofting.
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3. Detail Design
-Begins with actual pieces to be fabricated are designed.
-Like; Precise design of each individual rib, spar and section of
skin takes place.
-Involves in Production design.
- specialist determine how the aircraft will be fabricated
starting with the smallest & simple sub-assemblies
and building up to the final assembly process.
- Testing effort intensifies and actual structure of aircraft is
fabricated and tested.
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4. Preliminary Design
• Weight estimation
• Wing design
• Fuselage design
• Drag polar
• Maximum speed
• Thrust requirement and selection of engine
• Landing gear
• Tail surfaces design
• Airplane performance calculation
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5. Payload of 100 passenger
Range of 2000km
Cruise altitude = 12 km
Cruise Mach no = 0.7
Cruise speed = 743 km/hr =206.53 m/s
Landing distance = 1500 m
DESIGN REQUIREMENTS
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6. Crew members = 5
Weight of crew members (𝑊𝑐)= 5*110 = 550 kg
Weight of passenger(𝑊𝑝) = 100*110 = 11000 kg
Empty weight 𝑊𝑒
Weight of fuel 𝑊𝑓
Gross weight(𝑊0) = 𝑊𝑐+𝑊𝑝+𝑊𝑒+𝑊𝑓
= 𝑊𝑐 + 𝑊𝑝 +
𝑊 𝑒
𝑊0
𝑊0 +
𝑊 𝑓
𝑊0
𝑊0
Weight Estimation
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7. Weight Estimation
𝑊0 =
𝑊𝑐+𝑊𝑝
1−
𝑊𝑒
𝑊0
−
𝑊𝑓
𝑊0
Wf = 𝑊0-W5
Fuel fraction estimate
𝑊 𝑓
𝑊0
= 1.06 1 −
𝑊5
𝑊0
𝑊5
𝑊0
=
𝑊1
𝑊0
∗
𝑊2
𝑊1
∗
𝑊3
𝑊2
∗
𝑊4
𝑊3
∗
𝑊5
𝑊4
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8. Historical mission segment weight fraction
Takeoff
𝑊1
𝑊0
) 0.97
Climb (
𝑊2
𝑊1
) 0.985
Landing (
𝑊5
𝑊4
) 0.995
Cruise
𝑊3
𝑊2
= 𝑒𝑥𝑝
−𝑅 𝑐𝑡
𝑉𝑐𝑟
𝐿
𝐷 𝑐𝑟
= 0.90147
𝑉𝑐𝑟𝑢𝑖𝑠𝑒 = 0.7 ∗ 295.2 = 206.64 𝑚/𝑠
𝐿
𝐷 𝑚𝑎𝑥
= 15;
𝐿
𝐷 𝑐𝑟
= 0.866
𝐿
𝐷 𝑚𝑎𝑥
= 13 ; 𝑐 𝑡 = 0.5/hr
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9. No. Aircraft Type 𝐿
𝐷
𝑚𝑎𝑥
1. Sailplane (glider) 20-35
2. Jet transport 12-20
3. Subsonicmilitary 8-11
4. Supersonicfighter 5-8
5. Helicopter 2-4
6. Homebuilt 6-14
7. Ultralight 8-15
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10. Loiter

𝑊4
𝑊3
= 𝑒𝑥𝑝
−𝐸 𝑐𝑡
𝐿
𝐷 𝑚𝑎𝑥
= 0.9231
𝑐𝑡 = 0.4/hr and E =30 min.

𝑊5
𝑊0
=
𝑊1
𝑊0
∗
𝑊2
𝑊1
∗
𝑊3
𝑊2
∗
𝑊4
𝑊3
∗
𝑊5
𝑊4
= 0.791

𝑊 𝑓
𝑊0
= 0.22143
Weight Estimation
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11. 𝑊𝑒
𝑊0
= 1.02 𝑊0
−0.06
𝑊0 = 47350 kg
𝑊𝑒
𝑊0
= 0.53466
Weight Estimation
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12. 𝑆 𝐿 = 1500 m
𝑉𝑎 =
𝑆 𝐿
0.3
0.5 = 128.0788 knot
where 𝑆 𝐿 is in feet and 𝑉𝑎 is in knots
𝑉𝑎 = 65.88 m/s
𝑉𝑠𝑡𝑎𝑙𝑙 = 𝑉𝑎/1.3 = 50.68 m/s
 𝐶𝐿𝑚𝑎𝑥= 2.5(using fowler flap)

𝑊
𝑆 𝐿𝑎𝑛𝑑
= 3933 N/m2

𝑊
𝑆 𝑇𝑂
=
1
0.85
*
𝑊
𝑆 𝐿𝑎𝑛𝑑
= 4627 N/m2
Final choice of wing loading is =4346 N/m2 (+10 % variation in
𝑉𝑠𝑡𝑎𝑙𝑙
Choice Of Optimum Wing Loading
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13. Wing Design
Airfoil selection
-NASA SC(2) series airfoil to increase drag
divergence Mach no.
𝐶𝐿𝑐𝑟𝑢𝑖𝑠𝑒 =
𝑊
𝑆
𝑞 𝑐𝑟𝑢𝑖𝑠𝑒
= 0.564
Airfoil thickness ratio
-At root - 15.2 % & At tip - 10.3%
-Average t/c = 14 %
Wing Sweep(quarter chord Sweep) = 25º
Dihedral (Γ) = 5°
Wing twist = 3°
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14. Aspect Ratio(AR) = 9
 S =106.88 m2
Wing span = b = 𝐴𝑅. 𝑆
= 31.015 m
Taper Ratio(λ) = 0.24
Root chord(𝐶𝑟) =
2𝑆
𝑏 1+λ
= 5.558 m
Tip chord(𝐶𝑡) =𝑐 𝑟. 𝜆 = 1.3339 m
Mean Aerodynamic Chord((𝒄)
𝒄 =
2
3
1+𝜆+𝜆2
1+𝜆
. 𝑐 𝑟 = 3.877 m
Wing Design
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15. Fuselage comprises of nose and cockpit, passenger
compartment, and rear fuselage.
Fuselage Length(𝐿 𝑓) = 𝑎𝑊0
𝑏
; where a = 0.6 & b= 0.35
Fuselage Length(𝐿 𝑓) = 25.97m
Nose length(𝐿 𝑛𝑜𝑠𝑒)
-𝐿 𝑛𝑜𝑠𝑒 = 0.03𝐿 𝑓 = 0.7791m
Cockpit length = 2.5 m
Fuselage
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16. Fuselage
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17. Fuselage
Cabin parameters are chosen based on standards for similar airplanes
Parameter EconomyClass BusinessClass
Seat pitch(m) 0.8128 0.9652
Seat width(m) 0.508 0.5588
Aisle width(m) 0.5588 0.6096
Seats in a row 6 4
Numberofaisles 1 1
Max.height (m) 2.1 2.1
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18. Fuselage
Class No. of seats No. of rows Seat Pitch(m) Cabin
Length(m)
Economy 84 14 0.8128 11.3792
Business 16 4 0.9652 3.86
Cabin length = 11.3792+3.86 = 15.24 m
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19. Fuselage
Rear fuselage
Lrear= 0.25 Lf = 0.25 *25.97 = 6.501 m
Total fuselage length = 𝐿 𝑛𝑜𝑠𝑒+ 𝐿 𝑐𝑜𝑐𝑘𝑝𝑖𝑡 + 𝐿 𝑐𝑎𝑏𝑖𝑛 + 𝐿 𝑟𝑒𝑎𝑟
= 0.7791+2.5+ 15.24 + 6.501
= 25.02m
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20. Internal diameter of the cabin is calculated as:
- 𝐷𝑖 = Aisle width *1+(6*seat width) = 3.6068m =142inch
Structural thickness(in inches) is approximated by:
- t = 0.02*𝐷𝑖+1 = 3.84 inch = 0.097536m
The external diameter of the fuselage is obtained as:
- 𝐷0 =3.8018m= 149.68 inch
Fuselage
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21. Max Speed of aircraft
𝑀 𝑚𝑎𝑥 is taken as –
𝑀 𝑚𝑎𝑥 = 𝑀𝑐𝑟𝑢𝑖𝑠𝑒 + 0.04
𝑀 𝑚𝑎𝑥 = 0.74
𝑉𝑚 𝑎𝑥 = 0.74*295.2 = 218.448 m/s
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22. Drag Polar
Assuming -
e(for wing)=0.85 ; Dihedral angle = 5
e(airplane) = 0.74 ; Aspect ratio(AR)=9
k=
1
Π𝑒𝐴𝑅
= 0.047794 ;
𝐿
𝐷 𝑚𝑎𝑥
=15
𝐶 𝐷0 = .0233
So the drag polar equation is 𝐶 𝐷 = 𝐶 𝐷0+kCL
2
𝐶 𝐷 = .0233+.047794CL
2
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23. Thrust Requirement
Weight at mid cruise

𝑇
𝑊
= 0.5ρ𝑉2
𝑐𝑟𝑢𝑖𝑠𝑒
𝑆
𝑊
𝐶 𝐷0+
2𝐾𝑊
𝑉2
𝑐𝑟𝑢𝑖𝑠𝑒 𝑆
𝑊𝑚𝑐 = 𝑊2-0.5(𝑊2-𝑊3) = 45240.55- 0.5(45240.55-40783)
= 43011.77 kg

𝑊 𝑚𝑐
𝑆
= 3947.84. N/m2
Total thrust produced (Tcruise ) =62.96891KN
Tcruise each engine = 31.484 KN
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24. Thrust Requirement

𝑇 𝑆𝐿
𝑊
σ = 0.5ρ𝑉2
𝑚𝑎𝑥
𝑆
𝑊
𝐶 𝐷0+
2𝐾𝑊
ρ 𝑉2
max 𝑆
𝑇𝑆𝐿= 109.80KN
Thrust require maximum is 109.80KN
Based on the above calculations
ROLLS-ROYCE BR715/710 Turbofan
is selected.
Specification:-
• Thrust range = 14000lb to 21500 lb.
• 20% improved fuel burn
• Fan diameter 58 inch and a two-stage
booster driven by a three-stage LP turbine.
• single low emissions annular combustor with 20 burners.
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25. Landing Gear
 A conventional Tricycle landing gear is chosen based on the
trend followed by similar airplanes.
Parameter Value
Wheel base (in m) 13.2
Track length (in m) 5.8
Turning radius(in m) 19.3
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26. Tail Surfaces
Parameters HorizontalTail Vertical Tail
Area Ratio 0.31 0.21
AspectRatio 5 1.7
TaperRatio 0.26 0.31
Sh=0.31*106.88=33.1338 m2 bh = 12.87m
Sv=0.21*106.88=22.44 m2
bv = 6.1764 m
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27. Tail Arm
The values of the tail arm are chosen based on the data
collection as:
Lh = 45 % of Lf &
 Lv = 42% of Lf
Lh = 0.45 x 25.02=11.259 m
Lv = 0.42 x 25.02= 10.5084m
Vh =
𝑆ℎ 𝐿ℎ
𝑆𝐶
= 0.9
Vv =
𝑆 𝑉 𝐿 𝑉
𝑆𝐶
= .0.569
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28. Three view drawing of Boeing 737-200
Three view Drawing
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29. [1] D.P. Raymer “Aircraft design: a conceptual approach”AIAAeducational Series, Fourth
edition (2006).
[2] Harris C.H. “NASA supercritical airfoils – A matrix of family- related airfoils” NASA
TP 2969 , March 1990
[3] Jenkinson L. R., Simpkin P. and Rhodes D. “Civil jet aircraft design”Arnold,(1999).
[4]Introduction to Flight by John D.Anderson,2nd edition.
[5]Theory of wing sections by Ira. H. Abbott, Dover edition.
REFERENCES :-
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