Hi HST & STF
First of all, a little anecdotal information regarding chemi-milling of skins. We used French built aircraft (fighters and helicopters) that showed serious signs
of corrosion on such skins. The cause was eventually traced to the washing agent combining with the remnants of the chemicals used in the milling process. The
cure was to change the washing agent being used, as well as the OEM improving his manufacturing processes by washing away the residual products more effectively.
The dissimilar metals stack-up (as mentioned by HST) smacks of a serious lack of MSG-3 analysis in the company, w.r.t. airworthiness and DT inspection planning. This could cost the company dearly with through-life support due to poor design practices.
I have to admit, I'm a little confused by the exchange on this topic. HST seems to have access to the design load cases for the skin bay or panel in question. This does not sound like a back engineering job to determine the loads in the component(s). If he does (have the design L/Cs) then determining the delta in the MS for the critical case should be quite easy, I would have thought. A fuel slosh load case, is in my experience, a pressure load on the panel and the surrounding structure is reducing the stress on the access door or cover plate. (I'm assuming that we have a "framed" reinforcement around the hole, not a
thickening (pad-up) locally, like a donut integrally machined doubler) In the "framed" configuration, the surrounding integral stiffeners (blades or risers) will
then be bending with compression in the inner (w.r.t. the tank) fibres of the "blade" stiffeners. Determining the effective L or T section carrying the bending should have been done by the original stress engineer in the type record (check stress). I agree that changing these assumptions (effective skin/stiffener sections) now to satisfy this "crisis condition" will raise a few eyebrows with the certification authority, especially without resubmitting the original check stress to reflect same. I do agree however with STF, regarding the "reasonable" FAA official.
The discussion around the stress concentration factor for the circle and the ellipse, indicates that you wish to clear the panel for high tension loads (wing lower skin in tension) but mention is also made of the compression instability of the reduced thickness blade stiffener where the corrosion has been dressed out beyond design limits. The provision of the Kt values in Niu is to illustrate the stress raiser effect of such a cut out sans reinforcing. The stress is increased by the Kt factor very close to the hole's edge and covers a VERY small region extending radially from the hole. This causes a problem from a fatigue point of view (not a static one), and the "working stresses" may need to be reduced in an attempt to avoid crack initiation in this region. The overall panel tension
strength is replaced by the doubling around the hole (re-routing the load). This in turn lowers the stress due to the Kt effect, thereby reducing the possibility of crack initiation in that region. Remember that the Kt factors given by Niu are for pure tension only! For a more complex stress state the Kt distrubutions around the cut-outs need to be superimposed to obtain the worst combination. But again this is a fatigue consideration and not a static strength one. Net section analysis is valid for static strength.
The in-plane bending in the panel around the cut-out is due to the picture-frame bending (portal frames) of the reinforcing around the hole carrying the original shear in the panel, sans the cut-out. If the assumption has been made that the cut-out door carries no shear load, then the reinforcing usually is well "over-designed". Current design philosophy (in the interest of lighter weight design) is to allow the door (or cover plate) to carry between 20 and 40 percent of the original shear of the uncut panel. This depends on the number of fasteners around the door and the fastener (screws or bolts) tolerance fit in the holes (the ability of the fasteners to pick-up load in unison). The analysis of "modern" door cut-outs is more like the analysis of a flush patch. The high tension loads in the skin for an Nz max pos L/C will relieve the compression loads in the framing around the door resulting from the shear caused by the wing torsion at high speed, or control surface deflection. I can only deduce that the alternate load case considered by HST is for an Nz max neg L/C combined with high torsion.
Assuming that the wing bending test (160% of Limit up-bending result) was representative of the complete loading on the wing lower skin, i.e. tension and torsion, and represents the critical case for the panel, including the compression case, which is usually at lower stress levels, then an assessment of the panel (including the corrosion damage) relative to this test, should be possible. Bear in mind however, that 160% of Limit is only 6.7% higher than 150% of Limit, i.e. Ultimate. If 6% of "linear" material is removed then the margin is reduced to 0.7%, which may be OK, but is marginal. For a "non-linear" material reduction of 6%, i.e. bending or buckling, the margin will be negative by about 6%.
Without seeing all the loading conditions on the panel, and a clear description of the reinforcement around the cut-out, it is difficult to assess which L/C will become critical as a result of the reduced section. Perhaps a more detailed description of the L/Cs and the reinforcement configuration, on the panel would help us to further help HST with his problem.
Regards,
Ed.
