we have a general lack of awareness of the impact on fatigue life of flaps and associated structures of routinely operating close to the flap placard speeds.
The manuals are clear on a TYPICAL radar vectored approach... don’t use flaps as speedbrakes, descent clean where poss, second stage of flap a few miles before the loc etc.
I have calculated the basic loading on a retracted 787 flap system to be 20 Tonnes at 170 Tonnes landing weight. How do I calculate the increased loading at flap 30 / 170 knots. Rule of thumb is fine as I suspect this is hugely complicated in terms of certification type calls. Thanks in advance
Hmm, complicated. Could possibly calculate by working out the change in stab angle at set points and change in pitching moment about CofG. Suspect the data you will need will include CofG, stab angle and TAS clean, environmental data at the data point, aircraft weight.
You can then calculate the change in pitching moment by using a second data set as above, aircraft clean, settled, flaps posn x. You could then look at data to come up with a rough idea of how muc hof this pitching moment change is caused by the flap, although you will need pressure distribution graphs for the aerofoil with the flap deflected.
However, I would be careful stating that this is evidence increased likelihood of fatigue, which is presided upon cyclic load frequency often more than the scalar of force. Have a look at the Tacoma Narrows Bridge collapse to see what Im talking about. To figure out this, you would need to be a structural engineer, with access to the aeroelastic model they have for the aircraft as well as detailed composite layup data for the likes of flap panels and tracks. I would hazard a guess that there may well be a particular speed, which may well be lower than flap extended limit, which may cause the greatest fatigue damage due to the likes of vortex shedding from nacelles/vortilons or the complex interaction of other aerodynamic components. In short, this is not trivial and from my time in industry, aero(servo)elastics are the most difficult of problems to solve or even accurately predict, particularly with the massively flexible structures now in use to generate the largest Aspect Ratio possible.
However, this is well accounted for within the design spec and will be factored into aircraft cycle limits. With it being a new aircraft, Boeing may well have strain sensors on redundant load bearing components which will monitor usage to ensure that the cyclic loads are within predicted limits.
Final point, fatigue is usually not as big an issue in composite structural components as the likes of cut aluminium or titanium spars (as long as an ultimate load factor is not exceeded). One of the benefits of this material. Usually, degradation in resin will be the limiting factor in material life.
My other observation would be that jets like the 78 are so slippy that small flap deflections are useful (as long as you dont end up dragging her in at the bottom). Most of the chaps I know on the 78 have stated how slippery she is. Speed brake from my time in the boot is noisy, so I would argue, why not?