00001 ////////////////////////////////////////////////////////////////////////////////////////////////// 00002 /*! \file Keplerian.h 00003 * \brief Implementation of the Keplerian Orbit State Representation Class. 00004 * \author $Author: jayhawk_hokie $ 00005 * \version $Revision: 1.6 $ 00006 * \date $Date: 2007/05/21 17:25:29 $ 00007 *////////////////////////////////////////////////////////////////////////////////////////////////// 00008 /* 00009 * \todo include conversion functions 00010 */ 00011 ////////////////////////////////////////////////////////////////////////////////////////////////// 00012 00013 #ifndef __SSF_KEPLERIAN_H__ 00014 #define __SSF_KEPLERIAN_H__ 00015 #include <matrix/Matrix.h> 00016 #include <orbit/orbitstaterep/OrbitStateRepresentation.h> 00017 #include <string> 00018 #include <stdlib.h> 00019 #include <sstream> 00020 using namespace std; 00021 00022 // Defined Values 00023 #define MU 3.986012e5 // Gravitational Parameter for Earth (km^3/s^2) 00024 #define Re 6378.145 // Radius of the Earth (km) 00025 #define J2 0.00108263 // gravitational perturbation constant 00026 #define EARTH_ANGULAR_RATE 7.292115856e-5 // Earth's angular rate (rad/s) 00027 00028 #ifndef PI 00029 #include <math.h> 00030 #define PI M_PI 00031 #endif 00032 00033 namespace O_SESSAME { 00034 00035 /*! \ingroup OrbitStateRepresentation @{ */ 00036 #define NUM_KEPLERIAN_ELEMENTS 6 /*!< number of elements in a Keplerian state representation */ 00037 const int SEMIMAJOR_AXIS = VectorIndexBase + 0; /*!< position of the Semimajor axis, \f$a\f$, in the element vector */ 00038 const int ECCENTRICITY = VectorIndexBase + 1; /*!< position of the eccentricity, \f$e\f$, in the element vector */ 00039 const int INCLINATION = VectorIndexBase + 2; /*!< position of the inclination, \f$i\f$, in the element vector */ 00040 const int LONG_ASC_NODE = VectorIndexBase + 3; /*!< position of the longitude of the ascending node, \f$\Omega\f$, in the element vector */ 00041 const int ARG_PERIGEE = VectorIndexBase + 4; /*!< position of the argument of perigee, \f$\omega\f$, in the element vector */ 00042 const int TRUE_ANOMALY = VectorIndexBase + 5; /*!< position of the true anomaly, \f$\nu\f$, in the element vector */ 00043 00044 #define NUM_KEPLERIAN_PARAMETERS 7 /*!< number of parameters */ 00045 const int ECCENTRIC_ANOMALY = VectorIndexBase + 0; /*!< position of the Eccentric Anomaly, \f$E\f$, in the element vector */ 00046 const int ECCENTRIC_ANOMALY_0 = VectorIndexBase + 1; /*!< position of the Initial Eccentric Anomaly, \f$E_o\f$, in the element vector */ 00047 const int MEAN_ANOMALY = VectorIndexBase + 2; /*!< position of the Mean Anomaly, \f$M\f$, in the element vector */ 00048 const int MEAN_ANOMALY_0 = VectorIndexBase + 3; /*!< position of the Initial Mean Anomaly, \f$M\f$, in the element vector */ 00049 const int ARG_LATTITUDE = VectorIndexBase + 4; /*!< position of the Argument of Latitude, \f$\theta\f$, in the element vector */ 00050 const int LONG_PERIGEE = VectorIndexBase + 5; /*!< position of the Longitude of Perigee, \f$\Pi\f$, in the element vector */ 00051 const int TRUE_LONGITUDE = VectorIndexBase + 6; /*!< position of the True Longitude, \f$\ell\f$, in the element vector */ 00052 00053 /*! @} */ 00054 00055 typedef struct tleStruct { 00056 string satName; /*!< Satellite Name, from Line Zero */ 00057 int satNumber; /*!< Satellite Number, from Line One */ 00058 char satClassification; /*!< Almost always "U"nclassified (L1) */ 00059 int launchYear; /*!< Launched in... (L1) */ 00060 int launchNumber; /*!< Which launch of that year... (L1) */ 00061 string launchPiece; /*!< Which piece of that launch... (L1) */ 00062 int epochYear; /*!< Epoch year for this TLE (L1) */ 00063 double epochDay; /*!< Epoch day and time for this TLE (L1) */ 00064 double meanmotion1stDeriv; /*!< \f$\dot{meanMotion}\f$ (L1) */ 00065 double meanmotion2ndDeriv; /*!< \f$\ddot{meanMotion}\f$ (L1) */ 00066 double bstarDrag; /*!< B* drag term (L1) */ 00067 int ephemerisType; /*!< The type of ephemeris data (L1) */ 00068 int elementNumber; /*!< The TLE # (L1) */ 00069 int checksumLine1; /*!< Checksum for L1 */ 00070 double meanAnomaly; /*!< The TLE reports Mean Anomaly (L2) */ 00071 double eccentricAnomaly; /*!< We calculate Eccentric Anomaly */ 00072 double meanMotion; /*!< We calculate Mean Motion */ 00073 int revolutionNumber; /*!< How many orbits at epoch? (L2) */ 00074 int checksumLine2; /*!< Checksum for L2 */ 00075 }; 00076 00077 00078 00079 /*! \brief Keplerian orbital element representation of the orbital position. 00080 * \ingroup OrbitStateRepresentation 00081 * 00082 * The Keplerian class stores the orbital elements of an orbital position. 00083 * 00084 * \par Example: 00085 * 00086 */ 00087 class Keplerian : public OrbitStateRepresentation 00088 { 00089 public: 00090 00091 virtual ~Keplerian(); 00092 00093 virtual Keplerian* NewPointer(); 00094 00095 virtual Keplerian* Clone(); 00096 00097 Keplerian(); 00098 00099 Keplerian(const Vector& _Elements); 00100 00101 Keplerian KeplerianCopy( ); 00102 00103 void SetKeplerianRepresentationMeanAnomaly( const Vector& _OrbitalElements ); 00104 00105 void SetKeplerianRepresentationEccentricAnomaly( const Vector& _OrbitalElements ); 00106 00107 void SetKeplerianRepresentationTrueAnomaly( const Vector& _OrbitalElements ); 00108 00109 void SetPositionVelocity(const Vector& _Position, const Vector& _Velocity); 00110 00111 void SetPositionVelocity(const Vector& _PositionVelocity); 00112 00113 void SetPositionVelocity(const Vector& _Position, const Vector& _Velocity, const OrbitFrame& _OrbFrame); 00114 00115 void SetPositionVelocity(const Vector& _PositionVelocity, const OrbitFrame& _OrbFrame); 00116 00117 Vector GetPositionVelocity() const; 00118 00119 Vector GetPositionVelocity(const OrbitFrame& _TargetOrbFrame) const; 00120 00121 void GetPositionVelocity(Vector& _Position, Vector& _Velocity) const; 00122 00123 void GetPositionVelocity(Vector& _Position, Vector& _Velocity, const OrbitFrame& _TargetOrbFrame) const; 00124 00125 Vector GetPositionVelocityPQW() const; 00126 00127 double GetEccentricAnomalyFromMeanAnomaly(const double& _MeanAnomaly); 00128 00129 void GetTrueAnomalyFromEccentricAnomaly(const double& _EccentricAnomaly); 00130 00131 tleStruct ReadTwoLineElementSet(const string& _TwoLineElementSet); 00132 00133 double GetEccentricAnomalyFromTrueAnomaly(); 00134 00135 double GetMeanAnomalyFromEccentricAnomaly(); 00136 00137 double GetArgLattitude(); 00138 00139 double GetLongPerigee(); 00140 00141 double GetTrueLongitude(); 00142 00143 00144 /* ********* KEPLERIAN SPECIFIC FUNCTIONS ********* */ 00145 00146 /*! @defgroup ElementAccessFunctions Element Access functions 00147 * \ingroup OrbitStateRepresentation 00148 * \brief Functions to access the specific elements of the Keplerian Orbital Elements. 00149 * @{ 00150 */ 00151 /*! \brief Compute and return the Semiparameter \f$p\f$ of the orbit position. 00152 * @return the semiparameter: \f$p = a * (1-e^{2})\f$. (km) 00153 */ 00154 inline double GetSemiParameter() const {return GetSemimajorAxis() * (1 - pow(GetEccentricity(),2));}; 00155 00156 /*! \brief Return the Semimajor Axis \f$a\f$ of the orbit position. (km) */ 00157 inline double GetSemimajorAxis() const {return m_OrbitalElements(SEMIMAJOR_AXIS);}; 00158 00159 /*! \brief Return the Eccentricity \f$e\f$ of the orbit position. (-) */ 00160 inline double GetEccentricity() const {return m_OrbitalElements(ECCENTRICITY);}; 00161 00162 /*! \brief Return the Inclination \f$i\f$ of the orbit position. (-) */ 00163 inline double GetInclination() const {return m_OrbitalElements(INCLINATION);}; 00164 00165 /*! \brief Return the Longitude of the Ascending Node \f$\Omega\f$ of the orbit position. (rad) */ 00166 inline double GetLongAscNode() const {return m_OrbitalElements(LONG_ASC_NODE);}; 00167 00168 /*! \brief Return the Argument of Perigee \f$\omega\f$ of the orbit position. (rad) */ 00169 inline double GetArgPerigee() const {return m_OrbitalElements(ARG_PERIGEE);}; 00170 00171 /*! \brief Return the True Anomaly \f$\nu\f$ of the orbit position. (rad) */ 00172 inline double GetTrueAnomaly() const {return m_OrbitalElements( TRUE_ANOMALY );}; 00173 00174 /*! \brief Return the Mean orbital motion \f$n\f$ of the orbit position. (rad/s) */ 00175 inline double GetMeanMotion() const {return sqrt(MU/pow(GetSemimajorAxis(),3));}; 00176 00177 /*! \brief Return the Eccentric Anomaly \f$E\f$ of the orbit position. (rad) */ 00178 inline double GetEccentricAnomaly() const {return m_OrbitalParameters( ECCENTRIC_ANOMALY );}; 00179 00180 /*! \brief Return the Mean Anomaly \f$M\f$ of the orbit position. (-) */ 00181 inline double GetMeanAnomaly() const {return m_OrbitalParameters( MEAN_ANOMALY );}; 00182 /*! @} */ 00183 00184 /* ********* ********* ******** ********* ********* */ 00185 00186 00187 00188 00189 00190 /* ********* DEPRECATED FUNCTIONS ********* */ 00191 00192 virtual void SetState(const Vector& _Elements); 00193 00194 virtual Vector GetState() const; 00195 00196 virtual void GetState(Vector& _Elements) const; 00197 00198 /* ********* ********** ********* ********* */ 00199 00200 00201 00202 private: 00203 /*! 6x1 vector of Keplerian orbital elements [a, e, i, \f$\Omega\f$, \f$\omega\f$, \f$\nu\f$ ] (km, -, rad, rad, rad, rad) */ 00204 Vector m_OrbitalElements; 00205 00206 /*! nx1 vector of Keplerian orbital elements [E, E_o, M, Mo ] (rad, rad, -, - ) */ 00207 Vector m_OrbitalParameters; 00208 00209 /*! All of the non-orbital element data that's inside of a TLE. */ 00210 tleStruct m_tleData; 00211 }; 00212 } // close namespace O_SESSAME 00213 00214 #endif 00215 00216 // Do not change the comments below - they will be added automatically by CVS 00217 /***************************************************************************** 00218 * $Log: Keplerian.h,v $ 00219 * Revision 1.6 2007/05/21 17:25:29 jayhawk_hokie 00220 * *** empty log message *** 00221 * 00222 * Revision 1.5 2006/08/25 15:49:28 jayhawk_hokie 00223 * Added set Keplerian representation for True Anomaly, Eccentric Anomaly, and Mean Anomaly. 00224 * 00225 * Revision 1.4 2006/08/23 21:45:11 jayhawk_hokie 00226 * Updated ECI to Keplerian and Keplerian to ECI. Also tested functions for special eliptical orbits. 00227 * 00228 * Revision 1.3 2005/06/29 20:26:57 jayhawk_hokie 00229 * Fixed COE equations due to being incomplete. 00230 * 00231 * Revision 1.2 2005/06/10 12:53:28 jayhawk_hokie 00232 * Changed header file format to make it easier to link to the spacecraft code by only having to specify dsacss/dep/spacecraft/src (ex. "file.h" changed to <dir/file.h> where dir is a folder in src). 00233 * 00234 * Revision 1.1.1.1 2005/04/26 17:41:00 cakinli 00235 * Adding OpenSESSAME to DSACSS distrib to capture fixed version. 00236 * 00237 * Revision 1.12 2003/08/24 20:59:13 nilspace 00238 * Updated. 00239 * 00240 * Revision 1.11 2003/05/23 19:28:14 simpliciter 00241 * Moved comments to implementation file, basic housekeeping. 00242 * 00243 * Revision 1.10 2003/05/20 20:49:42 simpliciter 00244 * Added new functions for parsing TLEs: 00245 * - GetEccentricAnomalyFromMeanAnomaly, 00246 * - GetTrueAnomalyFromEccentricAnomaly, 00247 * - ReadTwoLineElementSet. 00248 * 00249 * Revision 1.9 2003/05/13 18:47:56 nilspace 00250 * Fixed comments for better formatting. 00251 * 00252 * Revision 1.8 2003/05/05 20:46:38 nilspace 00253 * Added inertial Get/SetPositionVelocity to conform to new OrbitStateRepresentation abstract class. 00254 * 00255 * Revision 1.7 2003/05/02 16:16:46 nilspace 00256 * Documented the API. 00257 * 00258 * Revision 1.6 2003/04/29 18:48:30 nilspace 00259 * Added NewPointer and Clone functions to help in getting the correct memory allocation. 00260 * 00261 * Revision 1.5 2003/04/24 21:14:23 nilspace 00262 * const'd all Get() functions. 00263 * 00264 * Revision 1.4 2003/04/24 20:10:47 nilspace 00265 * const'd all Get() functions. 00266 * 00267 * Revision 1.3 2003/04/23 18:52:29 nilspace 00268 * Updated to call correct OrbitFrame::GetRotation calls. 00269 * Added temporary PI and MU values. 00270 * Added K_Vector Values. 00271 * 00272 * Revision 1.2 2003/04/22 18:06:08 nilspace 00273 * Added math for Matthew Berry. 00274 * 00275 * Revision 1.1 2003/04/08 22:47:35 nilspace 00276 * Initial Submission. 00277 * 00278 ******************************************************************************/
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