﻿aeroPerformanceMap Element
 aeroPerformanceMap Element

Aerodynamic coefficients and derivatives

Namespace: Empty

Schema: Empty

Type
Parents
Children
NameOccurrencesDescription
All
altitude

Altitude [m]

angleOfAttack

Angle of attack [deg]

angleOfSideslip

Sideslip angle [deg]

cd[0, 1]

Drag coefficient in aerodynamic coordinates

cl[0, 1]

Lift coefficient in aerodynamic coordinates

cmd[0, 1]

Vector with semicolon separated values of type double

cml[0, 1]

Vector with semicolon separated values of type double

cms[0, 1]

Vector with semicolon separated values of type double

cs[0, 1]

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

dampingDerivatives[0, 1]

Damping derivatives for positive and negative rotation rates

incrementMaps[0, 1]

Increment maps for aerodynamic coefficients

machNumber

Mach number

Attributes
NameTypeRequiredDescription
externalDataNodePath string
externalFileName string
Remarks

#### Description

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.