Fatigue analysis optimization.

Lozici-Brinzei, Dorin ; Baran, Daniela ; Tataru, Simion 等

1. INTRODUCTION

The elementary steps in the fatigue analysis and damage tolerance evaluation are:

* Define the Aircraft Missions

* Develop the global load spectra

* Select critical locations for each Principal Structural Elements (PSE)

* Calculate nominal stress levels for PSEs and local stress levels at critical locations

* Calculate fatigue life and Margin of Safety

One of the most relevant questions arising now is about how well we are now endowed with the tools and knowledge to deal with the above steps.

If S-N data are available the Miner rule may be adopted to calculate the fatigue life under spectrum loading.

Figure 1 show the mission profile and stress distribution for IAR-99.

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Figure 2 shows the mission profile and stress distribution for civil aircrafts.

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2. GLOBAL LOAD SPECTRUM

Figure 3 shows several associated load spectrum examples [***CS-32,2003].

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3. FEM ANALYSIS

The stress analysis is performed to determine the stress distribution within a component, and usually involves relatively detailed models of airframe sub-components.

The principal normal stresses and maximum shear stress together with the angle of the principal axis can be determined from the applied stresses (fx, fy and fs) using the following equations [Bruhn, 1973], [Niu, 2005]:

[MATHEMATICAL EXPRESSION NOT REPRODUCIBLE IN ASCII]

The effective stress for von Mises is expressed as [Petre, 1984]:

[??] = [square root of 1/2 [[([[sigma].sub.x] - [[sigma].sub.y]).sup.2] + [([[sigma].sub.y] + [[sigma].sub.z]).sup.2] + [([[sigma].sub.z] - [[sigma].sub.x]).sup.2]] + 3([[tau].sup.2.sub.xy] + [[tau].sup.2.sub.yz] + [[tau].sup.2.sub.zx])]

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4. FATIGUE LIFE CALCULATION

Constant and variable amplitude loading may be considered in calculating fatigue life. By using an SN curve, designers can calculate the number of such cycles rapidly leading to the component failure.

This theory also assumes that the damage caused by a stress cycle is independent of where it occurs in the load history, and that the rate of the damage accumulation is independent of the stress level.

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The result, or Damage (D), is expressed as a fraction of the failure. The component failure occurs when D = 1.0, so, if D = 0.85 then 85% of the component's life has been consumed.

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5. CONCLUSIONS

The tools and approaches discussed in this review can help designers to improve the component safety while reducing over-engineered, heavy, and costly designs.

FEA provides excellent tools for studying fatigue by the SN approach, because the input consists of a linear elastic stress field, and FEA allows consideration of the possible interactions of multiple load cases. Because of its ease of implementation and the large amounts of available material data, the most commonly used method is SN. By making use of today's technology to avoid fatigue, catastrophes can often be prevented.

6. REFERENCES

*** (2003).CS-23 Certification Specifications for Normal, Utility, Aerobatic, and Commuter Category Aero planes, 14 Nov.

*** (2008), MIL HDBK 5J, Metallic Materials and Elements for Aerospace Vehicle Structures

Bruhn, E. F. (1973), Analysis and Design of Flight Vehicle Structures, Tri-State Offset Co, Cincinaty, ISBN-13: 978-0961523404

Niu, M. (2005), Airframe Stress Analysis and Sizing, Conmilet Press Ltd, 2nd Edition, New York, ISBN-13: 978-9627128083

Roark, R.J., Young, W.C. (2001), Formulas for Stress & Strains, MCGRAW-HILL International, 7th Edition, New York, ISBN: 978-0-07-072542-3

Petre A (1984), Calculul structurilor de aviatie, Editura Tehnica, Bucuresti.

Peterson, R.E., (2008), Stress Concentration Factors, John Wiley & Sons, 3rd Edition, San Francisco, ISBN 978-0-0470-04824-5