keplerianOrbit

class KeplerianOrbit

#include <keplerianOrbit.h>

The KeplerianOrbit class represents an elliptical orbit and provides a coherent set of common outputs such as position and velocity, orbital period, semi-parameter, etc. It uses the utility orbitalMotion to do orbital element to position and velocity conversion.

Public Functions

KeplerianOrbit(): This constructor initialized to an arbitrary orbit

KeplerianOrbit(classicElements oe, const double mu): The constructor requires orbital elements and a gravitational constant value

KeplerianOrbit(const KeplerianOrbit &orig): The copy constructor works with python copy

~KeplerianOrbit(): Generic Destructor

Eigen::Vector3d r_BP_P() const

body position vector relative to planet

Eigen::Vector3d v_BP_P() const

body velocity vector relative to planet

Eigen::Vector3d h_BP_P() const

angular momentum of body relative to planet

double M() const: return mean anomaly angle

double n() const: return mean orbit rate

double P() const

return mean orbit rate

return orbit period

double f() const

return orbital period

return true anomaly

double fDot() const

return true anomaly

return true anomaly rate

double RAAN() const: return right ascencion of the ascending node

double omega() const: return argument of periapses

double i() const: return inclination angle

double e() const: return eccentricty

double a() const: return semi-major axis

double h() const: return orbital angular momentum magnitude

double Energy(): return orbital energy

double r() const: return orbit radius

double v() const: return velocity magnitude

double r_a() const: return radius at apoapses

double r_p() const: return radius at periapses

double fpa() const: return flight path angle

double E() const: return eccentric anomaly angle

double p() const: return semi-latus rectum or the parameter

double rDot() const: return radius rate

double c3() const: return escape velocity

classicElements oe()

This method returns the orbital element set for the orbit

Returns:: classicElements oe

void set_mu(const double mu): This method sets the gravitational constants of the body being orbited

void set_a(double a): set semi-major axis

void set_e(double e): set eccentricity

void set_i(double i): set inclination angle

void set_omega(double omega): set argument of periapsis

void set_RAAN(double RAAN): set right ascension of the ascending node

void set_f(double f): set true anomaly angle

Private Functions

void change_orbit(): This method populates all outputs from orbital elements coherently if any of the classical orbital elements are changed

void change_f(): This method only changes the outputs dependent on true anomaly so that one orbit may be queried at various points along the orbit

Private Members

double mu = MU_EARTH

double semi_major_axis = 1E5

double eccentricity = 1E-5

double inclination = {}

double argument_of_periapsis = {}

double right_ascension = {}

double true_anomaly = {}

double true_anomaly_rate = {}

double orbital_period = {}

double orbital_energy = {}

double v_infinity = {}

double orbit_radius = {}

double radial_rate = {}

double r_apogee = {}

double r_perigee = {}

double semi_parameter = {}

double flight_path_angle = {}

double eccentric_anomaly = {}

double mean_motion = {}

double mean_anomaly = {}

Eigen::Vector3d orbital_angular_momentum_P

Eigen::Vector3d position_BP_P

Eigen::Vector3d velocity_BP_P