JSBSim Aerodynamics: Difference between revisions

==Frames==

JSBSim incorporates several frame of reference. The body frame and wind frame are the most important to the aerodynamic model.

* '''Body XYZ''' The body frame uses the X axis for forward and aft, with + to the front. The Y axis is side to side with + to the right. The Z axis is up and down, with + being up. The origin of this frame is the Center of Gravity (CG), about which the aircraft forces and moments are summed and the resulting accelerations are integrated to get velocities.  

* '''Wind XYZ''' The wind frame X-axis points directly into the relative wind. The Y-axis is perpendicular to the X-axis, and remains within the aircraft body axis XZ plane (also called the reference plane). The Z-axis completes a right hand coordinate system. The origin of this frame is the AeroRP.

* '''Wind UVW'''   This isn't really a frame, but it is the components of the wind velocity vector. The relative wind is imposed on the body frame using variables U,V, and W. U to represent the velocity of the wind flowing past the aircraft. U represents the wind blowing down the X or longitudinal axis of the aircraft. V represents the wind blowing down the body frame Y axis, that is, wind in your ear. W represents wind blowing down the body frame Z axis, or wind from above.

==Angles==

* '''Alpha''' Alpha is the angle between the X body axis and the X wind axis measured on the UV plane. In standard aerodynamics Alpha is often referenced to the main wing chord line, this is not the case in [[JSBSim]].

The metrics section provides a standard place to record common aircraft dimensions. [[Aeromatic]] built [[FDM]]s only use three of the defined properties.

* '''Lift: CL''' Lift is a function of QBar * Wing Area * Cl_lift. Cl_lift is generally derived from a 2D table as a function of AoA. In the real world it is also a function of Reynolds Number.

* '''Drag: CD''' It is important to ensure all coefficient functions in the drag section remain positive. When drag coefficient functions are negative, drag is effectively acting as thrust opposite the relative wind.

* '''Roll: Cl''' Aerodynamic rolling moment comes from multiple sources:

** One is the lift generated by the vertical tail. On most aircraft the center of pressure for the vertical tail is not on the X body axis. This sets up a moment arm creating a moment of (moment arm length)*(vertical tail lift force).

* '''Pitch: Cm''' The primary contributor to the pitching moment is the lift generated by the horizontal tail. The pitching moment of the main wing airfoil is a secondary contributor. Because of this fact it makes sense that the value for Cm will resemble the CL curve for the horizontal tail airfoil. The standard CL curve does not take into account the changes in tail Angle of Attack (AoA) and QBar due to main wing down-wash and rotational velocity around the Y axis. Using QBarUV is more appropriate for Cm than the generic QBar.

* '''Yaw:   Cn''' The primary contributor to the yawing moment is the lift generated by the vertical tail. The wind force on the fuselage is a secondary contributor. Because of this fact it makes sense that the value for Cn will resemble the CL curve for the vertical tail airfoil. Changes in QBar due to the vertical tail moving into the 'shadow' of a stalled main wing may need to be accounted for, as well. The source for vertical tail Angle of Attack (AoA) should be Beta and QBarUW is more appropriate for Cn than the generic QBar.

* Stall - A 'stall' is generally regarded as a loss of lift due to flow separation over the top of a wing, however, examination of lift polar for an airfoil over a full 360 degrees shows that significant amounts of lift are NOT lost as the stall occurs. The biggest aerodynamic effect of a stall is a large and rapid increase in drag.

* Spin - Spins are caused loss of stability in the Yaw Moment axis. A stock [[Aeromatic]] [[FDM]] yaw section does not take alpha into account when calculating the yaw moment.

Beyond +/- 22.5 degrees AoA we'll use 0.26 * sin(2*AoA) to approximate lift, and linear interpolate between 15 and 22.5 degrees AoA. (These numbers are rather arbitrary.) The resultant table returns the same value as alpha-rad for the "stable" flight regime so the function will accept the [[aeromatic]] coefficient. Outside of the "stable" flight regime a better number will be used allowing more realistic flight behavior at high alpha or beta angles.

[[Image:fakelifttable.png]]

* Calculating Stall Speed - For an aircraft in straight and level flight, lift is equal to the weight of the aircraft. If we plug some numbers into the equation