PREDICTION OF CRACK AND STRESS ANALYSIS … OF CRACK AND STRESS ANALYSIS OF A CUT ... ABHILASH V J1, LOHITESH JAGA KUMAR2, Dr. IRFAN G3, ... Figure 3.4 Standard atmospheric pressure

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DOI : 10.23883/IJRTER.2017.3360.TYCHV 285

PREDICTION OF CRACK AND STRESS ANALYSIS OF A CUT

OUT IN A AIRFRAME STRUCTURE

ABHILASH V J1, LOHITESH JAGA KUMAR2, Dr. IRFAN G3, THARA R4 1,2,3 Dept.of Mechanical Engineering, Akshaya Institute of Technology, Tumkur

4 Mechanical Engineering, CIT-Gubbi

Abstract—Aircraft is symbol of a high performance mechanical structure with a very high structural

safety record. Rarely an aircraft will fail due to a static overload during its service life. For the

continued airworthiness of an aircraft during its entire economic service life, fatigue and damage

tolerance design, analysis, testing and service experience correlation play a pivotal role.

A stiffened panel consisting of emergency exit door cutout is considered for the finite element

analysis. Emergency exit door cut out region is one of the stress concentration regions in the fuselage

structure. The fuselage cabin will be subjected to internal pressurization at higher altitudes. The

pressurization load case is one the critical load cases considered for the design of the aircraft

structure. The tensile loads at the edge of the panel corresponding to pressurization will be

considered for the linear static analysis of the panel. A location of maximum tensile stress and its

magnitude will be identified from the stress analysis. A fatigue life prediction for crack to initiate

will carried out for such a location.

I. INTRODUCTION

1.1 An Aircraft

An air ship is a mind boggling structure, however an exceptionally productive man-made

flying machine. Flying machines are by and large developed from the essential parts of wings,

fuselage, tail units and control surfaces. Every part has at least one particular capacities and must be

intended to guarantee that it can do these capacities securely. Any little disappointment of any of

these parts may prompt a disastrous fiasco creating gigantic obliteration of passengers and whole

plane. When outlining an air ship, it is finding the ideal extent of heaviness of vehicle and payload. It

should be solid and also stiffer to withstand the uncommon conditions which needs to work.

Solidness is essential component. Additionally, if a section comes up short, it doesn't really bring

about disappointment of the entire air ship. It is as yet feasible for the air ship to skim over a

sheltered landing place just if the streamlined shape is held auxiliary respectability is accomplished.

The bearing load individuals from the main sections and subjected to real powers, are known

as airframe. Airframe is the thing that remaining parts like hardware and frameworks are stripped

away. Nowadays airplanes, the skin assumes a vital part in conveying loads. Sheet metals can

typically just bolster strain. In any case, if the sheet is collapsed, it all of a sudden has the capacity to

convey compressive burdens. Stiffeners are utilized for that. A section of skin, joined with stiffeners,

called stringers, is named a thin-walled structure.

II. LITERATURE REVIEW

A.K.Vasudevan, K. Sadananda (2001) studied the influence of theCritical parameters for

fatigue damage. The fatigue failure examination is analyzed from an authentic point of view. The

examination shows that a portion of the issues concerning the fundamental aberrations between the

trial and model/elucidations. To help comprehend these issues, we have built up an approach with

two main thrust parameters to investigate the weakness conduct. Such an approach helps in review

International Journal of Recent Trends in Engineering & Research (IJRTER) Volume 03, Issue 07; July - 2017 [ISSN: 2455-1457]

@IJRTER-2017, All Rights Reserved 286

the harm regarding an inherent issue instead of an outward one. In the last examination one needs to

bring together the general harm procedures with the end goal that the depiction is finished from the

break start stage to short split to long split to definite disappointment. So as to bring together the

harm procedure, three essential parameters are presented for depicting the general weariness handle.

These are K, Kmax and interior anxiety commitment to Kmax. Likewise, there are different impacts

from condition and temperature that can add to these parameters. Specifically Kmax appears to

assume an imperative part in the general harm handle. We find that the inner anxiety is the missing

connection that can conquer any hindrance between the four principle phases of harm that lies

between the break nucleation stage to conclusive disappointment. Cases are sited in support of this

perspective of clarification. At long last, it is recommended that deliberate test information and

scientific displaying to depict the inside anxiety angles is required to help in framing a solid life

forecast philosophy.

H. Hosseini-Toudeshky, B. Mohammadi(2006)discussed on theA simple method to

calculate the crack growth life of adhesively repaired aluminium panels. Basic strategy is

produced to foresee the break development life of single-side repaired aluminum boards. The patches

are thought to be made of glass/epoxy, boron/epoxy and graphite/epoxy composite materials. For this

reason, split development lives of the repaired boards are computed utilizing two diverse

methodologies: (1) The break developments along the board thickness consistently. In this improved

approach, split bleeding edge stays opposite to the plate surfaces amid the break engendering. (2)

The split developments along the board thickness non-consistently. In this approach, beginning break

cutting edge changes to a bended shape as indicated by the variety of stress power figure amid the

split spread . At that point, another basic strategy is created in light of a parametric review utilizing

both methodologies. Utilizing this straightforward technique, it is not important to play out the mind

boggling break development investigation in view of the genuine split front shape.

Stevan Maksimović(2005) focuses the attention on Fatigue Life Analysis of Aircraft

Structural Components. This work characterizes a viable calculation technique that consolidates

Neuber`s Rule and the limited component strategy with strain–life paradigms to precisely foresee

weariness break start life and afterward build up an expected calendar of exhaustion life. The

proposed technique is connected to the delegate air ship auxiliary parts, for example, a plate with

opening to get anticipated lives. Mineworker's manage was utilized to figure the aggregate harm in

the weakness break start stage. This strategy is then connected to a plate with a focal gap. It is

imperative to stress that exhibited strategy for expectation of weakness life up to split start gives

great relationship exploratory information even with low and high exhaustion spaces. Also, the

previously mentioned strategy can beused for complex basic components. The characterized strategy,

for weakness life expectation of indented flying machine auxiliary segments upto split start, can mull

over uni-pivotal andmulti-hub stacking with consistent and variable plentifulness.

Dr. M. Neubauer, G. Güntherstudied the influence of the “Aircraft Loads AndStructural

Loads. The main to weariness and also erosion, contingent upon the utilization condition, are the

significant purpose behind debasement of structures. The a wide range of classes of burdens, the era

of stacking conditions amid the plan stage, as characterized in the weapons frameworks particular,

thought of static and weakness loads for basic lay-out and approval ideas are displayed. At the point

when life of airplanes are talked about, frequently the flight hours or number of flights are as yet

considered the overseeing variable, at times adjusted with elements on "harm hours" or "use", while

from an auxiliary building perspective the operational anxiety range and in this way the life on the

diverse flying machine segments are not just a matter of flight hours and range proportion

additionally determined by change status, basic weight status and part hardware. This paper portrays

loads-examination and confirmation exercises amid the significant periods of the life of an airframe.



III. STATEMENT OF THE PROBLEM

3.1 Introduction

Aircraft design is the most highly advanced design. Every minute detail has to be checked

before its flight. Primary aircraft structural component consists of fuselage, wing, horizontal and

vertical tail, landing gears, and attachment lug, cut outs etc. some critical components are subjected

to repeated fluctuating loads. One cannot really avoid fatigue loads as they are natural to any

machine component as breath to human body. We should learn about them and component so as to

avoid surprise failures.

In structure of airframe fatigue crack will appears in location of high tensile stress. A stiffened panel

consisting of emergency exit door cutout is considered for the finite element analysis. Emergency

exit door cut out region is one of the stress concentration regions in the fuselage structure. The

fuselage cabin will be subjected to internal pressurization at higher altitudes.

In the current scenario, an attempt has been made to predict the

1. Stress analysis of emergency exit door cutout in the airframe structure using FEA

2. The point of maximum tension and its magnitude identified by the stress analysis.

3. Fatigue life of emergency exit door cutout in aircraft structure.

4. Prediction of fatigue life to crack initiation.

Modeling is prepared in CATIA V5 and analysis done in MSC/PATRAN and MSC/NASTRAN.

3.2 Geometry Of The Emergency Exit Door Cut Out

Figure 3.1Geometry of the emergency exit door cut out



Figure 3.2 An isometric view of the emergency exit door cut out

Figure 3.1 shows geometry of the emergency exit door cut out and figure 3.2shows an

isometric view of the emergency exit door cut out. The assembly consists of a bottom plate, a Z-

stiffener, a cut out, a cut out web and rivet holes. Few analysis iterations are carried out to optimize

the design of the emergency exit door cut out.

3.3.2 Aircraft Cabin Pressurization

Airplanes are flown at high elevations for two reasons. Initial, an airplane flown at high

elevation expends less fuel for a given velocity than it accomplishes for a similar speed at a lower

height in light of the fact that the flying machine is more proficient at a high height. Second, awful

climate and turbulence might be evaded by flying in generally smooth air over the tempests.

Numerous advanced air ship are being intended to work at high heights, exploiting that condition. So

as to fly at higher heights, the flying machine must be pressurized. It is vital for pilots who fly these

flying machine to be comfortable with the fundamental working standards.

Figure 3.3 High performance airplane pressurization systems

A lodge pressurization framework regularly keeps up an internal compression elevation of

roughly 8,000 feet at the greatest planned cruising height of an air ship. This counteracts fast changes

of lodge height that might be awkward or make harm travelers and team. What's more, the

pressurization framework allows a sensibly quick trade of air from within to the outside of the lodge.

This is important to wipe out scents and to expel stale air.

http://www.flightlearnings.com/wp-content/uploads/2010/04/6-40.jpg



Figure 3.4 Standard atmospheric pressure chart.

Design pressure is selected as 62058 N/m2 (9 psi) by considering air pressure difference, if

critical condition aeries due to pressure control system failure.

3.3.3Load Calculation For The Emergency Exit Door Cut Out

The fuselage cabin will be subjected to internal pressurization of around 9 Psi,

1 Psi = lb

in2

= 0.4535

25.42

1 Psi =6.895 x 10-3 N/mm2 (7.029 x 10-4 kg/mm2)

9 Psi = 9 x 6.895 x 10-3

9 Psi =0.06205N/mm2(0.006326 kg/mm2)

σhoop =p x tWhere r = radius of fuselage i.e. 1500 mm

r t = thickness of the skin i.e. 2 mm

= (0.06205 x 1500)

2

Hoop stress = 46.54 N/mm2 (4.7445 kg/mm2)

Area= width of the emergency exit door cut out x skin thickness emergency exit door cut out

= 1000 x 2

Area = 2000 mm2

Stress = Load/Area

46.54= Load/2000

Load =93080 N(9489 kg)

This 93080 N load is converted into uniformly distributed load (UDL).

This UDL is applied at the one side of bottom skin.

UDL = 93080/width of bottom skin

=93080/ 1000

UDL=93.08 N/mm (9.49 kg/mm)

http://www.flightlearnings.com/wp-content/uploads/2010/04/6-41.jpg



3.3.4. Boundary Conditions

Kinematic boundary conditions are usually applied by suppressing / releasing the required

degree of freedom. The motion of each element is represented by three displacements Tx, Ty, Tz and

three rotations Rx, Ry, Rz. Each direction of motion is termed as degree of freedom.

In this project

1) Along the bottom plate all the six degree of freedom is constrained.

2) All the elements along the thickness direction is constrained to avoid the effect the buckling.

3.4 Material Used

The material considered for the structure is Aluminum Alloy – 2024-T351,

With the following properties.[7]

1. Young’s Modulus, E = 70,000 N/mm2

2. Poison's Ratio, μ = 0.3

3. Ultimate Tensile Strength, бu = 420 N/mm2

4. Yield Stress, бy = 350 N/mm2

The following table shows the composition of the material considered.

COMPOSITION Wt. % COMPOSITION Wt. %

Al 90.7-94.7 Mn 0.3-0.9

Cr Max. 0.1 Si Max. 0.5

Cu 3.8-4.9 Ti Max. 0.15

Fe Max. 0.5 Zn Max. 0.25

Mg 5.2-5.8 Others Max. 0.15

Table 3.1 composition of the Aluminum Alloy – 2024-T351

IV. FINITE ELEMENT ANALYSES (FEA)

4.1 Introduction

The FEM is a numerical method for taking care of issues which are portrayed by halfway

differential conditions or can be planned as utilitarian minimization. An area of intrigue is spoken to

as a get together of limited components. Approximating capacities in limited components are

resolved as far as nodal estimations of a physical field which is looked for. A persistent physical

issue is changed into a discretized limited component issue with obscure nodal values. For a direct

issue, an arrangement of straight mathematical conditions ought to be

With finite element method, modeling is critical because it establishes the structural locations where

stresses are evaluated, thus:

Modeling inadequacies include the incorrect placement of elements and attempting to define a

structure with an insufficient number of elements. Such errors can be avoided by anticipating areas

of maximum strain, but doing so requires engineering experience. In most cases, the finer the grid,

the more accurate the results. However, the computer capacity, time required and cost of analysis

increases with the numbers of elements used in the model. The efficiency can be increased by

concentrating elements in the interested areas of high stress while minimizing the number of

elements in low stress areas

It is not uncommon to develop finite element method for prototype design for which

experimental data can be obtained. Strain gauging is probably the most common method of obtaining

experimental data in structural tests. Once finite element method results and experimental data have

been correlated, design modifications can be made, and these subsequent changes are often tested

through finite element method before being implemented on the actual prototype.

Finite element method is useful in design work, such as structural repair or modification,

where a structural beef-up or change is contemplated. A finite element method baseline model can be



made for an existing structure for which stress and deflection data are known. A comparison is made

between finite element method results and known experimental data to calibrate the finite element

method results. Proposed design modifications can then be evaluated to the baseline modeling

knowing that the new finite element method results will have the same accuracy and requires the

same calibration as the baseline case.

V. RESULTS AND DISCUSSIONS

5.1 Stress Analysis of The Emergency Exit Door Cut Out

In stress analysis the meshing of CATIA model of emergency exit door cut out is done with the MSC

PATRAN and solved with MSC NASTRAN. In this section the maximum stress developed in the

panel is found out.The stress distribution around the rivet hole is captured through the repetitive

analysis by fine meshing near the rivet hole region.

5.1.1Finite Element Model Of The Emergency Exit Door Cut Out

As we have generated emergency exit door cut out model using preprocessor, then the model

is checked only for any overlapping of surfaces. This breaking up of surfaces helps in meshing of

emergency exit door cut out finely and gives good results at all points.

Figure 5.1Meshed And Assembled Emergency Exit Door Cut

FE model of the emergency exit door cut out is as shown in Figure 4.2 Meshing is done by

using QUAD and TRIA shell elements. To provide connectivity TRIA shell elements are used. Care

is taken such that less number of TRIA elements are used, as it has got more stiffness compared to

the QUAD element.

Fine meshing is done at the locations where the stress concentration is occurring, cores

meshing is done at rest of the locations. At cut out section riveting is done to connect to the bottom

skin.

Verification for the boundary, duplicates is carried out. Normal for each element is assigned.

Finely material properties are given to the model. A significant part of the analyst’s time is devoted

to describing the geometry of the model.

5.1.2Finite Element Model Summary

Total number of elements in the model = 36496

Total number of nodes in the model =36064

Total number of Quad elements =33369

Total number of tria elements =3126

5.1.3Quality Criteria For Elements

Once after meshing and before proceeding further, we check for elements so that no elements

fail during the application of the load. The elements which failed under the specific condition was

checked and modified so that all the elements fell under the pre-defined standards and hence we

ended up with a mesh where none of the elements failed.



The elements were checked for the following criteria:

Figure 5.2Quality Criteria For Elements

5.1.4 Loads And Boundary Conditions

Figure 5.3 Loads And Boundary Condition

5.2 Analysis

5.2.1. Displacements

The behavior of the structure is characterized by its displacements. If the displacement of the

structure is known, other important data can be calculated from it. The linear static analysis uses a

small displacement formulation in which the finite element equilibrium equations are written with

respect to the initial configuration of the structure.



Figure 5.4 Displacement Contour Of Emergency Exit Door Cut Out

The above figure 6.12 shows the displacement contour of emergency exit door cut out. The

maximum displacement is seen at the free end that too in the bottom skin. The maximum

displacement seen is 1.67 mm. Due to the presence of the stiffeners on either side less displacement

is observed.

6.1.7.2 Stress Contour

From the stress analysis of the emergency exit door cut out the maximum tensile stress

location is identified. This maximum stress is observed at the rivet hole location. Linear static stress

displayed in the software is shown in the Figure 6.13. Maximum stress displayed in the rivet location

is shown in the Figure 6.14.



Figure 5.5 Stress Contour Of Emergency Exit Door Cut Out

Figure 5.6 Closer View Of Maximum Stress Contour

The above figure shows the stress contour of emergency exit door cut out. The maximum

magnitude of stress in red colour and the dark blue colour showing minimum stress distribution.

Stress is maximum due to stress concentration resulting from the presence of rivet hole. The

maximum stress is observed from the analysis is 26.4 kg/mm2 (258.98 N/mm2).The stress

distribution around the rivet hole is captured through the repetitive analysis by fine meshing near the

rivet hole region.

Each elemental stress at the rivet hole is shown belowfigure 6.15. The maximum stress is

seen at the outterside of the rivet hole and it decrease inward.

The maximum stress developed from the analysis is 258.98 N/mm2 (26.4 kg/mm2).Which is less

than the ultimate stress 420 N/mm2(42.81 kg/mm2)

Hence the structure is safe.



REFERENCES 1. A.K.Vasudevan, K. Sadananda, G. Glinka (2001) focuses the attention on Critical parameters for fatigue damage.

2. R P G Richard Muller (1993) focuses the attention on fatigue crack initiation in riveted lap joints and in pressurized

fuselage.

3. H. Hosseini-Toudeshky, B. Mohammadi(2006)focuses the attention on A simple method to calculate the crack

growth life of adhesively repaired aluminum panels.

4. Jakub Kaczkowski Wojciech Potkanski focuses the attention on Analytical Support in Aircraft Certification.

5. Stevan Maksimović(2005) focuses the attention on Fatigue Life Analysis of Aircraft Structural Components.

Documents

PREDICTION OF CRACK AND STRESS ANALYSIS … OF CRACK AND STRESS ANALYSIS OF A CUT ... ABHILASH V J1, LOHITESH JAGA KUMAR2, Dr. IRFAN G3, ... Figure 3.4 Standard atmospheric pressure