U7AEB36-AIRCRAFT DESIGN PROJECT-II
MULTIROLE FIGHTER AIRCRAFT
A PROJECT REPORT
Submitted By COUTINHO VARNEY PLATO
13UEAE0018
K.ELUMALAI
13UEAE0021
S.GURUPRASAATH
13UEAE0026
D.DALJIT MAJIL
13UEAE0501
In partial fulfillment fulfillment for the award of the degree degree Of BACHELOR OF TECHNOLOGY IN AERONAUTICAL ENGINEERING
VEL TECH DR.RR & DR.SR TECHNICAL UNIVERSITY CHENNAI-6000062
DECEMBER 2016
Bonafide Certificate Certificate
This is to certify that the project work entitled “Multirole Fighter Aircraft” in partial fulfillment of the requirement of the award of Degree of Bachelor of Technology in Aeronautical Engineering of Vel Tech Dr. RR & Dr. SR Technical University, Chennai – 600 600 062, is an authentic work carried out by Coutinho Varney Plato (Reg. No. 13UEAE0018), K.Elumalai (Reg. No. 13UEAE0021), S.Guruprasaath (Reg. No.
our supervisions and guidance. 13UEAE0026) and D.Daljit Majil (Reg. No 13UEAE0501) under our supervisions To the best of my knowledge, the matter embodied in the project report has not been submitted to any other University/Institute for the award of any Degree or Diploma
R.Jaganraj
Kannan.G
Head of the Department,
Assistant Professor,
Dept of Aeronautical Engineering
Dept of Aeronautical Engineering
Vel Tech Dr RR & Dr SR
Vel Tech Dr RR & Dr SR
Technical University, Avadi,
Technical University, Avadi,
Chennai 600 062
Chennai 600 062
Certificate of Evaluation
University: Vel Tech Dr. RR & Dr.SR Technical Techn ical University Branch : Aeronautical Engineering Semester: VII
S.No
VTU NO
REG. NO
NAME
1.
4180
13UEAE0018
COUTINHO VARNEY PLATO
2.
4364
13UEAE0021
K.ELUMALAI
PROJECT TITLE
MULTIROLE FIGHER
3.
4094
13UEAE0026
S.GURUPRASAATH
4.
5726
13UEAE0501
D.DALJIT MAJIL
AIRCARFT
PROJECT GUIDE G.KANNAN,
ASST. PROFESSOR, DEPT. OF AERONAUTI CAL ENGIG
The report of the project work submitted by b y the above student in partial fulfillment for the award of Degree of Bachelor of Technology in Aeronautical Engineering of Vel Tech Dr. RR & Dr. SR Technical University was evaluated and confirmed to be the report of the work done by
the above student.
This project report was submitted for VIVA VOICE held on . . . . . . . . . . . . . . . . …………….. at VEL TECH Dr. RR & Dr. SR TECHNICAL UNIVERSITY, AVADI.
Internal Examiner
External examiner
Date …………
ACKNOWLEDGEMENT
First of all I would like to express my deepest gratitude to VEL TECH Dr. RR & Dr. SR TECHNICAL UNIVERSITY for giving me this tremendous opportunity. opportunity. I would like to express gratitude to Founder- President Prof Dr R Rangarajan B.E (Elec.), B.E (Mech.) M.S (Auto), D.Sc. for giving me the opportunity to be the part of this Institution. I would like to acknowledge Founder-Vice President and our Chancellor Dr.Sagunthala Rangarajan (MBBS) for her support. I would further like to express my gratitude to Chairperson and Managing Trustee Dr.Rangarajan Mahalakshmi Mahalakshmi K.B.E (IE) M.B.A (UK) Ph.D. I would also like to express my deepest thanks to Vice President Mr. K.V.D Kishore Kumar.
I would further like to thank our Vice- Chancellor Dr.Beela Satynarayan B.E (Mech.), M.E (MD), M.E (IE) M. Tech (CSE), Ph.D. (IIT Delhi) I would like to express my gratitude to our Registrar Dr.E.Kannan M.E, Ph.D., PGDSM (Hons.) I would like to thank Dr.A.T.Ravichandran Ph.D. Dean School of Mechanical for his Constant support. I would also like to express my deepest gratitude to Mr.R.JAGANRAJ Head of the Department (Aeronautical Department) for his valuable suggestions. Finally I would like to express my deepest gratitude to Mr.G.KANNAN Asst. Professor for helping me throughout the project and sharing his valuable knowledge.
CONTENTS INTRODUCTION ABSTRACT LIST OF SYMBOLS LIST OF TABLES LIST OF GRAPHS LIST OF FIGURES 1. REQUIRED DATA FROM ADP-1 2. V-N DIAGRAM 2.1 INTODUCTION 2.2 LOAD FACTOR 3. GUST ENVELOPE 4. CRITICAL PERFORMANCE PARAMETERS 4.1 CRUISING FLIGHT PERFORMANCE 4.2 TAKEOFF PERFORMANCE 4.3 LANDING PERFORMANCE 4.4 GROUND ROLL 5. COMBINED V-N DIAGRAM (OR) FLIGHT ENVELOPE 6. WING LOAD DISTRIBUTION 6.1 WING DESCRIPTION DESCRIPTION 6.2 LINEAR LIFT DISTRIBUTION 6.3 ELLIPTIC LIFT DISTRIBUTION DISTRIBUTION 6.4 SCHRENKS CURVE 6.5 LOAD ESTIMATION ON WINGS 6.6 FUEL WEIGHT 6.7 REACTION AND BENDING MOMENT CALCULATION 6.8 SHEAR FORCE 6.9 BENDING MOMENT 7. CG CALCULATION 8. MATERIAL SELECTION 8.1 DESCRIPTION 8.2 WOOD 8.3 ALUMINIUM ALLOY 8.4 EXTRUDED ALUMINIUM ALLOYS 8.5 STEEL 8.6 COMPOSITE MATERIALS 8.7 HEAVY AIRCRAFT RAW MATERIALS 9. WING DESIGN 9.1INTRODUCTION 9.2 AIRFOIL SELECTION 9.3 WING SLECTION 10. FUSELAGE DESIGN 11. DETAILED STRUCTURAL LAYOUT 11.1 FUNCTIONS OF THE STRUCTURE 11.2 WING STRUCTURAL LAYOUT 11.3 BASIC FUNCTIONS OF WING STRUCTURAL MEMBERS 11.4 FUSELAGE STRUCTURE 11.5 FUSELAGE LAYOUT CONCEPTS
1 3 3 5 6 10 10 12 13 14 15 16 16 16 18 20 21 23 25 26 27 29 31 31 31 32 32 32 33 33 34 34 34 39 42 46 46 46 47 48 48
12. THREE-VIEW DIAGRAM 13. CONCLUSION REFERENCES
50 52
INTRODUCTION Project Project Ai m:Main Objective of the Project is to design a Multi role fighter Aircraft that can perform different roles in combat. A term Multirole means for Aircraft designed for complete different tasks with same Airframe. Main motivation of Multirole fighter is to reduce the cost by using a Common airframe for different tasks. Multirole fighter aircraft will have tasks such as Aerial reconnaissance, Forward Air Control and Electronic Warfare Aircraft. Attack missions include the subtypes air interdiction, suppression interdiction, suppression of enemy air defense (SEAD), and close air support (CAS).It also have a capability of STOL(Short Range Takeoff and Landing) because of which Aircraft needs Shorter length of Runway.
Pre Pr esent F i ghter ghter s: Fighter Aircrafts are the aircrafts used only for the defense purpose of the country. There are different types of fighter aircrafts depending on the mission to accomplish some of them are Interceptor, Bomber, th th Dogfight, reconnaissance etc. The present time fighters are of 4 , 4.5th and 5 generation fighter Aircrafts. The Specialty of them is Stealth, Super cruise, STOL, Multirole etc. The fifth generation fighters are completely stealth fighters capable of operating at different atmospheric condition. Even though there are no bombers in the fifth generation the multirole fighters it acts as a bomber. The stealth Aircraft is an ideal Aircraft for reconnaissance. Some of the Fifth generation planes are F-35 lightening, F-22 Raptor, Su-30 etc. F-35 lightening is a VTOL aircraft with stealthy body whereas F-22 Raptor is a STOL aircraft with both stealth body and stealth coatings. The Stealth coating (radiation Absorbing paints makes the aircraft’s Maintenance charge more than anyone else of its kind.
Abstract The Current scenario in Aerial Combat requires an Aircraft that can perform multirole tasks to complete various missions with same airframe. The report summarizes the design of an aircraft with its design parameters and design considerations. The design includes the blend wing stealth technology which can perform multiple roles with greater flexibility. Aircraft is designed with a capability of carrying payload up to 8000kg that includes missile ( Air to Air & Air to ground), Bombs, Guns etc. The Huge amount of thrust allows the aircraft to attain STOL along TVC and Reach the cruise altitude in minimum time. The design has fully variable inlet and Nozzle for good good performance of the engine at various various speeds. speeds.
LIST OF SYMBOLS AR
Aspect ratio
B
Span
C
Chord
CG
Centre of Gravity
CD
Coefficient of Drag
CL
Coefficient of Lift
D
Drag
L
Lift
M
Mach Number
R
Range
S
Surface are of wing
T
Thrust
V
Velocity
W
Weight
Sg
Takeoff Distance
P
Power
LIST OF TABLES Table.1.1 Details weight from ADP-I
1
Table 1.2 Required Parameters from ADP-I
1
Table 2.1. Load factor for various types of aircraft
5
Table.6.1 Loads Simplified as Point loads
25
LIST OF GRAPHS Graph.2.1 V-n Diagram
4
Graph 3.1 Gust Load
9
Grapph 5.1 Combined v-n diagram
15
Graph 6.1 Linear Variation of lift along wing semi span
17
Graph6.2 Elliptic Variation of lift along wing semi span
19
Graph.6.3 Load distribution on wing
20
Graph.6.4 Self weight distribution on wing
22
Graph6.5 Fuel Distribution
24
Graph 6.6 Overall Weight Distribution
25
Graph 6.7 Shear Force on Wing
27
Graph 6.8Bending Moment on Wing
28
LIST OF FIGURES Fig 1.1 NACA 64A204 AIRFOIL
2
Fig 3.1 Gust in Aircraft
6
Fig.6.1 Slope of Fuel Weight
23
Fig 9.1 NACA 64A204 AIRFOIL
34
Fig.11.1 Structural components of Wing
46
1. REQUIRED DATA FROM ADP-I This aircraft design project-2 is basically a continuation of aircraft design project-1. In design project-1 the following were done,
Mission specifications Weight estimation Engine selection Airfoil selection Drag estimation Performance analysis
In aircraft design project-2 taking the values obtained in design project-1 as input, the load factors during various phases of flight is calculated and the V-n diagram is drawn, the load distribution on the wing and the fuselage is found and the shear force diagram and the bending moment diagram for the wing and the fuselage are drawn and the internal structure design is also done. The following values are taken from the aircraft design project-1 and are used in the th e aircraft design project-2Retrived data from aircraft design project-1.
Weight Empty Weight Fuel Weight Overall Weight Weight of Crew Payload weight
Unit(Kg)
Unit(N)
6247.5
61287.97
3421.25
33562.46
18159.54
178145.08
100
981
8300
81423
Table.1.1 Table.1.1 Details we weigh t fr om ADP-I
Estimation of critical performance measure: S.No 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
Content Aspect ratio(AR) Mach Angle(μ) Sweep angle (Ʌ) (Ʌ) Leading Edge Sweep Angle (ɅLE) Sweep Angle at half of chord(Ʌ chord(Ʌc/4 ) Rolling Moment (Clβ ) Wing Wing Area Area (S) (W/S)L (W/S)TO Wing Span(b) Root Chord (Croot) Tip Chord (Ctip) Equivalent Aspect Ratio (AR eq eq) Aerodynamic Chord (C) (t/C)root (t/C)tip Thickness Distribution(Y) Centre of Gravity (X)
Values 2.81 28.73 73.52° 74.29° 71.45° 0.518 53.38m 53.38m 5554. 5554.46 46Kg Kg/m /m 6534. 6534.6K 6Kg/ g/m m 12.26m 6.803m 1.90m 2.12 4.35m 0.06 0.3356 2.81m 7.68m
°
Table 1.2 1.2 Requi Requi r ed Paramete Parameters rs fr om AD P-I MULTIROLE FIGHTER AIRCRAFT
Page 1
Airfoil Characteristics, 1.Maximum coefficient of Lift(CLmax)= 1.45 2. Maximum Coefficient of Lift Wing (CLmax)W= 1.52 3.Gross maximum Lift coefficient (C Lmax)gross= 1.68(FLAP DOWN) 4.Net Maximum Lift Coefficient=1.72
F ig 1.1 1.1 NACA 64A204 64A204 AI RFOI L
MULTIROLE FIGHTER AIRCRAFT
Page 2
2. V-N DIAGRAM 2.1 2.1 I ntr oduction oduction Flight regime of any aircraft includes all permissible combinations of speeds, altitudes, weights, centers of gravity, and configurations. This regime is shaped by aerodynamics, propulsion, structure, and dynamics of aircraft. The borders of this flight regime are called flight envelope or maneuvering maneuvering envelope. The safety of human onboard is guaranteed by aircraft designer and manufacturer. Pilots are always trained and warned through flight instruction manual not to fly out of flight envelope, since the aircraft is not stable, or not controllable or not structurally strong enough outside the boundaries of flight envelope. A mishap or crash is expected, if an aircraft is flown outside flight envelope. The flight envelope has various types; each of which is usually the allowable variations of one flight parameter versus another parameter. These envelopes are calculated and plotted by flight mechanics engineers and employed by pilots and flight crews. For instance, the loadmasters of a cargo aircraft must pay extra caution to the center of gravity location whenever they distribute various loads on the aircraft. aircraft. There are several crashes and mishaps mishaps that safety board’s report indicated that load master are responsible, since they deployed more loads than allowed, or misplaced the load before take-off. Nose heavy and tail heavy are two flight concepts that pilots are familiar and experienced with, and are trained to deal with them safely. Pilots are using several graphs and charts in their flight operations. Four important envelopes are as follows: 1. Diagram of variations of aircraft lift coefficient versus Mach number (CL – M)2. M)2. Diagram of variations of airspeed versus altitude (V – (V – h) h) 3. Diagram of variations of center of gravity versus aircraft weight (Xcg – (Xcg – W) W) 4. Diagram of variations of airspeed versus load factor (V – (V – n) n) One of the most important impo rtant diagrams is referred to as flight envelope. This envelope demonstrates the variations of airspeed versus load factor (V – (V – n). n). In another word, it depicts the aircraft limit load factor as a function of airspeed. One of the primary reasons that this diagram is highly important is that, the maximum load factor; that is extracted from this graph; is a reference number in aircraft structural design. If the maximum load factor is under-calculated, the aircraft cannot withstand flight load safely. For this reason, it is recommended to structural engineers to recalculate the V -n diagram on their own as a safety factor. In this section, details of the technique to plot the V- n diagram in introduced. Shows a typical V-n diagram for a GA aircraft. This diagram is, in fact, a combination of two diagrams: 1. The V-n diagram without consideration of gust, 2. The V-n diagram on the effect of gust. In this section, we first have another look on the load factor and then present new concepts on load factor. Then the phenomena of gust and gust load are described. At the end of this section, the technique to plot V – V – n n diagram is completely described.
Calculation: Point A,
VS =
× ×. .×.
=183.51m/s
A(183.51)
MULTIROLE FIGHTER AIRCRAFT
Page 3
Point B, VS =510.89m/s Point C, VS =952.41m/s
C(183.51) Point F,
VS =
×−. ×. .×.
=371.2m/s
VS (371.2m/s,-4.5) Point E, N=-4.5
E(614.40,-4.5) Point G, N=-1
VS =
× ×. .×.
=175m/s
G(175,-1)
Graph.2.1 Graph.2.1 V-n Di agram agram MULTIROLE FIGHTER AIRCRAFT
Page 4
2.2 2.2 L oad F actor actor The load to the aircraft on the ground is naturally produced by the gravity (i.e. 1 times g). But, there are other sources of load to the aircraft during flight; one of which is the acceleration load. This load is usually normalized through load factor (i.e. "n" times g). In another word, aircraft load is expressed as a multiple of the standard acceleration due to gravity (g = 9.81 m/sec2 = 32.17 ft/sec2). In some instances of flight such as turn and pull-up, the aircraft must generate a lift force such that it is more than weight. For instance, load factor in a pull-up from equation 9.86 can c an be re-written as:
+ 1
n=
Where “a” is the centrifugal acceleration (V2/R). As this acceleration acceleration increases; i.e. airspeed increases or radius of turn decreases; the load factor will increase too. For other flight operations, similar expressions can be drawn. In some instances; especially for missiles; this load factor may get as high as 30. As the table 2.2 illustrates, a low load factor fighter may end up getting targeted by a high load factor missile. S.No
AIRCRAFT TYPE
1 2 3 4 5 6 7
Normal(non-acrobatic) Utility(Semi-acrobatic) Acrobatic Homebuilt Transport Highly maneuverable Bomber
MAXIMUM POSITIVE LOAD FACTOR 2.5-3.8 4.4 6 5 3-4 6.4-12 2-4
MAXIMUM NEGATIVE LOAD FACTOR -1 to -1.5 -1.8 -3 -2 -1 to -2 -3 to -6 -1to -2
Table 2.1. 2.1. Load factor f or vari ous types types of air craf t
MULTIROLE FIGHTER AIRCRAFT
Page 5
3. GUST ENVELOPE 3.1Description Gust loads are unsteady aerodynamic loads that are produced by atmospheric turbulence. They represent load factor that is added to the aerodynamic loads, which presented in the previous section. The effect of a turbulent gust is to produce a short time change chan ge in the effective angle an gle of attack. This can be either positive or negative, thereby producing an increase or decrease in the wing lift and a change in the load factor. The fig. shoes the model for the effect of gust on the aircraft in level flight. The aircraft has a forward velocity v the turbulent produce small velocity component in the aircraft υ and υ and U AT that instant, velocity component in the aircraft direction is υ+U.in level flight the mean velocity component normal to the flight direction is U= 0 Therefore total normal velocity is u
F ig 3.1 Gust Gust in Air craft
Where, K=Gust effictiveness CLaw= Slope of the wing curve, CLaw =
− √ −
W=18159.54Kg
MULTIROLE FIGHTER AIRCRAFT
2 S=53.38m cr =0.136668
3
kg/m
Page 6
The value of U is calculated by
×
U=K u Altitude range from(50000ft) S.No
V (m/s)
U (ft/s)
1
(VA) HIGH ANGLE OF ATTACK
38
2
(VCR CRUSEVE VELOCITY)
25
3
(VD) DIVE VELOCITY
12.5
1.03 .+1.03 =. w/s × o Cos Ʌ +
WHERE K=
SIMILARLY
Cla=
Ʌ
°
=73.52
. .×. ×. ×.×. =
=9.11
9.11×1.03 .+9.111.03
K=
K=0.58
a) U value for above 50,000 ft, VA-36ft/s
Vcr -25ft/s
VD-12.5ft/s
UV=KU MULTIROLE FIGHTER AIRCRAFT
Page 7
×
Uv=0.58 11.57
Uv=6.7106 m/s
Velocity at high AOA, From point B, VA=510.89 m/s
∆ ∆
.×.×.×.×. ×.
n=
n= 8.81
b) To find Gust Load to the VC (Cruise Velocity) for 25 ft
×
Uv=0.58 7.6175 Uv=4.418 m/s Velocity at cruise VC=614.46m/s
∆
.×.×.×.×. ×.
n=
∆
n= 6.975
c) To find Gust Load corresponding to the VD( Dive Velocity) U table is at 12.5ft=3.805m
×3.805
Uv=0.58
Uv=2.209m/s Velocity at dive condition VD=952.413m/s
MULTIROLE FIGHTER AIRCRAFT
Page 8
∆
.×.×.×.×. ×.
n=
∆
n= 5.406
Gr aph 3.1 Gust Gust L oad
MULTIROLE FIGHTER AIRCRAFT
Page 9
4. CRITICAL PERFORMANCE PARAMETERS 4.1C 4.1Cru ru isin isin g F li ght Per Per formance 4.1.1 Calculation of velocity at minimum thrust required: VTR (min) =
0.5 0.5
Where,
ρ = 0.1366 kg/m3 CDo = 0.054
= 340.19 kg/m at max.lift 3
K = 0.16 VTR (min) =
. .. 0.5340.1990.5
VTR (min) = 92.57 m/s
4.1.2Calculation of Max. Lift to Drag ratio (L/D)max : (L/D)max=
××
0.5
(L/D)max= 5.38
4.1.3Calculation of Velocity at Max. Lift to drag ratio V (L/D) max: V (L/D) max= VTR (min) V (L/D) max= 92.57m/s
4.1.4Thrust Required minimum Trmin: Trmin=
Trmin=
. × 9.81 .
Trmin= 33.11 kN
3.1.5Power Required minimum (Pr): For level unaccelerated flight power = MULTIROLE FIGHTER AIRCRAFT
= force × Page 10
=F×V Pr = Trmin× V = 33.11 × 614.46 Pr = 20.34 MNm/s
4.1.6Thrust Available: TA = 196KN (From Engine selections) 4.1.7Power available: PA = TA × V PA = 196× 614.46 3
PA = 196 × 10 × 614.46 PA = 120.43 MNm/s
4.1.8Max Rate of Climb(R/C)max : (R/C)max=
2 × × 1 × 2 2
Where R = 1 +
0.5
3/2
0.5
2 1 1 /
0.5
(L/D)max = 5.38 T/W = 1.02
ρ = 1.225 Kg/m3 CDo = 0.054
(
TO =
3
6534.6 Kg/m
Therefore, Z=1+
1 . 1.022
Z = 2.05 (R/C)max=
6534.6 . ×1.225 ×0.054 {1.02} 1. × 1.022 2.05 0.5
3/2
(R/C) max= 190.38m/s
MULTIROLE FIGHTER AIRCRAFT
Page 11
4.1.9Velocity at max.rate of climb:
/3×× {1.026534.6 ×2.05/3×1.225×0.054} 0.5
V(R/C)max=
0.5
=
V(R/C)max= 262.39 m/s
4.1.10 Level Turn: Turn Radius: R =
−.
Where, n=
= . = 9.08 .
v = 614.46 m/s R=
. ..− −.
R= 4263.33m
4.1.11Turn Rate (ɷ):
− −. ..− −. = .
ɷ =
ɷ = 0.144 rad
4.2Take 4.2Takeoff off per per f ormance 4.2.1. Takeoff distance:
. S = g
=
. . .×.× . ×.
Sg= 436.3 m 4.2.2. Flight path radius: R=
.
MULTIROLE FIGHTER AIRCRAFT
Page 12
.. .
=
R = 2666m
4.2.3 Flight path Angle:
) .) = Cos (1
OB =
OB
OB =
-1
Cos (1-1
°
6.129
4.2.4 Airborne Distance:
Sa= RSin
OB
°
= 2666 Sin6.129 = 284m 4.2.5 Total takeoff distance: Takeoff distance = Sg + Sa = 436.3 + 284.6 Takeoff distance = 720.9m
4.3 L andin g Per Per for mance 4.3.1 Flare Velocity: Vf = = 1.15 = 1.15
× ×
Vstall Vstall
Vf = 70.49 m/s
4.3.2 Flare Height:
Hf = R(1- Cos a) Where, R = 4263.33 m
− ] [.−.
-1 a =Sin
-1 a =Sin
MULTIROLE FIGHTER AIRCRAFT
Page 13
a =
°
-56.52
Hf = 4263.33(1- Cos (-56.52)) Hf = 1911.48 m
4.3.3 Approach Distance: Approach distance to clear the 50 feet distance is Sa =
−
Sub. All value in above equation, Sa =
−. −. −.
Sa = 1231.1m
4.3.4 Flare Distance: Sf = R
× × 56.52° Sin
Sf = = 4263.33
a
Sin
Sf = -3555.95m
4.4Grou 4.4Grou nd Roll Roll
Sg = jN × × + .. 2 × 18159.54 × 1 + Sg =1.1×3 1.225 53.38 1.45 .×.×.×.×.×. Sg = 82 m Total Landing Distance, Ld = Sa + Sg +Sf Ld = 1231.1 + 82 + (-3555.95) Ld = -2242.85m -ve sign indicates the direction of landing ( Fighter return) Ld= 2242.85m
MULTIROLE FIGHTER AIRCRAFT
Page 14
5. COMBINED V-N DIAGRAM (OR) FLIGHT ENVELOPE
Grapph 5.1 Combined v-n v-n di agram
MULTIROLE FIGHTER AIRCRAFT
Page 15
6. WING LOAD DISTRIBUTION 6.1 WING DESCRIPTION
Lift varies along the wing span due to the variation in chord length, angle of attack and Sweep along the span. Schrenk’s curve defines this lift distribution over the wing span of an Aircraft, also called simply as Lift Distribution Curve. Schrenk’s Curve is given by
+
Y=
Where
y1 is Linear Variation of lift along semi wing span also named as L1 y2 is Elliptic Lift Distribution along the wing span also named as L2 6.2 LINEAR LIFT DISTRIBUTION
By Schrenk’s Curve, Lift at Root,
Lroot =
×××
=
.×.×.×.
Lroot = 254487.8935N/m Lift at Tip,
Ltip =
××× =
.×.×.×.
Ltip = 71075.55N/m
MULTIROLE FIGHTER AIRCRAFT
Page 16
⎾ . = .
a=
a = 21.60m Equation of linear lift distribution for starboard wing Y1 = -mx + c Y1 = -(3289.19)x + 254487.8935 Equation of linear lift distribution for port wing Y1 = (3289.19)x + 254487.8935 For Half of the wing We Get, Y1/2 = -(1644.595)x + 127243.945
Graph 6.1 L in ear Var iation of l if t along wing se semi span span Coordinates for plotting the above is,
MULTIROLE FIGHTER AIRCRAFT
X
L1
L2
L
0
127243.9
5250.42
132494.4
1
125599.4
5244.79
130844.1
2
123954.8
5227.865
129182.6
3
122310.2
5199.533
127509.7
4
120665.6
5159.607
125825.2
5
119021
5107.815
124128.8
6
117376.4
5043.792
122420.2
7
115731.8
4967.064
120698.8 Page 17
8
114087.2
4877.032
118964.2
9
112442.6
4772.943
117215.5
10
110798
4653.855
115451.8
11
109153.4
4518.581
113672
12
107508.8
4365.616
111874.4
13
105864.2
4193.026
110057.2
14
104219.6
3998.27
108217.9
15
102575
3777.92
106352.9
16
100930.4
3527.185
104457.6
17
99285.83
3239.014
102524.8
18
97641.24
2902.279
100543.5
19
95996.64
2497.411
98494.05
20
94352.05
1983.111
96335.16
21
92707.45
1228.912
93936.36
21.6
91720.69
0
91720.69
6.3 ELLIPTIC LIFT DISTRIBUTION: Twice the area under u nder the curve cur ve or line will w ill give the lift li ft which will w ill be required requi red to overcome weight Considering an elliptic lift distribution we get,
==
A=
Where b1-is Actual lift at root And a- is wing semi span Lift at tip
b1 =
b1 =
.×. ××.
b1 = 5250.49N/m
Y2 = MULTIROLE FIGHTER AIRCRAFT
× √ Page 18
Y2 =
×. × .
.×√ .× √ . .
=
Graph6.2 Graph6.2 Ell iptic Vari ation of l if t along wing se semi span span Coordinates for plotting the above is,
MULTIROLE FIGHTER AIRCRAFT
X
L1
L2
L
0
127243.9
5250.42
132494.4
1
125599.4
5244.79
130844.1
2
123954.8
5227.865
129182.6
3
122310.2
5199.533
127509.7
4
120665.6
5159.607
125825.2
5
119021
5107.815
124128.8
6
117376.4
5043.792
122420.2
7
115731.8
4967.064
120698.8
8
114087.2
4877.032
118964.2
9
112442.6
4772.943
117215.5
10
110798
4653.855
115451.8
11
109153.4
4518.581
113672
12
107508.8
4365.616
111874.4
13
105864.2
4193.026
110057.2
14
104219.6
3998.27
108217.9
15
102575
3777.92
106352.9
16
100930.4
3527.185
104457.6
17
99285.83
3239.014
102524.8
Page 19
18
97641.24
2902.279
100543.5
19
95996.64
2497.411
98494.05
20
94352.05
1983.111
96335.16
21
92707.45
1228.912
93936.36
21.6
91720.69
0
91720.69
6.4 SCHRENKS CURVE
+
S=
. √ .
Y=-1644.595x+127243.945+243.075
Graph.6.3 Graph.6.3 L oad distribution distribution on win g Coordinates for plotting the above is,
x -21.6 -21 -20 -19 -18 -17 -16 -15 -14 -13 -12 -11 MULTIROLE FIGHTER AIRCRAFT
L 91720.7 93936.4 96335.2 98494.1 100544 102525 104458 106353 108218 110057 111874 113672 Page 20
-10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 21.6
115452 117216 118964 120699 122420 124129 125825 127510 129183 130844.1402 132494.365 130844.1402 129182.6196 127509.6928 125825.1719 124128.785 122420.1666 120698.8438 118964.217 117215.5334 115451.8498 113671.9805 111874.4212 110057.2361 108217.8847 106352.9404 104457.6099 102524.8443 100543.5139 98494.0508 96335.15567 93936.36181 91720.693
6.5 Load Estimation on Wings Self-Weight,
= 0.25 Wwing = 18159.54
×
9.81
= 44536.27N MULTIROLE FIGHTER AIRCRAFT
Page 21
W port wing = 22268.135N WStarboard wing = -22268.135N Assuming Parabolic Weight Distribution, Y3 = K (x-(b/2))
2
Y3 = K (x-(21.60/2))
2
−. ∫. 3 ∫. − =K
[[.−−. −−. ]
= K
Y3 =
−.
× .
K=
2
Y3 = -6.6289(x-21.60)
Graph.6.4 Se Self weigh weigh t distr distr ibu tion on wi ng Coordinates for plotting the above is,
X 0 1 2 3 MULTIROLE FIGHTER AIRCRAFT
Y3 -1557.5172 -1416.64098 -1282.44132 -1154.91826 Page 22
4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 21.6
-1034.07180 -919.901948 -812.408688 -711.592028 -617.451968 -529.988508 -449.201648 -375.091388 -307.657728 -246.900668 -192.820208 -145.416348 -104.689088 -70.638428 -43.264368 -22.566908 -8.546048 -1.201788 0
6.6 F uel Weigh Weigh t: Wfuel per wing = 33562.46Kg For 2 Wings= 67124.92 Kg
F ig.6.1 Slope Slope of F uel Weigh Weigh t
Using general formula for straight, Line y=mx+c dy = 1.5325 m Dy= 5250.42 – 5250.42 – 1.5325 1.5325 m MULTIROLE FIGHTER AIRCRAFT
Page 23
m=
.−. . 2
m = 1427.1 N/m
yf = 1427.10x-5250.42 Area of Ellipse,
= . =
X= 9.167
Trapezoidal,
= . = 0.9195 =0.9195 + 0.4195 = 1.839 m
Graph6.5 Graph6.5 Fu el Di stri bution
MULTIROLE FIGHTER AIRCRAFT
Page 24
Graph 6.6 Ove Overal l Weight Weight Di stri bution
Curve/Component
Ar ea enclosed/ nclosed/ Structur al Weight(N )
Centr Centr oid (F r om Win g r oot)
Y1/2
2364817572
7.2m
Y2/2
89064.1517
9.167m
Wing
22268.135
8.1m
Fuel
33562.46
1.839m
Tabl e.6.1 L oads Simpl if ied as as Poin t loads
6.7 Reac Reaction tion F orce and and B endin g M oment oment Calcul ations: ations: The wing is fixed at one end free at other end
Ʃ
F ig.4 Fi ghter ghter Ai rcraf t Win g as Cantileve Cantilever B eam
V, VA = 3364817.572 + 89064.1517 – 89064.1517 – 22268.135 – 22268.135 – 33562.46 33562.46
MULTIROLE FIGHTER AIRCRAFT
Page 25
VA =238051.129 N
Ʃ
M, MA -17026686.52 – -17026686.52 – 816451.0786 816451.0786 + 180371.8935 + 61721.36394 MA = 17601044.34 N/m
Now we know VA and MA, using this we can find out shear force and Bending moment. 6.8 Shear Force:
− = ∫ 3 − . +.+. . .− 177.231 = ∫ (
SFBC SFBC
5.70732) 2) 7325 73254.4.4366 436666 2
SFBC = -822.2975(x ) + 127243.9468 + [x
2
6.6289[ -6.13x +466.56x]-2398051.129
466.562 466.562
-1
+ 466.56 sin (x/6.13)] +
∫ ∫1427.10 5250.42
SFCD = SFBC +
dx
= SFBC +
2
SFCD =SFBC + (713.55 x – 5250.42x) 5250.42x) 2
SFDE = SFCD + (713.55x -5250.42x) + 33562.46
MULTIROLE FIGHTER AIRCRAFT
Page 26
Graph 6.7 Shear Shear F orce on Wi ng
6.9 Bending M oment oment
BMBC =
+ ∬{ 3 } dx +M 2
A
BMBC = -2398051.129x – -2398051.129x – 1096.3966x 1096.3966x3 + 127243.9468 x2 +
466.563 √ 466.563
) + 466.56 sin-1(x/6.13) +2.21(466.56-x2)1,5 –
121.5375[x(x
6.6289[(x4/12)+233.28x2-2.043x3]+`17601044.34
∬
BMCD= BMBC +
=
dx2
3
2
BMBC +4757x - 2625.21x
475.7 3
BMDE= BMCD +
MULTIROLE FIGHTER AIRCRAFT
-2625.21x2 Page 27
BMEA= BMDE-[4757x3-2625.21x2] + 33562.46 x
Graph 6.8Bending 6.8Bending M oment oment on Win g
MULTIROLE FIGHTER AIRCRAFT
Page 28
7. CG CALCULATION
MULTIROLE FIGHTER AIRCRAFT
Page 29
MULTIROLE FIGHTER AIRCRAFT
Page 30
8. MATERIAL SELECTION 8.1 Des Descri ption: Aircraft structures are basically unidirectional. This means that one dimension, the length, is much larger than the others - width or height. For example, the span of the wing and tail spars is much longer than their width and depth; the ribs have a much larger chord length than height and/or width; a whole wing has a span that is larger than its chords or thickness; and the fuselage is much longer than it is wide or high. Even a propeller has a diameter d iameter much larger than its blade width and thickness, etc.... For this simple reason, a designer chooses to use unidirectional material when designing designing for an efficient efficient strength to weight weight structure. structure. Unidirectional materials are basically composed of thin, relatively flexible, long fibers which are very strong in tension (like a thread, a rope, a stranded steel wire cable, etc.). An aircraft structure is also very close close to a symmetrical structure. Those mean the up and down loads are almost equal to each other. The tail loads may be down or up depending on the pilot raising or dipping the nose of the aircraft by pulling or pushing the pitch control; the rudder may be deflected to the right as well as to the left (side loads on the fuselage). The gusts hitting the wing may be positive or negative, giving the up or down loads which the occupant experiences by being pushed down in in the seat or hanging in in the belt. Because of these factors, factors, the designer has to use a structural material that can withstand both tension and compression. Unidirectional fibers may be excellent in tension, but due to their small cross section, they have very little inertia (we will explain inertia another time) and cannot take much compression. They will escape the load by bucking away. As in the illustration, you cannot load a string, or wire, or chain in compression. In order to make thin fibers strong in compression, they are "glued together" with some kind of an "embedding". In this way we can take advantage of their tension strength and are no longer penalized by their individual compression weakness because, as a whole, they become compression resistant resistant as they help each each other to not buckle away. The embedding is usually a lighter, softer "resin" holding the fibers together and enabling them to take the required compression loads. This is a very good structural material.
8.2 Wood: Wood: Historically, wood has been used as the first unidirectional structural raw material. They have to be tall and straight and their wood must be strong and light. The dark bands (late wood) contain many fibers, whereas the light bands (early wood) contain much more "resin". Thus the wider the dark bands, the stronger and heavier the wood. If the dark bands are very narrow and the light bands quite wide, the wood is light but not very strong. To get the most efficient strength to weight ratio for wood we need a definite numbers of bands per inch. Some of our aircraft structures are two-dimensional two -dimensional (length and width are large with respect to thickness). Plywood is often used for such structures. Several thin boards (foils) are glued together so that the fibers of the various layers cross over at different angles (usually 90 degrees today years back you could get them at 30 and 45 degrees as well). Plywood makes excellent "shear webs" if the designer knows how to use plywood efficiently. (We will learn the basis of stress analysis sometime later.)
MULTIROLE FIGHTER AIRCRAFT
Page 31
Today good aircraft wood is very hard to come by. Instead of using one good board for our spars, we have to use laminations because large pieces of wood are practically unavailable, and we no longer can trust the wood quality. From an availability point of view, we simply need a substitute for what nature has supplied us with until now.
8.3 8.3 Alumi ni um all oys oys: So, since wood may not n ot be as available as it was before, we look at another material which is strong, light and easily available at a reasonable price (there's no point in discussing Titanium - it's simply too expensive). Aluminum alloys are certainly one answer. We will discuss the properties of those alloys which are used in light plane construction in more detail later. For the time being we will look at Aluminum as a construction material.
8.4 8.4 Ex truded truded alumi ni um alloys: alloys: Due to the manufacturing manufacturing process for Aluminum we get a unidirectional material quite a bit stronger in the lengthwise direction than across. And even better, it is n ot only strong in tension but also in compression. Comparing extrusions to wood, the tension and compression characteristics are practically the same for aluminum alloys so that the linear stress analysis applies. Wood, on the other hand, has a tensile strength about twice as great as its compression strength; accordingly, special stress analysis methods must be used and a good understanding of wood under stress is essential if stress concentrations are to be avoided! Aluminium alloys, in thin sheets (.016 to .125 of an inch) provide an excellent two dimensional material used extensively as shear webs - with or without stiffeners - and also as tension/compression members when suitably formed (bent).It is worthwhile to remember that aluminium is an artificial metal. There is no aluminium ore in nature. Aluminium is manufactured by applying electric power to bauxite (aluminium oxide) to obtain the metal, which is then mixed with various strength-giving additives. ad ditives. (In a later article, we will see which additives are used, and why and how we can increase aluminum’s aluminum’s strength by cold work hardening or by tempering.) All the commonly used aluminium alloys al loys are available from the shelf of dealers. When requested with the purchase, you can obtain a "mill test report" that guarantees the chemical and physical properties as tested tested to accepted specifications. specifications. As a rule of thumb, aluminium is three times heavier, but also three times stronger than wood. Steel is again three times heavier and stronger than aluminium.
8.5 Stee Steel : The next material to be considered for aircraft structure will thus be steel, which has the same weight-tostrength ratio of wood or aluminium. Apart from mild steel which is used for brackets needing little strength, we are mainly using a chrome-molybdenum alloy called AISI 413ON or 4140. The common raw materials available are tubes and sheet metal. Steel, due to its high density, is not used as shear webs like aluminium sheets or plywood. Where we would need, say.100" plywood, a .032 inch aluminium sheet would be required, but only a .010 steel sheet would be required, which is just too thin to handle with any hope of a nice finish. That is why a steel fuselage uses tubes also as diagonals to carry carry the shear in compression or tension
MULTIROLE FIGHTER AIRCRAFT
Page 32
and the whole structure is then covered with fabric (light weight) to give it the required aerodynamic shape or desired look. It must be noted that this method involves two techniques: steel work and fabric covering.
8.6 Composi Composi te M ater ater i als: The designer of composite aircraft simply uses fibers in the desired direction exactly where and in the amount required. The fibers are embedded in resin to hold them in place and provide the required support against buckling. Instead of plywood or sheet metal which allows single curvature only, the composite designer uses cloth where the fibers are laid in two directions .(the woven thread and weft) also embedded in resin. This has the advantage of freedom of shape in double curvature as required by optimum aerodynamic shapes and for very appealing look (importance of aesthetics). Today's fibers (glass, nylon, Kevlar, carbon, whiskers or single crystal fibers of various chemical compositions) are very strong, thus the structure becomes very light. The drawback is very little stiffness. The structure needs stiffening which is achieved either by the usual discreet stiffeners, -or more elegantly with a sandwich structure: two layers of thin uni- or bi-directional fibers are held apart by a lightweight core (foam or "honeycomb"). This allows the designer to achieve the required inertia or stiffness. From an engineering en gineering standpoint, this method is very v ery attractive and supported by many authorities because it allows new developments which are required in case of war. But this method also has its drawbacks for homebuilding: A mold is needed, needed , and very strict quality control is a must for for the right amount of fibers fibers and resin and for good adhesion between both to prevent too "dry" or "wet" a structure. Also the curing of the resin is quite sensitive to temperature, humidity and pressure. Finally, the resins are active chemicals which will not only produce the well-known allergies but also the chemicals that attack our body (especially the eyes and lungs) and they have the unfortunate property of being cumulatively damaging and the result (in particular deterioration of the eye) shows up only years after initial contact. Another disadvantage of the resins is their limited shelf life, i.e., if the resin resin is not used within the specified time lapse after manufacturing, the results may be unsatisfactory and unsafe.
8.7 H eavy avy Air craft Raw M ater ater ial s: MAGNESIUM: An expensive expen sive material. Castings are the only on ly readily available forms. Special precaution must be taken when machining magnesium because this metal burns when hot. TITANIUM: A very expensive material. Very tough material and difficult to machine. CARBON FIBERS : Still very expensive materials. KEVLAR FIBERS : Very expensive and also critical to work with because it is hard to "soak" in the resin.
MULTIROLE FIGHTER AIRCRAFT
Page 33
9. WING DESIGN 9.1 INTRODUCTION:Wing is an important component in any Aircraft because of which Aerodynamic Lift force is generated that makes an aircraft to fly. And also it provides Stability for an aircraft. As we know that Airfoils is a cross section of wing. Selection airfoil should meet requirements with following calculations,
9.2 AIRFOIL SELECTION:Family
4-Digit
Advantag
Disadvantages
1.Goodstallcharacteristics
1.Lowmaximumliftcoefficient
2.Smallcenterofpressuremovement acrosslargespeedrange
2.Relativelyhighdrag
Applications
1.Generalaviation 2.Horizontaltails S Symmetrical: ymmet ric al:
3.Highpitchingmoment 3.Roughnesshaslittleeffect
5-Digit
1.Highermaximumliftcoefficient
1.Poorstallbehavior
2.Lowpitchingmoment
2.Relativelyhighdrag
3.Supersonicjets 4.Helicopterblades 5.Shrouds 1.Generalaviation 2.Pistonpoweredbombers, transports 3.Commuters
16-Series 1.Avoidslowpressurepeaks
1.Relativelylowlift
1.Aircraftpropellers 2.Shippropellers
6-Series
1.Highdragoutsideofthe optimumrangeofoperatin g conditions
1.Piston-poweredfighters 2.Businessjets 3.Jettrainers 4.Supersonicjets
1.Highmaximumliftcoefficient 2.Verylowdragoverasmallrangeof operatingconditions
2.Highpitchingmoment 3.Optimizedforhighspeed 3.Poorstallbehavior 7-Series
1.Verylowdragoverasmallrangeof operatingconditions
1.Reducedmaximumlif t coefficient
2.Lowpitchingmoment
2.Highdragoutsideofthe optimumrangeofoperatin g conditions
Seldomused
3.Poorstallbehavior 8-Series
Unknown
Unknown
Veryseldomused
64A204 Airfoil will be selected by following calculations, NA CA 64A204
F ig 9.1
MULTIROLE FIGHTER AIRCRAFT
Page 34
VApproach=1.3 Vstall VApproach=155knots
. . = .
Vstall=
(Knots)
(m/s)
Vstall =61.3 m/s
× ×× ×. = .×.×.
CLmax=
CLmax=1.45
CLmx=1.52 . CLmx= .=1.68 = . .
CLmax (wing) = CLmax (gross)
Selection of high lift devices:
MULTIROLE FIGHTER AIRCRAFT
Page 35
CLmax (net) =CLmax (gross) – ΔC – ΔCl(HLD) =1.68-0.48=1.2
V ) C ρ = × (61.3) ×1.45×1.225 =
(
2
stall
Lmax
2
= 3337.29 N/m
=340.19Kg/m ( =(V )× C 2
TO)
TO
Lmax (gross) ×
ρ
= (79.69)2 × 1.68 ×1.225
(TO)=6534.6
2
Kg/m
(TO))
(Landing) =
0.85 (
(Landing) =
5554.46 Kg/m
2
Win g Area
= . = 53.38m / .
S=
MULTIROLE FIGHTER AIRCRAFT
2
Page 36
Wi ng Span Span 0.5
b= (AR ×S)
0.5
b=(2.8175 × 53.38)
b = 12.26 m Root Chord
+ƛ ×. = .+.
CRoot =
CRoot = 6.803m Tip Chord
CTip= ƛ × × CRoot = 0.28 × 6.803 CTip= 1.90m Equivalent Aspect ratio C
Areq = A(Mmax)
Where, A=2.34 C= -0.13 Speed of sound at 56,025ft is 295.070m/s
. = 2.08 .
Mmax=
Areq= 2.34(2.08)
-0.13
Areq= 2.12 Wing Aerodynamic Chord (C)
C = × CRoot×
×ƛ×ƛ+ƛ +ƛ
MULTIROLE FIGHTER AIRCRAFT
Page 37
(OR)
. C= . C=
C= 4.35 Volume of the fuel in the wing
ℎ . = . × =
3
= 5.12m
Thickness ratio (t/C)
Volume of Fuel = × C × 0.5 × C × 0.5b × 0.75 × 2
5.12 = × 4.35 × 0.5 × 4.35 × 0.5b × 0.75 × 2
(root)=
0.06
(tip) =
0.3356
Thickness distribution
. = .
Y=( )
Y= 2.81m
MULTIROLE FIGHTER AIRCRAFT
Page 38
9.3 Wing Selection:POSITION OF WING
The location of the wing in the fuselage (along the vertical axis) is very important. Each configuration (Low, High and mid) has its own advantages but in this design, the mid wing Mid Wing In general, features of the mid-wing configuration (Figure 5.3-b, and Figure 5.4-4) stand Some what between features of high-wing configuration and features of low-wing configuration. The major difference lies in the necessity to cut the wing spar in two half in order to save the space inside the fuselage. However, another alternative is not to cut the wing spar and letting it to pass through the fuselage ;which leads to an occupied space of the fuselage. Both alternatives carry a few disadvantages. Other than those features that can be easily derived from two previous sections, some new features of amid-wing configuration are as follows:
1. The aircraft structure is heavier ,due to then ecessity of reinforcing reinforcing wing root at the intersection with with the fuselage. 2. The mid wing is more expensive compared with high and low-wing configurations. configurations. 3. The mid wing is more attractive attractive compared with two other configurations. 4. The mid wing is aerodynamically streamliner compared with two other configurations. 5. The strut is usually not used to reinforce the the wing structure. structure. 6. The pilot can get get into the cockpit cockpit using the the wing as a step in a small small GA aircraft. aircraft. 7. The mid-wing has less interference drag than low-wing and high-wing.
As per the requirement of multirole fighter aircraft design, Swept Back Wing selected by following calculation, SWEEP-BACK WING:
Sweep Angle Consider the top view of an aircraft .The angle between a constant percentage chord lines along.These misspend of the wing and the lateral axis perpendicular to the fuselage center line (y-axis) is called leading edges weep (LE).The angle between the wing leading edge and they -axis of the aircraft is called leading edge sweep (LE). Similarly, the angle between the wing trailing edge and the longitudinal axis (y-axis)of the aircraft is called trailing edge sweep(TE).In the same fashion, the angle between the wing quarter chord line and they-axis of the aircraft is called quarter chord sweep(C/4 ).Andfinally,theanglebetweenthewing50percentchordlineandthey-axis of the aircraft is 50percentchordsweep(C/2).
MULTIROLE FIGHTER AIRCRAFT
Page 39
Basically, a wing is being swept for the following five design goals: 1. 2. 3. 4. 5.
Improving the wing wing aerodynamic aerodynamic features (lift, drag ,pitching moment)at transonic, supersonic and hypersonic speeds by delaying the compressibility effects. Adjusting the aircraft center of gravity. Improving static lateral stability. Impacting longitudinal and directional stability. Increasing pilot view(especially for for fighter pilots. 1) Mach angle, μ= Sin
[ℎ .]
-1
Where,
Mach No. = 2.08
[.]
μ = Sin-1
MULTIROLE FIGHTER AIRCRAFT
Page 40
μ = 28.73° Swept Angle (Ʌ (Ʌ) = 1.2 × (90-28.73) Swept back wing angle is (Ʌ)= 73.52 ° 2) Effective Chord length of Swept wing,
= .
Ceff =
Ceff = 3.52 m
3) Leading edge Swept Back (ɅLE) -1
= Tan
-1
= Tan
[tan [tanɅ /−ƛ +ƛ ] −.] [tan [tan73.522 /.+. °
ɅLE = 74.39
4) Sweep angle at a reference or half of the chord
[tanɅ /−ƛ +ƛ ] −.] [tanɅ73.52 /.+.
-1
Ʌc/4 = Tan
-1
Ʌc/4 = Tan
°
Ʌc/4 = 71.45
5) The rolling moment due to aft sweep is proportional propo rtional to the sine of twice the leading edge sweep angle.
∞ ∞
Clβ
Sin (ɅLE)
Clβ 0.518 Oswald Span efficiency, η = 0.7
MULTIROLE FIGHTER AIRCRAFT
Page 41
10. FUSELAGE DESIGN FUSELAGE DESIGN: The fuselage (/ˈfjuːzəl /ˈfjuːzəlɑːʒ/; ɑːʒ/; from the French the French fuselé "spindle-shaped") is an aircraft' an aircraft'ss main body section that holds crew and passengers or cargo. cargo. In single-engine aircraft it will usually contain an engine, although in some amphibious aircraft the single engine is mounted on a pylon a pylon attached to the fuselage which in turn is used as a floating hull. floating hull. The The fuselage also serves to position control and stabilization surfaces in specific relationships to lifting surfaces, required for aircraft stability and maneuverability Mono coque shell
In this method, the exterior surface of the fuselage is also the primary structure. A typical early form of this (see the Lockheed the Lockheed Vega) was Vega) was built using molded plywood, molded plywood, where where the layers of plywood are formed over a "plug" or within a mold. A mold. A later form of this structure uses fiberglass uses fiberglass cloth impregnated with polyester or epoxy resin, instead of plywood, as the skin. A simple form of this used in some amateur-built aircraft uses rigid expanded foam plastic as the core, with a fiberglass covering, eliminating the necessity of fabricating molds, but requiring more effort in finishing (see the Rutan VariEze). VariEze). An example of a larger molded plywood aircraft is the de Havilland Mosquito fighter/light bomber of World World War II. No II. No plywood-skin fuselage is truly monocoque, since stiffening elements are incorporated into the structure to carry concentrated loads that would otherwise buckle the thin skin. The use of molded fiberglass using negative ("female") molds (which give a nearly finished product) is prevalent in the series production of many modern sailplanes. sailplanes. The use of molded composites for fuselage structures is being extended to large passenger aircraft such as the Boeing 787 Dreamliner (using pressure-molding on female molds). Semi-monocoque[edit] Semi-monocoque[edit]
This is the preferred method of constructing an all-aluminum all-aluminum fuselage. First, a series of frames frames in the shape of the fuselage cross sections are held in position on a rigid a rigid fixture. These fixture. These frames are then joined with lightweight longitudinal elements called stringers. stringers. These are in turn covered with a skin of sheet aluminum, attached by riveting or by bonding with special adhesives. The fixture is then disassembled and removed from the completed fuselage shell, which is then fitted out with wiring, controls, and interior equipment such as seats and luggage bins. Most modern large aircraft are built using this technique, but use several large sections constructed in this fashion which are then joined with fasteners with fasteners to form the complete fuselage. As the accuracy of the final product is determined largely by the costly fixture, this form is suitable for series production, where a large number of identical aircraft are to be produced. Early examples of this type include the Douglas Aircraft DC-2 and DC-3 civil aircraft and the Boeing Boeing B-17 Flying Fortress. Fortress. Most metal light aircraft are constructed using this process. Both monocoque and semi-monocoque are referred to as "stressed skin" structures as all or a portion of the external load (i.e. from wings and empennage, and from discrete masses such as the engine) is taken by the surface covering. In addition, all the load from internal pressurization internal pressurization is carried (as skin tension) by the external skin. The proportioning of loads between the components is a design choice dictated largely by the dimensions, strength, and elasticity of the components available for construction and whether or not a design is intended to be "self jigging", not requiring a complete fixture for alignment. MULTIROLE FIGHTER AIRCRAFT
Page 42
MULTIROLE FIGHTER AIRCRAFT
Page 43
Design calculation:
√1
Nsa=0.45
=0.45 Width of the seat – seat – 21 21 inch Internal fuselage diameter =df i Df i= 21in+0+2in Gap between seat and side wall 2inch+2inch=4inch Df i=width of the seat + gap between seat and side wall =21 inch+4inch =25 inches MULTIROLE FIGHTER AIRCRAFT
Page 44
=25*0.0254m =0.635m Fuselage wall thickness(left and right )
∆
d= df outer outer -df inner inner = 0.084m+(0.045*0.635)
=0.1126m Outer fuselage diameter
∆
df=df i+ d =0.635+0.1126 =0.7476 In fighter aircrafts there will be no cabins Fuselage length: lf= lcockpit+l tail =3.5+(1.6*0.7476) =4.69616 =4.7m
MULTIROLE FIGHTER AIRCRAFT
Page 45
11. DETAILED STRUCTURAL LAYOUT 11.1 11.1 Fu nction Of The Stru Stru cture: cture: The primary functions of an aircraft‟s structure can be basically basically broken down into the following:
To transmit and resist applied loads. To provide and maintain aerodynamic shape. To protect its crew, passenger, payload, systems, s ystems, etc.
For the vast majority majority of aircraft, this leads to use use of a semi-monocoque design (i.e. (i.e. a thin, stressed outer shell with additional stiffening members) for the wing, fuselage & empennage. These notes will discuss the structural layout possibilities for each of these main areas, i.e. wing, fuselage & empennage.
11.2 11.2 Wing Str Str uctur al L ayout: ayout: The specified structural roles of the wing (or main plane) are:
To transmit: wing lift to the root via the main span wise beam Inertia loads from the power plants, undercarriage, etc., to the main beam. Aerodynamic loads generated on the aerofoil, a erofoil, control surfaces & flaps to the main beam. To react against: Landing loads at attachment points Loads from pylons/stores Wing drag and thrust loads To provide: Fuel tank age space Torsional rigidity to satisfy stiffness and aero-elastic requirements. To fulfill these specific roles, a wing layout will conventionally compromise: Span wise members (known as spars or booms) Chord wise members(ribs) A covering skin Stringers
F ig.11.1 Structur al components components of Wi ng MULTIROLE FIGHTER AIRCRAFT
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11.3 11.3 Basic Basic F un ctions Of W in g Str Str uctur al M ember mber s The structural functions of each of these types of members may be considered independently as: 11.3.1 SPARS
Form the main span wise beam Transmit bending and torsional loads Produce a closed-cell structure to provide resistance to torsion, shear and tension loads.
In particular:
Webs – Webs – resist resist shear and torsional loads and help to stabilize the skin. Flanges - resist the compressive loads caused by wing bending.
11.3.2SKIN:
To form impermeable aerodynamics surface Transmit aerodynamic forces to ribs & stringers Resist shear torsion loads (with spar webs). React axial bending loads (with stringers).
11.3.3 STRINGERS:
Increase skin panel buckling strength by b y dividing into smaller length sections. React axial bending loads
11.3.4 RIBS:
Maintain the aerodynamic shape Act along with the skin to resist the distributed aerodynamic pressure loads Distribute concentrated loads into the structure & redistribute stress around any discontinuities Increase the column buckling strength of the stringers through end restraint Increase the skin panel buckling strength.
11.3.5 SPARS:
These usually comprise thin aluminum alloy webs and flanges, sometimes with separate vertical stiffeners riveted on to the webs. Types of spars:
In the case of a two or three spar box beam layout, the front spar should be located as far forward as possible to maximize the wing box size, though this is subject to there being:
Adequate wing depth for reacting vertical shear loads. Adequate nose space for LE devices, de-icing equipment, etc.
This generally results in the front spar being located at 12% to 18% of the chord length. For a single spar Dnose layout, the spar will usually located at the maximum thickness position of the aerofoil section (typically between 30% & 40% along the chord length). MULTIROLE FIGHTER AIRCRAFT
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For the standard box beam layout, the rear rear spar will be located as for aft as possible, possible, once again to maximize the wing box size, but positioning will be limited by various space requirements for flaps, control surfaces, spoilers etc. This usually results in a location somewhere between about 55%and 70% of the chord length. If any intermediate spars are used, the y would tend to be spaced uniformly unless there are specific pickup point requirements. 11.4 11.4 F uselage uselage Str Str uctur e:
The fundamental purpose of the fuselage structure is to provide an envelope to support the payload, crew, equipment, systems and (possibly) the power-plant. Furthermore, it must react against the in-flight maneuver, pressurization and gust loads; also the landing gear and possibly any an y power-plant loads. It must be also be able to transmit control and trimming loads from the stability and control surfaces throughout the rest of the structure Fuselage contributes very little to lift and produces more drag but it is an important structural member/component. It is the connecting member to all load producing components such as wing, horizontal tail, tail, vertical tail, landing gear etc. and thus redistributes redistributes the load. load. It It also serves the purpose of housing or accommodating practically all equipment, accessories and systems in addition to carrying the payload. Because of large amount of equipment inside the fuselage, it is necessary to provide sufficient number of cutouts in the fuselage for access and inspection purposes. These cutouts and discontinuities result in fuselage design being more complicated, less precise and often less efficient in design. As a common member to which other components are attached, thereby transmitting the loads, fuselage can be considered as a long hollow beam. The reactions produced by the wing, tail or landing gear may be considered as concentrated loads at the respective attachment points. The balancing reactions are provided by the inertia forces contributed by the th e weight of the fuselage structure and the various components inside the fuselage. These reaction forces are distributed all along the length of the fuselage, though need not be uniformly. Unlike the wing, which is subjected to mainly unsymmetrical load, the fuselage is much simpler for structural analysis due to its symmetrical cross-section and symmetrical loading. The main load in the case of fuselage is the shear load because the load acting on the wing is transferred to the fuselage skin in the form of shear only. The structural design of both wing and fuselage begin with shear force and bending moment diagrams for the respective members. The maximum bending stress produced in each of them is checked to be less than the t he yield stress of the material chosen for the respe ctive member..
11.5 F uselage uselage L ayout Con cepts cepts:: There are two main categories of layout la yout concept in common use;
Mass boom and longeron layout Semi-monocoque layout
11.5.1 MASS BOOM & LONGERON LAYOUT
This is fundamentally very similar to the mass-boom wing-box concept discussed in previous section. It is used when the overall structural structural loading is relatively low or when there are extensive cut-outs in the shell. The concept comprises four or more continuous heavy booms (longeron), reacting against any direct stresses caused by applied vertical and lateral bending loads. Frames or solid section 11.5.2 SEMI-MONOCOQUE LAYOUT
The semi-monocoque is the most often used construction for modern, high-performance aircraft. Semimonocoqueliterally means half a single shell. Here, internal braces as well as the skin itself carry the stress. The vertical structural members are referred to as bulkheads, frames, and formers. The heavier vertical members MULTIROLE FIGHTER AIRCRAFT
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are located at intervals to allow for concentrated loads. These members are also found at points where fittings are used to attach other other units, such as the wings and stabilizers. Primary bending loads are taken by the the longerons, which usually extend across several points of support. The longerons are supplemented by other longitudinal members known as stringers. stringers. Stringers are more numerous and lightweight than longerons. longerons. The stringers are smaller and lighter than longeron s and serve as fill-ins. They have some rigidity but are chiefly used for giving shape and for attachment of skin. The strong, heavy longerons hold the bulkheads and formers. The bulkheads and formers hold the stringers. All of these join together to form a rigid fuselage framework. Stringers and longerons prevent tension and compression stresses from bending the fuselage. The skin is attached to the longerons, bulkheads, and other structural members and carries part of the the load.
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12. THREE-VIEW DIAGRAM 12.1 Top Vi ew
F i g 12.1 12.1
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12.2 12.2 Fr ont Vi ew
F ig 12.2 12.2
12.3 Si Si de Vie Vi ew
F ig 12.3 12.3
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13. CONCLUSION Hence multi role Aircraft has been designed with various performance and aerodynamic parameters calculation, which can carry up to payload of 8000kg i.e. armaments (Missiles, bombs, Guns etc.) It can also perform multirole tasks with the capability of STOL (Short range Takeoff and Landing).This Aircraft can fly at higher altitudes with maximum cruising speed without caught in RADAR, because it is stealth which is invisible to RADAR. The Aircraft is installed with General Electric F414- GE400 which is a low by pass turbo fan engine that can produce sufficient thrust to reach Supersonic speed at altitude. Since the Engine will be equipped with Afterburner and Thrust Vectoring so it can escape from combat field quickly and highly maneuverability. Airfoil has been selected with various considerations and calculations. NACA 64A204 airfoil has been selected as per requirement. Since it is STOL with TVC so it does not need long range take off distance. It can have more combat radius because it have more fuel capacity with Drop tanks.
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References 1. Aircraft performance and design, “John D. Anderson, Jr. University University of Maryland” 2. Aircraft design – design – A A conceptual approach, “Daniel P. Raymer president Conceptual Research Cooperation, Sylmar Califor nia” nia” 3. An example of airplane airplane preliminary design procedure – procedure – Jet Jet Transport, “E. G. Tulapurkara, A. Venkattraman, V. Ganesh” 4. Aircraft Design A Systems Engineering Approach, “Mohammad H. Satrapy, Daniel Daniel Webster College, New Hampshire, USA 5. Design of Aircraft, Aircraft, “Thomas C. Corke, University of 6.
Notre Dame”
NPTEL Airplane Design (Aerodynamic), professor E.G. Tulapurkara. REFRENCES 7. Prof. Dieter Shoclz notes on Aircraft Design
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