mirror of https://github.com/gnss-sdr/gnss-sdr
295 lines
8.6 KiB
C++
295 lines
8.6 KiB
C++
/*!
|
|
* \file gps_cnav_ephemeris.cc
|
|
* \brief Interface of a GPS CNAV EPHEMERIS storage and orbital model functions
|
|
*
|
|
* See http://www.gps.gov/technical/icwg/IS-GPS-200E.pdf Appendix II
|
|
* \author Javier Arribas, 2015. jarribas(at)cttc.es
|
|
*
|
|
* -------------------------------------------------------------------------
|
|
*
|
|
* Copyright (C) 2010-2015 (see AUTHORS file for a list of contributors)
|
|
*
|
|
* GNSS-SDR is a software defined Global Navigation
|
|
* Satellite Systems receiver
|
|
*
|
|
* This file is part of GNSS-SDR.
|
|
*
|
|
* GNSS-SDR is free software: you can redistribute it and/or modify
|
|
* it under the terms of the GNU General Public License as published by
|
|
* the Free Software Foundation, either version 3 of the License, or
|
|
* (at your option) any later version.
|
|
*
|
|
* GNSS-SDR is distributed in the hope that it will be useful,
|
|
* but WITHOUT ANY WARRANTY; without even the implied warranty of
|
|
* MERCHANTABILITY or FITNESS FOR A PARTICULAR PURPOSE. See the
|
|
* GNU General Public License for more details.
|
|
*
|
|
* You should have received a copy of the GNU General Public License
|
|
* along with GNSS-SDR. If not, see <http://www.gnu.org/licenses/>.
|
|
*
|
|
* -------------------------------------------------------------------------
|
|
*/
|
|
|
|
#include "gps_cnav_ephemeris.h"
|
|
#include <cmath>
|
|
|
|
Gps_CNAV_Ephemeris::Gps_CNAV_Ephemeris()
|
|
{
|
|
i_satellite_PRN = 0;
|
|
|
|
d_Toe1 = -1;
|
|
d_Toe2 = -1;
|
|
|
|
d_TOW = 0;
|
|
d_Crs = 0;
|
|
d_M_0 = 0;
|
|
d_Cuc = 0;
|
|
d_e_eccentricity = 0;
|
|
d_Cus = 0;
|
|
|
|
d_Toc = 0;
|
|
d_Cic = 0;
|
|
d_OMEGA0 = 0;
|
|
d_Cis = 0;
|
|
d_i_0 = 0;
|
|
d_Crc = 0;
|
|
d_OMEGA = 0;
|
|
d_IDOT = 0;
|
|
|
|
i_GPS_week = 0;
|
|
|
|
d_TGD = 0; // Estimated Group Delay Differential: L1-L2 correction term only for the benefit of "L1 P(Y)" or "L2 P(Y)" s users [s]
|
|
|
|
d_A_f0 = 0; // Coefficient 0 of code phase offset model [s]
|
|
d_A_f1 = 0; // Coefficient 1 of code phase offset model [s/s]
|
|
d_A_f2 = 0; // Coefficient 2 of code phase offset model [s/s^2]
|
|
|
|
b_integrity_status_flag = false;
|
|
b_alert_flag = false; // If true, indicates that the SV URA may be worse than indicated in d_SV_accuracy, use that SV at our own risk.
|
|
b_antispoofing_flag = false; // If true, the AntiSpoofing mode is ON in that SV
|
|
|
|
d_satClkDrift = 0.0;
|
|
d_dtr = 0.0;
|
|
d_satpos_X = 0.0;
|
|
d_satpos_Y = 0.0;
|
|
d_satpos_Z = 0.0;
|
|
d_satvel_X = 0.0;
|
|
d_satvel_Y = 0.0;
|
|
d_satvel_Z = 0.0;
|
|
|
|
i_URA = 0;
|
|
i_signal_health = 0;
|
|
d_Top = 0.0;
|
|
d_DELTA_A = 0.0;
|
|
d_A_DOT = 0.0;
|
|
d_Delta_n = 0.0;
|
|
d_DELTA_DOT_N = 0.0;
|
|
d_DELTA_OMEGA_DOT = 0.0;
|
|
d_URA0 = 0.0;
|
|
d_URA1 = 0.0;
|
|
d_URA2 = 0.0;
|
|
d_ISCL1 = 0.0;
|
|
d_ISCL2 = 0.0;
|
|
d_ISCL5I = 0.0;
|
|
d_ISCL5Q = 0.0;
|
|
b_l2c_phasing_flag = false;
|
|
}
|
|
|
|
double Gps_CNAV_Ephemeris::check_t(double time)
|
|
{
|
|
double corrTime;
|
|
double half_week = 302400.0; // seconds
|
|
corrTime = time;
|
|
if (time > half_week)
|
|
{
|
|
corrTime = time - 2.0 * half_week;
|
|
}
|
|
else if (time < -half_week)
|
|
{
|
|
corrTime = time + 2.0 * half_week;
|
|
}
|
|
return corrTime;
|
|
}
|
|
|
|
|
|
// 20.3.3.3.3.1 User Algorithm for SV Clock Correction.
|
|
double Gps_CNAV_Ephemeris::sv_clock_drift(double transmitTime)
|
|
{
|
|
double dt;
|
|
dt = check_t(transmitTime - d_Toc);
|
|
d_satClkDrift = d_A_f0 + d_A_f1 * dt + d_A_f2 * (dt * dt) + sv_clock_relativistic_term(transmitTime);
|
|
|
|
// Correct satellite group delay
|
|
d_satClkDrift -= d_TGD;
|
|
|
|
return d_satClkDrift;
|
|
}
|
|
|
|
|
|
// compute the relativistic correction term
|
|
double Gps_CNAV_Ephemeris::sv_clock_relativistic_term(double transmitTime)
|
|
{
|
|
double tk;
|
|
double a;
|
|
double n;
|
|
double n0;
|
|
double E;
|
|
double E_old;
|
|
double dE;
|
|
double M;
|
|
const double GM = 3.986005e14; //!< Universal gravitational constant times the mass of the Earth, [m^3/s^2]
|
|
const double F = -4.442807633e-10; //!< Constant, [s/(m)^(1/2)]
|
|
const double A_REF = 26559710.0; // See IS-GPS-200H, pp. 163
|
|
double d_sqrt_A = sqrt(A_REF + d_DELTA_A);
|
|
|
|
// Restore semi-major axis
|
|
a = d_sqrt_A * d_sqrt_A;
|
|
|
|
// Time from ephemeris reference epoch
|
|
tk = check_t(transmitTime - d_Toe1);
|
|
|
|
// Computed mean motion
|
|
n0 = sqrt(GM / (a * a * a));
|
|
// Corrected mean motion
|
|
n = n0 + d_Delta_n;
|
|
// Mean anomaly
|
|
M = d_M_0 + n * tk;
|
|
|
|
// Reduce mean anomaly to between 0 and 2pi
|
|
//M = fmod((M + 2.0 * GPS_L2_PI), (2.0 * GPS_L2_PI));
|
|
|
|
// Initial guess of eccentric anomaly
|
|
E = M;
|
|
|
|
// --- Iteratively compute eccentric anomaly ----------------------------
|
|
for (int ii = 1; ii < 20; ii++)
|
|
{
|
|
E_old = E;
|
|
E = M + d_e_eccentricity * sin(E);
|
|
dE = fmod(E - E_old, 2.0 * PI);
|
|
if (fabs(dE) < 1e-12)
|
|
{
|
|
//Necessary precision is reached, exit from the loop
|
|
break;
|
|
}
|
|
}
|
|
|
|
// Compute relativistic correction term
|
|
d_dtr = F * d_e_eccentricity * d_sqrt_A * sin(E);
|
|
return d_dtr;
|
|
}
|
|
|
|
|
|
double Gps_CNAV_Ephemeris::satellitePosition(double transmitTime)
|
|
{
|
|
double tk;
|
|
double a;
|
|
double n;
|
|
double n0;
|
|
double M;
|
|
double E;
|
|
double E_old;
|
|
double dE;
|
|
double nu;
|
|
double phi;
|
|
double u;
|
|
double r;
|
|
double i;
|
|
double Omega;
|
|
|
|
const double A_REF = 26559710.0; // See IS-GPS-200H, pp. 170
|
|
double d_sqrt_A = sqrt(A_REF + d_DELTA_A);
|
|
|
|
const double GM = 3.986005e14; //!< Universal gravitational constant times the mass of the Earth, [m^3/s^2]
|
|
const double OMEGA_DOT_REF = -2.6e-9; // semicircles / s, see IS-GPS-200H pp. 164
|
|
const double OMEGA_EARTH_DOT = 7.2921151467e-5; //!< Earth rotation rate, [rad/s]
|
|
// Find satellite's position ----------------------------------------------
|
|
|
|
// Restore semi-major axis
|
|
a = d_sqrt_A*d_sqrt_A;
|
|
|
|
// Time from ephemeris reference epoch
|
|
tk = check_t(transmitTime - d_Toe1);
|
|
|
|
// Computed mean motion
|
|
n0 = sqrt(GM / (a*a*a));
|
|
|
|
// Mean motion difference from computed value
|
|
|
|
double delta_n_a = d_Delta_n + 0.5 * d_DELTA_DOT_N * tk;
|
|
|
|
// Corrected mean motion
|
|
n = n0 + delta_n_a;
|
|
|
|
// Mean anomaly
|
|
M = d_M_0 + n * tk;
|
|
|
|
// Reduce mean anomaly to between 0 and 2pi
|
|
//M = fmod((M + 2 * GPS_L2_PI), (2 * GPS_L2_PI));
|
|
|
|
|
|
// Initial guess of eccentric anomaly
|
|
E = M;
|
|
|
|
// --- Iteratively compute eccentric anomaly ----------------------------
|
|
for (int ii = 1; ii < 20; ii++)
|
|
{
|
|
E_old = E;
|
|
E = M + d_e_eccentricity * sin(E);
|
|
dE = fmod(E - E_old, 2 * PI);
|
|
if (fabs(dE) < 1e-12)
|
|
{
|
|
//Necessary precision is reached, exit from the loop
|
|
break;
|
|
}
|
|
}
|
|
|
|
// Compute the true anomaly
|
|
double tmp_Y = sqrt(1.0 - d_e_eccentricity * d_e_eccentricity) * sin(E);
|
|
double tmp_X = cos(E) - d_e_eccentricity;
|
|
nu = atan2(tmp_Y, tmp_X);
|
|
|
|
// Compute angle phi (argument of Latitude)
|
|
phi = nu + d_OMEGA;
|
|
|
|
// Reduce phi to between 0 and 2*pi rad
|
|
//phi = fmod((phi), (2*GPS_L2_PI));
|
|
|
|
// Correct argument of latitude
|
|
u = phi + d_Cuc * cos(2*phi) + d_Cus * sin(2*phi);
|
|
|
|
// Correct radius
|
|
r = a * (1 - d_e_eccentricity*cos(E)) + d_Crc * cos(2*phi) + d_Crs * sin(2*phi);
|
|
|
|
// Correct inclination
|
|
i = d_i_0 + d_IDOT * tk + d_Cic * cos(2*phi) + d_Cis * sin(2*phi);
|
|
|
|
// Compute the angle between the ascending node and the Greenwich meridian
|
|
double d_OMEGA_DOT = OMEGA_DOT_REF*PI + d_DELTA_OMEGA_DOT;
|
|
Omega = d_OMEGA0 + (d_OMEGA_DOT - OMEGA_EARTH_DOT)*tk - OMEGA_EARTH_DOT * d_Toe1;
|
|
|
|
// Reduce to between 0 and 2*pi rad
|
|
//Omega = fmod((Omega + 2*GPS_L2_PI), (2*GPS_L2_PI));
|
|
|
|
// --- Compute satellite coordinates in Earth-fixed coordinates
|
|
d_satpos_X = cos(u) * r * cos(Omega) - sin(u) * r * cos(i) * sin(Omega);
|
|
d_satpos_Y = cos(u) * r * sin(Omega) + sin(u) * r * cos(i) * cos(Omega);
|
|
d_satpos_Z = sin(u) * r * sin(i);
|
|
|
|
// Satellite's velocity. Can be useful for Vector Tracking loops
|
|
double Omega_dot = d_OMEGA_DOT - OMEGA_EARTH_DOT;
|
|
d_satvel_X = - Omega_dot * (cos(u) * r + sin(u) * r * cos(i)) + d_satpos_X * cos(Omega) - d_satpos_Y * cos(i) * sin(Omega);
|
|
d_satvel_Y = Omega_dot * (cos(u) * r * cos(Omega) - sin(u) * r * cos(i) * sin(Omega)) + d_satpos_X * sin(Omega) + d_satpos_Y * cos(i) * cos(Omega);
|
|
d_satvel_Z = d_satpos_Y * sin(i);
|
|
|
|
// Time from ephemeris reference clock
|
|
tk = check_t(transmitTime - d_Toc);
|
|
|
|
double dtr_s = d_A_f0 + d_A_f1 * tk + d_A_f2 * tk * tk;
|
|
|
|
/* relativity correction */
|
|
dtr_s -= 2.0 * sqrt(GM * a) * d_e_eccentricity * sin(E) / (SPEED_OF_LIGHT * SPEED_OF_LIGHT);
|
|
|
|
return dtr_s;
|
|
}
|