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Most Highly Stressed Structure Of An Airliner?  
User currently offlinefaro From Egypt, joined Aug 2007, 1630 posts, RR: 0
Posted (3 years 3 months 3 weeks ago) and read 4667 times:

What are the most highly stressed primary load-bearing structures in an airliner during T/O, climb, cruise, descent and landing? Per unit area, is the wing-body junction more or less stressed that the engine pylons-wing junction? During extreme attitude/overspeed maneouvring, what is likely to damage first, the horizontal stabiliser or the fin?

Faro


The chalice not my son
20 replies: All unread, jump to last
 
User currently offlineCitationJet From United States of America, joined Mar 2003, 2469 posts, RR: 3
Reply 1, posted (3 years 3 months 2 weeks 6 days 23 hours ago) and read 4603 times:

The structure that experiences the highest percentage of design limit (not ultimate) load on a given flight are the flaps. Flaps are typically deployed at the maximum or near maximum of the Vfe speed, which is the basis for limit load. The landing gear, fuselage, wings, and empennage do not come anywhere near design limit load (the envelope load case that occurs once in an aircraft's lifetime) on any given flight.


Boeing Flown: 701,702,703;717;720;721,722;731,732,733,734,735,737,738,739;741,742,743,744,747SP;752,753;762,763;772,773.
User currently offlineKPWMSpotter From United States of America, joined Dec 2006, 458 posts, RR: 2
Reply 2, posted (3 years 3 months 2 weeks 6 days 22 hours ago) and read 4568 times:

For all flight phases, the highest *loading* would be at the wing-to-body connection in the wing box. At this point, all of the flight loads from the wing will be transferred to the fuselage, at a bare minimum including the entire weight of the aircraft, thrust loads (for wing mounted powerplants only), and control loads from the flaps and ailerons. Additionally, this point will have the highest bending moment of any point on the wing (think of the moment as a summation of all of the lift loads across the span, multiplied by their respective moment arms, the distance from the imaginary point load to the wing root).

Depending on the aircraft type, the actual highest load point will vary, but typically the main spar of the wing will carry the largest share of the wing loading, so the most critical point would be where the main spar meets the wing box. This may vary for multi-sparred wings, but in general holds true.

Now, stress is dependent on area as well as the loading (σ = F/A). Wing spars are quite large, so even though the loads are tremendous, I doubt they carry the highest stress of any component. In fact, being so critical, wing spars and boxes are quite over-built, being able to carry their maximum flight loads plus an FAA mandated factor of safety of 1.5. Since most wing structure is constructed of aluminum rather than much stronger steel or Titanium, I'd rule out the spar/box for the most stressed structure.

For the highest stress on an individual component, my bet would be on the engine fuse pins (or equivalent structure) which connect the engines to their mountings. Engines are generally mounted to the airframe via only two or three attachment points, and these relatively small pins must carry all engine loads. Additionally, some airliners are designed with "fuse pins" which are designed to intentionally fail under extremely abnormal loads (i.e. they are designed to be intentionally near their failure point.) These connections are generally constructed of Steel or Titanium rather than Aluminum, which would seem to suggest that they are carrying greater stresses than other aluminum components. As a rough estimate, lets say a GE90 is connected to a pylon via two solid pins, ~2" in diameter. In that case, σ = 115,000lb / (2 x 3.14in^2) = 18,300 psi, well under the ultimate tensile strength of steel.

Another possible component would be the landing gear pylons. Often constructed of steel or titanium, the gear only carries their maximum loading for a small period of time, but carrying the entire weight of the aircraft plus any inertial loading through two or four small pylons must be pretty intense. Let's look at a 747-400 landing at its max landing weight of 650,000 lb. Assume a relatively tame landing of 1.5 Gs, so the force carried through the landing gear F = 975,000 lbf. Assuming the landing gear pylons to all be equal, solid cylinders of 5" diameter, σ = 975,000lb / (4 x 20in^2) = 12,800 psi. With a rough landing your value could easily double for an instantaneous load.

Now both of these estimates are pretty rough assumptions (landing gear pylons generally aren't solid, but are larger than I assumed, and I have no idea what the actual size of a fuse pin is), but I think these numbers give a rough idea that the two cases are on the same general order of magnitude. Calculating the stress on a spar attach point would be a bit more complicated, so I'm going to neglect the sample calculation for now.

Anyways, in general, for the strongest stresses I'd say look for the strongest structure. Anything which is important enough to warrant the weight of steel or titanium is probably seeing some pretty intense loads.

Quoting faro (Thread starter):
During extreme attitude/overspeed maneouvring, what is likely to damage first, the horizontal stabiliser or the fin?

I think that's a whole different can of worms...
Depends on the type of maneuver, where the flight loads are being applied, all sorts of things. In the case of China Airlines flight 006, where the aircraft was pulling out of a dive and pulled lots of vertical Gs, the horizontal stabilizer and elevator experienced the most damage. In the case of American 587, where the loads were due to an extreme yawing motion, it was the vertical tail which gave in first. I don't think there's any definite answer here...



I reject your reality and substitute my own...
User currently offlinefaro From Egypt, joined Aug 2007, 1630 posts, RR: 0
Reply 3, posted (3 years 3 months 2 weeks 6 days 21 hours ago) and read 4491 times:

Quoting CitationJet (Reply 1):
The structure that experiences the highest percentage of design limit (not ultimate) load on a given flight are the flaps. Flaps are typically deployed at the maximum or near maximum of the Vfe speed, which is the basis for limit load. The landing gear, fuselage, wings, and empennage do not come anywhere near design limit load (the envelope load case that occurs once in an aircraft's lifetime) on any given flight.

I find this very surprising and quite disturbing too. If operated so near their design limit load, and with repeated flight cycling, failure would be more likely statistically. If as is also likely this failure is asymmetric, one may lose or partially lose flight control...

Quoting KPWMSpotter (Reply 2):
Reply 2,

Thanx for the detailed reply, much appreciated.

Quoting KPWMSpotter (Reply 2):
As a rough estimate, lets say a GE90 is connected to a pylon via two solid pins, ~2" in diameter. In that case, σ = 115,000lb / (2 x 3.14in^2) = 18,300 psi, well under the ultimate tensile strength of steel.

Plus the weight of the engines/nacelle assembly too; somewhere around 19,000 lbs IIRC.

Faro



The chalice not my son
User currently offlineRoseFlyer From United States of America, joined Feb 2004, 9826 posts, RR: 52
Reply 4, posted (3 years 3 months 2 weeks 6 days 20 hours ago) and read 4424 times:

Stress is force divided by area. Structure that has higher forces is usually bigger and thus has more area.

The parts that are being talked about as having the highest loads and that are flight critical will have higher safety factors and larger areas. Safety factors for limit load range from about 1.2 to 4 depending on how critical they are for flight, the manufacturing process, the inspection techniques during manufacturing, the type of wear experienced during operation, typical abuse loads and if there is any warning prior to failure in operation.

Where you'll find the highest stress is in components that are not flight critical and have a lower safety factor. This is true in dual load path components, components with backups, and non critical components. If the component is not flight critical and has adequate warning of failure, then it will have a lower safety factor to keep the weight of the airplane down.



If you have never designed an airplane part before, let the real designers do the work!
User currently offlineCitationJet From United States of America, joined Mar 2003, 2469 posts, RR: 3
Reply 5, posted (3 years 3 months 2 weeks 6 days 20 hours ago) and read 4414 times:

Quoting faro (Reply 3):
I find this very surprising and quite disturbing too. If operated so near their design limit load, and with repeated flight cycling, failure would be more likely statistically.

That is addressed by the fatigue cyclic testing that is performed on the aircraft test article. The flight profiles and corresonding loads are applied to the aircraft structure to simulate the flight profiles (fuselage pressurization, takeoff, cruise, maneuvers, turbulence, descent, landing, taxi) for multiple lifetimes. The design is modified in those areas that exhibit cracks during fatigue testing.



Boeing Flown: 701,702,703;717;720;721,722;731,732,733,734,735,737,738,739;741,742,743,744,747SP;752,753;762,763;772,773.
User currently offlineKPWMSpotter From United States of America, joined Dec 2006, 458 posts, RR: 2
Reply 6, posted (3 years 3 months 2 weeks 6 days 19 hours ago) and read 4398 times:

Quoting RoseFlyer (Reply 4):
Where you'll find the highest stress is in components that are not flight critical and have a lower safety factor.

I wouldn't say that's necessarily true. While a flap or spoiler may be closer to its limit load, its limit load may be much lower. Most non-critical items such as fairings and spoilers are actually quite minimal in their construction, and while composites are strong, they still don't come close to the ultimate strengths of steel and titanium. Additionally, the loads carried are much, much smaller. Even at high speeds, aerodynamic loads on individual components will be much smaller than the total inertial and structural loads. Even though the parts operate closer to their limits, the limits are much smaller, therefore the total stress will generally be smaller.

I still maintain that the highest stress will occur in the components built of the strongest materials - carrying stress is what they're designed to do.



I reject your reality and substitute my own...
User currently offlineRoseFlyer From United States of America, joined Feb 2004, 9826 posts, RR: 52
Reply 7, posted (3 years 3 months 2 weeks 6 days 19 hours ago) and read 4357 times:

Quoting KPWMSpotter (Reply 6):

I still maintain that the highest stress will occur in the components built of the strongest materials - carrying stress is what they're designed to do.

I absolutely agree. I was comparing safety factors and ignoring the obvious. Components that are made of steel are likely in the most stress. Steel is rarely used in structure, but in components, it is heavily used. All the flight controls actuation equipment are steel, aluminum and titanium, but the original question was structure.

[Edited 2011-09-07 14:18:51]


If you have never designed an airplane part before, let the real designers do the work!
User currently offlinevikkyvik From United States of America, joined Jul 2003, 10350 posts, RR: 26
Reply 8, posted (3 years 3 months 2 weeks 6 days 16 hours ago) and read 4254 times:

Quoting faro (Reply 3):
I find this very surprising and quite disturbing too. If operated so near their design limit load, and with repeated flight cycling, failure would be more likely statistically. If as is also likely this failure is asymmetric, one may lose or partially lose flight control...

From http://rgl.faa.gov/Regulatory_and_Gu...C70B685256672004E98BB?OpenDocument :

Sec. 25.301

Loads.

(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.


So limit load is the max that the component is rated to handle in normal service and still function as proscribed. So you should be able to take said component up to the design limit load, and it'll be perfectly fine if you do whatever regular maintenance is required, if my understanding is correct.

Then on top of that, you have the safety factor, which then results in the ultimate load.

Quoting faro (Reply 3):

Plus the weight of the engines/nacelle assembly too; somewhere around 19,000 lbs IIRC.

Once the wings are bearing the weight of the airplane, wing-mounted engines actually have a beneficial role in partially balancing out the huge upward-bending moment on the wings. It could be that, keeping everything equal, an airplane with wing-mounted engines could require less wing structure than one with fuse-mounted engines.



How can I be an admiral without my cap??!
User currently offlinejetlife2 From United States of America, joined Jul 2006, 221 posts, RR: 25
Reply 9, posted (3 years 3 months 2 weeks 6 days 15 hours ago) and read 4195 times:

Widening the original question a little: The highest stresses in absolute terms are probably not in the aircraft structure, but in the engines' rotating parts. Stresses in excess of 150ksi (1000MPa) occur by design at local critical features. Materials sustaining these stresses are nickel based superalloys or certain flavors of titanium. Although they are certified to withstand limit cases, these often don't drive the design. Fatigue is often the limiting mode and so these stresses can be a high fraction of limit stresses due to the parts being managed to a known fixed life limit. For aerostructures these design considerations are often reversed.

User currently offlineKPWMSpotter From United States of America, joined Dec 2006, 458 posts, RR: 2
Reply 10, posted (3 years 3 months 2 weeks 6 days 14 hours ago) and read 4179 times:

Quoting jetlife2 (Reply 9):
The highest stresses in absolute terms are probably not in the aircraft structure, but in the engines' rotating parts. Stresses in excess of 150ksi (1000MPa) occur by design at local critical features. Materials sustaining these stresses are nickel based superalloys or certain flavors of titanium.

oooh, clever, I hadn't even thought of that. Yes, I revise my initial answer to concur.



I reject your reality and substitute my own...
User currently offline474218 From United States of America, joined Oct 2005, 6340 posts, RR: 9
Reply 11, posted (3 years 3 months 2 weeks 6 days 14 hours ago) and read 4161 times:

Quoting jetlife2 (Reply 9):
Stresses in excess of 150ksi (1000MPa) occur by design at local critical features.
Quoting KPWMSpotter (Reply 10):
oooh, clever, I hadn't even thought of that. Yes, I revise my initial answer to concur.

150ksi, pales in comparison with 260/280KSI vacuum melt 17-4 steel used in landing gear. Or even the 180/200ksi 4340 steel used in wing landing gear attached fittings.

As a rule of thumb, it is easy to determine what parts have the highest stresses, just look at the size of the attaching fasteners.


User currently offlinetdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 12, posted (3 years 3 months 2 weeks 6 days 12 hours ago) and read 4118 times:

Quoting faro (Thread starter):
During extreme attitude/overspeed maneouvring, what is likely to damage first, the horizontal stabiliser or the fin?

As others have noted, it depends on the maneuver. However, in general, you're more likely to overload the horizontal tail than the fin.

Quoting faro (Reply 3):
I find this very surprising and quite disturbing too. If operated so near their design limit load, and with repeated flight cycling, failure would be more likely statistically. If as is also likely this failure is asymmetric, one may lose or partially lose flight control...

No single failure would result in departure of a flap, which is what you'd need to get an asymmetry.

Quoting KPWMSpotter (Reply 6):
while composites are strong, they still don't come close to the ultimate strengths of steel and titanium.

CFRP laminate ultimate strength is about 60% higher than titanium and higher than all but the most exotic steels.

Tom.


User currently offlinedynamicsguy From Australia, joined Jul 2008, 887 posts, RR: 9
Reply 13, posted (3 years 3 months 2 weeks 6 days 8 hours ago) and read 4055 times:

Quoting faro (Reply 3):
I find this very surprising and quite disturbing too. If operated so near their design limit load, and with repeated flight cycling, failure would be more likely statistically. If as is also likely this failure is asymmetric, one may lose or partially lose flight control...

Since this load is driven by the flap angle and airspeed when they're deployed the load may get fairly close to limit but it is predictable, so it is difficult to go beyond the limit load. We analyse for Vf+gust for static analysis. And as already noted the fatigue loading spectrum accounts for the repeated load. High-speed flaps stowed cases are not consistently close to limit.

Quoting tdscanuck (Reply 12):
No single failure would result in departure of a flap, which is what you'd need to get an asymmetry.

  

We analyse for failure of any of the support or actuator locations. And there are features built into the design specifically to catch the flap if there is a failure, e.g. "catcher" fittings at hinge locations and redundant pins in case of pin failure. You would have to try extremely hard to lose a flap.


User currently offlineHAWK21M From India, joined Jan 2001, 31712 posts, RR: 56
Reply 14, posted (3 years 3 months 2 weeks 6 days 6 hours ago) and read 4023 times:

Looking at an An124 closely......The Wing root must definetly be the most stressed part considering the load it carries.


Think of the brighter side!
User currently offlinetdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 15, posted (3 years 3 months 2 weeks 6 days 2 hours ago) and read 3899 times:

Quoting HAWK21M (Reply 14):
Looking at an An124 closely......The Wing root must definetly be the most stressed part considering the load it carries.

I think for thread coherence we need to clarify what metric we're using:
-Highest load (i.e. maximum force, almost certainly the wing-body joint)
-Highest stress (i.e. maximum force per unit area, almost certainly in the landing gear)
-Highest % of limit load (i.e. closest to failure, usually the flaps)

It would be very unusual to see two or even three of these in the same component because of conflicting design drivers.

Tom.


User currently offlinefaro From Egypt, joined Aug 2007, 1630 posts, RR: 0
Reply 16, posted (3 years 3 months 2 weeks 5 days 22 hours ago) and read 3787 times:

Quoting dynamicsguy (Reply 13):
We analyse for Vf+gust for static analysis.

I don't quite understand; a gust is not static, is it? Isn't this analysis for dynamic loading rather?

BTW, is a wing's ultimate load determined in a like way by MTOW take-off loading + design (vertical) gust limit? What is the speed of this limiting gust and its impulse (time over which the gust is fully imparted to the wing)?

Finally, there are many accidents out there involving freak gust-induced snapping of fins/horizontal stabilisers, generally near Cb or mountain wave activity. Has there ever been a gust-induced snapping/fracturing of a wing?

Faro



The chalice not my son
User currently offlineCitationJet From United States of America, joined Mar 2003, 2469 posts, RR: 3
Reply 17, posted (3 years 3 months 2 weeks 5 days 22 hours ago) and read 3760 times:

Quoting faro (Reply 16):
Has there ever been a gust-induced snapping/fracturing of a wing?

There was a famous one of a B-52 vertical tail:
http://en.wikipedia.org/wiki/1964_Savage_Mountain_B-52_crash



Boeing Flown: 701,702,703;717;720;721,722;731,732,733,734,735,737,738,739;741,742,743,744,747SP;752,753;762,763;772,773.
User currently offlinefaro From Egypt, joined Aug 2007, 1630 posts, RR: 0
Reply 18, posted (3 years 3 months 2 weeks 5 days 20 hours ago) and read 3711 times:

Off-topic but judging by some of the aeroelastic fluttering videos out on the net, airliner structures are pretty flexible. Whilst holding rudder at engine-out cruise speed in a twinjet, how much sideways, lateral flex does the fuselage of, say, a 773 (74 m length) exhibit? More or less than one meter from the centerline?

Faro

[Edited 2011-09-08 13:01:46]


The chalice not my son
User currently offlinedynamicsguy From Australia, joined Jul 2008, 887 posts, RR: 9
Reply 19, posted (3 years 3 months 2 weeks 5 days 20 hours ago) and read 3696 times:

Quoting faro (Reply 16):
I don't quite understand; a gust is not static, is it? Isn't this analysis for dynamic loading rather?

It is both. You have to account for the variable loading from gusts in the fatigue analysis, but also the peak loading for the static analysis. If you are already at Vf and you get a head on gust then the peak load will be higher than the load at Vf.

I should correct myself. We analyse for Vfe+gust, not Vf+gust.


User currently offlinetdscanuck From Canada, joined Jan 2006, 12709 posts, RR: 80
Reply 20, posted (3 years 3 months 2 weeks 5 days 12 hours ago) and read 3550 times:

Quoting faro (Reply 16):
I don't quite understand; a gust is not static, is it? Isn't this analysis for dynamic loading rather?

It's static for the purposes of stress analysis. Dynamic in the stress analysis world means that you actually have to pay attention to strain rates...this is important in *very* fast events (think car crashes) but not relevant for flight load calculations.

Quoting faro (Reply 16):
BTW, is a wing's ultimate load determined in a like way by MTOW take-off loading + design (vertical) gust limit?

No. That's a load factor calculation. MTOW at 2.5g normal acceleration (usually).

Quoting faro (Reply 16):
Has there ever been a gust-induced snapping/fracturing of a wing?

Almost impossible to know for sure, but aircraft have been lost due to excessive turbulence so it's possible that's it's happened.

Quoting faro (Reply 18):
Whilst holding rudder at engine-out cruise speed in a twinjet, how much sideways, lateral flex does the fuselage of, say, a 773 (74 m length) exhibit? More or less than one meter from the centerline?

Less. Aerostructures tend to be relatively weak in torsion (which is why flutter is a problem in the first place) but very strong in bending. The bending moment on the fuselage while loaded with cargo and sitting on the ground would be considerably larger than the moment from the fin in an engine-out and you don't get anywhere close to 1m deflection sitting on the ground.

Tom.


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