RAF-6 Aerodynamics data

[mohr]

Hi...
I'm looking for aerodynamic data of the old RAF-6 airfoil, used in the propeller blade sections.
The only information source I have are the Naca TR-319 and TR-207.
Does anyone have knowledge of any more modern source of information to obtain data for different relative thicknesses?
The purpose of this request is to validate a program that I made for calculating propellers, based on the vortex theory.
Regards....

 

[Sparweb]

AFAIK air has the same properties as it did in the 1930's (just a wee bit more CO2).
Have you visited the Airfoiltools website? UIUC?
Take a look at Tech Report 640, it may better suit your needs. And TR-658... there are many more.
Why are you using the RAF-6 profile in your propeller analysis if you feel the available data is unsuitable?


No one believes the theory except the one who developed it. Everyone believes the experiment except the one who ran it.
STF

 

[mohr]

Hi spar web:
Tanks for your reply.
The reason for this data requirement was because I made a computer. program based on the ARC-R&M 1674 for the aerodynamic characteristics calculation of propellers blades and I wish to validate the results i.e. using the naca tr-378 specific for two-blade propellers. (CT. CP and efficiency). In this tr-378 are used clark-y and raf-6 airfoils.
I found only two reports about Clark-y and Raf-6 airfoil sections in NTRS : NACA´s TR-207 and TR-319. (nowhere else have I got them)
Both reports have appreciable difference between the aerodynamic results.
unfortunately UIUC only provides profile coordinates.

Anyway, I appreciate your reply.
Regards....

 

[Sparweb]

Try this:

http://lmgtfy.com/?q=NACA+technical+report+640

No one believes the theory except the one who developed it. Everyone believes the experiment except the one who ran it.
STF

 

[LiftDivergence]

mohr,

Could you possibly use the coordinates you can find on the UIUC datasite with XFOIL to get the results you are after?

The website Sparweb mentioned, airfoil tools, has a pretty convenient write-up on how to do basic runs in XFOIL (command prompts in caps):

http://airfoiltools.com/polar/index

LOAD dat/e1211-il.dat Load the dat file
MDES Go to the MDES menu
FILT Smooth any variations in the dat file data
EXEC Execute the smoothing
Back to main menu
PANE Set the number and location of the airfoil points for analysis
OPER Go to the OPER menu
ITER 70 Max number of iterations set to 70 for convergence
RE 50000 Set Reynolds number (required?)
VISC 50000 Set viscous calculation with Reynolds number
PACC Start polar output file
polar/e1211-il_50000.txt The output polar file name
No dump file
ALFA 0 Calculate lift and drag at 0° angle of attack
ALFA 0.25 ... 0.25°
ALFA 0.5 ... 0.5° ...
... ...more alpha calculations here ...
ALFA 3.5 At 3.5° no convergence
ALFA 3.5 ... try again ...
ALFA 3.5 ... and again
INIT Run INIT to reinitialise
ALFA 3.75 Skip to 3.75°
... ...rest of alpha calculations here ...
PACC Close polar file
VISC Reinitialise viscous calculation (required?)
Down to main menu
QUIT Exit Xfoil

XFLR5 will also give you Cp. Might be quicker depending on how long it would take to search it up.

Keep em' Flying
//Fight Corrosion!