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Aircraft Aerodynamics with LinAir - Aerodynamic Decelerator Systems Laboratory

ADSC - RPT - Aircraft Aerodynamics

Aircraft Aerodynamics with LinAir

Two software packages were used to characterize aerodynamics of each RPT. The first one, LinAir, is a simple and yet very reliable program that computes aerodynamic properties based on the model created from several individual elements (panels). It is capable of generating the effect of varying the angle of attack and sideslip angle (“virtual wind tunnel”), as well as introducing turn rates and deflecting the control surfaces. LinAir was developed back in 1982 by Ian Kroo from Stanford University and has been used since then by many companies such as Boeing, AeroVironment, Northrop, Lockheed, and NASA to quickly assess different novel design concepts.

As an example, Figs.D,E show a P10B RPT model composed of 22 panels

  • 3 panels to represent a vertical span of the fuselage (in the x-z plane)
  • 3 panels to represent a horizontal span of horizontal fuselage (in the x-y plane)
  • 2 panels to represent a left wing
  • 2 panels to represent a right wing (see a description of one of them in Fig.Db)
  • 1 panel for a left aileron
  • 1 panel for a right aileron
  • 2 panels for a horizontal stabilizer
  • 2 panels for a left elevator
  • 2 panels for a right elevator
  • 2 panels for a vertical stabilizer
  • 2 panels for a rudder
Nested Applications
ADSC - RPT - Aircraft Aerodynamics - Linair Pro 1
Linair Pro
ADSC - RPT - Aircraft Aerodynamics - Linair Pro 2
LinAir Pro
ADSC - RPT - Aircraft Aerodynamics - Figure D

Figure D. The P10B model in the LinAir environment.

ADSC - RPT - Aircraft Aerodynamics - Fig E
Figure E. Different views of the P10B LinAir model.

Figure E. Different views of the P10B LinAir model.

ADSC - RPT - Aircraft Aerodynamics - Fig. F

Figure F shows an aerodynamic force distribution for all 22 panels and serves as the good means of visual model verification (for a symmetric configuration the force contribution should be symmetric with most of the forces positive).

Figure F. Force contribution of the P10B model panels.

Figure F. Force contribution of the P10B model panels.

Once the base (nominal) model is created and verified, it is modified to create a few more models with the control surfaces (aileron, elevator, and rudder) deflected, so that the complete model is comprised of at least ten input files (one neutral plus three for each of the three control surfaces with different deflection angles).

Next, the angle of attack (alpha) and sideslip angle (beta) sweeps are executed for each of these input files. All data for each sweep are stored in a separate file in a table format. Figure G shows an example of visualizing these data in MATLAB.

Nested Applications
ADSC - RPT - Aircraft Aerodynamics - Fig Ga
Figure G. The P10B aerodynamic derivatives derived from the alpha and beta sweeps.
ADSC - RPT - Aircraft Aerodynamics - Fig Gb
Figure G. The P10B aerodynamic derivatives derived from the alpha and beta sweeps.
ADSC - RPT - Aircraft Aerodynamics - Fig G caption

Figure G. The P10B aerodynamic derivatives derived from the alpha and beta sweeps.

As an illustrative example, Table B provides aerodynamic and controls coefficients computed for P10B RPT with two wing configurations using LinAir.

Notation Coefficient Flat configuration 6° dihedral angle

Table B. Coefficients obtained for two configurations of P10B with LinAir.

CL0 lift coefficient at a = 0 0.172 0.252
CLalpha lift curve slope 5.04 5.04
CLa_dot lift due to angle of attack rate 4.95 4.89
CLq lift due to pitch rate 5.62 5.64
CLDe lift due to elevator 0.369 0.362
CD0 drag coefficient at CL = 0 0.0131 0.0129
A1 drag curve coefficient at CL 0.0002 0.0004
A2 drag curve coefficient at CL2 0.0777 0.0774
CDDe drag due to elevator 0.0424 0.0431
CYb side force due to sideslip -0.287 -0.292
CYDr side force due to rudder 0.157 0.159
Clb dihedral effect -0.103 -0.115
Clp roll damping -3.22 -4.05
Clr roll due to yaw rate 0.292 0.350
ClDa roll control power 0.137 0.451
ClDr roll due to rudder 0.0582 0.0052
Cm0 pitch moment at a = 0 0.0344 0.0974
Cma pitch moment due to angle of attack -2.29 -2.47
Cma_dot pitch moment due to angle of attack rate -1.97 -2.07
Cmq pitch moment due to pitch rate -7.71 -8.46
CmDe pitch control power -1.04 -1.03
Cnb weathercock stability 0.0573 0.0573
Cnp adverse yaw -0.206 -0.269
Cnr yaw damping -0.332 -0.378
CnDa aileron adverse yaw 0.0004 -0.0045
CnDr yaw control power -0.005 -0.059