Overview of one fatigue analysis approach for large aircraft | CDGudas
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Big Planes

General Overview of One Analysis Case for Production

For the last generation of the 747 transport aircraft, the fuselage was increased by 220 inch [5.6 m], with 160 inch in the so-called Section 42 and 60 inches aft of the wing in Section 46.   As result of the increased fuselage bending moments, initiation and propagation of fatigue cracks growing from the corners of the second door needed assessment and preventative measures that accounted for the several missions profiles of this aircraft.

The sizing of the aircraft was based on a “Loads Model” containing the average loads in a bay by bay arrangement, with a bay being generally defined between two consecutive frames and two adjacent stringers or at the middle of the uni dimensional elements (frame caps and stringers).

With the approximate dimensions for Door 2 of 54 inch in width and 82 inch in height, at the scale of the aircraft, the deformations of the door edges was assessed through thin markers (0.001 in2 sections area and small stiffness)

For the fatigue relevant loading situations, defined through over 400 primary external loads scenarios that include the ground operations, initial and final climbs, gust and manoeuvres and the descent portions of the flight, very localised stresses are required at the corners of the door and at the fastener holes.  Fastener loads are also required .

Section 42 Superelemet

(stingers & floor beams not shown)

Entire aircraft with new plugs highlighted
Stringers layout in Section 42

Top corner:       Location of Section 42 and added plugs – click side image to enlarge
Lower corner:   Overview of layout for fuselage reinforcing elements and floor beam arrangement

From Global Loads To Local Stresses

In order to estimate the edge stresses at the door corners a more detailed model of the door area was generated.   Several parts which were lumped in the skins or frames in the global model needed to be represented explicitly, especially that some of them are sandwiched between other parts where cracks are difficult to notice by visual or dye-penetrant inspections.

All the newly detailed parts were positioned at the layer of adjacent fuselage skins and the connecting fasteners were modelled with zero length CBUSH elements.

In order to establish the stresses along the edges of the door, very thin bar elements were added between consecutive nodes along the edges.   The material properties for these bars were identical with the material properties of the skins they bordered.   Depending on location, the length of these elements varied between 0.05 and 0.2 inch.   A similar approach was used for the other internal parts ending at or near the edges of the door cut-out.

In this new discretizatrion, the sub-model was solved using MSC Nastran solution 101 for all fatigue relevant load cases and the deformations at the door markers were compared against their values in  baseline configuration to ensure that any differences are within acceptable limits.  The stress data from the bar elements was next processed to generate the the fatigue stress spectra for the critical mission profile for different locations and the fatigue strength at the cut-out was conducted using company’s methods  and bespoke software.   The specific methodology and the details used during post-processing are now proprietary information of the Boeing Commercial Aircraft company.

Door 2 area in baseline configuration

Same area with refined disctretization

Task Related Details From Production Configuration

Currently two external doublers, at top-left and lower-right corners are being used to reinforce the edges of the fuselage skin at Door 2.  A process similar to that described above was used for the analysis of all the parts at this locations.

Supplementary, load transfer values were established at several dozen new fasteners and these locations were added to the evaluation of fatigue strength.   Crack growth analysis was conducted afterwards with using proprietary analysis methods and software.

Internal structure at Door 2

A

Door 2 straps

B

Edge stress captors

C

Doubler at upper corner

D

Doubler at lower corner

E

A. Door 2 analysis area looking from inside
B. Door 2 reinforcing straps (the upper and the lower straps overlap along 30% of their vertical branches)
C. Strain measuring bars along the edges
D. Upper doubler and fasteners
E. Lower doubler and fasteners

From Other Transport and Commuter Category Aircraft

Sizing of reinforcements for rectangular cut-out in the fuselage of Bombardier G 6000 (in the fuselage crown near the aft pressure bulkhead),
Determination of interface loads at the frames in Section 44 (above the wing) and Section 46 for the installation of Greenpoint Technologies relaxation units in the crown of 747-8,
Verification of the trailing edge panels (composite structure) and the flaps tracks for B 777 F,
Internal structure (spars & ribs) for the wing tip and fences of A 380,
Sizing of reinforcing structure for 14 in circular cutout, located inside the fuselage, aft of main spar (Beach 350)

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