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AIRCRAFT DESIGN PROJECT- I 7-15 SEATER BUSINESS JET

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ABSTRACT In this project we have designed a 7-15 seater Business jet. We

have taken the sufficient steps to make sure that the aircraft what we

are designing is in an optimum range. The aircraft parameters like

cruise velocity, cruise altitude, wing loading etc and weight estimation,

airfoil selection, wing selection, landing gear selection have been made

with extreme care. The adequate details have been collected to make

our calculation easier and to make design more precision. The details

have been collected from various sources which are given in the

bibliography.

Serial no TopicPage no

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1 Introduction 5

2 Classification of aircraft 9

3 The design 11

4 Comparative data sheet 12

5 Weight design 31

6 Wing design 35

7 Airfoil selection 37

8 Landing gear design 42

9 Performance calculations 53

10 3-D view Diagram 56

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INTRODUCTIONPurpose and scope of airplane design

An airplane is designed to meet the functional, operational and safety requirements set by or acceptable to the ultimate user. The actual process of design is a complex and long drawn out engineering task involving:

Selection of airplane type and shape Determination of geometric parameters Selection of power plant Structural design and analysis of various components and Determination of airplane flight and operational characteristics.

Over the year of this century, aircraft have evolved in many directions and the design of any modern plane is a joint project for a large body of competent engineers and technicians, headed by a chief designer. Different groups in the project specialize in the design of different components of the airplane, such as the wing, fuselage etc.

A new experimental plane has to meet higher performance requirements than similar planes already in service. Hence design laboratories involved in experimental and research work are indispensable adjuncts to a design office. These laboratories as well as allied specialized design offices and research institutions are concerned in helping the designer to obtain the best possible solutions for all problems pertaining to airplane design and construction and in the development of suitable components and equipment.

Airplane design procedure is basically a method of trial and error for the design of component units and their harmonization into a complete aircraft system. Thus each trial aims at a closer approach to the final goal and is based on a more profound study of the various problems involved. The three phases of aircraft design are

Conceptual design Preliminary design detail

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Phase of aircraft design

Conceptual design

Aircraft design can be broken into three major phases, as depicted in figure. Conceptual design is the primary focus of this book. It is in conceptual design that the basic questions of configuration arrangement, size and weight, and performance are answered.

The first question is “can an affordable aircraft be built that meets the requirements?” if not, the customer may wish to relax the requirements.

Conceptual design is a very fluid process. New ideas and problems emerge as a design is investigated in increasing detail. Each time the latest design is analyzed and sized, it must be redrawn to reflect the new gross weight, fuel weight, wing size, and other changes. Early wind tunnel test often revel problems requiring some changes to the configuration.

REQUIREMENTS

CONCEPTUAL DESIGN will it work?

What does it look like?

What requirements drive the design?

What trade-offs should be consider?

What should it weigh and cost?

PRELIMINARY DESIGN freeze the configuration

Develop lofting

Develop test and analytical base

Design major items

Develop actual cost estimation

DETAIL DESIGN design the actual piece to be built

Design the tooling and fabrication process

Test major items structure, landing gear

Finalize weight and performance estimate

FABRICATION

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Preliminary design

Preliminary design can be said to begin when the major changes are over. The big questions such as whether to use a canard or an aft tail have been resolved. The configuration arrangement can be expected to remain about as shown on current drawing, although minor revisions may occur. At some point late in preliminary design, even minor changes are stopped when a decision is made to freeze the configuration.

During preliminary design the specialists in area such as structure landing gear and control systems will design and analyze their portion of the aircraft. Testing is initiated in areas such as aerodynamics, propulsion, structures, and control. A mockup may be constructed at this point.

A key activity during preliminary design is “lofting”. Lifting is the mathematical modeling of the outside skin of the aircraft with sufficient accuracy to insure proper fit between its different parts, even if they are designed by different designers and possibly fabricated in different location. Lofting originated in shipyards and was originally done with long flexible rulers called” splines”. This work was done in a loft over the shipyard; hence the name.

The ultimate objective during preliminary design is to ready the company for the detail design stage, also called full-scale development. Thus, the end of preliminary design usually involves a full scale development proposal. In today’s environment, this can result in a situation jokingly referred to as “you-bet-your-company”. The possible loss on an overrun contrast o from lack of sales can exceed the net worth of the company! Preliminary design must establish confidence that the airplane can be built in time and at the estimated cost.

Detailed design

Assuming a favorable decision for entering full scale development, the detail deign phase begins in which the actual pieces to be fabricated are designed. For example, during conceptual and preliminary design the wing box will be designed and analyzed as a whole. During detail design, that whole will be broken down in to individual ribs, spars and skins, each of which must be separately designed and analyzed.

Another important part of detailed is called production design. Specialist determine how the airplane will be fabricated, starting with the smallest and simplest subassemblies and building up to the final assembly process. Production designers frequently wish to modify the design for ease of manufacture; that can have a major impact on performance or weight. Compromises are inevitable, but the design must still meet the original requirements.

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It is interesting to note that in the Soviet Union, the production design is done by a completely different design bureau than the conceptual and preliminary design, resulting in superior producibility at some expense in performance and weight.

During detail design, the testing effort intensifies. Actual structure of the aircraft is fabricated and tested. Control laws for the flight control system arte tested on an “iron-bird” simulator, a detailed working model of the actuator and flight control surfaces. Flight simulator are developed and flown by both company and customer test pilot.

Detail design ends with fabrication of the aircraft. Frequently the fabrication Begins on part of the aircraft before the entire detail-design effort is completed. Hopefully, changes to already- fabricated pieces can be avoided. The further along a design progresses, the more people are involved. In fact, most of the engineers who go to work for a major aerospace company will work in preliminary on detail design.

Classification of airplanes design

Functional classification:

The airplane today is used for a multitude of activities in civil and military fields. Civil applications include cargo transport, passenger travel, mail distribution, and specialized uses like agricultural, ambulance and executive flying. The main types of military airplane at the present time are fighters and bombers. Each of these types may be further divided into various groups, such as strategic fighters, interceptors, escort fighters, tactical bombers and strategic bombers. There are also special aircraft, such as ground attack planes and photo-re-connaisance planes. Sometimes more than one function may be combines so that we have multi-purpose airplanes like fighter-bombers. In addition to these, we have airplanes for training and sport.

Classification by power plants:

Types of engines used for power plant:

Piston engines (krishak, Dakota, super constellation) Turbo-prop engines ( viscount,friendship,An-102) Turbo-fan engines (HJT – 16, Boeing series, MIG-21) Ramjet engines Rockets (liquid and solid propellants) (X-15A)

Location of power plant:

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Engine ( with propeller) located in fuselage nose (single engine) (HT-2,Yak-9,A-109)

Pusher engine located in the rear fuselage (Bede XBD-2) Jet engines submerged in the wing

1. At the root(DH Comet, Tu-104,Tu-16)2. Along the span (Canberra, U-2, YF-12A)

Jet engines in nacelles suspended under the wing (pod mountings) (Boeing 707,DC-8,Convair 880)

Jet engines located on the rear fuselage (Trident, VC – 10 ,i1-62) Jet engines located within the rear fuselage (Hf – 24, lighting,MIG-

19)

Classification by configuration:

Airplanes are also classified in accordance with their shape and structural layout, which in turn contribute to their aerodynamic, tactical and operational characteristics. Classification by configuration is made according to:

Shape and position of the wing Type of fuselage Location of horizontal tail surfaces

Shape and position of the wing:

Braved biplane(D.H. Tiger moth) Braced sesquiplane (An-2) Semi-cantilever parasol monoplane (baby ace) Cantilever low wing monoplane (DC-3,HJT-16,I1-18,DH Comet) Cantilever mid wing monoplane (Hunter, Canberra) Cantilever high wing monoplane (An-22,Brequet 941 Fokker

Friendship) Straight wing monoplane (F-104 A) Swept wing monoplane (HF-24, MIG-21, Lighting) Delta monoplane with small aspect ratio (Avro-707, B-58 Hustler, Avro

Vulcan)

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Type of fuselage

Conventional single fuselage design ( HT-2,Boeing 707 Twin- fuselage design Pod and boom construction (Packet, Vampire)

Types of landing gear:

Retractable landing gear (DC-9,Tu-114,SAAB-35) Non- retractable landing gear (pushpak, An-14, Fuji KM-2) Tail wheel landing gear (HT-2,Dakota,Cessana J85 C) Nose wheel landing gear (Avro-748, Tu-134,F-5A) Bicycle landing gear (Yak-25,HS-P,112)

THE DESIGN

Design is a process of usage of creativity with the knowledge of science where we try to get the most of the best things available and to overcome the pitfalls the previous design has. It is an iterative process to idealism toward with everyone is marching still.

Design of any system is of successful application of fundamentals of physics. Thus the airplane design incorporates the fundamentals of aerodynamics, structures, performance and stability & control and basic physics. These are based on certain degree of judgment and experience. Every designer has the same technical details but each design prevails it own individuality and the mode of the designer.

Here the preliminary design has been done of an executive Transport Aircraft. The basic requirements are the safe, comfortable and economic transport mode with reasonable time period of flight. Here comfort and safety are given primary importance.

Here the most possible considerations have been taken. And the flight parameters and limitations are studied.

The modern day calls for the need of latest aircrafts for the use of passenger transport which aims mainly at improving the aerodynamic characteristics as well as the passenger comfort. This design project also looks at the above aspects in a lot more closer way. Also the design project has been classified into different stages in our design will be as follows.

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Collection of comparative data Selection of aircraft parameters Preliminary and second weight estimations Selection of power plant Airfoil selection, flaps, t/c, sweep, etc Layout of L/G, load, tires selection 3-view diagram Balance diagram

COMPARATIVE DATA SHEET

In the designers perspective it is necessary to compare the existing airplanes that are of same type as that of our desired airplane. Their important parameters, positive aspects to be considered and pitfalls to be overcome are taken into consideration.

The datas has been collected for various sites in the internet for 7-15 seater BUSINESS JET design

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Several parameters are compared for over 25 aircrafts and different critical parameters were plotted on graph. They are

Cruise Velocity Range Wing area Wing loading Aspect ratio Sweep angle Cruise altitude Thrust/weight ratio Maximum take off weight

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Parameters Cessna citation cj3Cessna citation excel

Cessna citation X

Maximum take off Weight(kg) 6,291 13,608 16,375

Design pay load(kg) 371 n/a 6,486

Design fuel load(kg) 2,136 3057 n/a

Length(m) 15.29 19.37 22

Area(m2) 27 34.35 50

Span(m) 16.26 19.24 19.4

Aspect ratio 10 8.4 7.527

Quarter chord Sweep(degrees) 0 0 37

Max wing loading(kg\m2) 230 264.1 334.45

Thrust\weight ratio 0.331 0.38 0.33

Length\span ratio0.94 1.006 1.13

Cruise velocity(km\hr) 773 848 1004

Design range(km) 3519 4,821 6,020

Altitude(m) 13,716 15,517 15,545

Capacity 9 10 8 to 12

Parameters Bombardier global express

Britten-Norman Islander

Cessna 421

Maximum take offWeight (kg) 44,500 2,994 3,103

Empty weight(kg) 22,600 1,627 2,132

Design pay load(kg) 805 n/a n/a

Design fuel load(kg)

19663 n/a n/a

Length(m) 30.3 10.86 10.29

Area(m2) 94.9 30.2 18.6

Span(m) 28.65 14.94 12.15

Aspect ratio 8.65 7.39 7.937

Quarter chord Sweep(degrees) 35 n/a n/a

Max wing loading(kg\m2) 468 99.14 166.8

Thrust\weight ratio 0.301 0.186 0.158

Length\span ratio 1.057 0.7269 0.847

Cruise velocity(km\hr) 950 n/a 420

Design range(km) 11,390 1,400 2,756

Altitude(m) 15,500 4,024 5,230

Capacity 8 to 19 9 8

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Parameters Gulf stream G100 Gulf stream G200 Hawker 400

Maximum take off Weight(kg)

11181 16079 7393

Empty weight(kg) 6214 8981 4785

Design pay load(kg)

363 1905 289

Design fuel load(kg)

3942 6804 2228

Length(m) 16.94 18.97 14.75

Area(m2) 29.41 34.28 22.43

Span(m) 16.64 17.7 13.25

Aspect ratio 8.8 9.1 7.8

Quarter chord Sweep(degrees)

34 34 20

Max wing loading(kg\m2)

380.3 469.1 329.67

Thrust\weight ratio 0.31 0.346 0.287

Length\span ratio1.01 1.07 1.11

Cruise velocity(km\hr)

850 880 778

Design range(km) 5461 6667 2883

Altitude(m)12497 13715 13243

Capacity9 8 8

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Parameters Hawker 800 Dassault Falcon 50Learjet 25

Maximum take off Weight(kg)

12701 18007 6803

Empty weight(kg) 7029 9603

Design pay load(kg)

1000 1710 n/a

Design fuel load(kg)

4536 7040 n/a

Length(m) 15.6 18.52 14.5

Area(m2) 34.75 46.83 21.53

Span(m) 15.66 18.86 10.84

Aspect ratio 7.1 7.6 5.4577

Quarter chord Sweep(degrees)

20 29 n/a

Max wing loading(kg\m2)

365.5 384.5 315.9

Thrust\weight ratio n/a n/a n/a

Length\span ratio0.99 0.98 1.3376

Cruise velocity(km\hr)

793 850 n/a

Design range(km) 4704 5602 2853

Altitude(m)11887 12497 13720

Capacity 14 9 8

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Parameters Learjet 45Cessna 680 citation sovereign

Cessna 550 citation

Maximum take off Weight(kg)

9298 13067 6713

Empty weight(kg) 5797 7892 3992

Design pay load(kg)

1030 726 370

Design fuel load(kg)

2750 4921 2204

Length(m) 17.81 19.35 14.4

Area(m2) 28.95 49.73 30

Span(m) 14.56 19.25 15.75

Aspect ratio 7.3 7.7 8.4

Quarter chord Sweep(degrees)

13 12 10

Max wing loading(kg\m2)

322.3 283.9 223.8

Thrust\weight ratio 0.383 0.33 0.38

Length\span ratio1.22 1.005 0.9

Cruise velocity(km\hr)

846 850 744

Design range(km) 3885 5056 1744

Altitude(m)15545 13105 8458

Capacity 9 12 10

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Parameters Learjet 35IAI WEST WIND GAF N22 NOMAD

Maximum take off Weight(kg)

8300 9390 3855

Empty weight(kg) 4590 n/a n/a

Design pay load(kg)

n/a n/a n/a

Design fuel load(kg)

n/a n/a n/a

Length(m) 14.83 15.93 12.56

Area(m2) 23.5 28.6 30.1

Span(m) 12.04 13.65 16.52

Aspect ratio 6.16 6.5 9

Quarter chord Sweep(degrees)

n/a n/a n/a

Max wing loading(kg\m2)

353.19 328.322 128.07

Thrust\weight ratio 0.383 0.3 0.192

Length\span ratio1.23 1.16 0.76

Cruise velocity(km\hr)

852 870 311

Design range(km) 4070 3410 1352

Altitude(m)13716 13716 6400

Capacity 8 10 12

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Parameters

RAYTHON BEECHCRAFT KING AIR

Learjet 23Cessna citation mustang

Maximum take off Weight(kg)

6350 6803 3807.97

Empty weight(kg) n/a n/a n/a

Design pay load(kg)

n/a n/a 281

Design fuel load(kg)

n/a n/a 1460

Length(m) 13.36 10.84 13.16

Area(m2) 28.2 21.46 22.3

Span(m) 16.61 13.18 13.16

Aspect ratio 9.78 8.09 7.7

Quarter chord Sweep(degrees)

n/a n/a n/a

Max wing loading(kg\m2)

225.18 317 215.6

Thrust\weight ratio 0.155 n/a n/a

Length\span ratio0.8 0.82246 1

Cruise velocity(km\hr)

583 n/a 630

Design range(km) 3630 1584 2310

Altitude(m)N\A 13715 12500

Capacity15

6 4 to 5

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Parameters North American sabreliner

Piaggio P180 Avanti

Lockheed jet star

Maximum take off Weight(kg)

8060 5262 20227

Empty weight(kg) 5000 3400 11477

Design pay load(kg)

n/a n/a n/a

Design fuel load(kg)

n/a 1271 n/a

Length(m) 13.41 14.4 18.42

Area(m2) 31.78 16 50.4

Span(m) 13.56 14.03 16.6

Aspect ratio 5.7858 12.3 5.46

Quarter chord Sweep(degrees)

n/a 1 n/a

Max wing loading(kg\m2)

253.6 327.4 401.3

Thrust\weight ratio 0.338 0.333 0.339

Length\span ratio0.9889 1.026 1.109

Cruise velocity(km\hr)

n/a 644 823

Design range(km) 4020 2595 4585

Altitude(m)n/a 12500 13135

Capacity 4 to 5 9 8to 10

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Parameters Learjet 60 Learjet 55Dassault falcon 7X

Maximum take off Weight(kg)

10660 9525 28,894

Empty weight(kg) 6641 5832 n/a

Design pay load(kg)

1070 n/a n/a

Design fuel load(kg)

3588 n/a 17735

Length(m) 17.88 16.8 23.19

Area(m2) 24.6 24.57 70.7

Span(m) 13.34 13.35 25.13

Aspect ratio 7.234 7.253 8.93

Quarter chord Sweep(degrees)

n/a n/a 34

Max wing loading(kg\m2)

433 387.66 435

Thrust\weight ratio 0.392 0.345 0.33

Length\span ratio1.34 1.258 0.923

Cruise velocity(km\hr)

782 744 685

Design range(km) 4628 4156 10,500

Altitude(m)15545 15545 15,500

Capacity 8 to 10 7 19

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s.no Design characteristics Values

1 Cruise velocity 800 Km/hr

2 Design range 3825 Km

3 Cruise altitude 12750ft

4 Aspect ratio 7.5

5 Sweep angle 31

6 Thrust /weight ratio .33

7 Wing loading 315 Kg/M2

8 Max take off weight 11250 kg

9 Wing area 30

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Weight EstimationThe weight of the aircraft (W) is the key factor in almost aircraft

performance problems. The gross weight is distributed in the following manner:

W = Wstruc + Wcrew + Wpass + Wfe + Wpp + Wf

Here,

Wstructure consists of the wing, fuselage, under-carriage & the empennage and accounts for about 32% of the gross weight, i.e., 0.32W.

Wfixed equipment includes the passenger seats, food, baggage racks, lavatories, air-conditioning, avionics and other passenger amenities. This adds to the weight by about 0.05W.

Wpowerplant is the weight of the engine and its systems. The initial assumption of engine weight is assumed to be 0.055W which may be modified later to suit thrust requirements.

Wfuel is the weight contribution of the fuel to the total weight. It depends on the range also includes the Reserve fuel that is used in case of an emergency. It adds to the gross weight by a factor of 0.3W.

Wcrew + Wpassengers accounts for the remaining weight. i.e., 0.275W. Taking passenger & baggage weight into consideration, a maximum of 1800N per passenger is permissible. As for a crew member, 1000N would suffice.

• W = 0.725W + (No. of passengers)*1800N + (Crew)*1000N

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• As for the aircraft to be designed, the Number of passengers will be 15 with a 5 member crew.

• W = 116363 N = 11861.7 kg (approx)

• The total weight of the power-plant (0.055W) requires being approx. 326 kg.

• The choice of a suitable engine, having been made, it is now possible to estimate the amount of fuel required for a flight at the given cruising speed for the given range.

• Wfuel = (no. of engines) x (thrust at altitude) x Range x SFC x 1.2

----------------------------------------------------------

Cruise velocity

• The factor of 1.2 is provided for reserve fuel

• Thrust at altitude is calculated using the relation

• Therefore the weight of the fuel

Wfuel = 3385.43 kg

Final weight estimation:

W = 0.425W + Wpayload +Wfuel

W = 0.425W + 3261.97 + 3385.43

Hence W = 10821.57 kg

2.10 * TT 0

alt

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This value closely matches with maximum take off weight obtained in the comparative data sheet obtained.

Engine selection:

Therefore the weight of the power plant = 652 kg

Choice of engine is a Turbofan for obvious reasons such as higher operating fuel economy & efficiency for high payloads.

Since it is a business jet of 7 to 15 seater hence two engines would be sufficient.

Engines can be used in combination of 2 x 326 kg engines

ENGINE NAME DRY WEIGHT (Kg) SFC (lb/hr/lbt) MAX. THRUST(KN)

TFE731-5A 400 0.469 20

J60-P-6 203 0.96 13.3

PW530A 275 0.468 18

CF700-202 349 0.65 20

JT15D-5 291 0.54 12.9

J85-GE-21 310 1.24 15.6

The power plant chosen is CF700-202

Wstruc = 0.32 Wg = 3600 kg

Wfixed equipment = 0.05 Wg = 562 kg

Wpower plant = 0.055 Wg = 652 kg

Wfuel = 3385.43 kg

Wcrew + Wpassengers = 3261.9 kg

Wtotal = 11460 kg

From data sheet we know that T/W = 0.33

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Hence thrust required = 37 K

Parameters :-

MTOW (N) 112435.6

Wpay (N) 32000

Wst (N) 35316

Wfe (N) 5513

Wpowerplant (N) 6396

Wfuel (N) 33211

Required Thrust (kN) 37

Thrust per engine (kN) 20

Cruise altitude (m) 11750

Sigma 0.26

Range (km) 3825

Cruise Velocity (kmph) 800

Time of cruise (hrs) 10

Power plant chosen CF700-202

Dry weight (kg) 349

Net engine weight (N) 3423.7

Thrust per engine (kN) 20

Net Thrust (kN) 40

Cruise SFC (lb/hr/lbt) 0.65

Calculated W fuel (N) 3375

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Wing selection Λ can be approx reasonably with more simply constructed trapezoidal using with a taper ratio of

λ=0.5(assuming)

S=30m2

A= (b2/S)

A=7.5= (b2/S)

b= = 15m

Formula to calculate root chord is given by

Cr=2b/ (A (1+λ))

=2 / (7.5(1+0.5)) =2.67m

Ct/Cr=0.31 Ct=2.67 0.5=1.33m

The mean aerodynamic chord length is given

C= 2Cr (1+λ+λ2)/ (3(1+λ))

=2 x 2.67(1+0.5+0.52)/ (3 1.5)

C=2.08m

The normalized span wise location of the mean aerodynamic chord from the centre span of the wing is

(Y/b)= (1+2λ)/ (6 (1+ λ))

(Y/b)= (1+1)/ (6 1.5) =0.222m

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Mach number for the a/c is given by

Mcruise =800/1188=0.673

McrD = Mcruise +0.071 = 0.744

McrDo = McrD + ΔMcrD (AR) - ΔMcrD (CL) = 0.748

From the historical data for M,

ΛLE=31⁰

Λc/4=tan-1(tan 31-((2 Cr (1- λ))/4b)

Cr =2.67m

b=15m

λ=0.5

Λc/4=29.5⁰

Aspect Ratio(AR) 7.5

Area(S) 30M2

Span(b) 15M

Cr 2.67M

Ct 1.33M

λ 0.5

Λc/4 29.5⁰

AIRFOIL SELECTION

During steady level flight, lift is equal to weight and hence

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L = W = 0.5 ×density × V2stall × S×CLmax

Density = 1.2256 kg/m3

Vstall = 1.25 × Vcruise

Vcruise = 222.22 m/s (from graph)

Vstall = 0.25×222.22 = 55.555 m/s

S = 30 m2 (from graph)

W = 11460kg

Therefore 11460 ×9.81 = 0.5×1.2256×55.5552 ×30× CLmax

CLmax = 1.98

Airfoil selection:

From the aerofoil data book various airfoils of required t/c are taken and are tabulated for maximum lift coefficient and minimum drag.

From the table the airfoil with optimum combination of maximum lift coefficient and minimum drag coefficient is selected.

Airfoil CLmax CDomin L/D

652015 1.4 0.004 350

652215 1.5 0.004 375

652415 1.6 0.004 400

652415 (a = 0.5) 1.6 0.004 400

Airfoil selected (root): NACA 652415

All airfoils listed below are from NACA series

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Airfoil CLmax CDomin L/D

63210 1.55 0.0045 344

63A010 1.2 0.0045 226

63A210 1.425 0.00425 335

64210 1.45 0.004 362.5

64110 1.4 0.004 350

64A010 1.225 0.00425 288

64A210 1.425 0.004 356.27

64A410 1.625 0.00425 382

65210 1.4 0.00375 373

65410 1.525 0.00375 406

66210 1.2 0.003 400

Selected airfoil (mean): NACA 64A410

6%Airfoil CLmax CDomin L/D

63006 0.9 0.004 22564006 0.86 0.00375 229.333364206 1.1 0.0038 289.473765006 0.95 0.0034 279.4118

65206 1.1 0.0035 314.2857

Airfoil selected (tip): NACA 65206

Cr =

8

= 2.666m

Ct = 0.5 × Cr

= 1.333 m

Cm =

= 2 m

parameter root mean Tip

Chord (m) 2.666 2.0 1.333

NACA 652415 64A410 65206

t/c 15 % 10 % 6%

Clmax 1.6 1.625 1.1

Cdomin 0.004 0.00425 0.0035

Volume of the fuel that can be stored in the wing

V = (0.5×Cm × t/c × Cm ×b/2 × 0.75) ×0.75×2

= 1.68m3

So volume of fuel that can be stored in wings is 1.68 m3

Total Volume of the fuel to be carried

Vt =

= 4.21 m3

Remaining Fuel = 4.218 –1.68 = 2.58 m3

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The remaining fuel is stored in the fuselage.

Flap selection:

With zero final velocity and with deceleration aided by thrust reverser, landing speed is calculated using the following equation. Since our aircraft is a 15 seater aircraft,

Runway length = 3100 m

Ground run= 60% of (Runway length) = 1850m

V2-VL2=2 a S

V=0, a= -0.25g s=1850m

VL = 95.25 m/s.

Vstall = VL /1.3= 73.27m/s

From the stalling speed the lift coefficient required to avoid stall is calculated as follows,

CL stall =

= 0.95

CLmaxavailable = 0.95

CLrequired = 1.98

CLreq = 1.03

Part span correction required for flap:

The flap on the wing of the aircraft does not run along the entire span. As a convention we can approximately take the flap is provided for quarter span of the wing i.e. flap span is b/4.

Hence CLactual with the part span correction is given by,

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CLactual =CLrequired /4 (since the flap is of quarter span)

Length = 50% of Semi span

Cf = 0.25 C

Split Flap:

Flap deflection

Full span Part span 50%

CL CDo CL CDo

0 0 0 0 0

10 0.22 0.018 0.11 0.009

20 0.42 0.048 0.21 0.024

30 0.6 0.09 0.3 0.045

40 0.65 0.136 0.325 0.068

50 0.79 0.18 0.395 0.09

60 0.825 0.22 0.4125 0.11

LANDING GEAR SELECTION

In aviation, the undercarriage or landing gear is the structure (usually wheels) that supports an aircraft and allows it to move across the surface of the Earth when it is not flying.

Overview

Landing gear usually includes wheels equipped with shock absorbers for solid ground, but some aircraft are equipped with skis for snow or floats for water, and/or skids or pontoons (helicopters).

Types of gear arrangements

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Wheeled undercarriages come in two types: conventional or "tail dragger" undercarriage, where there are two main wheels towards the front of the aircraft and a single, much smaller, wheel or skid at the rear; or tricycle undercarriage where there are two main wheels (or wheel assemblies) under the wings and a third smaller wheel in the nose. Most modern aircraft have tricycle undercarriages. Tail draggers are considered harder to land and take off, and usually require special pilot training. Sometimes a small tail wheel or skid is added to aircraft with tricycle undercarriage, in case of tail strikes during take-off. The Concorde, for instance, had a retractable tail “bumper” wheel (as delta winged aircraft need a high angle when taking off). Some aircraft with retractable conventional landing gear have a fixed tail wheel, which generate minimal drag and even improve yaw stability in some cases.

Retractable gear

To decrease drag in flight some undercarriages retract into the wings and/or fuselage with wheels flush against the surface or concealed behind doors; this is called retractable gear.

A design for retractable landing gear was first seen in 1876 in plans for an amphibious monoplane designed by Frenchmen Alphonse Pénaud and Paul Gauchot. Aircraft with at least partially retractable landing gear did not appear until 1917, and it was not until the late 1920s and early 1930s that such aircraft became common. By then, aircraft performance was improved to the point where the aerodynamic advantage of a retractable undercarriage justified the added complexity and weight. An alternate method of reducing the aerodynamic penalty imposed by fixed undercarriage is to attach aerodynamic fairings (often called "spats" or "pants") on the undercarriage, with only the bottoms of the wheels exposed.

Large aircraft

As aircraft grow larger, they employ more wheels to cope with the increasing weights. The Airbus A340-500/-600 has an additional four-wheel undercarriage bogie on the fuselage centerline. The Boeing 747 has five sets of wheels, a nose-wheel assembly and four sets of four-wheel bogies. A set is

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located under each wing, and two inner sets are located in the fuselage, a little rearward of the outer bogies.

Unusual types of gear

Some planes use wheels only for take off and drop them afterwards to gain the improved streamlining without the complexity, weight and space requirements of a retraction mechanism. In this case, landing is achieved on skids or similar simple devices. Historical examples include the Messerschmitt Me 163 and the Messerschmitt Me 321. A related contemporary example are the wingtip support wheels ("Pogos") on the U-2 reconnaissance aircraft, which fall away after take-off; the aircraft then relies on titanium skids on the wingtips for landing. Landing gear on an Airbus A310 an unusual undercarriage configuration is found on the Hawker Siddeley Harrier, which has two main wheels in line astern under the fuselage (called a bicycle or tandem layout) and a smaller wheel near the tip of each wing. On second generation Harriers, the wing is extended past the outrigger wheels to allow greater war loads to be carried.

A multiple tandem layout was used on some military jet aircraft during the 1950s such as the Lockheed U-2, Myasishchev M-4, Yakovlev Yak-25, Yak-28 and the Boeing B-47 because it allows room for a large internal bay between the main wheels. A variation of the multi tandem layout is also used on the B-52 Stratofortress which has four main wheel bogies underneath the fuselage and a small outrigger wheel supporting each wing-tip. The B-52's landing gear is also unique in that all four pairs of main wheels can be steered. This allows the landing gear to line up with the runway and thus makes crosswind landings easier (using a technique called crab landing).For light airplanes, a landing gear which is economical to produce is a simple wooden arch laminated from ash, as used on some homebuilt aircraft. A recent addition to this type of gear is the fixed-gear RJ.03 IBIS canard homebuilt aircraft.

Steering

The steering mechanism used on the ground with wheeled landing gear varies by aircraft, but there are several general types of steering. Tail dragger aircraft may be steered by rudder alone (depending upon the prop wash produced by the aircraft to turn it) with a freely-pivoting tail wheel, or by a steering linkage with the tail wheel, or by differential braking (the use of independent brakes on opposite sides of the aircraft to turn the aircraft by slowing one side more sharply

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than the other). Aircraft with tricycle landing gear usually have a steering linkage with the nose wheel (especially in large aircraft), but some allow the nose wheel to pivot freely and use differential braking and/or the rudder to steer the aircraft.

Virgin Atlantic Airbus A340-600 landing. This airliner has an undercarriage on the fuselage belly, as well as on the wings.

Some aircraft require that the pilot steer by using rudder pedals; others allow steering with the yoke or control stick. Some allow both. Still others have a separate control, called a tiller, used for steering on the ground exclusively.

Rudder steering

When an aircraft is steered on the ground exclusively using the rudder, turning the plane requires that a substantial airflow be moving past the rudder, which can be generated either by the forward motion of the aircraft or by thrust provided by the engines. Rudder steering requires considerable practice to use effectively. Although it requires air movement, it has the advantage of being independent of the landing gear, which makes it useful for aircraft equipped with fixed floats or skis.

Direct steering

Some aircraft link the yoke, control stick, or rudder directly to the wheel used for steering. Manipulating these controls turns the steering wheel (the nose wheel for tricycle landing gear, and the tail wheel for tail draggers). The connection may be a firm one in which any movement of the controls turns the steering wheel (and vice versa), or it may be a soft one in which a spring-like mechanism twists the steering wheel but does not force it to turn. The former provide positive steering but make it easier to skid the steering wheel; the latter provide softer steering (making it easy to over control) but reduce the probability of skidding the wheel used for steering. Aircraft with retractable gear may disable the steering mechanism wholly or partially when the gear is retracted.

Differential braking

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Differential braking depends on asymmetric application of the brakes on the main gear wheels to turn the aircraft. For this, the aircraft must be equipped with separate controls for the right and left brakes (usually on the rudder pedals). The nose or tail wheel usually is not equipped with brakes. Differential braking requires considerable skill. In aircraft with several methods of steering that include differential braking, differential braking may be avoided because of the wear it puts on the braking mechanisms. Differential braking has the advantage of being largely independent of any movement or skidding of the nose or tail wheel.

Tiller steering

A tiller in an aircraft is a small wheel or lever, sometimes accessible to one pilot and sometimes duplicated for both pilots, that controls the steering of the aircraft while it is on the ground. The tiller may or may not be designed to work in combination with other controls such as the rudder or yoke. In large airliners, for example, the tiller is often used as the sole means of steering during taxi, and then the rudder is used to steer during take-off and landing, so that both aerodynamic control surfaces and the landing gear can be controlled simultaneously when the aircraft is moving at aerodynamic rates of speed.

Maximum take off weight (from data sheet) = 10821.57 N= 25165 lbs

Tricycle-wheel arrangement

Nose- 1L/G

- 2 wheelsMain – 2 L/G

- 8 wheelsWeight taken by main landing gear, Wm = 0.9 ×W/8

= 12.6 KN/wheel = 2831 lbs/wheel

Weight taken by nose landing gear, Wn = 0.1 × W/2

= 5600 N/wheel = 1258.25 lbs/wheel

Tyre sizing:

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During landing and takeoff, the undercarriage supports the total weight of the airplane. Undercarriage is of three types:

1. Bicycle type2. Tricycle type3. Tricycle tail wheel type.

A tricycle wheel type needs more takeoff distance and floor is also needs to be inclined. So we have selected a tricycle nose wheel type. Also the types of runways are also be selected with due care since depending on this criterion, wheels and tires are selected.

Being an executive transport vehicle, it may land on grass fields and may even on beaches. For tricycle nose wheel type undercarriage, the nose wheel carries 10 % of the total load and the main undercarriage carries 90% of the total load.

For different runways, the allowable loadings are given by,

Grass 21.1 N/cm2

Grass Strip 36.9 N/cm2

Asphalt (Tar) 73.9 N/cm2

Concrete 116.1 N/cm2

Wheel configurations:

A BWheel diameter 1.51 0.349Wheel width 0.715 0.312

Main wheels:

Wheel diameter = A (Wm\2) B = 1.51× (2831\2) 0.349

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= 19 inches = 0.48 meters

Wheel width = A (Wm\2) B =0.715× (2831\2) 0.312

= 6.87 inches = 0.1746 meters

Nose wheels:

Wheel diameter = A (Wn)B = 1.51× (1258.25)0.349

= 18.23 inches = 0.463 meters

Wheel width = A (Wn)B = 0.715 × (1258.25)0.312

= 6.628 inches = 0.168 meters.

FUSELAGE SELECTION

Length of the fuselage = 16 m

Diameter of the fuselage = 1.74 m

HORIZONTAL TAILPLANE

From the standard design books =0.5 ARht= 4

Sht/s=0.15 S=30m2

Sht=4.5m2

bht=

bht=4.24m

Crt= (2Sht)/ ( +1) bnt) = (2 4.5)/(1.5 4.24)

Crt=1.415m

Ctt=0.5Crt=0.7m

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Yht= (bt (1 ))/(6(1+ )) =(4.24(1+2(0.5))/(6(1+0.5))=0.94m

Cht= ((2Crt (1+ + 2))/ (3(1+ ) = (2 1.415(1+0.5+0.52)/ (3 (1+0.5))Cht=1.1m

Vertical tail plane

From standard design books Svt/S=0.08 ARvt=1.5

Svt=2.4m

bvt=

bvt=1.879m

Crt= (2 Svt / ((1+ ) bvt)) = ((2 /(1+0.5) 1.897))

Ctt=Crt0.5=0.84m

Yvt = (2 bvt (1 ))/ (6(1+ )) =0.843m

Cvt= ((2Crt (1+ + 2))/ (3(1+ ) =1.3m

Horizontal Tail:

VHT = 0.7

ARht = 4

SHT = 4.5 m2

bht = 4.24 m

Cr = 1.415 m

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Cmht = 1.1 m

Ct = 0.7 m

Yht = 0.94 m

Vertical Tail:

VVT = 0.04

ARvt = 1.5

SVT = 2.4 m2

bvt = 1.897 m

Crvt = 1.68 m

Cmvt = 1.3 m

Ctvt = 0.84 m

Yvt = 0.843 m

Performance

Take off performance

The take off distance is given by

Sto = (1.44×W2)/ (g×ρ×S×CLmax[T-(D+μ (W-L))]

CLtakeoff = CLmax (Vstall/Vtakeoff)2

Assuming takeoff speed is 10 % greater than the stall speed

CLtakeoff = CLmax (1/1.1)2 = 1.945 × 0.826 = 1.61

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L= 0.5 × ρ×V2×S×CL

= 0.5×1.2256×55.5552×30×1.61

L= 91.35 KN

We know that W= 110362.5 N and T/W=0.33

Hence T=36419.6 N

Assuming L/D = 15

D=91350/15 = 6090 N

Thrust required = 40 KN

Therefore

Sto = (1.44×W2)/ (g×ρ×S×CLmax [T-(D+μ (W-L))]

Sto = 0.745 km

Landing performance:

The landing distance is given by

SLo = (1.69×W2)/ (g×ρ×S×CLmax (D+μ (W-L)))

Here WLO = Wg-0.3 Wg = 77.25 KN

Therefore

SLo = (1.69×W2)/ (g×ρ×S×CLmax (D+μ (W-L)))

SLo = 2.1 km

Climbing flight:

R/c = V sin γ

= V (T-D)/W

= (PA-PR)/W

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R/C = 68.28 m/s

68.28= V sin γ

γ = 17°53°

Horizontal turn:

Bank angle is Ø

Turn angle is θ

The sustained turn is considered for this:

nm = (T/W) × (L/D) = 4.95

nmax is the maximum sustained load factor.

But some factor of safety should be given hence n is assumed to be 3.5

L cos Ø = W

Sec Ø = nmax

Ø = sec-1(4.95)

Ømax = 78°30°

For normal horizontal turn n = 2.5

Hence Ø = 70°31°

Tan θ = (n2 – 1)0.5 = 4.85

(n^2 – 1)0.5 = V2/gR2 = 4.85

There fore R = 1038.43 m.

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Radius has to be decreased for better performance. It is decreased by decreasing the velocity.

ω = V/R = 0.21 m/s

Endurance:

The endurance for a jet engine is given by

E= (L/D) × (1/C) ×ln (Wi/Wf)

E = 9.59 hours.

Serial no characteristics value1 Take off distance 0.745 km2 Landing distance 2.1 km3 Rate of climb 68.28 m/s4 Climb angle 17°53°5 Turn radius 1038.43 m

AIRCRAFT DATA SHEET1 Cruise velocity 800 Km/hr

2 Design range 3825 Km

3 Cruise altitude 12750M

4 Aspect ratio 7.5

5 Sweep angle 31

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6 Thrust /weight ratio .33

7 Wing loading 315 Kg/M2

8 Max take off weight 11250 kg

9 Wing area 30

10 MTOW (N) 112435.6

11 Wpay (N) 32000

12 Wst (N) 35316

13 Wfe (N) 5513

14 Wpowerplant (N) 6396

15 Wfuel (N) 33211

16 Required Thrust (kN) 37

17 Thrust per engine (kN) 20

78 Cruise altitude (m) 11750

19 Sigma 0.26

20 Range (km) 3825

21 Cruise Velocity (kmph) 800

22 Time of cruise (hrs) 10

23 Power plant chosen CF700-202

24 Dry weight (kg) 349

25 Net engine weight (N) 3423.7

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26 Thrust per engine (kN) 20

27 Net Thrust (kN) 40

28 Cruise SFC (lb/hr/lbt) 0.65

29 Calculated W fuel (N) 3375

30 Aspect Ratio(AR) 7.5

31 Area(S) 30M2

32 Span(b) 15M

33 Cr 3.053M

34 Ct 0.947M

35 λ 0.31

36 Λc/4 29.5˚

AIRFOIL SELECTION

parameter root mean Tip

Chord (m) 2.666 2.0 1.333

NACA 652415 64A410 65206

t/c 15 % 10 % 6%

Clmax 1.6 1.625 1.1

Cdomin 0.004 0.00425 0.0035

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3-D VIEW DIAGRAM

1.74M

16M ZZZZZZZZZZZZZZZZZZZZZZZZZ

ZZZ

30 M2

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15 M

ConclusionThe aircraft is designed and the parameters like cruise velocity, wing loading, span etc have been selected for our aircraft. The weight estimation had been done to estimate the weight of our aircraft. The wings, airfoil, landing gear have been selected for our aircraft. The performance calculations were also made to estimate the performance. The aircraft parameters are in the optimum range and design characteristics have been found to be satisfactory

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BIBILOGRAPHY

1. “AIRCAFT PERFORMANCE AND DESIGN” by

John D.Anderson, Jr

2. “AIRCRAFT DESIGN” by Daniel P. Ramer

3. “ALL THE WORLD’S AIRCRAFT “ by janes

4. Web source: www.wikipedia.com

www.airliners.com

www.geae.com

www.pdas.com