So I am trying to extract aerodynamic data (Cm,Cn,Cl,CL,CN,CD) out of a CAD-model in both different aerodynamic angles and in different Mach's. Is there any tutorial or something out there that I have missed? I've been through the official ones (drag coefficient for the car and supersonic simulation for the bullet) but doing these generates data that doesnt match the data I already have from real wind-tunnel tests. Sometimes it matches somewhat but not always. I am currently wondering about the Torque-axis in the wall-calculator, is this to be set to 0 or use the suggested values to have it measure in the middle of the model? Whats odd is that changing the velocity gives different standardized coefficient. The formula I use is: Coefficient=Total Moment/(Dynamic Pressure * Real aircraft area * Total area as shown in Wall Calculator)
Hi RagnarDa !
I faced the same problem as Autodesk Simulation CFD doesn't provide direct aerodynamics data like Cd or Cl like ANSYS Fluent. You have to calculate those from Drag force and Lift Force using wall calculator of Simulation CFD.
I found this pdf that i attached helpful.
Hope it would help. Thanks!
The front area number is usually the chord length (in the flow direction) of the body or air-foil.
Hope this helps.
thanks! actually this sketch is not mine,took it from a video used in sustainabilityworkshop by autodesk!
Thanks for all the replys! The replies didn't help me though. When using the solutions that is suggested I get completely erronious results. An example:
If I test my model at 30 alpha 0 beta ~around Mach 0.5 (171.605 m/s) at sea level conditions i get the following results in the wall calculator (I removed the six walls of the bounding box from the calculations):
Total area 224.348 m^2
TOTAL FX -101726 Newton
TOTAL FY 2769050 Newton
TOTAL FZ -9520.9 Newton
Center of Force about X-Axis (Y-Z) 5.52947 6.87737 m
Center of Force about Y-Axis (X-Z) 1.19719 6.8917 m
Center of Force about Z-Axis (X-Y) 0.121217 5.70515 m
TOTAL MX -19096400 N-m
TOTAL MY -689668 N-m
TOTAL MZ 916020 N-m
Torque 916020 N-m
I am unsure which one of the moments is the pitch but if I calculate using MX (???) which is largest I get:
q = 0.5 * 1.2041 * (171.605 * 171.605) = 17729.3345808512
Cm = -19096400 / (q * 43 * 7.4) = -3.3850017048
which is far from my reference which is a wind tunnel testing data and is around -0.075. The 43 is the wing area and 7.4 is the chord length.
What is also odd is that if I change the velocity I get yet other values.
Note that you must be using the same moment reference axis to match published data. Most airoil data where a pitching moment is given has the reference at quarter chord; some use the leading edge (see notes here, for example: http://www.dept.aoe.vt.edu/~lutze/AOE3104/airfoilw
In SimCFD you do this in the wall calculator by selecting "Torque" and assigning a point on the centerline at the quarter chord location, and assign the axis to point in the direction of the pitch axis.