Three major types of airplane designs are 1. Conceptual design 2. Preliminary design 3. Detailed design

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1 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: Preliminary is based upon number of factors like Loitering. 1

2 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 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. Later several number of improvements were made. 2

3 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 Preferred Configuration: Figure 1.1 High wing Configuration with T tail 3

4 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 kg was collected from the appropriate resources. 4

5 2.1 Comparative configuration study of bomber airplanes: S.No 1 Name of the aircraft Collection of Comparative Data -Dimension 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) Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu- 142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.1 Collection of Comparative Data -Dimension 5

6 Collection of Comparative Data -Performance parameters S.No 1 2 Name of the aircraft Thrust to weight ratio Wing loading (N/m 2 ) Aspect ratio Power Plant Mirage IIIE SNECMA Atar 09C turbojet Mirage IVA F-111F F-111F Swept Tu-22R SNECMA Atar 9K-50[13] turbojets Dry thrust: kn (11,023 lbf) each Thrust with afterburner: kn (15,873 lbf) each Pratt & Whitney TF30-P-100 turbofans Thrust with afterburner: 25,100 lbf (112 kn) each Dry thrust: 17,900 lbf (79.6 kn) each Pratt & Whitney TF30-P-100 turbofans Thrust with afterburner: 25,100 lbf (112 kn) each Dry thrust: 17,900 lbf (79.6 kn) each Dobrynin RD-7M-2 turbojets Dry thrust: rated kn (24,250 lbf) each Thrust with afterburner: kn (36,376 lbf) each Dobrynin VD-4K turbo-compound radial 6 Tu-85/ engines, 3,200 kw (4,300 hp) each 7 YB Pratt & Whitney J57-P-3 turbojets, (38 kn) each 8 B-2A Tu- 142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept General Electric F118-GE-100 non-afterburning turbofans, 17,300 lbf (77 kn) each Kuznetsov NK-12MV turboprops, 11,033 kw (14,795 shp) each Kuznetsov NK-12M turboprops, 11,000 kw (14,800 shp)[23] each General Electric F101-GE-102 augmented turbofans Dry thrust: 14,600 lbf (64.9 kn) each Thrust with afterburner: 30,780 lbf ( kn) each General Electric F101-GE-102 augmented turbofans Dry thrust: 14,600 lbf (64.9 kn) each Thrust with afterburner: 30,780 lbf ( kn) each Pratt & Whitney TF33-P-3/103 turbofans, (76 kn) each Samara NK-321 turbofans Dry thrust: kn (30,865 lbf) each Thrust with afterburner: 245 kn (55,115 lbf) each Samara NK-321 turbofans Dry thrust: kn (30,865 lbf) each Thrust with afterburner: 245 kn (55,115 lbf) each Table 2.2 Collection of Comparative Data -Performance parameters 6

7 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 Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.3 Collection of Comparative Data -Performance parameters (Cont.) 7

8 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 In service Good 2 Mirage IVA Good 3 F-111F NACA root, NACA tip Good 4 F-111F Swept NACA root, NACA tip Good 5 Tu-22R Good 6 Tu-85/ Prototype Good 7 YB In service Good 8 B-2A In service Good 9 Tu-142M In service Awesome 10 Tu-95MS In service Awesome 11 B-1B NA In service Good 12 B-1B Swept NA In service Good 13 B-52H NACA 63A219.3 mod root, NACA 65A209.5 tip In service Awesome 14 Tu In service Good 15 Tu-160 Swept In service Good Table 2.4 Collection of Comparative Data -Performance parameters (Cont.) 8

9 Wing Loading (N/sq.m) 2.2 Comparative graphs for determining optimum value: S.No Name of the aircraft Wing loading (N/m 2 ) Maximum Speed (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.5 Wing loading vs. Maximum Speed Wing Loading (N/sq.m) Vs Maximum Speed (m/s) Wing loading (N/m2) Maximum Speed (m/s) Figure 2.1 Wing loading vs. Maximum Speed 9

10 Span to length Ratio S.No Name of the aircraft Span to length ratio Maximum Speed (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.6 Span to length Ratio vs. Maximum Speed 3.00 Span to length ratio Vs Maximum Speed (m/s) Span to length ratio Maximum Speed (m/s) Figure 2.2 Span to length Ratio vs. Maximum Speed 10

11 Aspect Ratio S.No Name of the aircraft Aspect ratio Maximum Speed (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.7 Aspect Ratio vs. Maximum Speed 14 Aspect ratio Vs Maximum Speed (m/s) Aspect ratio Maximum Speed (m/s) Figure 2.3 Aspect Ratio vs. Maximum Speed 11

12 Wing Area (sq.m) Maximum Speed S.No Name of the aircraft Wing Area (m 2 ) (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.8 Wing Area vs. Maximum Speed 600 Wing Area (sq.m) Vs Maximum Speed (m/s) Wing Area (m2) Maximum Speed (m/s) Figure 2.4 Wing area vs. Maximum Speed 12

13 Combat radius (km) S.No Name of the aircraft Combat Radius (km) Maximum Speed (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.9 Span to length Ratio vs. Maximum Speed 8000 Combat Radius (km) Vs Maximum Speed (m/s) Combat Radius (km) Maximum Speed (m/s) Figure 2.5 Combat radius vs. Maximum Speed 13

14 Payload Capacity (kg) S.No Payload Capacity Maximum Speed Name of the aircraft (kg)2 (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.10 Payload capacity vs. Maximum Speed Payload Capacity (kg) Vs Maximum Speed (m/s) Payload Capacity (kg) Maximum Speed (m/s) Figure 2.6 Payload Capacity vs. Maximum Speed 14

15 Thrust to weight ratio S.No Thrust to weight Maximum Speed Name of the aircraft ratio (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.11 Thrust to weight Ratio vs. Maximum Speed 0.70 Thrust to Weight Ratio Vs Maximum Speed (m/s) Thrust to weight ratio Maximum Speed (m/s) Figure 2.7 Thrust to weight Ratio vs. Maximum Speed 15

16 Span to Height Ratio S.No Name of the aircraft Span to height ratio Maximum Speed (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.12 Span to height Ratio vs. Maximum Speed Span to Height Ratio Vs Maximum Speed (m/s) Span to height ratio Maximum Speed (m/s) Figure 2.8 Span to height Ratio vs. Maximum Speed 16

17 Maximum Takeoff Weight (kg) S.No Name of the aircraft Maximum Weight (kg) Takeoff Maximum Speed (m/s) 1 Mirage IIIE Mirage IVA F-111F F-111F Swept Tu-22R Tu-85/ YB B-2A Tu-142M Tu-95MS B-1B B-1B Swept B-52H Tu Tu-160 Swept Table 2.13 Maximum takeoff weight vs. Maximum Speed Maximum Takeoff Weight (kg) Vs Maximum Speed (m/s) Maximum Takeoff Weight (kg) Maximum Speed (m/s) Figure 2.9 Maximum takeoff weight vs. Maximum Speed 17

18 2.2 Parameter Selection: From Comparison (Assumed and extrapolated values from graph) Maximum takeoff weight (kg) 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) Maximum speed (kmph) 1000 Service ceiling (m) Maximum speed (m/s) Table 2.14 Parameter Selection 18

19 3. Rough Weight Estimate Optimal values of mass fraction for bombers Parameter Range of values Notation Empty Mass Ratio M E /M TO Total Fuel Mass Ratio M F /M TO Payload Ratio M Pay /M TO Wing Loading N/m 2 W o /S Thrust to Weight Ratio T/W o Table 3.1 Mass Fraction Parameters for bomber 3.1 General rough weight estimate: 0.4 M Pay /M To Vs Maximum Takeoff weight (kg) M Pay /M To MPay/MTo Maximum Takeoff weight (kg) Figure 3.1 Payload mass Fraction vs. Maximum Takeoff weight 19

20 = kg 0.6 M E /M To Vs Maximum Takeoff weight (kg) M E /M To ME/Mto Maximum Takeoff weight (kg) Figure3.2 Empty mass Fraction vs. Maximum Takeoff weight 20

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 Table 3.2 Values from rough weight estimate 21

22 4. Redefined Mass Estimation 4.1 Mission profile analysis Profile 1: Strategic bombing mission h km km 2 R 1000 km km 7 1/2 hr Figure 4.1 Mission profile for Strategic bombing Analysis of Mission Profile: Warmup and takeoff Climb or descend Landing ( ) ( ) 22

23 Figure 4.2 Empty mass fraction vs takeoff mass -- taken from "Aircraft Design: A Conceptual Approach" by Daniel P.Raymer Analysis of Mission Profile Cruise TSFC values for Bomber Loiter 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) 23

24 ( ( ) ) ( ) ( ) 4.2: Before Refueling Part 1: Before refueling: h km 3 R 0 1 Figure 4.3 Mission profile 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 = (kg / N-s) 24

25 Range R 2-3 = 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 km can be interpreted as a combat radius of 5000 km 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 25

26 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. 26

27 4.4. After Refueling: Part 2: After refueling: h km km R 1000 km Figure 4.5 After refueling mission profile Cruise at 60% of maximum speed: (R-2) Descend: (2-3 ) Cruise: (3-4 ) Bombing (4 ) Climb: (4-5 ) 27

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: 28

29 Total fuel mass fraction after refueling: ( ) ( ) ( ) ( ) Hence ( ) is taken since the value turns out to be ( ) 29

30 Replacing as X in excel to solve the implicit function X ranges from to since initial mass estimate is kg X X f(x) f(x) X X f(x) f(x) X X f(x) f(x) X f(x) Thus mass of the aircraft is kg 30

31 Take-off Weight of the aircraft: 4.5.Thrust Estimation T= = N T= KN 31

32 T A = KN T = T= KN/Engine 32

33 5. Power Plant Selection 5.1. Comparative data of engines Thrust Estimation: ( ) ( ( For the chosen parameters: ( ) ( ) ) ) T= = N T= KN T A = KN T = T= KN/Engine 33

34 The performance data of various turbofan engines with thrust of range 330 kn to 500 kn were collected from the following resources Wet weight (kg) Dry Weight (kg) Name of Length Diameter S.No the Engine Manufacturer Type (m) (m) 1 Trent-900 Rolls Royce Turbofan 3 Shaft GP-7000 Engine Alliance Turbofan 2 Shaft GE 90-76B GE Turbofan 2 Shaft GE 90-92B GE Turbofan 2 Shaft GP 7270 Engine Alliance Turbofan 2 Shaft GE B1 GE Turbofan 2 Shaft GP 7277 Engine Alliance Turbofan 2 Shaft PW Pratt & Whitney Turbofan 2 Shaft GE 90-85B GE Turbofan 2 Shaft GE 90-94B GE Turbofan 2 Shaft GE 90-90B GE Turbofan 2 Shaft GE B GE Turbofan 2 Shaft 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 GP GE 90-76B GE 90-92B GP GE90-110B GP PW GE 90-85B GE 90-94B GE 90-90B GE B Table 5.2. Collection of Engine Comparative Data (Cont.) 34

35 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 = N/N - hr = = kg/n- hr TSFC = kg/n -s Service ceiling evaluation: By taking service ceiling as h=15 km 35

36 Number of Engines = 4 36

37 6. Airfoil selection and Wing Geometry estimates 6.1. Main Parameter Selection: Wing Loading: 6.2 Fuel volume consideration: ρ F can vary from 600 kg/m 3 to 800 kg/m 3. For ρ F = 800 kg/m 3 37

38 Volume of fuel accommodated in wing: ( (( ) )) ( ) ( ( )) ( ) Selecting NACA airfoil of fineness ratio (t/c ratio) as 0.18 ( ) ( ) ( ) 38

39 6.3 Takeoff Analysis: Figure 6.1 Runway length survey for military installations S R = 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 v i is initial velocity during takeoff. 39

40 ( ) ( ) Where value can be denoted as t also since we use MAC to obtain the thickness Thickness based Reynolds Number: 40

41 Figure 6.2 Cl vs Angle of attack curve for NACA at angle of attack 0.5 deg 41

42 Symbol Re x/c y/c Angle of attack Table 6.1 Airfoil data at various Re. At Re = by interpolating we get Location of aerodynamic center x/c = y/c = C L max = 1.48 α = Flap selection: C L max Flap Chosen is Triple slotted flap Wing setting angle or incidence angle i w = 3 degree Required Flap Deflection = 60 Change in C L due to flap deflection: 42

43 Figure 6.3 Drag polar curve fornaca at angle of attack 0.5 deg 43

44 6.5. Wing geometry: Sweep Analysis: For airfoil NACA At x/c = 0.46 ( ) Figure 6.4 Variation of local velocity with the free stream velocity 44

45 Hence by comparing v and a MSL formed. and a alt it is clear that the shock wave is 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 ( ) ( ) 45

46 Critical Mach number for the airfoil: Figure 6.5 Swept back wing By using a trapezoidal and sweepback we may get 46

47 For 1000 km/hr or 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 : ( ) 47

48 Where λ is taper ratio: λ can vary from 0 to 1. By evaluating the above four equations we get : λ c r (m) c t (m) (rad) (deg) Table 6.2 Angle of taper for various taper ratios Taking the value λ = 0.5 c r = m c t = m Span wise location of MAC: Hence: 48

49 Figure 6.6 Effect of aspect ratio on lift curve slope a= /degree for a = 0.15/ degree 49

50 7. Landing gear design 7.1. Tyre selection: 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 = = kn Load taken by wheels in main landing gear = = kn Nose Landing Gear Main Landing Gear Number of wheels 4 20 Total load supported (kn) Load taken by each wheel (kn) Tyre Pressure (psi) Tyre Pressure (bar) Table 7.1 Load Distribution The load is taken by the tyre due to internal pressure, 50

51 Figure 7.1 Typical tyre pressures - taken from Aircraft design: Conceptual Approach by Daniel P. Raymer Tyre Selection For Nose wheel: Wheel diameter = AW B w (A=1.63, B=0.315) d w =1.63 ( ) d w = in ( m) Wheel width = AW B w (A=0.1043, B=0.480) w w = ( ) w w = in (0.3604m) Figure 7.2 Emprical relations and constants for tyre selection - taken from Aircraft design: Conceptual Approach by Daniel P. Raymer 51

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 R t rolling radius for nose landing gear assembly has reduced by 6.667% of the wheel radius 52

53 Tyre Selection for Main landing gear: Wheel diameter = AW B w (A=1.63, B=0.315) d w =1.63 ( ) d w = in ( m) Wheel width = AW B w (A=0.1043, B=0.480) w w = ( ) w w = in (0.4787m) Contact area: ( ) ( ) ( ) 7.2. Runway Loading: Runway loading estimates for both Main and nose landing wheel: 53

54 For a rigid runway: Nominal working stress on Concrete pavement: 400 psi or 2.75 MN/m 2 Concrete Elastic modulus E= 27.5 GPa 54

55 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 55

56 8.2. Configuration of tail: Figure 8.1 T tail configuration Horizontal stabilizer: Airfoil used: NACA 0012 Horizontal stabilizer sizing: 15% of wing area A.R h =4.5 56

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 Figure 8.2 Local velocity vs free stream velocity for NACA 0012 airfoil 57

58 Figure 8.3 CL vs angle of attack curve for NACA

59 Figure 8.4 Drag polar curve for NACA

60 At x/c = ( ) M x/c= =1 ( ) Critical Mach number for the airfoil: By using a trapezoidal and sweepback we may get 60

61 For 1000 km/hr or 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: λ h c rh c th h (rad) h Table 8.1 Variation of taper angle of horizontal tail for various taper ratio Taking λ h as 0.4 c rh = m c th = m Span wise location of MAC: 61

62 Hence: Vertical Stabilizer Geometry: Vertical stabilizer sizing: 9% of wing area Airfoil used: NACA 0012 A.R v =0.9 62

63 λ v c rv c tv v (rad) v Table 8.2 Variation of taper angle of vertical tail for various taper ratio Taking λ t as 0.7 c rt = m c tt = m = c hr Span wise location of MAC: Hence: 63

64 9. Preparation of Layout Configuration: Anhedral high wing, T-Tail configuration. Nose radius r n = 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 X final is the location of cg from L.E at root. X final = 0.35 (Xc r -Xc t ) X final =11.59 m Substituting the values from condition 1 in the above equation: Fence the wing root L.E. has to be fixed at m from nose. That is from nose cg lies at = m 64

65 Figure 9.1 Wing Details for cg estimates 65

66 S.No S.No Condition 1 Full Payload and Full Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total 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 m 66

67 9.2. Three views of Aircraft: Figure 9.2 Top View 67

68 Figure 9.3 Front View 68

69 Figure 9.4 Side view 69

70 S.No Condition 2 Full Payload and Reserve Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line S.No Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total Table 9.2 Cg estimate for full payload and reserve fuel 70

71 S.No Condition 3 Half Payload and Full Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line S.No Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total Table 9.3 Cg estimate for half payload and full fuel 71

72 S.No Condition 4 Half Payload and Reserve Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line S.No Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total Table 9.4 Cg estimate for half payload and reserve fuel 72

73 S.No Condition 5 No Payload and Reserve Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line S.No Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total Table 9.5 Cg estimate for No payload and reserve fuel 73

74 S.No Condition 6 Full Payload and Half Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line S.No Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total Table 9.6 Cg estimate for full payload and half fuel 74

75 S.No Condition 7 Half Payload and Half Fuel Fuselage alone analysis Distance from reference line Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Crew Nose landing 2 gear Payload bay Fixed equipments Excess mass Fuselage mass Fuel in fuselage Main landing 8 gear assembly Main landing 9 gear assembly Payload bay Horizontal stabilizer Vertical Stabilizer Total Cg from Nose Wing alone analysis Distance from reference line S.No Components (m) Mass (kg) Weight (N) Moment (Nm) 1 Wing structure Wing fuel Wing fuel Power plant Power plant Total Cg from L.E Entire aircraft cg from L.E. Root Grand total Table 9.7 Cg estimate for half payload and half fuel 75

76 9.3. The variation of cg location is shown below S.No Details CG % of MAC Variation percentage in 1 Full payload + Full fuel No payload + Full fuel Full payload + Reserve 3 fuel Half payload + Full fuel Half payload + Reserve fuel No payload + Reserve 6 fuel Full payload + Half fuel Half payload + Half fuel Table 9.8 Cg estimate for various conditions 76

77 10. Drag Estimation Drag Equation for Entire Aircraft: Where Component Drag Estimates: Wetted surface area: Fuselage: Engine: Horizontal Stablizer: 77

78 Vertical Stablizer: Nose Landing Gear: Main Landing Gear 1: Main Landing Gear 2: Flap: 78

79 10.2. Total Drag Estimate: S.No Wetted Surface Components area Permanent components 1 Fuselage Engine Horizontal Stablizer Vertical Stablizer Temporary Components a Nose Landing Gear b Main Landing Gear c Main Landing Gear Landing Gear total i Flap at ii Flap at Table 10.1 Coefficient of Drag for different parts of aircraft Takeoff Performance: Landing Performance: Cruise Performance: 79

80 10.3. Drag Polar Drag Polar Analysis: Takeoff Landing Cruise C L 2 KC L C D Takeoff C D Landing C D Cruise (L/D) cruise Table 10.2 Coefficient of Drag for different flying conditions 80

81 L/D Drag Polar Curve: Drag Polar CL C D Takeoff Cruise Landing Figure 10.1 Drag Polar curve for Entire Aircraft during takeoff, cruise and Landing Lift to Drag Ratio: L/D 10 L/D CL Figure 10.2 L/D vs. CL for Entire aircraft at cruise (L/D) max = 22 81

82 Thrust in N Thousands 11. Performance Calculations: Thrust required and Thrust available analysis: W 1 = 25% of Fuel and 100 % of Payload W 1 = N W 2 = 50% of Fuel and 100 % of Payload W 2 = N W 3 = 75% of Fuel and 100 % of Payload W 3 = N 2500 Thrust values at sea level Velocity Thrust Available Tr at w1 Tr at w2 Tr at w3 Figure 11.1 Thrust scenarios at Sea level for different weights 82

83 Thrust in N Thousands Thrust in N Thousands Thrust values at 11 km Thrust avaliable Tr at w1 tr at w2 Tr at w Velocity Figure 11.2 Thrust scenarios at 11 km altitude for different weights Thrust values at 25 km Thrust avaliable Tr at w1 tr at w2 Tr at w3 Velocity Figure 11.3 Thrust scenarios at 25 km for different weights 83

84 Reference: Websites: 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 6. Irs.H. Abbott & Albert E Von Doenhoff (1949) Theory of Wing Sections 84