There is no polar graph about Aero derivatives with respect to body rates (p,q,r). Also in the stability analysis Lateral derivatives
Yv= -0.23503 CYb= -0.08318
Yp= -0.11732 CYp= -0.055361
Yr= 0.17284 CYr= 0.08156
Lv= -0.15326 Clb= -0.03616
Lp= -1.7779 Clp= -0.55931
Lr= 0.15593 Clr= 0.049055
Nv= 0.14247 Cnb= 0.033616
Np= -0.097373 Cnp= -0.030633
Nr= -0.092417 Cnr= -0.029074
there is only one Clp . I could not define analysis for different AOA for stability analysis.
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I still don’t get what you use as control parameter. A common choice is the elevator angle, when you do a sweep across control values this gives several different AoAs.
There is no polar graph about Aero derivatives with respect to body rates (p,q,r). Also in the stability analysis Lateral derivatives
Yv= -0.23503 CYb= -0.08318
Yp= -0.11732 CYp= -0.055361
Yr= 0.17284 CYr= 0.08156
Lv= -0.15326 Clb= -0.03616
Lp= -1.7779 Clp= -0.55931
Lr= 0.15593 Clr= 0.049055
Nv= 0.14247 Cnb= 0.033616
Np= -0.097373 Cnp= -0.030633
Nr= -0.092417 Cnr= -0.029074
there is only one Clp . I could not define analysis for different AOA for stability analysis.
Thanks for your answer. i dont use any control parameter . For example in the polar i can easily find the CL with respect to alpha . But there is no CL with respect to p ,q or r . I want to find for example Clp for different AOA without using control input.
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The reason why you have CL/alpha in Type 2 polars is that it is a parametric analysis on alpha. In stability analysis, you get to define what the parameter will be (it is called “control position”) and for each control position the program calculates the trim point so that the aircraft is balanced. Since you can decide what you want to use as parameter, you can use for instance the elevator rigging angle, or the wing rigging angle, and when you do a sweep across several control values you get corresponding AoA values.
“Without control input”, as you put it, there are no different AoAs.
But for example when there is no thrust, no elevator ,no aileron , no rudder deflection ; aircraft's AOA can be changing . After that change for example aircraft begin to descent and the Cm q value is changing with respect to alpha . So can't we find ?
In the simulink example aeroblk_HL20 , in the body rate body rate damping block ,they exactly make a Cmq table with respect to alpha.
So if we change the control parameters for finding derivatives for different AOA ,is it the right derivative ?
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If no control surfaces are deflected, and CG doesn’t change, how can the AoA change if the aircraft is trimmed (which is XFLR5’s boundary condition for stability analysis)?
Hi everybody .
I need to find Clp for different AOA . After stability analysis there is only one AOA and so one Clp. Can anybody help pls ?
Only one AOA? what are you using as your control for stability analysis?
There is no polar graph about Aero derivatives with respect to body rates (p,q,r). Also in the stability analysis Lateral derivatives
Yv= -0.23503 CYb= -0.08318
Yp= -0.11732 CYp= -0.055361
Yr= 0.17284 CYr= 0.08156
Lv= -0.15326 Clb= -0.03616
Lp= -1.7779 Clp= -0.55931
Lr= 0.15593 Clr= 0.049055
Nv= 0.14247 Cnb= 0.033616
Np= -0.097373 Cnp= -0.030633
Nr= -0.092417 Cnr= -0.029074
there is only one Clp . I could not define analysis for different AOA for stability analysis.
I still don’t get what you use as control parameter. A common choice is the elevator angle, when you do a sweep across control values this gives several different AoAs.
Thanks for your answer. i dont use any control parameter . For example in the polar i can easily find the CL with respect to alpha . But there is no CL with respect to p ,q or r . I want to find for example Clp for different AOA without using control input.
The reason why you have CL/alpha in Type 2 polars is that it is a parametric analysis on alpha. In stability analysis, you get to define what the parameter will be (it is called “control position”) and for each control position the program calculates the trim point so that the aircraft is balanced. Since you can decide what you want to use as parameter, you can use for instance the elevator rigging angle, or the wing rigging angle, and when you do a sweep across several control values you get corresponding AoA values.
“Without control input”, as you put it, there are no different AoAs.
Francesco
But for example when there is no thrust, no elevator ,no aileron , no rudder deflection ; aircraft's AOA can be changing . After that change for example aircraft begin to descent and the Cm q value is changing with respect to alpha . So can't we find ?
In the simulink example aeroblk_HL20 , in the body rate body rate damping block ,they exactly make a Cmq table with respect to alpha.
So if we change the control parameters for finding derivatives for different AOA ,is it the right derivative ?
If no control surfaces are deflected, and CG doesn’t change, how can the AoA change if the aircraft is trimmed (which is XFLR5’s boundary condition for stability analysis)?
i doubt some of the points but thanks for the answer. I really understand . I will make tables Cmq with different AOA by changing elevator gain .