56914974-Aircraft-Design-Project-1.pdf

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1. Introduction

1.1 Overview:

Three major types of airplane designs are 1. Conceptual design

2. Preliminary design 3. Detailed design

1. Conceptual design:

It depends on what are the major factors for designing the aircraft. (a) Power plant Location:

The Power plant must be located in the wings. (b) Selection of Engine:

The engine should be selected according to the power required i.e., thrust required.

(c) Wing selection:

The selection of wing depends upon the selection of (1) Low wing

(2) Mid wing (3) High wing

- For a bomber the wing is mostly high wing configuration and anhedral.

- Sweep may be required in order to reduce wave drag.

2. Preliminary design:

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3. Detailed design:

In the detailed design considers each & every rivets, bolts, paints etc. In this design the connection & allocations are made.

1.1

Bomber:

A bomber is a military aircraft designed to attack ground and sea targets, by dropping bombs on them, or – in recent years – by launching cruise missiles at them.

Strategic bombers are primarily designed for long-range bombing missions against strategic targets such as supply bases, bridges, factories, shipyards, and cities themselves, in order to damage an enemy's war effort.

Tactical bombing, aimed at enemy's military units and installations, is typically assigned to smaller aircraft operating at shorter ranges, typically along the troops on the ground or sea. This role is filled by various aircraft classes, as different as light bombers, medium bombers, dive bombers, fighter-bombers, ground-attack aircraft and multirole combat aircraft among others.

1.1.2 Origin of Bombers:

Bombers evolved at the same time as the fighter aircraft at the start of World War I. The first use of an air-dropped bomb however, was carried out by the Italians in their 1911 war for Libya.

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1.3 Project requirement

1. To design a bomber aircraft

2. Range of 20,000 km with refueling support & must carry 75,000+ kg of bombs & missiles (possibly nuclear warheads)

3. To operate at subsonic and transonic regimes

4. To operate at regional bases with low cost of operation & maintenance 5. The aircraft must also be capable of single pilot operation scenario. 6. Due to long range pilot work load must be reduced

7. The aircraft must be all weather, all terrain operation capable including the airbase.

8. To take up a load factor +7.5g to -3.5g.

1.4. Preferred Configuration:

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2. Comparative study of various bomber aircrafts

The first step in the design of aircraft is to collect data of existing aircraft of similar purpose i.e., bomber. This step is vital in aircraft design as it gives the designer an insight into the conventional trend in aircraft design.

The designer may, with the help of the data thus acquired, get an idea of the basic factors that affect the aircraft’s performance viz. Weight, Cruise velocity, Range, Wing area, Wingspan & Engine thrust. This database will also serve, during the design process, as a guide for validation of the design parameters that will be calculated, so that the designer does not deviate unduly from the conventional design.

The performance data of various bomber aircraft with payload capacity between 5000 and 56600 kg was collected from the appropriate resources.

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2.1 Comparative configuration study of bomber airplanes:

Collection of Comparative Data -Dimension

S.No Name of the aircraft Payload Capacity (kg) Length (m) Height (m) Wing span (m)2 No of Power Plant Loaded Weight (kg) Maximum Takeoff Weight (kg) Empty weight (kg) 1 Mirage IIIE 5000 15 4.5 8.22 1 12200 13500 7050 2 Mirage IVA 7264 23.49 5.4 11.85 2 31600 33475 14500 3 F-111F 14300 22.4 5.22 19.2 2 37600 45300 21400 4 F-111F Swept 14300 22.4 5.22 9.75 2 37600 45300 21400 5 Tu-22R 9000 41.6 10.13 23.17 2 85000 92000 6 Tu-85/1 18000 39.306 11.358 55.96 4 76000 107292 54711 7 YB-60 33000 52.1 18.4 62.8 8 73000 140000 69407 8 B-2A 23000 21 5.18 52.4 4 152200 170600 71700 9 Tu-142M3 15000 49.5 12.12 51.1 4 170000 185000 90000 10 Tu-95MS 15000 46.2 12.12 50.1 4 171000 188000 90000 11 B-1B 56600 44.5 10.4 41.8 4 148000 216400 87100 12 B-1B Swept 56600 44.5 10.4 24.1 4 148000 216400 87100 13 B-52H 31500 48.5 12.4 56.4 8 120000 220000 83250 14 Tu-160 40000 54.1 13.1 55.7 4 267600 275000 110000 15 Tu-160 Swept 40000 54.1 13.1 35.6 4 267600 275000 110000 Table 2.1 Collection of Comparative Data -Dimension

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Collection of Comparative Data -Performance parameters

S.No Name of the aircraft Thrust to weight ratio Wing loading (N/m2) Aspect

ratio Power Plant 1

Mirage

IIIE 0.50 3796.47 1.939 SNECMA Atar 09C turbojet 2

Mirage

IVA 0.60 5949.29 1.8

SNECMA Atar 9K-50[13] turbojets Dry thrust: 49.03 kN (11,023 lbf) each Thrust with afterburner: 70.61 kN (15,873 lbf) each

3 F-111F 0.61 6035.11 7.56

Pratt & Whitney TF30-P-100 turbofans Thrust with afterburner: 25,100 lbf (112 kN) each Dry thrust: 17,900 lbf (79.6 kN) each

4

F-111F

Swept 0.61 7563.51 1.95

Pratt & Whitney TF30-P-100 turbofans Thrust with afterburner: 25,100 lbf (112 kN) each Dry thrust: 17,900 lbf (79.6 kN) each

5 Tu-22R 0.38 5150.25 3.314

Dobrynin RD-7M-2 turbojets Dry thrust: rated 107.9 kN (24,250 lbf) each Thrust with afterburner: 161.9 kN (36,376 lbf) each

6 Tu-85/1 0.44 2717.37 11.45

Dobrynin VD-4K turbo-compound radial engines, 3,200 kW (4,300 hp) each

7 YB-60 0.44 1471.50 8.1 Pratt & Whitney J57-P-3 turbojets, (38 kN) each 8 B-2A 0.21 3227.49 5.74

General Electric F118-GE-100 non-afterburning turbofans, 17,300 lbf (77 kN) each 9 Tu-142M3 0.29 9112.02 8.394 Kuznetsov NK-12MV turboprops, 11,033 kW (14,795 shp) each 10 Tu-95MS 0.40 5944.86 8.097 Kuznetsov NK-12M turboprops, 11,000 kW (14,800 shp)[23] each 11 B-1B 0.38 8004.96 9.65

General Electric F101-GE-102 augmented turbofans Dry thrust: 14,600 lbf (64.9 kN) each Thrust with afterburner: 30,780 lbf (136.92 kN) each

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B-1B

Swept 0.38 8004.96 9.65

General Electric F101-GE-102 augmented turbofans Dry thrust: 14,600 lbf (64.9 kN) each Thrust with afterburner: 30,780 lbf (136.92 kN) each

13 B-52H 0.31 5836.95 8.56

Pratt & Whitney TF33-P-3/103 turbofans, (76 kN) each

14 Tu-160 0.37 7269.21 7.757

Samara NK-321 turbofans Dry thrust: 137.3 kN (30,865 lbf) each Thrust with afterburner: 245 kN (55,115 lbf) each

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Tu-160

Swept 0.37 7279.02 3.52

Samara NK-321 turbofans Dry thrust: 137.3 kN (30,865 lbf) each Thrust with afterburner: 245 kN (55,115 lbf) each

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Collection of Comparative Data -Performance parameters

S.No Name of the aircraft Maximum Speed (m/s) Range (km) Service ceiling (m) Rate of Climb (m/s) Combat Radius (km) Payload Capacity (kg)2 1 Mirage IIIE 652.8 2400 17000 83.3 1200 5000 2 Mirage IVA 650.0 4000 20000 43.13 1240 7264 3 F-111F 737.5 6760 20100 131.5 2140 14300 4 F-111F Swept 737.5 6760 20100 131.5 2140 14300 5 Tu-22R 419.4 4900 13300 2450 9000 6 Tu-85/1 177.2 12000 11700 17 5850 18000 7 YB-60 227.2 13000 16200 5.38 4700 33000 8 B-2A 270.0 11100 15200 5550 23000 9 Tu-142M3 256.9 12000 6500 15000 10 Tu-95MS 255.6 15000 13716 10 7500 15000 11 B-1B 372.2 11998 18000 5543 56600 12 B-1B Swept 372.2 11998 18000 5543 56600 13 B-52H 277.8 16232 15000 31.85 7210 31500 14 Tu-160 616.7 12300 15000 70 7300 40000 15 Tu-160 Swept 616.7 12300 15000 70 7300 40000

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Collection of Comparative Data -Performance parameters

S.No Name of the aircraft Airfoil Span to length ratio Span to height ratio Introduction Retiremen t Remarks

1 Mirage IIIE 0.30 1.83 1961 In service Good

2 Mirage IVA 0.23 2.19 1959 2005 Good

3 F-111F NACA 64-210.68 root, NACA 64-209.80 tip 0.23 3.68 1967 2010 Good 4 F-111F Swept NACA 64-210.68 root, NACA 64-209.80 tip 0.23 1.87 1967 2010 Good 5 Tu-22R 0.24 2.29 1962 1990 Good

6 Tu-85/1 0.29 4.93 Prototype Good

7 YB-60 0.35 3.41 1954 In service Good

8 B-2A 0.25 10.12 1997 In service Good 9 Tu-142M3 0.24 4.22 1953 In service Awesome 10 Tu-95MS 0.26 4.13 1956 In service Awesome 11 B-1B NA69-190-1 0.23 4.02 1986 In service Good 12 B-1B Swept NA69-190-2 0.23 2.32 1986 In service Good 13 B-52H

NACA 63A219.3 mod root, NACA

65A209.5 tip 0.26 4.55 1961 In service Awesome 14 Tu-160 0.24 4.25 2005 In service Good 15 Tu-160 Swept 0.24 2.72 2005 In service Good Table 2.4 Collection of Comparative Data -Performance parameters (Cont.)

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2.2 Comparative graphs for determining optimum value:

S.No Name of the aircraft Wing loading (N/m2) Maximum Speed (m/s)

1 Mirage IIIE 3796.47 652.8 2 Mirage IVA 5949.29 650.0 3 F-111F 6035.11 737.5 4 F-111F Swept 7563.51 737.5 5 Tu-22R 5150.25 419.4 6 Tu-85/1 2717.37 177.2 7 YB-60 1471.50 227.2 8 B-2A 3227.49 270.0 9 Tu-142M3 9112.02 256.9 10 Tu-95MS 5944.86 255.6 11 B-1B 8004.96 372.2 12 B-1B Swept 8004.96 372.2 13 B-52H 5836.95 277.8 14 Tu-160 7269.21 616.7 15 Tu-160 Swept 7279.02 616.7

Table 2.5 Wing loading vs. Maximum Speed

Figure 2.1 Wing loading vs. Maximum S peed

0.00 1000.00 2000.00 3000.00 4000.00 5000.00 6000.00 7000.00 8000.00 9000.00 10000.00 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 W ing L oa di ng ( N /s q. m ) Maximum Speed (m/s)

Wing Loading (N/sq.m) Vs Maximum Speed

(m/s)

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10 S.No Name of the aircraft Span to length ratio Maximum Speed (m/s)

Table 2.6 Span to length Ratio vs. Maximum Speed

Figure 2.2 S pan to length Ratio vs. Maximum S peed

0.00 0.50 1.00 1.50 2.00 2.50 3.00 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 Spa n to le ng th R at io Maximum Speed (m/s)

Span to length ratio Vs Maximum Speed (m/s)

Span to length ratio

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11 S.No Name of the aircraft Aspect ratio Maximum Speed (m/s)

Table 2.7 Aspect Ratio vs. Maximum Speed

Figure 2.3 Aspect Ratio vs. Maximum S peed

0 2 4 6 8 10 12 14 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 A spe ct R at io Maximum Speed (m/s)

Aspect ratio Vs Maximum Speed (m/s)

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12 S.No Name of the aircraft Wing Area (m2)

Maximum Speed (m/s) 1 Mirage IIIE 34.85 652.8 2 Mirage IVA 78 650.0 3 F-111F 61.07 737.5 4 F-111F Swept 48.77 737.5 5 Tu-22R 162 419.4 6 Tu-85/1 273.6 177.2 7 YB-60 486.7 227.2 8 B-2A 478 270.0 9 Tu-142M3 311.1 256.9 10 Tu-95MS 310 255.6 11 B-1B 180.2 372.2 12 B-1B Swept 181.2 372.2 13 B-52H 370 277.8 14 Tu-160 400 616.7 15 Tu-160 Swept 360 616.7

Table 2.8 Wing Area vs. Maximum Speed

Figure 2.4 Wing area vs. Maximum S peed

0 100 200 300 400 500 600 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 W ing A re a (s q. m ) Maximum Speed (m/s)

Wing Area (sq.m) Vs Maximum Speed (m/s)

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13 S.No Name of the aircraft Combat Radius (km) Maximum Speed (m/s)

1 Mirage IIIE 1200 652.8 2 Mirage IVA 1240 650.0 3 F-111F 2140 737.5 4 F-111F Swept 2140 737.5 5 Tu-22R 2450 419.4 6 Tu-85/1 5850 177.2 7 YB-60 4700 227.2 8 B-2A 5550 270.0 9 Tu-142M3 6500 256.9 10 Tu-95MS 7500 255.6 11 B-1B 5543 372.2 12 B-1B Swept 5543 372.2 13 B-52H 7210 277.8 14 Tu-160 7300 616.7 15 Tu-160 Swept 7300 616.7

Table 2.9 Span to length Ratio vs. Maximum Speed

Figure 2.5 Combat radius vs. Maximum S peed

0 1000 2000 3000 4000 5000 6000 7000 8000 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 C om ba t ra di us ( km ) Maximum Speed (m/s)

Combat Radius (km) Vs Maximum Speed (m/s)

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14 S.No Name of the aircraft

Payload Capacity (kg)2 Maximum Speed (m/s) 1 Mirage IIIE 5000 652.8 2 Mirage IVA 7264 650.0 3 F-111F 14300 737.5 4 F-111F Swept 14300 737.5 5 Tu-22R 9000 419.4 6 Tu-85/1 18000 177.2 7 YB-60 33000 227.2 8 B-2A 23000 270.0 9 Tu-142M3 15000 256.9 10 Tu-95MS 15000 255.6 11 B-1B 56600 372.2 12 B-1B Swept 56600 372.2 13 B-52H 31500 277.8 14 Tu-160 40000 616.7 15 Tu-160 Swept 40000 616.7

Table 2.10 Payload capacity vs. Maximum Speed

Figure 2.6 Payload Capacity vs. Maximum S peed

0 10000 20000 30000 40000 50000 60000 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 P ay loa d C apa ci ty ( kg ) Maximum Speed (m/s)

Payload Capacity (kg) Vs Maximum Speed (m/s)

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Thrust to weight ratio Maximum Speed (m/s) 1 Mirage IIIE 0.5 652.8 2 Mirage IVA 0.6 650.0 3 F-111F 0.61 737.5 4 F-111F Swept 0.61 737.5 5 Tu-22R 0.38 419.4 6 Tu-85/1 0.44 177.2 7 YB-60 0.44 227.2 8 B-2A 0.205 270.0 9 Tu-142M3 0.29 256.9 10 Tu-95MS 0.4 255.6 11 B-1B 0.38 372.2 12 B-1B Swept 0.38 372.2 13 B-52H 0.31 277.8 14 Tu-160 0.37 616.7 15 Tu-160 Swept 0.37 616.7

Table 2.11 Thrust to weight Ratio vs. Maximum Speed

Figure 2.7 Thrust to weight Ratio vs. Maximum S peed

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.0 200.0 400.0 600.0 800.0 Thr us t to w e ig ht r at io Maximum Speed (m/s)

Thrust to Weight Ratio Vs Maximum Speed

(m/s)

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16 S.No Name of the aircraft Span to height ratio Maximum Speed (m/s)

Table 2.12 Span to height Ratio vs. Maximum Speed

Figure 2.8 S pan to height Ratio vs. Maximum S peed

0.00 2.00 4.00 6.00 8.00 10.00 12.00 0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0 Spa n to H e ig ht R at io Maximum Speed (m/s)

Span to Height Ratio Vs Maximum Speed (m/s)

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Maximum Takeoff

Weight (kg) Maximum Speed (m/s)

1 Mirage IIIE 13500 652.8 2 Mirage IVA 33475 650.0 3 F-111F 45300 737.5 4 F-111F Swept 45300 737.5 5 Tu-22R 92000 419.4 6 Tu-85/1 107292 177.2 7 YB-60 140000 227.2 8 B-2A 170600 270.0 9 Tu-142M3 185000 256.9 10 Tu-95MS 188000 255.6 11 B-1B 216400 372.2 12 B-1B Swept 216400 372.2 13 B-52H 220000 277.8 14 Tu-160 275000 616.7 15 Tu-160 Swept 275000 616.7

Table 2.13 Maximum takeoff weight vs. Maximum Speed

Figure 2.9 Maximum takeoff weight vs. Maximum S peed

0 50000 100000 150000 200000 250000 300000 0.0 200.0 400.0 600.0 800.0 M axi m um T ak e of f W e ig ht ( kg ) Maximum Speed (m/s)

Maximum Takeoff Weight (kg) Vs Maximum

Speed (m/s)

Maximum Takeoff Weight (kg)

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2.2 Parameter Selection:

From Comparison (Assumed and extrapolated values from graph) Maximum takeoff weight (kg) 500000

Thrust to weight ratio 0.28

Aspect ratio 8.4

Wing loading (N/sq.m) 7848

Span to height ratio 5

Span to length ratio 1.5 Combat radius (km) 5000 Pay load capacity (kg) 75000 Maximum speed (kmph) 1000 Service ceiling (m) 15000 Maximum speed (m/s) 277.777 Table 2.14 Parameter Selection

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3. Rough Weight Estimate

Optimal values of mass fraction for bombers

Parameter Range of values Notation

Empty Mass Ratio 0.37-0.32 ME/MTO

Total Fuel Mass Ratio 0.40-0.62 MF/MTO

Payload Ratio 0.14-0.19 MPay/MTO

Wing Loading 4385-7848 N/m2 Wo/S

Thrust to Weight Ratio 0.26-0.40 T/Wo

Table 3.1 Mass Fraction Parameters for bomber 3.1 General rough weight estimate:

Figure 3.1 Payload mass Fraction vs. Maximum Takeoff weight

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0 50000 100000 150000 200000 250000 300000 MPa y /M To

Maximum Takeoff weight (kg)

M

_Pay

/M

_To

Vs Maximum Takeoff weight (kg)

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= 75000 kg

Figure3.2 Empty mass Fraction vs. Maximum Takeoff weight

0 0.1 0.2 0.3 0.4 0.5 0.6 0 50000 100000 150000 200000 250000 300000 ME /M To

Maximum Takeoff weight (kg)

M

_E

/M

_To

Vs Maximum Takeoff weight (kg)

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21 Final Values from rough weight estimate:

Mass Fraction Payload 0.15 Fuel 0.45 Structure 0.32 Power plant 0.07 Fixed equipments 0.01 Total 1.00

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4. Redefined Mass Estimation

4.1 Mission profile analysis

Profile 1: Strategic bombing mission

Analysis of Mission Profile: Warmup and takeoff

Climb or descend Landing ( ) ( ₎ 2’ 3’ 4’ 5’ 6’ 7’ 8’ 9’ 10’ 0 1 R 3 2 h 10000 km 1000 km 1000 km 9000 km 1/2 hr

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Figure 4.2 Empty mass fraction vs takeoff mass -- taken from "Aircraft Design: A Conceptual Approach" by Daniel P.Raymer

Analysis of Mission Profile

TSFC values for Bomber

Cruise Loiter

0.5 0.4

Table 4.1 TSFC values in lb/lbf-hr for bomber

Where A & c are constants from the historic data for bomber

c = -0.07; A = 0.93 (taken from Aircraft Design: A Conceptual Approach by Daniel P.Raymer)

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4.2: Before Refueling

Part 1: Before refueling:

Warm-up and Take-off: (0-1)

Climb: (1-2)

Cruise at 60% of maximum speed: (2-3)

For analysis (L/D) optimal = 17

Thrust Specific fuel Consumption C = 0.0001389 (kg / N-s)

0 1 R 3 2 h 10000 km

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25 Range R2-3 = 10000 km ⁄ Descend: (3-R) Total Mass fraction for first part of mission profile:

Fuel Mass fraction for first half of mission profile:

( )

Thus the range of 10000 km can be interpreted as a combat radius of 5000 km.

4.3. Refueling:

Operation:

The tanker aircraft flies straight and level and extends the hose/drogue which is allowed to trail out behind and below the tanker under normal aerodynamic forces. The pilot of the receiver aircraft extends his probe (if required) and uses normal flight controls to fly the refueling probe directly into the basket. This requires a closure rate of approximately two knots (walking

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speed) in order to establish solid probe/drogue couple and pushing the hose several feet into the HDU. Too little closure will cause an incomplete connection and no fuel flow (or occasionally leaking fuel). Too much closure is dangerous because it can trigger a strong transverse oscillation in the hose, severing the probe tip. Another significant danger is that the drogue may hit the recipient aircraft and damage it—instances have occurred in which the drogue has shattered the canopy of a fighter aircraft, causing great danger to its pilot.

Figure 4.4 A Tu-95MS simulating aerial refueling with an Ilyushin Il -78

The optimal approach is from behind and below (not level with) the drogue. Because the drogue is relatively light (typically soft canvas webbing) and subject to aerodynamic forces, it can be pushed around by the bow wave of approaching aircraft, exacerbating engagement even in smooth air. After initial contact, the hose and drogue is pushed forward by the receiver a certain distance (typically, a few feet), and the hose is reeled slowly back onto its drum in the HDU. This opens the tanker's main refueling valve allowing fuel to flow to the drogue under the appropriate pressure (assuming the tanker crew has energized the pump). Tension on the hose is aerodynamically balanced by a motor in the HDU so that as the receiver aircraft moves fore and aft, the hose retracts and extends, thus preventing bends in the hose that would cause undue side loads on the probe. Fuel flow is typically indicated by illumination of a green light near the HDU. If the hose is pushed in too far or not far enough, a cutoff switch will inhibit fuel flow, which is typically accompanied by amber light. Disengagement is commanded by the tanker pilot with a red light.

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4.4. After Refueling:

Part 2: After refueling:

Cruise at 60% of maximum speed: (R-2)

Descend: (2’-3’) Cruise: (3’-4’) Bombing (4’) Climb: (4’-5’) R 1000 km h 2’ 3’ _5’ 4’ 6’ 7’ 8’ 9’ 10’ 1000 km _{9000 km}

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28 Cruise: (5’-6’) Loiter: (6’-7’) Loiter time = ½ hr Descend: (7’-8’) Landing: (8’-9’)

Total Mass fraction for second part of mission profile:

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Total fuel mass fraction after refueling:

( ) ( ) ( ) ( ) Hence ( )

is taken since the value turns out to be ( )

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Replacing as X in excel to solve the implicit function

X ranges from 500000 to 520000 since initial mass estimate is 500000 kg

Thus mass of the aircraft is 508034.68 kg

X f(x) 500000 -9458.077 510000 2308.6163 520000 14026.793 X f(x) 501000 -8279.156 502000 -7100.742 503000 -5922.833 504000 -4745.425 505000 -3568.517 506000 -2392.106 507000 -1216.19 508000 -40.76524 509000 1134.1693 X f(x) 508000 -40.76524 508100 76.750211 X f(x) 508000 -40.76524 508010 -29.01347 508020 -17.26176 508030 -5.51009 508040 6.2415284 X f(x) 508030 -5.51009 508031 -4.334926 508032 -3.159763 508033 -1.984599 508034 -0.809437 508035 0.3657252 X f(x) 508034.6 -0.10434 508034.7 0.0131767 X f(x) 508034.67 -0.022078 508034.68 -0.010327 508034.69 0.001425 508034.7 0.0131767 508034.71 0.0249283

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31 Take-off Weight of the aircraft:

4.5.Thrust Estimation

T=0.28×4983820.113 =1395469.632N T=1395.469 KN

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32 ⁄ TA = 1395.469 KN T = T=348.86725 KN/Engine

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5. Power Plant Selection

5.1. Comparative data of engines

Thrust Estimation: ( ( ( ⁄ ) ( ) ( ) ) )

For the chosen parameters:

T=0.28×4983820.113 =1395469.632N T=1395.469 KN ⁄ TA = 1395.469 KN T = T=348.86725 KN/Engine

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The performance data of various turbofan engines with thrust of range 330 kN to 500 kN were collected from the following resources

 www.jet-engine.net  www.wikipedia.org

S.No

Name of

the Engine Manufacturer Type

Length (m) Diameter (m) Wet weight (kg) Dry Weight (kg) 1 Trent-900 Rolls Royce Turbofan 3 Shaft 4.55 2.94 6271 2 GP-7000 Engine Alliance Turbofan 2 Shaft 4.74 3.16 6800 6712 3 GE 90-76B GE Turbofan 2 Shaft 4.90 3.40 7540 7074 4 GE 90-92B GE Turbofan 2 Shaft 4.90 3.40 7648 7074 5 GP 7270 Engine Alliance Turbofan 2 Shaft 4.75 3.15 6800 6712 6

GE90-110B1 GE Turbofan 2 Shaft 4.90 3.40 8253 7550 7 GP 7277 Engine Alliance Turbofan 2 Shaft 4.75 3.15 6482 6033 8 PW 40477 Pratt & Whitney Turbofan 2 Shaft 4.87 3.01 6986 6598 9 GE 90-85B GE Turbofan 2 Shaft 4.90 3.40 7474 7074 10 GE 90-94B GE Turbofan 2 Shaft 4.90 3.40 8253 7550 11 GE 90-90B GE Turbofan 2 Shaft 4.90 3.40 7548 7074 12 GE 90-115B GE Turbofan 2 Shaft 4.90 3.40 8283 7550 Table 5.1. Collection of Engine Comparative Data

S.No Name of the Engine Maximum Thrust (kN) Overall Pressure Ratio Thrust to Weight Ratio Fan Diameter (m) 1 Trent-900 360 38 5.15 2 GP-7000 363 43.9 4.73 2.95 3 GE 90-76B 340 41.9 5.4 3.12 4 GE 90-92B 409 41.8 5.7 3.12 5 GP 7270 311 43.9 4.73 2.95 6 GE90-110B1 489 42.5 6.2 3.12 7 GP 7277 343 43 5.2 2.95 8 PW 40477 343 40 2.84 9 GE 90-85B 377 42 3.12 10 GE 90-94B 417 42 5.6 3.12 11 GE 90-90B 400 42 5.4 3.12 12 GE 90-115B 512 42 6.3 3.12

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5.2. Engine Selection:

From this we select Engine Alliance GP 7000 Specification of Engine

Name of the Engine GP-7000

Manufacturer Engine Alliance

Type Turbofan 2 Shaft

Length (m) 4.74

Diameter (m) 3.16

Wet weight (kg) 6800

Dry Weight (kg) 6712

Maximum Thrust (kN) 363

Overall Pressure Ratio 43.9

Thrust to Weight Ratio 4.73

Fan Diameter (m) 2.95

Table 5.3. Selected Engine Datas

5.3. Redefined Thrust to weight ratio:

Closer to initial value assumed value of 0.28 TSFC ≈ 0.8

TSFC = 0.7913 N/N - hr = 0.7913

= 0.080662 kg/N- hr TSFC = 0.02240627 kg/N -s Service ceiling evaluation:

By taking service ceiling as h=15 km

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36 √ √ Number of Engines = 4

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6. Airfoil selection and Wing Geometry estimates

6.1. Main Parameter Selection:

Wing Loading: ̂ ̂ ̂ ̂

6.2 Fuel volume consideration:

ρF can vary from 600 kg/m 3

to 800 kg/m3. For ρF = 800 kg/m

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38

Volume of fuel accommodated in wing:

( (( ) ̂ ))

( ) (

( ̂ )) ( )

Selecting NACA 653-418 airfoil of fineness ratio (t/c ratio) as 0.18 ( ) ( ) ( )

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6.3 Takeoff Analysis:

Figure 6.1 Runway length survey for military installations

SR = 2000 m for around 68% of airbase in the world.

Assuming take off at 60% of runway length and accelerating at 20% the gravitational attraction, where vi is initial velocity during takeoff.

(40)

40 ( ₎ ( ) ̂ ̂

Where ̂ can be denoted as t also since we use MAC to obtain the thickness value

Thickness based Reynolds Number:

̂

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41

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42

Symbol Re x/c y/c Angle of

attack

CL max

5.9×106 0.266 -0.052 18° 1.42

8.9×106 0.267 -0.047 18° 1.51

Table 6.1 Airfoil data at various Re. At Re = 6.3×106 by interpolating we get Location of aerodynamic center x/c = 0.2668 y/c = -0.0491 CL max = 1.48 α = 18°

6.4 Flap selection:

Flap Chosen is Triple slotted flap

Wing setting angle or incidence angle iw= 3 degree Required Flap Deflection = 60°

Change in CL due to flap deflection:

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43

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44

6.5. Wing geometry:

Sweep Analysis:

For airfoil NACA 653- 418 At x/c = 0.46

( )

Figure 6.4 Variation of local velocity with the free stream velocity

(45)

45 √ √ √ √

Hence by comparing v and aM SL and aalt it is clear that the shock wave is formed.

In order to avoid this unwanted phenomenon we need to sweep the wing. Critical Mach number:

( )

If the maximum velocity reached on the upper surface is equal to the lowest possible value of speed of sound then the velocity V∞ will be critical velocity which corresponds to critical Mach number

Or simply M x/c= 0.46 =1 ( ) ( )

(46)

46 √ Critical Mach number for the airfoil:

Figure 6.5 S wept back wing

By using a trapezoidal and sweepback we may get

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47

For 1000 km/hr or 277.77 m/s speed we get the sweep to avoid shock on upper surface as

Where is due to taper property

(

⁄ )

Mean Aerodynamic Chord (MAC) ( ̂) for swept back wing: ̂ (

)

Span wise location of MAC

On simplifying we get : ̂ ( )

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48

Where λ is taper ratio: λ can vary from 0 to 1. By evaluating the above four equations we get :

λ cr (m) ct (m) (rad) (deg) 1 8.695 8.695 0.000 0.000 0.9 9.153 8.237 0.025 1.435 0.8 9.661 7.729 0.053 3.028 0.7 10.229 7.161 0.084 4.802 0.6 10.869 6.521 0.118 6.786 0.5 11.593 5.797 0.157 9.015 0.4 12.421 4.969 0.201 11.529 0.3 13.377 4.013 0.251 14.372 0.2 14.492 2.898 0.307 17.595 0.1 15.809 1.581 0.371 21.251 0 17.390 0.000 0.443 25.384

Table 6.2 Angle of taper for various taper ratios Taking the value

λ = 0.5 cr = 11.593 m

ct = 5.797 m

Span wise location of MAC:

Hence:

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49

Figure 6.6 Effect of aspect ratio on lift curve slope

√ a= 0.1213507 /degree for a’ = 0.15/ degree

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7. Landing gear design

7.1. Tyre selection:

7.1.1. Load Distribution:

Typical load of aircraft while landing Possibility of aborting mission would lead to

And during static condition

Typically main landing gear takes around 90 % of load and the Nose landing gear takes around 10% of total load.

Load taken by wheels in nose landing gear = 0.1× 4983820.113 = 498.382 kN

Load taken by wheels in main landing gear = 0.9× 4983820.113 = 4485.438 kN Nose Landing Gear Main Landing Gear Number of wheels 4 20

Total load supported

(kN) 498.382 4485.438

Load taken by each

wheel (kN) 124.5955 224.2719

Tyre Pressure (psi) 200 200

Tyre Pressure (bar) 14.28 14.28

Table 7.1 Load Distribution The load is taken by the tyre due to internal pressure,

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Figure 7.1 Typical tyre pressures - taken from Aircraft design: Conceptual Approach by Daniel P. Raymer

7.1.2.Tyre Selection For Nose wheel: Wheel diameter = AWw B (A=1.63, B=0.315) dw =1.63 (27876.30) 0.315 dw = 40.96 in ( 1.0403 m) Wheel width = AWw B (A=0.1043, B=0.480) ww =0.1043 (27876.30) 0.480 ww = 14.190 in (0.3604m)

Figure 7.2 Emprical relations and constants for tyre selection - taken from Aircraft design: Conceptual Approach by Daniel P. Raymer

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52 Contact area:

Figure 7.3 Tyre contact area

From figure the contact area will be neither rectangular nor elliptic but a combination of both. √ ( ) ( √ ) ( √ )

Hence Rt rolling radius for nose landing gear assembly has reduced by 6.667% of the wheel radius

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53

7.1.3. Tyre Selection for Main landing gear: Wheel diameter = AWw B (A=1.63, B=0.315) dw =1.63 (50357.2) 0.315 dw = 49.36 in ( 1.253 m) Wheel width = AWw B (A=0.1043, B=0.480) ww =0.1043 (50357.2) 0.480 ww = 18.848 in (0.4787m) Contact area: √ ( ) ( √ ) ( √ ) 7.2. Runway Loading:

Runway loading estimates for both Main and nose landing wheel:

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54

For a rigid runway:

Nominal working stress on Concrete pavement: 400 psi or 2.75 MN/m2 Concrete Elastic modulus E= 27.5 GPa

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8. Dimensional estimates

8.1. Basic Dimensions:

Span to height ratio

Span to length ratio:

Where length is the length of the fuselage Total length m

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8.2. Configuration of tail:

Figure 8.1 T tail configuration

8.2.1. Horizontal stabilizer: Airfoil used: NACA 0012

Horizontal stabilizer sizing: 15% of wing area A.Rh =4.5

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57

̂

̂ Sweep analysis for the horizontal tail:

Tail will not be affected by downwash since we use T tail Horizontal tail geometry:

Sweep Analysis:

For airfoil NACA 0012

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58

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59

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60 At x/c = 0.125 ( ) M x/c= 0.125 =1 ( ) √ Critical Mach number for the airfoil:

By using a trapezoidal and sweepback we may get

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61

For 1000 km/hr or 277.77 m/s speed we get the sweep to avoid shock on upper surface as

Where is due to taper property

( ⁄ ) For horizontal tail:

By evaluating for λ which varies from 0 to 1 we get:

λh crh cth h (rad) h 1 4.60 4.60 0.00 0.00 0.9 4.84 4.36 0.01 0.76 0.8 5.11 4.09 0.03 1.60 0.7 5.41 3.79 0.04 2.55 0.6 5.75 3.45 0.06 3.60 0.5 6.13 3.07 0.08 4.80 0.4 6.57 2.63 0.11 6.16 0.3 7.08 2.12 0.13 7.72 0.2 7.67 1.53 0.17 9.53 0.1 8.37 0.84 0.20 11.64 0 9.20 0.00 0.25 14.13

Table 8.1 Variation of taper angle of horizontal tail for various taper ratio Taking λh as 0.4

crh = 6.5729 m cth = 5.797 m

Span wise location of MAC:

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62 Hence: 8.2.2. Vertical Stabilizer Geometry:

Vertical stabilizer sizing: 9% of wing area Airfoil used: NACA 0012

A.Rv =0.9 ̂ ̂

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63 λv crv ctv v (rad) v 1 7.981454716 7.981454716 0 0 0.9 8.40153128 7.561378152 0.02299242 1.317407458 0.8 8.868283017 7.094626414 0.048510003 2.779500429 0.7 9.389946724 6.572962707 0.076953376 4.40923372 0.6 9.976818394 5.986091037 0.108800855 6.23401364 0.5 10.64193962 5.32096981 0.144623174 8.286541884 0.4 11.40207817 4.560831266 0.185098527 10.60567718 0.3 12.2791611 3.68374833 0.231024637 13.23712703 0.2 13.30242453 2.660484905 0.283320456 16.23354517 0.1 14.51173585 1.451173585 0.343001791 19.6531346 0 15.96290943 0 0.411098972 23.55493075

Table 8.2 Variation of taper angle of vertical tail for various taper ratio Taking λt as 0.7

crt = 9.389 m

ctt = 6.5729 m = chr

Span wise location of MAC: Hence:

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64

9. Preparation of Layout

Configuration: Anhedral high wing, T-Tail configuration. Nose radius rn= 3m

9.1. Wing Location and CG Estimation:

Reference is taken from nose:

Where X is the location of wing root L.E. from the nose fuselage and Xfinal is the location of cg from L.E at root.

Xfinal = 0.35 (Xcr -Xct) Xfinal =11.59 m

Substituting the values from condition 1 in the above equation:

Fence the wing root L.E. has to be fixed at 14.645 m from nose. That is from nose cg lies at 14.645 +11.69 = 26.223 m

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65

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66 Condition 1 Full Payload and Full Fuel

Fuselage alone analysis

S.No Components

Distance from reference line

(m) Mass (kg) Weight (N) Moment (Nm)

1 Crew 3.043 270 2648.7 8059.9941 2 Nose landing gear 6.086 3600 35316 214933.176 3 Payload bay 1 12.172 37500 367875 4477774.5 4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182 5 Excess mass 22.641 21600 211896 4797537.336 6 Fuselage mass 22.641 43200 423792 9595074.672 7 Fuel in fuselage 22.641 95280.42 934700.9202 21162563.53 8 Main landing gear assembly 1 22.641 10800 105948 2398768.668 9 Main landing gear assembly 2 30.3211 7200 70632 2141639.935 10 Payload bay 2 30.321 37500 367875 11154337.88 11 Horizontal stabilizer 45.367 6400 62784 2848321.728 12 Vertical Stabilizer 48.501 3200 31392 1522543.392 Total 267699.61 2626133.174 60531705.13 Cg from Nose 23.04974695

Wing alone analysis

(m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure 16.35 64000 627840 10265184 2 Wing fuel 1 11.85 74567.285 731505.0659 8668335.03 3 Wing fuel 2 18.428 74567.285 731505.0659 13480175.35 4 Power plant 1 9.57 13600 133416 1276791.12 5 Power plant 2 15 13600 133416 2001240 Total 240334.57 2357682.132 35691725.5 Cg from L.E 15.13848072 Entire aircraft cg from L.E.

Root 11.60343233 Grand total 508034.18 4983815.306 96223430.63

Table 9.1 Cg estimate for fully loaded condition

Similarly for different cases i.e. conditions of loading must be evaluated by fixing the wing at the location x. For this case 1 the cg lies at 27.085 m

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9.2. Three views of Aircraft:

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68

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69

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70 Condition 2 Full Payload and Reserve Fuel

1 Crew 3.043 270 2648.7 8059.9941 2 Nose landing gear 6.086 3600 35316 214933.176 3 Payload bay 1 12.172 0 0 0 4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182 5 Excess mass 22.641 21600 211896 4797537.336 6 Fuselage mass 22.641 43200 423792 9595074.672 7 Fuel in fuselage 22.641 19056.084 186940.184 4232512.707 8 Main landing gear assembly 1 22.641 10800 105948 2398768.668 9 Main landing gear assembly 2 30.3211 7200 70632 2141639.935 10 Payload bay 2 30.321 0 0 0 11 Horizontal stabilizer 45.367 6400 62784 2848321.728 12 Vertical Stabilizer 48.501 3200 31392 1522543.392 Total 116475.274 1142622.438 27969541.93 Cg from Nose 24.47837623

Root 12.54876323 Grand total 237502.188 2329896.464 45942459.12 Table 9.2 Cg estimate for full payload and reserve fuel

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71 Condition 3 Half Payload and Full Fuel

1 Crew 3.043 270 2648.7 8059.9941 2 Nose landing gear 6.086 3600 35316 214933.176 3 Payload bay 1 12.172 18750 183937.5 2238887.25 4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182 5 Excess mass 22.641 21600 211896 4797537.336 6 Fuselage mass 22.641 43200 423792 9595074.672 7 Fuel in fuselage 22.641 95280.42 934700.9202 21162563.53 8 Main landing gear assembly 1 22.641 10800 105948 2398768.668 9 Main landing gear assembly 2 30.3211 7200 70632 2141639.935 10 Payload bay 2 30.321 18750 183937.5 5577168.938 11 Horizontal stabilizer 45.367 6400 62784 2848321.728 12 Vertical Stabilizer 48.501 3200 31392 1522543.392 Total 230199.61 2258258.174 52715648.94 Cg from Nose 23.34349967

Root 12.00007718 Grand total 470534.18 4615940.306 88407374.45 Table 9.3 Cg estimate for half payload and full fuel

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72 Condition 4 Half Payload and Reserve Fuel

Root 11.74118831 Grand total 275002.188 2697771.464 53758515.31 Table 9.4 Cg estimate for half payload and reserve fuel

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73 Condition 5 No Payload and Reserve Fuel

1 Crew 3.043 270 2648.7 8059.9941 2 Nose landing gear 6.086 3600 35316 214933.176 3 Payload bay 1 12.172 0 0 0 4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182 5 Excess mass 22.641 21600 211896 4797537.336 6 Fuselage mass 22.641 43200 423792 9595074.672 7 Fuel in fuselage 22.641 19056.084 186940.184 4232512.707 8 Main landing gear assembly 1 22.641 10800 105948 2398768.668 9 Main landing gear assembly 2 30.3211 7200 70632 2141639.935 10 Payload bay 2 30.321 0 0 0 11 Horizontal stabilizer 45.367 6400 62784 2848321.728 12 Vertical Stabilizer 48.501 3200 31392 1522543.392 Total 116475.274 1142622.438 27969541.93 Cg from Nose 24.47837623

Root 12.54876323 Grand total 237502.188 2329896.464 45942459.12 Table 9.5 Cg estimate for No payload and reserve fuel

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74 Condition 6 Full Payload and Half Fuel

1 Crew 3.043 270 2648.7 8059.9941 2 Nose landing gear 6.086 3600 35316 214933.176 3 Payload bay 1 12.172 37500 367875 4477774.5 4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182 5 Excess mass 22.641 21600 211896 4797537.336 6 Fuselage mass 22.641 43200 423792 9595074.672 7 Fuel in fuselage 22.641 47640.21 467350.4601 10581281.77 8 Main landing gear assembly 1 22.641 10800 105948 2398768.668 9 Main landing gear assembly 2 30.3211 7200 70632 2141639.935 10 Payload bay 2 30.321 37500 367875 11154337.88 11 Horizontal stabilizer 45.367 6400 62784 2848321.728 12 Vertical Stabilizer 48.501 3200 31392 1522543.392 Total 220059.4 2158782.714 49950423.36 Cg from Nose 23.13823575

Root 11.36246959 Grand total 385826.685 3784959.78 74567893.67 Table 9.6 Cg estimate for full payload and half fuel

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75 Condition 7 Half Payload and Half Fuel

Root 11.87233251 Grand total 348326.685 3417084.78 66751837.49 Table 9.7 Cg estimate for half payload and half fuel

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9.3. The variation of cg location is shown below

S.No Details CG % of MAC

Variation in percentage

1 Full payload + Full fuel 0.3500 0.0000 2 No payload + Full fuel 0.3794 7.7421 3

Full payload + Reserve

fuel 0.3819 8.3523

4 Half payload + Full fuel 0.3652 4.1671 5

Half payload + Reserve

fuel 0.3573 2.0531

6

No payload + Reserve

fuel 0.3819 8.3523

7 Full payload + Half fuel 0.3819 8.3523 8 Half payload + Half fuel 0.3613 3.1366

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10. Drag Estimation

Drag Equation for Entire Aircraft:

Where

10.1. Component Drag Estimates:

Wetted surface area:

Fuselage: Engine: Horizontal Stablizer:

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Vertical Stablizer:

Nose Landing Gear:

Main Landing Gear 1:

Main Landing Gear 2:

Flap:

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10.2. Total Drag Estimate:

S.No Components Wetted Surface area Permanent components 1 Fuselage 56 0.03 1.68 2 Engine 31.3706 0.03 0.941118 3 Horizontal Stablizer 95.256 0.0052 0.495331 4 Vertical Stablizer 57.154 0.0052 0.297201 3.41365 0.005375 Temporary Components

a Nose Landing Gear 1.4996 0.12 0.179952 b Main Landing Gear 1 7.197 0.12 0.86364 c Main Landing Gear 2 4.7984 0.12 0.575808

Landing Gear total 1.6194 0.00255 i Flap at 45° 75.645 0.016 1.21032 0.001906 ii Flap at 60° 75.645 0.02 1.5129 0.002382

Table 10.1 Coefficient of Drag for different parts of aircraft

Takeoff Performance: Landing Performance: Cruise Performance:

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10.3. Drag Polar

Drag Polar Analysis:

Takeoff Landing Cruise

CL KCL2 CD Takeoff CD Landing CD Cruise (L/D)cruise

-0.5 0.011841886 0.027172886 0.027648886 0.022716886 -22.0100592 -0.4 0.007578807 0.022909807 0.023385807 0.018453807 -21.6757443 -0.3 0.004263079 0.019594079 0.020070079 0.015138079 -19.8175741 -0.2 0.001894702 0.017225702 0.017701702 0.012769702 -15.662073 -0.1 0.000473675 0.015804675 0.016280675 0.011348675 -8.81160102 0 0 0.015331 0.015807 0.010875 0 0.1 0.000473675 0.015804675 0.016280675 0.011348675 8.81160102 0.2 0.001894702 0.017225702 0.017701702 0.012769702 15.662073 0.3 0.004263079 0.019594079 0.020070079 0.015138079 19.8175741 0.4 0.007578807 0.022909807 0.023385807 0.018453807 21.6757443 0.5 0.011841886 0.027172886 0.027648886 0.022716886 22.0100592 0.6 0.017052315 0.032383315 0.032859315 0.027927315 21.4843421 0.7 0.023210096 0.038541096 0.039017096 0.034085096 20.5368352 0.8 0.030315228 0.045646228 0.046122228 0.041190228 19.4220826 0.9 0.03836771 0.05369871 0.05417471 0.04924271 18.2768171 1 0.047367543 0.062698543 0.063174543 0.058242543 17.16958 1.1 0.057314727 0.072645727 0.073121727 0.068189727 16.1314621 1.2 0.068209262 0.083540262 0.084016262 0.079084262 15.1736891 Table 10.2 Coefficient of Drag for different flying conditions

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Drag Polar Curve:

Figure 10.1 Drag Polar curve for Entire Aircraft during takeoff, cruise and Landing

10.4. Lift to Drag Ratio:

Figure 10.2 L/D vs. CL for Entire aircraft at cruise

(L/D)max= 22 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 CL CD

Drag Polar

Takeoff Cruise Landing 0 5 10 15 20 25 0 0.5 1 1.5 2 2.5 3 L/ D CL

L/D

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11. Performance Calculations:

11.1. Thrust required and Thrust available analysis:

W1= 25% of Fuel and 100 % of Payload W1= 3185533.292 N

Figure 11.1 Thrust scenarios at S ea level for different weights

-500 0 500 1000 1500 2000 2500 0 200 400 600 Thr us t in N Thou sa nds Velocity

Thrust values at sea level

Thrust Available Tr at w1 Tr at w2 Tr at w3

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Figure 11.2 Thrust scenarios at 11 km altitude for different weights

Figure 11.3 Thrust scenarios at 25 km for different weights

0 100 200 300 400 500 600 700 800 900 0 100 200 300 400 500 Th ru st i n N Thou sa nds Velocity

Thrust values at 11 km

Thrust avaliable Tr at w1 tr at w2 Tr at w3 0 500 1000 1500 2000 2500 3000 0 100 200 300 400 500 Thr us t in N Thou sa nds Velocity

Thrust values at 25 km

Thrust avaliable Tr at w1 tr at w2 Tr at w3

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Reference:

Websites:  www.airminded.org  www.aviationexplorers.com  www.flightglobal.com  www.jet-engine.net  www.wikipedia.org  www.worldaircraftsearch.com  www.worldofkrauss.com Books:

1. Ajoy Kumar Kundu (2010) ―Aircraft Design – Cambridge Aerospace Series‖ 2. Daniel. P. Raymer (1989) ―Aircraft Design: A Conceptual Approach‖

3. Lloyd Jekinson & Jim Marchman (2003) ―Aircraft Design Projects For Engineering Students‖ 4. FAA ―Aircraft Weight and Balance Handbook‖

5. Jan Roskam (1985) ―Airplane Design-Part 1: Preliminary Design and Sizing and Part 4: Layout design of Landing Gear and system‖