CL_ex_secularJ2

Description

• Propagates orbital elements considering the mean secular effects due to J2 only.

  The jacobian of the transformation (i.e. the transition matrix) is also optionally computed.

• The type of orbital elements is the following:

  

• Notes:

  - There can be 1 or N initial times, and 1 or N final times.

  - This function works for elliptical orbits only.

Parameters

t1:

Initial time [days] (1x1 or 1xN)

kep_t1:

Orbital elements at time t1 (6x1 or 6xN).

t2:

Final time [days] (1xN or 1x1).

er:

(optional) Equatorial radius [m] (default is %CL_eqRad)

mu:

(optional) Gravitational constant [m^3/s^2] (default value is %CL_mu)

j2:

(optional) (Unnormalized) zonal coefficient (second zonal harmonic) (default is %CL_j1jn(2))

kep_t2:

Orbital elements propagated to time t2. (6xN)

jac:

Transition matrix = d(kep_t2)/d(kep_t1) (6x6xN)

Authors

• CNES - DCT/SB

See also

• CL_ex_propagate

• CL_op_driftJ2

Examples

// propagation of one satellite to several times: sma = 7.e6; ecc = 0.001; inc = CL_deg2rad(98); pom = CL_deg2rad(90); gom = 0; anm = 0; t1 = 21915; t2 = t1:0.1:t1+1; kep_t1 = [sma; ecc; inc; pom; gom; anm]; [kep_t2] = CL_ex_secularJ2(t1, kep_t1, t2); // propagation of 2 element sets to one final time: t2 = t1+1; t1 = [t1, t1+0.5]; kep_t1 = [kep_t1, [sma+1; ecc+0.1; inc; pom; gom; anm]]; [kep_t2] = CL_ex_secularJ2(t1, kep_t1, t2);