structural and modal analysis of a300 wing.pdf

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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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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 (m ) 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 Fi le 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 Operati on from Add Materi al 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 Par t and select Out er Surface.

Notice that Thick ness 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 Geometr y.

Delete any Connections

ANSYS may create connections automatically - however they are not required for thissimulation 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 Projec t

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 Outl ine 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 Geomet ry > App ly, 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

Divisio ns 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 Envi ronment 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 Def ine By > Componen ts . Define the Y Component as 6736.65 Pa.

We are now ready to set up the solution and solve.



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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 Out line window, go to Mesh > Edge Siz ing > Number of

Divisions > 40. Also, go to Mesh > Edge Si zing 2 > Number of Div isions > 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 Moda l(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 Elas tic and double click Iso tropic

Elastici ty to specify E and in the Proper ties 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 Proper ti es 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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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 Projec t

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 Outl ine window, expand Geometry and select Surface Body. In the

Deta ils 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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Again, in the Details window change the settings such that Type > Number of

Divisio ns 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 Outl ine 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 Defo rmation > To ta l. In

the Deta ils 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 Out line window, go to Mesh > Edge Siz ing > Number of

Divisions >20. Also, go to Mesh > Edge Sizing 2 > Number of Divisions > 40.



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Solution for case 2:

1) Mode shape 1

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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Final results

Stage 5:

Result for the static structural analysis:

Sr. no. Type of analysisCase 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



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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.



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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

passingBoard/University College/Institute

1 SSC 78.93% 2006Maharashtra State

Board

Holy Angels High School

Mumbai-81

2 HSC 75.67% 2008Maharashtra 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



Bibliography

Official Lecture Notes for INME 4717, 5717

- Aircraft Structures for EngineersVijay 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

https://confluence.cornell.edu/display/SIMULATION/ANSYS+WB+-+Modal+Analysis+of+a+Wing+-+Problem+Specification


https://confluence.cornell.edu/display/SIMULATION/ANSYS+WB+-+Wind+Turbine+Blade+-+Problem+Specification


http://www.airbus.com/fileadmin/media_gallery/files/tech_data/AC/AC_A300-600_20091201.pdf








structural and modal analysis of a300 wing.pdf

Documents