wing spar presentation
TRANSCRIPT
1CADES Proprietary Confidential
Design of Front and Rear Spars for The
Trainer Aircraft Wing.
2CADES Proprietary Confidential
TEAM
Team Members :Akshay A. Pradeep S. Shet
Pavan Kumar N. R.Raghunandan M.Lakshmana H. B.Chetan A. V.
Guide : Mr. H. N. Athavale
Co-ordinator : Mr. Umanath Nayak
CAE
CAD
3CADES Proprietary Confidential
OBJECTIVE
CAD
To generate the CAD model of wing using the available data and
prepare the assembly of all components
CAE
Determine the Spar locations with respect to chord length.
Determine the dimensions for flange and web of the spars.
Estimate the number of ribs and their positioning
4CADES Proprietary Confidential
SCOPE OF THE PROJECT
Estimation of spar position.
Dimension calculations of front and rear spars.
Calculations for number of ribs and their positions.
CAD
Profile creation of the wing using the given NACA standards.
Creation of the wing geometry
Use available data to develop CAD models for each individual component
Prepare an assembly of all components using CATIA
CAE
5CADES Proprietary Confidential
Root chord : 2400 mmTip chord : 700 mmSemi Span length : 5500 mmExposed Span : 4750 mmAirfoil (root) : NACA 64A
1215
(tip) : NACA 64A1210
Aircraft weight : 14000 NLift Load : 6gDesign Factor : 1.5Given Spar Position(in % of chord length)
Front Spar : 18-25Rear Spar : 62-70
INPUT
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DERIVED INPUT
●Limit load : 14000 * 6= 84000 N
●Design Load : 84000 * 1.5= 126000 N
●Load on semi-span : 126000 / 2= 63000 N
●Exposed wing area : 7.3625 E6 mm2
●Pressure load on wing : 63000 / 7.3625 E6 = 8556.87 E-6 N/mm2
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WING GEOMETRY
ALL DIMENSIONS ARE IN mm
RO
OT
CH
OR
D SWEEP AT ¼ CHORD
4750
700
2400
TIP
CH
OR
D
LEADING EDGE
TRAILING EDGE
Top View [RH]
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AIRFOIL
Generate the aerofoil section using the Coordinates of NACA 64A
1215 and NACA 64A
1210.
[source : http://www.pdas.com/sections6a.htm]
Generate the CAD model of the wing using CATIA- V5.
Aerofoil at Root NACA 64A1215Aerofoil at Tip NACA 64A
1210
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DESIGN PROCEDURE
Calculation of the Shear force, Bending moment & Torsion for the
given load.
Calculation of load distribution between the front and rear spar.
Estimation of spar positions.
Generation of CAD Model and Drafting.
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DESIGN PROCEDURE
Divide the wing area into number of divisions.
Calculate the chord length at each section.
Determine the C.G of each area.
Calculate the shear force, bending moment and Torque at the respective
sections.
Shear force =pressure*area.
Bending moment=shear force*CG distance.
Torque = Shear force*Distance b/w CG and CP.
11CADES Proprietary Confidential
METHODS AND METHODOLOGY
A10
A9
A8
A7
A6
A5
A4
A3
A2
A1
475
2400
700
ALL DIMENSIONS IN mm
L1
L2
L9
12CADES Proprietary Confidential
Chord Length, L1= L
root-((L
root-L
tip) / S) * x
At section 2, L1 = 2400-((2400-700)/4750)*4275
L1 = 870 mm
Area of Trapezium, A1 = 0.5*(L
1+L
tip)*h
A1 = 0.5*(870+700)*475
A1 = 373 E3 mm2
CG of Trapezoid Section = h/3*((Ltip
+2L1)/(L
tip+L
1))
CG=475/3*((700+2*870)/(700+870)) CG = 246 mm from L
tip
DESIGN PROCEDURE
S
xLroot
Ltip
L1
A1
h
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DESIGN PROCEDURE
Limit load = 84000 NDesign Load = Limit Load*Design factorDesign load on wing, = 84000*1.5 = 1,26,000 NDesign load on semi-span wing, = 63000 Npressure load on wing [P] = 8556.87 E-6 N/mm2
Load At Section 2, = P2+P
1 = P*A
2+P
1
= 8557 E-6 * 453625 + 3190.65 = 3881.6 + 3190.65 = 7072.25 N
Bending Moment At Section 2, M2 = P
2 * CG
2 + P
1 * (CG
1 + L
2)
M2 = 3881.6 * 230 + 3190.65 * (229 + 475)
M2 = 3248260 N-mm
14CADES Proprietary Confidential
SHEAR FORCE
Root 0 475 950 1425 1900 2375 2850 3325 3800 4275 TIP 4750
0
10000
20000
30000
40000
50000
60000
7000063000
53590.65
44872.25
36844.85
29508.4
22862.9
16908.39
11644.857072.25
3190.653190.65
Shear force diagram for the wing
Wing span [Root to tip] [mm]
She
ar f
orce
[N
]
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BENDING MOMENT
ROOT 0
475 950 1425 1900 2375 2850 3325 3800 4275 TIP 4750
0
20000000
40000000
60000000
80000000
100000000
120000000
140000000123020000
95259000
71809000
52341000
36527000
2403900014548000
77278603248260781700 0
Bending moment diagram for the wing span
Wing span [root to tip] [mm]
Ben
ding
mom
ent [
N-m
m]
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LOAD DISTRIBUTION
Centre of Pressure, CP = 45% of Chord Length (C) from LE [870mm] [1]Front Spar Position = 25% of C from LE [217.5mm]Rear Spar Position = 62% of C from LE [539.4mm]
Chord Length 'C'
45% of C
25% of C a b
c
62% of C
RA R
B
FS RS
ChordCP
a=174mm
b=148mm
c=322mm
C=870mm
17CADES Proprietary Confidential
Shear Force Distribution:
Shear Force on Front Spar, = Load * b/cAt Section 1, SF
FS = 3190.65 * (148/322)
SFFS
= 1465.974 N
Shear Force on Rear Spar SFRS
= 3190.65 - 1465.974 SF
RS = 1724.676 N
SF on Front Spar = 45.9% of total loadSF on Rear Spar = 54.1% of total load
18CADES Proprietary Confidential
Bending Moment Distribution:
Moment is distributed in same ratio as that of the Shear force.
Bending Moment on Front Spar, M
FS = 0.459 * 781700
MFS
= 359159 N-mm
Bending Moment on Rear Spar, M
RS = 781700 - 359159
MRS
= 422541N-mm
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Front Spar
Rear Spar
SHEAR FORCE & BENDING MOMENT
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MATERIAL
Material : AA 2024-T6 Ultimate tensile strength, σ : 427 MPa Shear strength : 283MPa Density : 2.79 E-6 kg/mm3
Young's Modulus, E : 72400 Mpa Poisson's Ratio : 0.33
[Aluminum Association, Inc]. [7]
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Moment of Inertia:
I = M*y/σ Where, I = Moment of Inertia, in mm4
M = Bending Moment, in N-mmy = distance b/w neutral axis to top surface, in mmσ = Tensile strength, in MPa
Moment of Inertia on Front Spar, IFS
= 359159 * 52.8 / 427 I
FS = 44412 mm4
Moment of Inertia on Rear Spar, IRS
= 422541 * 43.44 / 427 I
RS = 42987 mm4
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MOMENT OF INERTIA
Front Spar
Rear Spar
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TORSION
Area of Torque Box, A1 = 30980.3 mm2
CG of Torque Box = 165 mm From Rear sparDistance Between CG & CP = 18.268 mmTorque, T = Load*d = 3190.65 * 18.268
T = 58286 N-mm
Shear flow, q1 = T/(2*A1) [2]
q1
= 58286 / (2 * 30980.3) q
1= 0.941 N/mm
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CG OF TORQUE BOX
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Torque
Shear Flow
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TORQUE DIAGRAM
ROOT 0
475 950 1425 1900 2375 2850 3325 3800 4275 TIP 4750
0
2000000
4000000
6000000
8000000
10000000
1200000011857039.54
8689789.08
6187429.48
4252608.34
2795550.39
1734041.9992888.78
506210.31212789.9958285.91 0
Torque diagram for the wing span
Wing span [root to tip] [mm]
Tor
que
[N-m
m]
27CADES Proprietary Confidential
Shear force (SF) on Front SparSF
FS = q * h
FS
SFFS
= 0.941*105.6 = 99.34 NTotal SF on FS = 1465.974+99.34 = 1565.313 N
On Rear SparSF
RS = q*h
RS
SFRS
= 0.941*86.88SF
RS = 81.729 N
Total SF on RS = 1724.676+81.729 = 1806.405 N
SHEAR FORCE DUE TO TORSION
28CADES Proprietary Confidential
SHEAR FORCE DUE TO TORSION
Front Spar
Rear Spar
29CADES Proprietary Confidential
TOTAL SHEAR FORCE
Front Spar
Rear Spar
30CADES Proprietary Confidential
WEB THICKNESS
Thickness of the Web can be calculated from the following formula,
ح shear strength
= SFFS
/ A web
Where,
ح shear strength
= Shear strength of the material AA 2024-T6 in MPa
A web
= Area of the web = (height * thickness) in mm
283 = 1565.313 / (105.602 * t web
)
t web = 0.052 mm
Area of the web = height * thickness = 105.602 * 0.052
A web
= 5.531 mm2
Moment of Inertia of Web:Moment of Inertia of a rectangular section web is given by,
I web
= t web * (hFS
)3 / 12I
web = 0.052 * (105.602)3 / 12I
web = 5140.175 mm4
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Front Spar
Rear Spar
WEB
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FLANGE
MOIflange
= MOIFront Spar
- MOIWeb
I flange
= IFS
- Iweb
= 44411 - 5140.175
I flange
= 39270.825 mm4
Also Moment of Inertia of the flange is given by,
I flange
= Aflange
* (yFS
)2
Where, Iflange
= Moment of Inertia of flange in mm4
yFS
= height from neutral axis to top surface of the flange in mm
Hence, Aflange
= Iflange
/ (yFS
)2
= 39270.825 / (52.801)2
Aflange
= 14.086 mm2
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FLANGE
Front Spar
Rear Spar
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MASS CALCULATIONS
AFS
= Aflange
+ Aweb
AFS
= 14.09 + 5.53 = 19.62 mm2
VFS
= AFS
* 475 = 19.62 * 475
VFS
= 9318.3 mm3
Mass = Density * Total Volume
= 2.78 E-6 * 4218551.12
Mass = 11.73 kg
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62 63 64 65 66 67 68 69 7011.5
12
12.5
13
13.5
14
12.5042471206507
12.6220245652181
12.7465877815849
12.8783982887878
13.0179753996257
13.1658829246678
13.3227251692486
13.4892202653604
13.6661526949271
12.3881208913364
12.4999845046666
12.6186071735573
12.7444287056342
12.8779461219943
13.0197004118034
13.1702724904167
13.3303551420369
13.5007061649515
12.274714489472612.3808256430427
12.4936607138559
12.6136374240938
12.7412301397201
12.876956676848
13.0213742183015
13.1751498097018
13.3390141564684
12.163512014196312.2640129538278
12.3711945626529
12.4854520552158
12.6072367354625
12.7370428555719
12.8754034876527
13.0229595119557
13.1804140972234
12.053900196057912.1489117223174
12.2505530055126
12.3591962678982
12.4752692884713
12.5992423142542
12.7316238829751
12.8730282376668
13.0241305238174
11.945183922182312.0348041336471
12.130996070132712.2341084011342
12.344544850195
12.4627511576641
12.5892108491091
12.7245110055869
12.8719649469388
11.836630462490411.9209354659232
12.011747601955712.1093913423994
12.2142457502486
12.3267314916976
12.4473065436803
12.5945353217656
12.7744354934928
11.727572120147911.806620573670211.8921053901977
11.984326119573712.0836364901705
12.1904314665371
12.339144756082
12.5065546355284
12.6799534484769
1819202122232425
Rear spar position in %
Mas
s [k
g]
MASS CALCULATIONS
Fro
nt
spar
po
siti
on
Hence, from the Calculations it is found that (25% - 62%) combination of Spar Position was found suitable. The Mass of this combination is 11.73 Kg which is least than any other combinations
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BUCKLING
To Check whether the web fails under shear buckling.
Condition: Shear stressinduced
< Buckling stress (safe design)
●The thickness calculation is based on iterations,
Finduced
= q / tweb
Fcritical
= k*E*(tweb
/ b)2
where, q = shear flow, in N/mm
E = Young's Modulus, in MPa
b = height of spar, in mm
tweb
= web thickness, in mm
k = shear buckling coefficient from graph
[4]
[4]
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BUCKLING CALCULATIONSITERATION 1. RIB SPACING FOR EQUAL DISTANCE OF 475mm
Web thickness's of front spar at section 1 is as follows,F
induced = q
1 / t
web------------ (1)
= 0.941 / 0.052 F
induced = 18.09 N/mm2
Fallowable
= K * E * (t web
/ b)2-----------(2)18.09 = 5 * 72400 * (t
web / 105.602)2
The value calculated for tweb
is re substituted in Eqn.(1) and this loop will continue till we get equal consecutive thickness.
Hence, the thickness of the web is 0.30 mm at section 1. Same calculations were repeated for all sections of front spar to optimize the web thickness
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Front Spar
Rear Spar
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MASS CALCULATION
●Web design is safe under buckling.
●From buckling calculation the total mass of the spars is 16.14 kg.
●By this, mass of the spars got increased by 4.41 kg.
●To decrease the mass, one more iteration has been carried out.
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ITERATION-2● For optimum Rib spacing, (a/b) ratio >= 1
Rib no. Rib dist. From root Spar heights (a/b) ratioK from graph
Web thickness Web volumeFS RS FS RS FS RS FS RS
[mm] [mm] [mm] [mm] [mm]0 TIP4750 64.49 54.05 0 0.00 - - - -1 4440 81.97 68.4 3.78 4.53 5.10 5.00 0.22 0.2 5590.56 4240.82 4110 100.58 83.67 3.28 3.94 5.17 5.08 0.34 0.3 11285.3 8283.533 3780 119.19 98.94 2.77 3.34 5.30 5.15 0.44 0.39 17306.53 12734.094 3450 137.8 114.22 2.39 2.89 5.50 5.20 0.53 0.48 24101.05 18091.815 3120 156.41 129.49 2.11 2.55 5.75 5.40 0.64 0.57 33033.37 24356.696 2790 175.02 144.76 1.89 2.28 6.00 5.60 0.74 0.67 42739.15 32006.447 2470 193.06 159.57 1.66 2.01 6.30 5.80 0.84 0.76 51894.8 38807.188 2150 211.11 174.38 1.52 1.84 6.55 6.20 0.94 0.84 63500.68 46872.819 1830 229.15 189.19 1.40 1.69 6.90 6.25 1.03 0.94 75528.17 56907.75
10 1520 246.63 203.53 1.26 1.52 7.25 6.55 1.12 1.02 85630.63 64357.4511 1210 264.11 217.88 1.17 1.42 7.60 6.80 1.2 1.1 98250.04 74297.4212 900 281.59 232.23 1.10 1.33 7.80 7.00 1.31 1.19 114355.32 85668.5413 600 298.51 246.11 1.00 1.22 8.20 7.35 1.39 1.26 124479.09 93029.9614 300 315.43 259.99 1.05 1.15 8.00 7.60 1.51 1.35 142888.88 105297.5715 Root 0 332.35 273.88 1.11 1.10 7.80 7.80 1.63 1.43 162519.15 117494.52
Total volume 1053102.72 782446.56Web volume 1835549.28
[mm3] [mm3]FS RS
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WEIGHT CALCULATION
●Finally mass of the spars reduced by 0.89 kg when compared to 1st iteration.
● These dimensions are taken for modelling
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RESULTS AND DISCUSSION
Root 0 300 600 900 1210 1520 1830 2150 2470 2790 3120 3450 3780 4110 4440 47500.00
0.40
0.80
1.20
1.60
2.00
0.386 0.362 0.34 0.318 0.291 0.27 0.245 0.22 0.198 0.17 0.145 0.12 0.095 0.07 0.042
1.63
1.51
1.391.31
1.21.12
1.030.94
0.840.74
0.64
0.530.44
0.34
0.22
WEB THICKNESS FOR FRONT SPAR
ACTUALFROM BUCKLING
FROM ROOT TO TIP [mm]
TH
ICK
NE
SS
OF
WE
B[m
m]
43CADES Proprietary Confidential
Root 0 300 600 900 1210 1520 1830 2150 2470 2790 3120 3450 3780 4110 4440 47500.00
0.40
0.80
1.20
1.60
2.00
0.517 0.488 0.455 0.425 0.39 0.361 0.327 0.295 0.261 0.23 0.195 0.16 0.129 0.095 0.06
1.431.35
1.261.19
1.11.02
0.940.84
0.760.67
0.570.48
0.390.3
0.2
WEB THICKNESS FOR REAR SPAR
ACTUALFROM BUCKLING
FROM ROOT TO TIP
TH
ICK
NE
SS
OF
WE
B[m
m]
44CADES Proprietary Confidential
CONCLUSION
● Front Spar positioning is estimated to 25% and Rear Spar to 62% of the
Chord Length.
● Flange and web dimensions are calculated and suitable changes in
dimensions are incorporated from manufacturing point of view.
● Number of Ribs and their positioning for the prevention of bending and
buckling of Spars is calculated.
● Mass of the spars calculated from iterations is 15.25 kg.
● The Detail drawings for the front and rear spars are provided using CATIA V5.
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SCOPE FOR FURTHER WORK
● Spar position can be optimized based on buckling calculations.
● Further optimization of Rib is possible.--Varying number of Ribs and spacing of Ribs.
● Use of other materials for the design of spars can be thought of.
● Detail stress analysis of individual components and its validation with calculations can be carried out.
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CAD
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Taking values from NACA Standards
At Root: Profile: NACA 64A1215.
Leading Edge radius = 1.556% c.
Slope of mean line at leading edge = 0.0842.
At Tip: Profile: NACA 64A1210.
Leading Edge radius = 0.701% c.
Slope of mean line at leading edge = 0.0842.
CAD MODELING OF THE WING SPAR
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1. Generation of the profiles at the root and tip using the NACA profiles.
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● INCORPORATING THE LEADING EDGE RADIUS AS SPECIFIED IN THE PROFILE STANDARD.
1.Giving the slope in the sketcher mode
2.Creating the arc of the required dimension coming out of sketcher.
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● Using the connect curve option to join the leading edge radius and the aerofoil profile.
● Create the surface using multi section surface option.
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INTERSECTION OF THE PROFILES
● Creating the planes at the four sections at ½, ¼, ¾ of the span of the wing.● Intersecting the lofted surface on the planes creating unique sketches on them.
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ANGLE OF ATTACK
● Create a point at the quarter chord and draw a line for reference.
● Rotate the intersected profiles as 0.60 at the quarter, 1.10 at mid span, 1.60 at three
fourths and 20 at the tip.
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By considering the profiles generated with angle of attack at different sections,
the wing surface is created using multi-section surface option.
CREATE THE SURFACE USING MULTI SECTION SURFACE OPTION
Thus the surface is created as per the requirements incorporating all the necessary data.
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CREATION OF REFERENCE AEROFOIL SECTIONS
● 15 planes are created at rib positions along the wing span.● The intersections created are used as the reference for the creation of the spar.
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CONSIDERATIONS MADE DURING THE DESIGN OF SPAR
ELEMENTS
●The maintenance of the nose box is made easy.
● The front spar is I – section.
● The rear spar is C – section.
● Minimum distance required for a single row riveting is kept as 15 mm.
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DESIGNING OF SPAR ON MANUFACTURING BASIS
➢ The front spar is placed at 25% of chord length from leading edge.
➢ The rear spar is placed at 62% of chord length from leading edge.
➢ Thicknesses of the flanges and webs are different.
➢ The flanges are made of T-sections and L- sections.
➢ The webs are made with sheet metal.
➢ The thicknesses are optimized based on the availability of the standard gages of sheet metal.
➢ The final assembly of elements can be fastened with rivets.
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CROSS SECTION SPAR
Skin area, As = (b +2*20*t
s) mm2
Effective flange area = (Af- As)/2
Web thickness is altered as per the availability of sheet metal gages.
where , b= flange width in mm ts =skin thickness in mm Af =designed flange area in mm2
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Rib no. Dist. From root Flange Width Skin Area Available area Flange Thickness Effective Flange area Web thicknessFrom root (mm) (mm) (mm) (mm)
1 Root 0 70 220 266.94 3.81 266.94 1.632 300 70 220 215 3.07 215 1.633 600 70 220 175 2.5 175 1.634 900 65 210 150 2.31 150 1.295 1210 65 155.2 144.9 2.23 144.9 1.296 1520 60 147.2 116.4 2 120 1.297 1830 60 147.2 86.4 2 120 1.298 2150 55 139.2 60.4 2 110 0.919 2470 50 88.8 55.6 2 100 0.91
10 2790 45 82.8 36.1 2 90 0.9111 3120 40 76.8 16.6 2 80 0.6412 3450 35 70.8 -0.4 2 70 0.6413 3780 30 64.8 -12.4 2 60 0.6414 4110 30 64.8 -19.9 2 60 0.6415 4440 30 64.8 -24.9 2 60 0.6416 NO RIB 4750
(mm2) (mm2) (mm2)
FRONT SPAR DIMENSIONS
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REAR SPAR DIMENSIONS
Rib no. Dist. From root Flange Width Skin Area Available area Flange Thickness Effective Flange area Web thicknessFrom root (mm) (mm) (mm) (mm)
1 Root 0 90 260 415.01 4.61 415.01 1.452 300 90 260 340 3.78 340 1.453 600 80 240 295 3.69 295 1.454 900 75 230 257.5 3.43 257.5 1.155 1210 70 204 223 3.19 223 1.156 1520 65 194 180.5 2.78 180.5 1.157 1830 60 184 138 2.3 138 1.158 2150 55 174 100.5 2 110 0.919 2470 50 148 73.5 2 100 0.91
10 2790 45 138 48.5 2 90 0.9111 3120 40 128 21 2 80 0.6412 3450 35 118 4 2 70 0.6413 3780 30 108 -19 2 60 0.6414 4110 30 108 -36.5 2 60 0.6415 4440 30 108 -43 2 60 0.6416 NO RIB 4750
(mm2) (mm2) (mm2)
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CREATION OF THE SPAR SECTIONS
1. Two T sections for the flange, and web section for the front spar.2. Two L sections for the flange, and web section for the rear spar.
FRONT SPAR REAR SPAR FULL PROFILE
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GENERATING SPAR USINGDIFFERENT SECTIONS
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CRIMP HOLES OR LIGHTENING HOLES
The lightening holes are made in the element in order to reduce the weight of the element. the crimp holes are made to the web element of the spar. These holes provided in between the two successive rib locations.
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SPAR WITH LIGHTENING HOLES
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REPRESENTATION OF RIVET HOLES
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FINAL SPAR ASSEMBLY
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BIBLIOGRAPHY
1] Abbot & Albert,'Theory of wing sections',Dover publication,1949.
2] David J. Perry,'Aircraft structures',Mc-Graw Hill publication,1950.
3] E. F. Bruhn,'Analysis and design of flight vehicle structures',1973.
4] Michael C. Y. Niu, 'Airframe Stress Analysis and Sizing', 2001.
5] Michael C. Y. Niu, 'Airframe structural design', Conmilit press Ltd., 1989.
6] Kuethe and Schetzer, 'Foundations of Aerodynamics', 2nd Edition, John Wiley
and Sons, New York, 1959.
7] ASM Material Data Sheet
8] MIL Handbook.
&CADES Library.
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THANK YOU
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