(Aircraft Structural Integrity Program)
The ASIP dates back to a 1950's Air Force publication on structural integrity requirements. It was known from an early stage that ASIP was a vital program in prolonging the life and ensuring the structural safety of all aircraft. Meetings began in the 1970's, but it wasn't until 1984 that it was reshaped into the current conference format. Incidents like the 1988 Aloha Flight 243 Air Disaster highlighted the importance of ASIP requirements and the contributions of the ASIP community, to preclude the recurrence of such tragedies in the future. The ASIP Conference helps to accomplish this through the personal interactions of its attendees, resulting in the exchange of vital ideas and technology.
Stehlin Engineering Contribution to ASIP Conference 2014
Mr Thierry Stehlin has been selected to hold a presentation in name of RUAG during the ASIP Conference 2014. The presentation shows an application of the use of the pyNASSIF tool to perform CG analysis in a spar of the outer wing of the Swiss F/A-18 aircraft.
--------------------- Presentation Abstract ---------------------
From 2003 to 2005, RUAG Aviation performed a Full Scale Fatigue Test (FSFT) of the F/A-18 fighter aircraft in order to validate the Swiss-specific structural modifications implemented during the procurement phase of this aircraft.
The tear down inspection at the end of the FSFT revealed a 16in long crack in the outer wing spar 3 in spanwise direction along the web to bottom flange interface (about 1/4 of spar severed). The crack growth was not obvious to explain and no standard model was applicable to such a case. The use of pyNASSIF script combined with the R-Curve approach and AFGROW runs permitted to:
<![if !supportLists]>1. <![endif]>Understand the crack path and the final cracks lengths found in the test article. Note that all crack growth (CG) modes [I, II and III] were active at different stages of the CG and a method to assess the crack path is shown.
<![if !supportLists]>2. <![endif]>Explain partial static failure followed by subsequent crack growth found by quantitative fractography (QF).
<![if !supportLists]>3. <![endif]>Develop a total life model for fleet support.
The method permits to calculate the stress intensity factor (K) of up to 2 thru cracks fronts in virtually any thin structures. It automatically accounts for load redistribution between the parts and non-linear behaviour can also be taken into account if required. The cracks are implemented in a NASTRAN FEM model by unzipping the nodes along the crack path.
The process of running the model at several cracks sizes and extracting the necessary information (cracks lengths, crack increments, material properties, plate thickness and elastic-energy) to calculate the stress intensity is automatized in pyNASSIF, which main output is K in function of the crack length. In cases of two crack fronts the output is a set of two K matrices, one for each crack tip, with dependence on tip 1 and tip 2 lengths. K solutions can then be imported into a standard software to perform CG calculations.
<![if !supportLists]>Ø <![endif]>Introduction to RUAG Aviation.
<![if !supportLists]>Ø <![endif]>Introduction to the Swiss F/A-18 Full Scale Fatigue Test (FSFT).
<![if !supportLists]>Ø <![endif]>Presentation of one of the cracks discovered in the outer wing at the end of the FSFT.
<![if !supportLists]>Ø <![endif]>Description of pyNASSIF (theoretical background & implementation).
<![if !supportLists]>Ø <![endif]>Validation of pyNASSIF by comparing the results for standard solutions with other sources like NASGRO, AFGROW and FRANC2D.
<![if !supportLists]>Ø <![endif]>Application of pyNASSIF to F/A-18 outer wing spar 3 crack as part of total life model development.
<![if !supportLists]>Ø <![endif]>Comparison of the analytical results with the FSFT findings (strain gages and QF information).
<![if !supportLists]>Ø <![endif]>Impact on Swiss F/A-18 fleet management.
Conclusions and significance:
<![if !supportLists]>Ø <![endif]>A new application of the elastic energy release rate concept is presented.
<![if !supportLists]>Ø <![endif]>The method is straightforward and powerful.
<![if !supportLists]>Ø <![endif]>The method is validated against other sources.
<![if !supportLists]>Ø <![endif]>The method can be used to solve very complex cases, such as the F/A-18 outer wing spar crack.
<![if !supportLists]>Ø <![endif]>The method can be easily implemented by any company provided that a FEM solver is available.