|
SUPPORT SITE |
|
|
|
SUPPORT SITE |
|
Path: Orbit/Interplanetary
% Computes the heliocentric orbital elements after an Earth escape spiral.
Performs a simulation of the spiral out from Earth using ode113 and
FCRTBPRHS, and stops once the sun gravity acceleration dominates that
of Earth.
Since version 2014.1
--------------------------------------------------------------------------
Form:
[elHelio, rSEMI, vSEMI, t1, mass] = EarthOrbToHelioOrb( jD0, elEarth, SCParams)
--------------------------------------------------------------------------
------
Inputs
------
jD0 (1,:) Julian Date
elEarth (1,6) Initial orbital elements in the ECI frame
SCParams (.) Design parameters for the spacecraft
.mass0 Mass (kg)
.thrust Thrust (kN)
.Isp Specific impulse (sec)
.uE Exhaust velocity (km/s)
-------
Outputs
-------
elHelio (:,6) Orbital elements in the heliocentric frame
rSEMI (3,:) Position vector in the heliocentric frame
vSEMI (3,:) Velocity vector in the heliocentric frame
t1 (1,:) Time vector
--------------------------------------------------------------------------
See also SunEarthMoonSystemConstants, TransformECIToSEMR,
CRTBPRHS, FCRTBPRHS, LowThrustCRTBP_StopFcn, SEMRToSEMI
--------------------------------------------------------------------------
Common: Graphics/Plot2D Common: Time/Date2JD Math: Linear/Unit Orbit: OrbitCoord/SEMRToSEMI Orbit: OrbitCoord/TransformECIToSEMR Orbit: OrbitData/SunEarthMoonSystemConstants Orbit: ThreeBody/CRTBPRHS Orbit: ThreeBody/FCRTBPRHS Orbit: ThreeBody/LowThrustCRTBP_StopFcn SC: BasicOrbit/El2RV SC: BasicOrbit/RV2El
Back to the Orbit Module page