




is this the L/D plot of an aircraft obtained from the total drag of the aircraft (airfoils and body, parasitic and induced) versus lift and hence airspeed, or simply the Cl vs Cd curve for a particular airfoil? doesn't the former quite a lot of information about the aircraft and some experimental/guessimated data, and the later require wind tunnel data?



United States, CA, San Marcos
Joined Sep 2005
129 Posts

sligo565,
A drag polar generally refers to a plot of aircraft lift (vertical axis) versus drag (horizontal axis). The equation that you have written is part of the induced drag (drag due to lift), but it is incomplete. The induced drag is your equation, typically called k, times the lift coefficient squared. To compute the full drag polar, you need need to take into account all components of drag, primarily parasitic drag (form drag plus skin friction and interference) plus induced drag (drag due to lift), as well as have a good understanding of the lift curve of the design, which takes into account the 3d effects of a finite span on the slope of the lift curve, along with other factors. There are some decent websites available that can help you develop a first order model of your design, which will estimate the items above based on the geometry of your design. Hope that helps. Rich 


Joined Feb 2008
44 Posts

Hey,
I used the software JAVAFOIL to produce the following graph and numbers..do i just measure off the graph at the area of zero Lift? It's just that its giving me a very high number of 0.07....can i just use the equation: CD = CDmin + (CL  CLmindrag)^2/pi e A To find my Drag Coefficient instead of using this such high number of CDo?! 


Joined Feb 2008
44 Posts

Here is also a drawing of my conceptual design....its design is much like a sailplane....does anyone know an estimation of CDo of a sailplane that i could use?



United States, CA, San Marcos
Joined Sep 2005
129 Posts

Quote:
Rich 



Joined Dec 2012
1 Posts

I am trying to work out the drag polar of the Citation X, I have the following data. what I am confused about is that I when I wanted to work the Cd = Cd0 + KCl*2, I dont know the Cd0 and K , can someone please help me out with this. How do I find the Value of Cd0 and K>
MANY THANKS CL=2L/(ρ V^2*S) ρ=1.2252kg/m^3 V=146m^2 S = 50m^2 basic operating weight = 10189kg Max payload = 5865kg Max Fuel = 11330kg W_max=27384kg Cl= mg/(1⁄2ρ V^2 S) Cl=27384/(0.5*1.2252〖146〗^2*50)=0.4114463227 Drag = CD 1⁄2 ρV^1 S CD= CD0+K〖CLČ〗 



Quote:





I would suggest you use a program such as xfoil or avl to compute that. There are many things going on in that design that the simple equations you are using have trouble modeling. Even just a simplified planform in AVL will give you better accuracy and if you spend enough time building your model in AVL you can learn other interesting things about your design such as stability and trim states. Another program which might be nicer to a novice is XFLR5, its basically the same as AVL and xfoil but offers a GUI instead of a terminal based UI.
http://www.xflr5.com/xflr5.htm http://web.mit.edu/drela/Public/web/xfoil/ http://web.mit.edu/drela/Public/web/avl/ 