Hey Josh, Thank you for the really great videos. They explain things very clearly and It's pretty hard to find video's on these topics that are good. Your doing a great job! Kepp it up!
That's nice. Thank you for your videos. I'm working on airfoil optimization. I'm having some difficulties on using the softwares, cause I've never used MatLab or Xfoil, but I'm learning. I guess my idea is simple: I want to change the geometry of a 4-digit airfoil to get more specific and appropriate coefficients (cl and cd mainly). But I wanna take them as inputs. Do you have any idea of where I could start from? How can I make these two softwares get in touch?
Hey Josh, How can I get the profile equation? I want to find the pressure distribution and for that, I need the streamfunction (which is the airfoil profile) to find the velocities and then plug it into Bernoulli to find the pressure distribution. Let me know if there is a more straightforward method. Thanks.
Thank you! Yes, you can plot the pressure distribution if you use something called the Vortex Panel Method. That's a video I plan on making in the future, but there are already some good books and tutorials out there for how to code panel methods.
JoshTheEngineer Thanks for your reply.....Actually the eqn for pressure distribution is in differentiation from unlike aerofoil thickness eqn...Please make a video on it...Thanks
+JoshTheEngineer Thanks for this great video! Had a quick question though, how did you determine the constants (lines 23-27). I know you mentioned that they were from the other video but I still don't know from where you got them. Thanks in advance and nice job with the videos!
Thanks for watching! So first of all, if you don't already have it, I recommend getting the book Theory of Wing Sections by Abbott and Von Doenhoff if you're interested in wings/airfoils etc. In the section about the NACA 4-digit airfoils (section 6.4, page 113) they state the following: "When the NACA four-digit wing sections were derived, it was found that the thickness distributions of efficient wing sections such as the Gottingen 398 and the Clark Y were nearly the same when their camber was removed (mean line straightened) and they were reduced to the same maximum thickness. The thickness distribution for the NACA four-digit sections was selected to correspond closely to that for these wing sections and is given by the following equation" The equation stated is Eq. 6.2 on page 113, which includes these seemingly arbitrary coefficients that you're asking about. From the quoted selection, it seems like these coefficients were chosen to closely match the thickness distributions from other airfoils that were known to work well. You can mess around with the values and see how they change the geometry, keeping everything else the same. You can also load in the two airfoils they mention, remove their camber and compare the thickness distributions from these airfoils. It could be an interesting, informative experiment.
JoshTheEngineer Sorry my bad. Works fine. Thanks for the reply! Love all your videos. I usually stop by your videos for reference. You in the industry?
The best way to verify is to go to this website: m-selig.ae.illinois.edu/ads/coord_database.html You can download the coordinate files for the airfoils and check them against the ones produced by my code. If you choose the same number of points, the values should be identical (which I believe I check in this video for a certain airfoil). If you increase the number of points in my code, the shape should still be the same, it will just have a higher resolution.
Thanks! I'm sure it could be done, but I'm not familiar with the formatting of those files. For the CAD program though, you should be able to open up the text file with too much of a problem.
JoshTheEngineer yes by exporting the xls format is not a very big problem working with them, svg or dxf are only easier to work with :-) Thank you very much
Thanks a lot for video. İts helped to me make a decision shape. İf you have got videos for Naca5, Eppler and Gottingen you can share with us. Thank you.
You're welcome! In a semi-recent video that I posted, I showed how you can download the UIUC airfoil database, and then also how you can read those airfoils into MATLAB or Python. You might want to check out those two videos (ruclips.net/video/nILo18DlqAo/видео.html and ruclips.net/video/xJYxMfGFrk8/видео.html)
You're welcome! Since the code already calculates the camber line, you can just use the following code to plot it: plot(x,yc,'k-'); This plots a black line for the camber line.
Hello @JoshThe Engineer your video is so amazing and meaning full . But I am getting some problem while I inserting same code as yours , I got Error like Expected H to be a vector of handles to one or more MATLAB graphics object . Can you please help me out . Thanks and waiting for your positive response 👍🏻
You could use the xlswrite function (which apparently isn't recommended anymore, see the documentation), or you could use fprintf, which is what I typically do.
Hi :) First of all thanks for the video, it is very good. But i have a question. When i checked the wikipedia page of 4-digit airfoils, equations have chord length in it too. But you didn't use it. How could you do that or did i miss something? Thanks.
You're right, the Wikipedia equations have the chord length, 'c', in them. Now just plug in a value of one (1) for chord length into those equations, and they should be the same as the ones I'm using. It's really just a non-dimensionalized version of the Wikipedia page. Here's why the way I write it might be better. You can find the correct airfoil you want to use, and output the data coordinates just once. Then if you have a program that puts them together into a swept wing (for example, because it has varying chord length), then you can just multiply the X and Y values in the coordinate file by the chord length and you're good to go. You'll see in the airfoil coordinate files that the X values go from 0 to 1 (technically from 1 to 0 to 1, but that's to define both the upper and lower surfaces). Multiplying by the desired chord length makes the X values go from 0 to 'c', and since the thickness (Y values) was also defined with respect to a non-dimensional chord length of 1, they will change accordingly.
Thank you so much, that's a good idea. Actually i remembered we can use those coordinates in other programs as we took from matlab. Also i finished the series and that project was one of our teaching assistans suggested to me. Because i am an aerospace engineering student and i asked to him to how i can learn matlab depending on my education.Thank you so much again. And i want to ask if you have any suggestion to me like this project.
The best way to learn MATLAB is to just keep using it. It doesn't necessarily have to be aerospace related. Think of something that might be cool to have automated, and code it up. For example, I'm usually indecisive when picking an episode of my favorite TV series to re-watch. So I decided to make a random episode picker. Then I expanded it to all the shows I watch. Then I put it into a GUI. If you're looking for a couple aerospace related MATLAB programs, check out my GitHub account here: github.com/jte0419 In particular, the Rocket Nozzle Design and Taylor Maccoll codes are applicable to aerospace engineering. I'll be doing videos about the derivation and coding for those in the future, but you might be able to figure them out on your own.
hoa le dang You could take the basic outline of the code and use a language like C++ or Java if you are more familiar with them. You should also be able to make a spreadsheet in Excel that will calculate and plot an airfoil of your choosing.
Thank you for your videos. i am working on conformal mapping of an airfoil. I am having some difficulties i am not getting how to map NACA airfoil. Please could you help it.
You're welcome! I had planned on doing some conformal mapping videos in the future, but I don't know when that will be. The book Fundamental Fluid Mechanics by Currie has some good information on this topic though.
I'm not sure I can help you with this, as I've never looked into it before. You can probably take a NACA 0012 airfoil, change the chord length in some sort of sinusoidal manner, and then extrude along the wing span between the various airfoils. That should get you started.
hi! I loved your videos but I have a little question if I change my length chord, I mean I chose 1.47 instead 1 (m,ft) How can I fix it? I tried to change the limits and multiply by 1.47 but I feel that is something with ' ones' thanks for reading.
Premkumar P Pithiya Sorry, I didn't put units on the axes. It's just a unit length though. If you watch the video I posted in the description, I briefly mention that I'm using a length of one unit for the chord, and then the thickness is calculated accordingly. To get a different chord length, just multiply the X and Y values by whatever you want it to be.
![Explained: NACA 4-Digit GUI Part 1/10 [MATLAB]](http://i.ytimg.com/vi/iQ1-C6c3YyA/mqdefault.jpg)
![Explained: NACA 4-Digit GUI Part 1/10 [MATLAB]](/img/tr.png)
![Explained: NACA 4-Digit Airfoil [Airplanes]](http://i.ytimg.com/vi/6pt8Uolfj0M/mqdefault.jpg)
Hey Josh, Thank you for the really great videos. They explain things very clearly and It's pretty hard to find video's on these topics that are good. Your doing a great job! Kepp it up!
No problem! That's what I'm here for!
That's nice. Thank you for your videos.
I'm working on airfoil optimization. I'm having some difficulties on using the softwares, cause I've never used MatLab or Xfoil, but I'm learning. I guess my idea is simple: I want to change the geometry of a 4-digit airfoil to get more specific and appropriate coefficients (cl and cd mainly). But I wanna take them as inputs. Do you have any idea of where I could start from? How can I make these two softwares get in touch?
Help me please,
My program says Index Exceeds Matrix Dimension. Although I copied everything from your code, it says that error.
Hey Josh,
How can I get the profile equation? I want to find the pressure distribution and for that, I need the streamfunction (which is the airfoil profile) to find the velocities and then plug it into Bernoulli to find the pressure distribution. Let me know if there is a more straightforward method. Thanks.
Just what I was looking for...Ur explanation is v. Good....Can you pls suggest is it possible to programme Pressure distribution over the Aerofoil..
Thank you! Yes, you can plot the pressure distribution if you use something called the Vortex Panel Method. That's a video I plan on making in the future, but there are already some good books and tutorials out there for how to code panel methods.
JoshTheEngineer Thanks for your reply.....Actually the eqn for pressure distribution is in differentiation from unlike aerofoil thickness eqn...Please make a video on it...Thanks
It says INDEX EXCEEDS MATRIX error. 😭😭😭
+JoshTheEngineer Thanks for this great video! Had a quick question though, how did you determine the constants (lines 23-27). I know you mentioned that they were from the other video but I still don't know from where you got them. Thanks in advance and nice job with the videos!
Thanks for watching! So first of all, if you don't already have it, I recommend getting the book Theory of Wing Sections by Abbott and Von Doenhoff if you're interested in wings/airfoils etc. In the section about the NACA 4-digit airfoils (section 6.4, page 113) they state the following:
"When the NACA four-digit wing sections were derived, it was found that the thickness distributions of efficient wing sections such as the Gottingen 398 and the Clark Y were nearly the same when their camber was removed (mean line straightened) and they were reduced to the same maximum thickness. The thickness distribution for the NACA four-digit sections was selected to correspond closely to that for these wing sections and is given by the following equation"
The equation stated is Eq. 6.2 on page 113, which includes these seemingly arbitrary coefficients that you're asking about. From the quoted selection, it seems like these coefficients were chosen to closely match the thickness distributions from other airfoils that were known to work well.
You can mess around with the values and see how they change the geometry, keeping everything else the same. You can also load in the two airfoils they mention, remove their camber and compare the thickness distributions from these airfoils. It could be an interesting, informative experiment.
Hey Josh, I followed your code step by step and am not getting the airfoil as expected. I’m getting a half weird square
If you're using the same code as mine, then I don't know what would be going wrong. I would have to see the code to be able to check it for sure.
JoshTheEngineer Sorry my bad. Works fine. Thanks for the reply! Love all your videos. I usually stop by your videos for reference. You in the industry?
Thanks for the video. I have a question. Is there a way to verify if the aerofoil shape generated is correct or not?
The best way to verify is to go to this website: m-selig.ae.illinois.edu/ads/coord_database.html
You can download the coordinate files for the airfoils and check them against the ones produced by my code. If you choose the same number of points, the values should be identical (which I believe I check in this video for a certain airfoil). If you increase the number of points in my code, the shape should still be the same, it will just have a higher resolution.
can you wrire a program to caculation simulation aerodynamics of naca airfoil 2d by matlab ???
This is really an amazing program! Can you add the option of saving the airfoil as .svg or .dxf? Thank you very much
Thanks! I'm sure it could be done, but I'm not familiar with the formatting of those files. For the CAD program though, you should be able to open up the text file with too much of a problem.
JoshTheEngineer yes by exporting the xls format is not a very big problem working with them, svg or dxf are only easier to work with :-) Thank you very much
Thanks a lot for video. İts helped to me make a decision shape. İf you have got videos for Naca5, Eppler and Gottingen you can share with us. Thank you.
You're welcome! In a semi-recent video that I posted, I showed how you can download the UIUC airfoil database, and then also how you can read those airfoils into MATLAB or Python. You might want to check out those two videos (ruclips.net/video/nILo18DlqAo/видео.html and ruclips.net/video/xJYxMfGFrk8/видео.html)
Thanks!
Glad I found this channel :)
Contarius9
You're welcome! Thanks for watching.
Hello. Thank you for your video! How can I plot the camber line?
You're welcome! Since the code already calculates the camber line, you can just use the following code to plot it: plot(x,yc,'k-'); This plots a black line for the camber line.
Hello @JoshThe Engineer your video is so amazing and meaning full . But I am getting some problem while I inserting same code as yours , I got Error like
Expected H to be a vector of handles to one or more MATLAB graphics object .
Can you please help me out .
Thanks and waiting for your positive response 👍🏻
Help me please.
How to export that airfoil coordinate from Matlab to excel ??
You could use the xlswrite function (which apparently isn't recommended anymore, see the documentation), or you could use fprintf, which is what I typically do.
Hi :) First of all thanks for the video, it is very good. But i have a question.
When i checked the wikipedia page of 4-digit airfoils, equations have chord length in it too. But you didn't use it. How could you do that or did i miss something?
Thanks.
You're right, the Wikipedia equations have the chord length, 'c', in them. Now just plug in a value of one (1) for chord length into those equations, and they should be the same as the ones I'm using. It's really just a non-dimensionalized version of the Wikipedia page.
Here's why the way I write it might be better. You can find the correct airfoil you want to use, and output the data coordinates just once. Then if you have a program that puts them together into a swept wing (for example, because it has varying chord length), then you can just multiply the X and Y values in the coordinate file by the chord length and you're good to go. You'll see in the airfoil coordinate files that the X values go from 0 to 1 (technically from 1 to 0 to 1, but that's to define both the upper and lower surfaces). Multiplying by the desired chord length makes the X values go from 0 to 'c', and since the thickness (Y values) was also defined with respect to a non-dimensional chord length of 1, they will change accordingly.
Thank you so much, that's a good idea. Actually i remembered we can use those coordinates in other programs as we took from matlab.
Also i finished the series and that project was one of our teaching assistans suggested to me. Because i am an aerospace engineering student and i asked to him to how i can learn matlab depending on my education.Thank you so much again. And i want to ask if you have any suggestion to me like this project.
The best way to learn MATLAB is to just keep using it. It doesn't necessarily have to be aerospace related. Think of something that might be cool to have automated, and code it up. For example, I'm usually indecisive when picking an episode of my favorite TV series to re-watch. So I decided to make a random episode picker. Then I expanded it to all the shows I watch. Then I put it into a GUI.
If you're looking for a couple aerospace related MATLAB programs, check out my GitHub account here:
github.com/jte0419
In particular, the Rocket Nozzle Design and Taylor Maccoll codes are applicable to aerospace engineering. I'll be doing videos about the derivation and coding for those in the future, but you might be able to figure them out on your own.
That was nice. I will try to do some basic things which comes to my mind. Thank you for your kind answer.
thanks for your video. but where can i take the modul likes yours, because i don't know much about matlab , thank you
hoa le dang
You could take the basic outline of the code and use a language like C++ or Java if you are more familiar with them. You should also be able to make a spreadsheet in Excel that will calculate and plot an airfoil of your choosing.
thank you so much
Thank you for your videos.
i am working on conformal mapping of an airfoil. I am having some difficulties i am not getting how to map NACA airfoil. Please could you help it.
You're welcome! I had planned on doing some conformal mapping videos in the future, but I don't know when that will be. The book Fundamental Fluid Mechanics by Currie has some good information on this topic though.
How do you plot a 3 digit airfoil?
sir please provide me the code for how to make 3d wavy trailing edge for naca0012 aerofoil.
I'm not sure I can help you with this, as I've never looked into it before. You can probably take a NACA 0012 airfoil, change the chord length in some sort of sinusoidal manner, and then extrude along the wing span between the various airfoils. That should get you started.
can you please send me the sample code for that so that i can start
hi! I loved your videos but I have a little question if I change my length chord, I mean I chose 1.47 instead 1 (m,ft) How can I fix it? I tried to change the limits and multiply by 1.47 but I feel that is something with ' ones'
thanks for reading.
Yo te ayude wey
Please explain 5 digit and six digit NACA series with program....
Those videos are on my to-do list. It's been a crazy summer, but I'll be back to making videos soon.
whats on x-axis & Y-axis?
Premkumar P Pithiya
Sorry, I didn't put units on the axes. It's just a unit length though. If you watch the video I posted in the description, I briefly mention that I'm using a length of one unit for the chord, and then the thickness is calculated accordingly. To get a different chord length, just multiply the X and Y values by whatever you want it to be.