wing engineering stuff
 

Hello all,

First post on the new system for me... hope it works.

I was curious if anyone has looked into calculating the stress and bending loads on the RV-7(A) wing. Has anyone calculated the area moment of inertia for the main spar near the root or other places? I'm sure Van has all this stuff, but I'm just trying to see how and why the spars were chosen to be the size they are...

Thanks to all that reply,

Jonathan Cude

I am working on installing strain gages in the RV-7 wing (and maybe other locations) to do structural analysis on the RV-7. I don't have all the equipments and resources to do the analysis yet so it'll take me awhile before I can extract any useful data.

If anyone know the story/reason behind the wing spar design, I'm curious to know as well.

:confused:

G Meter Replacement?
 

I have spoke with a friend of mine about this very topic. Is this for a better G meter? I've always thought that G's were a poor (too conservative) way to limit an airplane because weight is unknown. Seems like you could 'easily' make a weight compensated G meter. You could also calculate torque by putting gages on the motor mount....

What's this for?

Chuck

purpose
 

Hey Chuck,

I was interested in studying the wing to find the physical numbers and math Van's must have used (if any!) when designing the wing. I'm not so interested in modifying the wing, increasing the g-load or gross weight. I was just curious about the loadings on the wing and how they are distributed through to the spar, especially near the root, where bending loads are the highest.

From calculating the area moment of inertia, I get a number around 42 in^4. Seems awful high and I question my formula seriously. Thought surely someone else in the engineering (graduated) world could shed some light.

I'm still in the learning process and my instructors have yet to show me the way... 4.5 years into aerospace engineering and I swear to god it was last week before I sat in a single class that even mentioned the word "AIRPLANE"!

Jonathan

area moments
 

Jonathan,

I did a very rough calculation (with guessed dimensions of the main spar) and I got around 38 in^4 so your calculation is in the ball park. The easiest way to do this calculation is to use a CAD model x-section. You can also just use the simple equation of 1/12*b*h^3 and use the parallel axis theroem to get the "spar caps" and then just add all the values together (spar web + each spar cap). Make sure you use the distance from the spar mid-plane to the spar cap centroid in your parallel axis equation.

As was hinted at earlier the spar does change alot near the root with alot more build up there. I don't have my spar in front of me so....

The spar is not the only member to take the load on a wing. In a wing with a rigid skin (not fabric) the skin also takes some of the bending loads.

Enjoy

Jon

When I was in grad school, I had to teach aircraft performance to my advisor professor (so he could teach the class). He wasn't even sure what the various parts of the airplane were called, let alone what they were for.

calculation
 

I pulled an equation out of the machinists holy bible... A big, old, green book we have in the shop on campus. It gave cross sectional information on many, many different shapes. I picked out a formula for a symmetric "C" channel, even though the top spar bar is about .125" shorter than the bottom. I used the top spar dimension as this would give a conservative (smaller) figure.

I=[(B*D^3)-H^3(B-T)]/12

B=.125+.063+1.25=1.438"
D=7.75"
H=7.75-(1.375)*2=5"
T=.125+.063=.188"

SUBSTITUTED IN: I = 42.75"^4

I assumed the skin to provide no additional support to the spar. This was done so that any additional strength from the skin would be left in the reserve margin of the structure (fatigue, corrosion, overloading, poor construction, etc.)

BENDING LOADS: Q=(M*Y)/I ALSO, Mmax=(Qmax*I)/Y

Q=yield strength=60ksi (2024-T3)
M=bending moment
Y=distance from neutral axis to extreme fiber
I=moment of inertia=42.75"^4

Mmax = [(60,000psi*42.75)/3.875"] = 661935 in*lb = 55160 ft*lb

So the wing has about a 10' arm, and if I assume a constant uniform lift distribution... (worse than real world bending loads?)

M=F*D and F=M/D 55160 FT*LB/5' = 11032 LB

SPAN LOADING: 11032#/10' = 1103.2# PER FT
WING LOADING: SPAN LOADING/CHORD = 1103.2/4.8' = 229.8#/FT^2

Now for the good stuff that makes a little more sense...

Wing loading * wing area = loading so, 229.8#/ft^2*120ft^2 = 27580#

and loading/gross weight = g forces so, 27580#/1800# = 15.3gs

Any comments? I'm in no way a structure genious and there have to be many things i've left out. I hope anyone interested could give there input here and maybe we could refine this number even further... Heck, I might even be so far off none of this is even worth typing!!! I just don't know.

Hope to hear more thoughts here,

Jonathan
RV-7A finishing Fuselage

loads
 

Jonathan,

Based off your "I" value and assuming "c" is 3.875 inches (stress=Mc/I) I calculate approximately 12.25 g at failure for the main spar at gross (1800 lb). This of course does not assume that the rear spar takes any load (which it does) and that the skin takes no load (which it does).

To determine how much load the rear spar takes you need to look at the airfoil plots to find the center of lift.

If you want a crash course in aircraft design look at the books on this web site http://www.aircraftdesign.com/

Before grad school I bought one of these (Aircraft Design, A Conceptual Approach) and read it. It is not very detailed but it is useful. Also get a good Mechanics of Materials text book for the stress calcs.

Jon

Design Books
 

The Raymer book you mention is a pretty good book. That and Roskam's Airplane Design series have been my stand-by books for aircraft design. For aircraft structures, the best book I've seen is the 1965 edition of Bruhn's _Analysis and Design of Flight Vehicle Structures_. I don't own a copy of it, but I wish I did!

Alex

Lift distribution
 

It's my understanding that the certification standards, which Van's uses as a guide, specify several "worst case" loading scenarios that are to be used for the structural calculations. There are low speed / high angle of attack cases and high speed / low angle of attack cases. Some or all cases specify control surface deflections. For example, if you have flaps up but a downward aileron deflection, that would shift the lift distribution toward the tip. Van's writeup on the RV-8 wing investigation describes some of the load testing they did.

Alex