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kusanagi |
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//
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// cNoradSGP4.cpp
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//
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// NORAD SGP4 implementation. See historical note in cNoradBase.cpp
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// Copyright (c) 2003 Michael F. Henry
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//
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// mfh 12/07/2003
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//
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#include "stdafx.h"
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#include "cNoradSGP4.h"
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#include "math.h"
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#include "cJulian.h"
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#include "cOrbit.h"
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#include "cVector.h"
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#include "coord.h"
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//////////////////////////////////////////////////////////////////////////////
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cNoradSGP4::cNoradSGP4(const cOrbit &orbit) :
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cNoradBase(orbit)
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{
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m_c5 = 2.0 * m_coef1 * m_aodp * m_betao2 *
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(1.0 + 2.75 * (m_etasq + m_eeta) + m_eeta * m_etasq);
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m_omgcof = m_Orbit.BStar() * m_c3 * cos(m_Orbit.ArgPerigee());
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m_xmcof = -TWOTHRD * m_coef * m_Orbit.BStar() * AE / m_eeta;
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m_delmo = pow(1.0 + m_eta * cos(m_Orbit.mnAnomaly()), 3.0);
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m_sinmo = sin(m_Orbit.mnAnomaly());
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}
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cNoradSGP4::~cNoradSGP4(void)
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{
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}
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//////////////////////////////////////////////////////////////////////////////
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// getPosition()
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// This procedure returns the ECI position and velocity for the satellite
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// in the orbit at the given number of minutes since the TLE epoch time
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// using the NORAD Simplified General Perturbation 4, near earth orbit
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// model.
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//
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// tsince - Time in minutes since the TLE epoch (GMT).
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// eci - ECI object to hold position information.
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// To convert the returned ECI position vector to km,
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// multiply each component by:
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// (XKMPER_WGS72 / AE).
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// To convert the returned ECI velocity vector to km/sec,
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// multiply each component by:
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// (XKMPER_WGS72 / AE) * (MIN_PER_DAY / 86400).
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bool cNoradSGP4::getPosition(double tsince, cEci &eci)
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{
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// For m_perigee less than 220 kilometers, the isimp flag is set and
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// the equations are truncated to linear variation in sqrt a and
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// quadratic variation in mean anomaly. Also, the m_c3 term, the
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// delta omega term, and the delta m term are dropped.
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bool isimp = false;
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if ((m_aodp * (1.0 - m_satEcc) / AE) < (220.0 / XKMPER_WGS72 + AE))
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{
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isimp = true;
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}
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double d2 = 0.0;
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double d3 = 0.0;
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double d4 = 0.0;
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double t3cof = 0.0;
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double t4cof = 0.0;
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double t5cof = 0.0;
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if (!isimp)
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{
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double c1sq = m_c1 * m_c1;
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d2 = 4.0 * m_aodp * m_tsi * c1sq;
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double temp = d2 * m_tsi * m_c1 / 3.0;
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d3 = (17.0 * m_aodp + m_s4) * temp;
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d4 = 0.5 * temp * m_aodp * m_tsi *
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(221.0 * m_aodp + 31.0 * m_s4) * m_c1;
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t3cof = d2 + 2.0 * c1sq;
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t4cof = 0.25 * (3.0 * d3 + m_c1 * (12.0 * d2 + 10.0 * c1sq));
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t5cof = 0.2 * (3.0 * d4 + 12.0 * m_c1 * d3 + 6.0 *
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d2 * d2 + 15.0 * c1sq * (2.0 * d2 + c1sq));
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}
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// Update for secular gravity and atmospheric drag.
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double xmdf = m_Orbit.mnAnomaly() + m_xmdot * tsince;
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double omgadf = m_Orbit.ArgPerigee() + m_omgdot * tsince;
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double xnoddf = m_Orbit.RAAN() + m_xnodot * tsince;
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double omega = omgadf;
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double xmp = xmdf;
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double tsq = tsince * tsince;
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double xnode = xnoddf + m_xnodcf * tsq;
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double tempa = 1.0 - m_c1 * tsince;
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double tempe = m_Orbit.BStar() * m_c4 * tsince;
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double templ = m_t2cof * tsq;
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if (!isimp)
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{
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double delomg = m_omgcof * tsince;
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double delm = m_xmcof * (pow(1.0 + m_eta * cos(xmdf), 3.0) - m_delmo);
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double temp = delomg + delm;
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xmp = xmdf + temp;
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omega = omgadf - temp;
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double tcube = tsq * tsince;
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double tfour = tsince * tcube;
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tempa = tempa - d2 * tsq - d3 * tcube - d4 * tfour;
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tempe = tempe + m_Orbit.BStar() * m_c5 * (sin(xmp) - m_sinmo);
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templ = templ + t3cof * tcube + tfour * (t4cof + tsince * t5cof);
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}
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double a = m_aodp * sqr(tempa);
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double e = m_satEcc - tempe;
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double xl = xmp + omega + xnode + m_xnodp * templ;
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double xn = XKE / pow(a, 1.5);
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return FinalPosition(m_satInc, omgadf, e, a, xl, xnode, xn, tsince, eci);
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}
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