I'm surprised that you haven't received very much publicity on this series. I would like to take a moment and thank you, as it is the least I can do. Thank you for sharing your knowledge, it was very informative and extremely helpful in understanding all the intricacies that go into aircraft design.
For these thin walled structures like the d tube and the fuselage tail boom, the failure mode would inevitably be local buckling on the compression side but I can't seem to find any robust mathematical framework to predict this buckling. Also to mitigate this buckling, you use ribs. But what should the spacing and configuration of the ribs be? Can you point me in the right direction? thanks
@@TheJustinJ Thank you, but suppose, I wanted to calculate the rib spacing, for say a tail boom with a circular cross section, how would I go about it? Thanks again
If the spar deflects too much, might it not affect the aerodynamics even though the yield strength of the material is not affected? Can a spar be considered to have failed if it deflects too much even though it is still structurally ok? Therefore shouldn't deflection be the design criterion for a spar rather than the yield strength or ultimate tensile strength? Thank you
Usually not, but stiffness is a design consideration because it affects flutter, especially on gliders with their long thin wings and fuselages. A structure that is too flexible can negatively affect flight characteristics. If an airfoil shape deforms, it probably creates more drag.
@@sonjaenglert I have one more question which I hope you won't mind answering; What values do you use for yield strength, ultimate tensile strength, compressive strength and elastic modulus for glass fiber reinforced epoxy resin? Thank you
@@SunilSundar E-glass has an elastic modulus of 10,600,000. It breaks around 477,000psi. Or 0.045 strain (4.5% strain). Epoxy is in the range of 500,000. It breaks around 8,000psi. Or 0.016 strain (160 micro-strain). Because epoxy cracks and disintegrates at 0.016 strain, that is the maximum strain that a laminate will be able to survive for more than one use. Most designers use 0.004-0.005 strain to allow plenty of safety margin, and to account for fatigue which will crack the laminate at a lower strain with repeated flexing. Now, multiply those above modulus's each by ratio of epoxy to glass in your layup. And then multiply by 0.005 for design allowable stress. Keep in mind, only the glass in the 0-direction (of the load) is counted. So a laminate of 0.5 glass/epoxy where the glass is 0/90/0/90 only half of the 0.5 glass will be counted for strength. So 0.25. Example: 10.6e6 x 0.4 Glass, x 1/2 in the 0 direction = 2,120,000. And 5e5 x 0.6 Epoxy = 300,000. Add together, = 2.42e6 Modulus. X 0.005 strain = 12,100psi allowable stress in the 0 direction. If 100% of fibers aligned to the 0 direction instead of 50%, use 24,200psi. This is tension. In compression, multiply by 0.8 for better estimate of strength. 24,200 x 0.8 = 19,360. This is hypothetical and not rigorous. There are many factors. This is probably conservative and useful for UAV or mode designs. Not for man rated aircraft. YMMV.
Hi Sonja, your explanation on small aircraft design is so good. We are a startup looking forward to benefit students in India with your expertise. It would be a privilege for us if you would tie up with us as industry expert. Can you share your email so I can send details. Thank you
Wow
I’ve watched all 17 videos at once!
Thank you
Excellent coverage of the aircraft structure. Very nicely explained.
Muy bueno. Sigan adelante con el tema. Es muy amplio pero vale la pena explicar. Excelente
I'm surprised that you haven't received very much publicity on this series. I would like to take a moment and thank you, as it is the least I can do. Thank you for sharing your knowledge, it was very informative and extremely helpful in understanding all the intricacies that go into aircraft design.
ive found your channel and it is very informative on this facinating field.Thankyou for all the effort and knowlege youve given .
Thank you! Very helpfull and inspiring
For these thin walled structures like the d tube and the fuselage tail boom, the failure mode would inevitably be local buckling on the compression side but I can't seem to find any robust mathematical framework to predict this buckling. Also to mitigate this buckling, you use ribs. But what should the spacing and configuration of the ribs be? Can you point me in the right direction? thanks
Depends on the construction method. Most aluminum designs allow post-buckling of the wing skin. The spar caps take all the bending loads.
@@TheJustinJ Thank you, but suppose, I wanted to calculate the rib spacing, for say a tail boom with a circular cross section, how would I go about it? Thanks again
If the spar deflects too much, might it not affect the aerodynamics even though the yield strength of the material is not affected?
Can a spar be considered to have failed if it deflects too much even though it is still structurally ok?
Therefore shouldn't deflection be the design criterion for a spar rather than the yield strength or ultimate tensile strength?
Thank you
Usually not, but stiffness is a design consideration because it affects flutter, especially on gliders with their long thin wings and fuselages. A structure that is too flexible can negatively affect flight characteristics. If an airfoil shape deforms, it probably creates more drag.
@@sonjaenglert thank you
@@sonjaenglert I have one more question which I hope you won't mind answering;
What values do you use for yield strength, ultimate tensile strength, compressive strength and elastic modulus for glass fiber reinforced epoxy resin?
Thank you
@@SunilSundar E-glass has an elastic modulus of 10,600,000. It breaks around 477,000psi. Or 0.045 strain (4.5% strain).
Epoxy is in the range of 500,000. It breaks around 8,000psi. Or 0.016 strain (160 micro-strain).
Because epoxy cracks and disintegrates at 0.016 strain, that is the maximum strain that a laminate will be able to survive for more than one use. Most designers use 0.004-0.005 strain to allow plenty of safety margin, and to account for fatigue which will crack the laminate at a lower strain with repeated flexing.
Now, multiply those above modulus's each by ratio of epoxy to glass in your layup.
And then multiply by 0.005 for design allowable stress. Keep in mind, only the glass in the 0-direction (of the load) is counted. So a laminate of 0.5 glass/epoxy where the glass is 0/90/0/90 only half of the 0.5 glass will be counted for strength. So 0.25.
Example: 10.6e6 x 0.4 Glass, x 1/2 in the 0 direction = 2,120,000. And 5e5 x 0.6 Epoxy = 300,000.
Add together, = 2.42e6 Modulus. X 0.005 strain = 12,100psi allowable stress in the 0 direction.
If 100% of fibers aligned to the 0 direction instead of 50%, use 24,200psi.
This is tension. In compression, multiply by 0.8 for better estimate of strength. 24,200 x 0.8 = 19,360.
This is hypothetical and not rigorous. There are many factors. This is probably conservative and useful for UAV or mode designs. Not for man rated aircraft. YMMV.
@@TheJustinJ This is the best answer I have gotten to this question! Thank you you Sir!
My email is on my web site caro-engineering.com
Hi Sonja, your explanation on small aircraft design is so good. We are a startup looking forward to benefit students in India with your expertise. It would be a privilege for us if you would tie up with us as industry expert. Can you share your email so I can send details. Thank you