WOLFRAM SYSTEM MODELER
AircraftBaseInterface for a complete aircraft model 
SystemModel["Aircraft.Physical.FixedWing.Interfaces.AircraftBase"]
This base model compiles a complete fixedwing aircraft model with a conventional wing configuration from the propulsion, surface and body components in the Parts package and takes input commands for the control actuator, i.e. commands for throttle position(s) as well as for the ailerons, rudder and elevator deflections. Additionally, all parameters of the aircraft other than the propulsionspecific parameters (except for the number of engines) are declared in this model. However, if the weight estimation method is used and any mass property is to be entered directly for any component instead of using the estimated value, it can be done in the corresponding component model in its Estimated Mass Properties parameter tab.
This base model is extended by the complete aircraft models found in the FixedWing package and can also be extended by the user's own fixedwing aircraft designs with conventional wing configuration.
Due to many couplings between the characteristics of different components of the aircraft, most of the aircraft parameters are declared on this toplevel aircraft model instead of declaring them lower at the component models. The parameters declared here are propagated down to the components where they are needed.
The first parameter in this tab, namely the weightEst parameter, declares whether the weight estimation method is used to estimate the mass properties of the aircraft component by component or if the known mass properties of the entire aircraft are entered directly. Depending on its value, the editing of the relevant mass property or design variable parameters is enabled. The theory behind the weight estimation method is described in the documentation of the corresponding components. The suggested workflow to be followed when using the weight estimation method in aircraft sizing is presented in GettingStarted.
If no weight estimation is used, the known mass properties to be entered are without the contribution of fuel (except for electric aircraft and gliders, for which they are the mass properties with the total mass), which all remain constant during a simulation. The mass of the initial fuel is defined separately with the initialMfuel parameter, and the consequent center of mass location and variable inertia tensor of the fuel system are estimated in the TankSystem component regardless of whether the weight estimation method is used or not. When the weight estimation method is used, the initial fuel mass is adjusted as a fraction of the estimated maximum fuel capacity in the Fuel System parameter tab of the propulsion component.
All parameters concerning propulsion and fuel or battery system are declared in the propulsion component, with the exception of the number of engines (nEng), which is declared in this base model. However, the choice of the propulsion type is done in this model by redeclaring the propulsion component into TurbofanPropulsion, TurbojetPropulsion, TurbopropPropulsion, PistonPropulsion, ElectricPropulsion or into a custom propulsion, as in Hawker Siddeley HS121 Trident 3B.
In this parameter tab, the coordinates of the wings with respect to the fuselage reference point, the dimensions of the fuselage and wings and the incidence, sweep and dihedral angles of the wings are to be entered. The fuselage reference point is defined at the halflength of all given maximum fuselage dimensions, as shown in Figure 1.
Figure 1: Definition of the fuselage reference point location. [1]
The required aerodynamic properties to be entered to the aircraft model are the properties of the twodimensional airfoils used in the main wing, horizontal tail and vertical tail and the surface roughness height on different surfaces. By default, the given surface roughness height for the entire aircraft (kSkinAC) is propagated to be used on the surfaces of all components, including the surface of the nacelles in the turbofan, turbojet and turboprop engines.
The derivation of the lift and drag coefficients of the wings and the fuselage from the given Aerodynamic Properties and Geometry parameters is described in the documentation for WingBody, HorizontalTail and VerticalTail components, and the derivation of the drag coefficients of the nacelles in the turbofan, turbojet and turboprop engines is described in their respective documentation.
The maximum control surface deflections given here are propagated to be used in the limiters in the Ailerons, HorizontalTail and VerticalTail components.
This tab includes the parameters that are derived based on the parameters given on the previous tabs. The equations used to derive and estimate these values may be bypassed by entering a known value in the parameter input field.
The parameters to define the initial position, orientation, and translational and rotational velocities are entered here. The initialization for the translational velocity is performed in this model from the Start Flight Conditions parameters, whereas the initialization for the position, orientation and angular velocity is performed in the Body model from the propagated Start Position and Start Orientation and Angular Velocities parameters.
By default, the Modelica Standard Library FixedShape primitives are used to visualize different components with simple box and cylinder shapes with the dimensions given in the Geometry parameters and by estimating them for certain components. However, if a CAD model of the complete aircraft model exists as an .obj file, it can be used by setting CADshapes as true and entering the path of the .obj file to CADpath field. This automatically disables the default animation in every component. The rotation and translation between the CAD object origin and the fuselage reference point can be entered to the translCADshape in the Body model.
This model contains real inputs for the control actuators, a frame connected to the fuselage reference point (aircraftRP), and a FlightDataOut connector for forwarding the variables of the flight data as a feedback to the autopilots. The aircraftRP frame can be used, for example, to create physical connection to other aircraft, as is done in the GliderTow example.
The numbered real inputs for the throttle position commands (deltaThrot1Cmd  deltaThrot5Cmd) follow the numbering of the engines presented in the ConventionalPropulsion model. If a CustomPropulsion model is used, the unnumbered deltaThrotCmd real input adjusts its dimension to the now unconstrained number of engines defined as nEng.
The variables of the complete aircraft that are solved for in this model by fetching the required variables from the subcomponents include:
As the mass properties of a rigid multibody system are not automatically solved, first the center of mass location for the entire aircraft including the contribution from the variable fuel mass (xCM, yCM, zCM) is calculated. Then, the moments of inertia around this center of mass (Ixx, Iyy, Izz) are solved by using the parallel axis theorem, and the total products of inertia (Ixy, Ixz, Iyz) are solved by summing the products of component masses and their coordinates with respect to the solved aircraft center of mass.
This method, however, is a minor simplification, as it considers the local frames of the main wing and horizontal tail surfaces to be parallel with the fuselage reference plane. In fact, the moments of inertia of the main wing and horizontal tail are solved in a frame parallel to their surfaces by considering the rotations around their incidence and dihedral angles.
[1] EräEsko, N. (2022). "Development and Use of System Modeler 6DOF Flight Mechanics Model in Aircraft Conceptual Design."
Available at: modelica://Aircraft/Resources/Documents/EraeEskoThesis.pdf.
weightEst 
Value: Type: Boolean Description: true, if weight estimation method is used for masses, center of mass location and inertia tensor 

mDry 
Value: Type: Mass (kg) Description: Aircraft dry mass (total mass for electric aircraft and gliders) 
initialMfuel 
Value: Type: Mass (kg) Description: Initial fuel mass 
xCMdry 
Value: Type: Length (m) Description: Aircraft center of mass xcoordinate w.r.t. fuselage reference point (with total mass for electric aircraft and gliders, positive xaxis towards nose) 
yCMdry 
Value: Type: Length (m) Description: Aircraft center of mass ycoordinate w.r.t. fuselage reference point (with total mass for electric aircraft and gliders, positive yaxis towards right) 
zCMdry 
Value: Type: Length (m) Description: Aircraft center of mass zcoordinate w.r.t. fuselage reference point (with total mass for electric aircraft and gliders, positive zaxis towards ground) 
IxxDry 
Value: Type: MomentOfInertia (kg·m²) Description: Aircraft moment of inertia about xaxis (with total mass for electric aircraft and gliders) 
IyyDry 
Value: Type: MomentOfInertia (kg·m²) Description: Aircraft moment of inertia about yaxis (with total mass for electric aircraft and gliders) 
IzzDry 
Value: Type: MomentOfInertia (kg·m²) Description: Aircraft moment of inertia about zaxis (with total mass for electric aircraft and gliders) 
IxyDry 
Value: Type: MomentOfInertia (kg·m²) Description: Aircraft xyproduct of moment of inertia (with total mass for electric aircraft and gliders) 
IxzDry 
Value: Type: MomentOfInertia (kg·m²) Description: Aircraft xzproduct of moment of inertia (with total mass for electric aircraft and gliders) 
IyzDry 
Value: Type: MomentOfInertia (kg·m²) Description: Aircraft yzproduct of moment of inertia (with total mass for electric aircraft and gliders) 
MTOMdes 
Value: Type: Mass (kg) Description: Design maximum takeoff mass 
nPax 
Value: Type: Integer Description: Design number of passengers 
mPLdes 
Value: Type: Mass (kg) Description: Design payload mass 
machDes 
Value: Type: Real Description: Design Mach number 
compMat 
Value: Type: Boolean Description: true, if composite materials are used in structures 
qMax 
Value: Type: Pressure (Pa) Description: Maximum dynamic pressure 
nMax 
Value: Type: Real Description: Maximum load factor 
nEng 
Value: Type: Integer Description: Number of engines 
wingMounted 
Value: if nEng == 0 then false else propulsion.wingMounted Type: Boolean 
engineType 
Value: if nEng == 0 then 4 else propulsion.engineType Type: Integer 
lFus 
Value: Type: Length (m) Description: Fuselage length 
wFus 
Value: Type: Length (m) Description: Fuselage maximum width 
hFus 
Value: Type: Length (m) Description: Fuselage maximum height 
dFusHT 
Value: Type: Length (m) Description: Fuselage diameter at horizontal tail 1/4 chord 
bWing 
Value: Type: Length (m) Description: Main wing span 
cWingRoot 
Value: Type: Length (m) Description: Main wing root chord (where wing intersects with fuselage) 
cWingTip 
Value: Type: Length (m) Description: Main wing tip chord 
tWingRoot 
Value: Type: Length (m) Description: Main wing root thickness 
tWingTip 
Value: Type: Length (m) Description: Main wing tip thickness 
xWingRootLE 
Value: Type: Length (m) Description: Main wing root leading edge xcoordinate w.r.t. fuselage reference point (positive xaxis towards nose) 
zWingRootLE 
Value: Type: Length (m) Description: Main wing root leading edge zcoordinate w.r.t. fuselage reference point (positive zaxis towards ground) 
lambdaWing 
Value: Type: Angle (rad) Description: Main wing sweep angle at 1/4 chord 
gammaWing 
Value: Type: Angle (rad) Description: Main wing dihedral angle 
iWing 
Value: Type: Angle (rad) Description: Main wing incidence angle 
cAil 
Value: Type: Length (m) Description: Aileron average chord 
yAilRoot 
Value: Type: Length (m) Description: Aileron root ycoordinate w.r.t. fuselage centerline 
yAilTip 
Value: Type: Length (m) Description: Aileron tip ycoordinate w.r.t. fuselage centerline 
bHT 
Value: Type: Length (m) Description: Horizontal tail span 
cHTroot 
Value: Type: Length (m) Description: Horizontal tail root chord 
cHTtip 
Value: Type: Length (m) Description: Horizontal tail tip chord 
tHTroot 
Value: Type: Length (m) Description: Horizontal tail root thickness 
tHTtip 
Value: Type: Length (m) Description: Horizontal tail tip thickness 
xHTrootLE 
Value: Type: Length (m) Description: Horizontal tail root leading edge xcoordinate w.r.t. fuselage reference point (positive xaxis towards nose) 
zHTrootLE 
Value: Type: Length (m) Description: Horizontal tail root leading edge zcoordinate w.r.t. fuselage reference point (positive zaxis towards ground) 
lambdaHT 
Value: Type: Angle (rad) Description: Horizontal tail sweep angle at 1/4 chord 
iHT 
Value: Type: Angle (rad) Description: Horizontal tail incidence angle 
Selv 
Value: Type: Area (m²) Description: Elevator area 
bVT 
Value: Type: Length (m) Description: Vertical tail span 
cVTroot 
Value: Type: Length (m) Description: Vertical tail root chord 
cVTtip 
Value: Type: Length (m) Description: Vertical tail tip chord 
tVTroot 
Value: Type: Length (m) Description: Vertical tail root thickness 
tVTtip 
Value: Type: Length (m) Description: Vertical tail tip thickness 
xVTrootLE 
Value: Type: Length (m) Description: Vertical tail root leading edge xcoordinate w.r.t. fuselage reference point (positive xaxis towards nose) 
zVTroot 
Value: Type: Length (m) Description: Vertical tail root zcoordinate w.r.t fuselage reference point 
lambdaVT 
Value: Type: Angle (rad) Description: Vertical tail sweep angle at 1/4 chord 
Srdr 
Value: Type: Area (m²) Description: Rudder area 
kSkinAC 
Value: Type: Length (m) Description: Surface roughness height (same value to be used for all components) 
kSkinFus 
Value: kSkinAC Type: Length (m) Description: Fuselage surface roughness height 
kSkinWing 
Value: kSkinAC Type: Length (m) Description: Main Wing surface roughness height 
ClAlphaWing2D 
Value: Type: CurveSlope (rad⁻¹) Description: Change in the section lift coefficient of the main wing airfoil (2D) due to alpha 
alpha0Wing2D 
Value: Type: Angle (rad) Description: Zerolift angle of attack of the main wing airfoil (2D) 
ClMaxWing2D 
Value: Type: Real Description: Maximum section lift coefficient of the main wing airfoil (2D) 
kSkinHT 
Value: kSkinAC Type: Length (m) Description: Horizontal tail surface roughness height 
ClAlphaHT2D 
Value: Type: CurveSlope (rad⁻¹) Description: Change in the section lift coefficient of the horizontal tail airfoil (2D) due to alpha 
alpha0HT2D 
Value: Type: Angle (rad) Description: Zerolift angle of attack of the horizontal tail airfoil (2D) 
ClMaxHT2D 
Value: Type: Real Description: Maximum section lift coefficient of the horizontal tail airfoil (2D) 
kSkinVT 
Value: kSkinAC Type: Length (m) Description: Vertical tail surface roughness height 
ClAlphaVT2D 
Value: Type: CurveSlope (rad⁻¹) Description: Change in the section lift coefficient of the vertical tail airfoil (2D) due to alpha 
deltaElvMax 
Value: Type: Angle (rad) Description: Maximum elevator deflection 
deltaAilMax 
Value: Type: Angle (rad) Description: Maximum aileron deflection 
deltaRdrMax 
Value: Type: Angle (rad) Description: Maximum rudder deflection 
sigmaBeta 
Value: max(3.06 * (SrefVT / SrefWing) / (1 + cos(lambdaWing)) + 0.4 * (tan(gammaWing) * (yWingAC  wFus / 2) + zWingRootLE) / wFus + 0.009 * (bWing ^ 2 / SrefWing)  0.276, 0) Type: Real Description: Change in sidewash due to beta 
rACcm 
Value: if weightEst then {xWingRootLE  lambdaWingLE * (yWingAC  wFus / 2)  0.15 * cWingMean, 0, 0} else {xCMdry, yCMdry, zCMdry} Type: Length[3] (m) Description: Aircraft dry center of mass w.r.t. fuselage reference point (estimated to be at 15% of MAC if weight estimation method is used) 
Cfus 
Value: Modelica.Constants.pi * (3 * (hFus / 2 + wFus / 2)  sqrt(10 * hFus / 2 * wFus / 2 + 3 * ((hFus / 2) ^ 2 + (wFus / 2) ^ 2))) Type: Length (m) Description: Fuselage circumference 
SwetFus 
Value: Cfus * lFus * (1  2 / (lFus / (Cfus / Modelica.Constants.pi))) ^ (2 / 3) * (1 + 1 / (lFus / (Cfus / Modelica.Constants.pi)) ^ 2) Type: Area (m²) Description: Fuselage wetted area 
FFfus 
Value: 1 + 0.0025 * (lFus / hFus) + 60 * (hFus / lFus) ^ 3 Type: Real Description: Fuselage form factor 
CDmaxFus 
Value: 0.8 * lFus * hFus / SrefWing Type: Real Description: Maximum drag coefficient of the fuselage 
nSeatAbs 
Value: if nPax > 180 then floor(0.9 * wFus / 0.5588)  1 else floor(0.9 * wFus / 0.5588) Type: Real Description: Number of seats abreast 
SrefWing 
Value: (cWingRoot + cWingTip) * (bWing  wFus) / 2 + cWingRoot * wFus Type: Area (m²) Description: Main wing reference area 
ARwing 
Value: bWing ^ 2 / SrefWing Type: Real Description: Main wing aspect ratio 
TRwing 
Value: cWingTip / cWingRoot Type: Real Description: Main wing taper ratio 
cWingMean 
Value: 2 / 3 * cWingRoot * ((1 + TRwing + TRwing ^ 2) / (1 + TRwing)) Type: Length (m) Description: Main wing mean chord length 
tWingMean 
Value: (yWingAC  wFus / 2) * (tWingTip  tWingRoot) / (bWing / 2  wFus / 2) + tWingRoot Type: Length (m) Description: Main wing mean thickness 
tauWing 
Value: tWingTip / cWingTip / (tWingRoot / cWingRoot) Type: Real Description: Ratio of thicknesstochord ratios at the main wing tip and root 
xWingAC 
Value: 0.25 * cWingMean Type: Length (m) Description: Main wing aerodynamic center from wing leading edge at mean chord (positive xaxis towards nose) 
yWingAC 
Value: (cWingRoot  cWingMean) / (cWingRoot  cWingTip) * (bWing  wFus) / 2 + wFus / 2 Type: Length (m) Description: Main wing aerodynamic center from fuselage centerline (ycoordinate w.r.t. fuselage centerline of mean chord) 
SwetWing 
Value: 2 * (SrefWing / cos(gammaWing)  cWingRoot * wFus) * (1 + 0.25 * (tWingRoot / cWingRoot) * (1 + tWingTip / cWingTip / (tWingRoot / cWingRoot) * TRwing) / (1 + TRwing)) Type: Area (m²) Description: Main wing wetted area 
lambdaWingLE 
Value: atan((cWingRoot / 4  cWingTip / 4 + tan(lambdaWing) * bWing / 2) / (bWing / 2)) Type: Angle (rad) Description: Main wing leading edge sweep angle 
lambdaWingHC 
Value: atan((cWingTip / 4  cWingRoot / 4 + tan(lambdaWing) * bWing / 2) / (bWing / 2)) Type: Angle (rad) Description: Main wing halfchord sweep angle 
lambdaWingTE 
Value: atan((lambdaWingLE * ((bWing  wFus) / 2) + cWingTip  cWingRoot) / ((bWing  wFus) / 2)) Type: Angle (rad) Description: Main wing trailing edge sweep angle (at ailerons location) 
CLmaxWing3D 
Value: 0.9 * ClMaxWing2D * cos(lambdaWing) Type: Real Description: Maximum lift coefficient of the main wing (3D) 
CDmaxWing3D 
Value: 1.98  0.81 * (1  Modelica.Constants.e ^ (20 / ARwing)) Type: Real Description: Maximum drag coefficient of the main wing (3D) 
FFwing 
Value: 0.421 * (2 + 4 * tWingMean / cWingMean + 240 * (tWingMean / cWingMean) ^ 4) Type: Real Description: Main wing form factor 
sdWing 
Value: 0.9998 + 0.0421 * (wFus / bWing)  2.6286 * (wFus / bWing) ^ 2 + 2 * (wFus / bWing) ^ 3 Type: Real Description: Fuselage drag factor for main wing 
kdWing 
Value: 3.333 * 10 ^ (4) * lambdaWing ^ 2 + 6.667 * 10 ^ (5) * lambdaWing + 0.38 Type: Real Description: Empirical constant for Oswald efficiency factor for main wing 
Sail 
Value: cAil * (yAilTip  yAilRoot) * 2 Type: Area (m²) Description: Aileron area (of both wings) 
xAilAC 
Value: (xCMdry  (xWingRootLE  lambdaWingLE * (yWingAC  wFus / 2)))  0.25 * cWingMean  tan(lambdaWing) * ((yAilRoot + yAilTip) / 2  yWingAC)  (cAilWingRoot + cAilWingTip) / 2 * 0.75 + cAil / 4 Type: Length (m) Description: Aileron aerodynamic center xcoordinate w.r.t. aircraft center of mass 
yAilAC 
Value: (yAilRoot + yAilTip) / 2 Type: Length (m) Description: Aileron aerodynamic center ycoordinate w.r.t. fuselage centerline 
cAilWingRoot 
Value: cWingRoot  (yAilRoot  wFus / 2) * 2 / (bWing  wFus) * (cWingRoot  cWingTip) Type: Length (m) Description: Local main wing chord at aileron root 
cAilWingTip 
Value: cWingRoot  (yAilTip  wFus / 2) * 2 / (bWing  wFus) * (cWingRoot  cWingTip) Type: Length (m) Description: Local main wing chord at aileron tip 
kCnDeltaAil 
Value: 0.350894  0.066355 * (yAilRoot / (bWing / 2)) ^ 4.15179 + 0.029308 * (bWing ^ 2 / SrefWing) Type: Real Description: Empirical factor for the yaw moment derivative due to ailerons 
tauAil 
Value: 1.129 * (Sail / SrefWing) ^ 0.4044  0.1772 Type: Real Description: Aileron effectiveness parameter 
SrefHT 
Value: (cHTroot + cHTtip) * (bHT  dFusHT) / 2 + cHTroot * dFusHT Type: Area (m²) Description: Horizontal tail reference area 
ARht 
Value: bHT ^ 2 / SrefHT Type: Real Description: Aspect ratio of horizontal tail 
TRht 
Value: cHTtip / cHTroot Type: Real Description: Taper ratio of horizontal tail 
cHTmean 
Value: 2 / 3 * cHTroot * (1 + TRht + TRht ^ 2) / (1 + TRht) Type: Length (m) Description: Horizontal tail mean chord 
tHTmean 
Value: bHT / 2 * (1 + 2 * TRht) / (3 + 3 * TRht) * (tHTtip  tHTroot) / (bHT / 2  dFusHT / 2) + tHTroot Type: Length (m) Description: Horizontal tail mean thickness 
lHTcm 
Value: abs(xHTrootLE  xCMdry  tan(lambdaHTle) * (bHT / 2  dFusHT / 2) * (1 + 2 * TRht) / (3 + 3 * TRht)  0.25 * cHTmean) Type: Length (m) Description: Horizontal tail arm length (from aircraft center of mass to horizontal tail 1/4 chord) 
lHTwingAC 
Value: abs(xHTrootLE  (xWingRootLE  lambdaWingLE * (yWingAC  wFus / 2))) + tan(lambdaHTle) * (bHT / 2  dFusHT / 2) * (1 + 2 * TRht) / (3 + 3 * TRht) + 0.25 * cHTmean  abs(xWingAC) Type: Length (m) Description: Horizontal tail arm length (from wing aerodynamic center to horizontal tail 1/4 chord) 
vHT 
Value: SrefHT / SrefWing * (lHTcm / cWingMean) Type: Real Description: Horizontal tail volume coefficient 
SwetHT 
Value: 2 * (SrefHT  cHTroot * dFusHT) * (1 + 0.25 * (tHTroot / cHTroot) * (1 + tHTtip / cHTtip / (tHTroot / cHTroot) * TRht) / (1 + TRht)) Type: Area (m²) Description: Horizontal tail wetted area 
lambdaHTle 
Value: atan((cHTroot / 4  cHTtip / 4 + tan(lambdaHT) * bHT / 2) / (bHT / 2)) Type: Angle (rad) Description: Horizontal tail leading edge sweep angle 
CLmaxHT3D 
Value: 0.9 * ClMaxHT2D * cos(lambdaHT) Type: Real Description: Maximum lift coefficient of the horizontal tail (3D) 
CDmaxHT3D 
Value: 1.98  0.81 * (1  Modelica.Constants.e ^ (20 / ARht)) Type: Real Description: Maximum drag coefficient of the horizontal tail (3D)*(SrefHT/SrefWing) 
FFht 
Value: 1 + 0.1 * (1  0.893 * abs(zHTrootLE / hFus)) * (2 + 4 * tHTmean / cHTmean + 240 * (tHTmean / cHTmean) ^ 4) Type: Real Description: Horizontal tail form factor 
sdHT 
Value: 0.9998 + 0.0421 * (dFusHT / bHT)  2.6286 * (dFusHT / bHT) ^ 2 + 2 * (dFusHT / bHT) ^ 3 Type: Real Description: Fuselage drag factor for horizontal tail 
kdHT 
Value: 3.333 * 10 ^ (4) * lambdaHT ^ 2 + 6.667 * 10 ^ (5) * lambdaHT + 0.38 Type: Real Description: Empirical constant for Oswald efficiency factor for horizontal tail 
tauElv 
Value: 1.129 * (Selv / SrefHT) ^ 0.4044  0.1772 Type: Real Description: Elevator effectiveness parameter 
SrefVT 
Value: (cVTroot + cVTtip) / 2 * bVT Type: Area (m²) Description: Vertical tail reference area 
ARvt 
Value: bVT ^ 2 / SrefVT Type: Real Description: Aspect ratio of vertical tail 
TRvt 
Value: cVTtip / cVTroot Type: Real Description: Taper ratio of vertical tail 
cVTmean 
Value: 2 / 3 * cVTroot * (1 + TRvt + TRvt ^ 2) / (1 + TRvt) Type: Length (m) Description: Vertical tail mean chord 
tVTmean 
Value: bVT * (1 + 2 * TRvt) / (3 + 3 * TRvt) * (tVTtip  tVTroot) / bVT + tVTroot Type: Length (m) Description: Vertical tail mean thickness 
lVTcm 
Value: abs(xVTrootLE  xCMdry  tan(lambdaVTle) * bVT * (1 + 2 * TRvt) / (3 + 3 * TRvt)  0.25 * cVTmean) Type: Length (m) Description: Vertical tail arm length (from aircraft center of mass to vertical tail 1/4 chord) 
lVTwingAC 
Value: abs(xVTrootLE  (xWingRootLE  lambdaWingLE * (yWingAC  wFus / 2))) + tan(lambdaVTle) * bVT * (1 + 2 * TRvt) / (3 + 3 * TRvt) + 0.25 * cVTmean  abs(xWingAC) Type: Length (m) Description: Vertical tail arm length (from wing aerodynamic center to vertical tail aerodynamic center) 
zVTac 
Value: zVTroot  bVT * (1 + 2 * TRvt) / (3 + 3 * TRvt) Type: Length (m) Description: Vertical tail center of pressure zcoordinate w.r.t. fuselage reference point 
vVT 
Value: SrefVT / SrefWing * (lVTcm / bWing) Type: Real Description: Vertical tail volume coefficient 
SwetVT 
Value: 2 * SrefVT * (1 + 0.25 * (tVTroot / cVTroot) * (1 + tVTtip / cVTtip / (tVTroot / cVTroot) * TRvt) / (1 + TRvt)) Type: Area (m²) Description: Vertical tail wetted area 
lambdaVTle 
Value: atan((tan(lambdaVT) * bVT  cVTtip / 4 + cVTroot / 4) / bVT) Type: Angle (rad) Description: Vertical tail leading edge sweep angle 
FFvt 
Value: 0.5 * (2 + 4 * tVTmean / cVTmean + 240 * (tVTmean / cVTmean) ^ 4) Type: Real Description: Vertical tail form factor 
sdVT 
Value: 0.9998 Type: Real Description: Fuselage drag factor for vertical tail 
kdVT 
Value: 3.333 * 10 ^ (4) * lambdaVT ^ 2 + 6.667 * 10 ^ (5) * lambdaVT + 0.38 Type: Real Description: Empirical constant for Oswald efficiency factor for vertical tail 
tauRdr 
Value: 1.129 * (Srdr / SrefVT) ^ 0.4044  0.1772 Type: Real Description: Rudder effectiveness parameter 
initialAltitude 
Value: 0 Type: Height (m) Description: Initial altitude 
initialLatPosition 
Value: {0, 0} Type: Position[2] (m) Description: Initial lateral position of the aircraft (x and y coordinates in world frame) 
initialVelocity 
Value: 0 Type: Velocity (m/s) Description: Initial velocity 
initialTrack 
Value: 0 Type: Angle (rad) Description: Initial track angle 
initialGamma 
Value: 0 Type: Angle (rad) Description: Initial flight path angle 
initialPhi 
Value: 0 Type: Angle (rad) Description: Initial roll angle 
initialTheta 
Value: 0 Type: Angle (rad) Description: Initial pitch angle 
initialPsi 
Value: 0 Type: Angle (rad) Description: Initial yaw angle (heading) 
initialAngularVelocity 
Value: {0, 0, 0} Type: AngularVelocity[3] (rad/s) Description: Initial {roll [p], pitch [q], yaw [r]} 
CADshapes 
Value: Type: Boolean Description: true, if external CAD files are used for animation 
CADpath 
Value: "modelica://Aircraft/Resources/CAD/PathName.obj" Type: String Description: Path for CAD file 
T0 
Value: atmos.T0 Type: Temperature (K) Description: Temperature at sealevel 
rho0 
Value: atmos.mAir * p0 / (atmos.R0 * T0) Type: Density (kg/m³) Description: Air density at sealevel 
p0 
Value: atmos.p0 Type: Pressure (Pa) Description: Static pressure at sealevel 
a0 
Value: sqrt(atmos.gammaAir * atmos.R0 / atmos.mAir * T0) Type: Velocity (m/s) Description: Speed of sound at sealevel 
gammaAir 
Value: atmos.gammaAir Type: Real Description: Adiabatic index for air 
xEtot 
Type: Real Description: State of aircraft total energy 

deltaAilCmd 
Type: RealInput Description: Ailerons deflection command 


deltaElvCmd 
Type: RealInput Description: Elevator deflection command 

deltaRdrCmd 
Type: RealInput Description: Rudder deflection command 

deltaThrot1Cmd 
Type: RealInput Description: Engine 1 throttle command 

deltaThrot2Cmd 
Type: RealInput Description: Engine 2 throttle command 

deltaThrot3Cmd 
Type: RealInput Description: Engine 3 throttle command 

deltaThrot4Cmd 
Type: RealInput Description: Engine 4 throttle command 

deltaThrot5Cmd 
Type: RealInput Description: Engine 5 throttle command 

aircraftRP 
Type: Frame_b Description: Connector to aircraft reference point 

flightDataOut 
Type: FlightDataOut 

deltaThrotCmd 
Type: RealInput[nEng] Description: Engine throttle commands when custom propulsion is used 
atmos 
Type: AtmosphericProperties 


flightData 
Type: FlightData 

body 
Type: Body 

propulsion 
Type: ConventionalPropulsion Description: Propulsion model 

surfaces 
Type: Conventional 
Aircraft.Physical.FixedWing Narrowbody turbofan airliner: Hawker Siddeley HS121 Trident 3B 

Aircraft.Physical.FixedWing Highspeed turboprop airliner: Saab 2000 

Aircraft.Physical.FixedWing Lightsport electric aircraft: Pipistrel Alpha Electro 

Aircraft.Physical.FixedWing Model of a general aviation aircraft 

Aircraft.Physical.FixedWing Midwing glider: Schweizer SGS 136 

Aircraft.Physical.FixedWing Narrowbody turbojet airliner: Douglas DC820 

Aircraft.Physical.FixedWing Narrowbody turbofan airliner: Boeing 737800 