INTELLIGENT WORK FORUMS
FOR ENGINEERING PROFESSIONALS

Log In

Come Join Us!

Are you an
Engineering professional?
Join Eng-Tips Forums!
  • Talk With Other Members
  • Be Notified Of Responses
    To Your Posts
  • Keyword Search
  • One-Click Access To Your
    Favorite Forums
  • Automated Signatures
    On Your Posts
  • Best Of All, It's Free!

*Eng-Tips's functionality depends on members receiving e-mail. By joining you are opting in to receive e-mail.

Posting Guidelines

Promoting, selling, recruiting, coursework and thesis posting is forbidden.

Jobs

Determining 7781 Fabric Material Properties

Determining 7781 Fabric Material Properties

(OP)
Hi,
This is in reference to an excellent response by Rpstress user in this thread.

I was wondering if the user Rpstress could elaborate more on how the Young's Modulus of the 7781 fiberglass in some sort of resin was determined. I especially did not get the origin of some of the knock down factors that were used in calculating the Young's Modulus.

On a more general note, how does one get reliable approved data about any ply (especially cloth)? Is Mil-17 Handbook Vol 2 a good source? For example, if I needed properties for a E-Glass fabric based facesheet in a honeycomb, I can pour over Hexcel website. But all I found was 7781 spec sheet which typically provides breaking strength in fill & warp direction apart from a few other properties.

Thanks in advance,
- B

RE: Determining 7781 Fabric Material Properties

Hi Burner2k.

The 0.5*0.5 was because I ignored the 50% of the fibers at 90° and the other 0.5 was because 50% of the area was considered to be resin (also ignored in that case).

On strength there was also a 0.8 to allow for the fact that basic fiber strength is always reduced in a real-world laminate (fibers are not in a straight line and so on) hence 0.5*0.5*0.8*fiber strength (I used 2200 MPa) is probably a passable guess at basic fabric strength at 440 MPa. Compare Hexcel's cloth strength of 570 or 450 lb/in, in a material 0.0086" thick that's 66 or 52 ksi, an average of 410 MPa. 440/410 means that estimate was only 10% high.

I should probably have had a factor on stiffness as well; that is usually 0.9 or 0.95 rather than 0.8. That would have made the estimate from fiber properties even lower at 16 or 17 GPa rather than 18. The 20 or 23 GPa is from a datasheet. Making more allowance for the fibres at 90° might have made the estimate more accurate. However, all that stuff is pretty much a guess and some actual test data is needed, especially since the elastomer had no data.

CMH-17 (and NIAR/NCAMP) is the best source for data as long it has the material of interest. The existence of B-basis information is quite important. Manufacturer's data is almost always average and may have been cherry-picked to look good (the majors generally won't do this but they always put warnings that the values are for comparison only or similar wording). If you have access to an airframer's data that should be about as good as CMH-17. It should also have all the values needed including bearing and effects of temperature.

The CMH-17 data is always for a flat laminate. A sheet on a honeycomb will telegraph (dimple into the honeycomb cells) if it is cobonded (facesheet resin cured at the same time as the bond to the honeycomb is made). This will happen a bit even if a caul plate is used. The allowable stress in the facesheet should be reduced from the CMH-17 value if that is the case. And yes, you need testing to quantify the effect although a knock-down of 0.8 times is a passable guess. Telegraphing affects compression strength more than tension and with glass the compression is sometimes a bit higher than the tension anyway. And of course also not in the CMH-17 data is any allowance for damage. Some allowance for damage should always be made. Open hole data if given can be appropriate, although for glass the filled hole strength will often be less than the OHT which may be less than the OHC. FHT is often the best (safest) value to use for glass. A lot of airframe is designed to compression after impact and this will often amount to 3500 microstrain for carbon laminates. I'm afraid I don't have any CAI data for glass. Your company or customer should have specific guidance as to allowance for damage.

Hope that clears stuff up a bit (and doesn't confuse matters more).

RE: Determining 7781 Fabric Material Properties

(OP)
Thanks a lot sir for the explanation. I may have to read it a few times to get a clear understanding.

Unfortunately as of now, we do not have access to any airframer datas (hopefully may get it in the near future if a project contract is finalized). CMH-17 seems like an evolution to Mil Hand Book-17. The latest rev of Mil HBK-17 was in 2002 (so its not that old) and it is available for free.

When you mention damage data, do you mean damage as in holes which are deliberately drilled for fastener installation or damage resulting from an accident? Is notched strength data for laminates available in public domain?

Let me ask you another question. I was under the impression that stressing of composites components is done by extracting ply failure indices. I got some sneak peak in to an excerpt of a major aircraft OEM stress report of a composite airframe component. To my surprise, they have based their results on a criteria (stress based) other than ply failure indices. During the training we received from MSC Corp, most of the stressing was done based on Failure indices (of course our examples were based on simple geometry & loading conditions). Do you have any thoughts based on your experience?

Thanks again for answering.
Regards,
- B

RE: Determining 7781 Fabric Material Properties

CMH-17 (currently issue -G (Mar-2012) for most volumes, of which there are six) is indeed the demilitarized version of MIL-HDBK-17. Like MIL-HDBK-5 turning into the MMPDS it now costs money. There is a bit of extra info compared with the last version of MIL-HDBK-17. If you work in a company they really should buy both MMPDS (now up to ver -08; I don't think -09 is out yet) and CMH-17 (and probably ESDU's MMDH if you do work in Europe, although that's more relevant to the UK; generally Airbus has its own stressing and materials data sheets like Boeing, Sikorsky, etc., so when working for a primary like those there's not so much need for these independent sources). Most of the CMH-17 strength and stiffness data is for UD S-glass. There is a bit for woven E-glass; my notes say three different resins were considered with 7781.

For secondary structure in carbon it was usual to use an open-hole strength to allow for damage. For glass we have been using filled hole data because the values are a bit lower than open hole strengths. Although impacts do a different sort of damage from drilled penetrations it was accepted that holes counted as damage. (You still had to do something like bearing/bypass analysis joints (holes with loaded fasteners)). For important structure compression after impact testing was needed. This sort of decision needs discussion with the certificating authorities.

Open hole compression strength was based on testing of quasi-isotropic laminates and classical laminate analysis was used to find (well, predict with some acceptable degree of error) the worst-case ply stresses at failure and after suitable statistical processing (as described in CMH-17) these were then used as B-basis allowables with ply-based analysis. A suitable failure criterion is needed for different combinations of directions of forces in plies (Tsai-Wu is common and pretty safe).

In the past it has been ok to use simple failure criteria and classical lamination theory (usually post-processing FE data) which gives a ply-based M.S., but recently (since the early 2000s) it's been more usual to use a laminate-based M.S. with strengths (or failure strains) based on tests of laminates with impact damage. This is for carbon. The laminate damaged and tested has to be very similar to the one under consideration meaning a large range of CAI tests were needed, making material qualification expensive (over a hundred CAI tests are generally needed and they are not a cheap test). The same should apply to glass, but it is rare to see compression after impact data for glass. It might pay to find out more from someone with a non-aerospace background; anyone worked on wind turbine blades or glass bits of civil structures like boats?

Note that currently the airframer has to meet the cost of qualification. Metal-suppliers usually pay for A- and B-basis tests for common static loading and often for S-N tests and even da/dN crack growth but composites suppliers do not. (This may be beginning to change; see the NIAR/NCAMP Wichita website for more discussion on this.)

RE: Determining 7781 Fabric Material Properties

(OP)
Rpstress,
Thanks for replying.

Your response is very valuable.

RE: Determining 7781 Fabric Material Properties

(OP)
Rpstress,
I am not sure if this question is pertinent here but for the sake of continuity, I will post here.

Based on your reply about stressing for composites, how are laminates defined in Nastran so as to obtain stress/strain outputs for the entire laminate & not plies? The procedure I've been using is to define individual plies using PCOMP and in the output, depending on output request, Nastran provides individual ply FI and/or Stress outputs. As far as I remember, it is not possible to obtain the output for the entire laminate if the laminate is defined using PCOMP.

So just curious on how one would go about FE modelling laminates if the outputs are needed for entire lamiantes.

Thanks,
- B

RE: Determining 7781 Fabric Material Properties

(OP)
Hi,
I may have found the answer to the above by meself.

I guess in NASTRAN, one has to turn ON (or enter) the option NOCOMPS <value>. Nastran will provide PSHELL equivalent laminate smeared output.

I hope the above procedure is the correct one!

RE: Determining 7781 Fabric Material Properties

Yes, that's how to get smeared stresses. The Quick Reference Guide says
"NOCOMPS controls the computation and printout of composite element ply stresses, strains and failure
indices. If NOCOMPS = 1, composite element ply stresses, strains and failure indices are printed. If
NOCOMPS = 0, the same quantities plus element stresses and strains for the equivalent homogeneous
element are printed. If NOCOMPS=-1, only element stresses and strains are printed.

Homogenous stresses are based upon a smeared representation of the laminate’s properties and in general
will be incorrect. Element strains are correct however."

Just because the stresses are based on a smeared representation doesn't make them totally useless. They should still tie up with the element forces and areas (thicknesses). Unless someone knows different?

Red Flag This Post

Please let us know here why this post is inappropriate. Reasons such as off-topic, duplicates, flames, illegal, vulgar, or students posting their homework.

Red Flag Submitted

Thank you for helping keep Eng-Tips Forums free from inappropriate posts.
The Eng-Tips staff will check this out and take appropriate action.

Reply To This Thread

Posting in the Eng-Tips forums is a member-only feature.

Click Here to join Eng-Tips and talk with other members!


Resources


Close Box

Join Eng-Tips® Today!

Join your peers on the Internet's largest technical engineering professional community.
It's easy to join and it's free.

Here's Why Members Love Eng-Tips Forums:

Register now while it's still free!

Already a member? Close this window and log in.

Join Us             Close