/// <summary>
/// 计算绕地航天器主要的加速度。
/// Computes the acceleration of an Earth orbiting satellite due to
/// - the Earth's harmonic gravity field,
/// - the gravitational perturbations of the Sun and Moon
/// - the solar radiation pressure and
/// - the atmospheric drag
/// </summary>
/// <param name="Mjd_TT"> Mjd_TT Terrestrial Time (Modified Julian Date)</param>
/// <param name="satXyzIcrf"> r 卫星在国际天球参考框架的位置 Satellite position vector in the ICRF/EME2000 system</param>
/// <param name="satVelocityIcrf"> v 卫星在国际天球参考框架的速度 Satellite velocity vector in the ICRF/EME2000 system</param>
/// <param name="Area"> Area Cross-section </param>
/// <param name="mass"> mass 质量 Spacecraft mass</param>
/// <param name="solarRadiPresCoeff"> coefOfRadiation 光压系数 Radiation pressure coefficient</param>
/// <param name="atmDragCoefficient"> coefOfDrag 大气阻力系数 Drag coefficient</param>
/// <returns> Acceleration (a=d^2r/dt^2) in the ICRF/EME2000 system</returns>
public static Geo.Algorithm.Vector AccelerOfMainForces(double Mjd_TT, Geo.Algorithm.Vector satXyzIcrf, Geo.Algorithm.Vector satVelocityIcrf, double Area, double mass, double solarRadiPresCoeff, double atmDragCoefficient)
{
double Mjd_UT1;
Geo.Algorithm.Vector a = new Geo.Algorithm.Vector(3), r_Sun = new Geo.Algorithm.Vector(3), r_Moon = new Geo.Algorithm.Vector(3);
Matrix T = new Matrix(3, 3), E = new Matrix(3, 3);
// Acceleration due to harmonic gravity field
Mjd_UT1 = Mjd_TT;
T = IERS.NutMatrix(Mjd_TT) * IERS.PrecessionMatrix(MJD_J2000, Mjd_TT);
E = IERS.GreenwichHourAngleMatrix(Mjd_UT1) * T;
a = AccelerOfHarmonicGraviFiled(satXyzIcrf, E, Grav.GM, Grav.R_ref, Grav.CS, Grav.n_max, Grav.m_max);
// Luni-solar perturbations
r_Sun = CelestialUtil.Sun(Mjd_TT);
r_Moon = CelestialUtil.Moon(Mjd_TT);
a += AccelerOfPointMass(satXyzIcrf, r_Sun, OrbitConsts.GM_Sun);
a += AccelerOfPointMass(satXyzIcrf, r_Moon, OrbitConsts.GM_Moon);
// Solar radiation pressure
a += CelestialUtil.Illumination(satXyzIcrf, r_Sun) * AccelerOfSolarRadiPressure(satXyzIcrf, r_Sun, Area, mass, solarRadiPresCoeff, OrbitConsts.PressureOfSolarRadiationPerAU, AU);
// Atmospheric drag
a += AccelerDragOfAtmos(Mjd_TT, satXyzIcrf, satVelocityIcrf, T, Area, mass, atmDragCoefficient);
// Acceleration
return(a);
}
/// <summary>
/// Computes the acceleration of an Earth orbiting satellite due to
/// - the Earth's harmonic gravity field,
/// - the gravitational perturbations of the Sun and Moon
/// - the solar radiation pressure and
/// - the atmospheric drag
/// </summary>
/// <param name="Mjd_TT">Mjd_TT Terrestrial Time (Modified Julian Date)</param>
/// <param name="r">r Satellite position vector in the ICRF/EME2000 system</param>
/// <param name="v">v Satellite velocity vector in the ICRF/EME2000 system</param>
/// <param name="Area"> Area Cross-section </param>
/// <param name="mass"> mass Spacecraft mass</param>
/// <param name="coefOfRadiation">coefOfRadiation Radiation pressure coefficient</param>
/// <param name="coefOfDrag">coefOfDrag Drag coefficient</param>
/// <param name="n"></param>
/// <param name="m"></param>
/// <param name="isConsiderSun">是否启用太阳质量影响</param>
/// <param name="isConsiderMoon">是否启用月亮质量影响</param>
/// <param name="isConsiderSolarRadi">是否启用太阳光压影响</param>
/// <param name="isConsiderDrag">是否考虑大气影响</param>
/// <returns>Acceleration (a=d^2r/dt^2) in the ICRF/EME2000 system</returns>
public Geo.Algorithm.Vector GetAcceleration(double Mjd_TT, Geo.Algorithm.Vector r, Geo.Algorithm.Vector v)
{
// Acceleration due to harmonic gravity field
var Mjd_UT1 = Mjd_TT + (IERS.GetUT1_UTC(Mjd_TT) - IERS.GetTT_UTC(Mjd_TT)) / 86400.0;
var T = IERS.NutMatrix(Mjd_TT) * IERS.PrecessionMatrix(OrbitConsts.MJD_J2000, Mjd_TT);
var E = IERS.GreenwichHourAngleMatrix(Mjd_UT1) * T;
var a = Force.AccelerOfHarmonicGraviFiled(r, E, Force.Grav.GM, Force.Grav.R_ref, Force.Grav.CS, Option.MaxDegree, Option.MaxOrder);
// Luni-solar perturbations
var r_Sun = CelestialUtil.Sun(Mjd_TT);
if (Option.EnableSun)
{
a += Force.AccelerOfPointMass(r, r_Sun, OrbitConsts.GM_Sun);
}
if (Option.EnableMoon)
{
var r_Moon = CelestialUtil.Moon(Mjd_TT);
a += Force.AccelerOfPointMass(r, r_Moon, OrbitConsts.GM_Moon);
}
// Solar radiation pressure
if (Option.EnableSolarRadiation)
{
a += Force.AccelerOfSolarRadiPressure(r, r_Sun, Option.Area, Option.Mass, Option.CoefOfRadiation, OrbitConsts.PressureOfSolarRadiationPerAU, OrbitConsts.AU);
}
// Atmospheric drag
if (Option.EnableDrag)
{
a += Force.AccelerDragOfAtmos(Mjd_TT, r, v, T, Option.Area, Option.Mass, Option.CoefOfDrag);
}
// Acceleration
return(a);
}
static void Main0(string[] args)
{
// Constants
const int N_Step = 8;
const double Step = 0.5; // [d]
// Sample Chebyshev coefficients from JPL DE405 for geocentric lunar
// coordinates in the ITRF(EME2000) frame valid from 2006/03/14 TDB
// to 2006/03/18 TDB
const double t1 = 53808.0; // MJD at start of interval
const double t2 = 53812.0; // MJD at end of interval
double[] Cx_moon = // Coefficients for x-coordinate N_coeff
{
-0.383089044877016277e+06, 0.218158411754834669e+05, 0.179067292901463843e+05,
-0.836928063411765777e+02, -0.628266733052023696e+02, -0.459274434235101225e+00,
0.491167202819885532e-01, 0.770804039287614762e-03, -0.125935992206166816e-03,
0.500271026610991370e-05, 0.107044869185752331e-05, 0.172472464343636242e-08,
-0.269667589576924680e-08
};
double[] Cy_moon = // Coefficients for y-coordinate N_coeff
{
-0.379891721705081436e+05, -0.143611643157166138e+06, 0.187126702787245881e+04,
0.112734362473135207e+04, 0.932891213817359177e+00, -0.191932684130578513e+01,
-0.266517663331897990e-01, 0.104558913448630337e-02, -0.359077689123857890e-04,
-0.123405162037249834e-04, 0.180479239596339495e-06, 0.525522632333670539e-07,
0.543313967008773005e-09
};
double[] Cz_moon = // Coefficients for z-coordinate
{
-0.178496690739133737e+05, -0.788257550331743259e+05, 0.880684692614081882e+03,
0.618395886330471512e+03, 0.103331218594995988e+01, -0.104949867328178592e+01,
-0.150337371962561087e-01, 0.569056416308259317e-03, -0.186297523286550968e-04,
-0.680012420653791955e-05, 0.902057208454410917e-07, 0.287891446432139173e-07,
0.319822827699973363e-09
};
// Chebyshev approximation of lunar coordinates
MultiDimChebshevFitter MoonCheb = new MultiDimChebshevFitter(t1, t2, // Order and interval
new Vector(Cx_moon), // Coefficients
new Vector(Cy_moon),
new Vector(Cz_moon));
// Variables
int i;
double Mjd0, Mjd_TT;
Vector r = new Vector(3);
// Epoch
Mjd0 = DateUtil.DateToMjd(2006, 03, 14, 00, 00, 0.0);
// Output
var info = "Exercise 3-2: Lunar Ephemerides " + "\r\n";
info += " Moon position from low precision analytical theory" + "\r\n";
info += " Date [TT] " + " Position [km] " + "\r\n";
Console.WriteLine(info);
for (i = 0; i <= N_Step; i++)
{
Mjd_TT = Mjd0 + i * Step;
r = CelestialUtil.Moon(Mjd_TT) / 1000.0;
info = " " + DateUtil.MjdToDateTimeString(Mjd_TT) + " " + r;
Console.WriteLine(info);
}
info = " Moon position from DE405" + "\r\n";
info += " Date [TT] " + " Position [km] " + "\r\n";
Console.WriteLine(info);
for (i = 0; i <= N_Step; i++)
{
Mjd_TT = Mjd0 + i * Step;
r = MoonCheb.Fit(Mjd_TT);
info = " " + DateUtil.MjdToDateTimeString(Mjd_TT) + " " + r;
Console.WriteLine(info);
}
Console.ReadKey();
}
