Aircraft design project - 1

AIRCRAFT DESIGN PROJECT

AIRCRAFT DESIGN PROJECT - 1 Department of Aerospace Engineering, SRM University, Kattankulathur, Chennai.

NAME: P H Abhishek Aerospace – A REG. NO: 1191110068

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AIRCRAFT DESIGN PROJECT

Acknowledgement

Foremost, I would like to thank Mr. Abdur Rasheed and Mr. Mohammed Ariff for their knowledge and guidance in completing this project. I am also thankful to our Head of the Department Mr. Vasudevan for his knowledge, insight and experience in diverse fields. Besides, I‘m also thankful to the rest of the faculty members of the department for their expertise and knowledge in various subjects. Finally, I would like to thank my friends for their continuous motivation and support.

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AIRCRAFT DESIGN PROJECT

Abbreviations = density at cruise altitude = cruising velocity = Reynolds number at cruising altitude = viscosity at cruising altitude = Aspect ratio of the wing = span of the wing = chord of the wing = viscosity at sea level = sweep angle = Taper ratio = Root chord = Tip chord = Mean chord = Thrust required = Thrust at cruising altitude = density at the specified altitude = density at sea level = Take – off velocity = Landing velocity = Stalling velocity = Gross weight of the airplane = Take – off distance = Landing distance = maximum rate of climb

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AIRCRAFT DESIGN PROJECT

Aim of the project The aim of this design project is to design a cargo aircraft capable of carrying a payload of 20,000 kgs. It covers both the design and the performance characteristics of the aircraft. The following design requirements and research studies are set for the project:    

To operate on an international scale. To use advanced and state of the art technologies to reduce operating costs. To offer a unique and competitive service to existing scheduled operations. To produce a commercial analysis of the aircraft project.

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AIRCRAFT DESIGN PROJECT

Abstract This project‘s motive is to develop a functional 20,000 kg payload cargo aircraft. The aircraft has a gross weight of 68554.56 kgs. It uses a NACA 2414 airfoil and has a high wing with slotted flaps. The engine selected has a thrust of 226 KN with one engine on either side. It is of the turbofan type such that it produces the adequate speed, range and fuel economy. The aircraft has a conventional type tail.

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AIRCRAFT DESIGN PROJECT

Basic design process

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AIRCRAFT DESIGN PROJECT

Introduction: An airplane design is both an art and a science. It is an intellectual engineering process of creating on paper a flying machine to  Meet specifications established by users.  Pioneer innovative, new ideas and technology. The design process is an intellectual activity developed via experience, by attention paid to successful airplane designs that have been used in the past and by design procedures and databases that are a part of every airplane manufacturer. Phases of airplane design: From time when an airplane materializes as a new thought to the time the finished product is ready, the complete design undergoes 3 distinct phases in perfect sequences which are:  Conceptual Design  Preliminary Design  Detail Design Conceptual Design: The design process starts with a set of specifications or much less frequently to desire to implement pioneering. There is a concrete goal where we designers are aiming at. The first step towards it is conceptual design. Within a fuzzy latitude, overall shape, size, weight are determined for the potential user. The product of the conceptual design phase is layout of the airplane configuration on paper. This drawing has flexible lines, which can be slightly changed. However when we get a detailed account of layout configuration at the end of this phase. The major drivers during conceptual design process are aerodynamics, propulsion and flight performance. Structural and control system considerations are not dealt in detail but however they are not totally absent. The designer is influenced by qualitative aspects. No part of the design process is carried out in total vacuum unrelated to other parts. Preliminary design: This phase includes only minor changes to be made in the configuration layout. There is serious control and structural system analysis and design takes place. During this phase substantial wind tunnel testing will be carried out and major computational fluid dynamics (CFD) calculation. At the end of this phase, the airplane configuration is frozen and defined. The drawing process is called lofting. This process makes precise shape of outside skin of airplane making certain all sections fit together.

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AIRCRAFT DESIGN PROJECT

The end of the phase is the decision if the plane is to be manufactured or not. It is no longer a critical condition where ―you- bet your company‖ on fuel scale development of a new airplane.

Detail Design: This phase is literally the ‗nuts and bolts‘ phase of the airplane design. The aerodynamic, propulsion, structures, performance and flight control analysis are over in the preliminary phase. The airplane is to be fabricated and machined. The size, number and location of rivets, fasteners are determined mow. Flight simulators are developed. At the end of this phase, the aircraft is ready to be fabricated. The seven intellectual pivot points for conceptual design The overall conceptual design is anchored seven intellectual ―pivot points‖ – seven factors that anchor the conceptual design thought process. They allow different, detailed thinking to reach out in all directions from each point. Requirements: The requirements are given by the people who are going to buy – the customers. For other aircrafts, these requirements are usually set by the manufacturer in full appreciation of needs of the owner. Requirements of one airplane are different form the other. There can be no stipulated specific standard. There must be established requirements that serve as impinge off point for design processes. The requirements that are frequently stipulated are:      

Range Take-off distance Stalling velocity Endurance Maximum velocity Rate of climb.

For dog fighting combat, maximum turn rate and minimum turn radius are required. The others are:     

Maximum load factor Service ceiling Cost Reliability and Maintainability Maximum size

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AIRCRAFT DESIGN PROJECT

Critical performance parameters: Requirements stipulate the performance of the new aircraft. The critical parameters are:  Maximum lift coefficient  Lift to drag ratio ( L/D )  Thrust to weight ratio ( T/W ) Therefore the next step is to make first estimates of W/S and T/W to achieve the performance as stipulated by the requirements.

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AIRCRAFT DESIGN PROJECT

Fig 1.1

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Configuration layout: The configuration layout is a drawing of the shape and size of the plane as evolved till stage. The critical performance parameters along with first estimate helps to draw the configuration and approximate the size of the aircraft. Better Weight Estimate: The overall size and shape of the airplane are better known now. There is now and improved estimate of weight based performance parameters. A more detailed estimate of fuel is required now.

Performance analysis: This is the point where the configuration is judged if it can meet all original specifications. An interactive process is initiated where the configuration in modified. The critical performance parameters are just adjusted for improving performance. In this stage, some mature decisions should be made as the specifications or cost or unavailable technology. Hence some specifications might be relaxed so that others might get higher priority. Optimization: When iterative process is over, it has produced a viable airplane. This leads to optimization. The optimization analysis is carried out, may be carried out by a systematic variation of different parameters T/W, W/S and plotting the performance of graphs which can be found using a sizing matrix or a carpet plot form which optimum design can be found.

Weight of the airplane – First estimate: No airplane can take off the ground unless it produces a lift greater than its weight. There should be a first estimate of gross take-off weight. The weight estimate is the next pivot point after the requirements. Lilienthal Langley and wright brothers knew more weight means more drag. This needed an engine with greater power and hence more weight. Constraint diagram: A constraint diagram is constructed which identifies allowable solution space for airplane design. A constraint diagram consists of plots of the sea – level thrust to take-off weight ratio versus wing loading at take-off weight ratio T0/W0 versus wing loading at takeoff W0/s determined by intellectual pivot point.

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AIRCRAFT DESIGN PROJECT

The design wheel

Fig 1.2

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Literature Survey

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The following tables include the geometric, performance, engine and weight parameters of a list of cargo aircrafts with a payload of 20,000 kg. GEOMETRIC PARAMETERS: S.No. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

14.

15. 16.

Name of the aircraft Antonov An – 12 Lockheed L – 100 Shaanxi Y – 8 Lockheed C – 130 Shaanxi Y – 9 Tupolev Tu – 204 Embraer KC – 390 Transall C – 160 UAC/HAL Il – 214 Airbus A310 MRTT Lockheed KC – 130 Boeing C – 40 Aero Spacelines Super Guppy Aero Spacelines Mini Guppy Boeing 727 – 200F Boeing 737 – 200F

Aspect Ratio 11.86:1

Wing Span (m) 38

Length (m)

Height (m)

33.1

10.53

Wing Area (m2) 121.7

10.06:1

40.4

34.35

11.66

162.1

11.84:1 10.06:1

38 40.4

34.02 29.8

11.16 11.6

121.9 162.1

9.48:1

40 41.8

36 46.14

11.3 13.9

184.2

35.06

33.91

10.26

40

32.4

11.65

30.1

33.2

10

43.9

47.4

15.8

10.07:1

40.41

29.79

11.84

162.1

12.42:1

34.32 47.625

33.63 43.84

12.55 14.78

182.51

11.53:1

43.1

33.7

11.7

161.1

7.08:1

32.92

40.6

10.36

153

7.87:1

28.35

28.65

11.23

102

10:1

160

Table 1.1

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AIRCRAFT DESIGN PROJECT

PERFORMANCE PARAMETERS: S.No.

Name of the aircraft

Range (Km)

Service Ceiling (m)

1.

Antonov An – 12 Lockheed L – 100 Shaanxi Y – 8 Lockheed C – 130 Shaanxi Y – 9 Tupolev Tu – 204 Embraer KC – 390 Transall C – 160 UAC/HAL Il – 214 Airbus A310 MRTT Lockheed KC – 130 Boeing C – 40 Aero Spacelines Super Guppy Aero Spacelines Mini Guppy Boeing 727 – 200F Boeing 737 – 200F

3600

2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

14.

15. 16.

10200

Max. Speed (km/hr ) 777

Cruising speed (km/hr ) 670

Rate of climb (m/min) 597.4

2470

7000

570

541

557.78

5615 3800

10400 10060

660 592

550 540

609.6 557.78

7800 4300

10400 12600

900

650 850

4815

10973

850

1853

8230

513

335

2500

12000

870

830

8889

396.24

978

5250

8615

671

643

5600 3219

12500 7620

990 463

407

6920

10670

603

482

4300

11000

1072.44

864

2850

10700

876

780

896.112

Table 1.2

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AIRCRAFT DESIGN PROJECT

ENGINE PARAMETERS: S.No. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

Name of the aircraft Antonov An – 12 Lockheed L – 100 Shaanxi Y – 8 Lockheed C – 130 Shaanxi Y – 9 Tupolev Tu – 204 Embraer KC – 390 Transall C – 160 UAC/HAL Il – 214 Airbus A310 MRTT Lockheed KC – 130

12.

Boeing C – 40

13.

Aero Spacelines Super Guppy Aero Spacelines Mini Guppy Boeing 727 – 200F

14. 15.

16.

Boeing 737 – 200F

Name of the engine Progress Al – 20 Allison 501 – D22A Zhuzhou WoJiang – 6 Allison T56A – 15 Zhuzhou WoJiang – 6C Aviadviagetel PS90 – A IAE – V2500 – E5 Rolls Royce Type MK22 Aviadviagetel PS90 – A – 76 GE – CF 6

Type of engine Turboprop Turboprop

No of engines 4 4

Thrust ( KN )

Turboprop

4

Turboprop

4

Turboprop

4

Turbofan

2

157

Turbofan

2

129

Turboprop

2

Turbofan

2

140

Turbofan

2

262

Dowty propellers Rolls Royce AE 2100 CFM international CM 56 – 7 Allison 501 – D22C Pratt and Whitney R4360 Pratt and Whitney JT8D – 1 Pratt and Whitney JT8D – 1

Turboprop

4

Turbofan

2

Turboprop

4

Radial

4

Turbofan

3

62

Turbofan

2

64

121

Table 1.3

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AIRCRAFT DESIGN PROJECT

WEIGHT PARAMETERS: S.No. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16.

Name of the aircraft Antonov An – 12 Lockheed L – 100 Shaanxi Y – 8 Lockheed C – 130 Shaanxi Y – 9 Tupolev Tu – 204 Embraer KC – 390 Transall C – 160 UAC/HAL Il – 214 Airbus A310 MRTT Lockheed KC – 130 Boeing C – 40 Aero Spacelines Super Guppy Aero Spacelines Mini Guppy Boeing 727 – 200F Boeing 737 – 200F

Empty Weight (Kg) 28000 35260

Max. Takeoff Weight (Kg) 61000 70300

Payload (Kg)

35490 34400

61000 70300

20000 20400

39000

77000 103000 81000

25000 21000 23600

51000 68000

16000 20000

113999

163998

28000

34274

79378

19090

57150 46039

78000 77110

18000 24720

37410

79370

17010

36560

77000

23500

28876

28100

50300

18140

17747.12

74000 29000

Fuel Weight (kg)

20000 23150

13500

Table 1.4

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AIRCRAFT DESIGN PROJECT

Cruising speed vs Range 10000 9000 8000 7000

Range

6000 5000 4000 3000 2000 1000 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.1

Cruising speed vs Length 50 45 40 35

Length

30 25 20 15 10 5 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.2 18

AIRCRAFT DESIGN PROJECT

Cruising speed vs Height 18 16 14

Height

12 10 8 6 4 2 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.3

Cruising speed vs Max. Take - off weight 180000 160000 140000 120000

Max. Takeoff weight

100000 80000 60000 40000 20000 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.4 19

AIRCRAFT DESIGN PROJECT

Cruising speed vs Wing loading 600

500

Wing loading

400

300

200

100

0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.5

Cruising speed vs Wing area 200 180 160

Wing area

140 120 100 80 60 40 20 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.6 20

AIRCRAFT DESIGN PROJECT

Cruising speed vs Service ceiling 14000

12000

Service ceiling

10000

8000

6000

4000

2000

0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.7

Cruising speed vs Span / Length 1.6 1.4

Span / length

1.2 1 0.8 0.6 0.4 0.2 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.8 21

AIRCRAFT DESIGN PROJECT

Cruising speed vs Payload 30000

25000

Payload

20000

15000

10000

5000

0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.9

Cruising speed vs Wingspan 60

50

Wingspan

40

30

20

10

0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.10

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AIRCRAFT DESIGN PROJECT

Cruising speed vs Rate of climb 1000 900 800

Rate of climb

700 600 500 400 300 200 100 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 1.11 Design Values:

Parameter Range Aspect ratio Rate of climb Max. Take-off Weight Service Ceiling Span/Length Length Height Wing Loading

Avg. Value 4400 km 10.52:1 770 m/min 70,000 kgs 11,000 m 1.02 36 m 11.52 m 400 Table 1.5

Conclusion: Thus the literature survey was performed on various aircrafts and the design values were obtained.

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AIRCRAFT DESIGN PROJECT

Weight Estimation

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AIRCRAFT DESIGN PROJECT

The second pivot point in our conceptual design analysis is the preliminary estimation of the gross weight of the airplane. The various components of an airplane are:  Crew weight (

):

The crew comprises the people necessary to operate the airplane in flight.  Payload weight (

):

The payload is what the airplane is intended to transport – passengers, baggage, freight etc. If the airplane is intended for military use, the payload includes bombs, rockets and other disposable ordnance.  Fuel weight(

):

This is the weight of the fuel in the fuel tanks. Since fuel is consumed during flight, is a variable, decreasing with time during flight.

):

 Empty weight (

This is the weight of everything else – the structure, engines, electronic equipment, landing gear, fixed equipment etc. The sum of these weights is the total weight of the airplane W. The design take off gross weight mission.

is the weight of the airplane the instant it begins its

Rewriting the above equation,

Or

Solving for

(

,

)

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AIRCRAFT DESIGN PROJECT

Estimation of

:

The historical, statistical data on previous airplanes provide a starting point for the conceptual design of a new airplane. By drawing a graph between

and

, we get the

required value. Estimation of

:

The amount of fuel required to carry out the mission depends critically on the efficiency of the propulsion device – the engine specific fuel consumption and propeller efficiency. It depends critically on the aerodynamic efficiency – the lift – to – drag ratio. By the Brequet Range equation, we have

The mission segment weight fraction is given by, Mission segment weight fraction = The mission profile, a conceptual sketch of altitude versus time is shown,

Fig 2.1 From the mission profile, the ratio of the airplane weight at the end of the mission to the initial gross weight is W5/W0. In turn,

If, at the end of the flight, the fuel tanks were completely empty, then

Having a 6% allowance, we get, 26

AIRCRAFT DESIGN PROJECT

(

)

(from table 3.1 – Airplane Design by Raymer, Ref. Annex. Table 7)

Calculation:

We / W0 vs W0 1 0.9 0.8 0.7

We / W0

0.6 0.5 0.4 0.3 0.2 0.1 0 0

20000

40000

60000

80000

W0

100000

120000

140000

160000

180000

Graph 2.1

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AIRCRAFT DESIGN PROJECT

(from table 2.1 – Airplane Design by Dr. Ian Roskam, Ref. Annex. Table 4)

For cruise,

(from table 2.1 – Airplane Design by Dr. Ian Roskam, Ref. Annex. Table 5)

For loiter,

(from table 2.1 –Airplane Design by Dr. Ian Roskam, Ref. Annex. Table 5)

Therefore,

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AIRCRAFT DESIGN PROJECT

Conclusion: Thus the weight estimation for the cargo aircraft with a payload of 20,000 kg was done and the gross weight was found to be

Also,

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Selection of airfoil

Introduction:

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Before the design layout can be started, values for a number of parameters which includes the airfoils must be chosen. The airfoil of a plane is chosen such that there is maximum lift and minimum drag. An airfoil is the shape of a wing as seen in cross section. An airfoil-shaped body moved through a fluid produces an aerodynamic force. The component of this force perpendicular to the direction of motion is called lift. The component parallel to the direction of motion is called drag. Subsonic flight airfoils have a characteristic shape with a rounded leading edge, followed by a sharp trailing edge, often with asymmetric camber. Importance of airfoil in an aircraft: The airfoil, in many respects, is the heart of the airplane. The shape of the airfoil affects the following factors:     

Cruising speed Takeoff and landing distance Stall speed Handling qualities (especially near the stall), and Overall aerodynamic efficiency during all phases of fight.

The P-51 was regarded as the finest fighter of world war 2 in part because of its radial laminar - flow airfoil. Schemes have been devised to define airfoils — an example is the NACA system. Various airfoil generation systems are also used. An example of a general purpose airfoil that finds wide application, and predates the NACA system, is the Clark-Y. Today, airfoils can be designed for specific functions using inverse design programs such as PROFOIL, XFOIL and Aerofoil. The NACA four-digit wing sections define the profile by: First digit describing maximum camber as percentage of the chord. Second digit describing the distance of maximum camber from the airfoil leading edge in tens of percents of the chord. Last two digits describing maximum thickness of the airfoil as percent of the chord. Next developed, was the five – digit series. In NACA 23012, the first digit when multiplied by 3/2, gives the design lift coefficient in tenths. The design lift coefficient is an index of the amount of camber. The second and third digits together are a number which, when multiplied by ½ gives the location of the maximum camber. The last two digits give the maximum thickness in percentage of the chord. The NACA 6 – digit series included the NACA 64 – 212, in which the first number simply gives the series designation. The second digit is the location of minimum pressure. The third digit gives the design lift coefficient in tenths. The last two digits, as usual, give the maximum thickness in percentage of the chord.

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AIRCRAFT DESIGN PROJECT

Fig 3.1 Airfoil nomenclature

Fig 3.2  The front of the airfoil is defined by a leading - edge radius which is tangent to the upper and lower surfaces.  An airfoil designed to operate in supersonic flow will have a sharp or nearly - sharp leading - edge to prevent a drag producing bow shock.  The chord of the airfoil is the straight line from the leading edge to the trailing edge.  The camber refers to the curvature characteristic of most airfoils. The mean camber line is the line equidistant from the upper and lower surfaces. Total airfoil camber is defined as the maximum distance of the mean camber line to the chord line.

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AIRCRAFT DESIGN PROJECT

Calculation: The Reynolds number at cruise is given by,

From graph, we know that,

Where

Service ceiling = 11,000 m From gas tables,

We know that,

(

)

(

)

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AIRCRAFT DESIGN PROJECT

S.No

Name of the airfoil

1 2 3 4 5 6 7

NACA 66 – 009 NACA 63 – 015 NACA 2414 NACA 2408 EPPLER 520 AIRFOIL LWK 80 – 150/K25 B737C – IL

1.01 0.85 0.8 1 0.75 0.65 0.9

0.06 0.018 0.01 0.015 0.01 0.015 0.012

66.2 47.22 80 66.67 75 43.3 75 Table 3.1

From the above data we can see that NACA 2414 has the maximum Cl and the minimum Cd for the reynolds number

.

Conclusion: The

airfoil is the best suited airfoil for the calculated reynolds number.

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AIRCRAFT DESIGN PROJECT

NACA 2414

Fig 3.3 Coordinates 1.00000 0.00147 0.99739 0.00210 0.98929 0.00396 0.97587 0.00700 0.95729 0.01112 0.93372 0.01620 0.90542 0.02207 0.87267 0.02857 0.83582 0.03552 0.79527 0.04274 0.75143 0.05004 0.70480 0.05723 0.65586 0.06412 0.60515 0.07053 0.55324 0.07629 0.50069 0.08120 0.44808 0.08512 0.39598 0.08787 0.34454 0.08913 0.29482 0.08866 0.24740 0.08645 0.20285 0.08255 0.16169 0.07707 0.12440 0.07014 0.09141 0.06198 0.06310 0.05281 0.03977 0.04289 0.02165 0.03245 0.00892 0.02171 0.00169 0.01085 0.00000 0.00000

0.00379 -0.01031 0.01293 -0.01956 0.02730 -0.02770 0.04669 -0.03471 0.07087 -0.04054 0.09957 -0.04516 0.13246 -0.04858 0.16918 -0.05082 0.20937 -0.05195 0.25260 -0.05208 0.29844 -0.05133 0.34644 -0.04987 0.39611 -0.04787 0.44739 -0.04537 0.49931 -0.04232 0.55129 -0.03886 0.60276 -0.03516 0.65316 -0.03132 0.70194 -0.02745 0.74857 -0.02365 0.79252 -0.01998 0.83331 -0.01650 0.87048 -0.01328 0.90360 -0.01035 0.93230 -0.00776 0.95626 -0.00557 0.97518 -0.00381 0.98886 -0.00252 1.00000 -0.00147

Table 3.1

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AIRCRAFT DESIGN PROJECT

Wing parameters

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AIRCRAFT DESIGN PROJECT

Sweep angle

:

The angle between the lateral arms and the quarter-chord line. It is also referred to as the leading-edge sweep. Referring to fig 4.2 in A conceptual approach by Raymer the quarter – Chord sweep (in degrees) = 7

Fig 4.1

Fig 4.2 37

AIRCRAFT DESIGN PROJECT

Taper ratio

:

It is the ratio between the tip chord and the centre-line root chord. Most wings of low sweep have a taper ratio of 0.4-0.6. Most swept wings have taper ratios of 0.2 to 0.4. For sweep angle= 7

, taper ratio= 0.35 (approx.)

Chord of the wing: The chord decreases along the length of the wing from wing root to wing tip. The chord at the root of the wing is called root chord and the chord at the wing tip, the tip chord. Root chord:

=

Position of wing:

Fig 4.3 High wing: It is mounted on the upper fuselage when contrasted to shoulder wing applies to awing mounted on a projector (carbon roof) above the top of the main fuselage. Advantages: 

Short landing distance



More ground clearance for engines



Easier passenger and cargo loading



Better view

Disadvantages: 

Poor visibility towards the top and rear of the aircraft. 38

AIRCRAFT DESIGN PROJECT



Placement of landing gear



Has more frontal area which increases drag

Applications: 

Antonov An-140



ATR-42

Mid-wing:

Fig 4.4 It is mounted approximately half-way up the fuselage. Advantages: 

Structurally efficient



Capable of performing extreme maneuvers and aerobatics



Space for armaments and cargo available



Less interference drag

Disadvantages: 

More expensive



Reinforcement at wing root not possible



Used in less aircrafts

Applications: 

Milla JM-2



Sawya Skyjacker-II

Low wing:

Fig 4.5 39

AIRCRAFT DESIGN PROJECT

It is mounted near the bottom of the fuselage. Advantages: 

Easy maintenance



Better visibility



Dihedral configuration makes it stable

Disadvantages: 

More interference drag



Susceptible to ground effects



More speed required at approach

Applications: 

Airbus A380

Conclusion: The following are the wing parameters chosen for a 20000 kg cargo aircraft.

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AIRCRAFT DESIGN PROJECT

Engine selection

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AIRCRAFT DESIGN PROJECT

To select the required engine, a graph between

and

is plotted.

Cruising speed vs T / Wtakeoff 8 7

T / Wtakeoff

6 5 4 3 2 1 0 0

100

200

300

400

500

Cruising speed

600

700

800

900

1000

Graph 5.1 From graph,

Now

The closest value we get from literature survey is chosen as the best engine for this aircraft. Name of the engine

CFM International CM 56- 7

No of the engines

2

Thrust( per engine) Kn By calculation From literature survey 113.115 121

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AIRCRAFT DESIGN PROJECT

Table 5.1

Fig 5.1 Engine specifications: General characteristics:    

Type: Twin-spool, high-bypass turbofan Length: 2.5 m Diameter: 1.55 m (fan) Dry weight: 2,366 kg (dry)

Components:   

Compressor: Single-stage fan, 3-stage low-pressure compressor, 9-stage high pressure compressor Combustors: annular Turbine: Single-stage high-pressure turbine, 4-stage low-pressure turbine

Performance:    

Maximum thrust: 121 kN Overall pressure ratio: 32.8:1 Bypass ratio: 5.5:1 Thrust-to-weight ratio: 3.7:1

Conclusion:

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The engine chosen for this aircraft from the literature survey was the which has a thrust of

Fuel Validation

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We know that the fuel weight can be calculated using the following formula,

Where

= ,

Substituting in

Therefore,

= =

Conclusion:

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The weight of fuel at cruise by calculation, for a 20000kg cargo was found to be From report 2,

=>

Flap selection, Lift distribution, Lift and drag estimation

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Lift is generated in accordance with the fundamental principles of physics. The most relevant physics reduce to three principles:  Conservation of Momentum, which is a direct consequence of Newton's laws of motion, especially Newton's second law which relates the net force on an element of air to its rate of momentum change,  Conservation of Mass, including the common assumption that the airfoil's surface is impermeable for the air flowing around, and  Conservation of Energy, which says that energy is neither created nor destroyed.

Fig 6.1 In addition, one needs an expression relating the fluid stresses (consisting of pressure and shear stress components) to the properties of the flow. The pressure depends on the other flow properties, such as its mass density, through the (thermodynamic) equation of state, while the shear stresses are related to the flow through the air's viscosity. The Prandtl lifting-line theory, is a mathematical model for predicting the lift distribution over a three-dimensional wing based on its geometry. The lifting-line theory makes use of the concept of circulation and of the Kutta–Joukowski theorem,

Where T is the circulation over the entire wing (m²/s) so that instead of the lift distribution function, the unknown effectively becomes the distribution of circulation over the span.

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Flap selection: When in cases where the is not enough for safe takeoff, we need to augment the lift produced and therefore, we find the use of ―high lift devices‖. They can be plain flap, split flap, leading edge slot, single slotted flap, double slotted flap and combinations each. A plain flap has a part of the airfoil is hinged to move, the split flap has a separate thinner flap selection along the rear end of the airfoil that is hinged and moves. A leading edge slat is a highly cambered airfoil kept at the tip of the airfoil (leading edge). A slotted flap is a plain flap with a slot in between the hinged portion and the airfoil.

Fig 6.2

Lift

:

Lift is the component of this force that is perpendicular to the oncoming flow direction. Drag

: It is the component of the surface force parallel to the flow direction.

Lift coefficient

:

The lift coefficient is a dimensionless coefficient that relates the lift generated by a lifting body to the density of the fluid around the body, its velocity and an associated reference area. 48

AIRCRAFT DESIGN PROJECT

Drag coefficient

:

The drag coefficient is a dimensionless quantity that is used to quantify the drag or resistance of an object in a fluid environment such as air or water.

Types of drag: Parasitic drag: Parasitic drag is drag that results when an object is moved through a fluid medium (in the case of aerodynamics, a gaseous medium, more specifically, the atmosphere). Parasitic drag is a combination of form drag, skin friction and interference drag. Form drag: Form drag or pressure drag arises because of the shape of the object. Skin friction drag: Skin friction arises from the friction of the fluid against the "skin" of the object that is moving through it. Interference drag: Interference drag results when airflow around one part of an object (such as a fuselage) must occupy the same space as the airflow around another part (such as a wing). Wave drag: Wave drag is a component of the drag on aircraft, blade tips and projectiles moving at transonic and supersonic speeds, due to the presence of shock waves.

Lift estimation: At ground,

At takeoff,

At landing, 49

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The values of

correspond to the slotted flap which are,

= =

Drag estimation:

(

Drag equation,

)

Calculation:

(

(

)

)

At takeoff,

Where

Therefore,

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At landing,

Drag estimation: At cruise,

For a jet transport the values of

(

and

are given as,

)

(

) (

)

(

)

At takeoff,

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At landing,

(

)

Result:

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Sizing of aircraft

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Fuselage sizing: Fuselage length From table 6.3 as seen in Aircraft design by Raymer, the values of a and c for a cargo aircraft with a payload of 20,000 kg was found to be a = 0.23 and c = 0.5 (Ref. Annex. Table 2) Therefore,

Tail sizing: Vertical tail: Vertical tail area,

Where Tail arm of vertical tail

Wing mean chord,

(from report 4)

Wing span, Wing surface area, Vertical tail volume coefficient, (from table 6.4 – Aircraft design by Raymer, Ref. Annex. Table 1)

Horizontal tail: Horizontal tail area,

Where 54

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Tail arm of horizontal tail = Horizontal tail volume coefficient, (from table 6.4 – Aircraft design by Raymer, Ref. Annex. Table 1)

Tail configuration: It may be characterized by  No. of tailplanes – from 0 (tailless or canard ) to 3 ( Roe triplane)  Location of tailplane – mounted high, mid, or low on the fuselag.  Fixed stabilizer and moveable elevator surfaces, or a single combined stabilizer. The major difference between a tail and a wing is that, the wing is designed to carry a substantial amount of lift, a tail is designed to operate normally at only a fraction of its lift potential. Tails provide trim, stability and control. Trim refers to the generation of a lift force that by acting through some moment about the center of gravity, balances some other moment produced by the aircraft. Conventional tails:

Fig 7.1  Provides adequate stability and control at lightest weight.  More than 70% of the aircrafts in service have this arrangement.

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T – tail:

Fig 7.2  Heavier than a conventional tail because the vertical must be strengthened in order to support the tail.  Vertical tail is smaller.  More efficient because it allows size reduction, thus reducing fatigue for the structure. V – Tail:

Fig 7.3  Intended to reduce wetted area of tail  Reduced interference drag but at the penality in control – actuation complexity. Y – Tail:

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Fig 7.4  Similar to v – tail, but a third suface is mounted vertically beneath the V.  Reduces interference drag when compared to the conventional tail.  It is used to primarily to keep the horizontal surfaces out of the wing wake at high angles of attack. Tail configuration used: The most suitable tail for a 20000 kg cargo aircraft is the conventional tail. It provides the following advantages:  Less weight.  Adequate stability.  Some of the aircrafts (from literature survey) had the same type of tail.

Conclusion: 57

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The following were found from this report.

Landing gear

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The landing gear is the structure that supports an aircraft on the ground and allows it to taxi, take-off, and land. In fact, landing gear design tends to have several interferences with the aircraft structural design.

Components: The landing gear usually includes wheels, but some aircraft are equipped with skis for snow or float for water. In the case of a vertical take-off and landing aircraft such as a helicopter, wheels may be replaced with skids. The figure below illustrates landing gear primary parameters. The descriptions of primary parameters are as follows. Landing gear height is the distance between the lowest point of the landing gear (i.e. bottom of the tire) and the attachment point to the aircraft. Since, landing gear may be attached to the fuselage or to the wing; the term height has different meaning. Furthermore, the landing gear height is a function of shock absorber and the landing gear deflection. The height is usually measured when the aircraft is on the ground; it has maximum take-off weight; and landing gear has the maximum deflection.

Fig 8.1 Thus, the landing gear when it has the maximum extension is still height, but is less important and application. The distance between the lowest point of the landing gear (i.e. ground) to the aircraft cg is also of significant importance and will be employed during 59

AIRCRAFT DESIGN PROJECT

calculations. Wheel base is the distance between main gear and other gear (from side view). The landing gear is divided into two sections: 1. Main gear or main wheel, 2. Secondary gear or secondary wheel. Main gear is the gear which is the closest to the aircraft center of gravity (cg). During the landing operation, the main wheel touches first with the point of contact to the ground. Furthermore, during the take-off operation, the main wheel leaves the ground last. On the other hand, main gear is carrying great portion of the aircraft load on the ground. Wheel track is the distance between two main gears (left and right) from front view. If a gear is expected to carry high load, it may have more than one wheel. In general, the landing gear weight is about 3% to 5% of the aircraft take-off weight. For instance, in the case of a Boeing 747 (Figures 3.7, 3.12 and 9.4), the landing gear assembly weight about 16,000 lb.

Functional Analysis and Design Requirements: It terms of design procedure, the landing gear is the last aircraft major component which is designed. In another word, all major components (such as wing, tail, fuselage, and propulsion system) must be designed prior to the design of landing gear. Furthermore, the aircraft most aft center of gravity (cg) and the most forward cg must be known for landing gear design. In some instances, the landing gear may drive the aircraft designer to change the aircraft configuration to satisfy landing gear design requirements. The primary functions of a landing gear are as follows: 1. To keep the aircraft stable on the ground and during loading, unloading, and taxi. 2. To allow the aircraft to freely move and maneuver during taxing. 3. To provide a safe distance between other aircraft components such as wing and fuselage while the aircraft is on the ground position to prevent any damage by the ground contact. 4. To absorb the landing shocks during landing operation. 5. To facilitate take-off by allowing aircraft acceleration and rotation with the lowest friction. In order to allow for a landing gear to function effectively, the following design requirements are established. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10.

Ground clearance requirement. Steering requirement. Take-off rotation requirement. Tip back prevention requirement. Overturn prevention requirement. Touch-down requirement. Landing requirement. Static and dynamic load requirement. Aircraft structural integrity. Ground lateral stability 60

AIRCRAFT DESIGN PROJECT

11. 12. 13. 14.

Low cost Low weight Maintainability Manufacturability

Landing Gear Configuration: The first job of an aircraft designer in the landing gear design process is to select the landing gear configuration. Landing gear functions may be performed through the application of various landing gear types and configuration. Landing gear design requirements are parts of the aircraft general design requirements including cost, aircraft performance, aircraft stability, aircraft contact, maintainability, productibility and operational considerations. In general, there are ten configurations for a landing gear as follows: 1. 2. 3. 4. 5. 6. 7. 8. 9.

Single main Bicycle Tail-gear Tricycle or nose-gear Quadricycle Multi-bogery Releasable rail Skid Seaplane landing device

Configuration Selection As aircraft grow larger, they employ more wheels to cope with the increasing weights. The earliest "giant" aircraft ever placed in quantity production, the Zeppelin-Staaken R.VI German World War I long-range bomber of 1916, used a total of eighteen wheels for its undercarriage, split between two wheels on its nose gear struts, and a total of sixteen wheels on its main gear units under each tandem engine nacelle, to support its loaded weight of almost 12 metric tons. The Boeing 747 has five sets of wheels: a nose-wheel assembly and four sets of four-wheel bogies. A set is located under each wing, and two inner sets are located in the fuselage, a little rearward of the outer bogies, adding up to a total of eighteen wheels and tires. The Airbus A380 also has a four-wheel bogie under each wing with two sets of six-wheel bogies under the fuselage. The enormous Ukrainian Antonov An-225 jet cargo aircraft has one of the largest, if not the largest, number of individual wheel/tire assemblies in its landing gear design – with a total of four wheels on the twin-strut nose gear units, and a total of 28 main gear wheel/tire units, adding up to a total of 32 wheels and tires. Retractable gear: To decrease drag in flight some undercarriages retract into the wings and/or fuselage with wheels ar against or concealed behind doors, this is called retractable gear. It was in late 61

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1920s and 1930s that such retractable landing gear became common. This type of gear arrangement increased the performance of aircraft by reducing the drag. Tire Sizing: Technically, the term ―wheel‖ refers to a circular metal plastic object around which the rubber ―tire‖ is mounted. The brake system is mounted inside the wheel to slow the aircraft during landing. However, in majority of cases, the entire wheel, tire, and brake system is also referred to as the wheel. The fundamental materials of modern tires are synthetic or natural rubber, fabric and wire, along with other compound chemicals. Today, the vast majority of tires is generally pneumatic inflatable and includes a doughnut-shaped body of cords and wires encased in rubber. So they consist of a tread and a body. Tires perform four important functions with the assistance of the air contained within them: 1. Tires support the aircraft structure off the ground. 2. They help absorb shocks from the runway surface. 3. They help transmit acceleration and braking forces to the runway surface. 4. They help change and maintain the direction of motion.

Fig 8.2 As a guideline, the following is the information about tires of a civil transport, a military fighter, and a GA aircraft. The transport aircraft Boeing 777-200 is employing (8) Goodyear main tires H49x19-22, and Michelin radial nose wheel tires 44x18-18. The fighter aircraft McDonnell Dougles F-15 Eagle is utilizing (8) Bendix wheels and Michelin AIR X with nose wheel tires size 22x7.75-9, and main wheel tires size 36x11-18 where tire pressure is 305 psi. The main wheel tire of business jet Cessna 650 Citation VII (8) is of size 22x5.75 (pressure of 168 psi), while the nose wheel tire size is 188x4.4 (140 psi). Generally speaking, for a tricycle configuration, nose tires may be assumed to be about 50100% the size of the main tires. For quadricycle and bicycle configurations, the front tires are often the same size as the main tires.

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Fuselage Design

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Introduction: The fuselage is an aircraft‘s main body section that holds crew and passengers or cargo. In single-engine aircraft it will usually contain an engine, although in some amphibious aircraft the single engine is mounted on a pylon attached to the fuselage which in turn is used as a floating hull. The fuselage also serves to position control and stabilization surfaces in specific relationships to lifting surfaces, required for aircraft stability and maneuverability. Common practice to modularize layout: 1. Crew compartment, power plant system, payload configuration, fuel volume, landing gear stowage, wing carry-through structure, empennage etc. 2. Or simply into front, centre and rear fuselage section designs. Functions of fuselage: 1. Provision of volume for payload. 2. Provide overall structural integrity. 3. Possible mounting of landing gear and power plant. Once fundamental configuration is established, fuselage layout proceeds almost independently of other design aspects. Primary considerations: Most of the fuselage volume is occupied by the payload, except for:    

Single and two-seat light aircraft. Trainer and light strike aircraft. Combat aircraft with weapons on outer fuselage & wing. High performance combat aircraft.

Structure: The primary concern in the development of a good structural arrangement is the provision of efficient ―load paths‖ – the structural elements by which opposing forces are connected. The primary forces to be resolved are the lift of the wing and the opposing weight or the 64

AIRCRAFT DESIGN PROJECT

major parts of the aircrafts, such as the engines and payload. The size and the weight of the structural members will be minimized by locating these opposing forces near to each other. The fighter skin is fastened by a longer or stringer or stiffener which is thin strip of material, to which the skin of the aircraft is fastened. In the fuselage, stringers are attached to formers (also called frames) and run the longitudinal direction of the fighter jet. They are primarily responsible for transferring the loads (aerodynamic) acting on the skin onto the frames/ formers. In the wing or horizontal stabilizer, longerons run span wise and attach between the ribs. The primary function here also is to transfer the bending loads acting on the wings onto the ribs and spar.

Fig 9.1 The ―ring-frame‖ approach relies upon large, heavy bulkhead to carry the bending moment through the fuselage. The wings panels are attached to fittings on the side of these fuselage bulkheads. While this approach is usually heavier from a structural viewpoint, the resulting drag reduction at high speeds has led to the use of the approach for most modern fighters. The fuselage width for the cargo aircraft would be around 8.38 m (B 747).

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Fig 9.2

Payload includes:     

Internal weapons (guns, free-fall bombs, bay-housed guided weapons) Crew (significant for anti-sub and early-warning aircraft) Avionics equipment. Fight test instrumentation (experimental aircraft) Fuel (often interchangeable with other payload items on a mass basis).

Fuselage Aerodynamics:     

Aim to achieve reasonable streamlined form together with minimum surface area to meet required internal volume. Both drag and mass heavily influenced by surface area. Require absence of steps and minimum number of excrescences Fundamental differences between subsonic and supersonic applications. Concerned with: cross-section shape, nose shape & length, tail shape/length, overall length.

Cross-Section Shape – supersonic Aircraft:   

Not too critical aerodynamically, but should: Avoid sharp corners Provide fairings for protuberances

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

Constant cross-section preferable for optimized volume utilization and ease of manufacture.

Nose Shape:     

Should not be unduly ―bluff‖. Local changes in cross-section needed to accommodate windscreen panels. Windscreen angle involves compromise between aerodynamics, bird-strike, reflection and visibility requirements. Windscreen panel sizes should be less than 0.5m2 each. Starting point for front fuselage layout is often satisfactory position for pilot‘s eye.

Reasonable nose length is about: 

4 x fuselage diameter (supersonic).

Tail Shape:    

Smooth change in section required, from maximum section area to ideally zero. Minimization of base area especially important for transonic/ supersonic aircraft. Important parameter for determining tail upsweep angle is ground clearance required for take-off and landing rotation. Typically 12‖ to 15‖.

Typical tail section lengths are: 

6 to 7 x diameter (supersonic)

Centre Fuselage & Overall Length – Supersonic Aircraft:     

Theoretically minimum drag for streamlined body with fineness ratio (length / diameter) of 3. In reality, typical value is around 10, due to: Need to utilize internal volume efficiently. Requirement for sufficiently large moment arm for stability / control purposes. Suitable placement of overall CG.

Wing Location – Aerodynamics Considerations:    

Mid-wing position gives lowest interference drag, especially well for supersonic stealth fighter aircraft. Top-mounted wing minimizes trailing vortex drag, especially well for low speed aircraft. Low wing gives improved landing gear stowage & more usable flap area. From the above given locations of wings, the one chosen is the Low wing configuration which gives improved landing gear stowage & more usable flap area.

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Empennage Layout: Vertical Surface: 

Single central fin most common arrangement, positioned as far aft as possible.

Horizontal Surface:  

Efficiency affected by wing downwash, thus vertical location relative to wing important. Usually mounted higher then wing except on high design or with small moment arm – low tail can give ground clearance problems.

Avionics: 1. Three X-band AESA radars located at the front and sides of the aircraft. These will be accompanied by L-band radars on the wing leading edges. L-band radars are proven to have increased effectiveness against very low observable, or stealthy, targets which are optimized only against X-band frequencies, but their longer wavelength reduce their resolution. 2. Full authority digital engine (or electronics) control (FADEC) is a system consisting of digital computer, called an electronic engine controller (EEC) or engine control unit (ECU), and its related accessories that control all aspects of aircraft engine performance. Advantages  Better fuel efficiency.  Automatic engine protection against out-of-tolerance operations  After as the multiple channel FADEC computer provides redundancy in ease of failure.  Care-free engine handling, with guaranteed thrust settings  Ability to use single engine type for wide thrust requirements by just reprogramming for FADECs  Provides semi-automatic engine starting. 3. An infra-red search and track (IRST) system (sometimes known as infra-red sighting and tracking) on the nose of fighter which is a method for detecting and tracking objects which give off infrared radiation such as jet aircraft and helicopters. 4. 3-D thrust vectoring along with all three aircraft axes: pitch, yaw and roll. Weapon carriage and missiles: Two internal bays estimated at 4.6-4.7 meters by 1-1.1 meters in an internal missile and bomb carriage tank. The missile is launched by ejection launch mechanism.

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The Zvezda Kh-35U (Russian: X-35Y; AS-20 ‗Kayak‘) is the jet-launched version of a Russian subsonic anti-ship missile. It is launched by the Ejection launch mechanism. Weight = 520 kg 4 * JDAM: The Joint Direct Attack Munition (JDAM) is a guidance kit that converts unguided bombs, or ―dumb bombs‖ into all-weather ―smart‖ munitions. JDAM-equipped bombs are guided by an integrated inertial guidance system coupled to a Global Positioning System (GPS) receiver, giving them a published range of upto 15 nautical miles (28km) = 910 kg.

Fig 9.3

External hardpoint missiles: The missiles are launched by the rail launch mechanism which is attached to pylons of the wing. The M61 Vulcan – It a hydraulically or pneumatically driven, six-barreled, air cooled, electrically fired Gatling-style rotary cannon which fires 20 mm rounds at an extremely high rate. Weight = 112 kg.

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70

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Performance Characteristics

Take - off Performance: Take-off distance,

, smooth panel surface , grass

71

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(

)

(

)

Climbing performance: Rate of climb,

Constant Altitude Bank Turn:



= 72

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Landing Performance: Ground roll landing distance,

Where W is in Newtons

73

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Result: The performance characteristics were found to be:

Conclusion Thus all the design and performance parameters were found to be as follows:. Geometric parameters

Wing parameters

Performance parameters

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Aspect ratio = 10.52 Max. Take – off weight = 70,000 kgs Service ceiling = 11,000 m Length = 36 m Height = 11.52 m = 27.252 m = 9901.439 (By calculation)

Wing loading = 400 Wingspan = 39.47 m Wing area = 150.06 m2

Range = 4400km Rate of climb = 770 m/min

Airfoil selected = NACA 2414 Sweep angle = 7 Taper ratio = 0.35 = 5.707 m = 4.1496 m = 1.99 m Position of wing = High wing Flap selected = Slotted flap

Table 11.1

Lift and Drag parameters

Tail parameters Tail selected = Conventional type

Engine parameters Engine selected = CFM International CM 56 – 7 No. of engines = 2 Thrust = 226.23 KN

Table 11.2

Bibliography  Aircraft design: A conceptual approach, Raymer, Daniel P, American Institute of Aeronautics & Astronautics; 5th Revised edition, ISBN-10: 1600869114

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 Theory of wing sections including a summary of airfoil data, Abott, Ira H, ISBN-10: 0486605868  Airplane Design, Roskam, Ian, Darcorporation; 2nd edition, ISBN-10: 1884885241  Aircraft Performance & Design, Anderson J D, Tata McGraw-Hill Education, ISBN 0070702454, 9780070702455  Airfoiltools.com  Wikipedia.org  web.mit.edu/drela/Public/web/xfoil/

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Annexure

Table 1

Table 2

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Table 3

Table 4 78

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Table 5

Type of aircraft

e

Piston Propeller

0.022-0.028

0.82

Large Turbo-prop

0.025

0.79

Small General Aviation

0.025

0.77

Small General Aviation

0.032

0.72

Agricultural A/C

0.75

0.68 79

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Subsonic Jet

0.014

0.78 Table 6

Table 7

80