Light flying wing reynold #?

[sloper steve]

HI I have a light flying (5.5oz) but at 43"
Around 400 sqin of wing area.

From this data is it possible to know the reynolds number?

Also,
I'm trying to figure out profili.
Could anyone explain cl and cd (coefficient of drag?) graphs at different reynolds numbers?

[Ollie]

The only other number you need is the design lift coefficient. That depends on the airfoil selected, the CG location and the twist used to achieve pitch trim.

Which coefficient of lift you choose depends on the purpose of the reynolds number calculation. If you want to calculate the performance envelope of the wing you could choose several different coefficients of lift within the range of angles of attack between positive and negative stall of the airfoil.

The tip and root ends of the wing will be operating at different reynolds numbers because of their different chord lengths. It is usual to take the average chord of the wing to represent an average reynolds number. The average wing chord is 400/43=9.3 inches average chord. Since the reynolds number in a standard atmosphere is a constant (6380) times the wing chord in feet (9.3/12=0.775) times the airspeed in feet per second, we need to find the airspeed associated with the chosen coefficient of lift. The wing loading in pounds per square foot is (5.5/16)/(400/144)=0.124.
The airspeed in feet per second is 29(0.124/Cl)^.5. For example, if the coefficient of lift is 0.5 then the air speed is 29(0.124/.5)^.5=14.4 feet per second. Therefore the average reynolds number of the wing in a standard atmosphere is 0.775 times 14.4 times 6380 equals 71,400.

To find the profile drag of the airfoil consult the Cl vs Cd polar diagram for a reynolds number of 71,400. You may have to interpolate between other reynolds numbers displayed on the polar to find the Cd associated with a Cl of 0.5 as in this example. To find the total drag of the model, the profile drag, you just found from the polar diagram of the airfoil, has to be added to the induced drag and the parasitic drag of any exposed linkages, tip fins or other protrusions from the wing. The minimum induced drag is the coefficient of lift squared divided by pi times the aspect ratio. Pi is 3.14. The aspect ratio is 43/9.3=4.62. The minimum induced drag for a coefficient of lift of 0.5 is 0.25/(3.14x4.62)=0.017.
The minimum induced drag must be divided by an efficiency factor if the lift distribution isn't elliptical. To determine the lift distribution of the configuration and the efficiency factor see:
http://aero.stanford.edu/WingCalc.html
The sweep angle to be used in this online program is the sweep of the 25%chord line of the wing. The Cl must be within the capabilities of the airfoil between its positive and negative stalling angles of attack. The twist used in this program is the aerodynamic twist which is measured between the zero lift angle of attack of the root and tip airfoils, not the geometric angles of attack.

Designing tailless aircraft is actually more complex than conventional configurations because there is so much additional interaction between design choices.

[raptor22]

You also need the speed you will travel at and your altitude. The following eq is used to calcualate the reynolds #:

RN=V/kv

RN= reynolds #
V= velocity (fps)
v= kinematic viscosity, sp ft per sec
x= distance from leading edge, ft (the location of RN that you are finding)

For x I'd probably just use the widdle of the chord.


--Alex

[Sail 'n Soar]

Originally posted by raptor22
You also need the speed you will travel at and your altitude. The following eq is used to calcualate the reynolds #:

RN=V/kv

RN= reynolds #
V= velocity (fps)
v= kinematic viscosity, sp ft per sec
x= distance from leading edge, ft (the location of RN that you are finding)

For x I'd probably just use the widdle of the chord.


--Alex

I'd not get too caught up in the altitude variation details and just run Ollie's calculations. The biggest drivers are CL, wing loading, and wing cord. There will be some variation with altitude - the density will go down, the speed up by the inverse of the square root of the density, and the viscosity will drop - at least to ~ 100,000 feet!. As a result, for the standard atmosphere and a given aircraft and trim (i.e., CL) the difference in Re from sea level to 5,000 feet will be less than 3%. The difference between sea level and 10,000 feet will be less than 8%. Compared to all the other variables, the aero impacts of these Re differences will be lost in the noise.

[sloper steve]

Any links to sites that explain CL and graphs?
For instance I ask profilis author about color pressure field and am able to show one on Java foil but when I try in profili I just see graphs. Still can't tell if one axis is chord length, pressure, which line is top airfoil. Spent a few weeks looking around the internet for beginners info on this stuff and find nothing.
The articles I'm reading for moment coefficient are so hard to understand. It sounds like it can be explained with references to fulcrums and teeter totters.

[Ollie]

B Squared published a book Understanding Airfoil Polars but it is out of print. That would have the explainations you are looking for if you can get a used copy. If not, get a copy of Model Aircraft Aerodynamics by Martin Simons. It is a good way to start your aerodynamic education. It will help you understand polar graphs and also put the understanding in the proper context.

The coefficient of lift is a dimensionless quantity that is a linear function of angle of attack up to the stall. It is used to compare similar flow conditions around different airfoils and to calculate the lift force. The lift force at a given reynolds number is proportional to the coefficient of lift times the wing area, times the velocity squared. With all the jargon it is necessatry to have some aerodynamic knowledge to understand what I am talking about. You can get that knowledge from the recommended reading. Aerodynamics is based on physics so, some knowledge of physical science will help greatly.

[sloper steve]

Alright! the book is on the way. Thanks.

If you know any sites that explain the graphs please let me know.

Model Aircraft Aerodynamics (http://www.amazon.com/exec/obidos/tg/stores/detail/-/books/1854861905/104-6818471-1538352)

[Ollie]

A Google search of "airfoil polar diagrams" turnd this up:
http://www.mh-aerotools.de/airfoils/hdipolar.htm

[Sail 'n Soar]

Originally posted by sloper steve
Alright! the book is on the way. Thanks.

If you know any sites that explain the graphs please let me know.

Model Aircraft Aerodynamics (http://www.amazon.com/exec/obidos/tg/stores/detail/-/books/1854861905/104-6818471-1538352)

I don't really understand your question. The following site gives polars for multiple airfoils at multiple reynolds numbers from about 40,000 to 500,000. http://www.nasg.com/afdb/search-airfoil-e.phtml

The charts include CL(lift coefficient) vs CD (drag coefficient) and CM (moment coefficient) vs alpha (angle of attack.) What is missing here is the relatiionship of CL with alpha. But in general these curves give you the basic information you need to choose an airfoil for a specific application.


At an elementary level, Lift = CL *A* V*V*rho/2, where A is area and rho is air density. Similarly, Drag = Lift = CL *A* V*V*rho/2. Your question appears to imply that you may be referring to charts that show the velocity distribution around the airfoil as a function of percent cord rather than CL vs CD charts. Unless you are trying to design an airfoil or understand what's accounting for the behavior of a specific airfoil, the charts of velocity vs. cord are more of an esoteric curiosity.

Can you attach one of the polars you are trying to understand so that we can give a more focused answer?

[Sail 'n Soar]

Similarly, Drag = Lift = CL *A* V*V*rho/2.

OOps, my cut and paste did me in. It should have read Drag = CD*A*V*V*rho/2 :o

[Ollie]

Farther down the Google search I found this 13 page PDF download that should be just what you are looking for:
http://www.google.com/url?sa=U&start=17&q=http://www.dreesecode.com/other/aflprimer.pdf&e=7830

[sloper steve]

I have made templates for flying wings which have produced a plane I like using a manual hot wire cutter.

Going from templates to CNC is a pretty big leap. Today the CNC followed wing paths and foamworks software seems to work good .

I have scanned in the templates for my flying wings and am eager for criticism then improvement/replacement.

They have been turned into .dat files using Profili. The end points are not closed and the leading edge could use smoothing.

There are some close airfoils in the database that can replace it but for passion of my current plane and comparison to the coming cuts I'd like to make one.

First they need to be cleaned up.
Im assuming I'd use profili to move the points by hand and also make the trailing edge sharp. Next use the smooth operations?

I have no problem posting the airfoils up bashing :)

Thanks for the help

[Ollie]

Don't throw away the inaccurate first core. Fill in the low spots around the leading edge with spackle and sand them smooth to near the desired contour. Cut the core bed in two for an upper and lower half. You can sand the trailing edge contour to a feather edge using a flat sanding block and very, very light sanding pressure while holding the core in one of its beds. It is slow work but you can get a functional core out of it.

[sloper steve]

It is all EPS and the nose is rather done :)

The templates that made it have been turned to dat files already.

I've tried profili replacements and get an eppler hydrofoil and am working on a way to smooth it out.

[Antonsoarer]

Steve, I have had similar problems getting rid the bumps in my scans. Can you just visually check candidates from profili by printing and comparing with your templates, by laying template on correctly sized printout?

I have done this to check out some of my 25 year old home-brew sections. It was amazing how close they were to existing ones, not very original or radical :(

Is that your site? Looks good, full of useful stuff.

Tony.

[JMP_blackfoot]

The problem is intermittent contact between carbon and/or metal parts.
One alternative option we found successful is to electrically connect the moving parts together.
Use thin flexible wire, scrape the carbon piece, wrap the flex wire around and apply conductive paste and let dry. We had the same problem with our Twisteuse horten wing, where the carbon tube lead edge rotates in a carbon tube bearing. We also applied some silicon grease to reduce friction between the carbon parts.

Sorry, wrong thread :-))