Dear Prof. Gudmundsson, thank you for your excellent work and channel, which I follow. At minute 3 you say "I want to keep this video short so that you will bother to watch it". First of all, watching your videos is not a bother but a pleasure, second, your audience is interested in the topic enough to watch a longer video. Please don't cut your videos short if you have something to say :)
Dear Dr. Gudmundsson, really was a pleasure to assist your videos, them are extremely clear and very well explained. Your book is an obligatory read for any one interested in aircraft design, I do spend countless hours reading and reading again, there is something new everything that I read. Please do not be worry about make longer videos, for us it is a pleasure to assist them. I hope someday have the honor to meet you personally.
Good day Prof.Snorri, thanks for your time sacrifice and help you've been giving sir. I ran a CFD analysis on a 3D airplane, i got the drag force, but am trying to get the area used so i can calculate CD. In the calculation of total 3D drag coefficient of an entire airplane, which area is used for calculations sir? do we use the reference area or projected area of the airplane relative to flow direction i.e flow direction is along the Zaxis and projected area is in the XandY direction or we use the total wetted area of the entire airplane or the planform area of the airplane. Thank you sir.
Hi Israel. I'm happy to help. For an airplane, always use the reference area. This is the planform area of the wing, from wingtip to wingtip including the area inside the fuselage. I hope this response is clear. Best wishes.
@@dr.gudmundssonaircraftdesign good day prof. Does this means the planform area of the fuselage and horizontal stab is not taken into account? Thanks sir.
@@israeloluwagbemi825 Yes, coefficients such as CL, CD, and CM always refer to the wing planform area, S. Thus, drag coefficient is CD=2D/(rho∙V²∙S). Of course, the fuselage, HT and VT contribute to the drag and, thus, make D larger. This contribution is captured when you calculate CD, even though their surface areas are omitted.
@@dr.gudmundssonaircraftdesign Thank you very much prof, am very grateful 🙏, I've been looking for this perfect explaination for a very long time. Thanks once more sir.
Good day Prof. Concerning the adjusted drag model, is CLminD for the 2Dairfoil section or the 3D wing... I tried calculating with the 2D airfoil, in this case i used a Clark Y airfoil of 13% thickness, the value i got for CLminD is extremely close to that of CL at 0° AOA let say 0.39 and 0.387 respectively at Re=4,700,000. When i used a lower Re of say 1,000,000 i got results that were far apart 0.381 and 0.5 . are the results Ok. thanks alot for the video sir.
Hi Israel. We have both Clmind (airfoils) and CLminD (3D geometry like wing or wing+fuselage, etc.). Your values sound reasonable, but I haven't checked specifically the Clark Y airfoil. Best wishes.
Dear Prof. Gudmundsson, thank you for your excellent work and channel, which I follow.
At minute 3 you say "I want to keep this video short so that you will bother to watch it". First of all, watching your videos is not a bother but a pleasure, second, your audience is interested in the topic enough to watch a longer video. Please don't cut your videos short if you have something to say :)
Wow, that's very nice of you to say. I'll definitely keep your words in mind. :-)
Yeah. This is it
Dear Dr. Gudmundsson, really was a pleasure to assist your videos, them are extremely clear and very well explained. Your book is an obligatory read for any one interested in aircraft design, I do spend countless hours reading and reading again, there is something new everything that I read. Please do not be worry about make longer videos, for us it is a pleasure to assist them.
I hope someday have the honor to meet you personally.
Thank you for the very kind words, Eduardo. I appreciate them very much. :-)
Very informative and well produced professor! I'm looking forward to applying this knowledge.
Thank you, Lucas. I appreciate it :-)
Hello Sir !
I will be grateful if you discuss the topic of control surface sizing in a video.
Hi Muhammad. Thanks for the suggestions. I'll consider adding such a video for sure. Best, Dr. Gudmundsson.
Good day Prof.Snorri, thanks for your time sacrifice and help you've been giving sir. I ran a CFD analysis on a 3D airplane, i got the drag force, but am trying to get the area used so i can calculate CD. In the calculation of total 3D drag coefficient of an entire airplane, which area is used for calculations sir? do we use the reference area or projected area of the airplane relative to flow direction i.e flow direction is along the Zaxis and projected area is in the XandY direction or we use the total wetted area of the entire airplane or the planform area of the airplane. Thank you sir.
Hi Israel. I'm happy to help. For an airplane, always use the reference area. This is the planform area of the wing, from wingtip to wingtip including the area inside the fuselage. I hope this response is clear. Best wishes.
@@dr.gudmundssonaircraftdesign thanks Prof, am very grateful 🙏.
@@dr.gudmundssonaircraftdesign
good day prof. Does this means the planform area of the fuselage and horizontal stab is not taken into account? Thanks sir.
@@israeloluwagbemi825 Yes, coefficients such as CL, CD, and CM always refer to the wing planform area, S. Thus, drag coefficient is CD=2D/(rho∙V²∙S). Of course, the fuselage, HT and VT contribute to the drag and, thus, make D larger. This contribution is captured when you calculate CD, even though their surface areas are omitted.
@@dr.gudmundssonaircraftdesign
Thank you very much prof, am very grateful 🙏, I've been looking for this perfect explaination for a very long time. Thanks once more sir.
amazing i need 1 book. hace you in spanish language?
Thanks, Hugo. No, unfortunately the book has not been translated to spanish. Sorry about that.
Awesome👌👌👍👍✈ ✈ 🙏🙏
Thanks ✌️
Good day Prof. Concerning the adjusted drag model, is CLminD for the 2Dairfoil section or the 3D wing... I tried calculating with the 2D airfoil, in this case i used a Clark Y airfoil of 13% thickness, the value i got for CLminD is extremely close to that of CL at 0° AOA let say 0.39 and 0.387 respectively at Re=4,700,000. When i used a lower Re of say 1,000,000 i got results that were far apart 0.381 and 0.5 . are the results Ok. thanks alot for the video sir.
Hi Israel. We have both Clmind (airfoils) and CLminD (3D geometry like wing or wing+fuselage, etc.). Your values sound reasonable, but I haven't checked specifically the Clark Y airfoil. Best wishes.
@@dr.gudmundssonaircraftdesign thank you sir, am very grateful for your help, i really appreciate your videos, they've been helpful.
@@israeloluwagbemi825 Thank you for your kind words. Best wishes.