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https://github.com/gnss-sdr/gnss-sdr
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3e87d1f204
- The GNSS-SDR code-based Observable generation algorithm for GPS L1 C/A was migrated from common TX time to common RX time. One of the main reasons is that the common TX algorithm was incompatible with the standard RINEX observables, wich requires the oservables at the same time of reception. - Using common Rx time, the code was simplified and now requires less memory due to the lack of symbol buffer - Now it is possible to use standard RINEX post processing software (gpstk and rtklib for instance) to get PVT and the obtained precision is comparable to the internal Least Squares solver. It is possible to use also the carrier phase observable, but is still experimental. - The default RINEX version output was set to 2.11 in this revision. RINEX 3.00 is under verification now. git-svn-id: https://svn.code.sf.net/p/gnss-sdr/code/trunk@364 64b25241-fba3-4117-9849-534c7e92360d
289 lines
8.0 KiB
C++
289 lines
8.0 KiB
C++
/*!
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* \file gps_ephemeris.cc
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* \brief Interface of a GPS EPHEMERIS storage and orbital model functions
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*
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* See http://www.gps.gov/technical/icwg/IS-GPS-200E.pdf Appendix II
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* \author Javier Arribas, 2013. jarribas(at)cttc.es
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*
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* -------------------------------------------------------------------------
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*
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* Copyright (C) 2010-2013 (see AUTHORS file for a list of contributors)
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*
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* GNSS-SDR is a software defined Global Navigation
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* Satellite Systems receiver
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*
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* This file is part of GNSS-SDR.
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*
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* GNSS-SDR is free software: you can redistribute it and/or modify
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* it under the terms of the GNU General Public License as published by
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* the Free Software Foundation, either version 3 of the License, or
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* at your option) any later version.
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*
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* GNSS-SDR is distributed in the hope that it will be useful,
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* but WITHOUT ANY WARRANTY; without even the implied warranty of
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* MERCHANTABILITY or FITNESS FOR A PARTICULAR PURPOSE. See the
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* GNU General Public License for more details.
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*
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* You should have received a copy of the GNU General Public License
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* along with GNSS-SDR. If not, see <http://www.gnu.org/licenses/>.
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*
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* -------------------------------------------------------------------------
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*/
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#include "gps_ephemeris.h"
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Gps_Ephemeris::Gps_Ephemeris()
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{
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i_satellite_PRN=0;
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d_TOW=0;
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d_Crs=0;
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d_Delta_n=0;
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d_M_0=0;
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d_Cuc=0;
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d_e_eccentricity=0;
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d_Cus=0;
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d_sqrt_A=0;
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d_Toe=0;
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d_Toc=0;
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d_Cic=0;
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d_OMEGA0=0;
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d_Cis=0;
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d_i_0=0;
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d_Crc=0;
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d_OMEGA=0;
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d_OMEGA_DOT=0;
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d_IDOT=0;
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i_code_on_L2=0;
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i_GPS_week=0;
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b_L2_P_data_flag=false;
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i_SV_accuracy=0;
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i_SV_health=0;
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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]
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d_IODC=0; //!< Issue of Data, Clock
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i_AODO=0; //!< Age of Data Offset (AODO) term for the navigation message correction table (NMCT) contained in subframe 4 (reference paragraph 20.3.3.5.1.9) [s]
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b_fit_interval_flag=false;//!< indicates the curve-fit interval used by the CS (Block II/IIA/IIR/IIR-M/IIF) and SS (Block IIIA) in determining the ephemeris parameters, as follows: 0 = 4 hours, 1 = greater than 4 hours.
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d_spare1=0;
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d_spare2=0;
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d_A_f0=0; //!< Coefficient 0 of code phase offset model [s]
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d_A_f1=0; //!< Coefficient 1 of code phase offset model [s/s]
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d_A_f2=0; //!< Coefficient 2 of code phase offset model [s/s^2]
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b_integrity_status_flag=false;
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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.
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b_antispoofing_flag=false; //!< If true, the AntiSpoofing mode is ON in that SV
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//Plane A (info from http://www.navcen.uscg.gov/?Do=constellationStatus)
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satelliteBlock[9] = "IIA";
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satelliteBlock[31] = "IIR-M";
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satelliteBlock[8] = "IIA";
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satelliteBlock[7] = "IIR-M";
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satelliteBlock[27] = "IIA";
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//Plane B
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satelliteBlock[16] = "IIR";
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satelliteBlock[25] = "IIF";
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satelliteBlock[28] = "IIR";
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satelliteBlock[12] = "IIR-M";
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satelliteBlock[30] = "IIA";
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//Plane C
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satelliteBlock[29] = "IIR-M";
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satelliteBlock[3] = "IIA";
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satelliteBlock[19] = "IIR";
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satelliteBlock[17] = "IIR-M";
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satelliteBlock[6] = "IIA";
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//Plane D
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satelliteBlock[2] = "IIR";
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satelliteBlock[1] = "IIF";
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satelliteBlock[21] = "IIR";
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satelliteBlock[4] = "IIA";
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satelliteBlock[11] = "IIR";
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satelliteBlock[24] = "IIA"; // Decommissioned from active service on 04 Nov 2011
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//Plane E
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satelliteBlock[20] = "IIR";
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satelliteBlock[22] = "IIR";
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satelliteBlock[5] = "IIR-M";
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satelliteBlock[18] = "IIR";
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satelliteBlock[32] = "IIA";
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satelliteBlock[10] = "IIA";
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//Plane F
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satelliteBlock[14] = "IIR";
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satelliteBlock[15] = "IIR-M";
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satelliteBlock[13] = "IIR";
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satelliteBlock[23] = "IIR";
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satelliteBlock[26] = "IIA";
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}
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double Gps_Ephemeris::check_t(double time)
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{
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double corrTime;
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double half_week = 302400; // seconds
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corrTime = time;
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if (time > half_week)
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{
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corrTime = time - 2*half_week;
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}
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else if (time < -half_week)
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{
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corrTime = time + 2*half_week;
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}
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return corrTime;
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}
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// 20.3.3.3.3.1 User Algorithm for SV Clock Correction.
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//clkDrift= af0 + af1.*(tTX - toc) + af2.*(tTX - toc).^2;
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double Gps_Ephemeris::sv_clock_drift(double transmitTime)
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{
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double dt;
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dt = check_t(transmitTime - d_Toc);
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d_satClkDrift = d_A_f0 + d_A_f1*dt + (d_A_f2 * dt)*(d_A_f2 * dt);
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return d_satClkDrift;
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}
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// compute the relativistic correction term
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double Gps_Ephemeris::sv_clock_relativistic_term(double transmitTime)
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{
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double tk;
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double a;
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double n;
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double n0;
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double E;
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double E_old;
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double dE;
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double M;
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// Restore semi-major axis
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a = d_sqrt_A*d_sqrt_A;
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// Time from ephemeris reference epoch
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tk = check_t(transmitTime - d_Toe);
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// Computed mean motion
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n0 = sqrt(GM / (a*a*a));
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// Corrected mean motion
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n = n0 + d_Delta_n;
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// Mean anomaly
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M = d_M_0 + n * tk;
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// Reduce mean anomaly to between 0 and 2pi
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M = fmod((M + 2*GPS_PI), (2*GPS_PI));
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// Initial guess of eccentric anomaly
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E = M;
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// --- Iteratively compute eccentric anomaly ----------------------------
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for (int ii = 1; ii<20; ii++)
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{
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E_old = E;
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E = M + d_e_eccentricity * sin(E);
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dE = fmod(E - E_old, 2*GPS_PI);
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if (fabs(dE) < 1e-12)
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{
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//Necessary precision is reached, exit from the loop
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break;
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}
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}
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// Compute relativistic correction term
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d_dtr = F * d_e_eccentricity * d_sqrt_A * sin(E);
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return d_dtr;
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}
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void Gps_Ephemeris::satellitePosition(double transmitTime)
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{
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double tk;
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double a;
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double n;
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double n0;
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double M;
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double E;
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double E_old;
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double dE;
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double nu;
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double phi;
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double u;
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double r;
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double i;
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double Omega;
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// Find satellite's position ----------------------------------------------
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// Restore semi-major axis
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a = d_sqrt_A*d_sqrt_A;
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// Time from ephemeris reference epoch
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tk = check_t(transmitTime - d_Toe);
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// Computed mean motion
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n0 = sqrt(GM / (a*a*a));
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// Corrected mean motion
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n = n0 + d_Delta_n;
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// Mean anomaly
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M = d_M_0 + n * tk;
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// Reduce mean anomaly to between 0 and 2pi
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M = fmod((M + 2*GPS_PI), (2*GPS_PI));
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// Initial guess of eccentric anomaly
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E = M;
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// --- Iteratively compute eccentric anomaly ----------------------------
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for (int ii = 1; ii<20; ii++)
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{
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E_old = E;
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E = M + d_e_eccentricity * sin(E);
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dE = fmod(E - E_old, 2*GPS_PI);
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if (fabs(dE) < 1e-12)
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{
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//Necessary precision is reached, exit from the loop
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break;
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}
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}
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// Compute relativistic correction term (now is in sepparated function)
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//d_dtr = F * d_e_eccentricity * d_sqrt_A * sin(E);
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// Compute the true anomaly
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double tmp_Y = sqrt(1.0 - d_e_eccentricity * d_e_eccentricity) * sin(E);
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double tmp_X = cos(E) - d_e_eccentricity;
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nu = atan2(tmp_Y, tmp_X);
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// Compute angle phi (argument of Latitude)
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phi = nu + d_OMEGA;
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// Reduce phi to between 0 and 2*pi rad
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phi = fmod((phi), (2*GPS_PI));
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// Correct argument of latitude
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u = phi + d_Cuc * cos(2*phi) + d_Cus * sin(2*phi);
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// Correct radius
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r = a * (1 - d_e_eccentricity*cos(E)) + d_Crc * cos(2*phi) + d_Crs * sin(2*phi);
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// Correct inclination
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i = d_i_0 + d_IDOT * tk + d_Cic * cos(2*phi) + d_Cis * sin(2*phi);
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// Compute the angle between the ascending node and the Greenwich meridian
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Omega = d_OMEGA0 + (d_OMEGA_DOT - OMEGA_EARTH_DOT)*tk - OMEGA_EARTH_DOT * d_Toe;
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// Reduce to between 0 and 2*pi rad
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Omega = fmod((Omega + 2*GPS_PI), (2*GPS_PI));
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// --- Compute satellite coordinates in Earth-fixed coordinates
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d_satpos_X = cos(u) * r * cos(Omega) - sin(u) * r * cos(i) * sin(Omega);
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d_satpos_Y = cos(u) * r * sin(Omega) + sin(u) * r * cos(i) * cos(Omega);
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d_satpos_Z = sin(u) * r * sin(i);
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// Satellite's velocity. Can be useful for Vector Tracking loops
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double Omega_dot = d_OMEGA_DOT - OMEGA_EARTH_DOT;
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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);
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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);
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d_satvel_Z = d_satpos_Y * sin(i);
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}
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