structural and modal analysis of a300 wing
DESCRIPTION
PROJECT REPORTONStructural and modal analysis of A300 wing structureBASP 002SUBMITTED IN THE PARTIAL FULFILLMENT OF THE DEGREE OF BACHELOR OF TECHNOLOGY IN AEROSPACE ENGINEERINGMr. Darshak Bhuptani Enrolment Number: 093574710 Year of Submission: December - 2012Indira Gandhi National Open University, New Delhi.Indian Institute for Aeronautical Engineering & Information Technology.S.No.85, SHASTRI CAMPUS, NDA ROAD,SHIVANE, PUNE-411023 (M.S) 1CERTIFICATEThis is to certify that DATRANSCRIPT
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PROJECT REPORT
ON
Structural and modal analysis of A300 wing structure
BASP 002
SUBMITTED IN THE PARTIAL FULFILLMENT OF THE DEGREE OF
BACHELOR OF TECHNOLOGY IN AEROSPACE ENGINEERING
Mr. Darshak Bhuptani
Enrolment Number: 093574710
Year of Submission: December - 2012
Indira Gandhi National Open University, New Delhi.
Indian Institute for Aeronautical Engineering
& Information Technology. S.No.85, SHASTRI CAMPUS, NDA ROAD,SHIVANE, PUNE-411023 (M.S)
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CERTIFICATE
This is to certify that DARSHAK BHUPTANI has successfully completed the project entitled
“Structural and modal analysis of A300 wing structure” in fulfillment, for the award of
B.Tech – Aerospace Engineering.
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INDIAN INSTITUTE FOR AERONAUTICAL ENGINEERING &
INFORMATION TECHNOLOGY
DEPARTMENT OF AEROSPACE ENGINEERING
CERTIFICATE
Certified that the project work entitled Structural and modal analysis of A300 wing structure
is a Bonified work done by DARSHAK BHUPTANI bearing Enrollment Number: 093574710
in the final year, eighth semester B.Tech in Aerospace Engineering from IGNOU, New Delhi.
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Acknowledgements
I owe a great many thanks to people who helped and supported me during this project. The
technical assistance, industrial exposure and advice provided by Mr. Gopal Belurkar, IGTR
(Indo-German Tool Room) Aurangabad, is greatly appreciated.
The author would like to express his gratitude to Mr. Ravindra Deb, Trainee engineer,
IGTR Aurangabad for the assistance provided with the modeling and designing part of a wing
structure. Thanks to Mr. Maruf Islam for providing all the required technical supports and
assistance for making this project possible.
I express my thanks to the Principal and H.O.D. of INDIAN INSTITUTE FOR
AERONAUTICAL ENGINEERING AND INFORMATION TECHNOLOGY, Pune for the
project approval and support.
Darshak Bhuptani
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Preface
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.
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Index
1. Objective of the study 06
2. Methodology of the study 07
3. Statement of the problem 08
4. Data 10
5. Analysis 39
6. Final results 55
7. Conclusion 56
8. Scope of the future study 57
9. Biodata 58
10. Bibliography 59
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Objective of the study
The objective of the project is to demonstrate a CAD model of a wing structure and importing it
to the ANSYS 14 workbench and then applying the necessary boundary conditions which
replicates the condition during the take-off of A300 aircraft normally from any airport. The
various structural parameters are determined with considerable assumptions taking into the
account to simplify the problem. The objective of project is to study structural analysis of a wing
structure of A300 aircraft series during take-off and climbing phase through finite element
analysis.
1. Study about the A300 wing design.
2. To study various conditions which can be replicated to get the most approximate results.
3. To create the CAD model most accurately to the dimensions with the limited information
available as any company does not release all the design data.
4. Investigate structural behavior of a wing during take-off and climbing phase.
5. Learn ANSYS and how to utilize it in the aerospace field.
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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) 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.
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b) Model experimental and model validation
The validation and updating of preliminary/sub-structural model are essential to 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
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Statement of the problem
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
variation in load factor which induces various types of stresses in various components of aircraft
body.
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 in at 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.
As all the details required for solving this problem are not available appropriate
assumptions are made wherever required to simplify the problem. To simplify the problem the
assumptions made are as follows:
The airfoil used in A300 is supercritical airfoil so we are using NACA 64-215 airfoil
throughout the wing structure.
The rib thickness is 100 mm, which is mirror extended from its mean position.
The diameter of the front spar is 300 mm and it is placed at 0.25 times the chord length at
each section from the leading edge.
The diameter of rear spar is 250 mm and it is placed at 0.7 times the chord length at each
section from the leading edge.
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The centre point of front and rear spar at the tip airfoil is at a distance of 12.54 m and
13.465 m from the reference point respectively.
The front spar is at 31° from the reference line while the rear spar is at 23° from the
reference line.
The holes are made in the ribs in order to save weight of the structure.
The material used for the whole structural and modal analysis purpose is aluminium alloy
with density of 2700 kg/m3, young modulus of 68300 MPa, and poison’s ratio of 0.34.
The boundary conditions applied to the FEA model is that the root section of the airfoil as
well as the spars is fixed so that the degree of freedom is restricted in all the six
directions.
The loading condition is found using the maximum take-off weight and maximum climb
angle which is allowed for this aircraft from any airport.
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Data for modeling and analyzing
Overall general data section
To complete this project, we will follow the following flow chart to do this project in a proper
sequence respectively.
The important stages are creating CATIA model, defining constrains, results, redefining the
mesh size and comparing results to the original results to validate the results and conclusion.
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Stage 1:
Data collection.
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Aircraft type model A300-600R
Wing area (m2) 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
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Stage 2:
CAD Modeling section
The first step is to get the airfoil shape in the CATIA V5 R18, part design.
With the help of software, commercially known as “designfoil software” which is available for 5
days as a trial version is used to create the airfoil shape by plotting all the co-ordinates in the
catia part design workbench.
The main benefit of this software is that all the co-ordinates are the function of the chord length,
that is (x/c, y/c).
NACA 64215 airfoil with the chord length of 2.78 m is exported to catia part design file.
For the trial version, 2.78 m is the maximum chord length available so it is selected.
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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 the mid wing span is 22.42 m we divide our airfoil in 23 sections each placed at an equal
distance from the reference airfoil. The distance between two airfoils is 1 m. the diameter of the
fuselage is 5.64 m, some part of our wing will be inside fuselage and this section is completely
rigid due to its wing box design. 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.
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.
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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 whose co-ordinates is (0, 0) 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 whose co-ordinate is (0, 0) 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
Local taper ratio =
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The following values are found with the help of geometry and trigonometry relations:
Section no. Local chord length c
(m) Local taper ratio
Distance of leading
edge point from the
reference line 1 (m)
Y
Distance of trailing
edge point from the
reference line 2 (m)
Z
1 9.4 1 0 4.89
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
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The above sketch is the result of the all the calculations which were carried out and 24 sections
of airfoil are being projected at regular interval from the reference plane.
In the wireframe and surface design workbench, the surface for the following sections has been
generated accordingly.
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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.
The spars and holes are being created in the wing design as per our assumptions respectively.
The complete design of the wing structure will be as shown below,
Before importing the .CAT file to the Ansys workbench, the file has to be converted into .IGS
format.
This conversion can be done by going to file option>save as> save as type: .igs format>save.
There will be some data loss during conversion and importing process resulting in approximate
results in ANSYS workbench.
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Analysis
FEA section for static 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 seen below
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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. Because 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.
Geometry
In Workbench in the Project Schematic window, go to File > Import. In
the Import window that opens, change the file type (next to the File Name text box)
to Geometry File. Select the downloaded geometry file and press Open. The geometry should
now be in the project schematic, as shown below.
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Generate the Geometry
Next, we will open the file to generate the geometry. Double click the imported
geometry to open the Design Modeler. When the Design Modeler opens, a
pop up window will ask us for the default units of measurement for the geometry.
Select Meter and then press OK. After you select the units, you will notice
the Graphics window is empty. We will fix this soon. First, click on in
the Outline window. In the Details window, change Operation from Add Material to Add
Frozen. Finally, generate the part by clicking Once you press , the
imported geometry should show in the Graphics window.
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Form 1 Part
Notice in the Outline window that the part has been imported as 2 separate parts.
We need to form 1 part from the 2 separate surfaces that were imported. While pressing Ctrl, left
click the two Surface Bodies, then right click one of the Surface Bodies and select Form
New Part.
Now, in the Outline window, you should see 1 Part with the 2 Surface Bodies as subordinates.
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While we are here, let's rename the surfaces so we may refer to them later. Name the outer
surface Outer Surface and name the spar Spar. To rename the surface, right click on it (the
surface you are renaming will be highlighted in the graphics window), and go to Rename.
Connect the Geometry
Next, we need to connect the geometry to our current project. Close the Design Modeler and
return to the project schematic. First click (and hold) on the imported geometry
box Drag and drop on . When you are finished, a line
should connect the two boxes showing that you have successfully linked them.
Now that the geometry is imported and generated, we are ready to mesh the geometry.
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Mesh
Initial Setup
Close the Design Modeler if you haven't already, and 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 and 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 next
to Thickness. Repeat with the value of 3e-2 for Spar to thickness.
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There should no longer be a question mark next to Geometry.
Delete any Connections
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,
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:
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In the details window, select Geometry > Apply, and select Type > Number of Divisions.
Change the Number of Divisions to 20. 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 40.
Generate the mesh by selecting Mesh > Generate Mesh
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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, 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.
Pressure Load
We want to apply a 6736.6 N/m2 upward force on the wing. To initialize a pressure load, in
the Environment menu bar select Loads > Force. Make sure the surface selection filter is
selected and choose the lower surface of the wing, as shown in the image below.
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The pressure load is determined by the calculating the load factor from the Arccosine of 17°.
The maximum climb angle for A300 from any airport is 17°.
n =
= 1.04569
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 semi wing area of 130 m2.
Therefore, the total pressure load applied from the bottom of the surface is 6736.65 Pa.
When the surfaces have been selected, press Geometry > Apply in the Details window. Next,
select Define By > Components. Define the Y Component as 6736.65 Pa.
We are now ready to set up the solution and solve.
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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-Mises). 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.
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
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Stage 4:
Solutions for static structural analysis:
Case 1:
1) Total deformation
2) Equivalent stress
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3) Shear stress
4) Stress intensity
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Save the file at proper place in your system. Make a copy of this whole folder and rename it
as “wing structure 2”.
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:
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Case 2: Solutions for static structural analysis:
1) Total deformation
2) Equivalent stresses
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3) Shear stress
4) Stress intensity
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FEA section for 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. This will open the start up
screen seen as seen below
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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.
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.
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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
Aluminium 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.
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).
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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 modeller. 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.
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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 and
drag . 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
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The geometry has been connected the project and we are ready for the next step.
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.
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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:
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:
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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.
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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.
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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.
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To solve the system, press
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Solution:
1) Mode shape1
2) Mode shape 2
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3) Mode shape 3
4) Mode shape 4
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5) Mode shape 5
6) Mode shape 6
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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.
52
Solution for case 2:
1) Mode shape 1
2) Mode shape 2
53
3) Mode shape 3
4) Mode shape 4
54
5) Mode shape 5
6) Mode shape 6
55
Final results
Stage 5:
Result for the static structural analysis:
Sr. no. Type of analysis Case 1: Edge sizing1- 20
Edge sizing 2- 40
Case 2: Edge sizing 2 - 40
Edge sizing 2 – 80
1. Total deformation 0.59828 m 0.5964 m
2. Equivalent stresses 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
no.
Case1 frequency
(Hz)
Case 1 maximum
amplitude (m)
Case 2 frequency
(Hz)
Case 2 maximum
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
56
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.
57
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.
58
Biodata
Name : Bhuptani Darshak Krishnkant
Fathers Name : Krishnkant Harilal Bhuptani
Date of Birth : 25th
of December 1990
Branch : B.Tech in Aerospace Engineering
Enrollment number: 093574710
Date of Enrollment: Jan, 2009
Educational Qualification:
Sr.
No. Qualification Score
Year of
passing Board/University College/Institute
1 SSC 78.93% 2006 Maharashtra State
Board
Holy Angels High School
Mumbai-81
2 HSC 75.67% 2008 Maharashtra State
Board
NES Ratnam Jr. college of
Science, Mumbai-78
3
B.Tech in
Aerospace
Engineering
Sem 7 75.08%
Sem 6 69.41%
Sem 5 79.62%
Sem 4 80.85%
Sem 3 84.11%
Sem 2 82.11%
Sem 1
71.28%
2012
2011
2011
2010
2010
2009
2009
Indira Gandhi
National Open
University
Indian Institute for
Aeronautical Engineering
and Information
Technology, Pune-52.
Darshak Bhuptani
59
Bibliography
Official Lecture Notes for INME 4717, 5717
- Aircraft Structures for Engineers
Vijay K. Goyal, Ph.D.
Associate Professor, Mechanical Engineering Department,
University of Puerto Rico at Mayaguez,
Mayaguez, Puerto Rico
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+WB+-
+Modal+Analysis+of+a+Wing+-+Problem+Specification
Static analysis tutorial for a wind turbine blade
https://confluence.cornell.edu/display/SIMULATION/ANSYS+WB+-
+Wind+Turbine+Blade+-+Problem+Specification
Airbus A300-600 series characteristic manual for airport planning
http://www.airbus.com/fileadmin/media_gallery/files/tech_data/AC/AC_A300-
600_20091201.pdf