Light Weight Diving Vehicle

Rick

Thanks for the tip. That thought did cross my mind once, but I wonder why I didnt worl on it. If I'm to slant the wings, I'll be needing the lift and drag coefficient data of the airfoil at different angles of slanting. I have the data for NACA 0015
http://www.aerospaceweb.org/question/airfoils/q0150b.shtml
Also, what do you think would be the Reynolds number for a speed of .6 m/sec?

Daniel

 

And yes, I do understand that you must have a little less time to devote now. And thank you for all the effort you are putting into the replies.
I really appreciate it.

 

And also Rick, how do I calculate the frontal area of the NACA 0015 aerofoil?

Right now, I am just calculating the projected area of the aerofoil, as in the trapezoidal area, with the top as 15 % of the top edge chord and the bottom as 15% of the bottom edge chord. with the span as the height. Is it correct?

 

Daniel
The area that is normally used for drag coeeficients on foils is the planar surface not the frontal area. Generally drag and lift coeeficients relate to this area. A rough estimation in the span times the chord. If it tapers then use the average chord.

Rick W

 

Rick,

The drag you are talking about is the drag for rotation right? :?:

I was thinking that for the lift, we will have to use the frontal area(with the airfoil(at an angle) advancing into the fluid with the frontal area.) Please correct me if I'm wrong.

Daniel

 

Daniel
No - The drag coefficient applies to the area parallel to the flow. Even a thin plate that has negligible thickness perefectly aligned with the flow will have drag. Hence the cefficient is applied to the planar surface, not the frontal area.

The force that prevents rotation is the lift.

If you have a NACA0010 foil that sticks out 4" and is 1" long angled at 5 degrees to the line of travel it will have a Cl of 0.38 and a Cd of 0.033 at Re# of 20,000.

To determine the rotating force you apply the Cl to the 4 square inches. To determine the resistance to motion you apply the Cd to the same 4 square inches. Nothing to do with frontal area in this calculation.

Rick W

 

Rick

Thanks for clarifying this.

Also, will the Cl and Cd remain same if I change the dimensions of the fin, keeping the aspect ratio same?

Daniel

 

Rick,

Also, can I please get the link to the source from where you got this NACA0010 data. I have been searching the internet like crazy but havent got anything so far.

Thank you

Daniel

 

Google JavaFoil and run the Applet. Select foil to NACA 4 series with 10% thickness. You need to set the option page for water and an aspect of 4. Then plot the polar curves.

The Re# for 0.6m/s on a 1" foil is about 20,000.

There are data bases with measured data but they will be for a foil with infinite length. You then have to apply the induced drag which is a bit complicated. JavaFoil will do this for you.

Rick W.

 

Rick

What settings did you exactly use for JAVAfoil? Because, I'm getting the value of Cl = 0.31 and Cd = 0.58.


The settings I'm using are:
Kinematic Viscosity = 10e-006 sq-m/sec
Velocity_sound = 1492 m/s
Mach = 0.58/1492
AR = 4

And on the POLAR tab,

TU = TL = 100 %


Daniel

 

And the stall model: CalcFoil
And Transition model : Eppler

 

Rick,

And also if I change the dimensions of the Aerofoil to

Span = 8 cm
Chord = 2 cm

How will the Reynolds number change?
Can you tell me how to calculate it depending on different dimensions?

 

Daniel
Your value of Cd is missing a decimal point. Otherewise I agree with the result. The figure given early was in fact for Re# of 40,000. I pressed the wrong tab.

Re# is a function of viscosity, density, velocity and the length in the direction of flow. So for 2cm it will be smaller. Also I did not actually calculate the 20,000 I just guessed that value as being in the ballpark. You can get the formula from many sites.

You will note that the L/D ratio deteriorates as the Re# reduces.

JavaFoil allows you to play with values of Re#, aspect ratio and angle of attack to see what gives the best result. You can also try different foils. The best will be one with camber in your case. So try about 4% camber and see how that goes.

Also note that you might need a different shape if the craft travels in both directions.

The other things you asked about regarding JavaFoil do not matter much for the operating conditions you have.

Rick W.

 

Thank you Rick.

 

Dear Rick,

I am using this formula to calculate the Reynolds number. Is this correct?

Rey. Num. = density*velocity*chord length/abs. viscosity

Daniel