/// <summary>
/// 计算大气延迟。 Computes the acceleration due to the atmospheric drag.
/// </summary>
/// <param name="Mjd_TT">Mjd_TT Terrestrial Time (Modified Julian Date)</param>
/// <param name="satXyz">r Satellite position vector in the inertial system [m]</param>
/// <param name="satVelocity"> v Satellite velocity vector in the inertial system [m/s]</param>
/// <param name="transMatrix"> T Transformation matrix to true-of-date inertial system</param>
/// <param name="Area">Area Cross-section [m^2]</param>
/// <param name="mass"> mass Spacecraft mass [kg]</param>
/// <param name="atmosDragCoeff"> coefOfDrag Drag coefficient</param>
/// <returns>Acceleration (a=d^2r/dt^2) [m/s^2]</returns>
public static Geo.Algorithm.Vector AccelerDragOfAtmos(double Mjd_TT, Geo.Algorithm.Vector satXyz, Geo.Algorithm.Vector satVelocity,
Matrix transMatrix, double Area, double mass, double atmosDragCoeff)
{
// Earth angular velocity vector [rad/s]
double[] Data_omega = { 0.0, 0.0, 7.29212e-5 };
Geo.Algorithm.Vector omega = new Geo.Algorithm.Vector(Data_omega, 3);
// Variables
double v_abs, dens;
Geo.Algorithm.Vector r_tod = new Geo.Algorithm.Vector(3), v_tod = new Geo.Algorithm.Vector(3);
Geo.Algorithm.Vector v_rel = new Geo.Algorithm.Vector(3), a_tod = new Geo.Algorithm.Vector(3);
Matrix T_trp = new Matrix(3, 3);
// Transformation matrix to ICRF/EME2000 system
T_trp = transMatrix.Transpose();
// Position and velocity in true-of-date system
r_tod = transMatrix * satXyz;
v_tod = transMatrix * satVelocity;
// Velocity relative to the Earth's atmosphere
v_rel = v_tod - omega.Cross3D(r_tod);
v_abs = v_rel.Norm();
// Atmospheric density due to modified Harris-Priester model
dens = TerrestrialUtil.AtmosDensity_HP(Mjd_TT, r_tod);
// Acceleration
a_tod = -0.5 * atmosDragCoeff * (Area / mass) * dens * v_abs * v_rel;
return(T_trp * a_tod);
}
/// <summary>
/// A_Diff: Function computes the difference of acceleration due to SRP and DRG
/// </summary>
/// <param name="h"> h Height [m]</param>
/// <returns>Difference of acceleration</returns>
public static double A_Diff(double h)
{
double CD = 2.3; // Spacecraft parameters
double CR = 1.3;
double Mjd_TT = 51269.0; // State epoch
Vector r = new Vector(1.0, 0.0, 0.0) * (Force.Grav.R_ref + h);
double dens = TerrestrialUtil.AtmosDensity_HP(Mjd_TT, r);
return((0.5 * CD * dens * Force.Grav.GM / (Force.Grav.R_ref + h)) - (OrbitConsts.PressureOfSolarRadiationPerAU * CR));
}