aeroPerformanceMap Element

Aerodynamic coefficients and derivatives

Namespace: Empty

Schema: Empty

All All
      Element altitude

Altitude [m]

      Element angleOfAttack

Angle of attack [deg]

      Element angleOfSideslip

Sideslip angle [deg]

      Element cd[0, 1]

Drag coefficient in aerodynamic coordinates

      Element cl[0, 1]

Lift coefficient in aerodynamic coordinates

      Element cmd[0, 1]

Vector with semicolon separated values of type double

      Element cml[0, 1]

Vector with semicolon separated values of type double

      Element cms[0, 1]

Vector with semicolon separated values of type double

      Element cs[0, 1]

Coefficient of the side force vector in aerodynamic coordinates (perpendicular to lift and drag)

      Element dampingDerivatives[0, 1]

Damping derivatives for positive and negative rotation rates

      Element incrementMaps[0, 1]

Increment maps for aerodynamic coefficients

      Element machNumber

Mach number

Attribute externalDataDirectorySimple Type string
Attribute externalDataNodePathSimple Type string
Attribute externalFileNameSimple Type string


The aeroPerformanceMap contains a map with aerodynamic data of the complete aircraft in the form of nondimensional coefficients. The force coefficients in i-direction (ci) are nondimensionalized by dynamic pressure and reference area, the moment coefficients (cmi) by dynamic pressure, reference area and reference length.

All coefficients in the aeroPerformanceMap relate to the aerodynamic coordinate system which is deducted from the CPACS coordinate system by the transformations of angle of attack and angle of yaw. See the documentation of the CPACS element for further details.

The dependend parameters of the aeroPerformanceMap are altitude, machNumber, angleOfSideslip and angleOfAttack. These elements are vectors of equal length, where values with identical indices belong together. The solution vectors ci and cmi have the same length as the input vectors. Shown below is an example where with 10 values per vector:

Example aeroPerformanceMap
<altitude mapType="vector">12e+02;12e+02;12e+02;12e+02;12e+02;12e+02;12e+02;12e+02;12e+02;12e+02</altitude>
<machNumber mapType="vector">0.2;0.2;0.2;0.2;0.2;0.2;0.2;0.2;0.2;0.2</machNumber>
<angleOfSideslip mapType="vector">0;0;0;0;0;2;2;2;2;2</angleOfSideslip>
<angleOfAttack mapType="vector">-2;0;2;4;6;-2;0;2;4;6</angleOfAttack>
<cd mapType="vector">0.056;0.094;0.132;0.17;0.208;0.072;0.11;0.148;0.186;0.224</cd>
<cs mapType="vector">0.;0.;0.;0.;0.;0.01;0.015;0.02;0.025;0.03</cs>
<cl mapType="vector">-0.1;0.04;0.18;0.32;0.46;-0.08;0.03;0.14;0.25;0.36</cl>

The aerodynamic coefficients for altitude=1200m, machNumber=0.2, angleOfSideslip=0° and angleOfAttack=6° can be found at the 5th index: cd=0.208, cs=0 and cl=0.46.

See Also