Struct GroundStation

pub struct GroundStation {
    pub name: String,
    pub elevation_mask_deg: f64,
    pub latitude_deg: f64,
    pub longitude_deg: f64,
    pub height_km: f64,
    pub frame: Frame,
    pub integration_time: Option<Duration>,
    pub light_time_correction: bool,
    pub timestamp_noise_s: Option<StochasticNoise>,
    pub range_noise_km: Option<StochasticNoise>,
    pub doppler_noise_km_s: Option<StochasticNoise>,
}

GroundStation defines a two-way ranging and doppler station.

Fields

name: String

elevation_mask_deg: f64

in degrees

latitude_deg: f64

in degrees

longitude_deg: f64

in degrees

height_km: f64

in km

frame: Frame

integration_time: Option<Duration>

Duration needed to generate a measurement (if unset, it is assumed to be instantaneous)

light_time_correction: bool

Whether to correct for light travel time

timestamp_noise_s: Option<StochasticNoise>

Noise on the timestamp of the measurement

range_noise_km: Option<StochasticNoise>

Noise on the range data of the measurement

doppler_noise_km_s: Option<StochasticNoise>

Noise on the Doppler data of the measurement

Implementations

impl GroundStation

pub fn from_point( name: String, latitude_deg: f64, longitude_deg: f64, height_km: f64, frame: Frame, ) -> Self

Initializes a point on the surface of a celestial object. This is meant for analysis, not for spacecraft navigation.

Examples found in repository

examples/01_orbit_prop/main.rs (lines 194-200)

fn main() -> Result<(), Box<dyn Error>> {
    pel::init();
    // Dynamics models require planetary constants and ephemerides to be defined.
    // Let's start by grabbing those by using ANISE's latest MetaAlmanac.
    // This will automatically download the DE440s planetary ephemeris,
    // the daily-updated Earth Orientation Parameters, the high fidelity Moon orientation
    // parameters (for the Moon Mean Earth and Moon Principal Axes frames), and the PCK11
    // planetary constants kernels.
    // For details, refer to https://github.com/nyx-space/anise/blob/master/data/latest.dhall.
    // Note that we place the Almanac into an Arc so we can clone it cheaply and provide read-only
    // references to many functions.
    let almanac = Arc::new(MetaAlmanac::latest().map_err(Box::new)?);
    // Define the orbit epoch
    let epoch = Epoch::from_gregorian_utc_hms(2024, 2, 29, 12, 13, 14);

    // Define the orbit.
    // First we need to fetch the Earth J2000 from information from the Almanac.
    // This allows the frame to include the gravitational parameters and the shape of the Earth,
    // defined as a tri-axial ellipoid. Note that this shape can be changed manually or in the Almanac
    // by loading a different set of planetary constants.
    let earth_j2000 = almanac.frame_from_uid(EARTH_J2000)?;

    let orbit =
        Orbit::try_keplerian_altitude(300.0, 0.015, 68.5, 65.2, 75.0, 0.0, epoch, earth_j2000)?;
    // Print in in Keplerian form.
    println!("{orbit:x}");

    // There are two ways to propagate an orbit. We can make a quick approximation assuming only two-body
    // motion. This is a useful first order approximation but it isn't used in real-world applications.

    // This approach is a feature of ANISE.
    let future_orbit_tb = orbit.at_epoch(epoch + Unit::Day * 3)?;
    println!("{future_orbit_tb:x}");

    // Two body propagation relies solely on Kepler's laws, so only the true anomaly will change.
    println!(
        "SMA changed by {:.3e} km",
        orbit.sma_km()? - future_orbit_tb.sma_km()?
    );
    println!(
        "ECC changed by {:.3e}",
        orbit.ecc()? - future_orbit_tb.ecc()?
    );
    println!(
        "INC changed by {:.3e} deg",
        orbit.inc_deg()? - future_orbit_tb.inc_deg()?
    );
    println!(
        "RAAN changed by {:.3e} deg",
        orbit.raan_deg()? - future_orbit_tb.raan_deg()?
    );
    println!(
        "AOP changed by {:.3e} deg",
        orbit.aop_deg()? - future_orbit_tb.aop_deg()?
    );
    println!(
        "TA changed by {:.3} deg",
        orbit.ta_deg()? - future_orbit_tb.ta_deg()?
    );

    // Nyx is used for high fidelity propagation, not Keplerian propagation as above.
    // Nyx only propagates Spacecraft at the moment, which allows it to account for acceleration
    // models such as solar radiation pressure.

    // Let's build a cubesat sized spacecraft, with an SRP area of 10 cm^2 and a mass of 9.6 kg.
    let sc = Spacecraft::builder()
        .orbit(orbit)
        .dry_mass_kg(9.60)
        .srp(SrpConfig {
            area_m2: 10e-4,
            cr: 1.1,
        })
        .build();
    println!("{sc:x}");

    // Set up the spacecraft dynamics.

    // Specify that the orbital dynamics must account for the graviational pull of the Moon and the Sun.
    // The gravity of the Earth will also be accounted for since the spaceraft in an Earth orbit.
    let mut orbital_dyn = OrbitalDynamics::point_masses(vec![MOON, SUN]);

    // We want to include the spherical harmonics, so let's download the gravitational data from the Nyx Cloud.
    // We're using the JGM3 model here, which is the default in GMAT.
    let mut jgm3_meta = MetaFile {
        uri: "http://public-data.nyxspace.com/nyx/models/JGM3.cof.gz".to_string(),
        crc32: Some(0xF446F027), // Specifying the CRC32 avoids redownloading it if it's cached.
    };
    // And let's download it if we don't have it yet.
    jgm3_meta.process(true)?;

    // Build the spherical harmonics.
    // The harmonics must be computed in the body fixed frame.
    // We're using the long term prediction of the Earth centered Earth fixed frame, IAU Earth.
    let harmonics_21x21 = Harmonics::from_stor(
        almanac.frame_from_uid(IAU_EARTH_FRAME)?,
        HarmonicsMem::from_cof(&jgm3_meta.uri, 21, 21, true).unwrap(),
    );

    // Include the spherical harmonics into the orbital dynamics.
    orbital_dyn.accel_models.push(harmonics_21x21);

    // We define the solar radiation pressure, using the default solar flux and accounting only
    // for the eclipsing caused by the Earth.
    let srp_dyn = SolarPressure::default(EARTH_J2000, almanac.clone())?;

    // Finalize setting up the dynamics, specifying the force models (orbital_dyn) separately from the
    // acceleration models (SRP in this case). Use `from_models` to specify multiple accel models.
    let dynamics = SpacecraftDynamics::from_model(orbital_dyn, srp_dyn);

    println!("{dynamics}");

    // Finally, let's propagate this orbit to the same epoch as above.
    // The first returned value is the spacecraft state at the final epoch.
    // The second value is the full trajectory where the step size is variable step used by the propagator.
    let (future_sc, trajectory) = Propagator::default(dynamics)
        .with(sc, almanac.clone())
        .until_epoch_with_traj(future_orbit_tb.epoch)?;

    println!("=== High fidelity propagation ===");
    println!(
        "SMA changed by {:.3} km",
        orbit.sma_km()? - future_sc.orbit.sma_km()?
    );
    println!(
        "ECC changed by {:.6}",
        orbit.ecc()? - future_sc.orbit.ecc()?
    );
    println!(
        "INC changed by {:.3e} deg",
        orbit.inc_deg()? - future_sc.orbit.inc_deg()?
    );
    println!(
        "RAAN changed by {:.3} deg",
        orbit.raan_deg()? - future_sc.orbit.raan_deg()?
    );
    println!(
        "AOP changed by {:.3} deg",
        orbit.aop_deg()? - future_sc.orbit.aop_deg()?
    );
    println!(
        "TA changed by {:.3} deg",
        orbit.ta_deg()? - future_sc.orbit.ta_deg()?
    );

    // We also have access to the full trajectory throughout the propagation.
    println!("{trajectory}");

    // With the trajectory, let's build a few data products.

    // 1. Export the trajectory as a CCSDS OEM version 2.0 file and as a parquet file, which includes the Keplerian orbital elements.

    trajectory.to_oem_file(
        "./01_cubesat_hf_prop.oem",
        ExportCfg::builder().step(Unit::Minute * 2).build(),
    )?;

    trajectory.to_parquet_with_cfg(
        "./01_cubesat_hf_prop.parquet",
        ExportCfg::builder().step(Unit::Minute * 2).build(),
        almanac.clone(),
    )?;

    // 2. Compare the difference in the radial-intrack-crosstrack frame between the high fidelity
    // and Keplerian propagation. The RIC frame is commonly used to compute the difference in position
    // and velocity of different spacecraft.
    // 3. Compute the azimuth, elevation, range, and range-rate data of that spacecraft as seen from Boulder, CO, USA.

    let boulder_station = GroundStation::from_point(
        "Boulder, CO, USA".to_string(),
        40.014984,   // latitude in degrees
        -105.270546, // longitude in degrees
        1.6550,      // altitude in kilometers
        almanac.frame_from_uid(IAU_EARTH_FRAME)?,
    );

    // We iterate over the trajectory, grabbing a state every two minutes.
    let mut offset_s = vec![];
    let mut epoch_str = vec![];
    let mut ric_x_km = vec![];
    let mut ric_y_km = vec![];
    let mut ric_z_km = vec![];
    let mut ric_vx_km_s = vec![];
    let mut ric_vy_km_s = vec![];
    let mut ric_vz_km_s = vec![];

    let mut azimuth_deg = vec![];
    let mut elevation_deg = vec![];
    let mut range_km = vec![];
    let mut range_rate_km_s = vec![];
    for state in trajectory.every(Unit::Minute * 2) {
        // Try to compute the Keplerian/two body state just in time.
        // This method occasionally fails to converge on an appropriate true anomaly
        // from the mean anomaly. If that happens, we just skip this state.
        // The high fidelity and Keplerian states diverge continuously, and we're curious
        // about the divergence in this quick analysis.
        let this_epoch = state.epoch();
        match orbit.at_epoch(this_epoch) {
            Ok(tb_then) => {
                offset_s.push((this_epoch - orbit.epoch).to_seconds());
                epoch_str.push(format!("{this_epoch}"));
                // Compute the two body state just in time.
                let ric = state.orbit.ric_difference(&tb_then)?;
                ric_x_km.push(ric.radius_km.x);
                ric_y_km.push(ric.radius_km.y);
                ric_z_km.push(ric.radius_km.z);
                ric_vx_km_s.push(ric.velocity_km_s.x);
                ric_vy_km_s.push(ric.velocity_km_s.y);
                ric_vz_km_s.push(ric.velocity_km_s.z);

                // Compute the AER data for each state.
                let aer = almanac.azimuth_elevation_range_sez(
                    state.orbit,
                    boulder_station.to_orbit(this_epoch, &almanac)?,
                    None,
                    None,
                )?;
                azimuth_deg.push(aer.azimuth_deg);
                elevation_deg.push(aer.elevation_deg);
                range_km.push(aer.range_km);
                range_rate_km_s.push(aer.range_rate_km_s);
            }
            Err(e) => warn!("{} {e}", state.epoch()),
        };
    }

    // Build the data frames.
    let ric_df = df!(
        "Offset (s)" => offset_s.clone(),
        "Epoch" => epoch_str.clone(),
        "RIC X (km)" => ric_x_km,
        "RIC Y (km)" => ric_y_km,
        "RIC Z (km)" => ric_z_km,
        "RIC VX (km/s)" => ric_vx_km_s,
        "RIC VY (km/s)" => ric_vy_km_s,
        "RIC VZ (km/s)" => ric_vz_km_s,
    )?;

    println!("RIC difference at start\n{}", ric_df.head(Some(10)));
    println!("RIC difference at end\n{}", ric_df.tail(Some(10)));

    let aer_df = df!(
        "Offset (s)" => offset_s.clone(),
        "Epoch" => epoch_str.clone(),
        "azimuth (deg)" => azimuth_deg,
        "elevation (deg)" => elevation_deg,
        "range (km)" => range_km,
        "range rate (km/s)" => range_rate_km_s,
    )?;

    // Finally, let's see when the spacecraft is visible, assuming 15 degrees minimum elevation.
    let mask = aer_df.column("elevation (deg)")?.gt(15.0)?;
    let cubesat_visible = aer_df.filter(&mask)?;

    println!("{cubesat_visible}");

    Ok(())
}

pub fn dss65_madrid( elevation_mask: f64, range_noise_km: StochasticNoise, doppler_noise_km_s: StochasticNoise, iau_earth: Frame, ) -> Self

pub fn dss34_canberra( elevation_mask: f64, range_noise_km: StochasticNoise, doppler_noise_km_s: StochasticNoise, iau_earth: Frame, ) -> Self

pub fn dss13_goldstone( elevation_mask: f64, range_noise_km: StochasticNoise, doppler_noise_km_s: StochasticNoise, iau_earth: Frame, ) -> Self

pub fn azimuth_elevation_of( &self, rx: Orbit, obstructing_body: Option<Frame>, almanac: &Almanac, ) -> AlmanacResult<AzElRange>

Computes the azimuth and elevation of the provided object seen from this ground station, both in degrees. This is a shortcut to almanac.azimuth_elevation_range_sez.

pub fn to_orbit(&self, epoch: Epoch, almanac: &Almanac) -> PhysicsResult<Orbit>

Return this ground station as an orbit in its current frame

Examples found in repository

examples/01_orbit_prop/main.rs (line 239)

fn main() -> Result<(), Box<dyn Error>> {
    pel::init();
    // Dynamics models require planetary constants and ephemerides to be defined.
    // Let's start by grabbing those by using ANISE's latest MetaAlmanac.
    // This will automatically download the DE440s planetary ephemeris,
    // the daily-updated Earth Orientation Parameters, the high fidelity Moon orientation
    // parameters (for the Moon Mean Earth and Moon Principal Axes frames), and the PCK11
    // planetary constants kernels.
    // For details, refer to https://github.com/nyx-space/anise/blob/master/data/latest.dhall.
    // Note that we place the Almanac into an Arc so we can clone it cheaply and provide read-only
    // references to many functions.
    let almanac = Arc::new(MetaAlmanac::latest().map_err(Box::new)?);
    // Define the orbit epoch
    let epoch = Epoch::from_gregorian_utc_hms(2024, 2, 29, 12, 13, 14);

    // Define the orbit.
    // First we need to fetch the Earth J2000 from information from the Almanac.
    // This allows the frame to include the gravitational parameters and the shape of the Earth,
    // defined as a tri-axial ellipoid. Note that this shape can be changed manually or in the Almanac
    // by loading a different set of planetary constants.
    let earth_j2000 = almanac.frame_from_uid(EARTH_J2000)?;

    let orbit =
        Orbit::try_keplerian_altitude(300.0, 0.015, 68.5, 65.2, 75.0, 0.0, epoch, earth_j2000)?;
    // Print in in Keplerian form.
    println!("{orbit:x}");

    // There are two ways to propagate an orbit. We can make a quick approximation assuming only two-body
    // motion. This is a useful first order approximation but it isn't used in real-world applications.

    // This approach is a feature of ANISE.
    let future_orbit_tb = orbit.at_epoch(epoch + Unit::Day * 3)?;
    println!("{future_orbit_tb:x}");

    // Two body propagation relies solely on Kepler's laws, so only the true anomaly will change.
    println!(
        "SMA changed by {:.3e} km",
        orbit.sma_km()? - future_orbit_tb.sma_km()?
    );
    println!(
        "ECC changed by {:.3e}",
        orbit.ecc()? - future_orbit_tb.ecc()?
    );
    println!(
        "INC changed by {:.3e} deg",
        orbit.inc_deg()? - future_orbit_tb.inc_deg()?
    );
    println!(
        "RAAN changed by {:.3e} deg",
        orbit.raan_deg()? - future_orbit_tb.raan_deg()?
    );
    println!(
        "AOP changed by {:.3e} deg",
        orbit.aop_deg()? - future_orbit_tb.aop_deg()?
    );
    println!(
        "TA changed by {:.3} deg",
        orbit.ta_deg()? - future_orbit_tb.ta_deg()?
    );

    // Nyx is used for high fidelity propagation, not Keplerian propagation as above.
    // Nyx only propagates Spacecraft at the moment, which allows it to account for acceleration
    // models such as solar radiation pressure.

    // Let's build a cubesat sized spacecraft, with an SRP area of 10 cm^2 and a mass of 9.6 kg.
    let sc = Spacecraft::builder()
        .orbit(orbit)
        .dry_mass_kg(9.60)
        .srp(SrpConfig {
            area_m2: 10e-4,
            cr: 1.1,
        })
        .build();
    println!("{sc:x}");

    // Set up the spacecraft dynamics.

    // Specify that the orbital dynamics must account for the graviational pull of the Moon and the Sun.
    // The gravity of the Earth will also be accounted for since the spaceraft in an Earth orbit.
    let mut orbital_dyn = OrbitalDynamics::point_masses(vec![MOON, SUN]);

    // We want to include the spherical harmonics, so let's download the gravitational data from the Nyx Cloud.
    // We're using the JGM3 model here, which is the default in GMAT.
    let mut jgm3_meta = MetaFile {
        uri: "http://public-data.nyxspace.com/nyx/models/JGM3.cof.gz".to_string(),
        crc32: Some(0xF446F027), // Specifying the CRC32 avoids redownloading it if it's cached.
    };
    // And let's download it if we don't have it yet.
    jgm3_meta.process(true)?;

    // Build the spherical harmonics.
    // The harmonics must be computed in the body fixed frame.
    // We're using the long term prediction of the Earth centered Earth fixed frame, IAU Earth.
    let harmonics_21x21 = Harmonics::from_stor(
        almanac.frame_from_uid(IAU_EARTH_FRAME)?,
        HarmonicsMem::from_cof(&jgm3_meta.uri, 21, 21, true).unwrap(),
    );

    // Include the spherical harmonics into the orbital dynamics.
    orbital_dyn.accel_models.push(harmonics_21x21);

    // We define the solar radiation pressure, using the default solar flux and accounting only
    // for the eclipsing caused by the Earth.
    let srp_dyn = SolarPressure::default(EARTH_J2000, almanac.clone())?;

    // Finalize setting up the dynamics, specifying the force models (orbital_dyn) separately from the
    // acceleration models (SRP in this case). Use `from_models` to specify multiple accel models.
    let dynamics = SpacecraftDynamics::from_model(orbital_dyn, srp_dyn);

    println!("{dynamics}");

    // Finally, let's propagate this orbit to the same epoch as above.
    // The first returned value is the spacecraft state at the final epoch.
    // The second value is the full trajectory where the step size is variable step used by the propagator.
    let (future_sc, trajectory) = Propagator::default(dynamics)
        .with(sc, almanac.clone())
        .until_epoch_with_traj(future_orbit_tb.epoch)?;

    println!("=== High fidelity propagation ===");
    println!(
        "SMA changed by {:.3} km",
        orbit.sma_km()? - future_sc.orbit.sma_km()?
    );
    println!(
        "ECC changed by {:.6}",
        orbit.ecc()? - future_sc.orbit.ecc()?
    );
    println!(
        "INC changed by {:.3e} deg",
        orbit.inc_deg()? - future_sc.orbit.inc_deg()?
    );
    println!(
        "RAAN changed by {:.3} deg",
        orbit.raan_deg()? - future_sc.orbit.raan_deg()?
    );
    println!(
        "AOP changed by {:.3} deg",
        orbit.aop_deg()? - future_sc.orbit.aop_deg()?
    );
    println!(
        "TA changed by {:.3} deg",
        orbit.ta_deg()? - future_sc.orbit.ta_deg()?
    );

    // We also have access to the full trajectory throughout the propagation.
    println!("{trajectory}");

    // With the trajectory, let's build a few data products.

    // 1. Export the trajectory as a CCSDS OEM version 2.0 file and as a parquet file, which includes the Keplerian orbital elements.

    trajectory.to_oem_file(
        "./01_cubesat_hf_prop.oem",
        ExportCfg::builder().step(Unit::Minute * 2).build(),
    )?;

    trajectory.to_parquet_with_cfg(
        "./01_cubesat_hf_prop.parquet",
        ExportCfg::builder().step(Unit::Minute * 2).build(),
        almanac.clone(),
    )?;

    // 2. Compare the difference in the radial-intrack-crosstrack frame between the high fidelity
    // and Keplerian propagation. The RIC frame is commonly used to compute the difference in position
    // and velocity of different spacecraft.
    // 3. Compute the azimuth, elevation, range, and range-rate data of that spacecraft as seen from Boulder, CO, USA.

    let boulder_station = GroundStation::from_point(
        "Boulder, CO, USA".to_string(),
        40.014984,   // latitude in degrees
        -105.270546, // longitude in degrees
        1.6550,      // altitude in kilometers
        almanac.frame_from_uid(IAU_EARTH_FRAME)?,
    );

    // We iterate over the trajectory, grabbing a state every two minutes.
    let mut offset_s = vec![];
    let mut epoch_str = vec![];
    let mut ric_x_km = vec![];
    let mut ric_y_km = vec![];
    let mut ric_z_km = vec![];
    let mut ric_vx_km_s = vec![];
    let mut ric_vy_km_s = vec![];
    let mut ric_vz_km_s = vec![];

    let mut azimuth_deg = vec![];
    let mut elevation_deg = vec![];
    let mut range_km = vec![];
    let mut range_rate_km_s = vec![];
    for state in trajectory.every(Unit::Minute * 2) {
        // Try to compute the Keplerian/two body state just in time.
        // This method occasionally fails to converge on an appropriate true anomaly
        // from the mean anomaly. If that happens, we just skip this state.
        // The high fidelity and Keplerian states diverge continuously, and we're curious
        // about the divergence in this quick analysis.
        let this_epoch = state.epoch();
        match orbit.at_epoch(this_epoch) {
            Ok(tb_then) => {
                offset_s.push((this_epoch - orbit.epoch).to_seconds());
                epoch_str.push(format!("{this_epoch}"));
                // Compute the two body state just in time.
                let ric = state.orbit.ric_difference(&tb_then)?;
                ric_x_km.push(ric.radius_km.x);
                ric_y_km.push(ric.radius_km.y);
                ric_z_km.push(ric.radius_km.z);
                ric_vx_km_s.push(ric.velocity_km_s.x);
                ric_vy_km_s.push(ric.velocity_km_s.y);
                ric_vz_km_s.push(ric.velocity_km_s.z);

                // Compute the AER data for each state.
                let aer = almanac.azimuth_elevation_range_sez(
                    state.orbit,
                    boulder_station.to_orbit(this_epoch, &almanac)?,
                    None,
                    None,
                )?;
                azimuth_deg.push(aer.azimuth_deg);
                elevation_deg.push(aer.elevation_deg);
                range_km.push(aer.range_km);
                range_rate_km_s.push(aer.range_rate_km_s);
            }
            Err(e) => warn!("{} {e}", state.epoch()),
        };
    }

    // Build the data frames.
    let ric_df = df!(
        "Offset (s)" => offset_s.clone(),
        "Epoch" => epoch_str.clone(),
        "RIC X (km)" => ric_x_km,
        "RIC Y (km)" => ric_y_km,
        "RIC Z (km)" => ric_z_km,
        "RIC VX (km/s)" => ric_vx_km_s,
        "RIC VY (km/s)" => ric_vy_km_s,
        "RIC VZ (km/s)" => ric_vz_km_s,
    )?;

    println!("RIC difference at start\n{}", ric_df.head(Some(10)));
    println!("RIC difference at end\n{}", ric_df.tail(Some(10)));

    let aer_df = df!(
        "Offset (s)" => offset_s.clone(),
        "Epoch" => epoch_str.clone(),
        "azimuth (deg)" => azimuth_deg,
        "elevation (deg)" => elevation_deg,
        "range (km)" => range_km,
        "range rate (km/s)" => range_rate_km_s,
    )?;

    // Finally, let's see when the spacecraft is visible, assuming 15 degrees minimum elevation.
    let mask = aer_df.column("elevation (deg)")?.gt(15.0)?;
    let cubesat_visible = aer_df.filter(&mask)?;

    println!("{cubesat_visible}");

    Ok(())
}

Trait Implementations

impl Clone for GroundStation

fn clone(&self) -> GroundStation

Returns a copy of the value.

fn clone_from(&mut self, source: &Self)

Performs copy-assignment from source.

impl ConfigRepr for GroundStation

fn load<P>(path: P) -> Result<Self, ConfigError>

where P: AsRef<Path>,

Builds the configuration representation from the path to a yaml

fn load_many<P>(path: P) -> Result<Vec<Self>, ConfigError>

where P: AsRef<Path>,

Builds a sequence of “Selves” from the provided path to a yaml

fn load_named<P>(path: P) -> Result<BTreeMap<String, Self>, ConfigError>

where P: AsRef<Path>,

Builds a map of names to “selves” from the provided path to a yaml

fn loads_many(data: &str) -> Result<Vec<Self>, ConfigError>

Builds a sequence of “Selves” from the provided string of a yaml

fn loads_named(data: &str) -> Result<BTreeMap<String, Self>, ConfigError>

Builds a sequence of “Selves” from the provided string of a yaml

impl Debug for GroundStation

fn fmt(&self, f: &mut Formatter<'_>) -> Result

Formats the value using the given formatter.

impl<'de> Deserialize<'de> for GroundStation

fn deserialize<__D>(__deserializer: __D) -> Result<Self, __D::Error>

where __D: Deserializer<'de>,

Deserialize this value from the given Serde deserializer.

impl Display for GroundStation

fn fmt(&self, f: &mut Formatter<'_>) -> Result

Formats the value using the given formatter.

impl<S: Interpolatable> EventEvaluator<S> for &GroundStation

where DefaultAllocator: Allocator<S::Size> + Allocator<S::Size, S::Size> + Allocator<S::VecLength>,

fn eval( &self, rx_gs_frame: &S, almanac: Arc<Almanac>, ) -> Result<f64, EventError>

Compute the elevation in the SEZ frame. This call will panic if the frame of the input state does not match that of the ground station.

fn epoch_precision(&self) -> Duration

Epoch precision of the election evaluator is 1 ms

fn value_precision(&self) -> f64

Angle precision of the elevation evaluator is 1 millidegree.

fn eval_string( &self, state: &S, almanac: Arc<Almanac>, ) -> Result<String, EventError>

Returns a string representation of the event evaluation for the given state

fn eval_crossing( &self, prev_state: &S, next_state: &S, almanac: Arc<Almanac>, ) -> Result<bool, EventError>

impl PartialEq for GroundStation

fn eq(&self, other: &GroundStation) -> bool

Tests for self and other values to be equal, and is used by ==.

fn ne(&self, other: &Rhs) -> bool

Tests for !=. The default implementation is almost always sufficient, and should not be overridden without very good reason.

impl Serialize for GroundStation

fn serialize<__S>(&self, __serializer: __S) -> Result<__S::Ok, __S::Error>

where __S: Serializer,

Serialize this value into the given Serde serializer.

impl TrackingDeviceSim<Spacecraft, RangeDoppler> for GroundStation

fn measure( &mut self, epoch: Epoch, traj: &Traj<Spacecraft>, rng: Option<&mut Pcg64Mcg>, almanac: Arc<Almanac>, ) -> Result<Option<RangeDoppler>, ODError>

Perform a measurement from the ground station to the receiver (rx).

fn measurement_covar( &mut self, epoch: Epoch, ) -> Result<OMatrix<f64, <RangeDoppler as Measurement>::MeasurementSize, <RangeDoppler as Measurement>::MeasurementSize>, ODError>

Returns the measurement noise of this ground station.

Methodology

Noises are modeled using a StochasticNoise process, defined by the sigma on the turn-on bias and on the steady state noise. The measurement noise is computed assuming that all measurements are independent variables, i.e. the measurement matrix is a diagonal matrix. The first item in the diagonal is the range noise (in km), set to the square of the steady state sigma. The second item is the Doppler noise (in km/s), set to the square of the steady state sigma of that Gauss Markov process.

fn name(&self) -> String

Returns the name of this tracking data simulator

fn location( &self, epoch: Epoch, frame: Frame, almanac: Arc<Almanac>, ) -> AlmanacResult<Orbit>

Returns the device location at the given epoch and in the given frame.

fn measure_instantaneous( &mut self, rx: Spacecraft, rng: Option<&mut Pcg64Mcg>, almanac: Arc<Almanac>, ) -> Result<Option<RangeDoppler>, ODError>

impl StructuralPartialEq for GroundStation