def test_nu2anom_zero():
"""
Make sure that at zero nu=E=M
"""
M_true = 0.0
ecc = 1.2
E_true = 0.0
nu_true = 0.0
E_python, M_python = kepler.nu2anom(nu_true,ecc)
np.testing.assert_allclose((E_python, M_python),(E_true,M_true))
def test_nu2anom_zero():
"""
Make sure that at zero nu=E=M
"""
M_true = 0.0
ecc = 1.2
E_true = 0.0
nu_true = 0.0
E_python, M_python = kepler.nu2anom(nu_true,ecc)
np.testing.assert_allclose((E_python, M_python),(E_true,M_true))
def test_nu2anom_pi():
"""
Make sure that at pi nu=E=M
"""
M_true = np.pi
ecc = 0.9
E_true = np.pi
nu_true = np.pi
E_python, M_python = kepler.nu2anom(nu_true,ecc)
np.testing.assert_allclose(E_python, E_true)
np.testing.assert_allclose(M_python, M_true)
def test_nu2anom():
"""
A matlab test case which is copied here
"""
M_matlab = np.deg2rad(110)
ecc = 0.9
E_matlab = 2.475786297687611
nu_matlab = 2.983273149717047
E_python, M_python = kepler.nu2anom(nu_matlab,ecc)
np.testing.assert_allclose(E_python, E_matlab)
np.testing.assert_allclose(M_python, M_matlab)
def test_nu2anom_pi():
"""
Make sure that at pi nu=E=M
"""
M_true = np.pi
ecc = 0.9
E_true = np.pi
nu_true = np.pi
E_python, M_python = kepler.nu2anom(nu_true,ecc)
np.testing.assert_allclose(E_python, E_true)
np.testing.assert_allclose(M_python, M_true)
def test_nu2anom():
"""
A matlab test case which is copied here
"""
M_matlab = np.deg2rad(110)
ecc = 0.9
E_matlab = 2.475786297687611
nu_matlab = 2.983273149717047
E_python, M_python = kepler.nu2anom(nu_matlab,ecc)
np.testing.assert_allclose(E_python, E_matlab)
np.testing.assert_allclose(M_python, M_matlab)
def test_m(self):
E_true, M_true = kepler.nu2anom(self.nu, self.ecc)
np.testing.assert_allclose(self.m, M_true)
def test_m(self):
E_true, M_true = kepler.nu2anom(self.nu, self.ecc)
np.testing.assert_allclose(self.m, M_true)
def tle_update(self, jd_span, mu=398600.5):
"""Update the state of satellite given a JD time span
This procedure uses the method of general perturbations to update
classical elements from epoch to a future time for inclined elliptical
orbits. It includes the effects due to first order secular rates
(second order for mean motion) caused by drag and J2.
Author: C2C Shankar Kulumani USAFA/CS-19 719-333-4741
Inputs:
deltat - elapsed time since epoch (sec)
n0 - mean motion at epoch (rad/sec)
ndot2 - mean motion rate divided by 2 (rad/sec^2)
ecc0 - eccentricity at epoch (unitless)
eccdot - eccentricity rate (1/sec)
raan0 - RAAN at epoch (rad)
raandot - RAAN rate (rad/sec)
argp0 - argument of periapsis at epoch (rad)
argdot - argument of periapsis rate (rad/sec)
mean0 - mean anomaly at epoch
Outputs:
n - mean motion at epoch + deltat (rad/sec)
ecc - eccentricity at epoch + deltat (unitless)
raan - RAAN at epoch + deltat (rad)
argp - argument of periapsis at epoch + deltat (rad)
nu - true anomaly at epoch + deltat (rad)
Globals: None
Constants: None
Coupling:
newton.m
References:
Astro 321 PREDICT LSN 24-25
"""
deltat = (jd_span - self.epoch_jd) * day2sec
# epoch orbit parameters
n0 = self.n0
ecc0 = self.ecc0
ndot2 = self.ndot2
nddot6 = self.nddot6
raan0 = self.raan0
argp0 = self.argp0
mean0 = self.mean0
inc0 = self.inc0
adot = self.adot
eccdot = self.eccdot
raandot = self.raandot
argpdot = self.argpdot
_, nu0, _ = kepler.kepler_eq_E(mean0, ecc0)
a0 = kepler.n2a(n0, mu)
M0, E0 = kepler.nu2anom(nu0, ecc0)
# propogated elements
a1 = a0 + adot * deltat # either use this or compute a1 from n1 instead
ecc1 = ecc0 + eccdot * deltat
raan1 = raan0 + raandot * deltat
argp1 = argp0 + argpdot * deltat
inc1 = inc0 * np.ones_like(ecc1)
n1 = n0 + ndot2 * 2 * deltat
# a1 = kepler.n2a(n1, mu)
p1 = kepler.semilatus_rectum(a1, ecc1)
M1 = M0 + n0 * deltat+ ndot2 *deltat**2 + nddot6 *deltat**3
M1 = attitude.normalize(M1, 0, 2 * np.pi)
E1, nu1, _ = kepler.kepler_eq_E(M1, ecc1)
# convert to vectors
r_eci, v_eci, _, _ = kepler.coe2rv(p1, ecc1, inc1, raan1, argp1, nu1, mu)
# save all of the variables to the object
self.jd_span = jd_span
self.coe = COE(n=n1, ecc=ecc1, raan=raan1, argp=argp1, mean=M1,
E=E1, nu=nu1, a=a1, p=p1, inc=inc1)
self.r_eci = r_eci
self.v_eci = v_eci
return 0
def tle_update(self, jd_span, mu=398600.5):
"""Update the state of satellite given a JD time span
This procedure uses the method of general perturbations to update
classical elements from epoch to a future time for inclined elliptical
orbits. It includes the effects due to first order secular rates
(second order for mean motion) caused by drag and J2.
Author: C2C Shankar Kulumani USAFA/CS-19 719-333-4741
Inputs:
deltat - elapsed time since epoch (sec)
n0 - mean motion at epoch (rad/sec)
ndot2 - mean motion rate divided by 2 (rad/sec^2)
ecc0 - eccentricity at epoch (unitless)
eccdot - eccentricity rate (1/sec)
raan0 - RAAN at epoch (rad)
raandot - RAAN rate (rad/sec)
argp0 - argument of periapsis at epoch (rad)
argdot - argument of periapsis rate (rad/sec)
mean0 - mean anomaly at epoch
Outputs:
n - mean motion at epoch + deltat (rad/sec)
ecc - eccentricity at epoch + deltat (unitless)
raan - RAAN at epoch + deltat (rad)
argp - argument of periapsis at epoch + deltat (rad)
nu - true anomaly at epoch + deltat (rad)
Globals: None
Constants: None
Coupling:
newton.m
References:
Astro 321 PREDICT LSN 24-25
"""
deltat = (jd_span - self.epoch_jd) * day2sec
# epoch orbit parameters
n0 = self.n0
ecc0 = self.ecc0
ndot2 = self.ndot2
nddot6 = self.nddot6
raan0 = self.raan0
argp0 = self.argp0
mean0 = self.mean0
inc0 = self.inc0
adot = self.adot
eccdot = self.eccdot
raandot = self.raandot
argpdot = self.argpdot
_, nu0, _ = kepler.kepler_eq_E(mean0, ecc0)
a0 = kepler.n2a(n0, mu)
M0, E0 = kepler.nu2anom(nu0, ecc0)
# propogated elements
a1 = a0 + adot * deltat # either use this or compute a1 from n1 instead
ecc1 = ecc0 + eccdot * deltat
raan1 = raan0 + raandot * deltat
argp1 = argp0 + argpdot * deltat
inc1 = inc0 * np.ones_like(ecc1)
n1 = n0 + ndot2 * 2 * deltat
# a1 = kepler.n2a(n1, mu)
p1 = kepler.semilatus_rectum(a1, ecc1)
M1 = M0 + n0 * deltat + ndot2 * deltat**2 + nddot6 * deltat**3
M1 = attitude.normalize(M1, 0, 2 * np.pi)
E1, nu1, _ = kepler.kepler_eq_E(M1, ecc1)
# convert to vectors
r_eci, v_eci, _, _ = kepler.coe2rv(p1, ecc1, inc1, raan1, argp1, nu1,
mu)
# save all of the variables to the object
self.jd_span = jd_span
self.coe = COE(n=n1,
ecc=ecc1,
raan=raan1,
argp=argp1,
mean=M1,
E=E1,
nu=nu1,
a=a1,
p=p1,
inc=inc1)
self.r_eci = r_eci
self.v_eci = v_eci
return 0