Library

Documentation for SatelliteToolbox.jl.

SatelliteToolbox.EOPData_IAU1980Type
EOPData_IAU1980{T}

EOP Data for IAU 1980.

Fields

  • x, y: Polar motion with respect to the crust [arcsec].
  • UT1_UTC: Irregularities of the rotation angle [s].
  • LOD: Length of day offset [s].
  • dPsi, dEps: Celestial pole offsets referred to the model IAU1980 [arcsec].
  • *_err: Errors in the components [same unit as the component].

Remarks

Each field will be an AbstractInterpolation indexed by the Julian Day. Hence, if one want to obtain, for example, the X component of the polar motion with respect to the crust at 19 June 2018, the following can be used:

x[DatestoJD(2018,19,06,0,0,0)]
source
SatelliteToolbox.EOPData_IAU2000AType
EOPData_IAU2000A{T}

EOP Data for IAU 2000A.

Fields

  • x, y: Polar motion with respect to the crust [arcsec].
  • UT1_UTC: Irregularities of the rotation angle [s].
  • LOD: Length of day offset [s].
  • dX, dY: Celestial pole offsets referred to the model IAU2000A [arcsec].
  • *_err: Errors in the components [same unit as the component].

Remarks

Each field will be an AbstractInterpolation indexed by the Julian Day. Hence, if one want to obtain, for example, the X component of the polar motion with respect to the crust at 19 June 2018, the following can be used:

x[DatestoJD(2018,19,06,0,0,0)]
source
SatelliteToolbox.J2_GravCteType
J2_GravCte{T}

Gravitational constants for J2 orbit propagator.

Fields

  • R0: Earth equatorial radius [m].
  • μm: √GM [er/s]^(3/2).
  • J2: The second gravitational zonal harmonic of the Earth.
source
SatelliteToolbox.J4_GravCteType
J4_GravCte{T}

Gravitational constants for J4 orbit propagator.

Fields

  • R0: Earth equatorial radius [m].
  • μm: √GM [er/s]^(3/2).
  • J2: The second gravitational zonal harmonic of the Earth.
  • J4: The fourth gravitational zonal harmonic of the Earth.
source
SatelliteToolbox.JB2008_OutputType
JB2008_Output

Output structure of the Jacchia-Bowman 2008.

Fields

  • nN2: Number density of N₂ [1/m³].
  • nO2: Number density of O₂ [1/m³].
  • nO: Number density of O [1/m³].
  • nAr: Number density of Ar [1/m³].
  • nHe: Number density of He [1/m³].
  • nH: Number density of H [1/m³].
  • rho: Total density [kg/m³].
  • T_exo: Exospheric temperature [K].
  • Tz: Temperature at the selected altitude [K].
source
SatelliteToolbox.JR1971_OutputType
JR1971_Output

Output structure of the Jacchia-Roberts 1971 model.

Fields

  • nN2: Number density of N₂ [1/m³].
  • nO2: Number density of O₂ [1/m³].
  • nO: Number density of O [1/m³].
  • nAr: Number density of Ar [1/m³].
  • nHe: Number density of He [1/m³].
  • nH: Number density of H [1/m³].
  • rho: Total density [kg/m³].
  • T_exo: Exospheric temperature [K].
  • Tz: Temperature at the selected altitude [K].
source
SatelliteToolbox.KeplerianElementsType
KeplerianElements{T1,T2}

This structure defines the orbit in terms of the Keplerian elements.

Fields

  • t: Epoch.
  • a: Semi-major axis [m].
  • e: Eccentricity [ ].
  • i: Inclination [rad].
  • Ω: Right ascension of the ascending node [rad].
  • ω: Argument of perigee [rad].
  • f: True anomaly [rad].
source
SatelliteToolbox.NRLMSISE00_FlagsType
NRLMSISE00_Flags

Flags to configure NRLMSISE-00.

Fields

  • output_m_kg
  • F107_Mean
  • time_independent
  • sym_annual
  • sym_semiannual
  • asym_annual
  • asyn_semiannual
  • diurnal
  • semidiurnal
  • daily_ap
  • all_ut_long_effects
  • longitudinal
  • ut_mixed_ut_long
  • mixed_ap_ut_long
  • terdiurnal
  • departures_from_eq
  • all_tinf_var
  • all_tlb_var
  • all_tn1_var
  • all_s_var
  • all_tn2_var
  • all_nlb_var
  • all_tn3_var
  • turbo_scale_height
  • use_ap_array
source
SatelliteToolbox.NRLMSISE00_OutputType
NRLMSISE00_Output

Output structure for NRLMSISE00 model.

Fields

  • den_N: Nitrogen number density [U].
  • den_N2: N₂ number density [U].
  • den_O: Oxygen number density [U].
  • den_aO: Anomalous Oxygen number density [U].
  • den_O2: O₂ number density [U].
  • den_H: Hydrogen number density [U].
  • den_He: Helium number density [U].
  • den_Ar: Argon number density [U].
  • den_Total: Total mass density [T/U] (this value has different meanings for routines gtd7 and gtd7d).
  • T_exo: Exospheric temperature [K].
  • T_alt: Temperature at the selected altitude [K].
  • flags: Flags used to compute NRLMSISE-00 model.

Notice that:

  • If flags.output_m_kg is false, then [U] is [cm⁻³] and [T] is [g/cm⁻³].
  • If flags.output_m_kg is true, then [U] is [m⁻³] and [T] is [kg/m⁻³].

Remarks

Anomalous oxygen is defined as hot atomic oxygen or ionized oxygen that can become appreciable at high altitudes (> 500 km) for some ranges of inputs, thereby affection drag on satellites and debris. We group these species under the term Anomalous Oxygen, since their individual variations are not presently separable with the drag data used to define this model component.

source
SatelliteToolbox.OrbitPropagatorType
OrbitPropagator{T}

Abstract type of the orbit propagator. Every propagator structure must be a subtype of this type and must implement the following API functions:

propagate!(orbp, t::Number)
propagate!(orbp, t::AbstractVector)
propagate_to_epoch!(orbp, JD::Number)
propagate_to_epoch!(orbp, JD::AbstractVector)
step!(orbp, Δt::Number)
source
SatelliteToolbox.OrbitPropagatorJ2Type
OrbitPropagatorJ2{T} <: OrbitPropagator{T}

Structure that holds the information related to the J2 orbit propagator.

Fields

  • orb: Mean orbital elements (see Orbit).
  • j2d: Structure that stores the J2 orbit propagator data (see J2_Structure).
source
SatelliteToolbox.OrbitPropagatorJ4Type
OrbitPropagatorJ4{T} <: OrbitPropagator{T}

Structure that holds the information related to the J4 orbit propagator.

Fields

  • orb: Mean orbital elements (see Orbit).
  • j4d: Structure that stores the J4 orbit propagator data (see J4_Structure).
source
SatelliteToolbox.OrbitPropagatorSGP4Type
OrbitPropagatorSGP4{T} <: OrbitPropagator{T}

Structure that holds the information related to the SGP4 propagator.

Fields

  • orb: Mean orbital elements (see Orbit).
  • sgp4_gc: Gravitational contents of the SGP4 algorithm (see SGP4_GravCte).
  • sgp4d: Structure that stores the SGP4 data (see SGP4_Structure).
source
SatelliteToolbox.OrbitPropagatorTwoBodyType
OrbitPropagatorTwoBody{T} <: OrbitPropagator{T}

Structure that holds the information related to the Two Body orbit propagator.

Fields

  • orb: Mean orbital elements (see Orbit).
  • tbd: Structure that stores the Two Body orbit propagator data (see TwoBody_Structure).
source
SatelliteToolbox.OrbitStateVectorType
OrbitStateVector{T}

Store the state vector representation of an orbit.

Fields

  • t: Epoch [Julian Day].
  • r: Position vector [m].
  • v: Velocity vector [m/s].
  • a: Acceleration vector [m/s²].
source
SatelliteToolbox._SOLFSMY_StructureType
_SOLFSMY_Structure

Structure to store the interpolations of the data in SOLFSMY.TXT file.

Fields

  • F10: 10.7-cm solar flux [10⁻²² W/(m² Hz)].
  • F81a: 10.7-cm averaged solar flux, 81-day centered on input time.
  • S10: EUV index.
  • S81a: EUV 81-day averaged centered index.
  • M10: MG2 index scaled to F10.
  • M81a: MG2 81-day averaged centered index.
  • Y81a: Solar X-ray & Lya 81-day averaged centered index.
  • Y81a: Solar X-ray & Lya 81-day averaged centered index.
source
SatelliteToolbox.DatetoJDMethod
DatetoJD(Y::Integer, M::Integer, D::Integer, h::Integer, m::Integer, s::Number)

Convert a date represented using the Gregorian Calendar (Year = y, Month = M (1-12), Day = D, Hour = h (0-24), minute = m, and second = s) to Julian Day.

Remarks

The algorithm was obtained from [2] (Accessed on 2018-04-11).

source
SatelliteToolbox.ECEFtoGeodeticMethod
ECEFtoGeodetic(r_e::AbstractVector)

Convert the vector r_e [m] represented in the Earth-Centered, Earth-Fixed (ECEF) reference frame into Geodetic coordinates (WGS-84).

Returns

  • Latitude [rad].
  • Longitude [rad].
  • Altitude [m].

Remarks

Based on algorithm in [3].

source
SatelliteToolbox.E_to_MMethod
E_to_M(e::Number, E::Number)

Compute the mean anomaly (0,2π) [rad] given the eccentricity e and the eccentric anomaly E [rad].

source
SatelliteToolbox.E_to_fMethod
E_to_f(e::Number, E::Number)

Compute the true anomaly (0,2π) [rad] given the eccentricity e and the eccentric anomaly E [rad].

source
SatelliteToolbox.GeodetictoECEFMethod
GeodetictoECEF(lat::Number, lon::Number, h::Number)

Convert the latitude lat [rad], longitude lon [rad], and altitude h [m] (WGS-84) into a vector represented on the Earth-Centered, Earth-Fixed (ECEF) reference frame.

Remarks

Based on algorithm in [3].

source
SatelliteToolbox.GeodetictoGeocentricMethod
GeodetictoGeocentric(ϕ_gd::Number, h::Number)

Compute the geocentric latitude and radius from the geodetic latitude ϕ_gd (-π/2,π/2) [rad] and height above the reference ellipsoid h [m] (WGS-84). Notice that the longitude is the same in both geocentric and geodetic coordinates.

Returns

  • Geocentric latitude [rad].
  • Radius from the center of the Earth [m].

Remarks

Based on algorithm in [4, p. 3].

source
SatelliteToolbox.J2000toGMSTMethod
J2000toGMST(J2000_UT1::Number)

Compute the Greenwich Mean Sideral Time (GMST) [rad] given the instant J2000_UT1 in J2000.0 reference [UT1].

Remarks

Based on algorithm in 2, accessed at 2015-12-01.

source
SatelliteToolbox.JD_TTtoUTCMethod
JD_TTtoUTC(JD_TT::Number, ΔAT::Number = 37)

Convert the Julian Day in TT JD_TT (Terrestrial Time) to the Julian Day in UTC (Terrestrial Time) using the accumulated difference ΔAT between UTC and the International Atomic Time (TAI). If no value is provided, then the leap seconds will be obtained from the table ΔAT_Data. Notice that, in this case, if a date previous to 1973 is provided, then a fixed value of 10 will be used, leading to wrong computations.

source
SatelliteToolbox.JD_UT1toUTCMethod
JD_UT1toUTC(JD_UT1::Number, ΔUT1::Number)

Convert the Julian Day in UT1 JD_UT1 to the Julian Day in UTC using the accumulated difference ΔUT1, which is provided by IERS EOP Data.

source
SatelliteToolbox.JD_UT1toUTCMethod
JD_UTCtoUT1(JD_UTC::Number, eop::Union{EOPData_IAU1980,EOPData_IAU2000A})

Convert the Julian Day in UT1 JD_UT1 to the Julian Day in UTC using the accumulated difference given by the EOP Data eop (see get_iers_eop). Notice that the accumulated difference will be interpolated.

source
SatelliteToolbox.JD_UTCtoTTMethod
JD_UTCtoTT(JD_UTC::Number [, ΔAT::Number])

Convert the Julian Day in UTC JD_UTC to the Julian Day in TT (Terrestrial Time) using the accumulated difference ΔAT between UTC and the International Atomic Time (TAI). If no value is provided, then the leap seconds will be obtained from the table ΔAT_Data. Notice that, in this case, if a date previous to 1973 is provided, then a fixed value of 10 will be used, leading to wrong computations.

source
SatelliteToolbox.JD_UTCtoUT1Method
JD_UTCtoUT1(JD_UTC::Number, ΔUT1::Number)

Convert the Julian Day in UTC JD_UTC to the Julian Day in UT1 using the accumulated difference ΔUT1, which is provided by IERS EOP Data.

source
SatelliteToolbox.JD_UTCtoUT1Method
JD_UTCtoUT1(JD_UTC::Number, eop::Union{EOPData_IAU1980,EOPData_IAU2000A})

Convert the Julian Day in UTC JD_UTC to the Julian Day in UT1 using the accumulated difference given by the EOP Data eop (see get_iers_eop). Notice that the accumulated difference will be interpolated.

source
SatelliteToolbox.JDtoDateMethod
JDtoDate([T,] JD::Number)

Convert a date represented in Julian Day JD to Gregorian Calendar. The optional parameter T defines the return type. If T is omitted, then it defaults to Int.

Returns

If T is omitted or Int, then a tuple with the following data will be returned:

  • Year.
  • Month (1 => January, 2 => February, ...).
  • Day.
  • Hour (0 - 24).
  • Minute (0 - 59).
  • Second (0 - 59).

Notice that if T is Int, then the seconds field will be Integer. Otherwise, it will be floating point.

If T is Date, then it will return the Julia structure Date. Notice that the hours, minutes, and seconds will be neglected because the structure Date does not handle them.

If T is DateTime, then it will return the Julia structure DateTime.

Remarks

The algorithm was obtained from [2] (Accessed on 2018-04-11). In [2], there is the following warning:

Note: This method will not give dates accurately on the Gregorian Proleptic Calendar, i.e., the calendar you get by extending the Gregorian calendar backwards to years earlier than 1582. using the Gregorian leap year rules. In particular, the method fails if Y<400.

source
SatelliteToolbox.JDtoGMSTMethod
JDtoGMST(JD_UT1::Number)

Compute the Greenwich Mean Sideral Time (GMST) [rad] for the Julian Day JD_UT1 [UT1].

Remarks

Based on algorithm in [1, pp. 188].

source
SatelliteToolbox.M_to_EFunction
M_to_E(e::Number, M::Number, tol::Number = 1e-10)

Compute the eccentric anomaly (0,2π) [rad] given the eccentricity e and the mean anomaly M [rad]. This function uses the Newton-Raphson algorithm and the tolerance to accept the solution is tol.

source
SatelliteToolbox.M_to_fFunction
M_to_f(e::Number, M::Number, tol::Number = 1e-10)

Compute the true anomaly (0,2π) [rad] given the eccentricity e and the mean anomaly M [rad]. This function uses the Newton-Raphson algorithm and the tolerance to accept the solution is tol.

source
SatelliteToolbox._ccorMethod
_ccor(alt::T, r::T, h1::T, zh::T) where T<:Number

Chemistry / Dissociation correction for MSIS models.

Args

  • alt: Altitude.
  • r: Target ratio.
  • h1: Transition scale length.
  • zh: Altitude of 1/2 r.

Returns

The chemistry / dissociation correction.

source
SatelliteToolbox._ccor2Method
_ccor2(alt::T, r::T, h1::T, zh::T, h2::T) where T<:Number

Chemistry / Dissociation correction for MSIS models.

Args

  • alt: Altitude.
  • r: Target ration.
  • h1: Transition scale length.
  • zh: Altitude of 1/2 r.
  • h2: Transition scale length 2.

Returns

The chemistry / dissociation correction.

source
SatelliteToolbox._densmMethod
_densm(re::T, gsurf::T, alt::T, d0::T, xm::T, tz::T, zn3::StaticVector{N3,T}, tn3::AbstractVector{T}, tgn3::AbstractVector{T}, zn2::StaticVector{N2,T}, tn2::AbstractVector{T}, tgn2::AbstractVector{T}) where {T<:Number,N2,N3}

Compute the temperature and density profiles for lower atmosphere.

Returns

  • The density.
  • The temperature.
source
SatelliteToolbox._densuMethod
_densu(re::T, gsurf::T, alt::T, dlb::T, tinf::T, tlb::T, xm::T, alpha::T, zlb::T, s2::T, zn1::StaticVector{N,T}, tn1::AbstractVector{T}, tgn1::AbstractVector{T}) where {T<:Number,N}

Compute the temperature and density profiles for MSIS models.

This algorithm uses new lower thermo polynomial.

Returns

  • The density.
  • The temperature.
source
SatelliteToolbox._dnetMethod
_dnet(dd::T, dm::T, zhm::T, xmm::T, xm::T) where T<:Number

Turbopause correction for MSIS models.

Args

  • dd: Diffusive density.
  • dm: Full mixed density.
  • zhm: Transition scale length.
  • xmm: Full mixed molecular weight.
  • xm: Species molecular weight.

Returns

The combined density.

source
SatelliteToolbox._get_degree_and_orderMethod
_get_degree_and_order(dP, P, n_max, m_max)

Return the maximum degree and order to compute the Legendre associated functions given the matrices dP, P, and the configuration values n_max and m_max.

source
SatelliteToolbox._get_degree_and_orderMethod
_get_degree_and_order(P, n_max, m_max)

Return the maximum degree and order to compute the Legendre associated functions given the matrix P and the configuration values n_max and m_max.

source
SatelliteToolbox._glob7sMethod
_glob7s(p::AbstractVector{T}, nrlmsise00d::NRLMSISE00_Structure{T}) where T<:Number

Version of Globe for lower atmosphere (1999-10-26).

Args

  • p: Vector with the coefficients.
  • nrlmsise00d: NRLMSISE-00 structure (see NRLMSISE00_Structure).

Returns

The temperature (?).

source
SatelliteToolbox._globe7!Method
_globe7!(p::AbstractVector{T}, nrlmsise00d::NRLMSISE00_Structure{T}) where T<:Number

Compute G(L) function.

Notice that the parameters apt and apdf of structure nrlmsise00d are modified.

Args

  • p: Vector with the coefficients.
  • nrlmsise00d: NRLMSISE-00 structure (see NRLMSISE00_Structure).

Returns

The temperature (?).

source
SatelliteToolbox._init_dtcfileMethod
_init_dctfile(;force_download = false, local_path = nothing)

Initialize the data in the file DTCFILE.TXT by creating _dtcfile_data. The initialization process is composed of:

  1. Download the file, if it is necessary;
  2. Parse the file;
  3. Create the interpolations and the structures.

If the keyword force_download is true, then the file will always be downloaded.

The user can also specify a location for the file using the keyword local_path. If it is nothing, which is the default, then the file will be downloaded.

source
SatelliteToolbox._init_fluxtableMethod
_init_fluxtable(;force_download = false, local_path = nothing)

Initialize the data in the file fluxtable.txt by creating _fluxtable_data. The initialization process is composed of:

  1. Download the file, if it is necessary;
  2. Parse the file;
  3. Create the interpolations and the structures.

If the keyword force_download is true, then the file will always be downloaded.

The user can also specify a location for the file using the keyword local_path. If it is nothing, which is the default, then the file will be downloaded.

source
SatelliteToolbox._init_solfsmyMethod
_init_solfsmy(;force_download = false, local_path = nothing)

Initialize the data in the file SOLFSMY.TXT by creating _solfsmy_data. The initialization process is composed of:

  1. Download the file, if it is necessary;
  2. Parse the file;
  3. Create the interpolations and the structures.

If the keyword force_download is true, then the file will always be downloaded.

The user can also specify a location for the file using the keyword local_path. If it is nothing, which is the default, then the file will be downloaded.

source
SatelliteToolbox._init_wdcfilesMethod
_init_wdcfiles(;force_download = false, local_dir = nothing, wdcfiles_oldest_year = year(now())-3)

Initialize the data in the WDC files by creating _wdcfiles_data. The initialization process is composed of:

  1. Download the files, if it is necessary;
  2. Parse the files;
  3. Create the interpolations and the structures.

If the keyword force_download is true, then the files will always be downloaded.

The user can also specify a location for the directory with the WDC files using the keyword local_dir. If it is nothing, which is the default, then the file will be downloaded.

The user can select what is the oldest year in which the data will be downloaded by the keyword wdcfiles_oldest_year. By default, it will download the data from 3 previous years.

The user can select what is the newest year in which the data will be downloaded by the keyword wdcfiles_newest_year. It it is nothing, which is the default, then it is set to the current year.

source
SatelliteToolbox._jb2008_TMethod
_jb2008_T(z::R, Tx::R, T∞::R) where R<:Number

Compute the temperature [K] at height z [km] given the temperature Tx [K] at the inflection point, and the exospheric temperature T∞ [K] according to the theory of the model Jacchia 1971 [3].

The inflection point is considered to by z = 125 km.

source
SatelliteToolbox._jb2008_highaltitudeMethod
_jb2008_highaltitude(h::Number, F10ₐ::Number)

Compute the high altitude exospheric density correction factor in altitude h [km] and the averaged 10.7-cm solar flux (81-day centered on input time) [10⁻²² W/(M² Hz)].

This function uses the model in Section 6.2 of [2].

source
SatelliteToolbox._jb2008_intMethod
_jb2008_int(z₀::Number, z₁::Number, R::Number, Tx::Number, T∞::Number, δf::Function)

Compute the integral of the function δf between z₀ and z₁ using the Newton-Cotes 4th degree method. R is a number that defines the step size, Tx is the temperature at the inflection point, and T∞ is the exospheric temperature.

The signature of the function δf is:

δf(z, Tx, T∞)

and it must be _jb2008_δf1 or _jb2008_δf2.

This function returns a tuple containing the integral and last value of z used in the numerical algorithm.

source
SatelliteToolbox._jb2008_semiannualMethod
_jb2008_semiannual(doy::Number, h::Number, F10ₐ::Number, S10ₐ::Number, M10ₐ::Number)

Compute the semiannual variation of the density considering the JB2008 model [1].

Args

  • doy: Day of the year + fraction of the day.
  • h: Height [km].
  • F10ₐ: Averaged 10.7-cm flux (81-day centered on input-time) [10⁻²² W/(M² Hz)].
  • S10ₐ: EUV 81-day averaged centered index.
  • M10ₐ: MG2 81-day averaged centered index.

Returns

  • Semiannual F(z) heigh function.
  • Semiannual G(t) yearly periodic function.
  • Semiannual variation of the density Δsalog₁₀ρ.
source
SatelliteToolbox._jb2008_ΔTcMethod
_jb2008_ΔTc(F10::Number, lst::Number, glat::Number, h::Number)

Compute the correction in the Tc for Jacchia-Bowman model.

This correction is mention in [2]. However, the equations do not seem to match those in the source-code. The ones implemented here are exactly the same as in the source-code.

Args

  • F10: F10.7 flux.
  • lst: Local solar time (0 - 24) [hr].
  • glat: Geocentric latitude [rad].
  • h: Altitude [km].

Returns

The correction ΔTc [K].

source
SatelliteToolbox._jr1971_MMethod
_jr1971_M(z::R) where R

Compute the mean molecular mass at altitude z [km] using the empirical profile in eq. 1 [3,4].

source
SatelliteToolbox._jr1971_TMethod
_jr1971_T(z::R, Tx::R, T∞::R) where R<:Number

Compute the temperature [K] at height z [km] given the temperature Tx [K] at the inflection point, and the exospheric temperature T∞ [K] according to the theory of the model Jacchia-Roberts 1971 [1,3,4].

The inflection point is considered to by z = 125 km.

source
SatelliteToolbox._jr1971_rootsMethod
_jr1971_roots(p::Polynomial{R}) where R

Compute the roots of the polynomial p necessary to compute the density below 125 km. It returns the value r₁, r₂, x, and y.

source
SatelliteToolbox._parse_dtcfileMethod
_parse_dtcfile(path::AbstractString)

Parse the DTCFILE.TXT file in path and return an instance of the structure _DTCFILE_Structure with the initialized interpolations.

The format of the file DTCFILE.TXT must be:

DTC YYYY DOY DTC_0h DTC_1h DTC_2h ... DTC_22h DTC_23h

in which DOY is the day of the year and DTC_Xh is the ΔTc at hour X.

source
SatelliteToolbox._parse_fluxtableMethod
_parse_fluxtable(path::AbstractString)

Parse the fluxtable.txt file in path and return an instance of the structure _fluxtable_Structure with the initialize interpolations.

source
SatelliteToolbox._parse_solfsmyMethod
_parse_solfsmy(path::AbstractString)

Parse the SOLFSMY.TXT file in path and retur an instance of the structure _SOLFSMY_Structure with the initialized interpolations.

The format of the file SOLFSMY.TXT must be:

YYYY DDD   JulianDay  F10   F81c  S10   S81c  M10   M81c  Y10   Y81c  Ssrc
source
SatelliteToolbox._parse_wdcfilesMethod
_parse_wdcfiles(filepaths::Vector{String}, years::Vector{Int})

Parse the WDC files with paths in filepaths related to the years in years.

Notice that the files must be sorted by the year!

source
SatelliteToolbox._prepare_wdc_remote_filesMethod
_prepare_wdc_remote_files(oldest_year::Number, newest_year::Number)

Configure all the WDC remote files between newest_year and oldest_year. Notice that previous years will never be updated whereas the current year will be updated daily.

If oldest_year is greater than current year, then only the files from the current year will be downloaded.

If newest_year is smaller than oldest_year, then only the files from the oldest_year will be downloaded.

This function modifies the global variable _wdcfiles.

source
SatelliteToolbox._splineMethod
_spline(x::StaticVector{N,T}, y::StaticVector{N,T}, yp1::T, ypn::T) where {T<:Number,N}

Compute the 2nd derivatives of cubic spline interpolation function tabulated by x and y given the 2nd derivatives values at x[1] (yp1) and at x[N] (ypn).

This function was adapted from Numerical Recipes.

Args

  • x: X components of the tabulated function in ascending order.
  • y: Y components of the tabulated function evaluated at x.
  • yp1: 2nd derivative value at x[1].
  • ypn: 2nd derivative value at x[N].

Returns

The 2nd derivative of cubic spline interpolation function evaluated at x.

Remarks

Values higher than 1e30 in the 2nd derivatives at the borders (yp1 and ypn) are interpreted as 0.

source
SatelliteToolbox._spliniMethod
_splini(xa::StaticVector{N,T}, ya::StaticVector{N,T}, y2a::StaticVector{N,T}, x::T) where {T<:Number,N}

Compute the integral of the cubic spline function from xa[1] to x.

Args

  • xa: X components of the tabulated function in ascending order.
  • ya: Y components of the tabulated function evaluated at xa.
  • y2a: Second derivatives.
  • x: Abscissa endpoint for integration.

Returns

The integral of cubic spline function from xa[1] to x.

source
SatelliteToolbox._splintMethod
_splint(xa::StaticVector{N,T}, ya::StaticVector{N,T}, y2a::StaticVector{N,T}, x::T) where {T<:Number,N}

Compute the cubic spline interpolation value at x.

This function was adapted from Numerical Recipes.

Args

  • xa: X components of the tabulated function in ascending order.
  • ya: Y components of the tabulated function evaluated at xa.
  • y2a: Second derivatives.
  • x: Abscissa endpoint for interpolation.

Returns

The cubic spline interpolation value at x.

source
SatelliteToolbox.adjacent_track_angle_grssMethod
adjacent_track_angle_grss(h::Number, T::Number, i::Number, To::Int, lat::Number)

Compute the angle between two adjacent ground tracks [rad] in a given latitude lat [rad] measured from the satellite position for a ground repeating, Sun-synchronous orbit with altitude in the Equator h [m], period T [s], inclination i [rad], and orbit cycle To [days].

Remarks

The functions does not check if the orbit is a GRSS orbit.

source
SatelliteToolbox.adjacent_track_angle_grssMethod
adjacent_track_angle_grss(h::Number, a::Number, e::Number, i::Number, To::Int, lat::Number)

Compute the angle between two adjacent ground tracks [rad] in a given latitude lat [rad] measured from the satellite position for a ground repeating, Sun-synchronous orbit with altitude in the Equator h [m], semi-major axis a [m], eccentricity e, inclination i [rad], and orbit cycle To [days].

Remarks

The functions does not check if the orbit is a GRSS orbit.

source
SatelliteToolbox.adjacent_track_distance_grssMethod
adjacent_track_distance_grss(T::Number, i::Number, To::Int, lat::Number)

Compute the distance between adjacent ground tracks [m] at a given latitude lat [rad] for a ground repeating, Sun-synchronous orbit with period T [s], inclination i [rad], and orbit cycle To [days].

Remarks

The functions does not check if the orbit is a GRSS orbit.

source
SatelliteToolbox.adjacent_track_distance_grssMethod
adjacent_track_distance_grss(a::Number, e::Number, i::Number, To::Int, lat::Number)

Compute the distance between adjacent ground tracks [m] at a given latitude lat [rad] for a ground repeating, Sun-synchronous orbit with semi-major axis a [m], eccentricity e, inclination i [rad], and orbit cycle To [days].

Remarks

The functions does not check if the orbit is a GRSS orbit.

source
SatelliteToolbox.angvelFunction
angvel(a::Number, e::Number, i::Number, pert::Symbol = :J2)
angvel(orb::Orbit, pert::Symbol = :J2)

Compute the angular velocity [rad/s] of an object in an orbit with semi-major axis a [m], eccentricity e, and inclination i [rad], using the perturbation terms specified by the symbol pert. The orbit can also be specified by orb (see Orbit).

pert can be:

  • :J0: Consider a Keplerian orbit.
  • :J2: Consider the perturbation terms up to J2.
  • :J4: Consider the perturbation terms J2, J4, and J2².

If pert is omitted, then it defaults to :J2.

source
SatelliteToolbox.angvel_to_aFunction
angvel_to_a(n::Number, e::Number, i::Number, pert::Symbol = :J2; μ::Number = m0, max_iter::Int = 20, tol::Number = 1e-10)

Compute the semi-major axis that will provide an angular velocity n [rad/s] in an orbit with eccentricity e and inclination i [rad], using the perturbation terms specified by the symbol pert.

Notice that the angular velocity n is related to the nodal period, i.e. the time between two consecutive passages by the ascending node.

pert can be:

  • :J0: Consider a Keplerian orbit.
  • :J2: Consider the perturbation terms up to J2.
  • :J4: Consider the perturbation terms J2, J4, and J2².

If pert is omitted, then it defaults to :J2.

Keyword

  • μ: Standard gravitational parameter for Earth [m^3/s^2]. (Default = m0)
  • max_iter: Maximum number of iterations allowed in the Newton-Raphson algorithm. (Default = 20)
  • tol: Tolerance to stop the Newton-Raphson algorithm. (Default = 1e-10)
source
SatelliteToolbox.beta_angleFunction
beta_angle(JD₀::Number, a::Number, e::Number, i::Number, RAAN::Number, Δt::Integer, pert::Symbol = :J2)

Compute the beta angle of an orbit with semi-major axis a [m], eccentricity e, inclination i [rad], and initial right ascension of the ascending node RAAN [rad]. The orbit epoch, which is also the day in which the analysis will begin, is JD₀ [Julian Day]. The analysis will be performed for each day during Δt days.

The argument pert can be used to select the perturbation terms that must be used when propagating the right ascencion of the ascending node. The possible values are:

  • :J0: Consider a Keplerian orbit.
  • :J2: Consider the perturbation terms up to J2.
  • :J4: Consider the perturbation terms J2, J4, and J2².

If pert is omitted, then it defaults to :J2.

Returns

An array with two columns. The first one contains the days of the analysis and the second one contains the beta angle [rad] for each day.

source
SatelliteToolbox.change_oe_frameMethod
change_oe_frame(a::Number, e::Number, i::Number, Ω::Number, ω::Number, f::Number, conv_args...)
change_oe_frame(oe::Orbit, conv_args...)

Change the reference frame of orbit elements. The orbit elements can be specified by a, e, i, Ω, ω, and f, or the structure oe (see Orbit). In the latter, the return value type will match the type of oe.

The conversion arguments conv_args are the same arguments that one should pass to the function rECItoECI to convert between the desired frames. For more information, see the documentation of the function rECItoECI.

Args

  • a: Semi-major axis [m].

  • e: Excentricity.

  • i: Inclination [rad].

  • Ω: Right-ascension of the ascending node [rad].

  • ω: Argument of perigee [rad].

  • f: True anomaly [rad].

  • conv_args...: Conversion arguments, which are the same arguments that one would pass to the function rECItoECI to convert between the desired frames.

  • oe: An instance of the structure Orbit with the orbit elements that will be converted [SI units].

Returns

Using the first signature, this function returns an instance of KeplerianElements. If the second signature is used, then the function return an element with the same type of the input.

Examples

julia> eop = get_iers_eop(:IAU1980);

julia> teme_epoch = DatetoJD(2016,6,1,11,0,0);

julia> tod_epoch  = DatetoJD(2016,1,1,0,0,0);

julia> k_teme     = KeplerianElements(teme_epoch,
                                      7130.982e3,
                                      0.001111,
                                      98.405*pi/180,
                                      227.336*pi/180,
                                      90*pi/180,
                                      320*pi/180)
KeplerianElements{Float64}:
           Epoch :    2.45754e6 (2016-06-01T11:00:00)
 Semi-major axis : 7130.98     km
    Eccentricity :    0.001111
     Inclination :   98.405    °
            RAAN :  227.336    °
 Arg. of Perigee :   90.0      °
    True Anomaly :  320.0      °

julia> k_j2000 = change_oe_frame(k_teme, TEME(), J2000(), teme_epoch, eop)
KeplerianElements{Float64}:
           Epoch :    2.45754e6 (2016-06-01T11:00:00)
 Semi-major axis : 7130.98     km
    Eccentricity :    0.001111
     Inclination :   98.3365   °
            RAAN :  227.134    °
 Arg. of Perigee :   90.0604   °
    True Anomaly :  320.0      °

julia> k_tod = change_oe_frame(k_teme, TEME(), teme_epoch, TOD(), tod_epoch, eop)
KeplerianElements{Float64}:
           Epoch :    2.45754e6 (2016-06-01T11:00:00)
 Semi-major axis : 7130.98     km
    Eccentricity :    0.001111
     Inclination :   98.4037   °
            RAAN :  227.331    °
 Arg. of Perigee :   90.0014   °
    True Anomaly :  320.0      °
source
SatelliteToolbox.cio_iau2006Method
cio_iau2006(JD_TT::Number)

Compute the coordinates X and Y of the Celestial Intermediate Pole (CIP) with respect to the Geocentric Celestial Reference Frame (GCRF), and the CIO locator s. The algorithm is based on the IAU-2006 theory.

The CIO locator s provides the position of the CIO on the Equator of the CIP corresponding to the kinematical definition of the non-rotation origin in the GCRS when the CIP is moving with respect to the GCRS between the reference epoch and the epoch due to precession and nutation [1, p. 214].

Returns

  • The coordinate X of the CIP w.r.t. the GCRF.
  • The coordinate Y of the CIP w.r.t. the GCRF.
  • The CIO locator s.
source
SatelliteToolbox.compute_RAAN_ltMethod
compute_RAAN_lt(JD::Number, asc_node_lt::Number)

Compute the RAAN (0,2π) [rad] so that the orbit plane local time is asc_node_lt [hour] at the Julian day JD.

source
SatelliteToolbox.compute_UMethod
compute_U(gm_coefs::GravityModel_Coefs{T}, r::AbstractVector, n_max::Number = -1, m_max::Number = -1) where T<:Number

Compute the gravitational potential [J/kg] at r (ITRF) [m] using the coefficients gm_coefs (see GravityModel_Coefs). The maximum degree that will be used while computing the spherical harmonics will be n_max and the maximum order is m_max.

If n_max is negative, then the maximum available degree will be used. If n_max is omitted, then it defaults to 0.

If m_max is negative or if it is greater than n_max, then it will be set to n_max. If m_max is omitted, then it defaults to 0.

Info

By convention, the result with n_max 0 and 1 will be the same.

source
SatelliteToolbox.compute_dUMethod
compute_dU(gm_coefs::GravityModel_Coefs{T}, r::AbstractVector, n_max::Number = -1, m_max::Number = -1) where T<:Number

Compute the derivatives w.r.t. the spherical coordinates of the gravitational field (∂U/∂r, ∂U/∂ϕ, ∂U/∂λ) defined by the coefficients gm_coefs (see GravityModel_Coefs) at the position r [m] in ITRF. The maximum degree that will be used while computing the spherical harmonics will be n_max and the maximum order is m_max.

If n_max is negative, then the maximum available degree will be used. If n_max is omitted, then it defaults to 0.

If m_max is negative or if it is greater than n_max, then it will be set to n_max. If m_max is omitted, then it defaults to 0.

Info

By convention, the result with n_max 0 and 1 will be the same.

Returns

  • The derivative of the gravitational field w.r.t. the radius (∂U/∂r).
  • The derivative of the gravitational field w.r.t. the latitude (∂U/∂ϕ).
  • The derivative of the gravitational field w.r.t. the longitude (∂U/∂λ).

Remarks

In this case, ϕ is the geocentric latitude and λ is the geocentric longitude.

source
SatelliteToolbox.compute_gMethod
compute_g(gm_coefs::GravityModel_Coefs{T}, r::AbstractVector, n_max::Number = -1, m_max::Number = -1) where T<:Number

Compute the gravitational acceleration (ITRF) [m/s²] at position r [m] (ITRF) using the coefficients gm_coefs (see GravityModel_Coefs). The maximum degree that will be used while computing the spherical harmonics will be n_max and the maximum order it m_max.

If n_max is negative, then the maximum available degree will be used. If n_max is omitted, then it defaults to 0.

If m_max is negative or if it is greater than n_max, then it will be set to n_max. If m_max is omitted, then it defaults to 0.

Info

By convention, the result with n_max 0 and 1 will be the same.

Remarks

Notice that this function computes the gravitational acceleration. Hence, the acceleration due to Earth rotation rate is not included.

source
SatelliteToolbox.compute_ss_orbit_by_ang_velMethod
compute_ss_orbit_by_ang_vel(n::Number, e::Number)

Compute the Sun-synchronous orbit given the angular velocity n [rad/s] and the eccentricity e.

Returns

  • The semi-major axis [m].
  • The inclination [rad].
  • The residues of the two functions.
  • A boolean variable that indicates if the numerical algorithm converged.
source
SatelliteToolbox.compute_ss_orbit_by_num_rev_per_dayMethod
compute_ss_orbit_by_num_rev_per_day(numRevPD::Number, e::Number)

Compute the Sun-synchronous orbit given the number of revolutions per day numRevPD and the eccentricity e.

Returns

  • The semi-major axis [m].
  • The inclination [rad].
  • The residues of the two functions.
  • A boolean variable that indicates if the numerical algorithm converged.
source
SatelliteToolbox.conf_nrlmsise00Function
conf_nrlmsise00(year::Int, doy::Int, sec::Number, alt::Number, g_lat::Number, g_long::Number, lst::Number, f107A::Number, f107::Number, ap::[Number, AbstractVector], flags::NRLMSISE00_Flags = NRLMSISE00_Flags())

Create the structure with the proper configuration to call the NRLMSISE-00 model.

Notice that the input variables have the same units of the original model.

Args

  • year: Year (currently ignored).
  • doy: Day of year.
  • sec: Seconds in day [UT].
  • alt: Altitude [km].
  • g_lat: Geodetic latitude [deg].
  • g_long: Geodetic longitude [deg].
  • lst: Local apparent solar time (hours).
  • f107A: 81 day average of F10.7 flux (centered on day of year doy).
  • f107: Daily F10.7 flux for previous day.
  • ap: Magnetic index (daily) if it is a number. If it is an array, then see Remarks.
  • flags: (OPTIONAL) An instance of the structure NRLMSISE00_Flags with the configuration flags for NRLMSISE00. If omitted, then the default configurations will be used.

Returns

An instance of the structure NRLMSISE00_Structure.

Remarks

If ap is a Vector, then it must be a vector with 7 dimensions as described below:

IndexDescription
1Daily AP.
23 hour AP index for current time.
33 hour AP index for 3 hours before current time.
43 hour AP index for 6 hours before current time.
53 hour AP index for 9 hours before current time.
6Average of eight 3 hour AP indices from 12 to 33 hours prior to current time.
7Average of eight 3 hour AP indices from 36 to 57 hours prior to current time.

Notes on input variables

UT, Local Time, and Longitude are used independently in the model and are not of equal importance for every situation. For the most physically realistic calculation these three variables should be consistent (lst=sec/3600 + g_long/15). The Equation of Time departures from the above formula for apparent local time can be included if available but are of minor importance.

f107 and f107A values used to generate the model correspond to the 10.7 cm radio flux at the actual distance of the Earth from the Sun rather than the radio flux at 1 AU. The following site provides both classes of values:

ftp://ftp.ngdc.noaa.gov/STP/SOLAR_DATA/SOLAR_RADIO/FLUX/

f107, f107A, and ap effects are neither large nor well established below 80 km and these parameters should be set to 150, 150, and 4 respectively.

source
SatelliteToolbox.dEps_dPsiMethod
dEps_dPsi(eop_iau2000a::EOPData_IAU2000A, JD::Number)

Returns the celestial pole offsets in obliquity (δϵ2000) and longitude (δΨ2000) [arcsec]. This function obtains those values by converting the celestial pole offsets with respect to the GCRS (dX and dY). These values are necessary in the equinox-based IAU-2006 theory.

The algorithm was obtained from [2, eq. 5.25] and [3, DPSIDEPS2000_DXDY2000].

source
SatelliteToolbox.dargpFunction
dargp(a::Number, e::Number, i::Number, pert::Symbol = :J2)
dargp(orb::Orbit, pert::Symbol = :J2)

Compute the time-derivative of the argument of perigee [rad/s] of an orbit with semi-major axis a [m], eccentricity e, and inclination i [rad], using the perturbation terms specified by the symbol pert. The orbit can also be specified by orb (see Orbit).

pert can be:

  • :J0: Consider a Keplerian orbit.
  • :J2: Consider the perturbation terms up to J2.
  • :J4: Consider the perturbation terms J2, J4, and J2².

If pert is omitted, then it defaults to :J2.

source
SatelliteToolbox.dlegendreFunction
dlegendre([N,] ϕ::Number, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false)

Compute the first-order derivative of the associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

The optional parameter N can be used to select the normalization. The following values are valid:

  • Val(:full): Compute the fully normalized associated Legendre function (see legendre_fully_normalized).
  • Val(:schmidt): Compute the Schmidt quasi-normalized associated Legendre function (see legendre_schmidt_quasi_normalized).
  • Val(:conv): Compute the conventional associated Legendre function (see dlegendre_conventional!).

If N is omitted, then the full normalization will be used (Val(:full)).

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

  • A matrix with the first-order derivative of the Legendre associated functions

P_n,m[cos(ϕ)].

  • A matrix with the Legendre associated functions P_n,m[cos(ϕ)].
source
SatelliteToolbox.dlegendre!Function
dlegendre!([N,] dP::AbstractMatrix, ϕ::Number, P::AbstractMatrix, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the first-order derivative of the associated Legendre function P_n,m[x] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The derivatives will be stored in the matrix dP. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix dP will be used.

This algorithm needs the matrix P with the associated Legendre function. This can be computed using the function legendre. Notice that this matrix must be computed using the same normalization (see N) as the one selected here.

The optional parameter N can be used to select the normalization. The following values are valid:

  • Val(:full): Compute the fully normalized associated Legendre function (see dlegendre_fully_normalized!).
  • Val(:schmidt): Compute the Schmidt quasi-normalized associated Legendre function (see dlegendre_schmidt_quasi_normalized!).
  • Val(:conv): Compute the conventional associated Legendre function (see dlegendre_conventional!).

If N is omitted, then the full normalization will be used (Val(:full)).

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

source
SatelliteToolbox.dlegendre_conventional!Function
dlegendre_conventional!(dP::AbstractMatrix, ϕ::Number, P::AbstractMatrix, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the first-order derivative of the conventional associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The derivatives will be stored in the matrix dP. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix dP will be used.

This algorithm needs the matrix P with the conventional associated Legendre function. This can be computed using the function legendre_conventional.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Remarks

The user is responsible to pass a matrix P with the correct values. For example, if ph_term is true, then P must also be computed with ph_term set to true.

source
SatelliteToolbox.dlegendre_conventionalMethod
dlegendre_conventional(ϕ::Number, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false)

Compute the first-order derivative of the conventional associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the first-order derivative of the Legendre associated functions P_n,m[cos(ϕ)].

source
SatelliteToolbox.dlegendre_fully_normalized!Function
dlegendre_fully_normalized!(dP::AbstractMatrix, ϕ::Number, P::AbstractMatrix, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the first-order derivative of the fully normalized associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The derivatives will be stored in the matrix dP. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix dP will be used.

This algorithm needs the matrix P with the fully normalized associated Legendre function. This can be computed using the function legendre_fully_normalized.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Remarks

The user is responsible to pass a matrix P with the correct values. For example, if ph_term is true, then P must also be computed with ph_term set to true.

source
SatelliteToolbox.dlegendre_fully_normalizedMethod
dlegendre_fully_normalized(ϕ::T, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false) where T<:AbstractFloat

Compute the first-order derivative of the Schmidt fully normalized associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the first-order derivative of the Legendre associated functions P_n,m[cos(ϕ)].

source
SatelliteToolbox.dlegendre_schmidt_quasi_normalized!Function
dlegendre_schmidt_quasi_normalized!(dP::AbstractMatrix, ϕ::Number, P::AbstractMatrix, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the first-order derivative of the Schmidt quasi-normalized associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The derivatives will be stored in the matrix dP. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix dP will be used.

This algorithm needs the matrix P with the Schmidt quasi-normalized associated Legendre function. This can be computed using the function legendre_schmidt_quasi_normalized.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Remarks

The user is responsible to pass a matrix P with the correct values. For example, if ph_term is true, then P must also be computed with ph_term set to true.

source
SatelliteToolbox.dlegendre_schmidt_quasi_normalizedMethod
dlegendre_schmidt_quasi_normalized(ϕ::T, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false) where T<:AbstractFloat

Compute the first-order derivative of the Schmidt quasi-normalized associated Legendre function P_n,m[cos(ϕ)] w.r.t. ϕ [rad]:

dP_n,m[cos(ϕ)]
--------------
      dϕ

The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the first-order derivative of the Legendre associated functions P_n,m[cos(ϕ)].

source
SatelliteToolbox.draanFunction
draan(a::Number, e::Number, i::Number, pert::Symbol = :J2)
draan(orb::Orbit, pert::Symbol = :J2)

Compute the time-derivative of the right ascension of the ascending node [rad/s] of an orbit with semi-major axis a [m], eccentricity e, and inclination i [rad], using the perturbation terms specified by the symbol pert. The orbit can also be specified by orb (see Orbit).

pert can be:

  • :J0: Consider a Keplerian orbit.
  • :J2: Consider the perturbation terms up to J2.
  • :J4: Consider the perturbation terms J2, J4, and J2².

If pert is omitted, then it defaults to :J2.

source
SatelliteToolbox.eclipse_time_summaryFunction
eclipse_time_summary(JD₀::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, Δd::Integer, relative::Bool = false, Δt₀::AbstractFloat = -1.0)

Compute the eclipse time of an orbit with semi-major axis a [m], eccentricity e, inclination i [rad], initial right ascension of the ascending node RAAN [rad], and initial argument of perigee w [rad]. The orbit epoch, which is also the day in which the analysis will begin, is JD₀ [Julian Day]. The analysis will be performed for each day during Δd days.

This function will compute the eclipse time of one orbit per day.

If the argument relative is true, then the computed times will be relative to the nodal period [%]. Otherwise, they will be computed in seconds. By default, relative = false.

The argument Δt₀ can be used to select the time step in which the orbit will be propagated. Notice that this algorithm performs a numerical search to find the beginning of each section (sunlight, penumbra, and umbra) with millisecond precision. Thus, selecting a high number for Δt₀ will make the analysis faster, but the accuracy is lost if a region time span is smalled than Δt₀. If this parameter is omitted or if it is negative, then the time step will be selected automatically to match a mean anomaly step of 5°.

All the analysis is performed using a J2 orbit propagator.

Returns

The following table:

    day | Sunlight Time | Penumbra Time | Umbra Time
   -----+---------------+---------------+------------
source
SatelliteToolbox.equation_of_timeMethod
equation_of_time(JD::Number)

Compute the difference between the Sun apparent local time and the Sun mean local time [rad], which is called Equation of Time, at the Julian Day JD. The algorithm was adapted from [1, p. 178, 277-279].

source
SatelliteToolbox.expatmosphereMethod
expatmosphere(h::Number)

Compute the atmospheric density [kg/m³] at the altitude h [m] (above the ellipsoid) using the exponential atmospheric model:

                ┌            ┐
                │    h - h₀  │
ρ(h) = ρ₀ ⋅ exp │ - ──────── │ ,
                │      H     │
                └            ┘

in which ρ₀, h₀, and H are parameters obtained from tables that depend only on h.

source
SatelliteToolbox.f_to_EMethod
f_to_E(e::Number,f::Number)

Compute the eccentric anomaly (0,2π) [rad] given the eccentricity e and the true anomaly f [rad].

source
SatelliteToolbox.f_to_MMethod
f_to_M(e::Number, f::Number)

Compute the mean anomaly (0,2π) [rad] given the eccentricity e and the true anomaly f [rad].

source
SatelliteToolbox.find_crossingMethod
find_crossing(f::Function, t₀::Number, t₁::Number, s₀, s₁; Δ = 1e-3, max = 100)

Return the crossing time tc in which the function f(t) goes from the state s₀ to the state s₁. It is assumed that f(t₀) = s₀ and f(t₁) = s₁.

If the computed interval is smalled than Δ or if the number of iterations is higher than max, then the algorithm stops.

source
SatelliteToolbox.geomag_dipoleFunction
geomag_dipole(r_e::AbstractVector, pole_lat::Number, pole_lon::Number, m::Number)

Compute the geomagnetic field [nT] using the simplified dipole model at position r_e (ECEF reference frame). This function considers that the latitude of the South magnetic pole (which lies in the North hemisphere) is pole_lat [rad] and the longitude is pole_lon [rad]. Furthermore, the dipole moment is considered to be m [A.m²].

geomag_dipole(r_e::AbstractVector, year::Number = 2019)

Compute the geomagnetic field [nT] using the simplified dipole model at position r_e (ECEF reference frame). This function uses the year year to obtain the position of the South magnetic pole (which lies in the North hemisphere) and the dipole moment. If year is omitted, then it will be considered as 2019.

Remarks

In both functions, the output vector will be represented in the ECEF reference frame.

source
SatelliteToolbox.get_ApMethod
get_Ap(JD::Number; mean::Tuple{Int} = (), daily = false)

Return the Ap index.

If mean is a tuple of two integers (hi, hf), then the average between hi and hf previous hours will be computed.

If mean is empty and daily is true, then the day average will be computed.

If mean keyword is empty, and daily keyword is false, then the Ap at Julian day JD will be computed.

By default, mean is empty and daily is false.

source
SatelliteToolbox.get_DstΔTcMethod
get_DstΔTc(JD::Number)

Get the value of the index DstΔTc at Julian Day JD.

This function requires the initialization of the variable _dtcfile_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_F10Method
get_F10(JD::Number)

Get the value of the index F10 at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_F81aMethod
get_F81a(JD::Number)

Get the value of the index F81a at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_M10Method
get_M10(JD::Number)

Get the value of the index M10 at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_M81aMethod
get_M81a(JD::Number)

Get the value of the index M81a at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_S10Method
get_S10(JD::Number)

Get the value of the index S10 at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_S81aMethod
get_S81a(JD::Number)

Get the value of the index S81a at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_Y10Method
get_Y10(JD::Number)

Get the value of the index Y10 at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_Y81aMethod
get_Y81a(JD::Number)

Get the value of the index Y81a at Julian Day JD.

This function requires the initialization of the variable _solfsmy_data. Otherwise, an exception will be raised. To initialize it, run init_space_indices().

source
SatelliteToolbox.get_iers_eopFunction
get_iers_eop(data_type::Symbol = :IAU1980; force_download = false)

Download and parse the IERS EOP C04 data. The data type is specified by data_type symbol. Supported values are:

  • IAU1980: Get IERS EOP C04 IAU1980 data.
  • IAU2000A: Get IERS EOP C04 IAU2000A data.

If data_type is omitted, then it defaults to IAU1980.

The files are downloaded using the RemoteFile package with daily updates. Hence, if one desires to force a download before the scheduled time, then set the keyword force_download to true.

Returns

A structure (EOPData_IAU1980 or EOPData_IAU2000A, depending on data_type) with the interpolations of the EOP parameters. Notice that the interpolation indexing is set to the Julian Day.

source
SatelliteToolbox.get_iers_eop_iau_1980Function
get_iers_eop_iau_1980(url::String = "https://datacenter.iers.org/data/latestVersion/223_EOP_C04_14.62-NOW.IAU1980223.txt")

Get the IERS EOP C04 IAU1980 data from the URL url. If url is omitted, then it defaults to https://datacenter.iers.org/data/latestVersion/223EOPC04_14.62-NOW.IAU1980223.txt

The file is downloaded using the RemoteFile package with daily updates. Hence, if one desires to force a download before the scheduled time, then set the keyword force_download to true.

Returns

The structure EOPData_IAU1980 with the interpolations of the EOP parameters. Notice that the interpolation indexing is set to the Julian Day.

Remarks

For every field in EOPData_IAU1980 to interpolation between two points in the grid is linear. If extrapolation is needed, then if will use the nearest value (flat extrapolation).

source
SatelliteToolbox.get_iers_eop_iau_2000AFunction
get_iers_eop_iau_2000A(url::String = "https://datacenter.iers.org/data/latestVersion/224_EOP_C04_14.62-NOW.IAU2000A224.txt"; force_download = false)

Get the IERS EOP C04 IAU2000A data from the URL url. If url is omitted, then it defaults to https://datacenter.iers.org/data/latestVersion/224EOPC04_14.62-NOW.IAU2000A224.txt

The file is downloaded using the RemoteFile package with daily updates. Hence, if one desires to force a download before the scheduled time, then set the keyword force_download to true.

Returns

The structure EOPData_IAU2000A with the interpolations of the EOP parameters. Notice that the interpolation indexing is set to the Julian Day.

Remarks

For every field in EOPData_IAU2000A to interpolation between two points in the grid is linear. If extrapolation is needed, then if will use the nearest value (flat extrapolation).

source
SatelliteToolbox.get_space_indexMethod
get_space_index(T, JD::Number; ...)

Return the space index T at the day JD [Julian Day]. T can be:

Daily 10.7-cm solar flux

The daily 10.7-cm solar flux can be obtained using:

  • F10(): 10.7-cm adjusted solar flux [10⁻²² W/(M² Hz)].
  • F10adj(): 10.7-cm adjusted solar flux [10⁻²² W/(M² Hz)].
  • F10obs(): 10.7-cm observed solar flux [10⁻²² W/(M² Hz)].

These indices require fluxtable (see init_space_indices).

Daily average 10.7-cm solar flux

The daily average 10.7-cm solar flux, centered at JD, can be obtained using:

  • F10M(): 10.7-cm adjusted solar flux [10⁻²² W/(M² Hz)].
  • F10Madj(): 10.7-cm adjusted solar flux [10⁻²² W/(M² Hz)].
  • F10Mobs(): 10.7-cm observed solar flux [10⁻²² W/(M² Hz)].

In this case, the keyword window::Int can be passed to select the size of the window. By default, it is selected as 81.

These indices require fluxtable (see init_space_indices).

Daily Kp and Ap

  • Kp(): Kp index (daily mean).
  • Kp_vect(): A vector containing the Kp index for the following hours of the day: 0-3h, 3-6h, 6-9h, 9-12h, 12-15h, 15-18h, 18-20h, 20-23h.
  • Ap(): Ap index (daily mean).
  • Ap_vect(): A vector containing the Ap index for the following hours of the day: 0-3h, 3-6h, 6-9h, 9-12h, 12-15h, 15-18h, 18-20h, 20-23h.

These indices require wdcfiles (see init_space_indices).

Daily S10, M10, and Y10

  • S10(): EUV index (26-34 nm) scaled to F10.7.
  • M10(): MG2 index scaled to F10.7.
  • Y10(): Solar X-ray & Lya index scaled to F10.7.

These indices require solfsmy (see init_space_indices).

81-day centered average of S10, M10, and Y10.

  • S81a: EUV 81-day averaged centered index.
  • M81a: MG2 81-day averaged centered index.
  • Y81a: Solar X-ray & Lya 81-day averaged centered index.

These indices require solfsmy (see init_space_indices).

Exospheric temperature variation due to Dst

  • DstΔTc: Exospheric temperature variation due to Dst [K].

This index requires dtcfile (see init_space_indices).

source
SatelliteToolbox.get_ΔATMethod
get_ΔAT(JD::Number)

Get the accumulated leap seconds (ΔAT) [s] between UTC and International Atomic Time (TAI) in the given JD. This function search for ΔAT in the array ΔAT_Data.

Remarks

If JD is before ΔAT_Data[1,1], then 10 will be returned. Notice that this can lead to errors.

If JD is after ΔAT_Data[end,1], then ΔAT_Data[end,2] will be returned, because it is not possible yet to predict when leap seconds will be added.

source
SatelliteToolbox.get_ΩFunction
get_Ω(orb)

Return the right ascention of the ascending node (RAAN) of the orbit representation orb [rad].

source
SatelliteToolbox.ground_station_accessesMethod
ground_station_accesses(orbp, vrs_e,     Δt, ECI, ECEF, vargs...; kwargs...)
ground_station_accesses(orbp, [(WGS84)], Δt, ECI, ECEF, vargs...; kwargs...)

Compute the accesses of a satellite with orbit propagator orbp (see init_orbit_propagator) to the ground stations defined in the vector vrs_e. The analysis interval begins in the propagator epoch and lasts Δt [s].

The ground stations can be specified by an array of 3×1 vectors describing the ground stations position in an ECEF frame vrs_e or by an array of tuples containing the WGS84 position of each ground station [(WGS84)]:

(latitude [rad], longitude [rad], altitude [m])

Args

  • ECI: Earth-Centered Inertial frame in which the state vector of the propagator is represented.
  • ECEF: Earth-Centered, Earth-fixed frame to be used for the analysis. It must be the same frame used to compute the ground station position vector.
  • vargs...: list of additional arguments to be passed to the function rECItoECEF when converting the ECI frame to the ECEF.

Keywords

  • θ: Minimum elevation angle for communication between the satellite and the ground stations [rad]. (Default = 10ᵒ)
  • reduction: A function that receives a boolean vector with the visibility between the satellite and each ground station. It must return a boolean value indicating if the access must be computed or not. This is useful to merge access time between two or more stations. (Default = v->|(v...) i.e. compute the access if at least one ground station is visible)
source
SatelliteToolbox.ground_station_gapsMethod
ground_station_gaps(args...; kwargs...)

Compute the gaps between the accesses of ground stations. The arguments and keywords are the same as the ones used in the function ground_station_accesses.

Notice that the gap analysis starts in the orbit propagator epoch and ends in the instant defined by the argument Δt.

source
SatelliteToolbox.ground_station_visibleMethod
ground_station_visible(r_e::AbstractVector, rs_e::AbstractVector, θ::Number)

Check if the satellite with position vector r_e (ECEF) is inside the visibility circle of a ground station with position vector rs_e (ECEF) and a minimum elevation angle of θ [rad].

Notice that r_e and rs_e must be represented in the same ECEF frame, and must have the same unit.

Returns true if the satellite is inside the visibility circle, or false otherwise.

source
SatelliteToolbox.ground_station_visibleMethod
ground_station_visible(r_e::AbstractVector, lat_s::Number, lon_s::Number, h_s::Number, θ::Number)

Check if the satellite with position vector r_e (ECEF) is inside the visibility circle of a ground station with latitude lat_s [rad], longitude lon_s [rad], altitude h_s (WGS-84), and a minimum elevation angle of θ [rad].

Notice that the units of r_e and h_s must be the same.

Returns true if the satellite is inside the visibility circle, or false otherwise.

source
SatelliteToolbox.ground_traceMethod
ground_trace(orbp::OrbitPropagator{N}, eop_data::Union{Nothing, EOPData_IAU1980, EOPData_IAU2000A} = nothing; ECI = TEME(), ECEF = PEF(), span = 1.0) where N

Compute the ground trace of the object with orbit defined by orbp.

By default, it considers that the orbit elements on the propagator are represented in the True Equator, Mean Equinox (TEME) reference frame and the ground trace will be computed in the Pseudo-Earth Fixed (PEF) reference frame. Hence, no EOP data is needed. However, this can be changed by the keywords presented as follows.

Keywords

  • eop_data: EOP data that will be used to convert the ECI reference frame to the ECEF reference frame. If nothing, then it will not be used (see rECItoECEF). (Default = nothing)
  • ECI: ECI frame in which the orbit elements in orbp are represented. (Default = TEME())
  • ECEF: ECEF frame that will be used to compute the ground trace. (Default = PEF())
  • span: Defines for how much time the ground trace will be computed. The unit is the orbit period. (Default = 1.0)
  • dt: Time interval between two samples [s]. (Default = 10.0)

Returns

A vector of tuples with the pairs (latitude,longitude) of the ground trace.

source
SatelliteToolbox.gtd7Method
gtd7(nrlmsise00d::NRLMSISE00_Structure{T}) where T<:Number

NRLMSISE-00

Neutral Atmosphere Empirical Model from the surface to lower exosphere.

This routine computes the NRLMSISE-00 outputs (see NRLMSISE00_Output) using the configurations in the structure nrlmsise00 (see NRLMSISE00_Structure).

Args

  • nrlmsise00d: An instance of NRLMSISE00_Structure.

Returns

An instance of structure NRLMSISE00_Output with the outputs.

In this case, the total mass den_Total (see NRLMSISE00_Output) is the sum of the mass densities of the species He, O, N₂, O₂, Ar, H, and N, but does not include anomalous oxygen.

Remarks

  1. The densities of O, H, and N are set to 0 below 72.5 km.
  2. The exospheric temperature T_exo is set to global average for altitudes below 120 km. The 120 km gradient is left at global average value for altitudes below 72.5 km.
  3. Anomalous oxygen is defined as hot atomic oxygen or ionized oxygen that can become appreciable at high altitudes (> 500 km) for some ranges of inputs, thereby affection drag on satellites and debris. We group these species under the term Anomalous Oxygen, since their individual variations are not presently separable with the drag data used to define this model component.
source
SatelliteToolbox.gtd7dMethod
gtd7d(nrlmsise00d::NRLMSISE00_Structure{T}) where T<:Number

NRLMSISE-00

Neutral Atmosphere Empirical Model from the surface to lower exosphere.

This routine computes the NRLMSISE-00 outputs (see NRLMSISE00_Output) using the configurations in the structure nrlmsise00 (see NRLMSISE00_Structure).

Args

  • nrlmsise00d: An instance of NRLMSISE00_Structure.

Returns

An instance of structure NRLMSISE00_Output with the outputs.

In this case, the total mass den_Total (see NRLMSISE00_Output) is the effective total mass density for drag and is the sum of the mass densities of all species in this model including the anomalous oxygen.

Remarks

  1. The densities of O, H, and N are set to 0 below 72.5 km.
  2. The exospheric temperature T_exo is set to global average for altitudes below 120 km. The 120 km gradient is left at global average value for altitudes below 72.5 km.
  3. Anomalous oxygen is defined as hot atomic oxygen or ionized oxygen that can become appreciable at high altitudes (> 500 km) for some ranges of inputs, thereby affection drag on satellites and debris. We group these species under the term Anomalous Oxygen, since their individual variations are not presently separable with the drag data used to define this model component.
source
SatelliteToolbox.gts7Method
gts7(nrlmsise00d::NRLMSISE00_Structure{T}) where T<:Number

Thermospheric portion of NRLMSISE-00. This function should not be called to compute NRLMSISE-00. Use gtd7 or gtd7d instead.

Args

  • nrlmsise00d: An instance of NRLMSISE00_Structure.

Returns

An instance of structure NRLMSISE00_Structure with the outputs.

source
SatelliteToolbox.igrfMethod
igrf(date::Number, [r,h]::Number, λ::Number, Ω::Number, T[, P, dP]; show_warns = true)

IGRF Model

Current version: v13

Compute the geomagnetic field vector [nT] at the date date [Year A.D.] and position (r, λ, Ω).

The position representation is defined by T. If T is Val(:geocentric), then the input must be geocentric coordinates:

  1. Distance from the Earth center r [m];
  2. Geocentric latitude λ (-π/2, +π/2) [rad]; and
  3. Geocentric longitude Ω (-π, +π) [rad].

If T is Val(:geodetic), then the input must be geodetic coordinates:

1 Altitude above the reference ellipsoid h (WGS-84) [m];

  1. Geodetic latitude λ (-π/2, +π/2) [rad]; and
  2. Geodetic longitude Ω (-π, +π) [rad].

If T is omitted, then it defaults to Val(:geocentric).

Notice that the output vector will be represented in the same reference system selected by the parameter T (geocentric or geodetic). The Y-axis of the output reference system always points East. In case of geocentric coordinates, the Z-axis points toward the center of Earth and the X-axis completes a right-handed coordinate system. In case of geodetic coordinates, the X-axis is tangent to the ellipsoid at the selected location and points toward North, whereas the Z-axis completes a right-hand coordinate system.

The optional arguments P and dP must be two matrices with at least 14x14 real numbers. If they are present, then they will be used to store the Legendre coefficients and their derivatives. In this case, no allocation will be performed when computing the magnetic field. If they are not present, then 2 allocations will happen to create them.

Keywords

  • show_warns: Show warnings about the data (Default = true).

Remarks

The date must be greater or equal to 1900 and less than or equal 2030. Notice that a warning message is printed for dates greater than 2025.

Disclaimer

This function is an independent implementation of the IGRF model. It contains a more readable code than the original one in FORTRAN, because it uses features available in Julia language.

source
SatelliteToolbox.igrf12synMethod
igrf12syn(isv::Int, date::Number, itype::Int, alt::Number, colat::Number, elong::Number; show_warns = true)

This is a Julia implementation of the official IGRF source code, which was written in Fortran [2]. The input and output variables are exactly the same as the ones described in the function igrf12syn in [2].

Args

  • isv: 0 if main-field values are required, 1 if secular variation values are required.
  • date: Year A.D.
  • itype: 1 if geodetic (spheroid), 2 if geocentric (sphere).
  • alt: Height above sea level [km] if itype = 1, or distance from the center of Earth [km] if itype = 2 (must be > 3485 km).
  • colat: Colatitude (0 - 180) [˚].
  • elong: East-Longitude (0 - 360) [˚].

Keywords

  • show_warns: Show warnings about the data (Default = true).

Returns

  • The north component [nT] if isv = 0, or [nT/year] if isv = 1.
  • The east component [nT] if isv = 0, or [nT/year] if isv = 1.
  • The vertical component [nT] if isv = 0, or [nT/year] if isv = 1.
  • The total intensity if isv = 0, or rubbish if isv = 1.

Remarks

  • The date must be greater or equal to 1900 and less than or equal 2025. Notice that a warning message is printed for dates grated than 2020.
source
SatelliteToolbox.igrf13synMethod
igrf13syn(isv::Int, date::Number, itype::Int, alt::Number, colat::Number, elong::Number; show_warns = true)

This is a Julia implementation of the official IGRF source code, which was written in Fortran [2]. The input and output variables are exactly the same as the ones described in the function igrf13syn in [2].

Args

  • isv: 0 if main-field values are required, 1 if secular variation values are required.
  • date: Year A.D.
  • itype: 1 if geodetic (spheroid), 2 if geocentric (sphere).
  • alt: Height above sea level [km] if itype = 1, or distance from the center of Earth [km] if itype = 2 (must be > 3485 km).
  • colat: Colatitude (0 - 180) [˚].
  • elong: East-Longitude (0 - 360) [˚].

Keywords

  • show_warns: Show warnings about the data (Default = true).

Returns

  • The north component [nT] if isv = 0, or [nT/year] if isv = 1.
  • The east component [nT] if isv = 0, or [nT/year] if isv = 1.
  • The vertical component [nT] if isv = 0, or [nT/year] if isv = 1.
  • The total intensity if isv = 0, or rubbish if isv = 1.

Remarks

  • The date must be greater or equal to 1900 and less than or equal 2030. Notice that a warning message is printed for dates grated than 2025.
source
SatelliteToolbox.igrfdMethod
igrfd(date::Number, [r,h]::Number, λ::Number, Ω::Number, T[, P, dP]; show_warns = true)

IGRF Model

Current version: v13

Compute the geomagnetic field vector [nT] at the date date [Year A.D.] and position (r, λ, Ω).

The position representation is defined by T. If T is Val(:geocentric), then the input must be geocentric coordinates:

  1. Distance from the Earth center r [m];
  2. Geocentric latitude λ (-90°, +90°); and
  3. Geocentric longitude Ω (-180°, +180°).

If T is Val(:geodetic), then the input must be geodetic coordinates:

1 Altitude above the reference ellipsoid h (WGS-84) [m];

  1. Geodetic latitude λ (-90°, +90°); and
  2. Geodetic longitude Ω (-180°, +180°).

If T is omitted, then it defaults to Val(:geocentric).

Notice that the output vector will be represented in the same reference system selected by the parameter T (geocentric or geodetic). The Y-axis of the output reference system always points East. In case of geocentric coordinates, the Z-axis points toward the center of Earth and the X-axis completes a right-handed coordinate system. In case of geodetic coordinates, the X-axis is tangent to the ellipsoid at the selected location and points toward North, whereas the Z-axis completes a right-hand coordinate system.

The optional arguments P and dP must be two matrices with at least 14x14 real numbers. If they are present, then they will be used to store the Legendre coefficients and their derivatives. In this case, no allocation will be performed when computing the magnetic field. If they are not present, then 2 allocations will happen to create them.

Keywords

  • show_warns: Show warnings about the data (Default = true).

Remarks

The date must be greater or equal to 1900 and less than or equal 2030. Notice that a warning message is printed for dates greater than 2025.

Disclaimer

This function is an independent implementation of the IGRF model. It contains a more readable code than the original one in FORTRAN, because it uses features available in Julia language.

source
SatelliteToolbox.init_orbit_propagatorMethod
init_orbit_propagator(Val(:J2), epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number, dn_o2::Number = 0, ddn_o6::Number = 0; j2_gc::J2_GravCte{T} = j2_gc_egm08) where T
init_orbit_propagator(Val(:J2), orb_0::Orbit, dn_o2::Number = 0, ddn_o6::Number = 0; j2_gc::J2_GravCte = j2_gc_egm08)

Initialize the J2 orbit propagator.

Args

  • epoch: Initial orbit epoch [Julian Day].
  • a_0: Initial mean semi-major axis [m].
  • e_0: Initial mean eccentricity.
  • i_0: Initial mean inclination [rad].
  • Ω_0: Initial mean right ascension of the ascending node [rad].
  • ω_0: Initial mean argument of perigee [rad].
  • f_0: Initial mean true anomaly [rad].
  • dn_o2: (OPTIONAL) First time derivative of mean motion divided by 2 [rad/s²] (Default = 0).
  • ddn_o6: (OPTIONAL) Second time derivative of mean motion divided by 6 [rad/s³] (Default = 0).
  • orb_0: Instance of the structure KeplerianElements with the initial mean orbital elements [SI].

Keywords

  • j2_gc: (OPTIONAL) J2 orbit propagator gravitational constants (Default = j2_gc_egm08).
source
SatelliteToolbox.init_orbit_propagatorMethod
init_orbit_propagator(Val(:J4), epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number, dn_o2::Number = 0, ddn_o6::Number = 0; j4_gc::J4_GravCte{T} = j4_gc_egm08) where T
init_orbit_propagator(Val(:J4), orb_0::Orbit, dn_o2::Number = 0, ddn_o6::Number = 0; j4_gc::J4_GravCte = j4_gc_egm08)

Initialize the J4 orbit propagator.

Args

  • epoch: Initial orbit epoch [Julian Day].
  • a_0: Initial mean semi-major axis [m].
  • e_0: Initial mean eccentricity.
  • i_0: Initial mean inclination [rad].
  • Ω_0: Initial mean right ascension of the ascending node [rad].
  • ω_0: Initial mean argument of perigee [rad].
  • f_0: Initial mean true anomaly [rad].
  • dn_o2: (OPTIONAL) First time derivative of mean motion divided by 2 [rad/s²] (Default = 0).
  • ddn_o6: (OPTIONAL) Second time derivative of mean motion divided by 6 [rad/s³] (Default = 0).
  • orb_0: Instance of the structure KeplerianElements with the initial mean orbital elements [SI].

Keywords

  • j4_gc: (OPTIONAL) J4 orbit propagator gravitational constants (Default = j4_gc_egm08).
source
SatelliteToolbox.init_orbit_propagatorMethod
init_orbit_propagator(Val(:sgp4), tle::TLE, sgp4_gc::SGP4_GravCte{T} = sgp4_gc_wgs84) where T

Initialize the SGP4 orbit propagator using the TLE tle.

Keywords

  • sgp4_gc: (OPTIONAL) Gravitational constants. (Default = sgp4_gc_wgs84)
source
SatelliteToolbox.init_orbit_propagatorMethod
init_orbit_propagator(::Val{:twobody}, epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number; μ::T = m0) where T

Initialize the two body orbit propagator.

Args

  • epoch: Initial orbit epoch [Julian Day].
  • a_0: Initial mean semi-major axis [m].
  • e_0: Initial mean eccentricity.
  • i_0: Initial mean inclination [rad].
  • Ω_0: Initial mean right ascension of the ascending node [rad].
  • ω_0: Initial mean argument of perigee [rad].
  • f_0: Initial mean true anomaly [rad].
  • orb_0: Instance of the structure KeplerianElements with the initial mean orbital elements [SI].

Keywords

  • μ: (OPTIONAL) Standard gravitational parameter of the central body [m^3/s^2] (Default = m0).
source
SatelliteToolbox.init_space_indicesMethod
init_space_indices(...)

Initialize all space indices. The files that will be initialized must be indicated by the array of symbols passed to the keyword argument enabled_files. If this is nothing, which is the default, then all files will be initialized. The symbol related to each file is described next.

Notice that the initialization process can be changed by a set of keywords as described next.

DTCFILE

Symbol: :dtcfile

This file contains the exospheric temperature variation caused by the Dst index. This is used for the JB2008 atmospheric model.

Keywords

  • dtcfile_path: Path for the file DTCFILE.TXT. If nothing, then it will be downloaded. (Default = nothing)
  • dtcfile_force_download: If true, then the file will always be downloaded if the path is not specified. (Default = false).

fluxtable

Symbol: :fluxtable

This file contains the F10.7 flux data in different formats.

Keywords

  • fluxtable_path: Path for the file fluxtable.txt. If nothing, then it will be downloaded. (Default = nothing)
  • fluxtable_force_download: If true, then the file will always be downloaded if the path is not specified. (Default = false).

SOLFSMY

Symbol: :solfsmy

This files contains the indices necessary for the JB2008 atmospheric model.

Keywords

  • solfsmy_path: Path for the file SOLFSMY.TXT. If nothing, then it will be downloaded. (Default = nothing)
  • solfsmy_force_download: If true, then the file will always be downloaded if the path is not specified. (Default = false).

WDC Files

Symbol: :wdcfiles

This set of files contain the Kp and Ap indices.

Keywords

  • wdcfiles_path: Path for the directory with the WDC files. If nothing, then they will be downloaded. (Default = nothing)
  • wdcfiles_force_download: If true, then the files will always be downloaded if the path is not specified. (Default = false).
  • wdcfiles_oldest_year: Oldest year in which the WDC file will be obtained. (Default = past 3 years).
  • wdcfiles_newest_year: Newest year in which the WDC file will be obtained. If it is nothing, then it defaults to the current year. (Default = nothing).
source
SatelliteToolbox.is_leap_yearMethod
is_leap_year(year::Integer)

Check if the year year is a leap year. It returns true if year is a leap year, or false otherwise.

Remarks

This algorithm was based on [3].

source
SatelliteToolbox.j2!Method
j2!(j2d::J2_Structure{T}, t::Number) where T

Propagate the orbit defined in j2d (see J2_Structure) until the time t [s]. Notice that the values in j2d will be modified.

Returns

  • The position vector represented in the inertial frame at time t [m].
  • The velocity vector represented in the inertial frame at time t [m/s]

Remarks

The inertial frame in which the output is represented depends on which frame it was used to generate the orbit parameters. Notice that the perturbation theory requires an inertial frame with true equator.

source
SatelliteToolbox.j2_initMethod
j2_init(epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number, dn_o2::Number = 0, ddn_o6::Number = 0; j2_gc::J2_GravCte{T} = j2_gc_egm08 ) where T

Initialize the data structure of J2 orbit propagator algorithm.

Args

  • epoch: Epoch of the initial mean orbital elements [Julian Day].
  • a_0: Initial mean semi-major axis [m].
  • e_0: Initial mean eccentricity.
  • i_0: Initial mean inclination [rad].
  • Ω_0: Initial mean right ascension of the ascending node [rad].
  • ω_0: Initial mean argument of perigee [rad].
  • f_0: Initial mean true anomaly [rad].
  • dn_o2: First time derivative of the mean motion divided by two [rad/s^2].
  • ddn_o6: Second time derivative of the mean motion divided by six [rad/s^3].

Keywords

  • j2_gc: J2 orbit propagator gravitational constants (see J2_GravCte). (Default = j2_gc_egm08)

Returns

The structure J2_Structure with the initialized parameters.

source
SatelliteToolbox.j2osc!Method
j2osc!(j2d::J2osc_Structure{T}, t::Number) where T

Propagate the orbit defined in j2oscd (see J2osc_Structure) until the time t [s]. Notice that the values in j2oscd will be modified.

Returns

  • The position vector represented in the inertial frame at time t [m].
  • The velocity vector represented in the inertial frame at time t [m/s]

Remarks

The inertial frame in which the output is represented depends on which frame it was used to generate the orbit parameters. If the orbit parameters are obtained from a TLE, then the inertial frame will be TEME. Notice, however, that the perturbation theory requires an inertial frame with true equator.

source
SatelliteToolbox.j2osc_initMethod
j2osc_init(j2_gc::J2_GravCte{T}, epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number, dn_o2::Number, ddn_o6::Number) where T

Initialize the data structure of J2 osculating orbit propagator algorithm.

Args

  • j2_gc: J2 orbit propagator gravitational constants (see J2_GravCte).
  • epoch: Epoch of the orbital elements [Julian Day].
  • a_0: Initial semi-major axis [m].
  • e_0: Initial eccentricity.
  • i_0: Initial inclination [rad].
  • Ω_0: Initial right ascension of the ascending node [rad].
  • ω_0: Initial argument of perigee [rad].
  • f_0: Initial true anomaly [rad].
  • dn_o2: First time derivative of the mean motion divided by two [rad/s^2].
  • ddn_o6: Second time derivative of the mean motion divided by six [rad/s^3].

Returns

The structure J2osc_Structure with the initialized parameters.

Remarks

The inputs are the mean orbital elements.

source
SatelliteToolbox.j4!Method
j4!(j4d::J4_Structure{T}, t::Number) where T

Propagate the orbit defined in j4d (see J4_Structure) until the time t [s]. Notice that the values in j4d will be modified.

Returns

  • The position vector represented in the inertial frame at time t [m].
  • The velocity vector represented in the inertial frame at time t [m/s]

Remarks

The inertial frame in which the output is represented depends on which frame it was used to generate the orbit parameters. Notice that the perturbation theory requires an inertial frame with true equator.

source
SatelliteToolbox.j4_initMethod
j4_init(epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number, dn_o2::Number = 0, ddn_o6::Number = 0; j4_gc::J4_GravCte{T} = j4_gc_egm08) where T

Initialize the data structure of J4 orbit propagator algorithm.

Args

  • epoch: Epoch of the initial mean orbital elements [Julian Day].
  • a_0: Initial mean semi-major axis [m].
  • e_0: Initial mean eccentricity.
  • i_0: Initial mean inclination [rad].
  • Ω_0: Initial mean right ascension of the ascending node [rad].
  • ω_0: Initial mean argument of perigee [rad].
  • f_0: Initial mean true anomaly [rad].
  • dn_o2: First time derivative of the mean motion divided by two [rad/s^2].
  • ddn_o6: Second time derivative of the mean motion divided by six [rad/s^3].

Keywords

  • j4_gc: J4 orbit propagator gravitational constants (see J4_GravCte). (Default = j4_gc_egm08)

Returns

The structure J4_Structure with the initialized parameters.

source
SatelliteToolbox.jb2008Method
jb2008(JD::Number, glat::Number, glon::Number, h::Number)
jb2008(JD::Number, glat::Number, glon::Number, h::Number, F10::Number, F10ₐ::Number, S10::Number, S10ₐ::Number, M10::Number, M10ₐ::Number, Y10::Number, Y10ₐ::Number, DstΔTc::Number)

Compute the atmospheric density using the Jacchia-Bowman 2008 (JB2008) model.

If the space indices are not provided (first call), then they will be obtained from the online database. In this case, the function init_space_indices() must be called first and the function will throw an exception if the selected JD is outside of the available data.

This model is a product of the Space Environment Technologies, more information can be seen in the websites:

http://sol.spacenvironment.net/jb2006/

http://sol.spacenvironment.net/jb2008/

Args

  • JD: Julian day.

  • glat: Geocentric latitude [rad].

  • glon: Geocentric longitude [rad].

  • h: Altitude [m].

  • F10: 10.7-cm solar flux [10⁻²² W/(M² Hz)] (Tabular time 1 day earlier).

  • F10ₐ: 10.7-cm averaged solar flux, 81-day centered on input time (Tabular time 1 day earlier).

  • S10: EUV index (26-34 nm) scaled to F10.7 (Tabular time 1 day earlier).

  • S10ₐ: EUV 81-day averaged centered index (Tabular time 1 day earlier).

  • M10: MG2 index scaled to F10.7 (Tabular time 2 days earlier).

  • M10ₐ: MG2 81-day averaged centered index (Tabular time 2 days earlier).

  • Y10: Solar X-ray & Lya index scaled to F10.7 (Tabular time 5 days earlier).

  • Y10ₐ: Solar X-ray & Lya 81-day averaged centered index (Tabular time 5 days earlier).

  • DstΔTc: Temperature variation related to the Dst.

Returns

An instance of the structure JB2008_Output with the computed values.

source
SatelliteToolbox.jr1971Method
jr1971(JD::Number, glat::Number, glon::Number, h::Number, F10::Number, F10ₐ::Number, Kp::Number)

Compute the atmospheric density using the Jacchia-Roberts 1971 model.

Args

  • JD: Julian day.
  • glat: Geodetic latitude [rad].
  • glon: Geodetic longitude [rad].
  • h: Altitude [m].
  • F10: 10.7-cm solar flux [10⁻²² W/(M² Hz)].
  • F10ₐ: 10.7-cm averaged solar flux, 81-day centered on input time.
  • Kp: Kp geomagnetic index (with a delay of 3 hours).

Returns

An instance of the structure JR1971_Output with the computed values.

source
SatelliteToolbox.kepler_to_rvMethod
kepler_to_rv(a::Number, e::Number, i::Number, Ω::Number, ω::Number, f::Number)
kepler_to_rv(k::KeplerianElements)

Convert the Keplerian elements (a, e, i, Ω, ω, and f) to a Cartesian representation (position vector r and velocity vector v). The Keplerian elements can also be passed inside an instance of the KeplerianElements structure.

Args

  • a: Semi-major axis [m].
  • e: Eccentricity.
  • i: Inclination [rad].
  • Ω: Right ascension of the ascending node [rad].
  • ω: Argument of perigee [rad].
  • f: True anomaly [rad].

Returns

  • The position vector represented in the inertial reference frame [m].
  • The velocity vector represented in the inertial reference frame [m].

References

This algorithm was adapted from [1] and [3, p. 37-38].

source
SatelliteToolbox.legendreFunction
legendre([N,] ϕ::Number, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false)

Compute the associated Legendre function P_n,m[cos(ϕ)]. The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

The optional parameter N can be used to select the normalization. The following values are valid:

  • Val(:full): Compute the fully normalized associated Legendre function (see legendre_fully_normalized).
  • Val(:schmidt): Compute the Schmidt quasi-normalized associated Legendre function (see legendre_schmidt_quasi_normalized).
  • Val(:conv): Compute the conventional associated Legendre function (see legendre_conventional).

If N is omitted, then the full normalization will be used (Val(:full)).

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the Legendre associated functions P_n,m[cos(ϕ)].

source
SatelliteToolbox.legendre!Function
legendre!([N,] P::AbstractMatrix, ϕ::Number, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the associated Legendre function P_n,m[cos(ϕ)]. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix P will be used.

The result will be stored at matrix P.

The optional parameter N can be used to select the normalization. The following values are valid:

  • Val(:full): Compute the fully normalized associated Legendre function (see legendre_fully_normalized!).
  • Val(:schmidt): Compute the Schmidt quasi-normalized associated Legendre function (see legendre_schmidt_quasi_normalized!).
  • Val(:conv): Compute the conventional associated Legendre function (see legendre_conventional!).

If N is omitted, then the full normalization will be used.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

source
SatelliteToolbox.legendre_conventional!Function
legendre_conventional!(P::AbstractMatrix, ϕ::Number, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the conventional associated Legendre function P_n,m[cos(ϕ)]. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix P will be used:

maximum degree -> number of rows
maximum order  -> number of columns

The result will be stored at matrix P.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

source
SatelliteToolbox.legendre_conventionalMethod
legendre_conventional(ϕ::T, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false) where T<:AbstractFloat

Compute the conventional associated Legendre function P_n,m[cos(ϕ)]. The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the Legendre associated functions P_n,m[cos(ϕ)].

source
SatelliteToolbox.legendre_fully_normalized!Function
legendre_fully_normalized!(P::AbstractMatrix, ϕ::Number, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the fully normalized associated Legendre function P_n,m[cos(ϕ)]. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix P will be used:

maximum degree -> number of rows
maximum order  -> number of columns

The result will be stored at matrix P.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Remarks

This algorithm was based on [1]. Our definition of fully normalized associated Legendre function can be seen in [2, p. 546]. The conversion is obtained by:

             _                     -
            |  (n-m)! . k . (2n+1)  |      k = 1 if m  = 0
K_n,m = sqrt| --------------------- |,     k = 2 if m != 0
            |         (n+m)!        |
             -                     -
_
P_n,m = P_n,m * K_n,m,

      _
where P_n,m is the fully normalized Legendre associated function.
source
SatelliteToolbox.legendre_fully_normalizedMethod
legendre_fully_normalized(ϕ::T, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false) where T<:AbstractFloat

Compute the fully normalized associated Legendre function P_n,m[cos(ϕ)]. The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the Legendre associated functions P_n,m[cos(ϕ)].

Remarks

This algorithm was based on [1]. Our definition of fully normalized associated Legendre function can be seen in [2, p. 546]. The conversion is obtained by:

             _                     -
            |  (n-m)! . k . (2n+1)  |      k = 1 if m  = 0
K_n,m = sqrt| --------------------- |,     k = 2 if m != 0
            |         (n+m)!        |
             -                     -
_
P_n,m = P_n,m * K_n,m,

      _
where P_n,m is the fully normalized Legendre associated function.
source
SatelliteToolbox.legendre_schmidt_quasi_normalized!Function
legendre_schmidt_quasi_normalized!(P::AbstractMatrix, ϕ::Number, ph_term::Bool = false, n_max::Integer = -1, m_max::Integer = -1)

Compute the Schmidt quasi-normalized associated Legendre function P_n,m[cos(ϕ)] [3,4]. The maximum degree and order that will be computed are given by the parameters n_max and m_max. If they are negative, then the dimensions of matrix P will be used:

maximum degree -> number of rows
maximum order  -> number of columns

The result will be stored at matrix P.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Remarks

This algorithm was based on [3,4]. The conversion is obtained by:

             _           -
            |     (n-m)!  |    k = 1 if m  = 0
K_n,m = sqrt| k. -------- |,   k = 2 if m != 0
            |     (n+m)!  |
             -           -

=
P_n,m = P_n,m * K_n,m,

      =
where P_n,m is the quasi-normalized normalized Legendre associated function.
source
SatelliteToolbox.legendre_schmidt_quasi_normalizedMethod
legendre_schmidt_quasi_normalized(ϕ::T, n_max::Integer, m_max::Integer = -1, ph_term::Bool = false) where T<:AbstractFloat

Compute the Schmidt quasi-normalized associated Legendre function P_n,m[cos(ϕ)]. The maximum degree that will be computed is n_max and the maximum order is m_max. Notice that if m_max is higher than n_max or negative, than it is set to n_max.

If ph_term is set to true, then the Condon-Shortley phase term (-1)ᵐ will be included. If ph_term is not present, then it defaults to false.

Returns

A matrix with the Legendre associated functions P_n,m[cos(ϕ)].

Remarks

This algorithm was based on [3,4]. The conversion is obtained by:

             _           -
            |     (n-m)!  |    k = 1 if m  = 0
K_n,m = sqrt| k. -------- |,   k = 2 if m != 0
            |     (n+m)!  |
             -           -

=
P_n,m = P_n,m * K_n,m,

      =
where P_n,m is the quasi-normalized normalized Legendre associated function.
source
SatelliteToolbox.list_ground_station_accessesMethod
list_ground_station_accesses(io, vargs...; kwargs...)

Print the ground station accesses to the io io. The arguments vargs... and keywords kwargs... are those of the function ground_station_accesses.

Additionally, the following keywords can be used to modify the behavior of this function:

  • format: If :pretty, then a formatted table will be printed. If :csv, then the access data will be printed using the CSV format. (Default = :pretty)
  • time_scale: Select the time scale of the access duration (:s for seconds, :m for minutes, and :h for hours). (Default = :m)
source
SatelliteToolbox.list_ground_station_gapsMethod
list_ground_station_gaps(io, vargs...; kwargs...)

Print the ground station gaps to the io io. The arguments vargs... and keywords kwargs... are those of the function ground_station_gaps.

Additionally, the following keywords can be used to modify the behavior of this function:

  • format: If :pretty, then a formatted table will be printed. If :csv, then the access data will be printed using the CSV format. (Default = :pretty)
  • time_scale: Select the time scale of the access duration (:s for seconds, :m for minutes, and :h for hours). (Default = :m)
source
SatelliteToolbox.list_ss_orbits_by_rep_periodFunction
list_ss_orbits_by_rep_period(minRep::Int, maxRep::Int, minAlt::Number=-1.0, maxAlt::Number=-1.0, e::Number=0.0)

Compute a list of repeating Sun-synchronous orbits.

Args

  • minRep: Minimum repetition time of the orbit [days].
  • maxRep: Maximum repetition time of the orbit [days].
  • minAlt: Minimum altitude of the orbits on the list [m].
  • maxAlt: Minimum altitude of the orbits on the list [m].
  • e: Eccentricity.

Returns

A matrix containing the orbits found with the following format:

Semi-major axis [m] | Altitude [m] | Inclination [rad] | Period [s] | Int | Num | Den
--------------------|--------------|-------------------|------------|-----|-----|-----

in which the period is Int + Num/Den.

Remarks

If minAlt or maxAlt is < 0.0, then the altitude will not be checked when a orbit is added to the list.

source
SatelliteToolbox.load_gravity_modelMethod
load_gravity_model(T)

Load an embedded gravity model coefficients T and return an instance of the structure GravityModel_Coefs with the parsed values.

The current supported values for T are:

TModel NameMaximum Degree
EGM96()Earth Gravitational Model 1996360
JGM2()Joint Earth Gravity Model 270
JGM3()Joint Earth Gravity Model 370
–––––-––––––––––––––––––––––––

For other models, you can downlad the gfc file at

http://icgem.gfz-potsdam.de/home

and load it using the functions parse_icgem and create_gravity_model_coefs.

source
SatelliteToolbox.luni_solar_args_iau2006Method
luni_solar_args_iau2006(JD_TT::Number)

Compute the fundamental arguments related to the luni-solar effect for the IAU-2006 theory [1, p. 211].

The returned values are in [rad].

source
SatelliteToolbox.mean_obliquity_iau2006Method
mean_obliquity_iau2006(JD_TT::Number)

Compute the mean obliquity of the ecliptic [rad] using the equinox-based IAU-2006 theory in the Julian day JD_TT [Terrestiral Time].

The algorithm was obtained in [3].

source
SatelliteToolbox.minimum_half_FOV_grssMethod
minimum_half_FOV_grss(h::Real, T::Real, i::Real, To::Integer)

Compute the minimum half FOV of a ground repeating Sun-synchronous (GRSS) orbit to cover the entire Equator within the revisit interval.

Args

  • h: Orbit altitude in the Equator [m].
  • T: Orbit period [s].
  • i: Inclination [rad].
  • To: Orbit cycle [days].

Returns

The minimum half FOV [rad].

source
SatelliteToolbox.minimum_half_FOV_grssMethod
minimum_half_FOV_grss(h::Real, a::Real, e::Real, i::Real, To::Integer)

Compute the minimum half FOV of a ground repeating Sun-synchronous (GRSS) orbit to cover the entire Equator within the revisit interval.

Args

  • h: Orbit altitude in the Equator [m].
  • a: Semi-major axis [m].
  • e: Eccentricity.
  • i: Inclination [rad].
  • To: Orbit cycle [days].

Returns

The minimum half FOV [rad].

source
SatelliteToolbox.minimum_swath_grssMethod
minimum_swath_grss(T::Real, i::Real, To::Integer)

Compute the minimum swath of a ground repeating Sun-synchronous (GRSS) orbit to cover the entire Equator within the revisit interval.

Args

  • T: Orbit period [s].
  • i: Inclination [rad].
  • To: Orbit cycle [days].

Returns

The minimum swath [m].

source
SatelliteToolbox.minimum_swath_grssMethod
minimum_swath_grss(a::Real, e::Real, i::Real, To::Integer)

Compute the minimum swath of a ground repeating Sun-synchronous (GRSS) orbit to cover the entire Equator within the revisit interval.

Args

  • a: Semi-major axis [m].
  • e: Eccentricity.
  • i: Inclination [rad].
  • To: Orbit cycle [days].

Returns

The minimum swath [m].

source
SatelliteToolbox.moon_position_iMethod
moon_position_i(JD_TDB::Number)

Compute the Moon position represented in the IAU-76/FK5 (mean-equator, mean-equinox), also called as J2000, at the Julian Day JD. The algorithm was adapted from [1, p. 288].

source
SatelliteToolbox.nrlmsise00Method
nrlmsise00(JD::Number, alt::Number, g_lat::Number, g_long::Number [, f107A::Number, f107::Number, ap::Union{Number,AbstractVector}]; output_si::Bool = true, dversion::Bool = true)

NRLMSISE-00

Neutral Atmosphere Empirical Model from the surface to lower exosphere.

This routine computes the NRLMSISE-00 outputs (see NRLMSISE00_Output) using the configurations in the structure nrlmsise00 (see NRLMSISE00_Structure).

Notice that the NRLMSISE-00 will be run using the default flags (see NRLMSISE00_DEFAULT_FLAGS). The user can only change the value of flags[:output_m_kg] using the keyword output_si to select whether the output must be converted to SI units. If more control is needed, then the user must manually call the function conf_nrlmsise00 and then call gtd7 or gtd7d with the desired flags.

If the space indices f107A, f107, and ap are missing, then they will be obtained from the online databases (see init_space_indices()).

Args

  • JD: Julian Day [UTC].
  • alt: Altitude [m].
  • g_lat: Geodetic latitude [rad].
  • g_long: Geodetic longitude [rad].
  • f107A: 81 day average of F10.7 flux (centered on day of year JD).
  • f107: Daily F10.7 flux for previous day.
  • ap: Magnetic index (daily) if it is a number. If it is an array, then see Remarks.

Keywords

  • output_si: (OPTIONAL) If true, then the output units will be [m⁻³] for species number density and [kg/m⁻³] for the total density. Otherwise, the units will be [cm⁻³] and [g/cm⁻³], respectively.
  • dversion: (OPTIONAL) If true, run gtd7d. Otherwise, run gtd7 (see Remarks).

Returns

An instance of the structure NRLMSISE00_Output. The result in variable den_Total depends on the value of dversion (see Remarks, Notes on input variables).

Remarks

  1. The densities of O, H, and N are set to 0 below 72.5 km.
  2. The exospheric temperature T_exo is set to global average for altitudes below 120 km. The 120 km gradient is left at global average value for altitudes below 72.5 km.
  3. Anomalous oxygen is defined as hot atomic oxygen or ionized oxygen that can become appreciable at high altitudes (> 500 km) for some ranges of inputs, thereby affection drag on satellites and debris. We group these species under the term Anomalous Oxygen, since their individual variations are not presently separable with the drag data used to define this model component.

AP

If ap is a Vector, then it must be a vector with 7 dimensions as described below:

IndexDescription
1Daily AP.
23 hour AP index for current time.
33 hour AP index for 3 hours before current time.
43 hour AP index for 6 hours before current time.
53 hour AP index for 9 hours before current time.
6Average of eight 3 hour AP indices from 12 to 33 hours prior to current time.
7Average of eight 3 hour AP indices from 36 to 57 hours prior to current time.

Notes on input variables

f107 and f107A values used to generate the model correspond to the 10.7 cm radio flux at the actual distance of the Earth from the Sun rather than the radio flux at 1 AU. The following site provides both classes of values:

ftp://ftp.ngdc.noaa.gov/STP/SOLAR_DATA/SOLAR_RADIO/FLUX/

f107, f107A, and ap effects are neither large nor well established below 80 km and these parameters should be set to 150, 150, and 4 respectively.

If dversion is true, then the total mass den_Total (see NRLMSISE00_Output) is the sum of the mass densities of the species He, O, N₂, O₂, Ar, H, and N, but does not include anomalous oxygen.

If dversion is false, then total mass den_Total (see NRLMSISE00_Output) is the effective total mass density for drag and is the sum of the mass densities of all species in this model including the anomalous oxygen.

source
SatelliteToolbox.nutation_eo_iau2006Function
nutation_eo_iau2006(JD_TT::Number)

Compute the nutation parameters and the Equation of Origins (EO) at the Julian Day JD_TT [TT] using the equinox-based 2006 IAU Theory of Nutation. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_2000) [rad] and in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

Returns

  • The mean obliquity of the ecliptic [rad].
  • The nutation in obliquity of the ecliptic [rad].
  • The nutation in longitude [rad].
  • The Equation of Origins (EO) [rad].
source
SatelliteToolbox.nutation_fk5Function
nutation_fk5(JD_TT::Number, n_max::Number = 106, nut_coefs_1980::Matrix = nut_coefs_1980)

Compute the nutation parameters at the Julian Day JD_TT [Terrestrial Time] using the 1980 IAU Theory of Nutation. The coefficients are nut_coefs_1980 that must be a matrix in which each line has the following syntax [1, p. 1043]:

an1  an2  an3  an4  an5  Ai  Bi  Ci  Di

where the units of Ai and Ci are [0.0001"] and the units of Bi and Di are [0.0001"/JC]. The user can also specify the number of coefficients n_max that will be used when computing the nutation. If n_max is omitted, the it defaults to 106.

Returns

  • The mean obliquity of the ecliptic [rad].
  • The nutation in obliquity of the ecliptic [rad].
  • The nutation in longitude [rad].
source
SatelliteToolbox.orbsvMethod
orbsv(t::T1, r::AbstractVector{T2}, v::AbstractVector{T3} = [0,0,0], a::AbstractVector{T4} = [0,0,0]) where {T1<:Number, T2<:Number, T3<:Number, T4<:Number}
orbsv(t::T1, vec::AbstractVector{T2}) where {T1<:Number, T2<:Number}

Create a new satellite state vector (see OrbitStateVector) using the position r, velocity v, and acceleration a. It is also possible to pass a vector vec with the information concatenated.

Info

The vectors r, v, and a must have at least 3 elements. In the case more elements are available, they will be neglected. On the other hand, the vector v must have 6 or 9 dimensions, indicating [r;v], or [r;v;a].

source
SatelliteToolbox.parse_icgemMethod
parse_icgem(filename::AbstractString)

Parse the ICGEM file filename and return an instance of the structure ICGEM with the parsed data.

source
SatelliteToolbox.periodFunction
period(a::Number, e::Number, i::Number, pert::Symbol = :J2)

Compute the period [s] of an object in an orbit with semi-major axis a [m], eccentricity e, and inclination i [rad], using the perturbation terms specified by the symbol pert. The orbit can also be specified by orb (see Orbit).

pert` can be:

  • :J0: Consider a Keplerian orbit.
  • :J2: Consider the perturbation terms up to J2.
  • :J4: Consider the perturbation terms J2, J4, and J2².

If pert is omitted, then it defaults to :J2.

source
SatelliteToolbox.planetary_args_iau2006Method
planetary_args_iau2006(JD_TT::Number)

Compute the fundamental arguments related to the planetary effects for the IAU-2006 theory [1, p. 211].

The returned values are in [rad].

source
SatelliteToolbox.precession_fk5Method
precession_fk5(JD_TT::Number)

Compute the angles related to the precession movement in the Julian Day JD_TT [Terrestrial Time] using the theory IAU-76/FK5.

Returns

The angles (ζ, Θ, z) as described in [1, p. 226-228].

source
SatelliteToolbox.precession_iau2006Method
precession_iau2006(JD_TT::Number)

Compute the precession angles [rad] according to equinox-based IAU-2006 theory in the Julia day JD_TT [Terrestrial Time].

This algorithm was obtained from [3, p. 49].

source
SatelliteToolbox.propagate!Function
propagate!(orbp::OrbitPropagator{T}, t::Number) where T
propagate!(orbp::OrbitPropagator{T}, t::AbstractVector) where T

If t is a number, then propagate orbp by t [s] from the orbit epoch. Otherwise, if t is an array, then propagate the orbit in orbp using the time instants defined in the vector t [s].

In both cases, the orbit propagator algorithm is the one related to the structure orbp.

The structure orbp will contain the elements at the last propagation instant.

Returns

  • The Keplerian elements represented in inertial frame in each time instant (see KeplerianElements) [SI units].
  • The position vector represented in inertial frame in each time instant [m].
  • The velocity vector represented in inertial frame in each time instant [m].

If t is an array, then those values will be an array containing the information related to each epoch in t.

source
SatelliteToolbox.propagate_to_epoch!Method
propagate_to_epoch!(orbp::OrbitPropagator{T}, JD::Number) where T
propagate_to_epoch!(orbp::OrbitPropagator{T}, JD::AbstractVector) where T

If t is a number, then propagate orbp until the epoch JD [Julian Day]. Otherwise, if JD is an array, then propagate the orbit in orbp using the epochs defined in the vector t [Julian Day].

In both cases, the orbit propagator algorithm is the one related to the structure orbp.

The structure orbp will contain the elements at the last propagation instant.

Returns

  • The Keplerian elements represented in inertial frame in each time instant (see KeplerianElements) [SI units].
  • The position vector represented in inertial frame in each time instant [m].
  • The velocity vector represented in inertial frame in each time instant [m].

If JD is an array, then those values will be an array containing the information related to each epoch in JD.

source
SatelliteToolbox.rCIRStoGCRF_iau2006Function
rCIRStoGCRF_iau2006([T::Type,] JD_TT::Number, dX::Number = 0, dY::Number = 0)

Compute the rotation that aligns the Celestial Intermediate Reference System (CIRS) with the Geocentric Celestial Reference Frame (GCRF) at the Julian Day JD_TT [TT] and considering the IERS EOP Data dX [rad] and dY [rad] (see get_iers_eop). This algorithm uses the IAU-2006 theory.

The IERS EOP Data dX and dY accounts for the free-core nutation and time dependent effects of the Celestial Intermediate Pole (CIP) position with respect to the GCRF.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the CIRS frame with the GCRF frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rCIRStoTIRS_iau2006Method
rCIRStoTIRS_iau2006([T::Type,] JD_UT1::Number)

Compute the rotation that aligns the Celestial Intermediate Reference System (CIRS) with the Terrestrial Intermediate Reference System (TIRS) at the Julian Day JD_UT1 [UT1]. This algorithm uses the IAU-2006 theory.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the CIRS frame with the TIRS frame. The rotation representation is selected by the optional parameter T.

Remarks

The reference frames TIRS and CIRS are separated by a rotation about the Z-axis of the Earth Rotation Angle, which is the angle between the Conventional International Origin (CIO) and the Terrestrial Intermediate Origin (TIO) [1]. The latter is a reference meridian on Earth that is located about 100m away from Greenwich meridian along the equator of the Celestial Intermediate Pole (CIP) [1].

source
SatelliteToolbox.rECEFtoECEFMethod
rECEFtoECEF([T,] ECEFo, ECEFf, JD_UTC::Number, eop_data)

Compute the rotation from an Earth-Centered, Earth-Fixed (ECEF) reference frame to another ECEF reference frame at the Julian Day [UTC] JD_UTC. The rotation description that will be used is given by T, which can be DCM or Quaternion. The origin ECEF frame is selected by the input ECEFo and the destination ECEF frame is selected by the input ECEFf. The model used to compute the rotation is specified by the selection of the origin and destination frames. Currently, there are two models supported: IAU-76/FK5 and IAU-2006 with 2010 conventions (CIO approach only).

Rotation description

The rotations that aligns the origin ECEF frame with the destination ECEF frame can be described by Direction Cosine Matrices or Quaternions. This is selected by the parameter T.

The possible values are:

  • DCM: The rotation will be described by a Direction Cosine Matrix.
  • Quaternion: The rotation will be described by a Quaternion.

If no value is specified, then it falls back to DCM.

Conversion model

The model that will be used to compute the rotation is automatically inferred given the selection of the origin and destination frames. Notice that mixing IAU-76/FK5 and IAU-2006/2010 frames is not supported yet.

ECEF Frame

The supported ECEF frames for both origin ECEFo and destination ECEFf are:

  • ITRF(): ECEF will be selected as the International Terrestrial Reference Frame (ITRF).
  • PEF(): ECEF will be selected as the Pseudo-Earth Fixed (PEF) reference frame.
  • TIRS(): ECEF will be selected as the Terrestrial Intermediate Reference System (TIRS).

EOP Data

The conversion between the supported ECEF frames always depends on EOP Data (see get_iers_eop and read_iers_eop). If IAU-76/FK5 model is used, then the type of eop_data must be EOPData_IAU1980. Otherwise, if IAU-2006/2010 model is used, then the type of eop_data must be EOPData_IAU2000A.

Returns

The rotation description represented by T that rotates the ECEF reference frame into alignment with the ECI reference frame.

Examples

julia> eop_IAU1980 = get_iers_eop(:IAU1980);

julia> rECEFtoECEF(PEF(), ITRF(), DatetoJD(1986,6,19,21,35,0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
  1.0          0.0         4.35684e-7
  0.0          1.0         1.44762e-6
 -4.35684e-7  -1.44762e-6  1.0

julia> rECEFtoECEF(Quaternion, PEF(), ITRF(), DatetoJD(1986,6,19,21,35,0), eop_IAU1980)
Quaternion{Float64}:
  + 0.9999999999997147 - 7.236343481310813e-7.i + 2.1765518308012794e-7.j + 0.0.k

julia> eop_IAU2000A = get_iers_eop(:IAU2000A);

julia> rECEFtoECEF(TIRS(), ITRF(), DatetoJD(1986,6,19,21,35,0), eop_IAU2000A)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
  1.0          3.08408e-11  -4.3531e-7
 -3.14708e-11  1.0          -1.44727e-6
  4.3531e-7    1.44727e-6    1.0

julia> rECEFtoECEF(Quaternion, TIRS(), ITRF(), DatetoJD(1986,6,19,21,35,0), eop_IAU2000A)
Quaternion{Float64}:
  + 0.9999999999997146 - 7.236343481345639e-7.i + 2.176551830689726e-7.j + 1.5577911634233308e-11.k
source
SatelliteToolbox.rECEFtoECIMethod
rECEFtoECI([T,] ECEF, ECI, JD_UTC::Number [, eop_data])

Compute the rotation from an Earth-Centered, Earth-Fixed (ECEF) reference frame to an Earth-Centered Inertial (ECI) reference frame at the Julian Day [UTC] JD_UTC. The rotation description that will be used is given by T, which can be DCM or Quaternion. The ECEF frame is selected by the input ECEF and the ECI frame is selected by the input ECI. The possible values are listed below. The model used to compute the rotation is specified by the selection of the origin and destination frames. Currently, there are two models supported: IAU-76/FK5 and IAU-2006 with 2010 conventions (CIO and equinox approaches).

Rotation description

The rotations that aligns the ECEF with ECI can be described by Direction Cosine Matrices or Quaternions. This is selected by the parameter T. The possible values are:

  • DCM: The rotation will be described by a Direction Cosine Matrix.
  • Quaternion: The rotation will be described by a Quaternion.

If no value is specified, then it falls back to DCM.

Conversion model

The model that will be used to compute the rotation is automatically inferred given the selection of the origin and destination frames. Notice that mixing IAU-76/FK5 and IAU-2006/2010 frames is not supported yet.

ECEF Frame

The ECEF frame is selected by the parameter ECEF. The possible values are:

  • ITRF(): ECEF will be selected as the International Terrestrial Reference Frame (ITRF).
  • PEF(): ECEF will be selected as the Pseudo-Earth Fixed (PEF) reference frame.
  • TIRS(): ECEF will be selected as the Terrestrial Intermediate Reference System (TIRS).

ECI Frame

The ECI frame is selected by the parameter ECI. The possible values are:

  • TEME(): ECI will be selected as the True Equator Mean Equinox (TEME) reference frame.
  • TOD(): ECI will be selected as the True of Date (TOD).
  • MOD(): ECI will be selected as the Mean of Date (MOD).
  • J2000(): ECI will be selected as the J2000 reference frame.
  • GCRF(): ECI will be selected as the Geocentric Celestial Reference Frame (GCRF).
  • CIRS(): ECI will be selected as the Celestial Intermediate Reference System (CIRS).
  • ERS(): ECI will be selected as the Earth Reference System (ERS).
  • MOD06(): ECI will be selected as the Mean of Date (MOD) according to the definition in IAU-2006/2010 theory.
  • MJ2000(): ECI will be selected as the J2000 mean equatorial frame (MJ2000).
Note

The frames MOD() and MOD06() are virtually the same. However, we selected different names to make clear which theory are being used since mixing transformation between frames from IAU-76/FK5 and IAU-2006/2010 must be performed with caution.

EOP Data

The conversion between the frames depends on EOP Data (see get_iers_eop and read_iers_eop). If IAU-76/FK5 model is used, then the type of eop_data must be EOPData_IAU1980. Otherwise, if IAU-2006/2010 model is used, then the type of eop_data must be EOPData_IAU2000A. The following table shows the requirements for EOP data given the selected frames.

ModelECEFECIEOP Data
IAU-76/FK5ITRFGCRFEOP IAU1980
IAU-76/FK5ITRFJ2000EOP IAU1980
IAU-76/FK5ITRFMODEOP IAU1980
IAU-76/FK5ITRFTODEOP IAU1980
IAU-76/FK5ITRFTEMEEOP IAU1980
IAU-76/FK5PEFGCRFEOP IAU1980
IAU-76/FK5PEFJ2000Not required¹
IAU-76/FK5PEFMODNot required¹
IAU-76/FK5PEFTODNot required¹
IAU-76/FK5PEFTEMENot required¹
IAU-2006/2010 CIO-basedITRFCIRSEOP IAU2000A
IAU-2006/2010 CIO-basedITRFGCRFEOP IAU2000A
IAU-2006/2010 CIO-basedTIRSCIRSNot required¹
IAU-2006/2010 CIO-basedTIRSGCRFNot required¹ ²
IAU-2006/2010 Equinox-basedITRFERSEOP IAU2000A
IAU-2006/2010 Equinox-basedITRFMOD06EOP IAU2000A
IAU-2006/2010 Equinox-basedITRFMJ2000EOP IAU2000A
IAU-2006/2010 Equinox-basedTIRSERSNot required¹ ³
IAU-2006/2010 Equinox-basedTIRSMOD06Not required¹ ³
IAU-2006/2010 Equinox-basedTIRSMJ2000Not required¹ ³

¹: In this case, the Julian Time UTC will be assumed equal to Julian Time UT1 to compute the Greenwich Mean Sidereal Time. This is an approximation, but should be sufficiently accurate for some applications. Notice that, if EOP Data is provided, the Julian Day UT1 will be accurately computed.

²: In this case, the terms that account for the free core nutation and time dependent effects of the Celestial Intermediate Pole (CIP) position with respect to the GCRF will not be available, reducing the precision.

³: In this case, the terms that corrects the nutation in obliquity and in longitude due to the free core nutation will not be available, reducing the precision.

MOD and TOD

In this function, if EOP corrections are not provided, then MOD and TOD frames will be computed considering the original IAU-76/FK5 theory. Otherwise, the corrected frame will be used.

Returns

The rotation description represented by T that rotates the ECEF reference frame into alignment with the ECI reference frame.

Examples

julia> eop_IAU1980 = get_iers_eop(:IAU1980);

julia> rECEFtoECI(DCM, ITRF(), GCRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267      0.78518     -0.00132979
 -0.78518      -0.619267     3.33492e-5
 -0.000797313   0.00106478   0.999999

julia> rECEFtoECI(ITRF(), GCRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267      0.78518     -0.00132979
 -0.78518      -0.619267     3.33492e-5
 -0.000797313   0.00106478   0.999999

julia> rECEFtoECI(PEF(), J2000(), DatetoJD(1986, 06, 19, 21, 35, 0))
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619271      0.785176    -0.00133066
 -0.785177     -0.619272     3.45854e-5
 -0.000796885   0.00106622   0.999999

julia> rECEFtoECI(PEF(), J2000(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267      0.78518     -0.00133066
 -0.78518      -0.619267     3.45854e-5
 -0.000796879   0.00106623   0.999999

julia> rECEFtoECI(Quaternion, ITRF(), GCRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
Quaternion{Float64}:
  + 0.4363098936462618 - 0.0005909969666939257.i + 0.00030510511316206974.j + 0.8997962182293519.k

julia> eop_IAU2000A = get_iers_eop(:IAU2000A);

julia> rECEFtoECI(ITRF(), GCRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU2000A)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267      0.78518     -0.00132979
 -0.78518      -0.619267     3.33502e-5
 -0.000797312   0.00106478   0.999999

julia> rECEFtoECI(TIRS(), GCRF(), DatetoJD(1986, 06, 19, 21, 35, 0))
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619271      0.785176    -0.00133066
 -0.785177     -0.619272     3.45884e-5
 -0.000796885   0.00106623   0.999999

julia> rECEFtoECI(Quaternion, ITRF(), GCRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU2000A)
Quaternion{Float64}:
  + 0.4363098936309669 - 0.000590996988144556.i + 0.0003051056555230158.j + 0.8997962182365703.k
source
SatelliteToolbox.rECItoECEFMethod
rECItoECEF([T,] ECI, ECEF, JD_UTC::Number [, eop_data])

Compute the rotation from an Earth-Centered Inertial (ECI) reference frame to an Earth-Centered, Earth-Fixed (ECEF) reference frame at the Julian Day [UTC] JD_UTC. The rotation description that will be used is given by T, which can be DCM or Quaternion. The ECI frame is selected by the input ECI and the ECEF frame is selected by the input ECEF. The possible values are listed below. The model used to compute the rotation is specified by the selection of the origin and destination frames. Currently, there are two models supported: IAU-76/FK5 and IAU-2006 with 2010 conventions (CIO and equinox approaches).

Rotation description

The rotations that aligns the ECI with ECEF can be described by Direction Cosine Matrices or Quaternions. This is selected by the parameter T. The possible values are:

  • DCM: The rotation will be described by a Direction Cosine Matrix.
  • Quaternion: The rotation will be described by a Quaternion.

If no value is specified, then it falls back to DCM.

Conversion model

The model that will be used to compute the rotation is automatically inferred given the selection of the origin and destination frames. Notice that mixing IAU-76/FK5 and IAU-2006/2010 frames is not supported yet.

ECI Frame

The ECI frame is selected by the parameter ECI. The possible values are:

  • TEME(): ECI will be selected as the True Equator Mean Equinox (TEME) reference frame.
  • TOD(): ECI will be selected as the True of Date (TOD).
  • MOD(): ECI will be selected as the Mean of Date (MOD).
  • J2000(): ECI will be selected as the J2000 reference frame.
  • GCRF(): ECI will be selected as the Geocentric Celestial Reference Frame (GCRF).
  • CIRS(): ECEF will be selected as the Celestial Intermediate Reference System (CIRS).
  • ERS(): ECI will be selected as the Earth Reference System (ERS).
  • MOD06(): ECI will be selected as the Mean of Date (MOD) according to the definition in IAU-2006/2010 theory.
  • MJ2000(): ECI will be selected as the J2000 mean equatorial frame (MJ2000).
Note

The frames MOD() and MOD06() are virtually the same. However, we selected different names to make clear which theory are being used since mixing transformation between frames from IAU-76/FK5 and IAU-2006/2010 must be performed with caution.

ECEF Frame

The ECEF frame is selected by the parameter ECEF. The possible values are:

  • ITRF(): ECEF will be selected as the International Terrestrial Reference Frame (ITRF).
  • PEF(): ECEF will be selected as the Pseudo-Earth Fixed (PEF) reference frame.
  • TIRS(): ECEF will be selected as the Terrestrial Intermediate Reference System (TIRS).

EOP Data

The conversion between the frames depends on EOP Data (see get_iers_eop and read_iers_eop). If IAU-76/FK5 model is used, then the type of eop_data must be EOPData_IAU1980. Otherwise, if IAU-2006/2010 model is used, then the type of eop_data must be EOPData_IAU2000A. The following table shows the requirements for EOP data given the selected frames.

ModelECIECEFEOP Data
IAU-76/FK5GCRFITRFEOP IAU1980
IAU-76/FK5J2000ITRFEOP IAU1980
IAU-76/FK5MODITRFEOP IAU1980
IAU-76/FK5TODITRFEOP IAU1980
IAU-76/FK5TEMEITRFEOP IAU1980
IAU-76/FK5GCRFPEFEOP IAU1980
IAU-76/FK5J2000PEFNot required¹
IAU-76/FK5MODPEFNot required¹
IAU-76/FK5TODPEFNot required¹
IAU-76/FK5TEMEPEFNot required¹
IAU-2006/2010 CIO-basedCIRSITRFEOP IAU2000A
IAU-2006/2010 CIO-basedGCRFITRFEOP IAU2000A
IAU-2006/2010 CIO-basedCIRSTIRSNot required¹
IAU-2006/2010 CIO-basedGCRFTIRSNot required¹ ²
IAU-2006/2010 Equinox-basedERSTIRSEOP IAU2000A
IAU-2006/2010 Equinox-basedMOD06ITRFEOP IAU2000A
IAU-2006/2010 Equinox-basedMJ2000ITRFEOP IAU2000A
IAU-2006/2010 Equinox-basedERSTIRSNot required¹ ³
IAU-2006/2010 Equinox-basedMOD06TIRSNot required¹ ³
IAU-2006/2010 Equinox-basedMJ2000TIRSNot required¹ ³

¹: In this case, the Julian Time UTC will be assumed equal to Julian Time UT1 to compute the Greenwich Mean Sidereal Time. This is an approximation, but should be sufficiently accurate for some applications. Notice that, if EOP Data is provided, the Julian Day UT1 will be accurately computed.

²: In this case, the terms that account for the free-core nutation and time dependent effects of the Celestial Intermediate Pole (CIP) position with respect to the GCRF will not be available, reducing the precision.

MOD and TOD

In this function, if EOP corrections are not provided, then MOD and TOD frames will be computed considering the original IAU-76/FK5 theory. Otherwise, the corrected frame will be used.

Returns

The rotation description represented by T that rotates the ECI reference frame into alignment with the ECEF reference frame.

Examples

julia> eop_IAU1980 = get_iers_eop(:IAU1980);

julia> rECItoECEF(DCM, GCRF(), ITRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267    -0.78518     -0.000797313
  0.78518     -0.619267     0.00106478
 -0.00132979   3.33492e-5   0.999999

julia> rECItoECEF(GCRF(), ITRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267    -0.78518     -0.000797313
  0.78518     -0.619267     0.00106478
 -0.00132979   3.33492e-5   0.999999

julia> rECItoECEF(J2000(), PEF(), DatetoJD(1986, 06, 19, 21, 35, 0))
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619271    -0.785177    -0.000796885
  0.785176    -0.619272     0.00106622
 -0.00133066   3.45854e-5   0.999999

julia> rECItoECEF(J2000(), PEF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267    -0.78518     -0.000796879
  0.78518     -0.619267     0.00106623
 -0.00133066   3.45854e-5   0.999999

julia> rECItoECEF(Quaternion, GCRF(), ITRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU1980)
Quaternion{Float64}:
  + 0.4363098936462618 + 0.0005909969666939257.i - 0.00030510511316206974.j - 0.8997962182293519.k

julia> eop_IAU2000A = get_iers_eop(:IAU2000A);

julia> rECItoECEF(GCRF(), ITRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU2000A)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619267    -0.78518     -0.000797312
  0.78518     -0.619267     0.00106478
 -0.00132979   3.33502e-5   0.999999

julia> rECItoECEF(GCRF(), TIRS(), DatetoJD(1986, 06, 19, 21, 35, 0))
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 -0.619271    -0.785177    -0.000796885
  0.785176    -0.619272     0.00106623
 -0.00133066   3.45884e-5   0.999999

julia> rECItoECEF(Quaternion, GCRF(), ITRF(), DatetoJD(1986, 06, 19, 21, 35, 0), eop_IAU2000A)
Quaternion{Float64}:
  + 0.4363098936309669 + 0.000590996988144556.i - 0.0003051056555230158.j - 0.8997962182365703.k
source
SatelliteToolbox.rECItoECIMethod
rECEFtoECI([T,] ECIo, ECIf, JD_UTC::Number [, eop_data])
rECEFtoECI([T,] ECIo, JD_UTCo::Number, ECIf, JD_UTCf::Number [, eop_data])

Compute the rotation from an Earth-Centered Inertial (ECI) reference frame to another ECI reference frame. If the origin and destination frame contain only one of date frame, then the first signature is used and JD_UTC is the epoch of this frame. On the other hand, if the origin and destination frame contain two of date frame[1], e.g. TOD => MOD, then the second signature must be used in which JD_UTCo is the epoch of the origin frame and JD_UTCf is the epoch of the destination frame.

The rotation description that will be used is given by T, which can be DCM or Quaternion. The origin ECI frame is selected by the input ECIo and the destination ECI frame is selected by the input ECIf. The model used to compute the rotation is specified by the selection of the origin and destination frames. Currently, there are two models supported: IAU-76/FK5 and IAU-2006 with 2010 conventions (CIO and equinox approaches).

Rotation description

The rotations that aligns the origin ECI frame with the destination ECI frame can be described by Direction Cosine Matrices or Quaternions. This is selected by the parameter T.

The possible values are:

  • DCM: The rotation will be described by a Direction Cosine Matrix.
  • Quaternion: The rotation will be described by a Quaternion.

If no value is specified, then it falls back to DCM.

Conversion model

The model that will be used to compute the rotation is automatically inferred given the selection of the origin and destination frames. Notice that mixing IAU-76/FK5 and IAU-2006/2010 frames is not supported yet.

ECI Frame

The supported ECI frames for both origin ECIo and destination ECIf are:

  • TEME(): ECI will be selected as the True Equator Mean Equinox (TEME) reference frame.
  • TOD(): ECI will be selected as the True of Date (TOD).
  • MOD(): ECI will be selected as the Mean of Date (MOD).
  • J2000(): ECI will be selected as the J2000 reference frame.
  • GCRF(): ECI will be selected as the Geocentric Celestial Reference Frame (GCRF).
  • CIRS(): ECEF will be selected as the Celestial Intermediate Reference System (CIRS).
  • ERS(): ECI will be selected as the Earth Reference System (ERS).
  • MOD06(): ECI will be selected as the Mean of Date (MOD) according to the definition in IAU-2006/2010 theory.
  • MJ2000(): ECI will be selected as the J2000 mean equatorial frame (MJ2000).
Note

The frames MOD() and MOD06() are virtually the same. However, we selected different names to make clear which theory are being used since mixing transformation between frames from IAU-76/FK5 and IAU-2006/2010 must be performed with caution.

EOP Data

The conversion between the frames depends on EOP Data (see get_iers_eop and read_iers_eop). If IAU-76/FK5 model is used, then the type of eop_data must be EOPData_IAU1980. Otherwise, if IAU-2006/2010 model is used, then the type of eop_data must be EOPData_IAU2000A. The following table shows the requirements for EOP data given the selected frames.

ModelECIoECIfEOP DataFunction Signature
IAU-76/FK5GCRFJ2000EOP IAU1980First
IAU-76/FK5GCRFMODEOP IAU1980First
IAU-76/FK5GCRFTODEOP IAU1980First
IAU-76/FK5GCRFTEMEEOP IAU1980First
IAU-76/FK5J2000GCRFEOP IAU1980First
IAU-76/FK5J2000MODNot requiredFirst
IAU-76/FK5J2000TODNot requiredFirst
IAU-76/FK5J2000TEMENot requiredFirst
IAU-76/FK5MODGCRFEOP IAU1980First
IAU-76/FK5MODJ2000Not requiredFirst
IAU-76/FK5MODTODNot requiredSecond
IAU-76/FK5MODTEMENot requiredSecond
IAU-76/FK5TODGCRFEOP IAU1980First
IAU-76/FK5TODJ2000Not requiredFirst
IAU-76/FK5TODMODNot requiredSecond
IAU-76/FK5TODTEMENot requiredSecond
IAU-76/FK5TEMEGCRFEOP IAU1980First
IAU-76/FK5TEMEJ2000Not requiredFirst
IAU-76/FK5TEMEMODNot requiredSecond
IAU-76/FK5TEMETODNot requiredSecond
IAU-2006/2010 CIO-basedGCRFCIRSNot required¹First
IAU-2006/2010 CIO-basedCIRSCIRSNot required¹Second
IAU-2006/2010 Equinox-basedGCRFMJ2000Not requiredFirst²
IAU-2006/2010 Equinox-basedGCRFMOD06Not requiredFirst
IAU-2006/2010 Equinox-basedGCRFERSNot required³First
IAU-2006/2010 Equinox-basedMJ2000GCRFNot requiredFirst²
IAU-2006/2010 Equinox-basedMJ2000MOD06Not requiredFirst
IAU-2006/2010 Equinox-basedMJ2000ERSNot required³First
IAU-2006/2010 Equinox-basedMOD06GCRFNot requiredFirst
IAU-2006/2010 Equinox-basedMOD06MJ2000Not requiredFirst
IAU-2006/2010 Equinox-basedMOD06ERSNot required³First
IAU-2006/2010 Equinox-basedERSGCRFNot required³First
IAU-2006/2010 Equinox-basedERSMJ2000Not required³First
IAU-2006/2010 Equinox-basedERSMOD06Not required³First

¹: In this case, the terms that account for the free-core nutation and time dependent effects of the Celestial Intermediate Pole (CIP) position with respect to the GCRF will not be available, reducing the precision.

²: The transformation between GCRF and MJ2000 is a constant rotation matrix called bias. Hence, the date does not modify it. However, this signature was kept to avoid complications in the API.

³: In this case, the terms that corrects the nutation in obliquity and in longitude due to the free core nutation will not be available, reducing the precision.

MOD and TOD

In this function, if EOP corrections are not provided, then MOD and TOD frames will be computed considering the original IAU-76/FK5 theory. Otherwise, the corrected frame will be used.

Returns

The rotation description represented by T that rotates the origin ECI reference frame into alignment with the destination ECI reference frame.

Examples

julia> eop_IAU1980 = get_iers_eop(:IAU1980);

julia> rECItoECI(DCM, GCRF(), J2000(), DatetoJD(1986, 6, 19, 21, 35, 0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
  1.0          -2.45469e-12   4.56602e-10
  2.45466e-12   1.0          -1.84455e-9
 -4.56602e-10   1.84455e-9    1.0

julia> rECItoECI(Quaternion, TEME(), GCRF(), DatetoJD(1986, 6, 19, 21, 35, 0), eop_IAU1980)
Quaternion{Float64}:
  + 0.9999986335698654 + 1.8300414020900853e-5.i + 0.0006653038276169474.j - 0.0015132396749411375.k

julia> rECItoECI(TOD(), DatetoJD(1986,6,19,21,35,0), TOD(), DatetoJD(1987,5,19,3,0,0), eop_IAU1980)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 1.0          -0.000224087  -9.73784e-5
 0.000224086   1.0          -5.79859e-6
 9.73797e-5    5.77677e-6    1.0

julia> rECItoECI(Quaternion, TOD(), JD_J2000, MOD(), JD_J2000, eop_IAU1980)
Quaternion{Float64}:
  + 0.9999999993282687 - 1.400220690336851e-5.i + 1.3473593746216003e-5.j - 3.107834312843103e-5.k

julia> rECItoECI(J2000(), TEME(), DatetoJD(1986,6,19,21,35,0))
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
  0.999995    0.0030265    0.00133055
 -0.00302645  0.999995    -3.86125e-5
 -0.00133066  3.45854e-5   0.999999

julia> eop_IAU2000A = get_iers_eop(:IAU2000A);

julia> rECItoECI(CIRS(), GCRF(), DatetoJD(1986,6,19,21,35,0), eop_IAU2000A)
3×3 StaticArrays.SArray{Tuple{3,3},Float64,2,9}:
 0.999999     3.88379e-8  -0.00133066
 7.18735e-9   1.0          3.45882e-5
 0.00133066  -3.45882e-5   0.999999

julia> rECItoECI(Quaternion, CIRS(), GCRF(), DatetoJD(1986,6,19,21,35,0), eop_IAU2000A)
Quaternion{Float64}:
  + 0.9999997785177528 + 1.7294102099105917e-5.i + 0.0006653310148723835.j + 7.912627369563795e-9.k
source
SatelliteToolbox.rERStoMOD_iau2006Function
rERStoMOD_iau2006([T::Type,] JD_TT::Number, δΔϵ_2000::Number = 0, δΔΨ_2000::Number = 0)

Compute the rotation that aligns the Earth Reference System (ERS) with the Mean of Date (MOD) reference frame at Julian day JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

Notice that one can provide corrections for the nutation in obliquity (δΔϵ_2000) and in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop and dEps_dPsi). This corrections are related to Free Core Nutation (FCN) that models the effect of a liquid Earth core.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the ERS frame with the MOD frame. The rotation representation is selected by the optional parameter T.

Remarks

The reference systems ERS and MOD are separated by the nutation of the pole.

source
SatelliteToolbox.rERStoTIRS_iau2006Function
rERStoTIRS_iau2006(JD_UT1::Number, JD_TT::Number, δΔΨ_2000::Number = 0)

Compute the rotation that aligns the Earth Reference System (ERS) with the Terrestrial Intermediate Reference System (TIRS) at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

Notice that one can provide corrections for the nutation in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop and dEps_dPsi). This corrections are related to Free Core Nutation (FCN) that models the effect of a liquid Earth core.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the ERS frame with the TIRS frame. The rotation representation is selected by the optional parameter T.

Remarks

The reference frames TIRS and ERS are separated by a rotation about the Z-axis of the Greenwhich apparent sidereal angle (GAST). This angle is computed using the IAU-2006 theory, which consist of obtaining the Earth Rotation Angle (ERA) and subtracting the result of the Equation of Origins (EO).

source
SatelliteToolbox.rGCRFtoCIRS_iau2006Function
rGCRFtoCIRS_iau2006([T::Type,] JD_TT::Number, dX::Number = 0, dY::Number = 0)

Compute the rotation that aligns the Geocentric Celestial Reference Frame (GCRF) with the Celestial Intermediate Reference System (CIRS) at the Julian Day JD_TT [TT] and considering the IERS EOP Data dX [rad] and dY [rad] (see get_iers_eop). This algorithm uses the IAU-2006 theory.

The IERS EOP Data dX and dY accounts for the free-core nutation and time dependent effects of the Celestial Intermediate Pole (CIP) position with respect to the GCRF.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the GCRF frame with the CIRS frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rGCRFtoITRF_fk5Function
rGCRFtoITRF_fk5([T,] JD_UT1::Number, JD_TT::Number, x_p::Number, y_p::Number [, δΔϵ_1980::Number, δΔψ_1980::Number])

Compute the rotation that aligns the Geocentric Celestial Reference Frame (GCRF) with the International Terrestrial Reference Frame (ITRF) at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time], and considering the IERS EOP Data x_p [rad], y_p [rad], δΔϵ_1980 [rad], and δΔψ_1980 [rad] (see get_iers_eop). This algorithm uses the IAU-76/FK5 theory.

x_p is the polar motion displacement about X-axis, which is the IERS Reference Meridian direction (positive south along the 0˚ longitude meridian). y_p is the polar motion displacement about Y-axis (90˚W or 270˚E meridian). δΔϵ_1980 is the nutation in obliquity. δΔψ_1980 is the nutation in longitude.

The Julian Day in UT1 is used to compute the Greenwich Mean Sidereal Time (GMST) (see JDtoGMST), whereas the Julian Day in Terrestrial Time is used to compute the nutation in the longitude. Notice that the Julian Day in UT1 and in Terrestrial Time must be equivalent, i.e. must be related to the same instant. This function does not check this.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the GCRF frame with the ITRF frame. The rotation representation is selected by the optional parameter T.

Remarks

The EOP data related to the polar motion (x_p and y_p) is required, since this is the only way available to compute the conversion ITRF <=> PEF (the models are highly imprecise since the motion is still not very well understood [1]). However, the EOP data related to the nutation of the obliquity (δΔϵ_1980) and the nutation of the longitude (δΔψ_1980) can be omitted. In this case, the GCRF frame is what is usually called J2000 reference frame.

source
SatelliteToolbox.rGCRFtoMJ2000_iau2006Function
rGCRFtoMJ2000_iau2006([T::Type,] JD_TT::Number = 0)

Compute the rotation that aligns the Geocentric Celestial Reference Frame (GCRF) with the J2000 mean equatorial frame. This algorithm uses the IAU-2006 theory. Notice that this rotation is just a bias matrix that does not depend on the date. However, this function receives the argument JD_TT just to keep the API compatibility.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MJ2000 frame with the MOD frame. The rotation representation is selected by the optional parameter T.

Remarks

According to [1], the frame bias that converts MJ2000 <=> GCRF is not a precise transformation for all the times.

source
SatelliteToolbox.rGCRFtoMOD_fk5Method
rGCRFtoMOD_fk5([T,] JD_TT::Number)

Compute the rotation that aligns the Geocentric Celestial Reference Frame (GCRF) with the Mean of Date (MOD) frame at the Julian Day [Terrestrial Time] JD_TT. This algorithm uses the IAU-76/FK5 theory.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the GCRF frame with the MOD frame. The rotation representation is selected by the optional parameter T.

Remarks

The Geocentric Celestial Reference Frame (GCRF) is rotated into the Mean of Date (MOD) frame considering the IAU 1976 Precession model.

Notice that if the conversion MOD => TOD is performed without considering the EOP corrections, then the GCRF in this rotation is what is usually called the J2000 reference frame.

source
SatelliteToolbox.rGCRFtoTEMEFunction
rGCRFtoTEME([T,] JD_TT::Number [, δΔϵ_1980::Number = 0, δΔψ_1980::Number = 0])

Compute the rotation that aligns the GCRF frame with the True Equator Mean Equinox (TEME) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233]. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the GCRF frame with the TEME frame. The rotation representation is selected by the optional parameter T.

Remarks

The EOP data related to the nutation of the obliquity (δΔϵ_1980) and the nutation of the longitude (δΔψ_1980) can be omitted. In this case, the GCRF frame is what is usually called J2000 reference frame.

source
SatelliteToolbox.rITRFtoGCRF_fk5Function
rITRFtoGCRF_fk5([T,] JD_UT1::Number, JD_TT::Number, x_p::Number, y_p::Number [, δΔϵ_1980::Number, δΔψ_1980::Number])

Compute the rotation that aligns the International Terrestrial Reference Frame (ITRF) with the Geocentric Celestial Reference Frame (GCRF) at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time], and considering the IERS EOP Data x_p [rad], y_p [rad], δΔϵ_1980 [rad], and δΔψ_1980 [rad] (see get_iers_eop). This algorithm uses the IAU-76/FK5 theory.

x_p is the polar motion displacement about X-axis, which is the IERS Reference Meridian direction (positive south along the 0˚ longitude meridian). y_p is the polar motion displacement about Y-axis (90˚W or 270˚E meridian). δΔϵ_1980 is the nutation in obliquity. δΔψ_1980 is the nutation in longitude.

The Julian Day in UT1 is used to compute the Greenwich Mean Sidereal Time (GMST) (see JDtoGMST), whereas the Julian Day in Terrestrial Time is used to compute the nutation in the longitude. Notice that the Julian Day in UT1 and in Terrestrial Time must be equivalent, i.e. must be related to the same instant. This function does not check this.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the ITRF frame with the GCRF frame. The rotation representation is selected by the optional parameter T.

Remarks

The EOP data related to the polar motion (x_p and y_p) is required, since this is the only way available to compute the conversion ITRF <=> PEF (the models are highly imprecise since the motion is still not very well understood [1]). However, the EOP data related to the nutation of the obliquity (δΔϵ_1980) and the nutation of the longitude (δΔψ_1980) can be omitted. In this case, the GCRF frame is what is usually called J2000 reference frame.

source
SatelliteToolbox.rITRFtoPEF_fk5Method
rITRFtoPEF_fk5([T,] x_p::Number, y_p::Number)

Compute the rotation that aligns the International Terrestrial Reference Frame (ITRF) with the Pseudo-Earth Fixed (PEF) frame considering the polar motion represented by the angles x_p [rad] and y_p [rad] that are obtained from IERS EOP Data (see get_iers_eop).

x_p is the polar motion displacement about X-axis, which is the IERS Reference Meridian direction (positive south along the 0˚ longitude meridian). y_p is the polar motion displacement about Y-axis (90˚W or 270˚E meridian).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the ITRF frame with the PEF frame. The rotation representation is selected by the optional parameter T.

Remarks

The ITRF is defined based on the International Reference Pole (IRP), which is the location of the terrestrial pole agreed by international committees [1]. The Pseudo-Earth Fixed, on the other hand, is defined based on the Earth axis of rotation, or the Celestial Intermediate Pole (CIP). Hence, PEF XY-plane contains the True Equator. Furthermore, since the recovered latitude and longitude are sensitive to the CIP, then it should be computed considering the PEF frame.

source
SatelliteToolbox.rITRFtoTIRS_iau2006Method
rITRFtoTIRS_iau2006([T::Type,] JD_TT::Number, x_p::Number, y_p::Number)

Compute the rotation that aligns the International Terrestrial Reference Frame (ITRF) with the Terrestrial Intermediate Reference System (TIRS) considering the polar motion represented by the angles x_p [rad] and y_p [rad] that are obtained from IERS EOP Data (see get_iers_eop).

x_p is the polar motion displacement about X-axis, which is the IERS Reference Meridian direction (positive south along the 0˚ longitude meridian). y_p is the polar motion displacement about Y-axis (90˚W or 270˚E meridian).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the ITRF frame with the TIRS frame. The rotation representation is selected by the optional parameter T.

Remarks

The ITRF is defined based on the International Reference Pole (IRP), which is the location of the terrestrial pole agreed by international committees [1]. The Terrestrial Intermediate Reference Frame (TIRS), on the other hand, is defined based on the Earth axis of rotation, or the Celestial Intermediate Pole (CIP). Hence, TIRS XY-plane contains the True Equator. Furthermore, since the recovered latitude and longitude are sensitive to the CIP, then it should be computed considering the TIRS frame.

The TIRS and PEF (IAU-76/FK5) are virtually the same reference frame, but according to [1] it is convenient to separate the names as the exact formulae differ.

source
SatelliteToolbox.rMJ2000toGCRF_iau2006Function
rMJ2000toGCRF_iau2006([T::Type,] JD_TT::Number = 0)

Compute the rotation that aligns the J2000 mean equatorial frame with the Geocentric Celestial Reference Frame (GCRF). This algorithm uses the IAU-2006 theory. Notice that this rotation is just a bias matrix that does not depend on the date. However, this function receives the argument JD_TT just to keep the API compatibility.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MJ2000 frame with the MOD frame. The rotation representation is selected by the optional parameter T.

Remarks

According to [1], the frame bias that converts MJ2000 <=> GCRF is not a precise transformation for all the times.

source
SatelliteToolbox.rMJ2000toMOD_iau2006Method
rMJ2000toMOD_iau2006([T::Type,] JD_TT::Number)

Compute the rotation that aligns the J2000 mean equatorial frame with the Mean of Date (MOD) reference frame with the at Julian day JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MJ2000 frame with the MOD frame. The rotation representation is selected by the optional parameter T.

Remarks

The J2000 reference frame here is not equal to the previous definition in FK5 theory. It is the reason why it is internally called MJ2000. According to [3]:

The mean equinox of J2000.0 to be considered is not the “rotational dynamical mean equinox of J2000.0” as used in the past, but the “inertial dynamical mean equinox of J2000.0” to which the recent numerical or analytical solutions refer. The latter is associated with the ecliptic in the inertial sense, which is the plane perpendicular to the angular momentum vector of the orbital motion of the Earth-Moon barycenter as computed from the velocity of the barycenter relative to an inertial system. The rotational equinox is associated with the ecliptic in the rotational sense, which is perpendicular to the angular momentum vector computed from the velocity referred to the rotating orbital plane of the Earth-Moon barycenter. (The difference between the two angular momenta is the angular momentum associated with the rotation of the orbital plane.)

source
SatelliteToolbox.rMODtoERS_iau2006Function
rMODtoERS_iau2006([T::Type,] JD_TT::Number, δΔϵ_2000::Number = 0, δΔΨ_2000::Number = 0)

Compute the rotation that aligns the Mean of Date (MOD) reference frame with the Earth Reference System (ERS) at Julian day JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

Notice that one can provide corrections for the nutation in obliquity (δΔϵ_2000) and in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop and dEps_dPsi). This corrections are related to Free Core Nutation (FCN) that models the effect of a liquid Earth core.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MOD frame with the ERS frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rMODtoGCRF_fk5Method
rMODtoGCRF_fk5([T,] JD_TT::Number)

Compute the rotation that aligns the Mean of Date (MOD) frame with the Geocentric Celestial Reference Frame (GCRF) at the Julian Day [Terrestrial Time] JD_TT. This algorithm uses the IAU-76/FK5 theory.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MOD frame with the GCRF frame. The rotation representation is selected by the optional parameter T.

Remarks

The Mean of Date (MOD) frame is rotated into the Geocentric Celestial Reference Frame (GCRF) considering the IAU 1976 Precession model.

Notice that if the conversion TOD => MOD is performed without considering the EOP corrections, then the GCRF obtained by this rotation is what is usually called the J2000 reference frame.

source
SatelliteToolbox.rMODtoMJ2000_iau2006Method
rMODtoMJ2000_iau2006([T::Type,] JD_TT::Number)

Compute the rotation that aligns the Mean of Date (MOD) reference frame with the J2000 mean equatorial frame at Julian day JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MOD frame with the MJ2000 frame. The rotation representation is selected by the optional parameter T.

Remarks

The J2000 reference frame here is not equal to the previous definition in FK5 theory. It is the reason why it is internally called MJ2000. According to [3]:

The mean equinox of J2000.0 to be considered is not the “rotational dynamical mean equinox of J2000.0” as used in the past, but the “inertial dynamical mean equinox of J2000.0” to which the recent numerical or analytical solutions refer. The latter is associated with the ecliptic in the inertial sense, which is the plane perpendicular to the angular momentum vector of the orbital motion of the Earth-Moon barycenter as computed from the velocity of the barycenter relative to an inertial system. The rotational equinox is associated with the ecliptic in the rotational sense, which is perpendicular to the angular momentum vector computed from the velocity referred to the rotating orbital plane of the Earth-Moon barycenter. (The difference between the two angular momenta is the angular momentum associated with the rotation of the orbital plane.)

source
SatelliteToolbox.rMODtoPEF_fk5Function
rMODtoPEF_fk5([T,] JD_UT1::Number, JD_TT::Number [, δΔϵ_1980::Number, δΔψ_1980::Number])

Compute the rotation that aligns the Mean of Date (MOD) reference frame with the Pseudo-Earth Fixed (PEF) frame at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The Julian Day in UT1 is used to compute the Greenwich Mean Sidereal Time (GMST) (see JDtoGMST), whereas the Julian Day in Terrestrial Time is used to compute the nutation in the longitude. Notice that the Julian Day in UT1 and in Terrestrial Time must be equivalent, i.e. must be related to the same instant. This function does not check this.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MOD frame with the PEF frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rMODtoTEMEFunction
rMODtoTEME([T,] JD_TT::Number [, δΔϵ_1980::Number = 0, δΔψ_1980::Number = 0])

Compute the rotation that aligns the Mean of Date (MOD) frame with the True Equator Mean Equinox (TEME) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233]. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop). .

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MOD frame with the TEME frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rMODtoTIRS_iau2006Function
rMODtoTIRS_iau2006([T::Type,] JD_UT1::Number, JD_TT::Number, δΔϵ_2000::Number = 0, δΔΨ_2000::Number = 0)

Compute the rotation that aligns the Mean of Date (MOD) reference frame with the Terrestrial Intermediate Reference System (TIRS) at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

Notice that one can provide corrections for the nutation in obliquity (δΔϵ_2000) and in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop and dEps_dPsi). This corrections are related to Free Core Nutation (FCN) that models the effect of a liquid Earth core.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TIRS frame with the ERS frame. The rotation representation is selected by the optional parameter T.

Remarks

This composed rotation TIRS <=> ERS <=> MOD is implemented as a new function because the single rotations TIRS <=> ERS and ERS <=> MOD call the function nutation_eo, which has a high computational burden. In this case, the composed algorithm is about 2x faster than calling those function separately.

source
SatelliteToolbox.rMODtoTOD_fk5Function
rMODtoTOD_fk5([T,] JD_TT::Number [, δΔϵ_1980::Number, δΔψ_1980::Number])

Compute the rotation that aligns the Mean of Date (MOD) frame with the True of Date (TOD) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the MOD frame with the TOD frame. The rotation representation is selected by the optional parameter T.

Remarks

The Mean of Date (MOD) frame is rotated into the True of Date (TOD) frame considering the 1980 IAU Theory of Nutation. The IERS EOP corrections must be added if one wants to make the rotation consistent with the Geocentric Celestial Reference Systems (GCRS).

source
SatelliteToolbox.rPEFtoITRF_fk5Method
rPEFtoITRF_fk5([T,] x_p::Number, y_p::Number)

Compute the rotation that aligns the Pseudo-Earth Fixed (PEF) with the International Terrestrial Reference Frame (ITRF) considering the polar motion represented by the angles x_p [rad] and y_p [rad] that are obtained from IERS EOP Data (see get_iers_eop).

x_p is the polar motion displacement about X-axis, which is the IERS Reference Meridian direction (positive south along the 0˚ longitude meridian). y_p is the polar motion displacement about Y-axis (90˚W or 270˚E meridian).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the PEF frame with the ITRF. The rotation representation is selected by the optional parameter T.

Remarks

The ITRF is defined based on the International Reference Pole (IRP), which is the location of the terrestrial pole agreed by international committees [1]. The Pseudo-Earth Fixed, on the other hand, is defined based on the Earth axis of rotation, or the Celestial Intermediate Pole (CIP). Hence, PEF XY-plane contains the True Equator. Furthermore, since the recovered latitude and longitude are sensitive to the CIP, then it should be computed considering the PEF frame.

source
SatelliteToolbox.rPEFtoMOD_fk5Function
rPEFtoMOD_fk5([T,] JD_UT1::Number, JD_TT::Number [, δΔϵ_1980::Number, δΔψ_1980::Number])

Compute the rotation that aligns the Pseudo-Earth Fixed (PEF) frame with the Mean of Date (MOD) at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The Julian Day in UT1 is used to compute the Greenwich Mean Sidereal Time (GMST) (see JDtoGMST), whereas the Julian Day in Terrestrial Time is used to compute the nutation in the longitude. Notice that the Julian Day in UT1 and in Terrestrial Time must be equivalent, i.e. must be related to the same instant. This function does not check this.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the PEF frame with the TOD frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rPEFtoTEMEMethod
rPEFtoTEME([T,] JD_TT::Number)

Compute the rotation that aligns the Pseudo-Earth Fixed (PEF) frame with the True Equator Mean Equinox (TEME) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233].

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the PEF frame with the TEME frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rPEFtoTOD_fk5Function
rPEFtoTOD_fk5([T,] JD_UT1::Number, JD_TT::Number [, δΔψ_1980::Number])

Compute the rotation that aligns the Pseudo-Earth Fixed (PEF) frame with the True of Date (TOD) frame at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory. Notice that one can provide correction for the nutation in longitude (δΔψ_1980) [rad] that is usually obtained from IERS EOP Data (see get_iers_eop).

The Julian Day in UT1 is used to compute the Greenwich Mean Sidereal Time (GMST) (see JDtoGMST), whereas the Julian Day in Terrestrial Time is used to compute the nutation in the longitude. Notice that the Julian Day in UT1 and in Terrestrial Time must be equivalent, i.e. must be related to the same instant. This function does not check this.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the PEF frame with the TOD frame. The rotation representation is selected by the optional parameter T.

Remarks

The Pseudo-Earth Fixed (PEF) frame is rotated into the True of Date (TOD) frame considering the 1980 IAU Theory of Nutation. The IERS EOP corrections must be added if one wants to make the rotation consistent with the Geocentric Celestial Reference Systems (GCRS).

source
SatelliteToolbox.rTEMEtoGCRFFunction
rTEMEtoGCRF([T,] JD_TT::Number [, δΔϵ_1980::Number = 0, δΔψ_1980::Number = 0])

Compute the rotation that aligns the True Equator Mean Equinox (TEME) frame with the Geocentric Celestial Reference Frame (GCRF) at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233]. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TEME frame with the GCRF frame. The rotation representation is selected by the optional parameter T.

Remarks

The EOP data related to the nutation of the obliquity (δΔϵ_1980) and the nutation of the longitude (δΔψ_1980) can be omitted. In this case, the GCRF frame is what is usually called J2000 reference frame.

source
SatelliteToolbox.rTEMEtoMODFunction
rTEMEtoMOD([T,] JD_TT::Number [, δΔϵ_1980::Number = 0, δΔψ_1980::Number = 0])

Compute the rotation that aligns the True Equator Mean Equinox (TEME) frame with the Mean of Date (MOD) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233]. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TEME frame with the MOD frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rTEMEtoPEFMethod
rTEMEtoPEF([T,] JD_TT::Number)

Compute the rotation that aligns the True Equator Mean Equinox (TEME) frame with the Pseudo-Earth Fixed (PEF) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233].

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TEME frame with the PEF frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rTEMEtoTODFunction
rTEMEtoTOD([T,] JD_TT::Number [, δΔϵ_1980::Number = 0, δΔψ_1980::Number = 0])

Compute the rotation that aligns the True Equator Mean Equinox (TEME) frame with the True of Date (TOD) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233]. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TEME frame with the TOD frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.rTIRStoCIRS_iau2006Method
rTIRStoCIRS_iau2006([T::Type,] JD_UT1::Number)

Compute the rotation that aligns the Terrestrial Intermediate Reference System (TIRS) with the Celestial Intermediate Reference System (CIRS) at the Julian Day JD_UT1 [UT1]. This algorithm uses the IAU-2006 theory.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TIRS frame with the CIRS frame. The rotation representation is selected by the optional parameter T.

Remarks

The reference frames TIRS and CIRS are separated by a rotation about the Z-axis of the Earth Rotation Angle, which is the angle between the Conventional International Origin (CIO) and the Terrestrial Intermediate Origin (TIO) [1]. The latter is a reference meridian on Earth that is located about 100m away from Greenwich meridian along the equator of the Celestial Intermediate Pole (CIP) [1].

source
SatelliteToolbox.rTIRStoERS_iau2006Function
rTIRStoERS_iau2006([T::Type,] JD_UT1::Number, JD_TT::Number, δΔΨ_2000::Number = 0)

Compute the rotation that aligns the Terrestrial Intermediate Reference System (TIRS) with the Earth Reference System (ERS) at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

Notice that one can provide corrections for the nutation in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop and dEps_dPsi). This corrections are related to Free Core Nutation (FCN) that models the effect of a liquid Earth core.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TIRS frame with the ERS frame. The rotation representation is selected by the optional parameter T.

Remarks

The reference frames TIRS and ERS are separated by a rotation about the Z-axis of the Greenwhich apparent sidereal angle (GAST). This angle is computed using the IAU-2006 theory, which consist of obtaining the Earth Rotation Angle (ERA) and subtracting the result of the Equation of Origins (EO).

source
SatelliteToolbox.rTIRStoITRF_iau2006Method
rTIRStoITRF_iau2006([T::Type,] JD_TT::Number, x_p::Number, y_p::Number)

Compute the rotation that aligns the Terrestrial Intermediate Reference System (TIRS) with the International Terrestrial Reference Frame (ITRF) considering the polar motion represented by the angles x_p [rad] and y_p [rad] that are obtained from IERS EOP Data (see get_iers_eop).

x_p is the polar motion displacement about X-axis, which is the IERS Reference Meridian direction (positive south along the 0˚ longitude meridian). y_p is the polar motion displacement about Y-axis (90˚W or 270˚E meridian).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TIRS frame with the ITRF frame. The rotation representation is selected by the optional parameter T.

Remarks

The ITRF is defined based on the International Reference Pole (IRP), which is the location of the terrestrial pole agreed by international committees [1]. The Terrestrial Intermediate Reference Frame (TIRS), on the other hand, is defined based on the Earth axis of rotation, or the Celestial Intermediate Pole (CIP). Hence, TIRS XY-plane contains the True Equator. Furthermore, since the recovered latitude and longitude are sensitive to the CIP, then it should be computed considering the TIRS frame.

The TIRS and PEF (IAU-76/FK5) are virtually the same reference frame, but according to [1] it is convenient to separate the names as the exact formulae differ.

source
SatelliteToolbox.rTIRStoMOD_iau2006Function
rTIRStoMOD_iau2006([T::Type,] JD_UT1::Number, JD_TT::Number, δΔϵ_2000::Number = 0, δΔΨ_2000::Number = 0)

Compute the rotation that aligns the Terrestrial Intermediate Reference System (TIRS) with the Mean of Date (MOD) reference frame at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-2006 theory.

Notice that one can provide corrections for the nutation in obliquity (δΔϵ_2000) and in longitude (δΔψ_2000) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop and dEps_dPsi). This corrections are related to Free Core Nutation (FCN) that models the effect of a liquid Earth core.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TIRS frame with the ERS frame. The rotation representation is selected by the optional parameter T.

Remarks

This composed rotation TIRS <=> ERS <=> MOD is implemented as a new function because the single rotations TIRS <=> ERS and ERS <=> MOD call the function nutation_eo, which has a high computational burden. In this case, the composed algorithm is about 2x faster than calling those function separately.

source
SatelliteToolbox.rTODtoMOD_fk5Function
rTODtoMOD_fk5([T,] JD_TT::Number [, δΔϵ_1980::Number, δΔψ_1980::Number])

Compute the rotation that aligns the True of Date (TOD) frame with the Mean of Date (MOD) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TOD frame with the MOD frame. The rotation representation is selected by the optional parameter T.

Remarks

The True of Date (TOD) frame is rotated into the Mean of Date (MOD) frame considering the 1980 IAU Theory of Nutation. The IERS EOP corrections must be added if one wants to make the rotation consistent with the Geocentric Celestial Reference Systems (GCRS).

source
SatelliteToolbox.rTODtoPEF_fk5Function
rTODtoPEF_fk5([T,] JD_UT1::Number, JD_TT::Number [, δΔψ_1980::Number])

Compute the rotation that aligns the True of Date (TOD) frame with the Pseudo-Earth Fixed (PEF) frame at the Julian Day JD_UT1 [UT1] and JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory. Notice that one can provide correction for the nutation in longitude (δΔψ_1980) [rad] that is usually obtained from IERS EOP Data (see get_iers_eop).

The Julian Day in UT1 is used to compute the Greenwich Mean Sidereal Time (GMST) (see JDtoGMST), whereas the Julian Day in Terrestrial Time is used to compute the nutation in the longitude. Notice that the Julian Day in UT1 and in Terrestrial Time must be equivalent, i.e. must be related to the same instant. This function does not check this.

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TOD frame with the PEF frame. The rotation representation is selected by the optional parameter T.

Remarks

The True of Date (TOD) frame is rotated into the Pseudo-Earth Fixed (PEF) frame considering the 1980 IAU Theory of Nutation. The IERS EOP corrections must be added if one wants to make the rotation consistent with the Geocentric Celestial Reference Systems (GCRS).

source
SatelliteToolbox.rTODtoTEMEFunction
rTODtoTEME([T,] JD_TT::Number [, δΔϵ_1980::Number = 0, δΔψ_1980::Number = 0])

Compute the rotation that aligns the True of Date (TOD) frame with the True Equator Mean Equinox (TEME) frame at the Julian Day JD_TT [Terrestrial Time]. This algorithm uses the IAU-76/FK5 theory and TEME definition in [1, p. 233]. Notice that one can provide corrections for the nutation in obliquity (δΔϵ_1980) [rad] and in longitude (δΔψ_1980) [rad] that are usually obtained from IERS EOP Data (see get_iers_eop).

The rotation type is described by the optional variable T. If it is DCM, then a DCM will be returned. Otherwise, if it is Quaternion, then a Quaternion will be returned. In case this parameter is omitted, then it falls back to DCM.

Returns

The rotation that aligns the TOD frame with the TEME frame. The rotation representation is selected by the optional parameter T.

source
SatelliteToolbox.read_iers_eopFunction
read_iers_eop(filename::String, data_type::Symbol = :IAU1980)

Read IERS EOP Data from the file filename. The user must specify if the data is related to the model IAU 1980 (data_type = :IAU1980), which is the default, or to the model IAU 2000A (data_type = :IAU2000A).

Returns

A structure (EOPData_IAU1980 or EOPData_IAU2000A, depending on data_type) with the interpolations of the EOP parameters. Notice that the interpolation indexing is set to the Julian Day.

Remarks

The input file must be exactly the same as provided by IERS. One can download it using the following commands:

  • IAU 1980

      curl -O https://datacenter.iers.org/data/latestVersion/223_EOP_C04_14.62-NOW.IAU1980223.txt
      wget https://datacenter.iers.org/data/latestVersion/223_EOP_C04_14.62-NOW.IAU1980223.txt
  • IAU 2000A

      curl -O https://datacenter.iers.org/data/latestVersion/224_EOP_C04_14.62-NOW.IAU2000A224.txt
      wget https://datacenter.iers.org/data/latestVersion/224_EOP_C04_14.62-NOW.IAU2000A224.txt
source
SatelliteToolbox.rv_to_keplerFunction
rv_to_kepler(r_i::AbstractVector, v_i::AbstractVector, t::Number = 0)
rv_to_kepler(x::Number, y::Number, z::Number, vx::Number, vy::Number, vz::Number, t::Number = 0)

Convert a Cartesian representation (position vector r_i [m] and velocity vector v_i [m/s²]) to the Keplerian elements. Optionally, the user can specify the epoch of the returned elements using the parameter t. It it is omitted, then it default to 0.

The input vectors can also be passed component by component:

r_i = [x,   y,  z]
v_i = [vx, vy, vz]

Returns

An instance of the structure KeplerianElements [SI units].

Remarks

The special cases are treated as follows:

  • Circular and equatorial: the right ascension of the ascending node and the argument of perigee are set to 0. Hence, the true anomaly is equal to the true longitude.
  • Elliptical and equatorial: the right ascension of the ascending node is set to 0. Hence, the argument of perigee is equal to the longitude of periapsis.
  • Circular and inclined: the argument of perigee is set to 0. Hence, the true anomaly is equal to the argument of latitude.

References

The algorithm was adapted from [1].

source
SatelliteToolbox.rv_to_mean_elements_sgp4Method
rv_to_mean_elements_sgp4(vJD::AbstractVector{T}, vr::AbstractVector{Tv}, vv::AbstractVector{Tv}, W = I; estimate_bstar::Bool = true, mean_elements_epoch::Symbol = :end, max_it::Int = 50, sgp4_gc = sgp4_gc_wgs84, atol::Number = 2e-4, rtol::Number = 2e-4) where {T,Tv<:AbstractVector}

Compute the mean elements for SGP4 based on the position vr and velocity vectors vr represented in TEME reference frame. The epoch of those measurements [Julian Day] must be in vJD.

The matrix W defined the weights for the least-square algorithm.

Keywords

  • estimate_bstar: If true, then the BSTAR parameters of the TLE will be estimated.
  • mean_elements_epoch: If it is :end, the epoch of the mean elements will be equal to the last value in vJD. Otherwise, if it is :begin, the epoch will be selected as the first value in vJD.
  • max_it: The maximum allowed number of iterations.
  • sgp4_gc: SPG4 constants (see SGP4_GravCte).
  • atol: The tolerance for the absolute value of the residue. If, at any iteration, the residue is lower than atol, then the iterations stop.
  • rtol: The tolerance for the relative difference between the residues. If, at any iteration, the relative difference between the residues in two consecutive iterations is lower than rtol, then the iterations stop.

Returns

  • The epoch of the elements [Julian Day].
  • The mean elements for SGP4 algorithm:
    • Semi-major axis [m];
    • Eccentricity [ ];
    • Inclination [rad];
    • Right ascension of the ascending node [rad];
    • Argument of perigee [rad];
    • True anomaly [rad];
    • BSTAR (0 if estimate_bstar is false).
  • The covariance matrix of the mean elements estimation.
source
SatelliteToolbox.rv_to_tleMethod
rv_to_tle(args...; name::String = "UNDEFINED", sat_num::Int = 9999, classification::Char = 'U', int_designator = "999999", elem_set_number::Int = 0, rev_num, kwargs...)

Convert a set of position and velocity vectors represented in TEME reference frame to a TLE. The arguments args and keywords kwargs are the same as those described in the function rv_to_mean_elements_sgp4.

Additionally, the user can specify some parameters of the generated TLE.

This function returns the TLE and the covariance of the estimated elements (state vector).

source
SatelliteToolbox.satellite_check_BrazilMethod
satellite_check_Brazil(lat::Number, lon::Number)

Verify if a point described by latitude lat [rad] and longitude lon [rad] is inside Brazil. Returns true if the point is inside Brazil, of false otherwise.

Remarks

This function was based on the algorithm sent by Renato Branco to Ronan Arraes by e-mail at 2016-02-16.

source
SatelliteToolbox.satellite_lighting_conditionMethod
satellite_lighting_condition(r_i::AbstractVector, s_i::AbstractVector)

Compute the satellite lighting condition given the Sun unitary vector s_i [m] and the satellite position vector r_i [m].

Returns

  • SAT_LIGHTING_SUNLIGHT: Satellite is under sunlight.
  • SAT_LIGHTING_PENUMBRA: Satellite is at penumbra region.
  • SAT_LIGHTING_UMBRA: Satellite is at umbra region.
source
SatelliteToolbox.satellite_position_iMethod
satellite_position_i(a::Number, e::Number, i::Number, RAAN::Number, w::Number, f::Number)

Compute the satellite position in the Earth-Centered Inertial (ECI) reference frame given the orbital elements a, e, i, RAAN, w, and f.

Notice that the ECI frame used will be the same as the frame of the orbital elements.

Args

  • a: Semi-major axis.
  • e: Eccentricity.
  • i: Inclination [rad].
  • RAAN: Right ascension of the ascending node [rad].
  • w: Argument of perigee [rad].
  • f: True anomaly [rad].

Returns

  • The satellite position vector represented in the ECI reference frame.
  • The unit vector perpendicular to the satellite position vector that lies on the orbit plane represented in the ECI reference frame.

Remarks

The satellite position vector will have the same unit of the semi-major axis.

source
SatelliteToolbox.satellite_sun_angle_earth_pointingFunction
satellite_sun_angle_earth_pointing(JD0::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, numDays::Integer, N::AbstractVector, step::Number = 0.1*pi/180.0)

Compute the Sun angle on a satellite surface for an Earth-pointing mission.

Args

  • JD0: Initial instant for the analysis [Julian day].
  • a: Semi-major axis of the orbit [m].
  • e: Orbit eccentricity.
  • i: Orbit inclination [rad].
  • w: Argument of perigee [rad].
  • RAAN: Right ascension of the ascending node at JD0 [rad].
  • numDays: Number of days for the analysis.
  • N: Vector normal to the surface represented in the body reference frame.
  • meanAnomaly: (OPTIONAL) If true, compute using angular steps in the mean anomaly instead of in the orbit latitude (Default: false).
  • step: (OPTIONAL) Mean anomaly step (Default: 0.1 deg).

Returns

A matrix containing the Sun angle for each position in orbit for each day.

NOTE: if the Sun angle is larger than 90 deg or if the satellite is in eclipse, then NaN is returned in the matrix.

Remarks

The body reference frame is defined as:

  • Z axis points towards the center of Earth;
  • Y axis points towards the negative direction of orbit normal;
  • X axis completes the right-hand reference frame.

If the mean anomaly is used, then the average value of the output is the average sun radiation received by the satellite surface, because every angular steps have a fixed time interval.

If the mean anomaly is used, then the angle interval is [0, 2π]. Otherwise, the angle interval is [-π,π].

source
SatelliteToolbox.satellite_sun_angle_earth_pointingFunction
satellite_sun_angle_earth_pointing(JD0::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, numDays::Integer, fN_k::Function, meanAnomaly::Bool = false, step::Number = 0.1*pi/180.0)

Compute the Sun angle on a satellite surface for an Earth-pointing mission.

Args

  • JD0: Initial instant for the analysis [Julian day].
  • a: Semi-major axis of the orbit [m].
  • e: Orbit eccentricity.
  • i: Orbit inclination [rad].
  • w: Argument of perigee [rad].
  • RAAN: Right ascension of the ascending node at JD0 [rad].
  • numDays: Number of days for the analysis.
  • fN_k: Function f(s_b) that describes the solar panel normal at each k-th sampling step. Notice that s_b is the Sun vector represented in the body coordinate frame.
  • meanAnomaly: (OPTIONAL) If true, compute using angular steps in the mean anomaly instead of in the orbit latitude (Default: false).
  • step: (OPTIONAL) Mean anomaly step (Default: 0.1 deg).

Returns

A matrix containing the sun angle [rad] for each position in orbit for each day.

NOTE: if the Sun angle is larger than 90 deg or if the satellite is in eclipse, then NaN is returned in the matrix.

Remarks

The body reference frame is defined as:

  • Z axis points towards the center of Earth;
  • Y axis points towards the negative direction of orbit normal;
  • X axis completes the right-hand reference frame.

If the mean anomaly is used, then the average value of the output is the average sun radiation received by the satellite surface, because every angular steps have a fixed time interval.

If the mean anomaly is used, then the angle interval is [0, 2π]. Otherwise, the angle interval is [-π,π].

source
SatelliteToolbox.satellite_sun_radiation_earth_pointingFunction
satellite_sun_radiation_earth_pointing(JD0::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, numDays::Integer, N::Vector, meanAnomaly::Bool = false, step::Number = 0.1*pi/180.0)

Compute the Sun radiation on a satellite surface for an Earth-pointing mission.

Args

  • JD0: Initial instant for the analysis [Julian day].
  • a: Semi-major axis of the orbit [m].
  • e: Orbit eccentricity.
  • i: Orbit inclination [rad].
  • w: Argument of perigee [rad].
  • RAAN: Right ascension of the ascending node at JD0 [rad].
  • numDays: Number of days for the analysis.
  • N: Vector normal to the surface represented in the body reference frame.
  • meanAnomaly: (OPTIONAL) If true, compute using angular steps in the mean anomaly instead of in the orbit latitude (Default: false).
  • step: (OPTIONAL) Mean anomaly step (Default: 0.1 deg).

Returns

A matrix containing the Sun radiation [W/m²] for each position in orbit for each day.

NOTE: if the Sun angle is larger than 90 deg or if the satellite is in eclipse, then NaN is returned in the matrix.

Remarks

The body reference frame is defined as:

  • Z axis points towards the center of Earth;
  • Y axis points towards the negative direction of orbit normal;
  • X axis completes the right-hand reference frame.

If the mean anomaly is used, then the average value of the output is the average sun radiation received by the satellite surface, because every angular steps have a fixed time interval.

If the mean anomaly is used, then the angle interval is [0, 2π]. Otherwise, the angle interval is [-π,π].

source
SatelliteToolbox.satellite_sun_radiation_earth_pointingFunction
satellite_sun_radiation_earth_pointing(JD0::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, numDays::Integer, fN_k::Function, meanAnomaly::Bool = false, step::Number = 0.1*pi/180.0)

Compute the Sun radiation on a satellite surface for an Earth-pointing mission.

Args

  • JD0: Initial instant for the analysis [Julian day].
  • a: Semi-major axis of the orbit [m].
  • e: Orbit eccentricity.
  • i: Orbit inclination [rad].
  • w: Argument of perigee [rad].
  • RAAN: Right ascension of the ascending node at JD0 [rad].
  • numDays: Number of days for the analysis.
  • fN_k: Function f(s_b) that describes the solar panel normal at each k-th sampling step. Notice that s_b is the Sun vector represented in the body coordinate frame.
  • meanAnomaly: (OPTIONAL) If true, compute using angular steps in the mean anomaly instead of in the orbit latitude (Default: false).
  • step: (OPTIONAL) Mean anomaly step (Default: 0.1 deg).

Returns

A matrix containing the Sun radiation [W/m²] for each position in orbit for each day.

NOTE: if the Sun angle is larger than 90 deg or if the satellite is in eclipse, then NaN is returned in the matrix.

Remarks

The body reference frame is defined as:

  • Z axis points towards the center of Earth;
  • Y axis points towards the negative direction of orbit normal;
  • X axis completes the right-hand reference frame.

If the mean anomaly is used, then the average value of the output is the average sun radiation received by the satellite surface, because every angular steps have a fixed time interval.

If the mean anomaly is used, then the angle interval is [0, 2π]. Otherwise, the angle interval is [-π,π].

source
SatelliteToolbox.satellite_sun_radiation_earth_pointing_meanFunction
satellite_sun_radiation_earth_pointing_mean(JD0::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, numDays::Integer, fN_k::Function, step::Number = 0.1*pi/180.0)

Compute the mean Sun radiation on a satellite surface for an Earth-pointing mission.

Args

  • JD0: Initial instant for the analysis [Julian day].
  • a: Semi-major axis of the orbit [m].
  • e: Orbit eccentricity.
  • i: Orbit inclination [rad].
  • w: Argument of perigee [rad].
  • RAAN: Right ascension of the ascending node at JD0 [rad].
  • numDays: Number of days for the analysis.
  • fN_k: Function f(s_b) that describes the solar panel normal at each k-th sampling step. Notice that s_b is the Sun vector represented in the body coordinate frame.
  • meanAnomaly: (OPTIONAL) If true, compute using angular steps in the mean anomaly instead of in the orbit latitude (Default: false).
  • step: (OPTIONAL) Mean anomaly step (Default: 0.1 deg).

Returns

The mean Sun radiation on a surface [W/m²].

Remarks

For more details, see satellitesunradiationearthpointing.

source
SatelliteToolbox.satellite_sun_radiation_earth_pointing_meanFunction
satellite_sun_radiation_earth_pointing_mean(JD0::Number, a::Number, e::Number, i::Number, RAAN::Number, w::Number, numDays::Integer, N::AbstractVector, step::Number = 0.1*pi/180.0)

Compute the mean Sun radiation on a satellite surface for an Earth-pointing mission.

Args

  • JD0: Initial instant for the analysis [Julian day].
  • a: Semi-major axis of the orbit [m].
  • e: Orbit eccentricity.
  • i: Orbit inclination [rad].
  • w: Argument of perigee [rad].
  • RAAN: Right ascension of the ascending node at JD0 [rad].
  • numDays: Number of days for the analysis.
  • N: Vector normal to the surface represented in the body reference frame.
  • meanAnomaly: (OPTIONAL) If true, compute using angular steps in the mean anomaly instead of in the orbit latitude (Default: false).
  • step: (OPTIONAL) Mean anomaly step (Default: 0.1 deg).

Returns

The mean Sun radiation on a surface [W/m²].

Remarks

For more details, see satellitesunradiationearthpointing.

source
SatelliteToolbox.sim_RAAN_J2Method
sim_RAAN_J2(a::Number, e::Number, i::Number, RAAN_0::Number, numDays::Integer)

Simulate the RAAN of an orbit with semi-major axis a [m], eccentricity e, inclination i [rad], and initial RAAN RAAN_0 [rad] considering J2 perturbations. The analysis is performed for numDays days.

Returns

A numDays × 2 matrix in which the i-th line is:

| day | RAAN (0,2π) [rad] |
source
SatelliteToolbox.step!Function
step!(orbp::OrbitPropagator{T}, Δt::Number)

Propagate the orbit in orbp by Δt [s] using the algorithm of orbp. The new parameters will be written in orbp.

Returns

  • The Keplerian elements represented in the inertial frame after the step (see KeplerianElements) [SI units].
  • The position vector represented in the inertial frame after the step [m].
  • The velocity vector represented in the inertial frame after the step [m].
source
SatelliteToolbox.sun_position_iMethod
sun_position_i(JD::Number)

Compute the Sun position represented in the Mean Equinox of Date (MOD) at the Julian Day JD. The algorithm was adapted from [3, p. 277-279].

source
SatelliteToolbox.sun_velocity_iMethod
sun_velocity_i(JD::Number)

Compute the Sun velocity represented in the Mean Equinox of Date (MOD) at the Julian Day JD. The algorithm was obtained by computing the time derivative of the Sun position in [3, p. 277-279].

source
SatelliteToolbox.svECEFtoECEFMethod
svECEFtoECEF(sv::OrbitStateVector, args...)

Convert the orbit state vector sv from an ECEF frame to another ECEF frame. The arguments args... must match those of the function rECEFtoECEF wihtout the rotation representation.

source
SatelliteToolbox.svECEFtoECIMethod
svECEFtoECI(sv::OrbitStateVector, ECEF, ECI, JD_UTC [, eop_data])

Convert the orbit state vector sv from the Earth-Centered, Earth-Fixed (ECEF) reference frame ECEF to the Earth-Centered Inertial (ECI) reference frame at the Julian day JD_UTC [UTC]. The eop_data may be required depending on the selection of the input and output reference system. For more information, see the documentation of the function rECEFtoECI.

Info

It is assumed that the input velocity and acceleration in sv are obtained by an observer on the ECEF frame. Thus, the output will contain the velocity and acceleration as measured by an observer on the ECI frame.

source
SatelliteToolbox.svECItoECEFMethod
svECItoECEF(sv::OrbitStateVector, ECI, ECEF, JD_UTC [, eop_data])

Convert the orbit state vector sv from the Earth-Centered Inertial (ECI) reference frame ECI to the Earth-Centered, Earth-Fixed (ECEF) reference frame at the Julian day JD_UTC [UTC]. The eop_data may be required depending on the selection of the input and output reference system. For more information, see the documentation of the function rECItoECEF.

Info

It is assumed that the input velocity and acceleration in sv are obtained by an observer on the ECI frame. Thus, the output will contain the velocity and acceleration as measured by an observer on the ECEF frame.

source
SatelliteToolbox.svECItoECIMethod
svECItoECI(sv::OrbitStateVector, args...)

Convert the orbit state vector sv from an ECI frame to another ECI frame. The arguments args... must match those of the function rECItoECI wihtout the rotation representation.

source
SatelliteToolbox.sv_to_keplerMethod
sv_to_kepler(sv::OrbitStateVector)

Convert the state vector sv to Keplerian elements represented by an instance of the structure KeplerianElements.

source
SatelliteToolbox.swath_widthMethod
swath_width(h::real, HalfFOV::real)

Compute the swath width given the orbit altitude and the half FOV.

Args

  • h: Orbit altitude [m].
  • HalfFOV: Half field of view [rad].

Returns

The swath width [m].

source
SatelliteToolbox.twobody!Method
twobody!(tbd::TwoBody_Structure, t::Number)

Propagate the orbit defined in tbd (see TwoBody_Structure) until the time t [s]. Notice that the values in tbd will be modified.

Returns

  • The position vector represented in the inertial frame at time t [m].
  • The velocity vector represented in the inertial frame at time t [m/s]

Remarks

The inertial frame in which the output is represented depends on which frame it was used to generate the orbit parameters.

source
SatelliteToolbox.twobody_initMethod
twobody_init(epoch::Number, a_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, f_0::Number; μ::T = m0) where T

Initialize the data structure of two body orbit propagator algorithm.

Args

  • epoch: Epoch of the initial mean orbital elements [s].
  • a_0: Initial mean semi-major axis [m].
  • e_0: Initial mean eccentricity.
  • i_0: Initial mean inclination [rad].
  • Ω_0: Initial mean right ascension of the ascending node [rad].
  • ω_0: Initial mean argument of perigee [rad].
  • f_0: Initial mean true anomaly.

Keywords

  • μ: Standard gravitational parameter of the central body [m^3/s^2]. (Default = m0)

Returns

The structure TwoBody_Structure with the initialized parameters.

source
SatelliteToolbox.@_keyword_foundMacro
@_keyword_found(keyword, keywords_found, current_line)

Check if the keyword exists in the list keywords_found. If true, then throw an error indicating that the problem occurred on the current_line.

source
SatelliteToolbox.@check_orbitMacro
@check_orbit(a, e)

Verify if the orbit with semi-major axis a [m] and eccentricity e is valid. This macro throws an exception if the orbit is not valid.

Return true is the orbit is valid, and false otherwise.

source
SatelliteToolbox.SGP4.SGP4_GravCteType
SGP4_GravCte{T<:Real}

Gravitational constants for SGP4.

Fields

  • R0: Earth equatorial radius [km].
  • XKE: √GM [er/s]^(3/2).
  • J2: The second gravitational zonal harmonic of the Earth.
  • J3: The thrid gravitational zonal harmonic of the Earth.
  • J4: The fourth gravitational zonal harmonic of the Earth.
source
SatelliteToolbox.SGP4.J2000toGMSTMethod
J2000toGMST(J2000_UT1::Number)

Compute the Greenwich Mean Sideral Time (GMST) [rad] given the instant J2000_UT1 in J2000.0 reference [UT1].

Remarks

Based on algorithm in 2, accessed at 2015-12-01.

source
SatelliteToolbox.SGP4.JDtoGMSTMethod
JDtoGMST(JD_UT1::Number)

Compute the Greenwich Mean Sideral Time (GMST) [rad] for the Julian Day JD_UT1 [UT1].

Remarks

Based on algorithm in [1, pp. 188].

source
SatelliteToolbox.SGP4.dsinitMethod
dsinit(epoch::T, nll_0::T, all_0::T, e_0::T, i_0::T, Ω_0::T, ω_0::T, M_0::T, dotM::T, dotω::T, dotΩ::T) where T<:Number

Initialize the deep space structure. This function performs the initial computations and save the values at an instance of the structure SGP4_DeepSpace. Those will be used when calling the functions dsper! and dpsec!.

Args

  • epoch: Epoch of the initial orbit [Julian Day].
  • nll_0: Initial mean motion [rad/min].
  • all_0: Initial semi-major axis [ER].
  • e_0: Initial eccentricity.
  • i_0: Initial inclination [rad].
  • Ω_0: Initial right ascencion of the ascending node [rad].
  • ω_0: Initial argument of perigee [rad].
  • M_0: Initial mean motion [rad].
  • dotM: Time-derivative of the mean motion [rad/min].
  • dotω: Time-derivative of the argument of perigee [rad/min].
  • dotΩ: Time-derivative of the RAAN [rad/min].

Returns

An instance of the structure SGP4_DeepSpace with the initalized values.

source
SatelliteToolbox.SGP4.dsper!Method
dsper!(sgp4_ds::SGP4_DeepSpace{T}, e_k::T, i_k::T, Ω_k::T, ω_k::T, M_k::T, Δt:Number) where T<:Number

Compute the effects caused by Lunar-Solar periodics.

Notice that the values in the structure sgp4_ds will be modified.

Args

  • sgp4_ds: Deep space structure (see SGP4_DeepSpace).
  • e_k: Current eccentricity.
  • i_k: Current inclination [rad].
  • Ω_k: Current right ascension of the ascending node [rad].
  • ω_k: Current argument of perigee [rad].
  • M_k: Current mean anomaly [rad].
  • Δt: Time interval since the epoch [min].

Returns

The following elements perturbed by lunar-solar periodics.

  • Eccentricity.
  • Inclination [rad].
  • Right ascension of the ascending node [rad].
  • Argument of perigee [rad].
  • Mean anomaly [rad].
source
SatelliteToolbox.SGP4.dssec!Method
dssec!(sgp4_ds::SGP4_DeepSpace{T}, nll_0::T, e_0::T, i_0::T, ω_0::T, Ω_k::T, ω_k::T, M_k::T, dotω::T, Δt::Number) where T<:Number

Compute the secular effects.

Notice that the values in the structure sgp4_ds will be modified.

Args

  • sgp4_ds: Deep space structure (see SGP4_DeepSpace).
  • nll_0: Initial mean motion [rad/min].
  • e_0: Initial eccentricity.
  • i_0: Initial inclination [rad].
  • ω_0: Initial argument of perigee [rad].
  • Ω_k: Current right ascension of the ascending node [rad].
  • ω_k: Current argument of perigee [rad].
  • M_k: Current mean anomaly [rad].
  • dotω: Time-derivative of the argument of perigee [rad/min].
  • Δt: Time interval since the epoch [min].

Returns

The following elements perturbed by the secular effects:

  • Mean motion [rad/min].
  • Eccentricity.
  • Inclination [rad].
  • Right ascension of the ascending node [rad].
  • Argument of perigee [rad].
  • Mean anomaly [rad].
source
SatelliteToolbox.SGP4.sgp4!Method
sgp4!(sgp4d::SGP4_Structure{T}, t::Number) where T

Propagate the orbit defined in sgp4d (see SGP4_Structure) until the time t [min]. Notice that the values in sgp4d will be modified.

Returns

  • The position vector represented in TEME frame at time t [km].
  • The velocity vector represented in TEME frame at time t [km/s].
source
SatelliteToolbox.SGP4.sgp4Method
sgp4(Δt, args...)

Function that initialize the SGP4 structure and propagate the orbit until the time Δt.

Returns

  • The position vector [km].
  • The velocity vector [km/s].
  • The SGP4 structure (see SGP4_Structure).
source
SatelliteToolbox.SGP4.sgp4_initFunction
sgp4_init(spg4_gc::SGP4_GravCte{T}, epoch::Number, n_0::Number, e_0::Number, i_0::Number, Ω_0::Number, ω_0::Number, M_0::Number, bstar::Number) where T
sgp4_init(tle::TLE, sgp4_gc::SGP4_GravCte = sgp4_gc_wgs84)

Initialize the data structure of SGP4 algorithm.

Args

  • spg4_gc: SPG4 gravitational constants (see SGP4_GravCte).
  • epoch: Epoch of the orbital elements [Julian Day].
  • n_0: SGP type "mean" mean motion at epoch [rad/min].
  • e_0: "Mean" eccentricity at epoch.
  • i_0: "Mean" inclination at epoch [rad].
  • Ω_0: "Mean" longitude of the ascending node at epoch [rad].
  • ω_0: "Mean" argument of perigee at epoch [rad].
  • M_0: "Mean" mean anomaly at epoch [rad].
  • bstar: Drag parameter (B*).
  • tle: TLE to initialize the SPG4 (see TLE).

Returns

The structure SGP4_Structure with the initialized parameters.

source
SatelliteToolbox.SatelliteToolboxTLE.TLEType
TLE

This structure contains the same elements of the TLE with the same units.

Fields

  • name: Name of the satellite.

First line

  • sat_num: Satellite number.
  • classification: Classification ('U', 'C', or 'S').
  • int_designator: International designator.
  • epoch_year: Epoch year (two digits).
  • epoch_day: Epoch day (day + fraction of the day).
  • epoch: The epoch represented in Julian Day.
  • dn_o2: 1st time derivative of mean motion / 2 [rev/day²].
  • ddn_o6: 2nd time derivative of mean motion / 6 [rev/day³].
  • bstar: B* drag term.
  • elem_set_number: Element set number.
  • checksum_l1: Checksum of the line 1 (modulo 10).

Second line

  • i: Inclination [deg].
  • Ω: Right ascension of the ascending node [deg].
  • e: Eccentricity.
  • ω: Argument of perigee [deg].
  • M: Mean anomaly [deg].
  • n: Mean motion [rev/day].
  • rev_num: Revolution number at epoch [rev].
  • checksum_l2: Checksum of the line 2 (modulo 10).
source
SatelliteToolbox.SatelliteToolboxTLE.compute_checksumMethod
compute_checksum(str::AbstractString)

Compute the checksum of the line str modulo 10.

The algorithm is simple: add all the numbers in the line, ignoring letters, spaces, periods, and plus signs, but assigning +1 to the minus signs. The checksum is the remainder of the division by 10.

source
SatelliteToolbox.SatelliteToolboxTLE.read_tleFunction
read_tle(tle_filename::String, verify_checksum::Bool = true)

Read the TLEs in the file tle_filename and return an array of TLE with the parsed TLEs.

If verify_checksum if true, then the checksum of both TLE lines will be verified. Otherwise, the checksum will not be checked. If verify_checksum is omitted, then it defaults to true.

source
SatelliteToolbox.SatelliteToolboxTLE.read_tle_from_stringFunction
read_tle_from_string(tles::String, verify_checksum::Bool = true)
read_tle_from_string(tle_l1::String, tle_l2::String, verify_checksum::Bool = false)

Parse a set of TLEs in the string tles or one TLE with first line tle_l1 and second line tle_l2. This function returns an array of TLE with the parsed TLEs.

If verify_checksum if true, then the checksum of both TLE lines will be verified. Otherwise, the checksum will not be checked. If verify_checksum is omitted, then it defaults to true.

source
SatelliteToolbox.SatelliteToolboxTLE.tle_to_strMethod
tle_to_str(tle::TLE; recompute_checksum = true, bstar_exp_le = true)

Convert the TLE tle to a string. If recompute_checksum is true, then the checksums in tle will be ignored and they will be computed considering the TLE data.

The keyword bstar_exp_le selects if the BSTAR exponent signal will be + or - when BSTAR is zero. This is required for the tests because it is not standardized in TLE generation. If it is true, then the exponent signal will be - when BSTAR is zero.

source
SatelliteToolbox.SatelliteToolboxTLE.@tle_strMacro
@tle_str(str)

Parse a set of TLEs in the string str and return as an array of TLE. This version verifies the checksum of the TLE. If the checksum verification is not desired, see @tlenc_str.

Example

julia> tles = tle"""
       CBERS 4
       1 40336U 14079A   18166.15595376 -.00000014  00000-0  10174-4 0  9993
       2 40336  98.4141 237.7928 0001694  75.7582 284.3804 14.35485112184485
       SCD 1
       1 22490U 93009B   18165.62596833  .00000225  00000-0  11410-4 0  9991
       2 22490  24.9690 231.7852 0042844 200.7311 292.7198 14.44524498338066
       SCD 2
       1 25504U 98060A   18165.15074951  .00000201  00000-0  55356-5 0  9994
       2 25504  24.9961  80.1303 0017060 224.4822 286.6438 14.44043397 37312
       """
source
SatelliteToolbox.SatelliteToolboxTLE.@tlenc_strMacro
@tlenc_str(str)

Parse a set of TLEs in the string str and return as an array of TLE. This version does not verify the checksum of the TLE. If the checksum verification is required, see @tle_str.

Example

julia> tles = tlenc"""
       CBERS 4
       1 40336U 14079A   18166.15595376 -.00000014  00000-0  10174-4 0  9993
       2 40336  98.4141 237.7928 0001694  75.7582 284.3804 14.35485112184485
       SCD 1
       1 22490U 93009B   18165.62596833  .00000225  00000-0  11410-4 0  9991
       2 22490  24.9690 231.7852 0042844 200.7311 292.7198 14.44524498338066
       SCD 2
       1 25504U 98060A   18165.15074951  .00000201  00000-0  55356-5 0  9994
       2 25504  24.9961  80.1303 0017060 224.4822 286.6438 14.44043397 37312
       """
source
  • 1TEME is an of date frame.