Struct OrbitalDynamics 

pub struct OrbitalDynamics {
    pub accel_models: Vec<Arc<dyn AccelModel + Sync>>,
}

OrbitalDynamics provides the equations of motion for any celestial dynamic, without state transition matrix computation.

Fields

accel_models: Vec<Arc<dyn AccelModel + Sync>>

Implementations

impl OrbitalDynamics

pub fn point_masses(celestial_objects: Vec<i32>) -> Self

Initializes the point masses gravities with the provided list of bodies

Examples found in repository

examples/03_geo_analysis/stationkeeping.rs (line 77)

fn main() -> Result<(), Box<dyn Error>> {
    pel::init();
    // Set up the dynamics like in the orbit raise.
    let almanac = Arc::new(MetaAlmanac::latest().map_err(Box::new)?);
    let epoch = Epoch::from_gregorian_utc_hms(2024, 2, 29, 12, 13, 14);

    // Define the GEO orbit, and we're just going to maintain it very tightly.
    let earth_j2000 = almanac.frame_info(EARTH_J2000)?;
    let orbit = Orbit::try_keplerian(42164.0, 1e-5, 0., 163.0, 75.0, 0.0, epoch, earth_j2000)?;
    println!("{orbit:x}");

    let sc = Spacecraft::builder()
        .orbit(orbit)
        .mass(Mass::from_dry_and_prop_masses(1000.0, 1000.0)) // 1000 kg of dry mass and prop, totalling 2.0 tons
        .srp(SRPData::from_area(3.0 * 6.0)) // Assuming 1 kW/m^2 or 18 kW, giving a margin of 4.35 kW for on-propulsion consumption
        .thruster(Thruster {
            // "NEXT-STEP" row in Table 2
            isp_s: 4435.0,
            thrust_N: 0.472,
        })
        .mode(GuidanceMode::Thrust) // Start thrusting immediately.
        .build();

    // Set up the spacecraft dynamics like in the orbit raise example.

    let prop_time = 30.0 * Unit::Day;

    // Define the guidance law -- we're just using a Ruggiero controller as demonstrated in AAS-2004-5089.
    let objectives = &[
        Objective::within_tolerance(
            StateParameter::Element(OrbitalElement::SemiMajorAxis),
            42_165.0,
            20.0,
        ),
        Objective::within_tolerance(
            StateParameter::Element(OrbitalElement::Eccentricity),
            0.001,
            5e-5,
        ),
        Objective::within_tolerance(
            StateParameter::Element(OrbitalElement::Inclination),
            0.05,
            1e-2,
        ),
    ];

    let ruggiero_ctrl = Ruggiero::from_max_eclipse(objectives, sc, 0.2)?;
    println!("{ruggiero_ctrl}");

    let mut orbital_dyn = OrbitalDynamics::point_masses(vec![MOON, SUN]);

    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.
    };
    jgm3_meta.process(true)?;

    let harmonics = GravityField::from_stor(
        almanac.frame_info(IAU_EARTH_FRAME)?,
        GravityFieldData::from_cof(&jgm3_meta.uri, 8, 8, true)?,
    );
    orbital_dyn.accel_models.push(harmonics);

    let srp_dyn = SolarPressure::default_flux(EARTH_J2000, almanac.clone())?;
    let sc_dynamics = SpacecraftDynamics::from_model(orbital_dyn, srp_dyn)
        .with_guidance_law(ruggiero_ctrl.clone());

    println!("{sc_dynamics}");

    // Finally, let's use the Monte Carlo framework built into Nyx to propagate spacecraft.

    // Let's start by defining the dispersion.
    // The MultivariateNormal structure allows us to define the dispersions in any of the orbital parameters, but these are applied directly in the Cartesian state space.
    // Note that additional validation on the MVN is in progress -- https://github.com/nyx-space/nyx/issues/339.
    let mc_rv = MvnSpacecraft::new(
        sc,
        vec![StateDispersion::zero_mean(
            StateParameter::Element(OrbitalElement::SemiMajorAxis),
            3.0,
        )],
    )?;

    let my_mc = MonteCarlo::new(
        sc, // Nominal state
        mc_rv,
        "03_geo_sk".to_string(), // Scenario name
        None, // No specific seed specified, so one will be drawn from the computer's entropy.
    );

    // Build the propagator setup.
    let setup = Propagator::rk89(
        sc_dynamics.clone(),
        IntegratorOptions::builder()
            .min_step(10.0_f64.seconds())
            .error_ctrl(ErrorControl::RSSCartesianStep)
            .build(),
    );

    let num_runs = 25;
    let rslts = my_mc.run_until_epoch(setup, almanac.clone(), sc.epoch() + prop_time, num_runs);

    assert_eq!(rslts.runs.len(), num_runs);

    rslts.to_parquet("03_geo_sk.parquet", ExportCfg::default())?;

    Ok(())
}

examples/03_geo_analysis/raise.rs (line 95)

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)?);
    // Fetch the EME2000 frame from the Almabac
    let eme2k = almanac.frame_info(EARTH_J2000).unwrap();
    // Define the orbit epoch
    let epoch = Epoch::from_gregorian_utc_hms(2024, 2, 29, 12, 13, 14);

    // Build the spacecraft itself.
    // Using slide 6 of https://aerospace.org/sites/default/files/2018-11/Davis-Mayberry_HPSEP_11212018.pdf
    // for the "next gen" SEP characteristics.

    // GTO start
    let orbit = Orbit::keplerian(24505.9, 0.725, 7.05, 0.0, 0.0, 0.0, epoch, eme2k);

    let sc = Spacecraft::builder()
        .orbit(orbit)
        .mass(Mass::from_dry_and_prop_masses(1000.0, 1000.0)) // 1000 kg of dry mass and prop, totalling 2.0 tons
        .srp(SRPData::from_area(3.0 * 6.0)) // Assuming 1 kW/m^2 or 18 kW, giving a margin of 4.35 kW for on-propulsion consumption
        .thruster(Thruster {
            // "NEXT-STEP" row in Table 2
            isp_s: 4435.0,
            thrust_N: 0.472,
        })
        .mode(GuidanceMode::Thrust) // Start thrusting immediately.
        .build();

    let prop_time = 180.0 * Unit::Day;

    // Define the guidance law -- we're just using a Ruggiero controller as demonstrated in AAS-2004-5089.
    let objectives = &[
        Objective::within_tolerance(
            StateParameter::Element(OrbitalElement::SemiMajorAxis),
            42_165.0,
            20.0,
        ),
        Objective::within_tolerance(
            StateParameter::Element(OrbitalElement::Eccentricity),
            0.001,
            5e-5,
        ),
        Objective::within_tolerance(
            StateParameter::Element(OrbitalElement::Inclination),
            0.05,
            1e-2,
        ),
    ];

    // Ensure that we only thrust if we have more than 20% illumination.
    let ruggiero_ctrl = Ruggiero::from_max_eclipse(objectives, sc, 0.2).unwrap();
    println!("{ruggiero_ctrl}");

    // Define the high fidelity dynamics

    // 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 = GravityField::from_stor(
        almanac.frame_info(IAU_EARTH_FRAME)?,
        GravityFieldData::from_cof(&jgm3_meta.uri, 8, 8, true).unwrap(),
    );

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

    // 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_flux(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 sc_dynamics = SpacecraftDynamics::from_model(orbital_dyn, srp_dyn)
        .with_guidance_law(ruggiero_ctrl.clone());

    println!("{orbit:x}");

    // We specify a minimum step in the propagator because the Ruggiero control would otherwise drive this step very low.
    let (final_state, traj) = Propagator::rk89(
        sc_dynamics.clone(),
        IntegratorOptions::builder()
            .min_step(10.0_f64.seconds())
            .error_ctrl(ErrorControl::RSSCartesianStep)
            .build(),
    )
    .with(sc, almanac.clone())
    .for_duration_with_traj(prop_time)?;

    let prop_usage = sc.mass.prop_mass_kg - final_state.mass.prop_mass_kg;
    println!("{:x}", final_state.orbit);
    println!("prop usage: {prop_usage:.3} kg");

    // Finally, export the results for analysis, including the penumbra percentage throughout the orbit raise.
    traj.to_parquet("./03_geo_raise.parquet", ExportCfg::default())?;

    for status_line in ruggiero_ctrl.status(&final_state) {
        println!("{status_line}");
    }

    ruggiero_ctrl
        .achieved(&final_state)
        .expect("objective not achieved");

    Ok(())
}

examples/02_jwst_covar_monte_carlo/main.rs (line 99)

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.
    // For details, refer to https://github.com/nyx-space/anise/blob/master/data/latest.dhall.

    // Download the regularly update of the James Webb Space Telescope reconstucted (or definitive) ephemeris.
    // Refer to https://naif.jpl.nasa.gov/pub/naif/JWST/kernels/spk/aareadme.txt for details.
    let mut latest_jwst_ephem = MetaFile {
        uri: "https://naif.jpl.nasa.gov/pub/naif/JWST/kernels/spk/jwst_rec.bsp".to_string(),
        crc32: None,
    };
    latest_jwst_ephem.process(true)?;

    // Load this ephem in the general Almanac we're using for this analysis.
    let almanac = Arc::new(
        MetaAlmanac::latest()
            .map_err(Box::new)?
            .load_from_metafile(latest_jwst_ephem, true)?,
    );

    // By loading this ephemeris file in the ANISE GUI or ANISE CLI, we can find the NAIF ID of the JWST
    // in the BSP. We need this ID in order to query the ephemeris.
    const JWST_NAIF_ID: i32 = -170;
    // Let's build a frame in the J2000 orientation centered on the JWST.
    const JWST_J2000: Frame = Frame::from_ephem_j2000(JWST_NAIF_ID);

    // Since the ephemeris file is updated regularly, we'll just grab the latest state in the ephem.
    let (earliest_epoch, latest_epoch) = almanac.spk_domain(JWST_NAIF_ID)?;
    println!("JWST defined from {earliest_epoch} to {latest_epoch}");
    // Fetch the state, printing it in the Earth J2000 frame.
    let jwst_orbit = almanac.transform(JWST_J2000, EARTH_J2000, latest_epoch, None)?;
    println!("{jwst_orbit:x}");

    // Build the spacecraft
    // SRP area assumed to be the full sunshield and mass if 6200.0 kg, c.f. https://webb.nasa.gov/content/about/faqs/facts.html
    // SRP Coefficient of reflectivity assumed to be that of Kapton, i.e. 2 - 0.44 = 1.56, table 1 from https://amostech.com/TechnicalPapers/2018/Poster/Bengtson.pdf
    let jwst = Spacecraft::builder()
        .orbit(jwst_orbit)
        .srp(SRPData {
            area_m2: 21.197 * 14.162,
            coeff_reflectivity: 1.56,
        })
        .mass(Mass::from_dry_mass(6200.0))
        .build();

    // Build up the spacecraft uncertainty builder.
    // We can use the spacecraft uncertainty structure to build this up.
    // We start by specifying the nominal state (as defined above), then the uncertainty in position and velocity
    // in the RIC frame. We could also specify the Cr, Cd, and mass uncertainties, but these aren't accounted for until
    // Nyx can also estimate the deviation of the spacecraft parameters.
    let jwst_uncertainty = SpacecraftUncertainty::builder()
        .nominal(jwst)
        .frame(LocalFrame::RIC)
        .x_km(0.5)
        .y_km(0.3)
        .z_km(1.5)
        .vx_km_s(1e-4)
        .vy_km_s(0.6e-3)
        .vz_km_s(3e-3)
        .build();

    println!("{jwst_uncertainty}");

    // Build the Kalman filter estimate.
    // Note that we could have used the KfEstimate structure directly (as seen throughout the OD integration tests)
    // but this approach requires quite a bit more boilerplate code.
    let jwst_estimate = jwst_uncertainty.to_estimate()?;

    // Set up the spacecraft dynamics.
    // We'll use the point masses of the Earth, Sun, Jupiter (barycenter, because it's in the DE440), and the Moon.
    // We'll also enable solar radiation pressure since the James Webb has a huge and highly reflective sun shield.

    let orbital_dyn = OrbitalDynamics::point_masses(vec![MOON, SUN, JUPITER_BARYCENTER]);
    let srp_dyn = SolarPressure::new(vec![EARTH_J2000, MOON_J2000], almanac.clone())?;

    // Finalize setting up the dynamics.
    let dynamics = SpacecraftDynamics::from_model(orbital_dyn, srp_dyn);

    // Build the propagator set up to use for the whole analysis.
    let setup = Propagator::default(dynamics);

    // All of the analysis will use this duration.
    let prediction_duration = 6.5 * Unit::Day;

    // === Covariance mapping ===
    // For the covariance mapping / prediction, we'll use the common orbit determination approach.
    // This is done by setting up a spacecraft Kalman filter OD process, and predicting for the analysis duration.

    // Build the propagation instance for the OD process.
    let odp = SpacecraftKalmanOD::new(
        setup.clone(),
        KalmanVariant::DeviationTracking,
        None,
        BTreeMap::new(),
        almanac.clone(),
    );

    // The prediction step is 1 minute by default, configured in the OD process, i.e. how often we want to know the covariance.
    assert_eq!(odp.max_step, 1_i64.minutes());
    // Finally, predict, and export the trajectory with covariance to a parquet file.
    let od_sol = odp.predict_for(jwst_estimate, prediction_duration)?;
    od_sol.to_parquet("./02_jwst_covar_map.parquet", ExportCfg::default())?;

    // === Monte Carlo framework ===
    // Nyx comes with a complete multi-threaded Monte Carlo frame. It's blazing fast.

    let my_mc = MonteCarlo::new(
        jwst, // Nominal state
        jwst_estimate.to_random_variable()?,
        "02_jwst".to_string(), // Scenario name
        None, // No specific seed specified, so one will be drawn from the computer's entropy.
    );

    let num_runs = 5_000;
    let rslts = my_mc.run_until_epoch(
        setup,
        almanac.clone(),
        jwst.epoch() + prediction_duration,
        num_runs,
    );

    assert_eq!(rslts.runs.len(), num_runs);
    // Finally, export these results, computing the eclipse percentage for all of these results.

    rslts.to_parquet("02_jwst_monte_carlo.parquet", ExportCfg::default())?;

    Ok(())
}

examples/05_cislunar_spacecraft_link_od/main.rs (line 82)

fn main() -> Result<(), Box<dyn Error>> {
    pel::init();

    // ====================== //
    // === ALMANAC SET UP === //
    // ====================== //

    let manifest_dir =
        PathBuf::from(std::env::var("CARGO_MANIFEST_DIR").unwrap_or(".".to_string()));

    let out = manifest_dir.join("data/04_output/");

    let almanac = Arc::new(
        Almanac::new(
            &manifest_dir
                .join("data/01_planetary/pck08.pca")
                .to_string_lossy(),
        )
        .unwrap()
        .load(
            &manifest_dir
                .join("data/01_planetary/de440s.bsp")
                .to_string_lossy(),
        )
        .unwrap(),
    );

    let eme2k = almanac.frame_info(EARTH_J2000).unwrap();
    let moon_iau = almanac.frame_info(IAU_MOON_FRAME).unwrap();

    let epoch = Epoch::from_gregorian_tai(2021, 5, 29, 19, 51, 16, 852_000);
    let nrho = Orbit::cartesian(
        166_473.631_302_239_7,
        -274_715.487_253_382_7,
        -211_233.210_176_686_7,
        0.933_451_604_520_018_4,
        0.436_775_046_841_900_9,
        -0.082_211_021_250_348_95,
        epoch,
        eme2k,
    );

    let tx_nrho_sc = Spacecraft::from(nrho);

    let state_luna = almanac.transform_to(nrho, MOON_J2000, None).unwrap();
    println!("Start state (dynamics: Earth, Moon, Sun gravity):\n{state_luna}");

    let bodies = vec![EARTH, SUN];
    let dynamics = SpacecraftDynamics::new(OrbitalDynamics::point_masses(bodies));

    let setup = Propagator::rk89(
        dynamics,
        IntegratorOptions::builder().max_step(0.5.minutes()).build(),
    );

    /* == Propagate the NRHO vehicle == */
    let prop_time = 1.1 * state_luna.period().unwrap();

    let (nrho_final, mut tx_traj) = setup
        .with(tx_nrho_sc, almanac.clone())
        .for_duration_with_traj(prop_time)
        .unwrap();

    tx_traj.name = Some("NRHO Tx SC".to_string());

    println!("{tx_traj}");

    /* == Propagate an LLO vehicle == */
    let llo_orbit =
        Orbit::try_keplerian_altitude(110.0, 1e-4, 90.0, 0.0, 0.0, 0.0, epoch, moon_iau).unwrap();

    let llo_sc = Spacecraft::builder().orbit(llo_orbit).build();

    let (_, llo_traj) = setup
        .with(llo_sc, almanac.clone())
        .until_epoch_with_traj(nrho_final.epoch())
        .unwrap();

    // Export the subset of the first two hours.
    llo_traj
        .clone()
        .filter_by_offset(..2.hours())
        .to_parquet_simple(out.join("05_caps_llo_truth.pq"))?;

    /* == Setup the interlink == */

    let mut measurement_types = IndexSet::new();
    measurement_types.insert(MeasurementType::Range);
    measurement_types.insert(MeasurementType::Doppler);

    let mut stochastics = IndexMap::new();

    let sa45_csac_allan_dev = 1e-11;

    stochastics.insert(
        MeasurementType::Range,
        StochasticNoise::from_hardware_range_km(
            sa45_csac_allan_dev,
            10.0.seconds(),
            link_specific::ChipRate::StandardT4B,
            link_specific::SN0::Average,
        ),
    );

    stochastics.insert(
        MeasurementType::Doppler,
        StochasticNoise::from_hardware_doppler_km_s(
            sa45_csac_allan_dev,
            10.0.seconds(),
            link_specific::CarrierFreq::SBand,
            link_specific::CN0::Average,
        ),
    );

    let interlink = InterlinkTxSpacecraft {
        traj: tx_traj,
        measurement_types,
        integration_time: None,
        timestamp_noise_s: None,
        ab_corr: Aberration::LT,
        stochastic_noises: Some(stochastics),
    };

    // Devices are the transmitter, which is our NRHO vehicle.
    let mut devices = BTreeMap::new();
    devices.insert("NRHO Tx SC".to_string(), interlink);

    let mut configs = BTreeMap::new();
    configs.insert(
        "NRHO Tx SC".to_string(),
        TrkConfig::builder()
            .strands(vec![Strand {
                start: epoch,
                end: nrho_final.epoch(),
            }])
            .build(),
    );

    let mut trk_sim =
        TrackingArcSim::with_seed(devices.clone(), llo_traj.clone(), configs, 0).unwrap();
    println!("{trk_sim}");

    let trk_data = trk_sim.generate_measurements(almanac.clone()).unwrap();
    println!("{trk_data}");

    trk_data
        .to_parquet_simple(out.clone().join("nrho_interlink_msr.pq"))
        .unwrap();

    // Run a truth OD where we estimate the LLO position
    let llo_uncertainty = SpacecraftUncertainty::builder()
        .nominal(llo_sc)
        .x_km(1.0)
        .y_km(1.0)
        .z_km(1.0)
        .vx_km_s(1e-3)
        .vy_km_s(1e-3)
        .vz_km_s(1e-3)
        .build();

    let mut proc_devices = devices.clone();

    // Define the initial estimate, randomized, seed for reproducibility
    let mut initial_estimate = llo_uncertainty.to_estimate_randomized(Some(0)).unwrap();
    // Inflate the covariance -- https://github.com/nyx-space/nyx/issues/339
    initial_estimate.covar *= 2.5;

    // Increase the noise in the devices to accept more measurements.

    for link in proc_devices.values_mut() {
        for noise in &mut link.stochastic_noises.as_mut().unwrap().values_mut() {
            *noise.white_noise.as_mut().unwrap() *= 3.0;
        }
    }

    let init_err = initial_estimate
        .orbital_state()
        .ric_difference(&llo_orbit)
        .unwrap();

    println!("initial estimate:\n{initial_estimate}");
    println!("RIC errors = {init_err}",);

    let odp = InterlinkKalmanOD::new(
        setup.clone(),
        KalmanVariant::ReferenceUpdate,
        Some(ResidRejectCrit::default()),
        proc_devices,
        almanac.clone(),
    );

    // Shrink the data to process.
    let arc = trk_data.filter_by_offset(..2.hours());

    let od_sol = odp.process_arc(initial_estimate, &arc).unwrap();

    println!("{od_sol}");

    od_sol
        .to_parquet(
            out.join("05_caps_interlink_od_sol.pq"),
            ExportCfg::default(),
        )
        .unwrap();

    let od_traj = od_sol.to_traj().unwrap();

    od_traj
        .ric_diff_to_parquet(
            &llo_traj,
            out.join("05_caps_interlink_llo_est_error.pq"),
            ExportCfg::default(),
        )
        .unwrap();

    let final_est = od_sol.estimates.last().unwrap();
    assert!(final_est.within_3sigma(), "should be within 3 sigma");

    println!("ESTIMATE\n{final_est:x}\n");
    let truth = llo_traj.at(final_est.epoch()).unwrap();
    println!("TRUTH\n{truth:x}");

    let final_err = truth
        .orbit
        .ric_difference(&final_est.orbital_state())
        .unwrap();
    println!("ERROR {final_err}");

    // Build the residuals versus reference plot.
    let rvr_sol = odp
        .process_arc(initial_estimate, &arc.resid_vs_ref_check())
        .unwrap();

    rvr_sol
        .to_parquet(
            out.join("05_caps_interlink_resid_v_ref.pq"),
            ExportCfg::default(),
        )
        .unwrap();

    let final_rvr = rvr_sol.estimates.last().unwrap();

    println!("RMAG error {:.3} m", final_err.rmag_km() * 1e3);
    println!(
        "Pure prop error {:.3} m",
        final_rvr
            .orbital_state()
            .ric_difference(&final_est.orbital_state())
            .unwrap()
            .rmag_km()
            * 1e3
    );

    Ok(())
}

examples/06_lunar_orbit_determination/main.rs (line 92)

fn main() -> Result<(), Box<dyn Error>> {
    pel::init();

    // ====================== //
    // === ALMANAC SET UP === //
    // ====================== //

    // Dynamics models require planetary constants and ephemerides to be defined.
    // Let's start by grabbing those by using ANISE's MetaAlmanac.

    let data_folder: PathBuf = [
        env!("CARGO_MANIFEST_DIR"),
        "examples",
        "06_lunar_orbit_determination",
    ]
    .iter()
    .collect();

    let meta = data_folder.join("metaalmanac.dhall");

    // Load this ephem in the general Almanac we're using for this analysis.
    let almanac = MetaAlmanac::new(meta.to_string_lossy().as_ref())
        .map_err(Box::new)?
        .process(true)
        .map_err(Box::new)?;

    // Lock the almanac (an Arc is a read only structure).
    let almanac = Arc::new(almanac);

    // Build a nominal trajectory
    // TODO: Switch this to a sequence once the OD over a spacecraft sequence is implemented.

    let epoch = Epoch::from_gregorian_utc_at_noon(2024, 2, 29);
    let moon_j2000 = almanac.frame_info(MOON_J2000)?;

    // To build the trajectory we need to provide a spacecraft template.
    let orbiter = Spacecraft::builder()
        .mass(Mass::from_dry_and_prop_masses(1018.0, 900.0))
        .srp(SRPData {
            area_m2: 3.9 * 2.7,
            coeff_reflectivity: 0.96,
        })
        .orbit(Orbit::try_keplerian_altitude(
            150.0, 0.00212, 33.6, 45.0, 45.0, 0.0, epoch, moon_j2000,
        )?) // Setting a zero orbit here because it's just a template
        .build();

    // ========================== //
    // === BUILD NOMINAL TRAJ === //
    // ========================== //

    // Set up the spacecraft dynamics.

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

    // We want to include the spherical harmonics, so let's download the gravitational data from the Nyx Cloud.
    // We're using the GRAIL JGGRX model.
    let mut jggrx_meta = MetaFile {
        uri: "http://public-data.nyxspace.com/nyx/models/Luna_jggrx_1500e_sha.tab.gz".to_string(),
        crc32: Some(0x6bcacda8), // Specifying the CRC32 avoids redownloading it if it's cached.
    };
    // And let's download it if we don't have it yet.
    jggrx_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 Moon principal axes frame.
    let moon_pa_frame = MOON_PA_FRAME.with_orient(31008);
    let sph_harmonics = GravityField::from_stor(
        almanac.frame_info(moon_pa_frame)?,
        GravityFieldData::from_shadr(&jggrx_meta.uri, 80, 80, true)?,
    );

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

    // We define the solar radiation pressure, using the default solar flux and accounting only
    // for the eclipsing caused by the Earth and Moon.
    // Note that by default, enabling the SolarPressure model will also enable the estimation of the coefficient of reflectivity.
    let srp_dyn = SolarPressure::new(vec![MOON_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}");

    let setup = Propagator::rk89(dynamics.clone(), IntegratorOptions::default());

    let truth_traj = setup
        .with(orbiter, almanac.clone())
        .for_duration_with_traj(Unit::Day * 2)?
        .1;

    // ==================== //
    // === OD SIMULATOR === //
    // ==================== //

    // Load the Deep Space Network ground stations.
    // Nyx allows you to build these at runtime but it's pretty static so we can just load them from YAML.
    let ground_station_file = data_folder.join("dsn-network.yaml");
    let devices = GroundStation::load_named(ground_station_file)?;

    let proc_devices = devices.clone();

    // Typical OD software requires that you specify your own tracking schedule or you'll have overlapping measurements.
    // Nyx can build a tracking schedule for you based on the first station with access.
    let configs: BTreeMap<String, TrkConfig> =
        TrkConfig::load_named(data_folder.join("tracking-cfg.yaml"))?;

    // Build the tracking arc simulation to generate a "standard measurement".
    let mut trk = TrackingArcSim::<Spacecraft, GroundStation>::with_seed(
        devices.clone(),
        truth_traj.clone(),
        configs,
        123, // Set a seed for reproducibility
    )?;

    trk.build_schedule(almanac.clone())?;
    let arc = trk.generate_measurements(almanac.clone())?;
    // Save the simulated tracking data
    arc.to_parquet_simple("./data/04_output/06_lunar_simulated_tracking.parquet")?;

    // We'll note that in our case, we have continuous coverage of LRO when the vehicle is not behind the Moon.
    println!("{arc}");

    // Now that we have simulated measurements, we'll run the orbit determination.

    // ===================== //
    // === OD ESTIMATION === //
    // ===================== //

    let sc = SpacecraftUncertainty::builder()
        .nominal(orbiter)
        .frame(LocalFrame::RIC)
        .x_km(0.5)
        .y_km(0.5)
        .z_km(0.5)
        .vx_km_s(5e-3)
        .vy_km_s(5e-3)
        .vz_km_s(5e-3)
        .build();

    // Build the filter initial estimate, which we will reuse in the filter.
    let initial_estimate = sc.to_estimate()?;

    println!("== FILTER STATE ==\n{orbiter:x}\n{initial_estimate}");

    // Build the SNC in the Moon J2000 frame, specified as a velocity noise over time.
    let process_noise = ProcessNoise3D::from_velocity_km_s(
        &[1e-13, 1e-13, 1e-13],
        1 * Unit::Hour,
        10 * Unit::Minute,
        None,
    );

    println!("{process_noise}");

    // We'll set up the OD process to reject measurements whose residuals are move than 3 sigmas away from what we expect.
    let odp = SpacecraftKalmanScalarOD::new(
        setup,
        KalmanVariant::ReferenceUpdate,
        Some(ResidRejectCrit::default()),
        proc_devices,
        almanac.clone(),
    )
    .with_process_noise(process_noise);

    let od_sol = odp.process_arc(initial_estimate, &arc)?;

    let final_est = od_sol.estimates.last().unwrap();

    println!("{final_est}");

    let ric_err = truth_traj
        .at(final_est.epoch())?
        .orbit
        .ric_difference(&final_est.orbital_state())?;
    println!("== RIC at end ==");
    println!("RIC Position (m): {:.3}", ric_err.radius_km * 1e3);
    println!("RIC Velocity (m/s): {:.3}", ric_err.velocity_km_s * 1e3);

    println!(
        "Num residuals rejected: #{}",
        od_sol.rejected_residuals().len()
    );
    println!(
        "Percentage within +/-3: {}",
        od_sol.residual_ratio_within_threshold(3.0).unwrap()
    );
    println!("Ratios normal? {}", od_sol.is_normal(None).unwrap());

    od_sol.to_parquet(
        "./data/04_output/06_lunar_od_results.parquet",
        ExportCfg::default(),
    )?;

    let od_trajectory = od_sol.to_traj()?;
    // Build the RIC difference.
    od_trajectory.ric_diff_to_parquet(
        &truth_traj,
        "./data/04_output/06_lunar_od_truth_error.parquet",
        ExportCfg::default(),
    )?;

    Ok(())
}

examples/03_geo_analysis/drift.rs (line 76)

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_info(EARTH_J2000)?;

    // Placing this GEO bird just above Colorado.
    // In theory, the eccentricity is zero, but in practice, it's about 1e-5 to 1e-6 at best.
    let orbit = Orbit::try_keplerian(42164.0, 1e-5, 0., 163.0, 75.0, 0.0, epoch, earth_j2000)?;
    // Print in in Keplerian form.
    println!("{orbit:x}");

    let state_bf = almanac.transform_to(orbit, IAU_EARTH_FRAME, None)?;
    let (orig_lat_deg, orig_long_deg, orig_alt_km) = state_bf.latlongalt()?;

    // 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)
        .mass(Mass::from_dry_mass(9.60))
        .srp(SRPData {
            area_m2: 10e-4,
            coeff_reflectivity: 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 = GravityField::from_stor(
        almanac.frame_info(IAU_EARTH_FRAME)?,
        GravityFieldData::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 and Moon.
    let srp_dyn = SolarPressure::new(vec![EARTH_J2000, MOON_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(epoch + Unit::Century * 0.03)?;

    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}");

    println!("Spacecraft params after 3 years without active control:\n{future_sc:x}");

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

    // 1. Export the trajectory as a parquet file, which includes the Keplerian orbital elements.

    let analysis_step = Unit::Minute * 5;

    trajectory.to_parquet(
        "./03_geo_hf_prop.parquet",
        ExportCfg::builder().step(analysis_step).build(),
    )?;

    // 2. Compute the latitude, longitude, and altitude throughout the trajectory by rotating the spacecraft position into the Earth body fixed frame.

    // We iterate over the trajectory, grabbing a state every two minutes.
    let mut offset_s = vec![];
    let mut epoch_str = vec![];
    let mut longitude_deg = vec![];
    let mut latitude_deg = vec![];
    let mut altitude_km = vec![];

    for state in trajectory.every(analysis_step) {
        // Convert the GEO bird state into the body fixed frame, and keep track of its latitude, longitude, and altitude.
        // These define the GEO stationkeeping box.

        let this_epoch = state.epoch();

        offset_s.push((this_epoch - orbit.epoch).to_seconds());
        epoch_str.push(this_epoch.to_isoformat());

        let state_bf = almanac.transform_to(state.orbit, IAU_EARTH_FRAME, None)?;
        let (lat_deg, long_deg, alt_km) = state_bf.latlongalt()?;
        longitude_deg.push(long_deg);
        latitude_deg.push(lat_deg);
        altitude_km.push(alt_km);
    }

    println!(
        "Longitude changed by {:.3} deg -- Box is 0.1 deg E-W",
        orig_long_deg - longitude_deg.last().unwrap()
    );

    println!(
        "Latitude changed by {:.3} deg -- Box is 0.05 deg N-S",
        orig_lat_deg - latitude_deg.last().unwrap()
    );

    println!(
        "Altitude changed by {:.3} km -- Box is 30 km",
        orig_alt_km - altitude_km.last().unwrap()
    );

    // Build the station keeping data frame.
    let mut sk_df = df!(
        "Offset (s)" => offset_s.clone(),
        "Epoch (UTC)" => epoch_str.clone(),
        "Longitude E-W (deg)" => longitude_deg,
        "Latitude N-S (deg)" => latitude_deg,
        "Altitude (km)" => altitude_km,

    )?;

    // Create a file to write the Parquet to
    let file = File::create("./03_geo_lla.parquet").expect("Could not create file");

    // Create a ParquetWriter and write the DataFrame to the file
    ParquetWriter::new(file).finish(&mut sk_df)?;

    Ok(())
}

Additional examples can be found in:

• examples/01_orbit_prop/main.rs
• examples/04_lro_od/main.rs

pub fn two_body() -> Self

Initializes a OrbitalDynamics which does not simulate the gravity pull of other celestial objects but the primary one.

pub fn new(accel_models: Vec<Arc<dyn AccelModel + Sync>>) -> Self

Initialize orbital dynamics with a list of acceleration models

pub fn from_model(accel_model: Arc<dyn AccelModel + Sync>) -> Self

Initialize new orbital mechanics with the provided model. Note: Orbital dynamics always include two body dynamics, these cannot be turned off.

impl OrbitalDynamics

pub fn dual_eom( &self, _delta_t_s: f64, osc: &Orbit, almanac: Arc<Almanac>, ) -> Result<(Vector6<f64>, Matrix6<f64>), DynamicsError>

Trait Implementations

impl Clone for OrbitalDynamics

fn clone(&self) -> OrbitalDynamics

Returns a duplicate of the value.

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

Performs copy-assignment from source.

impl Display for OrbitalDynamics

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

Formats the value using the given formatter.