Composite wing spar design 2

[Samp51]

I am building a 3/4 P-51 mustang and the original design was of wood. I am bulding it out of composite materials and want to replace the wooden spar with a composte one, made out of carbon fiber and either a honeycomb core or foam core. I need help with the structural sizing of the spar. If anyone can help I can provide the needed load information.
Thanks
Sam

 

[MikeHalloran]

Rent an engineer.



Mike Halloran
Pembroke Pines, FL, USA

 

[mudflaps]

3/4 p-51?

Do you mean 3/4" to 12" or 9" to 12"? You said the "original" had a wood spar. I'm guessing you want 3/4" to 12", especially if you want a foam core. Probably not a real good forum for RC plane design.

 

[40818]

Some questions for you....

First question: Why would you want to??

Second question: I'm not familiar with the cross secion of the spar (front spar i take it?), but i would have thought it was a "C" section, so how do you intend to bend your sandwich panel to the required shape?

Third question: Is the sandwich panel pre-made? or are you intending to make it yourself?

Fourth question: Have you thought of problems due to thermal differentials and fasteners?

Loads more questions, but you really dont want to start flying in something with little confidence in its spars, its a long way down from 10000ft.........

 

[GBor]

I'm guessing 3/4 scale P-51? What kind of wood and what are its dimensions? Build a composite spar of equivalent strength and the load information doesn't matter...if the wood meets it, then the composite will meet it if properly equivalent.

 

[GBor]

Oh, unless damage tolerance is an issue...then you may have other considerations.

 

[berkshire]

Samp51 (Aerospace)
If the original was built of wood you will have a hard time coming in under the original weights with composite materials. Wood has a good stiffness to weight ratio.
How is the original wooden spar drawn? Is it a box spar, an I beam spar, or a solid block of wood like a champ or piper cub spar, also is this a single spar or two spar wing?

Each one of these will require a different approach when fabricating from composites compared to the original wood.
B.E.

 

[Samp51]

Berkshire,
You are the only one who answered with sensible questions. The original design is made from spruce, birch and mahogany. It is a box spar. The spar is designed with an upper and lower superior and inferior spar with diphrams in between the upper and lower superior and inferior spars. The upper and lower sections are than boxed with plywood. The reason I wish to use composite materials is to save some weight while adding strength. The wing needs to be stronger to increase the cruise speed and the never exceed speeds of the aicraft. Also the materials I wish to use I am familiar with. I am just not an engineer. I wish to build an I beam spar with the caps being carbon and the shear web being either laminated carbon honeycomb or divinycell HT foam. Thanks for your input. Sam

 

[Samp51]

The overall structure will be built with carbon fiber and a combination of composite materials that suite the different areas of the aircraft. The airplane has a gross weight of 3200lbs. and should meet the requirements of FAR 23

 

[berkshire]

Samp51 (Aerospace)
At this stage I would suggest that if you have not already done so, you find and join a local chapter of the EAA and talk to the chapters DAR.
Also if you beef up the wing, you then need to look at the rest of the systems in the aircraft.
I was involved in an accident investigation on an aircraft where the wing had been beefed up, but the tailplane attachment brackets had not, guess what failed.

Another thing to consider, it is better to just use the shape of the aircraft and design from scratch in advanced composites than to build a "black" aluminum or wooden aircraft.
B.E.
Building modifying and repairing plastic aircraft since 1970

 

[40818]

Samp51, i take some umbrage at your comment stating that my (and others) answers were not sensible.
I have carried out stress analysis for both composite and metallic wing spars for both small and very large aircraft for quite some time now and am more than qualified to give advice.
I shall therefore try to enlighten you with some further advice which you are welcome to take onboard or ignore...

If you get it wrong, then you will be in serious trouble. I take it that you wont be qualifying your wing structure by full scale testing? And therefore you will only really find out if your wing is statically strong enough when your pulling a high g inertial manoeuvre and the wing falls off.
You will probably not carrying out any fatigue & Damage tolerance analysis on your new spar, so you will have no idea as to when the spar will have cracks of a detectable length, and whether the existing inspection interval is sufficent to pick them up. Also your going to exceed the current flight loading envelope which will impact onto the whole aircraft (both static and fatigue). Thoughts like this need to be addressed at an early stage.

You intend to create an I-beam using different materials between the caps and web, have you though about the method for attaching these together. Which ever method you choose then these need to be able to carry the shear loading developed during wing bending. Also as you will be using composites, the load distribution for loaded fasteners will have peak critical locations.

You are familar with sandwich panel construction, and therefore you will also know that the method of joining the parts is so critical in the final strength of the assembly, so you need to make sure everything is tip-top in that respect, otherwise you adhesive bond strength drops through the floor.

Fasteners will also be something you need to pay attention to, especially on the new spar to cover and rib interfaces.

You also need to have experience of the analysis proceedures to stress the whole wing. You will inevitably change the stiffnesses of the structure and therefore the load distribution, and you cannot then rely on existing calculations to give you guidance. A large Boeing passenger aircraft fell out of the sky many years ago due to changing the upper stabiliser cover from aluminium to steel and didnt think about stiffnesses.
Your changing stiffnesses will also impact onto the response of the aircraft due to decreased (most probably) flexibility, and its natural frequencies. It is driven by a prop aircraft and therefore you may change the local frequency responses and have parts resonating and ergo low fatigue lives. Other wing loadings must also be catered for, brazier, fuel pressures, fuel systems overloads, local load inputs etc

Your spar also acts as a boundary for the integral fuel tank, so does the composite spar have adequate drainage (if its honeycomb), or sufficent capability againt leakage.
Other items that need thinking about are thermal differentials, material compatabilities, material behaviours. I would recommend getting some early professional advice regarding your project. otherwise if it goes through to a DER clearing house then it wont get very far i'm afraid.

 

[SWComposites]

Samp51, you say "The wing needs to be stronger to increase the cruise speed and the never exceed speeds of the aircraft." Seems like a lot more would require redesign to increase speeds than just the spar. You also say "The overall structure will be built with carbon fiber and a combination of composite materials that suit the different areas of the aircraft"; are you redesigning the entire aircraft to be built with composite materials?

Is this a one-off experimental aircraft, or do you intend to try to certify it to FAR23?

Can you post a cross-section of the current spar?

 

[rb1957]

wasn't the original P51 close to trans-sonic ?

i think so thought needs to go into flying this 75% model faster than 75% of the original P51 (geometric scaling, reynolds numbers).

i think you can design and build the structure, then test it on the ground ... sandbags (lots of them) would give you a good idea as to the strength of the wing.

do you Really want to certify to FAR25, (I don't think you can without a test) or is this "just" the design guide-line ?

 

[plasgears]

I had some experience trying to qualify an extruded carbon filament beam mfr. Our first interest was for armored car work. These beams were resin rich, and even a 22 cal bullet penetrated without any trouble. I rejected this approach, and I got an angry visit from these people in force. There were loud voices raised by the mfr in my office. He will never darken my facilities again.

Your 3/4 scale airplane needs the attention of an aero with experience in composites: glass, carbon, aramid, boron, etc. The glider community probably has the right combination of experiences you need.

 

[Aerodesign]

I recommend using the Martin Hollmann guides and software -

http://www.aircraftdesigns.com.

You will need to satisfy your local inspector that the design you have come up with is sufficient - do this by stress analysis or test a typ section of spar. In the USA that inspectorate will be the EAA, in the UK its the LAA. Generally you will have significant trouble clearing the design in the UK, unless you demonstrate that you have taken into consideration the appropriate material 'special factors'. For composites these special factors are usually covered by a design factor of 1.5. So for example, its its a aerobatic aircraft stressed to +6/-3g you will need to design to +6 x 1.5 (for ultimate load) x 1.5 (composite special factor) = 13.5g
Martin Hollmann's publications have some allowables for composite materials that are a good starting point. If you deviate significantly from the grades he uses you will need to justify the numbers (maybe by reference to AGATE or other documents). Fatigue is not usually an issue with composites, however, with carbon you should take into account local strains and strain cencentrations. Excessive localised strains will not be re-distributed as in a metallic (alloy) structure, carbon fails spectacularly in a brittle mode of failure at a relatively low strain.
I would therefore recommend that you base your spar design on a hybrid design of S-glass with carbon caps. You will need to consider the problems of high differenctial thermal expansion (owing to different cte values) but this combination can be made to work very well.
Lastly, when you lay up your spar, ensure that exotherm (heat built up) is kept under control and l would therefore recommend a vinylester resin system. This may depend upon your locale (geographically) and ambient conditions.
Dont forget that the spar needs to be attached to the rest of the wing - so dont go drillign holes in it please, design teh structure with bonded joints.
Best of luck.

 

[jim787]

My two cents… I think your making a big mistake! You might be decreasing the elastic response behavior of the wing as a whole which may in turn increase the local stresses where the other wing shape components attach to the spar system. Who designed this aircraft and what is their background! I have experience in building composite aircraft and I’ve seen some pretty scaring engineering and workmanship. That thing your building is not a real Mustang and no matter what you do it never will be, build an aircraft that already does what you want and avoid this all too common trap.

 

[zerosum]

I tend to agree with jim787. A test pilot friend died in a midget air racer whose wing spar had been very slightly modified. Flutter was the result.

 

[berkshire]

The word " natural frequency" comes to mind here.
When you change that, as in by building a stiffer wing than the original designer other things change too.
even things you do not think about.

I was in a hanger one day while vibration tests were being done on an airframe. All of a sudden the tester found a frequency that coupled with the ailerons, the plane which had been quietly sitting on the suspension slings, started dancing, the ailerons violently flapping up and down until the mag pots were shut off.
The guy standing next to me looked at me and said " I am glad I was not flying that when that happened." He was the test pilot.
B.E.

 

[dgapilot]

Samp51,
Much of the advice given here is correct. From the sounds of what you are talking about, you are looking at designing a completely new aircraft with the outward appearance of a 3/4 scale P51. If you are changing materials and or performance characteristics of an existing design, a complete engineering analysis should be performed. As a DAR that can issue airworthiness certificates for homebuilts, I would want to see the full engineering analysis before signing my name. As the speed/weight/complexity increase, I want to see more and more data to support the fact that the aircrqaft can fly safely.

If built to plans (that have several examples already flying) or from a kit and to the kit plans, I have a much better feeling about issuing the certificate. If an original design, I have the option to issue any operating limitations I see fit to insure the safety of the aircraft, its occupants and the public. For a high performance aircraft of untested structure, you can bet that the limitations will not be the ones published in 8130-2F.

Listen to what Berkshire and others are saying. Brian was working with composite aircraft back in the early days of glass sailplanes so he knows what he is talking about.

DS