Structures of an aircraft can be categorised as primary structural components and secondary structure components. Primary structure components are the components which lead to failure of the aircraft if such component is failed during the flight cycle. Secondary components are load sharing components in an aircraft but will not pave the way to catastrophic failure. Designing aircraft structures should follow several strategies to assure safety. For that, there are three main methods used in designing and maintenance procedures. First one is the safe flight, which an aircraft component has a lifetime. That component is not used beyond that limit and should replace though it is not failed. The fail-safe method is another one that redundant systems or components are there to ensure there is another way to carry the load or do necessary control. The final one is the damage tolerance which measures the current damages are within acceptable limit and carry out the main functions until the next main maintenance process. To determine the safety of a structure component load distribution, stress and strain variation, deflection can be used as parameters to make sure that component can withstand maximum allowable load with safety factor. There are several techniques used to get accurate results as numerical methods, Finite Element Method (FEM) and experimental methods. In the design process, those three steps are followed in an orderly manner to ensure the safety of an aircraft.

1. Experimental and Numerical Stress Analysis of a
Rectangular Wing Structure

2. 1
1 Table of Contents
1 Table of Contents.............................................................................................................................. 1
2 Table of figures................................................................................................................................. 2
3 Introduction....................................................................................................................................... 3
4 Experimental setup and procedure.................................................................................................... 4
5 Experimental results analysis............................................................................................................ 6
6 FEM method ................................................................................................................................... 10
7 FEM results analysis....................................................................................................................... 15
8 Discussion....................................................................................................................................... 28
9 Conclusion ...................................................................................................................................... 30
10 References ................................................................................................................................... 31

3. 2
2 Table of figures
Figure 1 - Loosen friction lock to............................................................................................................. 4
Figure 2 - Load wheel .............................................................................................................................. 5
Figure 3 - Strain gauge arrangement........................................................................................................ 7
Figure 4 - New static structural system.................................................................................................. 10
Figure 5 - Define new material .............................................................................................................. 11
Figure 6 - Import geometry .................................................................................................................... 11
Figure 7 - Select material properties ...................................................................................................... 12
Figure 8 - Define fixed points ................................................................................................................ 12
Figure 9 - Define acting forces............................................................................................................... 12
Figure 10 – Define the required results.................................................................................................. 12
Figure 11 - Mesh view of ribs and spars arrangement ........................................................................... 13
Figure 12 - Mesh view of the skin.......................................................................................................... 14
Figure 13 - Mesh view of the assembly ................................................................................................. 14

4. 3
3 Introduction
Structures of an aircraft can be categorised as primary structural components and secondary
structure components. Primary structure components are the components which lead to failure of the
aircraft if such component is failed during the flight cycle. Secondary components are load sharing
components in an aircraft but will not pave the way to catastrophic failure.
Designing aircraft structures should follow several strategies to assure safety. For that, there are
three main methods used in designing and maintenance procedures. First one is the safe flight, which an
aircraft component has a lifetime. That component is not used beyond that limit and should replace
though it is not failed. The fail-safe method is another one that redundant systems or components are
there to ensure there is another way to carry the load or do necessary control. The final one is the damage
tolerance which measures the current damages are within acceptable limit and carry out the main
functions until the next main maintenance process.
To determine the safety of a structure component load distribution, stress and strain variation,
deflection can be used as parameters to make sure that component can withstand maximum allowable
load with safety factor. There are several techniques used to get accurate results as numerical methods,
Finite Element Method (FEM) and experimental methods. In the design process, those three steps are
followed in an orderly manner to ensure the safety of an aircraft.
FEM is a widely used technique that relies on computational power and modelling techniques. That
method is fast and can get accurate results to ensure the strength and other mechanical properties are in
the required region. But for the realistic results, should move on to the experimental methods.
In this report, FEM is used to analyse the aircraft wing test specimen and the experimental procedure
is carried out to get results from the experimental test rig. Both these results analyse and compare those
two at the end of the report. Finally, the accuracy of these two methods, the importance of these two
methods and further improvements to get accurate results are discussed at the end.

5. 4
4 Experimental setup and procedure
There are procedures to take measurements from the wing box in the lab.
Step 1: The complete unit was moved to the required position and waited a couple of minutes until
it acclimates the room temperature is the temperature is changed.
Step 2: The wing was positioned to align fixture holes at the end of the spar to the fixture.
Step 3: The knobs were tightened to fix the wing to the fixture. To remove any sag in the system,
the wing structure was lifted from the cantilevered end. The knobs were tightened gradually while
continuing to lift the end of the wing.
Step 3: The strain gauges were connected to strain bridge controller to get the readings.
Step 4: Turn on the strain bridge controller and keep it several minutes to acclimate.
Step 5: After the system gets stable, record the rest voltages that indicated in the strain bridge
controller for all strain gauges. To that, turn the channel knob to each strain gauge channel manually and
note down those values.
Step 6: Lose the friction lock and slide the load cell point load applicator unit under the required
position. In this practical, front spar and rear spar are the two locations. First, the load cell was placed
under the middle of the front spar.
Step 7: The friction lock was tightened up.
Step 8: Force is gradually applied to the wing box by rotating crank handle clockwise. The crack
handle was rotated until the required force value was applied to the system. Force values were displayed
in the load panel meter.
Step 9: Allow the system to settle down for about minute before taking the measurement. Then
voltage values of each strain gauge were taken down, indicated in the strain bridge controller.
Step 10: All the voltage values that were taken down, entered to the given software with other
necessary data.
Step 11: Calculated strain values were taken from the software.
Step 12: Repeat steps from step 6 and corresponding voltage values were taken down for rear spar.
Figure 1 - Loosen friction lock to
position load cell

6. 5
Figure 2 - Load wheel

7. 6
5 Experimental results analysis
Experimental data
Material – 6061-T6 Aluminium
Material properties
• Density – 2700 kg/m3
• Tensile yield strength – 276 MPa
• Ultimate tensile strength – 310 MPa
• Modulus of Elasticity – 68.9 GPa
• Bearing yield stress – 386 MPa
• Shear modulus – 36 GPa
• Shear strength – 207 MPa
• Fatigue strength – 96.5 MPa
• Poisson’s Ratio – 0.33
Wing Front Spar
Strain gauge
number
Output Voltage
Strain Stress (MPa)Unloaded
Condition
Loaded
Condition
A 24.52 23.39 -0.000144157 -9.932
B 31.42 32.45 0.000131438 9.056
C 19.24 19.4 0.000020415 1.407
D 1.67 21.66 -0.000001276 -0.088
E 23.44 23.2 -0.000030621 -2.110
F 7.5 7.3 -0.000025518 -1.758
G 22.74 21.94 -0.000102062 -7.032
H 17.44 16.81 -0.000080376 -5.538
I 17.47 17.83 0.000045935 3.165
J 6.99 7.56 0.000072733 5.011
K 25.99 26.09 0.000012759 0.879
L 36.4 36.31 -0.000011483 -0.791

8. 7
Wing Rear Spar
Strain gauge
number
Output Voltage
Strain Stress (MPa)Unloaded
Condition
Loaded
Condition
A 24.37 23.63 -0.000094409 -6.505
B 31.38 32.06 0.00008677 0.598
C 19.16 19.27 0.000014035 0.967
D 21.6 21.57 -0.000003828 -0.264
E 23.4 23.2 -0.000025518 -1.758
F 7.47 7.01 -0.000058689 -4.044
G 22.85 22.17 -0.000086754 -5.977
H 17.47 17.24 -0.000029345 -2.022
I 17.38 17.5 0.000016587 1.143
J 6.67 7.15 0.000061248 4.220
K 25.86 26.16 0.000038279 2.637
L 36.37 36.35 -0.000002552 -0.176
Figure 3 - Strain gauge arrangement

9. 8
Loaded Spar Maximum strain value in
spar
Maximum stress value in spar
Front spar 0.000131438 9056078.2 Pa
Rear spar 0.00008677 5978453 Pa
Loaded Spar Minimum strain value in
spar
Minimum stress value in spar
Front spar -0.000144157 -9932417 Pa
Rear spar -0.000094409 -6504780 Pa
Loaded Spar Maximum strain value in
skin
Maximum stress value in skin
Front spar 0.000072733 5011303.7 Pa
Rear spar 0.000061248 4219987 Pa
Loaded Spar Minimum strain value in
skin
Minimum stress value in skin
Front spar -0.000102062 -7032072 Pa
Rear spar -0.000086754 -5977351 Pa
As the directions of each strain gauges are different, the comparison of strain values is not possible.
So that comparison in each loading conditions cannot be compared. The two different loading conditions,
front spar loading, and rear spar loading can be compared. That gives what spar is highly stressed after
applying the load.
When applied load to the front spar, the stress values of the front spar are higher than the rear spar
values. That result is trivial as the result of applying force directly to the front spar. In wing designing,
front spar bears most of the load in the aircraft so that this is a critical part. In this practical also there
are several strain gauges placed on the front spar. The lack of the number of strain gauges in the rear
spar makes some difficulties to compare strain and stress values after the load is applied.
Comparison of the skin stress with the different loading points can be carried out because there are
enough strain gauges located in the skin. Similar strain gauges arrangements can be compared to get
good results.
Strain gauge Front spar loading – stress Rear spar loading - stress
G -7.032 -5.977
J 5.011 4.220

10. 9
Strain gauge Front spar loading – stress Rear spar loading - stress
F -1.758 -4.044
K 0.879 2.637
Strain gauge Front spar loading – stress Rear spar loading - stress
H -5.538 -2.022
I 3.165 1.143
Comparing the stress values using the above results, we can see front spar loading will cause to
make higher skin stress values. So that additional load on the front spar will cause to increase the stress
values in the skin compared with the front spar loading.
Basically, the upper part of the skin has compressive stress and the bottom part of the skin is
designed to bear tension loads. so that there are two different material properties can be seen in the
commercial aircraft wing. In this practical, use the same material is used in the upper and bottom of the
skin. So, the optimum design of the skin cannot be estimated. Also, strain gauges are mounted in the
bottom of the skin only. So, top surface stress cannot be calculated using this experiment.
To apply more load rather than these applied loads, we can say rear spar is the better place to apply
those loads. Front spar gets highly stressed than rear spar so rear spar can bear more load than front spar.
When considering the landing gear mounting, that area should be capable of bear impact load in the
landing and take-off.
Front spar loading Rear spar loading
Deflection 34 mm 41 mm
Front spar is lesser deflected and rear spar is deflected more than that. Based on that fact, more load
can be carried if the load is applied to the front spar. So that, impact loads applied to the front spar.

11. 10
6 FEM method
Finite element method is the most used mathematical model to solve engineering and mathematical
problems. The wing box structure is modelled with CAD and then used FEM method to analyse the
deflection, stress and strain variation. FEM model is developed using ANSYS static structural model
and then evaluate the forces acting on different positions in the wing structure. The steps are taken to
model the wing box structure and method of analysis is as follows.
Step 1: Dimensions of the wing box structure was measured using Vernier calliper, meter ruler.
Wing box skin thickness, the thickness of the rib, fixing hole diameters were measured using Vernier
calliper and length of the wing box, dimensions of the spar and width of the wing box structure were
measured using meter ruler.
Step 2: The geometry was modelled using SOLIDWORKS 2018. There are two different methods
used to model the wing box as,
1. Model the wing box only based on measured dimensions previously.
2. The photo was taken and trace the image with SOLIDWORKS Autotrace Add On. Here
scaled the image with measured data and modelled the parts with dimensions that were
measured in the previous step. The skin was modelled using auto trace due to the complexity
of the geometry.
Step 3: Start ANSYS workbench and make new Static Structural Analysis System (Figure 4)
Step 4: Define new material in Engineering data section. In the material library, there is no material
for 6061 T6 Aluminium. So that new material was declared in the Engineering data section (figure 5)
Step 5: Import the modelled part to the ANSYS (figure 6)
Step 6: Open model in ANSYS Static Structural. Define materials in the Geometry tree (figure 7).
Step 7: Create a mesh and define fine the mesh to get accurate results.
Step 8: Define fixed support of the system (figure 8).
Step 9: Apply force to the corresponding point. First, apply force to the main spar (figure 9).
Step 10: Define required solution under the Solution in the design tree (ex: deflection, stress, etc.).
(Figure 10 )
Step 11: Repeat the procedure to apply the load to the rear spar.
Figure 4 - New static structural system

12. 11
Figure 5 - Define new material
Figure 6 - Import geometry

13. 12
Figure 7 - Select material properties
Figure 9 - Define acting forces
Figure 10 – Define the required results.
Figure 8 - Define fixed points

14. 13
Mesh details
Figure 11 - Mesh view of ribs and spars arrangement

15. 14
Figure 12 - Mesh view of the skin
Figure 13 - Mesh view of the assembly

16. 15
7 FEM results analysis
Results were compared for deformation; stress values and strain values as follows.
Front spar loading – Total deformation
Rear spar loading – Total deformation

17. 16
Front spar loading – Total deformation
Rear spar loading – Total deformation

18. 17
Front spar loading – Normal stress Y-axis
Rear spar loading – Normal stress Y-axis

19. 18
Front spar loading – Shear stress (XY plane)
Rear spar loading – Shear stress (XY plane)

20. 19
Front spar loading – Shear stress (YZ plane)
Rear spar loading – Shear stress (YZ plane)

21. 20
Front spar loading – Shear stress (XZ plane)
Rear spar loading – Shear stress (XZ plane)

22. 21
Front spar loading – Maximum shear stress
Rear spar loading – Maximum Shear stress

23. 22
Front spar loading – Maximum principal stress
Rear spar loading – Maximum principal stress

24. 23
Front spar loading – Normal elastic strain (Y-axis)
Rear spar loading – Normal elastic strain (Y-axis)

25. 24
Front spar loading – Total deformation
Rear spar loading – Total deformation

26. 25
Front spar loading – Shear stress (XY plane)
Rear spar loading – Shear stress (XY plane)

27. 26
Front spar loading – Maximum shear stress
Rear spar loading – Maximum Shear stress

28. 27
Front spar loading – Maximum principal stress
Rear spar loading – Maximum principal stress
With these figures and comparison, the bottom of the skin and the spars are stressed more compared
with the top surface. As the bottom of the wing get tension and top side of the wing get compression,
tension gets more impact here. Stress distribution in the spars is higher than the skin stress distribution.

29. 28
8 Discussion
There are several steps should be followed by the operator before set up and continue the
experiment.
• The test rig should be moved to an appropriate place before starting the experiment. Test rig
should stand still for a couple of minutes before continuing the experiment as it should be
acclimated with the environment. Because strain gauges are sensitive for the temperature
variation. So that there should be enough time for the setup to adapt to the environmental
conditions.
• The test section, wing box will have a play at the end of the section after it fixed to the test
rig. Wing box will have sag. If the experiment was carried out with such deflection, the
results will be erroneous. So, before starting the test, the free end of the wing should list to
remove any sag. After that load is applied to the test section.
• After connecting the wing section to the strain bridge controller, that should wait a couple
of minutes before taking any reading. As same as strain gauges, the controller is sensitive to
the temperature variation. To get accurate results, these steps should be followed.
When modelling the wing box structure, there are two different methods used in this project to
model. One is model the wing box structure using solely based on dimensions that measured using
Vernier calliper a diameter ruler. Another method is to model the wing box structure using Autotrace
feature in SOLIDWORKS to get accurate modelling of the skin contour as that cannot be correctly
measured using Vernier calliper or meter ruler.
In the first method, the wing box was modelled as two parts, spars and ribs structure and skin. Then
these two were assembled using SOLIDWORKS assembly file. All these models were converted to IGS
file format before import to the simulations as the versions of the SOLIDWORKS will make some
difficulties. Assembly mechanism was simple. There are no mechanical fasteners or adhesives used to
assemble the different parts. So that general mating conditions were used for simplicity. If there are nut
bolt mechanical fasteners in ANSYS simulations, the assembly gets complicated. So, all the mechanical
fastens were removed and the common mating system was used.
In the Autotrace method, skin and spar arrangement were modelled as a single part to remove the
mating requirement. As there is no mechanical mating in the ANSYS simulation, single modelling can
be used. But the isolation of the rib arrangement and skin cannot be viewed.
Both these methods were used to evaluate the stress, stress variation of the wing box structure and
both results not deviated much from the other. So, we can say both these methods can be used to get the
results.
Spars were modelled neglecting the small holes and defects because of the dimensions of those are
small compared with the other dimensions. Also, the endplate that is in the wing was neglected as it is
mounted at the edge of the wing.

30. 29
Mesh for the FEM was refined to get more accurate values and that will help to get mesh
independent results. The mesh for the skin was structured and that will give better results. But due to the
complexity of the mesh for the spars, the structured mesh cannot be generated. So, the mesh was refined
as much as possible to get accurate results.
By comparing the experimental and computational results, experimental results give higher
deflection values and higher stress values. There can be a human error while doing the experimental
process and the instrumental error is the reasons for the deviation of results. Because of the deflection
difference, stresses are also varied. But using both these results, we can say that more loads can be
applied to front spar rather than rear spar as rear spar is stressed and deformed than front spar.
[1] [2]

31. 30
9 Conclusion
Landing gear makes higher stresses on the spars and on the skin due to the impact load on the
structure. So, the attachment point of such components should be analysed well. From experimental
analysis, the front spar is highly stressed than the rear spar in most of the shear stress values though the
complete deformation is high when the rear spar is loaded. Due to the low deformation of the front spar,
the landing gear should mount to the front spar.
Human errors, instrumental errors, mesh refinements, geometry modelling errors are caused by the
deviation of the results. But comparing the results, we can conclude that front spar can absorb impact
load rather than rear spar.

32. 31
10 References
[1] “Strain Analysis System - Operator's Manual and Sample Lab Procedure,” Turbine technologies,
2008.
[2] “Strain Analysis System - Operator's Manual and Sample Lab Procedure,” Turbine Technologies,
2015.