DESIGN PROJECT ON REINCARNATION OF CONCORDE A PROJECT REPORT
Submitted by,
ANANTHA RAMAN.L ASHOK KUMAR BALASUBRAMANIAM
97605101003 97605101011 97605101012
GOPALSAMY.M
97605101018
HERBERT JAYARAJ.J
97605101019
KARTHICK.S
94605101025
KAUSHIK.M.B
97605101026
LIJOMON.H.M
97605101029
MANIKANDAN.K
97605101030
in partial fulfillment for the
AIRCRAFT DESIGN PROJECT PHASE-1 INFANT JESUS COLLEGE OF ENGINEERING, KEELAVALLANADU, TUTICORIN.
ANNA UNIVERSITY : CHENNAI 600 025
ANNA UNIVERSITY: CHENNAI 600 025
BONAFIDE CERTIFICATE
Certified that this report “DESIGN PROJECT ON REINCARNATION
OF CONCORDE ” is the bonafide work of project members members Who carried out the project work under my supervision.
SIGNATURE
SIGNATURE
Prof. S.C.GHOSH
Mr.KARTHIKEYAN
HEAD OF THE DEPARTMENT
PROJECT GUIDE
Aeronautical Aeronautical engineering,
Lecturer, Lecturer,
Infant jesus college of engineering,
Aeronautical Aeronautical engineering,
keelavallanadu. 628851.
Infant jesus college of engineering, Keelavallanadu Keelavallanadu -628851
Internal Examiner
External Examiner Examiner
2
CONTENTS
CHAPTER NO
TITLE
PAGE
i)
ABSTRACT
6
ii)
LIST OF SYMBOLS
7
iii)
LIST OF GRAPHS
9
iv)
LIST OF DIAGRAMS
9
v)
INTRODUCTION
10
1. COMPARITIVE STUDY OF CONCORDE AIRCRAFT SPECFICATION
1.1 DIMENSIONS
12
1.2 WEIGHT SPECFICATIONS
13
1.3 PERFORMANCE PERFORMANCE SPECFICATIONS
14
2. SELECTION OF MAIN PARAMETERS
2.1 Selection Of Airfoil
15
2.2 Wing Configuration
15
2.3 Landing Gear Selection
15
2.4 Location Of Cg
15
2.5 Co-Efficient Of Lift Vs Mach Number
15
2.6 Max.L/D Vs Velocity Or Mach No
16
2.7 Weight Vs Velocity
17
2.8 Velocity Vs Wing Loading(W/S):
18
2.9 Specific Fuel Consumption Vs Vs Mach No
20
2.10 Airfoil Selection
20
2.11 Coefficient Of Lift Vs Angle Of Attack
20
2.12 Coefficient Of Lift Vs Coefficient Of Drag(Cl Vs Cd)
20
3
2.13 Maximum L/D Vs Velocity
20
2.14 Dihedral Effect
21
2.15 Velocity Vs Range: 2.16 Coefficient Of Lift Vs Coefficient Of Drag
22 23
2.17 Velocity Vs Aspect Ratio:
24
2.18 Velocity Vs Altitute 2.19 Co-Efficient Of Lift Vs Angle Of Attack (Cl Vs α ) 2.20 conclusion
25 26 27
3. WEIGHT ESTIMATION
3.1 Mission Profile
28
3.2 Approximate Weight Estimation
28
3.3 Actual Weight Estimation
30
3.4 calculation of take off weight
30
3.5 % Of Error Calculation
31
3.6 Iteration
32
3.7 conclusion
32
4. ENGINE SELECTION
4.1 Location Of Engine
33
4.2 Thrust Calculation
33
4.3 Advantages Of Low Wing
33
4.4 Disadvantages Of Low Wing
33
4.5 Thrust Vs Sfc
34
4.6 Thrust Matching
34
4.7 Calculation Of L/D
34
4.8 Conclusion
35
5. AIRFOIL SELECTION
5.1 Co-Efficient Of Lift
36
5.2 without flap
36
5.2 Drag polar
37
5.3 With flap deflection
37
5.4 conclusion
43
4
6. WING SELECTION
6.1 Equivalent Aspect Ratio
44
6.2 conclusion
45
7. WETTED SURFACE AREA AND DRAG ESTIMATION
7.1 Drag polar for cruise condition 7.2 Drag polar 7.3 Calculation of drag 7.4 conclusion
46 48 48 51
8. ESTIMATION OF RATE OF CLIMB
8.1 CALCULATION OF RATE OF CLIMB
52
8.1.1 At sea level, 8.1.2 At h=2.46km, 8.1.3 At h=4.92km 8.1.4 At h=7.38km 8.2 CONCLUSION
54
9. HORIZONTAL & VERTICAL TAIL SIZING
9.1 Horizontal Tail Sizing
55
9.2 Vertical Tail Sizing
55
9.3 Load Considerations
56
9.4 volume consideration
56
9.5 Aerodynamic Considerations
56
9.6 Drag consideration
56
9.7 correctness of ∆clmax
9.8
57
Conclusion
57
10. CALCULATION OF TAKE-OFF & LANDING DISTANCE
10.1 Length Of Take-Off Distance
58
10.2 Length Of The Landing Distance
60
10.3 Conclusion
60
11. CALCULATION OF CENTRE OF GRAVITY
61
12. THREE VIEW DIAGRAM
12
5
13. BIBILIOGRAPHY
13
ABSTRACT
As we know concorde was the only one supersonic transport aircraft. That was so famous because of the time consuming ability by it’s supersonic speed. By the way it had a performance in it’s accidents during flying.
In this design project we are going to concentrate on the possible modification which is suitable to low SFC with relative high range and endurance. Our main concentration is on the speed of the aircraft. The supersonic speed is reduced by removing the afterburners.
According to the following conclusion, we also going to do the modification in power plants and airfoil, etc. Finally we designed a 3-D view of concorde, according to the calculation made by us.
6
LIST OF SYMBOLS USED
W
Weight of aircraft
W0
Overall weight
Wf
Weight of fuel
We
Empty weight
L
Lift of aircraft
D
Drag of the aircraft
CL
Coefficient of lift
CD
Coefficient of drag
S
Wing area
b
Wing span T
Thrust
T/W
Thrust loading
W/S
Wing loading
A.R
Aspect ratio
Cr ,C ,Ct
Chord length of root,tip
Tr ,t ,tt
thickness of root, tip
Sπ
Wetted surface area
CDπ ΛL.E
Coefficient of drag of wetted surface area Sweep angle of the leading edge
ß
Dihedral angle
α
Angle of attack 7
ρ
Density(kg/m3)
C
Wing mean chord
μ
Ground friction
ν
Kinematics viscosity λ
Taper ratio
C.G
Center of gravity
R
range
E
Endurance
V∞
Free stream velocity
C
Chord
Lf
Length of fuselage
VT
Vertical tail
HT
Horizontal tail
θ
Angle of flap deflection
η0,ηi
Span station of flap
g
Gravity
s
Distance
H
Height
h
altitude
8
LIST OF GRAPHS
s.no
title
Graph 1 Graph 2 Graph 3 Graph 4 Graph 5 Graph 6 Graph 7 Graph 8 Graph 9 Graph 10
mach no vs Cl velocity vs L/D velocity vs weight velocity vs w/s velocity vs range velocity vs weight velocity vs T/w velocity vs aspect ratio velocity vs altitude Cl vs α
16 17 18 19 21 22 23 24 25
Graph 11
SFC vs thrust
34
Graph 12 Graph 13 Graph 14 Graph 15 Graph 16 Graph 17
page no
2
26
x percent vs (u/v) α vs Cl
38 39
u/v vs Y(per cent c) station vs ordinate α vs Cl
40 42 42
Cl vs Cd
43
LIST OF DIAGRAMS
s.no
title
page no
1
mission profile
2
Centre of gravity
61
3
Front view
62
4
Top view
5
Side view
9
28
63 64
INTRODUCTION
Airplane Design – Introduction Three major types of airplane design are 1. Conceptual design 2. Preliminary design 3 Detailed designs
1. CONCEPTUAL DESIGN:
It depends on what are the major factors for the designing the aircraft
A. powerplant location The power plant location is either padded or buried type engines are more preferred .Rear location is preferred for low drag, reduced shock and to use whole thrust.
B. Selection of engine: The engine to be used is selected according to the power required.
C. Wing selection: The selection of wing depends upon the selection of
low wing
mid wing
high wing
2. PRELIMINARY DESIGN: Preliminary design is based only on loitering; U is the mathematical method of skinning the aircraft after skinning the aircraft looks like a masked body. Preliminary design is done with the help of FORTRAN software.
10
2. DETAILED DESIGN: In the detailed design considers each and every rivets, nuts, bolts, paints, etc. In this design the connection and allocation are made.
11
1.COMPARATIVE STUDY
TABLE 1.1 DIMENSION:
S.no: Aircraft name
Length Height Crew Wing span (m) (m) (m)
1
Boeing 2707-sst
93.27 m
2
3
Wing area (m2)
Aspect ratio
14.1 m
3
32.23meters 358.25 m²
AEROSPATIALE- 62.10 BAC meters CONCORDE
11.40 meters
3
25.56 meters
385.25 1.6 sq_meters
TUPOLEV TU144
12.85 meters
3
28.80 meters
438.00 1.8 sq_meters
65.70 meters
12
2.8
TABLE 1.2 WEIGHT:
S.n o:
Aircraft name
Empty weight
Loaded Weight
Maximum takeoff weight
1
Boeing 2707SST
287,500 lb (130308 kg)
75,000 lb (34020 kg)
675,000 lb (306175 kg)
2
AEROSPATIAL E-BAC CONCORDE
78,700kg (173,500lb),
12,700kg (28,000lb).
185,065kg (408,000lb).
3
TUPOLEV TU144
85,000 kg
19,500 kg
180,000 kg
13
TABLE 1.3
PERFORMANCE:
S. No:
Aircraft name
Speed Mach (km/hr No: )
Range (km)
Service ceiling (m)
Rate of climb (m/s)
W/S (kg/m2)
T/W
1
Boeing 2707
2900
2.7
6840
18,300
25.40
854.64
.15
2
AEROSPATI ALE-BAC CONCORDE
2180
2.2
6580
18,290
25.40
217.85
0.374
3
TUPOLEV TU-144
2500
2.4
6500
18300
25.40
410
.110
14
2. SELECTION OF MAIN PARAMETERS FOR AIRCRAFT DESIGN
2.1 SELECTION OF AIRFOIL :
Selection of airfoil is depend up on the need of the weight of the aircraft . The airfoil selection is an very important in the a/c design.
2.2 WING CONFIGURATION :
The dihedral effect is created by wing dihedral angle гo , which is positive for tip chord above the root chord.
2.3 LANDING GEAR SELECTION :
The landing gear selection is depend upon the types of aircrafts. For our aircraft we use tricycle type landing gear. So the visibility of the pilot will be h igh because of the use of nose wheel.
2.4 LOCATION OF cg :
Location of cg is the important factor which responds to the stability of the aircraft . It has some limits and thus these both are inter related.
2.5
MACH NO Vs Cl :
As the mach no increases, the value of Cl also increases because of airflow velocity past over the surface of wing increases ,this will gradually increases the coefficient of lift. when the velocity reaches the stalling velocity ,the value of Cl started declines. The graph between mach no vs Cl,
15
2.6 VELOCITY Vs L/D:
The graph is plotted between velocity and L/D.
16
The above plot is drawn between (L/D) and Velocity.From the above graph we get the optimum velocity as 605m/s.
2.7 VELOCITY Vs WEIGHT :
The graph is drawn between veloicity & weight .It is plotted between the overall weight of similar type of subsonic twintail fighter aircraft and the velocity of the corresponding aircraft for our specification of aircraft, the weight of aircraft is 80,500kg in the corresponding velocity of 605 m/s.
17
2.8 VELOCITY VS WING LOADING (W/S) :
The graph is drawn between wing loading & velocity.wing loading is the ratio of weight to the wing span.
18
The above graph is plotted between velocity and w/s (wing loading).from the above graph we get the optimum value of w/s as 210kg/m2 and the optimum velocity is 605 m/s.
19
2.9 SPECIFIC FUEL CONSUMPTION Vs MACH NO :
The weight of fuel consumed per unit thrust per unit time. Mach number is the ratio between velocities of aircraft to velocity of sound. The variation of thrust with subsonic mach no is drawn for ratio as altitude.
2.10 AIRFOIL SELECTION :
The general dynamics designers examined two class of configuration. 1. The conventional wing body arrangement and 2. The blended wing body arrangement. The blended wing body configuration provides two important advantages. It was relatively natural to includes fore body strakes in such a blended configuration and the area ruling was more easily carried out. So I select the blended wing body. 2.11 COEFFICIENT OF LIFT Vs ANGLE OF ATTACK :
The experimental data indicate that coefficient of lift varies linearly with angle of attack. Thin airfoil theory which is the subject of more advanced book of aerodynamics also predicts the same type of linear variation. The slope of the linear portion of lift curve is designed as α =
dC L
dα
= lift slope
2.12 COEFFICIENT OF LIFT Vs COEFFICIENT OF DRAG :
For every aerodynamic body there is a relation between coefficient of lift and drag that can be graph. Both equation and graph is called drag polar. 2.13 MAXIMUM L/D Vs VELOCITY :
Speed, altitude, range were the primary performance goals. For supersonic fighter aircraft high value of L/D and W/S were important. The variation of L/D max with mach no is shown in fig. Here we see example of how dramatically the aerodynamic characteristic of and an air plane change we can go from subsonic to supersonic speeds. The value of L/D max is almost in half of drag divergence/wave drag effects at supersonic speeds on the other hand the resulting value of L/D max is 6.5 at mach3.
20
2.14 VELOCITY Vs RANGE :
Range is the total distance traversed ofan airplane on one load of fuel.we denote range R. R=2/Cl*
2 / ρ α * S *(Cl/CD)*(wo.5-w1.5)
The above plot is drawn between Range and Velocity.From the above graph we get the optimum velocity as 605 m/s and the optimum Range as 6500km2.15
21
2.15 VELOCITY Vs WEIGHT :
The graph is plotted between the velocity and the overall weight of similar type of twintail subsonic fighter aircraft.by that graph we get the optimum value of overall weight of the aircraft.
The above plot is drawn between Weight and Velocity.From the above graph we get the optimum velocity as 605 m/s and the optimum Weight as 185065 kg.
22
2.16 VELOCITY Vs T/W :
The graph is drawn between Thrust/Weight & velocity. In addition to Clmax ,the other important parameter affecting take-off & ranging distance is T/W. The choice of a too high T/W is determinant to efficient cruise. The value T/W is 0.355 in the corresponding velocity of 605m/s.
The optimum T/W from the above graph is 0.355 in the corresponding velocity of 605 m/s.
23
2.17 VELOCITY Vs ASPECT RATIO The graph is drawn between the aspect ratio & velocity, the choice of
low aspect ratio the wing having full span leading edge flaps, the vertical tails are casted outward by 28o & incorporate conventional rudders
The above plot is drawn between AR (aspect ratio) and Velocity.From the above graph we get the optimum velocity as 605 m/s and the optimum Aspect ratio as 6.
24
2.18 VELOCITY Vs ALTITUDE : The graph is drawn between the altitude &velocity. It is main design
parameter. The optimum altitude is 9.850 km in the corresponding velocity of 605 m/s
25
The above plot is drawn between Altitude and Velocity.From the above graph we get the optimum velocity as 605 m/s and the optimum Altitude as 9.85km.
2.19 COEFFICIENT OF LIFT Vs ANGLE OF ATTACK (Cl vs
α
):
The experimental data indicated that Cl varies linearly with α over a large range of angle of attack. Thin airfoil theory which is the subject of more advanced book on aerodynamics also predicts the same type of linear variation, slope of the linear portion of the lift curve is designed as α o = ∂cl / ∂α =lift slope .at the angle of attack=12o,the Clmax is 1.4.
26
2.20 CONCLUSION:
SL.NO
PARAMETER
OPTIMUM VALUES
1
Altitude
18,900 m
2
Velocity
605 m/s
3
Range
6500 km
4
Weight
1,70,095 kg
5
Aspect Ratio
1.628
27
3. WEIGHT ESTIMATION
3.1 MISSION PROFILE:
The mission profile for our aircraft is as follows,
1-2:warm-up and take-off 2-3:climbing 3-4:cruising 4-5:descending 5-6:landing
3.2 APPROXIMATE WEIGHT ESTIMATION :
overall weight of the aircraft, wo=wcrew+w payload+wfuel+wempty The mission profile of the fighter aircraft the loitering is neglected (fighter aircraft loitering is 10 minutes allowed).
Mission profile segment
a/c weight at the end of the mission segment
28
weight fraction
=
a/c wt. at the beginning of the mission segment
= Wi / W(i-1) Range, R=L/D* ln(Wi / W(i-1) )*V/C
In fighter aircraft , W10
W2*W3*W4*W5*W6* =
W1 W1*W2*W3*W4*W5* In take off, W2/W1 = 0.99 In climbing flight mission, W3/W2=1.0065-0.0325*0.56 =0.9883 In cruising, W4/W3=exp(-RC/V*(L/D)max) R- the range in nautical mile C- SFC in lb/lb*hr V- velocity in knots W4/W3= e-(291*0.6/458*7) W4/W3 = 0.95 Decending , W5/W4= 0.99 Landing and shut down, W6/W5= 0.995 W6/W1= 0.768 Then the fuel weight fraction is , Wf /Wo = 1-(W6/W1) = 1- (0.768) Wf /Wo = 0.238 Wcrew + W payload
29
W0 = 1-(Wf /Wo) – (We/Wo) =
(128+12700)/(1-0.232-0.55)
W0 = 58844.03 kg
3.3 ACTUAL WEIGHT CALCULATION :
In warm up and take off, W2/W1= 0.99 Climbing flight mission, W3/W2= 1.0065 – 0.0325 * 0.33 W3/W2 = 0.9957
In cruising, W4/W3= e-(204.95*0.6/233.4*7) W4/W3= 0.93 In decending, W5/W4=0.99 In landing, W6/W1= 0.99 Therefore, W6/W1 = 0.761 Then the fuel fraction is, Wf /W0= 1- (W10/W1) =1 - 0.761 Wf /W0 =0.238 Wactual= 0.238*1.06 = 0.25288 As we know, We/Wo = A*W0c
3.4 CALCULATION OF TAKE-OFF WEIGHT (T/W):
30
(1)cruise(T/W)=1/ (
L
D
For cruise flight (
L
D
(
) max cruise = (
L
D
L
D
(
) max cruise )max *0.866
) max cruise = 7.8 * 0.866 = 6.7548
T W
)cruise =
1/ (
L
D
)max cruise = 1/ 6.5748 = 0.1480
(2)For loitering (
T W
)loitering = 1/ (
L
D
=
)max
0.1282
(3)For take-off (
T W
)take −off = (
T W
Wcruise
)cruise * (
WT .O
) *(
T T .O T cruise
)
= 0.148 * (8500/12,500 ) * (52.0 /81.0 ) = 0.064 Thus the calculated (T/W) ratio and optimum (T/W) ratio values are approximately met each other.
For supersonic aircraft
A = 1.02,C =-0.06 Hence We/Wo = 1.02*30500-0.06 = 0.548
Then, Wtotal = Wcrew + W payload + W empty + W fuel Wcrew + W payload = 1 – Wf /W0 – We/W0 W0 = (128+ 12700/1-0.238-0.548) =59906.54kg Wo = 59906.54kg
31
3.5 % OF ERROR CALCULATION :
error % =
W actual
– Wapprox Wactual
= (59906.05-58844.03 /58844.03)*100 error % = 1.172%
3.6 CALCULATION OF MAXIMUM WEIGHT USING ITERATION METHOD :
i) We/Wo=1.02*59906.05-0.06=0.516 ii) We/Wo=1.02*58844.03
-0.06
=0.517
iii) We/Wo=1.02*58843-0.06=0.527 From ii & iii The gross weight of the aircraft is 58844.03 kg
3.7 CONCLUSION:
Wo = 59906.54kg
This is the total weight of the aircraft which we considered from the above calculation
32
4. ENGINE SELECTION 4.1 LOCATION OF ENGINE:
Two engine configuration was selected. This type of engine is podded engine.
4.2 THRUST CALCULATION:
T= W0*(T/W) = 33732.71*0.355 = 11975kg = 26400.41lb Thrust per each engine = 13217.56lb
4.3 ADVANTAGES OF LOW WING TYPE AIRCRAFT :
1) Integrated structure of wing having maximum strength for carrying the maximum wing loading 2) Because of low wing type,the downwash to the horizontal stabilizer is greatly reduced. 3) Maintenance of engine in low-wing type is easily possible.
4.4 DISADVANTAGES OF LOW WING TYPE AIRCRAFT :
1 )Requires long landing gear for maintaining the optimum clearance between ground and engine. 2)chance for entering the dust particles into engine which seriously affect the engine efficiency.
33
4.5 THRUST Vs SFC :
The above engine meet the thrust requirement of our aircraft with minimum sfc. The configuration of the engine is podded engine. So from this above graph at 0.789 is the optimum sfc at 553450 N of thrust.
4.6 THRUST MATCHING:
For further selection of aircraft parameters we consider the thrust matching between the optimum T/W by plotting graph and the T/W ratio obtained by using wetted aspect ratio.
4.7 CALCULATION OF L/D:
34
aspect ratio Wetted aspect ratio = (wetted surface area / reference surface area) Here the wetted surface area represents the wing area and the reference surface area represents the extra projection from fuselage or wing like canard surface etc. From the historical data , S wet = 3 8.4/11.2 = 3.4285 S ref wetted aspect ratio= 1.75/3.4285 = 0.511 From the graph, for wetted aspect ratio=1.967 (L/D)max = 7.8
(from aircraft data book)
4.8 CONCLUSION
Hence from the above calculations
Thrust per each engine = 13217.56lb
Hence turbojet engine is used for propulsion, with the absence of the after burner to reduce the speed to subsonic level.
35
5. AIRFOIL SELECTION The optimum altitude = 9.85km The density at this altitude = 0.43966 kg/m3
5.1 COEFFICIENT OF LIFT (CLmax)
Vstal = 0.25*Vcruise =
0.25 * 552.8 = 138.2
2(w/s) CLmax = Ρalt*v2stall = 2*221.4 *9.81/(0.43966* 138.22) = 0.5173 5.2 WITHOUT FLAP ANGLE OF ATTACK(α)
CO-EFFICIENT OF LIFT (CL)
CO-EFFICIENT OF DRAG (CD)
-8 -6 -4 -2 0 2 4 6 8 10 12 16 17
-7 -5 -3 -0.1 0.1 0.3 0.5 0.75 1 0.75 0.45 0.35 0.25
0.045 0.0289 0.0229 0.01287 0.01287 0.01823 0.02895 0.04988 0.0792 0.1086 0.14352 0.17316 0.20583
36
5.3 DRAG POLAR (CD Vs CL) :
The drag polar is the drag minimum at the graph CL and CD . The graph is plotted CL and CD. This graph is used to find the value of minimum drag co-efficient is 0.006647. Station and ordinates given in % of airfoil chord (x/c,y/c) The graph is plotted between stations in percent of chord x/c and y/c. The leading edge radius is 0.256m. 5.4 WITH FLAP DEFLECTION :
ANGLE OF ATTACK (α) -12 -10 -8 -6 -4 -2 0 1 2 4 6 7
CO-EFFICIENT OF LIFT
-0.9 -0.7 0.5 0.7 0.9 1.1 1.3 1.5 1.6 1.8 1.7 1.65
Landing CLmax landing = 1.75*(25% of CLmax take off ) =0.765625 CD = CDO +KCL2 CDO = Cfe (Swet/Sref ) ; K = (1/ π *AR*e) ; e = 0.79 Reynolds number, ρ Vstall Cm Re = μ Cm = Cv + Ct 2 = 3.57m
37
CO-EFFCIENT OF DRAG 0.006647 0.04503 0.02845 0.045 0.066 0.093 0.125 0.163 0.184 0.229 0.205 0.1946
Vstall = 30m/s Re = 10.928*106 From the above data we selected aerofoil NACA 64006 for the required co-efficient of lift.then the plots are as follows
X (percent c)
(u/v)2
0 0.5 0.75 1.25 2.5 5 7.5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 100
0 0.912 1.016 1.084 1.127 1.152 1.167 1.179 1.195 1.208 1.217 1.225 1.230 1.235 1.220 1.119 1.163 1.133 1.102 1.069 1.033 0.995 0.957 0.918 0.878 0.839
38
y (percent c)
(u/v)
0
0
0.658
0.955
0.794
1.008
1.005
1.041
1.365
1.062
1.875
1.073
2.259
1.008
2.574
1.093
3.069
1.099
39
3.437
1.103
3.704
1.107
3.884
1.109
3.979
1.111
3.992
1.105
3.883
1.091
3.684
1.078
3.411
1.064
3.081
1.054
2.704
1.034
2.291
1.016
1.854
0.997
1.404
0.978
0.961
0.958
0.550
0.937
0.206
0.916
0
0.901
40
For the selected aerofoil,
Upper surface
Lower surface
Station
ordinate
Station
ordinate
0
0
0
0
0.459
0.542
0.541
-0.442
0.704
0.664
0.796
-0.524
1.198
0.859
1.302
-0.645
2.440
1.208
2.560
-0.836
4.934
1.719
5.066
-1.087
7.432
2.115
7.568
-1.267
9.933
2.444
10.067
-1.140
14.937
2.970
15.063
-1.624
19.943
3.367
20.057
-1.775
24.952
3.667
25.048
-1.877
29.961
3.879
25.087
-1.877
34.971
4.011
35.029
-1.951
39.981
4.066
40.019
-1.924
41
44.991
4.014
45.009
-1.824
50.000
3.878
50.000
-1.672
55.008
3.670
54.992
-1.480
60.015
3.402
59.992
-1.480
65.020
3.080
64.980
-1.020
70.023
2.713
69.977
-0.768
75.025
2.307
74.975
-0.517
80.024
1.868
79.976
-0.276
85.020
1.410
84.980
-0.064
90.015
0.940
89.985
0.094
95.007
0.473
940993
0.159
100
0
100
0
L.E. radius:0.256 Slope of radius through L.E.:0.084
42
43
For flap deflection = 60o ; CLmax = 1.75 Here the CL available without flap is 0.9 . Then the ∆ CLmax required from the flap is 0.16 . so 10-15 o deflection is enough for our lift requirement . CLref = CLavailable + ∆ CLmax
∆ CLmax = CLreq – CLwithout flap = 1.06-0.9 =0.16 with 60o flap deflection = 1.75 we need 15o flap deflection for our aircraft
5.5 CONCLUSION
We plot the graphs between cl and cd, α and cl and hence concluded that the above calculated airfoil is suitable for our aircraft.
44
6.WING SELECTION 6. 1 EQUIVALENT ASPECT RATIO :
A.req = a*Mcmax a=4.11,c=-0.622 A.req=7.71 Optimum w/s=210 kg/m3 S = w/210=33732.71/210=160.63m2 AR = b2/s b2 = AR*S = 6.2*160.63 = 995.906 b = 31.5 m c r = 2b/AR(1+ λ )
λ = 0.4 for AR(6-10) cr = 7.2m
λ =ct/cr => ct=0.4*7.2 ct=2.88 m Mean Aerodynamic chord=2cr [(1+ λ + λ 2)/1+ λ ]/3 C
=5.34 m
Volume of fuel=Wt.of fuel/ρfuel Density of fuel=800 kg/m3 Wf /W0=0.238 Wf =0.238*33732.71 Wf =8028.38 kg
45
Total volume of fuel=8028.38*9.81/800*9.81 =10.03 m3 Assume 80% of fuel carring in the wing, V=[(t/c)*c(0.5*c)*0.5*b*0.75]*2 8.028=[(t/c)*5.34*0.5*5.34*0.5*15.5*0.75]*2 t/c=0.0486
∴ t/cr =0.0486=> troot=34.8 cm t/c tip=0.0486=>ttip=9.9 cm For subsonic, Sweep angle tan λ LE=tan Λ c/4+[(1- λ )/AR(1+ λ )] Here Λ c/4 =34 o Tan λ LE=0.6745+0.07142 λ LE=36.72 o
6.2 CONCLUSION
From the above calculations we concluded that the low wing is suitable for the designed aircraft.
46
7. ESTIMATION OF WETTED SURFACE AREA AND DRAG 7.1 DRAG POLAR FOR CRUISE CONDITION :
1)fuselage length = awoc from historic data,a=0.366;c=0.42
lf =0.366*(33732)0.42 =29.91m lf /df =7-11 df =29.19/7 df =4.17m fuselage s π = π /4*df 2 s π =9.12m2 2) wing area=bw*tw =34.87*10-2*31.5 sw π =10.98m2 3)horizontal tail: sht =tht*spanht bht =
(sht *A.R)
s=690ft2 ;A.R=3.5 A.R=b2/s b=14.978m tht=t/c*ctip =0.0486*2.88 tht=13.99cm -2 sht=13.99*10 *14.978 sht π =2.096m2 4)vertical tail(twin tail):
47
svt =tvt *bvt A.R=0.6-1.4 A.R=1 A.R=b2/s ,s=650ft2 =60.385m2 1=b2/60.385 b=7.77m svt=13.99*10-2*7.77*2 svt π =2.174m2 5)Engine: a = π /4*d2 = π /4*1.192 for 2 engines, a=2.224m2 6)under carriage, s=(2.224*0.1)+2.224 s=2.4464m2 7)1/4 flap(15o) s π = θ /360* π *r 2 r= 0.2*cr =1;s π =15/360* π *1 s π =0.1308m2 8)full flap:(25o) s π =25/360* π *(0.2*5.34) 2 s π =0.2488 m2
7.2 DRAG POLAR :
48
cDt=cDo+cDo(others)+k k=1/( π A.Re)= 0.115 cruise: 4
cD(others) =
∑=
(cD π *s π )/sw
i 1
=(0.2736+0.0167+0.017+0.08787)/160.63 cD(others) = 0.00246 Take-off: 7
cDo(others) =
∑=
(cD π *s π )/sw
i 1
cD(others) = 0.0323 Landing: 6
cD(others) =
∑=
(cD π *s π )/sw+(cD π *s π )8/sw
i 1
cD(others)= 0.00328
7.3 CALCULATION OF DRAG :
At h=0;T=288.16K;ρ=1.225kg/m3 a= 1.4*287*288.16 =340.26m/s 2(w/s) cl= ρv2 2(33732.68/160.63)*9.81 = 1.225*v2 cl =3363.3/v2
C D
= ( C DT . O + C DW )/ 1 − M ∞2
C D
= ( C D0 +K C L )*1.05
0
T
2
49
D=( C DT * W 0 )/ CL
V(m/s)
CL
M=v/a
C D
C D
C D
C D
D (KN)
44
1.737
0.129
0.00323
0.359
0.365
0.747
142.3
88
0.4340
0.258
0.00323
0.03386
0.03838
0.063
48.03
132
0.193
0.388
0.00323
0.0344
0.03835
0.0447
76.64
176
0.1085
0.517
0.00323
0.0136
0.01692
0.02918
78.49
220
0.0694
0.646
0.00323
0.01275
0.0161
0.017486
83.37
T .O
W
0
T
At h=2.46km T=272.57k ρ=0.9784 kg/m3 V(m/s)
CL
M=v/a
C D
C D
C D
C D
D(KN)
44
2.175
0.1329
0.00323
0.5562
0.5803
1.135
146.68
88
0.543
0.2659
0.00323
0.0461
0.0537
0.089
54.23
132
0.2416
0.3988
0.00323
0.0188
0.0340
0.0267
36.57
176
0.1359
0.5318
0.00323
0.0142
0.0199
0.0231
56.24
220
0.0870
0.6647
0.00323
0.0130
0.0206
0.0222
78.35
T .O
At h=4.92km T=254.05k ρ=0.0.7214 kg/m3
50
W
0
T
V(m/s)
CL
M=v/a
C D
C D
C D
C D
D(KN)
44
2.88
0.137
0.00323
0.666
0.6784
1.42
166.13
88
0.72
0.274
0.00323
0.072
0.075
0.144
66.38
132
0.3209
0.411
0.00323
0.024
0.0287
0.04254
43.89
176
0.2805
0.549
0.00323
0.0159
0.0228
0.0278
51.1
220
0.1155
0.686
0.00323
0.0137
0.0232
0.0259
71.23
T .O
W
0
T
At h=7.38km T=240.12k ρ=0.5635 kg/m3 V(m/s)
CL
M=v/a
C D
C D
C D
C D
D(KN)
44
3.776
0.141
0.00323
1.65
1.669
2.47
166.54
88
0.944
0.283
0.00323
0.1146
0.123
0.236
32.72
132
0.419
0.425
0.00323
0.0323
0.04
0.063
47.75
176
0.236
0.566
0.00323
0.0186
0.0269
0.0349
43.93
220
0.151
0.708
0.00323
0.0148
0.0258
0.0298
65.306
T .O
W
51
0
T
At h=9.85km T=224.23k ρ=0.4673 kg/m3 V(m/s)
CL
M=v/a
C D
C D
C D
C D
D(KN)
44
4.55
0.146
0.00323
2.392
2.421
2.44
174.5
88
1.138
0.293
0.00323
0.1611
0.1718
0.2367
28.84
132
0.506
0.439
0.00323
0.0416
0.0498
0.0732
34.91
176
0.284
0.586
0.00323
0.0214
0.0303
0.0415
48.35
220
0.182
0.7329
0.00323
0.0160
0.0282
0.0336
61.09
T .O
W
0
T
7.4 CONCLUSION
Hence the wetted surface area and the drag area is calculated from the above calculations.
8 RATE OF CLIMB ESTIMATION
52
Thrust available = 680.8 KN Thrust required = F. σ 1.15 =90*((20-h)/(20+h)) 1.15 At sea level, F=680.8 KN At 2.462km, F=965.47 KN At 4.924 km, F=914.40 KN At 7.368 km, F=830.84 KN At 9.85 km, F=713.87 KN
8.1 CALCULATION OF RATE OF CLIMB
8.1.1 At sea level, V(m/s)
D(KN)
T(KN)
R.C=(T-D)*V*60/W0(Km/min)
-
-
680.8
-
805
48.03
680.8
16.84
605
76.64
680.8
12.08
694
78.49
680.8
13.83
8.1.2 At h=2.46km,
53
V(m/s)
D(KN)
T(KN)
R.C=(T-D)*V*60/W0(Km/min)
-
146.68
680.8
-
805
54.23
680.8
16.6
605
36.57
680.8
12.88
694
56.24
680.8
14.32
8.1.3 At h=4.92km
V(m/s)
D(KN)
T(KN)
R.C=(T-D)*V*60/W0(Km/min)
-
166.13
680.8
-
805
66.38
680.8
16.35
605
43.89
680.8
12.73
694
51.1
680.8
14.40
8.1.4 At h=7.38km V(m/s)
D(KN)
T(KN)
R.C=(T-D)*V*60/W0(Km/min)
-
166.54
680.8
-
805
32.72
680.8
17.24
605
47.75
680.8
12.6
694
43.93
680.8
14.6
54
8.2 CONCLUSION
The rate of climb is calculated for the different values of h.
55
9 HORIZONTAL AND VERTICAL TAIL SIZING 9.1 HORIZONTAL TAIL :
The horizontal tail dihedral angle =5 o The root chord = 4.27m The tip chord = 1.71m Tapper ratio =0.4 Sweep angle = 41 o SHT=CHT* C w *SW/LHT Where,
.
SHT-horizontal tail surface area C w -wing mean chord sw-wing area LHT-distance from ¼ chord of the horizontal stabilizer to the wing ¼ chord 0.4*5.34*160.63 LHT = 64.1 LHT = 5.35m
9.2 VERTICAL TAIL SIZING :
Vertical tail area for vertical tail,
L VT
b w × Sw
= CVT ×
56
SVT
L VT Distance
from
¼
chord
of
the
vertical
stabilizer
to
the
wing
1/4chord. LVT
= 0.04*8.20*38.4/8 =
1.5744 m
9.3 LOAD CONSERVATION :
a. Air loads:
Maneuver, component of interaction, gust load, and control deflection buffet load. b. Inertial loads:
Acceleration, rotation, vibration, flutter and other dynamic loads.
c. Landing loads: Breaking loads, vertical load, factors, skin up, spring back, and arrested loads. d. Take-off loads:
Aborted load, catapult loads. e. Power-plant loading:
Thrust, torque, hammer shock, vibration, duct pressure. f.
Other loadings:
Bird strike, pressurization, actualization, fuel pressure & crash.
9.4 VOLUME CONSIDERATION :
i. ii.
passenger requirement crew requirement
iii.
fuel storage requirement
iv.
buried engine and their inlets
v. vi.
wing loading gear attachments
9.5 AERODYNAMIC CONSIDERATION :
57
i. ii.
fuselage shapes fuselage fineness ratio
9.6 DRAG CONSIDERATION :
i.
fuselage
ii.
wing & horizontal, vertical stabilizer
iii.
engine
iv.
landing gear
At root Reynolds number,
ρ× V × c r µ 0.19475 × 315 × 3.813 = 1.42 ×10 -5 Rer = 16.44 ×10 6 Rer =
At the tip Reynolds number, Re t Re t
ρ× V × c µ = 6.5776 ×10 =
t
6
9.7 CORRECTNESS OF ∆clmax.
Take-off:
∆CLmax
take−off
= 1.05 CLmax
take−offreq
− CLmax
available
= 1.05[ 1.36-1.1]
∆CLmax
take−off
= 0.273
Landing:
∆C
L max Landing
= 1.05 C
L maxLandingfreq
−C
= 1.05[ 1.616-1.1]
∆C
Lmax Landing
= 0.5418 58
L maxavailable
9.8 CONCLUSION
The horizontal and vertical tail sizing calculations are calculated above.
10. CALCULATION OF TAKE-OFF AND LANDING DISTANCE
10.1 LENGTH OF TAKE OFF DISTANCE : (1)Ground run:
T-D=μ(W-L)+(W/g)*(
dv / dt )
------------(1)
(T-D)-μ(W-L)= (W/g)*V*( dv / ds ) S 1
V 1
∫ ds =
S1=
(W/g)
0
∫ VdV 0
[(T-D)-μ(W-L)] 2
W(V 1 /2g) [(T-D)-μ(W-L)] V1=1.2*Vstall S1=
=1.2* 2 * ( w / s ) /(ρ * C l maxinf lap ) =1.2* v1=1.0829 km (2)Transition run :(S2)
μ(W-L)=0 (T-D)=(W/g)*V* ( dv / ds ) ds =(W/g)*V* dv (T-D) S 1
S2=
∫ ds 0
59
--------------- (2)
V 2
(W/g)
∫ VdV
V 1
= (T-D) = W(V2 − V 1 ) 2g(T-D) 2
2
V1=44m/s;V2=88m/s;D=48.03KN S2= 599.254*(882-442) 2*9.81(132.17-48.03) S2=1.59 km (3)Climb: (T-D)-Wsinθ=0 cot θ=S3/H ; tanθ=H/S3 cotθ=cosθ/sinθ = 1 − sin 2 θ sin θ 1− (
−D
T
W
)2
= T
−D W
H*
( W2 − ( T −
= (T-D) 2463* S3 = (132.17-48.03) S3=7.052 km S=S1+S2+S3 = 1.082 + 1.59 +7.015 = 9.68 km
60
2
D)
10.2 LENGTH OF LANDING DISTANCE :
(1)Descend: H* (W 2 − ( D − T )2 ) S1= D-T 2463* = 48.03-132.17 S1=-11.1 km
(2)Transition: 2 2 W* (V2 − V 1 )
S2 = 2g(D-T)
2 2 599.25 * (88 − 44 )
S2 = 2*9.81*(48.03-132.17) S2= -2.11 km (3)Ground run : 2
W(V 1 /2g) S3= [(D-T)-μ(W-L)] S3= 0.69 km S=S1+S2+S3
61
= 11.1-2.11+0.69 S = 9.68km 10.3 CONCLUSION :
TAKE-OFF DISTANCE = 9.68 km ; LANDING DISTANCE = 9.68 km
11. CALCULATION OF CENTER OF GRAVITY
The major weight components for which we have some idea of this locations are the engines,the crew and payload.Using the information we can make a very preliminary estimation of the location of the center of gravity.The tail,fuselage and wing also contribute the location of ‘cg’.we can take them in to account later,when there was a better design made for than for now,ever if they are taken into account,they will give an approximate value of cg.
62
12 THREE VIEWS DIAGRAM
63
64
65
13 BIBLIOGRAPHY
66
