Aircraft Design Project i

7/27/2019 Aircraft Design Project i

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ANNA UNIVERSITY CHENNAI-600025

MADHA ENGINEERING COLLEGE

KUNDRATHUR, CHENNAI-600069.

Department of Aeronautical Engineering

Aircraft design project 1

Long range business jet aircraft

Submitted by

KANMANI RAJA T 41108101018

SARAVANA KUMAR N 41108101042

SYEDHALEEM M 41108101052

Guided by,

Mr. R K MUTHURAMAN B.E. (MBA)

Lecturer,

Department of aeronautical engineering,

Madha engineering college,

Chennai-69.


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ANNA UNIVERSITY::CHENNAI 600025

BONAFIDE CERTIFICATE

This is to certify that this project DESIGN OF LONG RANGE BUSINESS JET

AIRCRAFTis the bonafide work of

KANMANI RAJA T 41108101018

SARAVANA KUMAR N 41108101042

SYEDHALEEM M 41108101052

SIGNATURE OF GUIDE, SIGNATURE OF HOD,

MR. R. K. MUTHURAMAN B.E., (MBA) MR. J. KUMARAGURUBARAN M.E.

LECTURER, HEAD OF THE DEPARTMENT,AERONAUTICAL DEPARTMENT, AERONAUTICAL DEPARTMENT,

MADHA ENGINEERING COLLEGE, MADHA ENGINEERING COLLEGE,

CHENNAI-69. CHENNAI-69.

Viva voce held on___________

INTERNAL EXAMINER EXTERNAL EXAMINER


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ACKNOWLEDGEMENT

We would like to thank our chairman and founder of Madha group of

Academic institutions Dr. Ln. S. Peter for his Excellent contribution

towards the department.

We would also like to thank our kind principal Dr. C. B. Lakshmikantha

B.Tech., M.Tech., Ph.D., MISTE, MTAI., for his extended support and

motivation.

We would also like to thank our beloved HoD, Mr. J. Kumaragurubaran

M.E., for helping us in times of need and guiding us and maintaining the

department in an excellent manner.

We would like to thank our guide, class in charge, Mr. R.K. Muthuraman

B.E. (MBA), for his contribution towards making this project into a

successful one and guiding and for motivating us.

Finally, we would like to thank the staff members of the department of

aeronautical engineering and our beloved friends who stood by us and

helped us in the completion of the project.


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CONTENTS

Expt.

No.

Date

of

Expt.

Name of the Experiment Page

no

Guides initial

1 04:01:11 The design process 4

2 18:01:11 Literature survey 13

3 1801:11 Comparative study 24

4 25:01:11 Selection of main parameters 27

5 02:02:11 Weight estimation 29

6 09:02:11 Selection of airfoil 37

7 09:02:11 Estimation of Maximum Cl 44

8 17:02:11Selection of wing & control

surfaces 47

9 24:02:11 Estimation of wing loading 55

10 03:03:11 Estimation of thrust to weight ratio 62

11 10:03:11 Selection of powerplant 66

12 17:03:11 Performance curves 70

13 24:03:11 3 view diagram 79


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INDEX

Description page no.

1. Abstract12. Introduction to aircraft design project23. Introduction to design process4

Starting a design process.4 Phases of airplane design5 Conceptual design...5 Preliminary design...5 Detail design6 Requirements...7 Weight of the airplane.7 Critical performance parameters7 Configuration layout...8 Better weight estimate.8 Performance analysis......8 Optimization....8 New design..9 Design aspects9 Performance aspects...9

4. Literature survey.13 Classification of airplanes..14 Based on operation.14 Based on configuration..14 Based on position of wings....15 Based on shape of wings....16 Based on engines....19 Based on fuselage...21


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Based on landing gear21 Consideration for own aircraft..22

5. Comparative study.24 Comparative data sheet.25

6. Selection of main parameters277. Weight estimation..29

Fuel fraction estimation30 Estimation of empty weight ratio.34 Iteration table35

8. Selection of airfoil.37 Airfoil families..37 NACA 4 series..37 NACA 5 series..38 NACA 1 series (16 series)....39 NACA 6 series..39 Selection of airfoil....41 Characteristic curves of selected airfoils....41

9. Estimation of maximum lift coefficient..44 Average maximum lift coefficient..44 Landing maximum lift coefficient..44 Take off maximum lift coefficient..44

10.Selection of wing and control surfaces...47 Calculation of wing dimensions.47 Dihedral and sweepbackeffect..49 Selection of control surfaces..52

11.Estimation of wing loading.55 Stall velocity constraint..55 Landing distance constraint...56

12.Estimation of thrust to weight ratio...6213.Powerplant selection..66

Specifications.68


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14.Performance curves..70 Drag polar.71 V vs. L/D..72 V vs. Treq.73 V vs. Preq.73 V vs. Tav..74 V vs. Pav..74 V vs. T..75 V vs. P..75 V vs. R/C..76

15.Three view diagram.7816.Conclusion...7917.References80


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ABSTRACT

The ultimatum of this project is the design of a long-range business jet aircraft with the

desired specifications. This aircraft design project is nothing but designing our own

imaginative aircraft with some help from Existing available data of similar types of

aircrafts. This project boosts up the innovative and creative part of the mind. During the

design process, the Existing theoretical formulae, concepts, basics are scrutinized to aid

in the design process, thus developing ones mind and a better understanding capacity.

This design process also helps in developing the potential in ourselves. This project is just

a basic design with the use of basic formulae. However, designs based on these formulae

have been found to comply to the desired specifications with minimal variations. This

design process includes some of the basic estimations like weight, wing parameters and

airfoil, critical performance parameters and the powerplant and finally the three-view

diagram of the aircraft for which the design is being proposed.


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INTRODUCTION TO AIRPLANE DESIGN

Airplane design is both an art and a science. We can see by the name itself, airplane

design involves Experience and practice rather than just a book. However, theory and

Experimentation are interconnected, so we ought to go through the available books on

design before starting a detailed design procedure. Airplane design is the intellectual

process of creating on paper or on a computer screen a flying machine to meet certain

specifications and requirements established by us, the designers or the concerned person

or the firm for whom we are designing the aircraft. Airplane design is said to have started

from the ages of Leonardo Da Vinci, Sir George Cayley, Otto Lilienthal, Alexander

Mozhaiski, Felix Du Temple, Langley and the Wright brothers. In Indian Mythology, the

demon king Ravan was believed to be fond of science. He is said to have designed an

aircraft called the Pushpak Vaaghan that he used to abduct Sita, the wife of lord Ram, the

prince of Ayodhya. The design process is thus a process that does not stop. Even though,

many modern configuration aircrafts Exist today and yet more are to come, the design

process will never stop until the human desires of 100% efficiency, comfort, convenience

and luxury are met.


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EX. No. 1 THE DESIGN PROCESS 04:01:11

Aim

To make a brief study of the airplane designing process to aid in the designing of long

range business jet aircraft.

Introduction

Those involved in design can never quite agree as to just where the design process

begins. The designer thinks that it starts with a new airplane concept. The sizing

specialist knows that nothing can begin until an initial estimate of weight is made. The

customer feels that the design begins with requirements. They all are correct. The design

process is actually an iterative process.

Starting a design process

The start of the design process requires the identification of need. It is essential to

understand at the start of the study where the project originated and to recognize what

External factors are influential to the design before the design process is started. The

design process never ends as the designers continuously provide many modifications to

the aircraft to improve its safety and performance, services and any repairs, maintenance

instructions etc that are necessary to keep the aircraft in an airworthy condition.

Many airplanes never make it beyond the initial or preliminary design phase. In fact most

dont. What happens beyond the preliminary design phase depends largely on the results

obtained during the preliminary design and on the real or perceived market interest

afterwards.

If, because of the preliminary design studies a specific need can be met, then the full-

scale development of the aircraft can follow. If, because of the preliminary design phase

certain problem areas are discovered and then a research and development program can


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be initiated aimed at overcoming these problems. Eventually, with the problems solved

which then can lead to a full-scale development.

Phases of airplane design

The complete design process involves three distinct phases that are carried out in

sequence. The phases are,

Conceptual Design Preliminary Design Detail Design

Conceptual Design

Usually, the design process starts with a set of preliminary requirements or specifications

for the new airplane or with the desire to introduce some innovative ideas and

technology. In this the overall shape, weight size and performance of the new design are

determined. The result of conceptual design is the layout on a sheet of paper or on the

computer screen of the airplane configuration. However, these drawings have to be

visualized as flexible lines as they are prone to modifications during the second phase of

the design process. It gives out some data like the wing and tail dimensions, position and

type of the engine etc. During the conceptual design phase, the designer is influenced by

qualitative aspects such as the increased structural loads imposed by a high horizontal T-

tail versus a conventional tail location through the fuselage, and the difficulties associated

with the cut outs in the wing structure if the landing gears are to retract into the wing

rather than the fuselage or engine nacelle.

Preliminary Design

Minor changes are made to the configuration layout from the conceptual design in thispreliminary design process. It is in this process serious structural and control system

analysis and design takes place. In addition, during this phase substantial wind tunnel

testing will be carried out and CFD (Computational Fluid Dynamics) calculations of the

complete flow over the airplane configuration will be made. These tests will uncover

some undesirable aerodynamic interference, or some Expected stability problems. At the


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end of preliminary design, the airplane configuration is fixed. It will not undergo any

further modifications. The drawing process called lofting is carried out that

mathematically models the precise shape of the outer skin of the airplane, making sure

that all the sections of the aircraft properly fit together. The future of the design rests in

the result of the preliminary design process whether to commit to the manufacture of the

airplane or not.

Detail Design

The detail design is literally the nuts and bolts phase of the airplane design. The

aerodynamics, propulsion, structures and flight control analyses all have been finished

with the preliminary design phase. In this, the airplane is simply a machine waiting to be

fabricated. The precise design of each individual component like the ribs, spars, section

of skin etc, takes place. The size, number and location of fasteners are determined.

Manufacturing tools and jigs are designed. The flight simulators for the aircraft are

developed. At the end of this phase, the aircraft is ready to be fabricated.


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Requirements

Where and how to start the design? Clearly, with a list of requirements for the airplane,

we can start the design process. The requirements might be from the designer himself or

the customer for whom he is designing the airplane. Like the fingerprints for every

human being differ, so do the requirements of every new airplane. Frequently stipulated

aspects of requirements include the following,

Range Take off distance Stalling velocity Endurance Maximum velocity Rate of climb For combat type of aircraft maximum turn rate and minimum turn radius Maximum load factor Service ceiling Cost Reliability and maintainability Reasonable size

Weight of the airplane

For an airplane to get off the ground, it must be able to produce a lift greater than its

weight. Therefore, the estimation of weight is an important step in the design process. It

is known to all as the weight of airplane increases, so does the lift required to overcome it

and consequently the drag of the airplane increases. Therefore, the estimation of weight

of the airplane is the first step in the design process.

Critical performance parameters

It focuses on the estimation of some critical performance parameters like,

Maximum lift coefficient CLmax Lift to drag ratio L/D (usually at cruise) Wing loading W/S


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Thrust to weight ratio T/W

Configuration layoutThe configuration layout is the drawing of the airplanes shape and size. The critical

performance parameters along with the initial estimate of weight give enough information

to approximately size the aircraft and draw the configuration.

Better weight estimate

The pivot point at this stage is the improved estimate of weight based on the performance

parameters. A detailed component weight breakdown based on the configuration layoutand a more detailed estimate of the fuel weight necessary to meet the requirements.

Performance analysis

At this pivot point, the design of the aircraft from the previous stage is put through a

preliminary performance analysis. The configuration is judged whether it can meet all the

original specifications set forth. This is obviously critical point in conceptual design

process. An iterative process is initiated wherein the configuration is modified, with the

Expectation of coming closer to meeting the requirements. The iteration is repeated until

the resulting airplane meets the requirements.

Optimization

After the iteration process, the next question that arises in the designers mind is that Is

it the best design? This leads to optimization analysis. The optimization is carried out by

plotting the performance of different airplanes on graphs that provide a sizing matrix or a

carpet plot from which the optimum design can be found.


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New design

The following are some basic areas where we should concentrate to

design a new aircraft

Aerodynamics Propulsion Light weight structures Controls

The above areas involve some parameters like

Size Shape Weight Performance

Of these parameters, we should use the optimized value for our new design, and it should

be selected on the basis that it would not affect the other parameters.

Design aspects

For passenger aircraft

High AR wings High wing loading, in order to minimize lift and induced drag for efficient cruise.

For fighter aircraft

Low AR wings Low wing loading

Structure factor

It is defined as the ratio between empty weight to the total take off weight.


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Performance aspects

Aircraft purpose Type of payload

Cruise and maximum speeds Maximum cruise altitude Endurance Range Take off distance at maximum weight Landing distance with 50% of maximum fuel weight Purchase cost Other requirements

Aircraft purpose

Our design of aircraft starts with deciding the purpose of the aircraft. There are three

major purposes for use of aircrafts

Military aircrafts [fighter & bomber] Passenger aircrafts Cargo aircrafts

Payload

The material, which is carried onboard and delivered as a part of the mission, is called the

payload. There are two types of payloads

Non-Expendable payloads

Expendable payloads

Non-Expendable payloads are Expected to be transported during the complete duration of

flight plan. E.g. Passengers and cargo.

At some point in the flight plan, it permanently leaves the aircraft. E.g. bombs, rockets,

missiles.

For business jet aircrafts the payload includes the passengers, passenger baggage and

crew members.


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Cruise and maximum speeds

The mission of the aircraft is usually determined through the range of speed of the

aircraft.

The propeller driven aircrafts are usually designed to cruise at speeds of 150-300 knots.

The jet-powered aircraft has higher cruise speeds than that of propeller driven aircrafts.

The speed of jet propelled aircraft is given in terms of mach no. for business and

commercial jet aircrafts 0.8


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EX. No. 2 LITERATURE SURVEY 18:01:11

Aim

To do a brief literature study on the existing airplanes and about their merits and

demerits.

Introduction

It is essential to go through some existing data on different types of aircrafts and the

components used in them and their evolution with time, before starting the design

process. Literature survey is one such thing. It consists of configuration studies, design

trades etc. first, a basic understanding of the types of flying machines present is

necessary. The aircrafts are classified according to their weight, propulsion system, and

place of use. Lets see the classification of airplane,

Aircrafts

Lighter than air Heavier than air

Airships Free Balloons Captive Balloons Power Non power Man powerdriven driven driven

Gliders Seaplanes Kites

Airplane Rotorcraft Ornithopters

Land plane Sea plane Amphibian

Helicopter Gyroplane Cyclogyroplane


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Float plane Flying boat

Classification of airplanes

Based on operation

Subsonic

Transonic Supersonic Hypersonic

Based on configuration

Monoplane

Merits

Simple to construct Less interference drag Less induced drag Higher aspect ratio

Demerits

Heavier structural weightsBiplane

Merits

Lower structural weights

Demerits

Lower aspect ratio Higher induced drag Complex to construct


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Triplane

Merits

Lower structural weightDemerits

Higher interference drag Lower aspect ratio Higher induced drag Complex to construct

Position of wings

High wing

Merits

More stable Lesser interference of fuselage on wing flow Easy to fix wing to fuselage Larger height and larger clearance for service vehicles Podded engines are away from the ground.

Demerits

Inspection of top surface is difficult More stick force is required

Low wing

Merits

Easy to attach to the fuselage Smaller stick force Inspection of top surface is easy


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Demerits

Smaller clearance for service vehicles Podded engines are closer to the runway Less stable More interference drag

Mid wing

Merits

Provides the lowest drag of any three locations of the wing body Interference is minimized Fillet is not required to decrease interference Wing bending moment can be transmitted across the fuselage by a series of heavy

ring frames in the fuselage shell

Demerits

The bending moment due to wing lift will be carried through the fuselage thatimposes structural limitations

There would be an unacceptable obstruction through the middle of the fuselage

Shape of the wing

Rectangular wing

Merits

Higher aspect ratio

Smaller induced drag Easy to construct

Demerits

Larger structural weight


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Tip stall occurs

Tapered wingMerits

Lower structural weight Lesser induced drag No tip stall occurs

Demerits

Difficult to construct Root stall occurs Small aspect ratio

Canard wing

Merits

Tail will be more effective and is not in the effect of the wing More lift Lower structural weight Stability increases with Mach number

Demerits

Less stable Large control forces at higher Mach number Small tail lever arm Shock stall can occur


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Elliptical wing

Merits

Least induced drag

Smaller structural weight Higher drag divergence Mach number

Demerits

Difficult to construct Smaller aspect ratio

Delta wing

Merits

Smaller structural weight Higher drag divergence Mach number Not very difficult to manufacture

Demerits

Smaller aspect ratio Higher induced drag Large area of wing controls

Swept back wings

Merits

Higher drag divergence Mach number

Smaller structural weight More stable

Demerits

Lower aspect ratio


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Higher induced drag Ailerons are less effective

Swept forward wings

Merits

Higher drag divergence Mach number Cross flow could be omitted

Demerits

Higher structural weight

Smaller aspect ratio Less stable

Blended wing or flying wing

Merits

Structurally strong Able to carry more payload Lift generation is more Takes advantage in wing boundary layer ingestion Tail drag is absent

Demerits

Large in shape and size High cost Complicated constructions

Engines

Types of engines used for power plant

Piston engines Turbo prop engines


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Turbojet engines Turbofan engines Ramjet engines Rocket

Number of engines

Single engine Twin engine Multi engine

Location of engines

In modern transport, pylons can hold engines

Merits

Weight decreases by 15-20% Wing space can be utilized for fuel Maintenance, inspection and replacement are facilitated Wing structure is free from the heat of the engines that improves fire safety

Demerits

Failure of outboard engine creates a large yawing moment This Moment has to be countered by rudder deflection that results in higher drag High acoustic stresses on the ailerons and load bearing members in the lower part

of the wing call for an increase in wing rigidity and weight

Noise level in cabin is 5 dB higher as compared to aircraft having engines o therear fuselage

Smaller ground clearance increases chances of FODEngines located in wing root

Merits

Very little increase in frontal area Entire wing span can be utilized for ailerons and high lift devices


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Demerits

Weight is more due to compensation of cuts in wing spars Space in the root section of the wing cannot be utilized for the storage for fuel Intake is located at a place where the airflow is not clean

Engines located on rear fuselage

Merits

Less noise in the cabin Entire wing space can be used to store fuel Whole wingspan can be used for ailerons and high lift devices Fire hazard is a minimum

Demerits

Fuel is located far from the engines which increases the length of fuel linesrequired and special fuel pumps

Due to weight at the tail, large horizontal and vertical tail surface areas arerequired

Fuselage

Conventional single fuselage design Twin fuselage design Pod and boom construction type

Landing gear

Retractable landing gear Non-retractable landing gear Nose wheel landing gear Bicycle landing gear


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Consideration for own aircraft

From the above types, we have selected a conventional single fuselage design with twin

engines mounted on the rear fuselage and a sweptback tapered low wing configuration.

The aim of this project is to design a long-range business jet i.e. having a range of >5000

km. The maximum speed to be in the range of 800-950 kmph, service ceiling around

14000 m. since, it is business jet it will be carrying not more than 20 passengers. Most of

the business jet aircrafts have an average accommodation for 12 passengers.

Conclusion

Thus, the brief study on available literature was done and some preliminary factors were

considered for our aircraft.


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EX No. 3 COMPARATIVE STUDY 18:01:11

Aim

To do a comparative study on similar types of existing aircrafts and to prepare a

comparative data sheet.

Introduction

The comparative study consists of preparing a comparative data sheet using data

collected from Existing aircrafts similar to the aircraft that is to be designed. The data

sheet focuses attention on some of the important parameters of the aircraft like,

Empty weight Fuel weight MTOW Thrust Powerplant Maximum speed Range Service ceiling T/W ratio Wing loading Passengers and crew Wing area Wingspan

We are considering five aircrafts for our comparative study. They are,

Name of the aircraft Manufacturer Country

Legacy 600 Executive Embraer Brazil

Boeing business jet Boeing USA

Falcon 7X Dassault aviations France

Gulf stream V (C-37A) Gulfstream aerospace USA

Bombardier Global 5000 Bombardier aerospace Canada


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Comparative data sheet

Aircraft/Char Boeing

Business Jet

Legacy 600

Executive

Bombardier

Global 5000

Falcon 7X Gulf Stream

V

Crew 2+2 flightattendants

2+1 flightattendant

optional

2+1 flightattendant

2+1 flightattendant

2+0 to 2flight

attendants

Length 39.47 m 26.33 m 29.49 m 23.19 m 29.4 m

Wingspan 35.79 m 21.17 m 25.15 m 29.69 m

Wing area 70.7 m 105.6 m

Empty weight 43082 kg 13250 kg 22838 kg 15456 kg 17917 kg

Loaded weight 57155 kg 16000 kg 25401 kg 21682 kg

MTOW 77565 kg 22500 kg 39780 kg 28893 kg 38600 kg

Fuel weight 32825 kg 8140 kg 16329 kg 13109 kg 15966 kg

Power plant 2*CFM 56-7turbofans

2*RollsRoyce

AE3007/A1

P turbofans

2*BMWRolls Royce

BR710

turbofans

3*Pratt &Whitney

Canada

PW307Aturbofans

2*RollsRoyce

Deutschland

BR710-48turbofans

Thrust 121.4 kN 2*37.1 kN 3*21.7 kN 2*68.4 kN

Max. speed 890 kmph 834 kmph 945 kmph 685 kmph 1056 kmph

Range 11482 to

9936 km

5926 km 8889 km 10556 km 10742 km

Service ceiling 12496 m 11885 m 15545 m 15545 m 15545 m

Wing loading 620.5 kg/m 439.6 kg/m 419 kg/m 408.7

kg/m

2365.4 kg/m

T/W ratio 0.52:1 0.36:1

Payload/passeng

ers

8 to 50 13+1 in

cockpit jump

seat

19 14 14 to 19

This comparative data sheet will help us in determining the main parameters and limits

for our aircraft.

Conclusion

Thus, the data from similar types of existing aircrafts was studied, and the comparative

data sheet was prepared and parameters were compared.


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Ex No. 4 SELECTION OF MAIN 25:01:11

PARAMETERS

Aim

To fix or select primary requirements for our aircraft.

Introduction

Before designing the aircraft, the main parameters of the aircraft have to be selected. The

customer provides them in case the aircraft is being designed for him. The main

parameters are simply the aircrafts requirements. They are as follows,

Criteria Value

Payload-Passengers 12 at 100 kg each (220.46 lb)

Payload-Baggage 20 kg each (44.09 lb)

Crew 2 pilots, 1 cabin attendant at 100 kg each with 20 kg baggage

Range 7000 km (3779.7 nm)

Reserve fuel 100 nm (185.2 km) followed by 1 hour loiter

Cruise altitude 12000 m (40000 ft)

Cruise mach no. 0.8

Powerplant 2 turbofans

Take off field length 800 m (2624.47 ft)

Landing distance 1000 m (3280.84 ft)

Stalling velocity 50 ms-

(164.04 fts-

or 111.85 mph)

Estimated weight 38000 kg (83775.66 lb)

Thrust 70 kN (15730 lbf)

Conclusion

Thus, the primary parameters required for our aircraft were selected with help from the

comparative data sheet.


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EX No. 5 WEIGHT ESTIMATION 02:02:11

Aim

To do a conceptual weight estimation of the aircraft. Its fuel weight, empty weight and

total weight.

Introduction

Design takeoff gross weight is the total weight as the aircraft begins its mission for which

it was designed. This is not necessarily same as the MTOW or the maximum takeoff

weight for all aircrafts.

The design takeoff gross weight can be broken down into crew weight, payload or

passengers weight, fuel weight and the remaining or empty weight. The empty weight

includes the structure, engines, landing gear, fixed equipments, avionics and anything

else that is not considered as a part of the crew, payload or fuel. The following equation

summarizes the design takeoff weight,

The crew and the payload weights are known. The only unknowns are the fuel weight and

empty weight. The following iterative equation is used to calculate the weight,


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Now, W0 can be determined if Wf/W0 and We/W0 are known.

In order to meet the mission requirements the aircraft must be able to take off and climb

to the desired cruising altitude and then cruise for the desired range and land. There

should be enough additional fuel for emergency climb & cruise and loitering before

landing.

Fuel fraction estimation (Wf/W0)

The amount of fuel required to carry out the mission effectively depends upon the

efficiency of the powerplant and efficient aerodynamics of the airplane. The total fuel

consumed during the mission is the amount that is consumed from engine start, taxi, take

off, cruise, descent, landing, and taxi and finally shut down. Sometimes, there is some

loitering when the aircraft is put into a holding pattern by the ATC prior to landing. The

fuel fraction for each segment can be calculated from Existing graphs or from some other

formulae. Generally, Breguets range and endurance equations are used. For loitering

segment, a usual timing of 20 min is sufficient. However, we are selecting a loitering

time of 1 hour. Further, a reserve of fuel for a cruise for 100 nm is considered. The

calculations then are as follows,

Engine start or warm upW1/W0 = 0.99

W1=37620 kg

Taxi

W2/W1=0.995

W2=37431.9 kg

Take off

W3/W1=0.995


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W3=37244.74 kg

Climb

W4/W3=0.98

W4=36499.84 kg

Cruise

Using Breguets range formula,

R-range in m

Ct-TSFC in s-1

W0- initial weight

W1- final weight

Value of Ct

For high bypass turbofan business jets

For cruise Ct=0.5 lb/lbf-h = 14.16*10-6

kg/N-s

For loiter Ct=0.4 lb/lbf-h = 11.32*10-6

kg/N-s

Value of L/D

(L/D)max = 10+AR

=10+9 = 19

L/D = 0.866*(L/D)max = 0.866*19 = 16.454


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Rearranging Breguets range equation

W5/W4 = Exp (-RCt g/ (VL/D))

=Exp (-7000*103*14.16*10

-6*9.81/ (236.039*16.454))

=0.778516

W5 = 28415.709 kg

Emergency climb

W6/W5 = 0.98

W6 = 27847.395 kg

Emergency cruise (for 100 nm = 185.2 km)

W7/W6 = Exp (-185.2*103*14.16*10

-6*9.81/ (236.039*16.454))

= 0.993398

W7 = 27663.544 kg

Loitering (for 1 hour = 3600 seconds)

Using the endurance formula

Rearranging the above equation


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W8/W7 = Exp (-E*Ct*g/ (L/D))

During loitering L/D is (L/D) max

W8/W7 = Exp (-3600*9.81*11.32*10-6

/19)

= 0.97918

W8 = 27087.559 kg

Descent

W9/W8 = 0.990

W9 = 26816.683 kg

Landing, taxi and shut down

W10/W9 = 0.992

W10

= 26602.1499 kg

Fuel fraction

W10/W0 = (W1/W0)*(W2/W1)*(W3/W2)**(W10/W9)

= 0.99*0.995*0.995*0.98*0.778516*0.98*0.993398*0.97918*0.99*0.992

= 0.70006

Therefore, the fuel fraction is

Wf/W0 = (1+mfres)*(1-W10/W0)


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Where mfres is the weight fraction of reserved and trapped fuel. It is usually about 6% of

the total fuel weight.

Wf/W0 = (1+0.06)*(1-0.70006)

= 0.3179

Estimation of We/W0

A new design always has an evolutionary change from an Existing aircraft. Therefore, we

can assume a value for this ratio from Existing data from similar aircraft. The graphs

yield an equation for the calculation of empty weight ratio. It is,

We/W0 = AW0cKvs

A & c are constants for particular type of aircraft, for business jets A = 1.02 and c = -0.06

Kvs has a value of 1 for fixed sweep and 1.04 for variable sweep

We/W0 = 1.02W0-0.06

Weight of crew and payload (passengers)

Wcrew = (100*3) + (20*3) = 360 kg

Wpayload = (100*12) + (20*12) = 1440 kg

Substituting the values in the iteration equation


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W0 = 1800/ (1-0.3179-(1.02W0-0.06

))

Iteration table

38000 12826.703

12826.703 17332.164

17332.164 15761.304

15761.304 16222.459

16222.459 16079.369

16079.369 16123.032

16123.032 16109.639

16109.639 16113.741

16113.741 16112.484

16112.484 16112.869

16112.869 16112.751

16112.751 16112.787

16112.787 16112.77

16112.77 16112.77

Result

Thus, the conceptual weight of the aircraft was estimated. The values of empty weight

and fuel weight were found from the iteration table the initial estimate of take of gross

weight is

W0 = 16112.77 kg

Wf= 5122.249 kg

We = 9190.521 kg


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Ex No. 6 SELECTION OF AIRFOIL 09:02:11

Aim

To study different types of airfoils and select an appropriate airfoil for the airplane

design.

Introduction

After the estimation of weight, a compatible airfoil for the type of aircraft that would be

able to generate enough lift to overcome the weight efficiently has to be selected. First, a

study of the Existing airfoils is suggested before proceeding to airfoil selection. NACA

provides a wide variety of airfoils, each having its own different merits and demerits.

AIRFOIL Families

NACA Four-Digit Series

Around 1932, NACA tested a series of airfoil shapes known as the four-digit sections.The four-digit airfoil geometry is defined, as the name implies, by four digits; the first

gives the maximum camber in percent of chord, the second the location of the maximum

camber in tenths of chord, and the last two the maximum thickness in percent of chord.

For Example, the 2412 airfoil is a

12% thick airfoil having a 2% camber located 0.4 from the leading edge.


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Advantages

Good stall characteristics Small centre of pressure movement across large speed range Roughness has little effect

Disadvantages

Low CLmax Relatively high drag High pitching moment

NACA five-Digit Series

The NACA five-digit series developed around 1935 uses the same thickness distribution

as the four-digit series.

The numbering system for the five-digit series is not as straightforward as for the four-

digit series. The first digit multiplied by 3/2 gives the design lift coefficient of the airfoil.

The next two digits are twice the position of maximum camber in percent of chord. The

last two digits give the percent thickness. For Example, the 23012 airfoil is a 12% thick

airfoil having a design Cl of 0.3 and a maximum camber located 15% of c back from the

leading edge.

Advantages

Higher CLmax Low pitching moment Roughness has little effect

Disadvantages

Poor stall behaviour Relatively high drag


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NACA l-Series (Series 16)

The NACA 1-series of wing sections developed around 1939 was the first series based on

theoretical considerations. The most commonly used 1-series airfoils have the minimum

pressure located at the 0.6 and are referred to as series-16 airfoils. The camber line for

these airfoils is designed to produce a uniform chord wise pressure difference across it. In

the thin airfoil theory to follow, this corresponds to a constant chord wise distribution of

vortices.

Operated at its design Cl, the series-16 airfoil produces its lift while avoiding low-

pressure peaks corresponding to regions of high local velocities.

Thus the airfoil has been applied extensively to both marine and aircraft propellers. In the

former application, Low-pressure regions are undesirable from the standpoint of

cavitation (the formation of vaporous cavities in a flowing liquid). In the latter, the use of

series-16 airfoils delays the onset of deleterious effects resulting from shock waves being

formed locally in regions of high velocities.

Series-1 airfoils are also identified by five digits as, for Example, the NACA 16212

section. The first digit designates the series; the second digit designates the location of the

minimum pressure in tenths of chord. Following the dash, the first number gives the

design Cl in tenths. As for the other airfoils, the last two digits designate the maximum

thickness in percent of chord.

Advantages

Avoids low pressure peaks Low drag at high speed

Disadvantages

Relatively low lift

NACA 6 Series

The 6 series airfoils were designed to achieve desirable drag, compressibility, and C l,

performance. These requirements are somewhat conflicting, and it appears that the


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motivation for these airfoils was primarily the achievement of low drag. The chord wise

pressure distribution resulting from the combination of thickness and camber is

conducive to maintaining Extensive laminar flow over the leading portion of the airfoil

over a limited range of G values. Outside of this range, C d and Clm values are not too

much different from other airfoils.

The mean lines used with the 6-series airfoils have a uniform loading back to a distance

of x/c = 2. Aft of this location the load decreases linearly.

The = 1 mean line corresponds to the uniform loading for the series-16 airfoils.

There are many perturbations on the numbering system for the 6 series airfoils. The later

series is identified, for Example, as

NACA 651-212 a = 0.6

Here 6 denotes the series; the numeral 5 is the location of the minimum pressure in tenths

of chord for the basic thickness and distribution; and the subscript 1 indicates that low

drag is maintained at Cl, values of 0.1 above and below the design C l, of the 0.2, denoted

by the 2 following the dash. Again, the last two digits specify the percentage thickness. If

the fraction, a, is not specified, it is understood to equal unity.

Advantages

High CLmax Very low drag over a small range of operating conditions Optimized for high speed

Disadvantages

High drag outside of optimum range of operating conditions Higher pitching moment Poor stall behaviour Very susceptible to roughness


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From the collected data, the airfoils that we have selected are

Root section: NACA 23021 (CLmax = 1.5)

Tip section : NACA 23012 (CLmax = 1.8)

Average CLmax = (1.5+1.8)/2 = 1.65

Characteristic lift versus angle of attack plots.


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Conclusion

The different types of airfoils were studied and the appropriate type of airfoils for the

airplane design were selected.


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Ex No. 7 ESTIMATION OF MAXIMUM CL 09:02:11

AimTo estimate the values of maximum lift coefficient during landing and takeoff for the

selected airfoils.

From the collected data, the airfoils that we have selected are

Root section: NACA 23021 (CLmax = 1.5)

Tip section : NACA 23012 (CLmax = 1.8)

Average CLmax = (1.5+1.8)/2 = 1.65

For ease of calculations, a plain flap is used. Flaps are high lift devices that are used

temporarily to increase the lift during takeoff and landing. Other high lift devices are slots

and slats.

To aid in landing a flap deflection of 450 will yield an increase in CLmax of

CLmax =0.9

Therefore, CLmax = 1.65+0.9 = 2.55

For finite wings with aspect ratio greater than 5, the CLmax is 0.9 times of the previous

CLmax

CLmax = 0.9*2.55 = 2.295

To aid in take off, a flap deflection of 200

is provided that yields an increase in CLmax of

CLmax = 0.5

Therefore, CLmax = 1.65+0.5 = 2.15


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For finite wings with aspect ratio greater than 5, the CLmax is 0.9 times of the previous

CLmax

Therefore, CLmax = 0.9*2.15 = 1.935

Therefore, the airfoils that we have selected yield the following values of maximum lift

coefficients

Result

The values of maximum lift coefficient during takeoff and landing were calculated for the

selected airfoils sections.

Average maximum lift coefficient = 1.65

Landing maximum lift coefficient = 2.295

Takeoff maximum lift coefficient = 1.935


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Ex No. 8 SELECTION OF WING 17:02:11

AND CONTROL SURFACES

Aim

To select the appropriate wing and its dimensions and the control surfaces

Description

The wing is the lift generating component and the selection and estimation of its

dimensions and the control surfaces is a crucial step in the design process as it would

affect the performance and stability of the aircraft.

We can start designing the wing by the assumed value of our aspect ratio and the span

area obtained from the wing loading value.

First we will select the type of wing for our aircraft. We have already discussed the

effects of wing shapes and their positions in the airplane in the literature study. So, from

that we are selecting a swept back dihedral tapered low wing. Now, let us proceed with

the wing calculations.

From the wing loading value,

Calculation of wing dimensions

Wing area

S = W0/ (W/S)

S = 16112.77/358.228 = 44.9791 m2


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b = sqrt(44.9791*9) = 20.1199 m

Therefore, the wingspan obtained by wing loading is b = 20.1199 m. Ifctbe the tip chord

and crbe the root chord and taper ratio be = ct/cr

For most of the commercial aircrafts, a taper ratio of 0.3 is taken

cr= 2*44.9791/ ((1+0.3)*20.1199)

Root chord = 3.439 m

Tip chord = 1.0318 m (where tip chord = 0.3*root chord)

The thickness of the wing at the root (NACA 23021) = 3.439*0.21 = 0.722 m

The thickness of the wing at the tip (NACA 23012) = 1.0318*0.12 = 0.124 m


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To find mean aerodynamic chord

Y = (20.1199/6)*((1+2*0.3)/(1+0.3))

Y = 4.127 m

C = (2*3.439/3)*((1+0.3+0.32)/ (1+0.3)) = 2.4514 m

The mean aerodynamic chord has a value of 2.4514 m at a distance of 4.127 m from the

root.

Dihedral and sweep back effect

The dihedral and sweep back can be seen in almost all of the heavy aircrafts. The dihedral

improves the lateral stability of the aircraft and the sweep back of the wings allows the

wing to have a critical mach number higher than that of a relatively straight wing.

Busemann, a German aerodynamicist, proposed the concept of swept wing. The main

reason for applying wing sweep is to increase the drag rise or drag divergence Mach

number and consequently, the critical Mach number. However, the wing sweep also

affects other aerodynamic parameters like the lift slope curve.

In swept wings, the pressure distribution is due to the effective velocity component

perpendicular to the wing and not the freestream velocity. Therefore, the components like

lift drag etc have to be calculated in terms of this effective velocity instead of the

freestream velocity. This gives us the following relationships,


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Me = Mscos

CPe = CPs /cos2

CLe = CLs /cos2

e= s/cos

(Z/C)e = (Z/C)s/ cos

Where e stands for effective and s stands for streamwise and is the sweep angle.

From the above relations we can see that sweep decreases the lift coefficient. Therefore,

an aircraft with sweep has to be careful during low speed regimes and would probably

require high lift device to take off and land satisfactorily. Another disadvantage of swept

wings is their characteristic tip stall behavior because of the outboard spanwise flow

causing the boundary layer to thicken as it approaches the tips.

There are several ways of preventing tip stall on swept wings. Most measures, such as

shark or dog teeth (local leading-edge extensions), saw cuts or leading edge

boundary layer fences create at high angles-of-attack a streamwise vortex such that the

boundary-layer cross flow on the inboard wing is swept inboard thus relieving the

boundary layer on the outer wing.

The oldest device for preventing tip stall is a vertical plate fitted on the wing upper

surface in a streamwise direction thus forming a physical barrier for the boundary-layer

cross flow, the full-chord fence.

A disadvantage of wing fences is the increase in drag. For this reason on modern aircraft

wing fences are only applied when at a late stage in the development or during flight

testing stalling characteristics are found to be unsatisfactory.

In swept back wings the wingtips are located behind the centre of gravity. Therefore, any

loss of lift in the wingtips causes the centre of pressure to move forwards. This in turn

will cause the airplane nose to come up and consequently the angle of attack will be


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increased that results in loss of control. Therefore, the useful lift coefficient actually

decreases with increasing sweep angle. This can be avoided by use of stall fences which

prevents outboard spanwise flow.

Therefore, keeping in mind the above restrictions, we are selecting a sweep angle of 250

Wing dihedral improves the lateral stability of the aircraft. Dihedral is the upward angle

of wing along the span against the horizon. Due to the dihedral effect, if the airplane goes

into a roll because of a gust or some aileron input, a restoring force will be generated

which tends to bring the aircraft back to the steady level position. Another stabilizing

effect is that though the airplane is steadied after gust or aileron movement but due to

inertia, the airplane continues to rotate. Under this condition, the up going wing has

decreased angle of attack and the down going wing has increased angle of attack that

generates a restoring force inducing counter rotation that brings the aircraft back to steady

level flight. This counter rotational force ceases after steady level condition is achieved.

However, every good thing has something bad associated with it. Too much dihedral

results in a characteristic movement called Dutch roll. Therefore, keeping in mind the

limitations, we are selecting a wing dihedral angle of =30

The contributions of both dihedral and sweepback of the wing are enhanced if the center

of lift of each wing is far out along the wingspan.

Along with these, we are adding a small winglet at the wingtips in order to reduce the

wingtip vortices.

For the horizontal tail, we are selecting a sweepback of 300

and a dihedral of 50. For the

vertical tail, we are selecting a sweepback of 450

Selection of control surfaces

The primary control surfaces aileron, elevator and rudder are essential for stability and

control. Besides the primary control surfaces, there are secondary and auxiliary control

surfaces. Some of them are slats, slots, flaps, spoilers, trim tabs, spring tabs etc.


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We have already selected a plain flap for our aircraft. Flaps are high lift devices to

increase the lift during takeoff and landing. Spoilers are used to increase the drag to slow

down for landing or to overcome over speeding. There are two types of spoilers. Flight

spoilers and ground spoilers. The ground spoilers are automatically deflected up after

touch down if they had been engaged by the pilot.

The tabs are auxiliary control surfaces that help in the movement of primary control

surfaces. The tabs are deflected opposite to the direction of movement of the control

surface so that the wind incident on it would produce a force that aids in the movement ofthe control surface in the intended direction.


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Finally, the control surfaces that we have selected for our aircraft are,

Aileron Elevator Rudder Plain flap Spoilers Trim tabs

Conclusion

Thus, the wing, tail and control surfaces were selected and the wing dimensions were

estimated.


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Ex No. 9 ESTIMATION OF WING LOADING 24:02:11

Aim

To estimate the value of wing loading for the airplane design based on stall velocity and

landing distance constraints.

Introduction

Wing loading plays an important role in aircraft performance. The related parameters

such as wingspan and chord calculations are based on the value of wing loading. The

wing loading for most of the airplanes is determined by the considerations of V stall and

landing distance. The Vmax of an airplane increases as W/S increases. We are considering

Vstall and landing distance as our primary constraints.

First, let us consider the constraint imposed by stall velocity. This can be given by the

relation,

The above equation can be rewritten as

From our requirements, the stalling velocity is not to Exceed 50 ms-1

W/S = 0.5*1.225*502*2.295

W/S = 3514.219 N/m2

W/S = 358.228 kg/m2


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This is the value of wing loading constrained by stall velocity.

Now, let us examine the constraints imposed by landing distance.

Total landing distance = approach distance + flare distance + ground roll

SLD = Sa + Sf+ Sg

For commercial airplanes,

Approach velocity Va = 1.3Vstall


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Touchdown velocity VTD = 1.15Vstall

Flare velocity occurs between approach and touchdown velocity. Therefore, it is

acceptable to consider flare velocity as an average of touchdown and approach velocities.

Flare velocity Vf= Va + VTD

Vf= 1.3Vstall + 1.15Vstall

Vf= 1.23Vstall

Vf= 1.23*50 = 61.5 ms-1

Flight path radius during flare is given by,

During landing velocity is flare velocity and the load factor has an approximate value of n

= 1.2

R = 61.5

2/ (0.2*9.81)

R = 1927.75 m

Flare height


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For most commercial airplanes, the approach angle a = 30

hf= 1927.75*(1-cos30)

Flare height = 2.642 m

Approach distance

The approach distance required to clear a 50 ft obstacle is

Sa = (15.24-2.642)/tan30

Sa = 240.386 m

Flare distance

Sf= 1927.75*sin30

Sf= 100.89 m

From our specifications, the total landing distance is not to exceed 1000 m

Sa + Sf+ Sg = 1000 m

Sg = 1000100.89240.386


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Sg = 658.724 mHowever, Sg can also be found out using the equation,

For commercial airplanes, the value of j is taken as 1.15. N is the time increment for free

roll immediately after touchdown and it is taken as 3 s. r is the coefficient of friction of

the runway. Assuming a concrete runway that has a friction coefficient of 0.4. We have,

Sg = 1.15*3*(2*W/S/(1.225*2.295))1/2

+ 1.152W/S/(9.81*2.295*1.225*0.4)

Sg = 2.91sqrt(W/S) + 0.1199 W/S

Solving this equation, we have,

W/S = 3965.57 N/m

2

W/S = 404.237 kg/m2

Clearly, if the wing loading is less than 404.237, the landing distance will be less than

1000 m. therefore, the wing loading obtained by the stall velocity as constraint is

considered as the wing loading of complete airplane as it is lesser than 404.237 kg/m2

Therefore the wing loading W/S = 358.228 kg/m2


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Result

Thus, the value of wing loading constrained by stall velocity and landing distance was

found out,

Wing loading constrained by stall velocity = 358.228 kg/m2

Wing loading constrained by landing distance = 404.237 kg/m2


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Ex No. 10 ESTIMATION OF 03:03:11

THRUST TO WEIGHT RATIO

AimTo estimate the value of thrust to weight ratio for the airplane design

Introduction

Like the wing loading, the thrust to weight ratio also plays an important role in affecting

the performance of the airplane. It is the ratio of instantaneous thrust to weight. It is used

as a figure of merit for quantitative comparison of engine or vehicle design.

The value of T/W ratio determines in part the take off distance, rate of climb and

maximum velocity. First, let us consider the take off distance that is taken as 800 m in our

requirements.

Total take off distance = ground roll + airborne distance

Sg = 1.21*358.228*9.81/ (9.81*1.225*1.935*(T/W))

Sg = 182.864/ (T/W)

Vstall = sqrt (2*3514.219/ (1.225*1.935))

Vstall = 54.45 ms-1


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Radius of path during lift off

Where velocity is equal to 1.15Vstall and the load factor of n = 1.19

R = 6.69*54.452/9.81

R = 2103.627 m

Included flight path angle

= cos-1

(1-15.24/2103.627)

OB = 6.9010

The airborne distance

Sa = 2103.627sin 6.9010

Sa = 252.757 m

Sg + Sa = 800

Sg = 800-252.757 = 547.243 m

182.664/(T/W) = 547.243


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T/W = 0.3341

This is the value of thrust required at a velocity of

V = 0.7VLO = 0.7(1.1Vstall)

V = 0.7(1.1*54.45) = 41.927

V = 41.927 ms-1

or 93.79 mph or 137.55 fts-1

At this velocity, the power required is,

PR= TV = (T/W)*W0*V

= 0.3341*16112.77*9.81*41.927

PR= 2.215*106

watt or 2.215 MW

Power = 2.215*106/746 = 3711.4611 hp

Result

The value of thrust to weight ratio was estimated.

Thrust to weight ratio = 0.3341

Power required = 2.215 MW = 3711.4611 hp


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Ex No. 11 POWERPLANT SELECTION 10:03:11

Aim

To select a powerplant that seems to meet the aircrafts requirements.

Introduction

Selection of powerplant is an important step in design process because one has to

compromise with weight of the engine, thrust provided by it, TSFC etc. different types of

powerplants are available,

Turbojet Turboprop Turbofan Ramjet

Turbojet aircraft was the earliest form of jet engine. It can be classified into two types

based upon the type o compressor used

Axial flow compressor Centrifugal flow compressor

Though the turbojet engine has certain advantages like easy construction, weight andsize, it also has disadvantages like low efficiency and disturbing noise of high dB.

Turboprop is a combination of the propeller and the jet engine. It has good advantage of

efficiency but has altitude and speed limitations due to the propeller effects. It is better

suited for medium altitude and medium speed cruise.

Turbofan is a modified form of jet engine. It has a comparatively large fan attached ahead

of the compressor. A turbofan engine has two types of thrust like the turboprop engine,

fan (propeller) thrust and jet thrust. There are two flows in a turbofan engine. The outer

flow or cold flow and the inner or core or hot flow. The ratio of mass flow rate of outer

flow to that of core or inner flow is known as the bypass ratio. Based on this parameter,

the turbofan engines can be classified as,


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Low bypass turbofan Medium bypass turbofan High bypass turbofan

Higher the bypass ratio, higher the propulsive efficiency. This is the reason why most

transport and business aircrafts utilize high bypass turbofan engines. The TSFC of a

turbofan engine is almost half of that of a conventional turbojet engine.

Keeping in mind the above data we have selected Rolls Royce AE3007 turbofan engines.

It is the most suited engine to meet our specifications.

ROLLS ROYCE AE3007 TURBOFAN ENGINE


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Specifications

Description Specification

Thrust 8900 lb or 39.6 kNBypass ratio 4.8

Fan diameter 38.5 in or 0.9779 m

Length 115.1 in or 2.923 m

Weight 1586 lb or 719.397 kg

Inlet mass flow 240-280 lbs-

108.86-127 kgs-1

Stages Fan ; 14 HPC ; 2 HPT ; 3LPT

Overall pressure ratio 18-20:1

Turbine inlet temperature 994 C

Conclusion

The powerplant with appropriate requirements was selected.


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Ex No. 12 PERFORMANCE CURVES 17:03:11

Aim

To draw the performance curves for the design.

Introduction

The performance parameters that are so far discussed are tabulated and the curves are

plotted to correlate the stability and performance control that will be discussed in our

future calculation.

The performance curves are

Cl vs. Cd

V vs. L/D

V vs. Treq

V vs. Tav

V vs. T

V vs. Preq

V vs. PavV vs. P

V vs. R/C

Estimation of CL value

CL= 2W/(*S*V2)

2*16112.77*9.81/(1.225*44.9791*502) = 2.295

Estimation of CD value

CD = CD,0 + KCL2

where CD,0 0.02 and K = 1/(AR*e*) = 1/(9*0.8*) = 0.0442

0.02 + 0.0442*2.2952 = 0.2529


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Vinf Cl Cd L/D

50 2.295 0.2529 9.076

60 1.5937 0.1323 12.05

75 1.02 0.066 15.46

100 0.5737 0.0346 16.6

125 0.3672 0.026 14.14

150 0.255 0.0229 11.15

175 0.1873 0.0216 8.693

200 0.1434 0.0209 6.86

225 0.1133 0.0206 5.51

250 0.0918 0.0204 4.506

0

0.05

0.1

0.15

0.2

0.25

0.3

0 0.5 1 1.5 2 2.5

Drag Polar Cl vs Cd

Cd


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Estimation of Treq

Treq = 0.5**V2*S*CD = 0.5*1.225*25

2*44.9791*0.2529

17415.18 N

Estimation of Preq

Preq = Treq*V

17415.18*50 = 870759.242 WVinf Treq Preq

50 17415.18 870759.242

60 13120.87 787252.194

75 10227.21 767040.58

100 9519.365 951936.472

125 11175.31 1396914.11

150 14179.33 2126899.66

175 18183.39 3182093.33

200 23042.12 4608423.05

225 28686.06 6454362.41

250 35078.63 8769657.86

0

2

4

6

8

10

12

14

16

18

0 50 100 150 200 250 300

V vs L/D

L/D


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Estimation of Tav

Tav = 72000 N

Estimation of Pav

Pav = Tav*V = 72000*50 = 3600000 W = 3.6 MW

0

5000

10000

15000

20000

25000

30000

35000

40000

0 50 100 150 200 250 300

V vs Treq

Treq

0

2000000

4000000

6000000

8000000

10000000

12000000

14000000

16000000

0 50 100 150 200 250 300 350

V vs Preq

Preq


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Vinf Tav Pav

50 72000 3600000

60 72000 4320000

75 72000 5400000

100 72000 7200000

125 72000 9000000

150 72000 10800000

175 72000 12600000

200 72000 14400000

225 72000 16200000

250 72000 18000000

0

10000

20000

30000

40000

50000

60000

70000

80000

0 50 100 150 200 250 300

V vs Tav

Tav

0

2000000

4000000

6000000

8000000

10000000

12000000

14000000

16000000

18000000

20000000

0 50 100 150 200 250 300

V vs Pav

Pav


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V vs T

V vs P

Rate of climb

R/C = (Pav-Preq)/W0

(3600000-870759.242)/(16112.77*9.81) = 17.266

0

20000

40000

60000

80000

100000

120000

0 50 100 150 200 250 300 350

Treq

Tav

0

5000000

10000000

15000000

20000000

25000000

0 50 100 150 200 250 300 350

Preq

Pav


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Conclusion

Thus, the performance curves were plotted.

-10

0

10

20

30

40

50

60

70

0 50 100 150 200 250 300 350

V vs R/C

R/C

Vinf Preq Pav R/C

50 870759.242 3600000 17.26643322

60 787252.194 4320000 22.34978862

75 767040.58 5400000 29.31023369

100 951936.472 7200000 39.52812565

125 1396914.11 9000000 48.10062081

150 2126899.66 10800000 54.87002466

175 3182093.33 12600000 59.58201232

200 4608423.05 14400000 61.94602253

225 6454362.41 16200000 61.65538897

250 8769657.86 18000000 58.3953928


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Ex No. 13 THREE VIEW DIAGRAM 24:03:11

Aim

To draw the three view diagram of the design aircraft.

The three view diagram is nothing but the result or outcome of the conceptual design

process. The configuration or layout helps in proceeding to the next level in design

process i.e. the preliminary design and then over to the detailed design.

Conclusion

This three-view diagram is the outcome of the conceptual design process.


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CONCLUSION

The design of the selected aircraft long range business jet was done in a step by step

method sticking to the basic rules. Calculations were performed with respect to data andformulae obtained from available design books. Some problems aroused during the

process. However, they were solved by good teamwork. The conceptual design done in

this project meets the initial requirements that were set by us.. It also helped us to

understand some basic things about the aircraft and its design. We saw how the weight of

an aircraft plays an important role in the design process. All the other following

parameters vary with weight. The estimation of weight was a very crucial step and was

also interesting. Then followed by the selection of an appropriate airfoil and estimating

the lift that could be obtained using it. Then the estimation of some critical performance

parameters to finalize the conceptual design process and selection of an appropriate

powerplant. Finally, it ended with the three view diagram. We would like to conclude

saying that the experience that we got during the process will be helpful and the moments

will always be remembered by us.


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REFERENCES

Books

Aircraft Performance and Design by John D. Anderson Jr, Tata McGraw HillEdition 2010

Aircraft Design, a conceptual Approach by Daniel P Raymer, AIAA Educationseries 2

ndedition

Airplane Design by Dr. Jan Roskam, Roskam Aviation and EngineeringCorporation, 1985

Design of the Aeroplane by Darrol Stinton, BSP Professional Books Introduction to Flight by John D. Anderson Jr, Tata McGraw Hill Edition 2009 Airplane Aerodynamics and Performance by Dr. Jan Roskam & Dr. Chuan Tau

Edward Lan, DAR Corporation 1997

Theory of Wing Sections by Ira H. Abott & Albert E. Von Doenhoff, DoverPublications 1959

Internet

Wikipedia Google Rolls Royce

Documents

Aircraft Design Project i