Skin reinforcement for Ellipse Cut-out

[HSThompson]

Skin reinforcement / Ellipse Cut-out
I have monitored this forum for quite awhile and this is my first post so I will try to be as clear as possible.

Situation is:

I have a lower wing skin with a fuel bay cut-out shaped like an ellipse. It has an integral chem-milled cut-out reinforcement.

The case I have is that I have locally blended out corrosion on the integral stiffener reducing it's thickness. To numerically substantiate the repair I have chosen to use a unit shear-flow in the before and after picture and determine the delta in the margin of safety.

The problem is:

There are plenty of analysis methods for ring (circular) and rectangular reinforced cut-outs in webs, but nothing on the ellipse. (i.e. Red Niu(chap. 13), Flabel (chap. 9), Bruhn D3.7))

My proposition / question is:

If I treat the reinforcing ring as a column with 15t web thickness, (addition of effective web), and use the unit shear flow to determine buckling/crippling in the before and after,.. is this accurate?

My back-up is to just do a FEM, but it seems to me that there should be a simple, logical correlation here???

Your discussion and help is greatly appreciated.

H. S. Thompson

 

[Sparweb]

How about treating it like a rectangular cut-out, and the radius of the corners equal to the local radius of the ellipse (at the middle of the blend).

I'm just speculating... I'd be more comfortable with specific data, too. I remember seeing elliptical cut-outs in Niu's green book, but I don't have it with me now. I will get back to you shortly.


STF

 

[aviat]

Often those panels vary in thickness from root to tip and even though they are interchangeable physical, they are not structurally. The reinforcement thickness could also vary with location on wing.

However, that is a common corrosion site, and there is usually a limitation written somewhere. It may be in your SRM or if it is an aged A/C in a Supplemental Inspection Documents (SID), a Corrosion Prevention and Control Program (CPCP) document or a SB/AD etc.

 

[Sparweb]

Niu's green book (Chapter 6) covers cut-outs, and makes some mention of elliptical cut-outs. He gives a K factor based on a particular a/b ratio, but if you read further you might come up with something.

Are you sure it's elliptical? I've seen cheaters use four arcs to get the same effect.


STF

 

[HSThompson]

For SparWeb:
Yeah, It looks about as elliptical as you can get.

I will be checking Green Niu for further information.

For aviat:
I fully understand what your saying. Unfortunately this is a rather new a/c, <400hrs. I am the stress guy that defines what the limits are and I need a clarification of the limits for the following reason. The previous &quot;stress guy&quot; defined the limits based on the entire door land being blended to the same remaining thickness. His critical cased was fuel pressure, &quot;slosh&quot;, and blowout of the door. In my situation I have locally, 1.0&quot;, blended beyond his limits. I don't believe that I have the same failure mode as he did, I believe I potentially have buckling of the wing skin bay do to reduced cut-out reinforcement effectiveness.
I need to substantiate this for the possibility that FAA asks; &quot;hey what gives, before you guys said this was the limit, now your saying this is the limit&quot;? I am confident in the repair, I just need to have my ducks in a row, (CYA)


Thanks guys!
HST

 

[Sparweb]

Any reasonable engineer (FAA or otherwise) would follow your argument as long as he's comfortable with the supporting data.

Is it correct to assume that a doubler is out of the question? This is a &quot;wet wing&quot;, I understand?

Given the apparent speed of the corrosion, I wonder if this is potentially a chronic problem for the type.


STF

 

[HSThompson]

SparWeb,

No I wouldn't say an internal doubler is out of the quesiton, but I just don't believe it is necessary. The reason I justify that is tha the wing was tested to 160% of limit load. That being the critical case of upbending, which would give peak tension and shear flow around the skin panel. The justification is then the standard statement; &quot;well, I have removed less than 6% material and my MS therefore is still positive&quot;.

The problem then becomes: &quot;well, that's the logic, but where's the math?&quot;

As to your other questions, yes this is a wet wing, and yes it is a crappy design and the model does have a corrosion problem.

You might be able to tell that I am extremly irritated with our designers and the original stress folks that signed off on this.

To give you another example. In another location on the wing there is a stack-up of; Titanium fitting, Stainless splice plate, Aluminum Skin with a Cad plate hi-lock. The joke is that will just solder a wire to the fitting and call it our back-up electrical power supply.

 

[HSThompson]

Actually guys, I think I have an inspiration of sorts...

I think I will analyze this as a circular cut-out with the diameter equal to the major axis of the ellipse.

My justification is two fold:

1. The direction of the ellipse major axis is parallel to the axial tension in the skin. This results in a Kt of 3 for an ellipse vs. Kt of 2 for a circle. That makes the analysis more conservative.

2. By treating it as a circle with diameter equivalent to the major axis of the ellipse I am reducing the amount of structure. Again, more conservative.

what do you think.

 

[Sparweb]

Given what's in Niu's green book, I'd say that's plenty conservative. If it gives you a MS>0 result, fine. I would emphasize that you should beg, borrow, or steal a copy of Niu's Aircraft Structural Design, to have a more authoritative source.


STF

 

[edbgtr]

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 &quot;framed&quot; reinforcement around the hole, not a
thickening (pad-up) locally, like a donut integrally machined doubler) In the &quot;framed&quot; 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 &quot;blade&quot; 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 &quot;crisis condition&quot; 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 &quot;reasonable&quot; 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 &quot;working stresses&quot; 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 &quot;over-designed&quot;. 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 &quot;modern&quot; 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 &quot;linear&quot; material is removed then the margin is reduced to 0.7%, which may be OK, but is marginal. For a &quot;non-linear&quot; 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.

 

[HSThompson]

Ed and all,

I apologize for not responding sooner, recently had a death in the family and things are crazy.

Anyway, Yes I do have access to all the stress notes, test data, report, etc. since I am the OEM. (Note: This is an upper mid-sized business jet.)

To restate here: while the previous analysis case, submitted to the FAA via 8110-3, for minimum blend was for the fuel &quot;slosh&quot; load case, I do not consider this to be the critical case for the current situation. The reason I say this is that I am only locally blending in a 1&quot; area.

Again, I believe that critical load case on the lower wing skin fuel door cut-out to be maximum upbending. Result: Bending and shear in the skin.

I agree with the 6.7% minimum margin due to the 160% of limit load test.

My intent is to0 logically substantiate this repair: Since the c/s area removed constitutes less than 6.7% of the reinforcement plus effective skin then I am ok. The cross section is roughly shown below. If you require dimensions I will be happy to provide them, however what you see is &quot;approximately&quot; full scale->



_______________
Reinforcement ?| |
/__________| +---+? Door Land
/----------------------------+ ?
Blend Area

 

[edbgtr]

HST
For the upbending case you mention, and thus bending (tension) and shear in the panel, the material blended out is &quot;linear&quot; and your margin is still positive based on the static test.
I suggest you also look at the DT/fatigue work done on the detail (if it is an SSI) and decide whether it warrants a change to the inspection plan providing for the continued airworthiness of the airframe.
Ed