PROJECT REPORT
ON
WING STRUCTURE ANALYSIS OF AIRBUS A300
BASP 001
SUBMITTED IN THE PARTIAL FULFILLMENT OF THE DEGREE
OF BACHELOR OF TECHNOLOGY IN AEROSPACE
ENGINEERING
Mr. Kamal Kumar
Enrolment Number: 118526393
Year of Submission: October-2014
Indira Gandhi National Open
University, New Delhi
Hindustan Aviation Academy
Bangalore
1
Project Report – BASP 001
P.P. No. 6026
WING STRUCTURE ANALYSIS OF AIRBUS A300
Submitted as a partial fulfillment of requirements of the bachelor degree in
Aerospace Engineering
Under IGNOU
Submitted by
Kamal Kumar
Enrollment No.:- 118526393
Under the supervision of
Mr. Prashant Sinha
October- 2014
October- 2014
Hindustan Aviation Academy
(AEP 004)
Chinnapanahalli, Marathahalli, Bangalore
2
CERTIFICATE OF ORIGINALITY
I hereby declare that the project work titled “WING STRUCTURE ANALYSIS OF
AIRBUS A300” as my own work and to the best of my knowledge. It contains no
materials previously published or written by another person, nor material which to a
substantial extent has been accepted for the award of any degree or diploma at
IGNOU or any other educational institute, except where due acknowledgment, is
made in the project. Any contribution made to the research by others, with whom I
have worked or discussed, is explicitly acknowledged.
I also declare that the intellectual content of this project work is the product of my
own work, except to the extent that assistance from others in the project’s design
and conception or in style, presentation and linguistic is acknowledged. I also verified
the originality of contents through plagiarism software.
---------------------------------------Kamal Kumar
Enrollment No. 118526393
This is to certify that that the above statement made by the candidate is correct to
the best of my knowledge and belief.
_________________________________
Mr. PRASHANT SINHA
(HOD, Aeronautical Engineering,
HEA Polytechnic)
3
Every project big or small is successful largely due to the effort of a number of wonderful
people who have always given their valuable advice or lent a helping hand. I sincerely
appreciate the inspiration; support and guidance of all those people who have been
instrumental in making this project a success.
At this juncture I feel deeply honored in expressing my sincere thanks to my project
guide Mr. PRASHANT SINHA for making the resources and information available at right
time and injecting me with the knowledge of required software leading to the successful
completion of my project.
I also express my gratitude to my honorable Principal Mr. Sajjad Rahim and Head of
Department Mr. Ramakrishna Reddy who were always faithful to me in their academic
guidance, fairness and responsiveness to all of my project queries..
I would also like to thank all the faculty members of Hindustan Aviation Academy for their
critical advice and guidance without which this project would not have been possible.
Last but not the least I place a deep sense of gratitude to my family members and my friends
who have been constant source of inspiration during the preparation of this project work.
4
Abstract
In this project Structural and modal analysis of A300 wing structure, we aim to learn the
process to solve many engineering problems with the help of a solver commonly known as
SPARSE direct solver which is the default solver in ANSYS without preparing the prototype
model and caring the actual experiments.
The methods used in solving any problem in ANSYS vary from person to person. One may
take assumptions to solve the problem with a unique approach towards the problem. So the
results obtained may vary from person to person. The results obtained by ANSYS software
are just approximate results which accounts for various conditions which cannot be
considered in analytical method and cannot give 100% accurate results as experimental
values are obtained.
In this project we aim at developing a CAD model of A300 (Airbus-300) using the modeling
software CATIA V5 R18. The main purpose of the project shall be to determine the structural
parameters such as total deformation, equivalent stresses which is also known as Von-Mises
stress, shear stress, shear intensity on the skin of the aircraft wing which has a thickness of
10mm. The modal analysis will be carried out to find out the first 6 modes of vibrations and
the different mode shape in which wing can deform without the application of load. The
outcomes and shortcomings if any will be analyzed and suitable mitigation measures will be
presented.
5
Index
CONTENT
Page No.
1. Title of the Project…………………………………………..8
2. Objective of the study ……………………………………….8
3. Methodology of the study…………………………………...9
4. Statement of the problem……………………………………11
5. Input…………………………………………………………12
6. Description…………………………………………………..14
7. Airfoil Generation……………………………………………22
8. Flow Analysis over Airfoil in Virtual Wind tunnel…………23
9. Pre-Design Calculation………………………………………28
10. CAD Modelling of Wing…………………………………….31
11. Structural Analysis…………………………………………..36
a. Problem Specifications
b. Pre-Analysis and start up
c. Model
d. Meshing
e. Solution
f. Verification and Validation
12. Model Analysis………………………………………………50
a. Problem Specification
b. Pre-Analysis and Start up
c. Mesh
d. Solution
e. Verification and Validation
6
13. Final Result……………………………………………….66
a. Flow Analysis Result
b. Structural Analysis result
c. Model Analysis Result
14. Conclusion………………………………………………..68
15. Scope of Further Studies…………………………………69
16. Bibliography………………………………………………70
17. Bio-data of the Student……………………………………71
7
1. TITLE OF THE PROJECT
WING STRUCTURE ANALYSIS OF AIRBUS A300
2. OBJECTIVE OF THE STUDY
 Study about the A300 wing design and the airfoil used in it.
 To study various conditions, which may be applied to get the most appropriate result
 To create the CAD model of wing structure with the limited dimensional information
available (company does not provide full and accurate dimensions)
 Investigate structural behavior of a wing during take-off and climbing phase
8
3. METHODOLOGY OF THE STUDY
To do such a project we need enough information about the wing dimensions. So collecting
details of A300 is the first step. All the required design parameters are not available but with
the help of the airplane characteristics manual provided by the manufacture for the airport
planning will be used as it provides the external features of the aircraft. The internal structure
of the wing will be assumed and it will be simplified to get approximate result on the skin of
the aircraft wing.
For analyzing in ANSYS, it is important that we make a model of wing in CATIA V5 R18.
For making this model dimensions are necessary. After making surface model of wing in
CATIA this work will be saved and import to ANSYS.
By using ANSYS software we can calculate and analyze structural loads on wings during
take-off and climbing phase. Because of highly classified information we can only take
approximate dimensions of wing. So we cannot say the result will be highly accurate.
a) Airfoil plotting and analyzing flow properties
b) CAD modelling
The CAD model of a wing is established by using CATIA. The full model consists of
ribs, front and aft spars and the skin. The various design parameters are taken directly
from the airplane characteristics manual and for internal structure suitable
assumptions and simplifications will be done.
c) Model experimental and model validation
The validation and updating of preliminary/sub-structural model are essential to
9
assure the accuracy of full/system model. The general procedure of model validation
and model updating is to develop a simple FE model at the beginning of the process to
stimulate the behavior of the system. The preliminary result is used to define the test
conditions by optimizing and refine the mesh size. The updating process, including
model correlation and model updating, uses the data obtained from the model
experiment to refine the FE model. Finally, the updated model is expected to
represent the behavior of the structure in a more accurate way. This validated FE
model can be used later for the various types of different analysis subjected to thrust
and aerodynamic loads
Airbus A300 dimensional
data collection
NACA 64-215 airfoil plotting
Flow analysis over obtained
airfoil
Preparing CAD model of Wing
Importing CAD model to ANSYS
and solving it
Redefine the MESH
Obtaining the result
Conclusion
10
4. STATEMENT OF THE PROBLEM
The aircraft is a highly sophisticated and scientific machine, which requires the expertise of
hundreds of engineers and scientists working in harmony to design and produce a successful
product.
One of the major design factors in the development of any new aircraft is to ensure that the
wing should be able to take the load of whole aircraft when in flight. The wing structure
should also be of such a parameter that the air flow over it should provide maximum support
in its flight and stability. The wing structure experience various types of loads during each
phase of the flight which includes take-off, climb, cruise, loiter, landing, touch-down. In each
segment there is a variation in load factor which induces various types of stresses in various
components of aircraft body.
Hence, here we are interested to find out the various types of stresses and its intensity
induced in the skin of the aircraft during take-off. We are also interested to find the first six
modes of vibration which are possible when the aircraft is on ground.
This problem can be simplified by considering it as a cantilever beam whose one end is fixed
in the fuselage and the tip end is free. The loading condition on a wing is equivalent to the
uniform varying load throughout the wing. The wing design here we use is as per the
technical data issued by AIRBUS team in May 2014 revised publication.
11
5. INPUT
All the data applicable to this analysis will be taken from the technical publications of
AIRBUS with reference to A300.
Figure 1: Important dimensions of AIRBUS A300
Some required data which are necessary in the process of creating the model were missing,
which has been calculated geometrically and has been taken forward with some realistic
assumptions. The loading condition is found using the maximum take-off weight and
maximum climb angle which is allowed for this aircraft from any airport.
12
Some of the required data which will be used in the CATIA modelling of wing structure are
as shown below:
13
Aircraft Type Model
A 300B4-600R
Airfoil
NACA 64-215
Wing Area (meter square)
260
Wing span (m)
44.84
MAC (m)
6.44
Aspect Ratio
7.73
Taper Ratio
0.3
Average thickness (t/c %)
10.5
¼ Chord sweep angle ()
28
Rib Thickness
100mm
6. DESCRIPTION
AIRCRAFT is a weight carrying structure that can travel through the air supported by its
buoyancy and by the dynamic action of the air against its surfaces. The method by which the
aircraft flies is based on BERNOULLI’SPRINCIPLE; Bernoulli’s principle applies to any
fluid, and since air is a fluid, it applies to air.
The camber of an airfoil causes an increase in the velocity of the air passing over the airfoil.
This results in a decrease in the pressure in the stream of air moving over the airfoil. This
decrease in pressure on the top of the airfoil causes lift.
Aircrafts are kept in the air and can fly because of flight produced by the wing. If an aircraft
is flying straight at a level with constant speed the lift produced by the wing must be equal to
the weight of the aircraft. If the lift exceeds the weight, the aircraft will be pushed higher (i.e.
pulled up) and if the weight exceeds the lift the aircraft will sink (pulled down).
As speed is constant the lift is maintained at the correct amount by keeping the angle of the
wing to the airflow to the correct amount. In the conventional aircraft this is achieved by the
action of tail plane.
The airplane structure consists of a skeleton framework with a covering to give its shape. The
various parts of the aircraft structure are as shown below in figure:
14
Figure: Aircraft Parts
Figure: Aircraft parts
15
FORCES ACTING ON AN AIRPLANE
An airplane in flight is the center of a continuous tug of war between four forces: lift, weight,
thrust, and drag.
Lift and Drag are considered aerodynamic forces because they exist due to the movement of
the aircraft through the air. The weight pulls down on the plane opposing the lift created by
air flowing over the wing. Thrust is generated by the propeller and opposes drag caused by
air resistance to the frontal area of the airplane.
During take-off, thrust must overcome drag and lift must overcome the weight before the
airplane can become airborne. In level flight at constant speed, thrust exactly equals drag and
lift exactly equals the weight or gravity force. For landing, thrust must be reduced below the
level of drag and lift below the level of the gravity force or weight.
Figure: Four forces acting on an airplane
16
Aircraft structure
Airframe Components: Almost any airframe may be split into four main components:

The fuselage or body

The tail unit (or fore-planes for a canard-type aircraft)

Wing

Mountings for all other systems (undercarriage, engines, etc.)
As my project is completely focusing towards the structure of wing and its basic unit i.e.
airfoil, our discussion will move straightly towards wing and its unit structure.
Wing
As Jack Moran (Network Engineer, The Boeing Company) says, wings are a thrust
amplifier. Their magic is in their ability to defy gravity without using raw thrust from a fuelguzzling rocket. Instead, wings use the air flowing past to create a vertical force called lift
which seems to defy gravity.
Hence, the function of the aircraft wing is to
provide lift to the aircraft when it moves in air
(i.e. to provide flight to the aircraft). The cross
section of the wing is in form of an airfoil and
these sections provide the lift to the aircraft on
the basis of lift and drag forces.
Figure: Cross section of wing
The wing must generate lift from the airflow over it to support the aircraft in flight. The
amount of lift required depends on how the aircraft is flying or maneuvering.
17

To take off and climb, the required lift must be developed at a low airspeed.
If the aircraft is to fly in very tight turns, the wing must produce lift equal to perhaps
eight times the aircraft weight.

For straight and level flight, the total lift produced must be equal to the weight of the
aircraft.

For landing, the slowest possible forward speed is required, and enough lift must be
produced to support the aircraft at these low speeds.
For take-off and landing, lift-augmenting devices like flaps, leading-edge slats, etc. are
normally added.
The wing needs to be stiff and strong to resist high lift forces, and the drag forces associated
with them. So it could be argued that the wing is the most essential component of an
airframe.
The three primary loads that acts on the aircraft wing are aerodynamic lift, load due to wing
structure weight, and load due to the weight of the fuel contained in the wing. These loads act
perpendicular to the wing surface, and their magnitude varies along the length of the wing as
shown in figures below:
Figure: loads on aircraft wing due to aerodynamic lift
18
Figure: loads on aircraft wing due to wing structure weight
Figure: loads on aircraft wing due to weight of the fuel contained in the wing
Airfoil
An airfoil is a 2D (outline) shape of the wing .The ultimate objective is to obtain the lift
necessary to keep airplane in the air .The desirable shape of airfoil is round in front and sharp
at the tip for a subsonic aircraft.
19
The following six terms are essential in determining the shape of the typical airfoil:

The leading edge

The trailing edge

The chord line

The camber line

The upper surface

The lower surface
The first step in designing any wing is to plot an airfoil from which the rib pattern is created
.It is generally wise to place the main spars at the thickest section of the airfoil.
The desired length of the airfoil section is determined by placing the leading and the trailing
edge with desired distance apart. The chord line is drawn connecting these two points
together. The camber line determines the amount of curvature. This curvature greatly aids an
airfoil section’s lifting abilities, a thickness function is “wrapped” about and below the
camber line; this thickness determines the upper and lower surfaces. Finally airfoil shape with
all aerodynamic characteristics can be obtained.
Parametric Description of AIRFOIL
The NACA airfoils are named over their parametric specifications and it also gives some
basic physical properties related to it, like boundary layer formation.
20
The NACA 6-Series numbers have many variations and it can better explained with the help
of an example, as below:
NACA 64-215
1. NACA 64-215 : The first digit designates that this is a "Six Series" type of airfoil.
2. NACA 64-215 : This series produces laminar boundary layers. The second digit
designates how far back behind the leading edge in tenths that the airfoil will maintain
a laminar boundary layer. Older airfoils were lucky to reach 15%. These airfoils are
designed for much longer regions of laminar flow. In this example, the laminar
boundary layer will stretch back 40% from the leading edge.
3. NACA 64-215 : The "dash" is sometimes replaced with an A. its importance can be
best understood on looking at the razor-sharp trailing edge of the NACA 6-Series
airfoils. They were so thin and brittle that designers requested the series be modified
to have an actual thickness at the trailing edge. The result is the "A" modification
where the back slopes of the airfoil are straight and the trailing edge is thick.
4. NACA 64-215 : The third digit designates the design Lift Coefficient for this airfoil.
It typically references the Lift Coefficient found in the center of the drag bucket. A
value of zero means that the airfoil is symmetric.
5. NACA 64-215 : As with all of the NACA airfoils, the last two digits combine
together to designate how thick the airfoil is as a percentage of chord length. In this
example, our airfoil will be 15% of the chord length thick. With this class of airfoils,
the maximum thickness occurs at a distance of 30% from the leading edge.
21
7. AIRFOIL GENERATION
Open designfoil workshop
Go to airfoil > NACA Airfoil > NACA 6 series > airfoil generator window opens, as shown
in the figure
Set the following Data:

Minimum pressure location = 40

Design lift coefficient = 0.2

Maximum airfoil thickness as percentage of chord = 15

Mean line parameter (a) or (-) =1.0
Next, after setting the above data in “Designfoil software”, Click on the “modified 6A Series”
and then go to “Generate / Restore View”
We get a plot of 71 points, constituting airfoil “NACA 64-215”, as shown below in figure:
22
8. FLOW ANALYSIS OVER AIRFOIL IN VIRTUAL WINDTUNNEL
Problem specification:
In flow analysis, we are interested to find out the flow property over the obtained airfoil to
check if its property are good enough to support the flight of the suggested aircraft i.e. airbus
A300. We will be checking the drag polar plot for the airfoil.
23
Steps :Go to Tools analysis and open Virtual Wind Tunnel >
On the left side, a window is available “Boundary Layer / Mach Parameters”
Here, we will be testing the lift and drag conditions as
per sea level conditions.
Hence to update the Reynolds number, click on
to open a new pop up window “Standard Atmosphere
Model”, where we should provide the following intake conditions,
Altitude = o feet
Reference Length =2.72feet
Flight Speed = 273.99 feet per second
24
Click on to
Click on
25
get following data
Now, on the right side of the window, we can see
Next we click on
to set the boundary layer prediction criterion;

Transition
Criterion
is
Schlicting

Transition to be natural.

Drag polar angle range(17 to
-10)

Prediction on for CL max using
“Proprietary Dreese Method”
Hence, Re = 4739873
Now, click
26
,
we get following data
AOA
Cl
-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
27
Cd
-0.984
-0.865
-0.747
-0.627
-0.508
-0.389
-0.269
-0.149
-0.029
0.091
0.211
0.331
0.451
0.571
0.69
0.81
0.929
1.037
1.13
1.211
1.277
1.331
1.372
1.4
1.415
1.417
1.406
Cm
0.0111
0.0101
0.0091
0.0082
0.0077
0.0072
0.0067
0.0063
0.0056
0.0051
0.0049
0.0052
0.0057
0.0067
0.0075
0.008
0.0089
0.0095
0.0103
0.0113
0.013
0.0145
0.016
0.0178
0.0198
0.0219
0.0243
-0.041
-0.041
-0.041
-0.042
-0.042
-0.043
-0.043
-0.044
-0.044
-0.045
-0.045
-0.046
-0.046
-0.046
-0.047
-0.047
-0.048
-0.048
-0.049
-0.049
-0.05
-0.05
-0.051
-0.051
-0.051
-0.052
-0.052
9. Pre-designing Calculation for wing
As the mid wing span is 22.42 m we divide total wing in 23 sections each placed at a certain
distance (derived from its geometry) from the reference airfoil. The distance between two
airfoils is 1m. Also, the diameter of the fuselage is 5.64 m and some part of our wing will be
inside fuselage and which is completely rigid due to its wing box design as shown in figure 1.
The section which is completely rigid is 2.82 m.
From the section placed at a distance of 2.82 m from the reference plane, the airfoil shape is
scaled appropriately to get the desired wing profile. While after that 21 more planes are to be
plotted, which in total makes the wing span of 23m in total 23 planes including reference
plane, as shown in figure below.
The above sketch is the conceptual sketch of a wing which will be created with the help of
basic geometry and trigonometric relations in CAD software.
28
Calculations of the required values:
The formula for calculating the distance of leading edge point whose co-ordinate is (0,0)
from reference line 1 using similarities of the triangle concept is given by,
Where,
Y = distance of a point on leading edge from the reference line 1
a =distance of the section from the root chord
From the drafting we came to know that the trailing edge makes an angle of 20.1035° with
the reference line 2.
So distance of trailing edge point whose co-ordinate is (0, 0) from the reference line 2 is
given by the formula,
Z = b.tan (20.1035)
Where,
Z = distance of a point on a trailing edge from the reference line 2.
b = distance of a section from the tip chord up to section 9.
Calculation of the local chord length can be done using the formula,
c = 11.54 + 2.75 - Y- Z
Calculation of local taper ratio is given by
29
The following values are found with the help of geometry and trigonometry relations:
Local chord length
Section no.
Distance of leading
edge point from top
Distance of trailing
edge point from bottom
reference line i.e. line 2
X (m)
4.89
Local taper ratio
C (m)
1
9.4
1
reference line i.e. line 1
Y (m)
0
2
9.4
1
0
4.89
3
9.2941
0.9887
0.1059
4.89
4
8.7053
0.9260
0.6947
4.89
5
8.1165
0.8634
1.2835
4.89
6
7.5277
0.8008
1.8723
4.89
7
6.939
0.7381
2.4610
4.89
8
6.3502
0.6755
3.0498
4.89
9
5.7614
0.6129
3.6386
4.89
10
5.5167
0.5868
4.2274
4.5459
11
5.294
0.5631
4.8161
4.1799
12
5.0712
0.5394
5.4049
3.8139
13
4.8485
0..5157
5.9937
3.4478
14
4.6257
0.4920
6.5825
3.0818
15
4.403
0.4684
7.1712
2.7158
16
4.1802
0.4447
7.7600
2.3498
17
3.9574
0.421
8.3488
1.9838
18
3.7347
0.3973
8.9376
1.6177
19
3.512
0.3736
9.5263
1.2517
20
3.2892
0.3499
10.1151
0.8857
21
3.2892
0.3499
10.1151
0.5197
22
3.0664
0.3262
10.7039
0.1537
23
2.8436
0.30251
11.2927
0.0534
24
2.75
0.2925
11.54
0
30
10.CAD- MODELLING OF WING
Import the co-ordinate of airfoil section into CATIA V5 having chord length of 2.72m
As we are considering that wing is designed with only one airfoil throughout, it has to be
scaled down accordingly to get the required shape of a wing profile as per calculation shown
above.
31
Projection of plane is as shown,
Next, the airfoil will be projected and re-scaled as per geometry
32
Finally, we obtain
Now, each section is padded 50 mm mirror extended so that the airfoil section is converted
into the rib section with a thickness of 100 mm each
33
Next, join all the airfoil section by multi-section surface definition
We obtain,
The spars and holes are being created in the wing design,
34
So, final design with the spars and holes in wing structure is
Before importing the .CAT file to the ANSYS workbench, the file has to be converted into
.stp format.
This conversion can be done by going to file option>save as> save as type: .stp format>save.
There may be some data loss during conversion and importing process resulting in
approximate results in ANSYS workbench.
35
11. Structural Analysis
Problem specification:
In static structural analysis we are interested in the total deformation, Von Misses stress
which is also known as equivalent stress, shear stress and stress intensity induced in the skin
structure of the wing.
Pre-Analysis and Start-Up :
Open ANSYS Workbench
We are ready to do a simulation in ANSYS Workbench. Open ANSYS Workbench by going
to Start > ANSYS > Workbench. This will open the start-up screen seen as shown:
To begin, we need to tell ANSYS
what kind of simulation we are doing.
On the left of the start-up window, we
can see the Toolbox Window. Take a
look through the different selections.
As we are only doing a force loading,
we will be doing a Static Structural
simulation.
Load the Static Structural tool box by dragging and dropping it into the Project Schematic.
Name the Project Wing structure by doubling clicking {Static Structural (ANSYS)}}
underneath the project schematic.
36
Engineering Data
Double click on the engineering data, design modeler will open and a pop up window will
ask for the default unit of measurement for the geometry. Select meter and press OK.
Geometry
Go to the Workbench in the Project Schematic window, double click on Geometry to open
Design modeler.
In Design Modeler, click on File >Import External Geometry File
Select the saved .stp file of the designed CAD model and then click on Generate.
Go to detail window,
Select Operation > Add frozen
The imported geometry should show in the Graphics window as below,
37
Conversion of 1 part from 2 part
Notice in the Outline window that the part has been imported as 2 separate parts and 2
bodies.
We need to convert 2 separate surfaces into 1 part surfaces. For this we need to select both
the Surface Bodies and right click on one of the Surface Bodies and select Form New Part.
Now, in the Outline window, we can see 1 Part
and 2 Surface Bodies as subordinates.
Here we can rename the surfaces as Outer
Surface and the Spar.
38
Now that the geometry is imported and generated, we are ready to mesh the geometry.
MODEL
Open ANSYS Mechanical by double clicking
.
When ANSYS Mechanical opens, notice that there is a question mark next to Geometry in
the Project Outline - this means that there is something missing in this section. Expand
Geometry > expand Part > select Outer Surface.
Notice that Thickness is highlighted as it does
not have a value specified.
We will specify a thickness so the geometry will mesh correctly. For the Outer Surface, enter
1e-2 (10mm) next to Thickness. Repeat with the value of 3e-2 (30mm) for Spar to thickness.
Once the value is given, question mark is
removed by default.
39
CONNECTIONS
Go to Connections > Contacts >Delete
ANSYS may create connections automatically - however they are not required for this
simulation and will cause problems when meshing. Expand Connections and delete the folder
titled Contacts by right clicking and selecting Delete.
MESH
Body Sizing
For this geometry, we will be using a body sizing. Click on Mesh in the Project Outline
window to open up the Meshing Menu in the menu bar. To create a new sizing, go to Mesh
Control > Sizing. Next, we need to select the geometry that the sizing will affect. We want to
select the entire geometry.
Mapped Face Meshing
To apply a mapped face meshing, first click on Mesh in the Outline window. This will bring
up the Meshing Menu Bar at the top of the screen. Next, select Mesh Control > Mapped Face
Meshing. Select the 2 faces of the mesh by holding down the left mouse button and dragging
over the entire geometry. In the Details window, click Geometry > Apply - it should say 2
faces are selected.
Edge Sizing
In the Meshing Menu, click Meshing Control > Sizing. Click the edge selection filter
.
Select the 4 curved edges on the outside of the geometry that make up the shape of the
NACA 64215 Airfoil as the picture shows:
40
In the details window,
Select Geometry > Apply,
Select Type > Number of Divisions > Change the Number of Divisions to 20.
Change Behavior > Hard.
Next, create another Edge Sizing, and this time, select the 2 edges at the very front and very
back of the airfoil that run along the wingspan, as the picture shows:
Again, in the Details window Select Geometry > Apply,
Select Type > Number of Divisions > Change the Number of Divisions to 40.
Change Behavior > Hard.
Generate the mesh by selecting Mesh > Generate Mesh
41
Setup
Fixed Support
Next, we will apply the boundary conditions to the geometry. In the
graphics window, click the positive Z-Axis on the compass to look
at one side of the airfoil.
In the Outline window, select static structural to bring up the Setup Menu. In the Setup Menu,
select Supports > Fixed Supports. Make sure the Edge Selection Filter is selected, Ctrl+ left
mouse click to select the upper and lower edges of the airfoil you are looking at. In the details
window, select Geometry > Apply.
42
Pressure Load
We want to apply an upward force on the wing, which can be in terms of pressure load and
can be determined by calculating the load factor.
Load factor (n) can be found from the maximum climb angle for A300 from any airport is
17°.
i.e.
The maximum take-off weight of A300-600 R is around 170,000 kg.
From the basic aerodynamics,
Lift force = load factor * weight of an aircraft.
As we are interested to calculate the structural parameters during take-off and climbing
phase, lift must be greater than weight of an aircraft.
Thus the total lift force required to climb through 17°, the aircraft should be able to generate
the lift force 1751.531 KN. This is the total lift which has to be generated by the sets of its
wing. Thus the force developed by each wing is 875.765 KN. This force is converted into the
pressure load, which is in the form of uniformly distributed load by dividing this force by the
2
semi wing area of 130 m .
Therefore, the total pressure load applied from the bottom of the surface is 6736.65 Pa.
2
Hence, we want to apply a 6736.6 N/m force on the wing.
To initialize a pressure load,
Go to Environment menu bar > select Loads > Force.
Make sure the surface selection filter
wing, as shown in the image below.
43
is selected and choose the lower surface of the
When the surfaces have been selected, go to detail window
Geometry > Apply
Define By > Components > Define the Y Component as 6736.65 Pa.
We are now ready to set up the solution and solve.
Solution
Deformation
To add deformation to the solution, first click
to add the solution sub
menu to menu bar. Now in the solution sub menu click Deformation > Total to add the total
deformation to the solution. It should appear in the outline tree.
Equivalent Stress
In the solution sub menu, select stress > Equivalent (Von-Miss). In the details pane, ensure
Geometry is set to ALL BODIES.
Shear Stress
In the solution sub menu, select stress > shear stress. In the details pane, ensure Orientation is
set to X Axis, and Geometry is set to All Bodies. Rename the Stress to Stress XX by right
clicking Shear Stress in the Outline window and selecting Rename.
44
Stress intensity
In the solution sub menu, select stress > stress intensity. In the detail pane, ensure Geometry
is set to All Bodies.
We are ready to solve the simulation. Press
Case 1: Solutions for static structural analysis:
1) Total deformation
2) Equivalent stress
45
3) Shear stress
4) Stress intensity
Save the file at proper place in your system. Make a copy of this whole folder and rename it
as “wing structure 2”.
46
Verification & Validation
Refine the Mesh:
One of the ways we can check the validity of our analysis is by refining our mesh. If the
values for our results approach a limit, then we have arrived at our answer. If the values
change drastically when we refine the mesh, then we need to refine the mesh further and we
have not yet found an acceptable solution. We will refine the mesh by increasing the number
of divisions in our edge sizing. In the Outline window, go to Mesh > Edge Sizing > Number
of Divisions > 40. Also, go to Mesh > Edge Sizing 2 > Number of Divisions > 80. Our new
mesh looks like this:
47
Case 2: Solutions for static structural analysis:
1) Total deformation
2) Equivalent stresses
48
3) Shear stress
4) Stress intensity
49
12.Modal analysis:
Section 2
In modal analysis we are interested to find the first six modes of shape of vibration. The first
six natural frequency of the system will be found out using ANSYS workbench which will
serve as a base for us for transient and Vibrational analysis of the system.
Problem Specification
A wing with a NACA 64-215 airfoil section is supported such that one end is fixed and the
other end is free. The wing has a root chord of 9.4 meter and tip chord of 2.75 meter, sweep
angle of 28° at quarter chord length, mid-span of 22.42 meters, and a thickness of 0.01
meters. The wing is Aluminium 6061-T6. Find the first 6 modes of vibration of the airfoil
using ANSYS Workbench.
Pre-Analysis & Start-Up
Open ANSYS Workbench
Open ANSYS Workbench by going to Start > ANSYS > Workbench.
To begin, we need to tell ANSYS what kind of
simulation we are doing. If you look to the left of
the start-up window, you will see the Toolbox
Window. Take a look through the different
selections. We are doing a modal analysis
simulation. Load the Modal (ANSYS) box by
dragging and dropping it into the Project
Schematic.
50
Name the project Modal Analysis.
Engineering Properties
Now we need to specify what type of material we are working with. Double click
Engineering Data and it will take you to the Engineering Data Menus.
If you look under the Outline of Schematic A2: Engineering Data Window, you will see that
the default material is Structural Steel. The Problem Specification states we will be using
Aluminum 6061-T6. To add a new material, click in an empty box labelled Click here to add
a new material and give it a name. Our Material is Al 6061-T6. On the left hand side of the
screen in the toolbox window, expand linear Elastic and double click Isotropic Elasticity to
specify E and in the properties of Al 6061-T6 window, Set the Elastic Modulus units to Pa.,
set the magnitude as 1E7, and set the Poisson Ratio to 0.33. We will need to define the
density as well. Expand Physical Properties and double click density. In the Properties of: Al
6061-T6 window, a density bar will have appeared. Define it as being 2700 kg/m^3.
Now that the Material has been specified, we are ready to load the geometry in ANSYS.
51
Geometry
To open the file in ANSYS go to File > Import. Browse to the geometry location on your
computer. If you do not see the file, make sure you are browsing for geometry files (the pull
down menu at the bottom right of the browsing window for computers running Windows 7).
Select the Geometry and click Open. This will import your geometry into ANSYS. Your
project window should now include the main project, and the newly imported geometry (see
below).
Now that the geometry has been imported, let's open the file and make sure everything is in
order! Double click
. This will open the design modeler. When you are
prompted, select Meter as your standard unit of measurement. The first thing you should
notice is that the geometry is not there, so click
to generate the geometry. When
the geometry finally generates, you should see the screen below.
52
Once we are satisfied with our geometry, we can close the design modeler. Now, we should
be looking at the Project window. To connect the geometry to the project, click anddrag
. As soon as you drag the box, ANSYS will highlight the geometry and model
boxes in the main project.
Drag and drop the geometry box
onto
The geometry has been connected the project and we are ready for the next step.
53
Mesh
Open the Mesher
To open the mesher, double click the Model box
in the Project
Outline window. This will load ANSYS Mechanical. You should now be able to see the
airfoil geometry.
The first thing we are going to need to do when the mesher opens is specify the thickness of
the airfoil walls. In the Outline window, expand Geometry and select Surface Body. In the
Details window, change the thickness to 0.01 m. We also need to specify the material. In the
Outline window. In the Details window, select Material > Assignment > Al 6061 -T6. The
material has now been specified.
Mapped Face Meshing
To apply a mapped face meshing, first click on Mesh in the Outline window. This will bring
up the Meshing Menu Bar at the top of the screen. Next, select Mesh Control > Mapped Face
Meshing. Select the 2 faces of the mesh by holding down the left mouse button and dragging
over the entire geometry. In the Details window, click Geometry > Apply - it should say 2
faces are selected.
Edge Sizing
In the Meshing Menu, click Meshing Control > Sizing. Click the edge selection filter
.
Select the 4 curved edges on the outside of the geometry that make up the shape of the
NACA 64-215 Airfoil as the picture shows:
54
In the details window, select Geometry > Apply, and select Type > Number of Divisions .
Change the Number of Divisions to 10. Also, change Behavior > Hard.
Next, create another Edge Sizing, and this time, select the 2 edges at the very front and very
back of the airfoil that run along the wingspan, as the picture shows:
Again, in the Details window change the settings such that Type > Number of Divisions and
Behavior > Hard. This time, change the Number of Divisions to 20. Generate the mesh by
selecting Mesh > Generate Mesh.
55
Setup
Fixed Support
Next, we will apply the boundary conditions to the geometry. In the graphics window, click
the positive Z-Axis on the compass to look at one side of the airfoil.
In the Outline window, select Modal to bring up the
Setup Menu. In the Setup Menu, select Supports > Fixed
Supports. Make sure the Edge Selection Filter is
selected, hold down Ctrl, and left mouse click the upper
and lower edges of the airfoil you are looking at. In the
details window, select Geometry > Apply.
This is all we have to do to setup
this problem.
56
Solution
ANSYS will by default solve for the frequencies of the first 6 vibration modes; however, we
would also like to see how this affects the geometry. We can accomplish this task by looking
at the total deformations of the airfoil to see where the nodes occur and how the geometry
deforms. To tell ANSYS to solve for the deformation, first select Solution in the Outline
window to bring up the Solution Menu bar. In the Solution Menu, select Deformation >
Total. In the Details Window, notice that the deformation is solving for Mode 1. Rename
Total Deformation to Mode Shape 1.
Create another instance total deformation and rename it Mode Shape 2. Select it, and change
Mode > 2. Now, you will be solving for the deformation of the 2nd Mode. Repeat this step
until you are solving for the total deformation of all 6 modes.
57
To solve the system, press
58
Solution:
1) Mode shape1
2) Mode shape 2
59
3) Mode shape 3
4) Mode shape 4
60
5) Mode shape 5
6) Mode shape 6
61
Verification & Validation
Refine the Mesh
Case 2:
One of the ways we can check the validity of our analysis is by refining our mesh. If the
values for our frequencies approach a limit, then we have arrived at our answer. If the values
change drastically when we refine the mesh, then we need to refine the mesh further and we
have not yet found an acceptable solution. We will refine the mesh by increasing the number
of divisions in our edge sizing. In the Outline window, go to Mesh > Edge Sizing > Number
of Divisions >20. Also, go to Mesh > Edge Sizing 2 > Number of Divisions > 40.
62
Solution for case 2:
1) Mode shape 1
2) Mode shape 2
63
3) Mode shape 3
4) Mode shape 4
64
5) Mode shape 5
6) Mode shape 6
65
13.Final results
2
1.5
Coefficient Value
1
-15
0.5
CL
CD
0
-10
-5
0
5
10
15
20
CM
-0.5
-1
-1.5
Angle Of Attack
Following graph shows that 16 is the stall angle. While the drag polar plot shown below
suggests that the airfoil has very good flow property.
CL Vs Cd
0.03
Coefficient of Drag (CD)
0.025
0.02
0.015
Y-Values
0.01
0.005
0
-1.5
66
-1
-0.5
0
0.5
1
Coefficient of lift (CL)
1.5
2
Result for the static structural analysis:
Sr. No.
Types of analysis
Case 1 :
Case 2 :
Edge sizing 1 = 20
Edge sizing 1 = 40
Edge sizing 2 =40
Edge sizing 2 = 80
1
Total deformation
.59828m
.5964m
2
Equivalent stress
1.8001 e8 Pa
2.341 e8 Pa
3
Shear Stress
7.2481 e7 Pa
7.0918 e7 Pa
4
Stress intensity
2.0371 e8 Pa
2.6882 e8 Pa
Result for the modal analysis:
Sr. no.
Mode shape
Case1 frequency
Case 1 maximum Case 2 frequency
Case 2 maximum
no.
(Hz)
amplitude (m)
(Hz)
amplitude (m)
1.
1
1.0008
0.011158
1.0033
0.011169
2.
2
4.0656
0.01237
4.0758
0.012396
3.
3
7.7655
0.010838
7.7923
0.0108957
4.
4
9.1175
0.013119
9.1361
0.013154
5.
5
9.2167
0.1448
9.4477
0.14296
6.
6
9.4769
0.14529
9.5261
0.14311
67
14. Conclusion
From the above results we can conclude that the difference between the values of case 1 and
case 2 i.e. unrefined and refined mesh sizes respectively are minimal. So the results obtained
are validated and verified.
If the difference between the two result values would be considerable, than we have to go for
the fine refined meshing in order to get more accurate results. Although all the results through
ANSYS are approximate and one can get close to that approximate result by decreasing the
mesh size.
If the stresses induced in the body exceeds the ultimate strength of the material than there are
chances that the material will fail. ANSYS will not show that at what instant of stress, the
material will break. Its users part to analysis the reading and compare it with some standard
reference data and arrive at some reasonable conclusion with the help of some considerable
assumptions.
The ultimate strength of the material used above is Aluminum alloy T6 6061 is 290 Mpa.
From the above table we can observe that the all the values of stresses are below 290 Mpa.
Thus, we can conclude that at the above assumed loading conditions and constraints our
wing structure will not fail due to material properties.
68
15. Scope of further studies
The problems solved above are very simple in nature. In actual practice, the problems, loading
conditions, constrains encountered are very different and even more complex in nature.
This is just a basic approach which can be applied to solve the more problems which are
complex in nature and problems from other domain such as thermal analysis, fluid flow
analysis, non-linear analysis, transient analysis, combination of the two or more analysis at a
same time can be solved with some extra efforts.
69
16.Bibliography
 Official Lecture Notes for INME 4717, 5717
 Aircraft Structures for Engineers

Vijay K. Goyal
 Aircraft structures for Engineering students

T.H.Megson
 Aircraft Structures

D.J.Peery
 Designfoil software.

www.designairfoil.com
 Modal analysis tutorial for an aircraft wing

https://confluence.cornell.edu/display/SIMULATION/ANSYS++Modal+Analysis+of+a+Wing
 Static analysis tutorial for a wind turbine blade

https://confluence.cornell.edu/display/SIMULATION/ANSYS++Wind+Turbine+Blade
 Airbus A300-600 series characteristic manual for airport planning

70
http://www.airbus.com/fileadmin/media_gallery/files/tech_data