Analytical Evaluation of the Wing Box Splice Joint for Static & Fatigue Loading

Analytical Evaluation of the Wing Box Splice Joint for Static &

Fatigue Loading

G Krishnaveni

, G Kavitha

, E.S.Elumalai

D Dominic Xavier

R Sarath kumar

1, 2, 3,4,5 Asst. professor, school of Aeronautical sciences, Hindustan institute of

technology and science

[email protected]

Abstract:

Aircraft is symbol of a high performance mechanical structure with high structural safety record. Safety and the structural weight are important parameters to be considered during the design phase. Analysis is carried out to find all of the theoretical stresses within and predict failure due to unknown stresses by showing stress concentrated areas in the material. Further, fatigue life to crack initiation can also be estimated by local stress strain approach. Based on the analysis the structural design, sizing and optimization is performed to meet the best feasible design. This ensures the safety of an aircraft keeping in account its structural weight. Analysis of the wing box is carried out for static and fatigue loading using finite element analysis

Keywords: wing box, Nastran, Patran, Fatigue, Splice Joint

1.

INTRODUCTION :

Several analysis methods are presently available for fatigue life evaluations. The process can be broadly classified as either “crack initiation” or “crack propagation approaches”. In the recent years it has been recognized that the fatigue failure process involves three phases. A crack initiation phase occurs first, followed by a crack propagation phase;

finally, when the crack reaches a critical size, the final phase of unstable rapid crack growth to fracture components, the failure process. The modelling of each of these phases has been under intense scrutiny, but the models have not yet been developed in a coordinated way to provide a widely accepted engineering design tool. Fatigue design against crack initiation may lead to different material selection criteria and structural design from fatigue design against crack propagation. The aim of the study is to define a complete procedure for fatigue life prediction of structural elements up to crack initiation.

The procedure for fatigue life estimation is based on combining computation stress analysis with strain-life methods. Methods for stress analysis that will be used here are analytical and FE method. The analytical method proved to be easier while FEM is adequate for application with complex structures. FEM is favourable fo

r detection of critical locations.

2 METHODOLOGY

The work involves analysing the wing box for static and fatigue loading using finite element analysis and determining the magnitude and location of the maximum stressed developed. Distribution of fasteners loads and local stress field at rivet locations will be studied using finite element analysis. The work also involves the modifications required to correct the boundary effects of the panel. Further the fatigue life to crack initiation is evaluated under a realistic service load spectrum using a local stress strain approach.

Stress distribution and deformation in the wing box is identified with the help of analysis

results. Areas with high stress concentration are also found. Repeated finite element analysis will be carried out to get the response of the parent structure (wing) at the joint location. The response of the splice joint will be evaluated. The splice joint is one of the critical locations for fatigue crack to initiate. In this project prediction of fatigue life for crack initiation will be carried out at maximum stress location. A wing box of a twelve seater aircraft is analysed for the given loading condition and stresses are estimated. Areas with concentrated stresses are identified. Later, calculations are made to estimate the fatigue life to crack initiation. Wing box of this aircraft acts like a cantilever beam. To simulate the same condition, one edge of the box is fixed and load is applied to the other 3. GEOMETRIC SPECIFICATION:

The wing box consists of two essential parts, the internal wing structure, consisting of spars, ribs and stringers, and the external wing, which is the skin. In this section all the components of the wing box are briefly explained and their dimensions are specified. The figure below illustrates the CATIA model of the wing box analyzed

.

Figure 1: catia model of wingbox and rivet

Figure 2: Meshed wing box

Thickness of all the components is 3 mm other than top and bottom skin being 2 mm in thickness. All the components of the wing box are joined to each other using rivets with diameter 5mm.

4 MATERIAL SPECIFICATIONS

3. Modulus of Elasticity: 72 GPa 4. Poisson’s Ratio: 0.33

5 LOAD SPECIFICATION

Uniformly distributed load is applied at the other end of the structure. The load to be applied on the wing box is calculated as below.

1. The all-up weight of the aircraft (12-seater) is 6000 kg Hence the weight of the aircraft = 6000 kg.

• Design load factor considered = 3.2 g.

• Total load acting on the aircraft = 6,000 = 19,200 kg-f

• Factor of safety considered = 1.5

• The design load = 19,200×1.5= 28,800 kg-f

• Lift load is experienced by both fuselage and wing.

• Lift load on the wing = 80% of total load =0.8×28,800= 23,040 kg-f

• Load acting on each wing = 0.5×23,040 = 11,520 kg

• Total span of the wing = 9500 mm

Figure 3: Rough sketch of loaded wing

Coloured region of the figure shows the loaded wing box. To avoid the boundary effects, Rib 1 is fixed and Rib 2 is loaded.

Bending moment at the section A-A(Rib1)towards the root = 11,520×(300+1366) = 19192320 kg-mm

• The transverse load to be applied at the free end of the box to create the same bending moment at the other end of the box = 19192320/1229 = 15,616.21 kg-f

• L=1500 mm ; B=950 mm; D=404 mm

• The uniformly distributed load can be calculated as

15616.2083/3744 = 4.17099 N/mm

Loads and boundary conditions are set. Wings of an aircraft act like a cantilever beam.

Cantilever refers to a loading condition where one end is fixed and the other end is loaded vertically. Under such loading condition wings tend to bend. Fixing the ends refer to constraining all the degrees of freedom. Hence, to fix the wing box at one end, rotation and translation in all the directions are restricted. Lastly analysis is carried out to find out the stress distribution

.

6 RESULTS AND DISCUSSION:

The stress distribution on wing box is obtained. The maximum stresses are observed near the splice joint of a wing box. The magnitude of stress is 45.4 kg/mm^2.

Figure 5: maximum stress developed

Figure 6: Concentrated stress near splice joint

Figure 7: stress distribution in wing box

6.1 FATIGUE CALCULATION

After the magnitude and location is found, we proceed to fatigue calculations.These calculations are made to estimate fatigue life to crack initiaton.Load is applied to wing box corresponding to 4.8 g condition i.e (3.2g * 1.5g).

Stress generated corresponding to 4.8 g = 45.4 n/mm^2.

hence for 1g condition the stress developed= 45.4/4.8 =9.45kg/mm2

Table 1 Stress corresponding to different g conditions

G Corresponding stress

0 0

0.5 4.9583

0.75 7.4375

1 9.45

1.25 12.3958

1.50 14.875

1.75 17.3542

2 19.8333

A realistic service load spectrum is used for finding the number of cycles; an aircraft undergoes in particular g condition. This data is used in further calculations. The load spectrum of the aircraft under analysis is tabulated below

Table 2 Number of cycle corresponding to different g conditions

G condition No. Of cycles (ni)

0.5 g to 0.75 g 48,000

0.75 g to 1 g 33,000

1 g to 1.25 g 26,000

1.25 g to 1.50 g 22,000

0 g to 1.75 g 45

0 g to 2 g 1

-0.5 g to 1.5 g GAG (ground air ground)

The values of stress ratio and stress amplitude corresponding to different g ranges are tabulated below.

Table 3 stress amplitude and stress ratio

Magnitude of the factors affecting fatigue is listed below

• Surface roughness : 0.8

• Types of loading : 1

• Stress concentration factor : 1

• Mean stress effect : 0.5

• Reliability in design : 0.897 total factor = 0.8*1*1*0.5*0.897 = 0.3588

Stress amplitude and stress ratio is calculated again considering the factors affecting fatigue. These calculations are tabulate below

Table 4 stress amplitude and stress ratio G range Stress (min) ,

σ1 Stress (max) ,

σ2 R = σ1 / σ2 σa=(σ2-σ1)/2 σa (ksi)

0.5 – 0.75 4.95833 7.4375 0.667 1.2395 1.735

0.75 – 1 7.4375 9.9167 0.75 1.2395 1.735

1 – 1.25 9.9166 12.3958 0.8 1.2395 1.735

1.25 – 1.5 12.3958 14.875 0.83 1.25 1.75

0 – 1.75 0 17.3541 0 8.677 12.147

0 – 2 0 19.8333 0 9.9165 13.883

-0.5 – 1.5 4.9583 14.875 -0.333 9.916 13.882

G range σ1 σ2 R=σ1/σ2 σa=(σ2-

σ1)/2

σa ( ksi)

0.5- 0.75 13.818 20.727 0.667 3.454 4.836

0.75-1 20.727 27.636 0.75 3.454 4.836

1- 1.25 27.636 34.545 0.8 3.454 4.836

1.25-1.5 34.5545 41.457 0.833 3.483 4.877

0-1.75 0 48.366 0 24.183 33.856

0-2 0 55.275 0 27.637 38.691

6.1.a S- N curve for aluminium

In fatigue testing, a curve showing the relation between the value of stress and the number of cycles at that value of stress required to produce failure in the test specimen is known as S N curve. the figure below shows the S N curve for aluminium. Damage caused is now evaluated using this curve. Ni refers to the number of induced cycles and nf

refers to the number of cycles leading to fatigue. Value of Induced number of cycles is provided by the manufacturer whereas the number of cycles leading to fatigue is obtained by the SN curve.Ratio of number of induced cycles to that of number of cycles leading to fatigue gives the value of damage occurred.

When the ratio reaches 1, crack is said to be initiated. Corresponding to 100 flight hours, the damage in the wing box is calculated as below

∑ (ni/nf) = 0 + 0 + 0 + 0+ (45/7*10⁴) +(1/1.75*10⁴) +( 100/3.25*10⁴) = 0.0037

Figure 8. S.N Curve for Aluminium

7 CONCLUSION

In order to improve the aircraft life and make it defect free, stress analysis and fatigue life prediction is carried out on aircraft structure through FEM approach. A FEM approach is followed by the stress analysis of aircraft structure. Wings of an aircraft are one of the most important structures. They generate lift, hold the aircraft and hence Stress analysis is carried on structure to identify maximum stress location in the structure. Depending on the values of stress developed design, weight, performance manufacturing, life, maintenance schedule are defined.The wing box of the twelve seater aircraft considered develops a maximum stress of 45.4 kg/mm². This stress concentration is located in the splice joint of the wing box and is around the rivets.The reasons for this stress concentration are estimated as follow

 The skin is not continuous; two sheets of metals are spliced together. The structure at this location is therefore highly vulnerable.

 The structure at this location is riveted. This is similar to plate with a hole where

Thickness can be increased around these rivets to reduce the concentration of stresses.

Further the damage caused in hundred flights is calculated. This aircraft develops a very little damage of 0.37. Hence it can be concluded that the in 100 flights the crack is not initiated.

REFERENCE

[1] Adarsh Adeppa, Patil M S and Girish K E (2012), “Stress Analysis and Fatigue Life Prediction for Splice Joint in an Aircraft Fuselage Through an FEM Approach”, International Journal of Engineering and Innovative Technology (IJEIT), Vol. 1,No. 4, pp.

142-144.

[2] Amarendra Atre (2006), “A Finite Element and Experimental Investigation on the Fatigue of Riveted Lap Joints in Aircraft Applications”, May, Dissertation, Georgia Institute of Technology.

[3] Bhaumik, M. Sujata(2008),Fatigue failure of aircraft components, journal of engineering failure analysis, Elsevier, Vol.15,No. 19,pp.675-694.