pub struct SpacecraftDynamics {
pub orbital_dyn: OrbitalDynamics,
pub force_models: Vec<Arc<dyn ForceModel>>,
pub guid_law: Option<Arc<dyn GuidanceLaw>>,
pub decrement_mass: bool,
}
Expand description
A generic spacecraft dynamics with associated force models, guidance law, and flag specifying whether to decrement the fuel mass or not. Note: when developing new guidance laws, it is recommended to not enable fuel decrement until the guidance law seems to work without proper physics. Note: if the spacecraft runs out of fuel, the propagation segment will return an error.
Fields§
§orbital_dyn: OrbitalDynamics
§force_models: Vec<Arc<dyn ForceModel>>
§guid_law: Option<Arc<dyn GuidanceLaw>>
§decrement_mass: bool
Implementations§
Source§impl SpacecraftDynamics
impl SpacecraftDynamics
Sourcepub fn from_guidance_law(
orbital_dyn: OrbitalDynamics,
guid_law: Arc<dyn GuidanceLaw>,
) -> Self
pub fn from_guidance_law( orbital_dyn: OrbitalDynamics, guid_law: Arc<dyn GuidanceLaw>, ) -> Self
Initialize a Spacecraft with a set of orbital dynamics and a propulsion subsystem. By default, the mass of the vehicle will be decremented as propellant is consumed.
Sourcepub fn from_guidance_law_no_decr(
orbital_dyn: OrbitalDynamics,
guid_law: Arc<dyn GuidanceLaw>,
) -> Self
pub fn from_guidance_law_no_decr( orbital_dyn: OrbitalDynamics, guid_law: Arc<dyn GuidanceLaw>, ) -> Self
Initialize a Spacecraft with a set of orbital dynamics and a propulsion subsystem. Will not decrement the fuel mass as propellant is consumed.
Sourcepub fn new(orbital_dyn: OrbitalDynamics) -> Self
pub fn new(orbital_dyn: OrbitalDynamics) -> Self
Initialize a Spacecraft with a set of orbital dynamics and with SRP enabled.
Sourcepub fn from_model(
orbital_dyn: OrbitalDynamics,
force_model: Arc<dyn ForceModel>,
) -> Self
pub fn from_model( orbital_dyn: OrbitalDynamics, force_model: Arc<dyn ForceModel>, ) -> Self
Initialize new spacecraft dynamics with the provided orbital mechanics and with the provided force model.
Examples found in repository?
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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_from_uid(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)
.dry_mass_kg(1000.0) // 1000 kg of dry mass
.fuel_mass_kg(1000.0) // 1000 kg of fuel, totalling 2.0 tons
.srp(SrpConfig::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::SMA, 42_164.0, 5.0), // 5 km
Objective::within_tolerance(StateParameter::Eccentricity, 0.001, 5e-5),
Objective::within_tolerance(StateParameter::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 = Harmonics::from_stor(
almanac.frame_from_uid(IAU_EARTH_FRAME)?,
HarmonicsMem::from_cof(&jgm3_meta.uri, 8, 8, true)?,
);
orbital_dyn.accel_models.push(harmonics);
let srp_dyn = SolarPressure::default(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 = MultivariateNormal::new(
sc,
vec![StateDispersion::zero_mean(StateParameter::SMA, 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);
// For all of the resulting trajectories, we'll want to compute the percentage of penumbra and umbra.
rslts.to_parquet(
"03_geo_sk.parquet",
Some(vec![
&EclipseLocator::cislunar(almanac.clone()).to_penumbra_event()
]),
ExportCfg::default(),
almanac,
)?;
Ok(())
}
More examples
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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_from_uid(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)
.dry_mass_kg(1000.0) // 1000 kg of dry mass
.fuel_mass_kg(1000.0) // 1000 kg of fuel, totalling 2.0 tons
.srp(SrpConfig::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::SMA, 42_165.0, 20.0),
Objective::within_tolerance(StateParameter::Eccentricity, 0.001, 5e-5),
Objective::within_tolerance(StateParameter::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 = Harmonics::from_stor(
almanac.frame_from_uid(IAU_EARTH_FRAME)?,
HarmonicsMem::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(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!("{:x}", orbit);
// 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 fuel_usage = sc.fuel_mass_kg - final_state.fuel_mass_kg;
println!("{:x}", final_state.orbit);
println!("fuel usage: {:.3} kg", fuel_usage);
// Finally, export the results for analysis, including the penumbra percentage throughout the orbit raise.
traj.to_parquet(
"./03_geo_raise.parquet",
Some(vec![
&EclipseLocator::cislunar(almanac.clone()).to_penumbra_event()
]),
ExportCfg::default(),
almanac,
)?;
for status_line in ruggiero_ctrl.status(&final_state) {
println!("{status_line}");
}
ruggiero_ctrl
.achieved(&final_state)
.expect("objective not achieved");
Ok(())
}
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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(SrpConfig {
area_m2: 21.197 * 14.162,
cr: 1.56,
})
.dry_mass_kg(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 OD process, and predicting for the analysis duration.
let ckf = KF::no_snc(jwst_estimate);
// Build the propagation instance for the OD process.
let prop = setup.with(jwst.with_stm(), almanac.clone());
let mut odp = SpacecraftODProcess::ckf(prop, ckf, None, almanac.clone());
// Define the prediction step, i.e. how often we want to know the covariance.
let step = 1_i64.minutes();
// Finally, predict, and export the trajectory with covariance to a parquet file.
odp.predict_for(step, prediction_duration)?;
odp.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.
// For all of the resulting trajectories, we'll want to compute the percentage of penumbra and umbra.
let eclipse_loc = EclipseLocator::cislunar(almanac.clone());
let umbra_event = eclipse_loc.to_umbra_event();
let penumbra_event = eclipse_loc.to_penumbra_event();
rslts.to_parquet(
"02_jwst_monte_carlo.parquet",
Some(vec![&umbra_event, &penumbra_event]),
ExportCfg::default(),
almanac,
)?;
Ok(())
}
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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)?;
// 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)
.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 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",
Some(vec![
&EclipseLocator::cislunar(almanac.clone()).to_penumbra_event()
]),
ExportCfg::builder().step(analysis_step).build(),
almanac.clone(),
)?;
// 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(())
}
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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", "04_lro_od"]
.iter()
.collect();
let meta = data_folder.join("lro-dynamics.dhall");
// Load this ephem in the general Almanac we're using for this analysis.
let mut almanac = MetaAlmanac::new(meta.to_string_lossy().to_string())
.map_err(Box::new)?
.process(true)
.map_err(Box::new)?;
let mut moon_pc = almanac.planetary_data.get_by_id(MOON)?;
moon_pc.mu_km3_s2 = 4902.74987;
almanac.planetary_data.set_by_id(MOON, moon_pc)?;
let mut earth_pc = almanac.planetary_data.get_by_id(EARTH)?;
earth_pc.mu_km3_s2 = 398600.436;
almanac.planetary_data.set_by_id(EARTH, earth_pc)?;
// Save this new kernel for reuse.
// In an operational context, this would be part of the "Lock" process, and should not change throughout the mission.
almanac
.planetary_data
.save_as(&data_folder.join("lro-specific.pca"), true)?;
// Lock the almanac (an Arc is a read only structure).
let almanac = Arc::new(almanac);
// Orbit determination requires a Trajectory structure, which can be saved as parquet file.
// In our case, the trajectory comes from the BSP file, so we need to build a Trajectory from the almanac directly.
// To query the Almanac, we need to build the LRO frame in the J2000 orientation in our case.
// Inspecting the LRO BSP in the ANISE GUI shows us that NASA has assigned ID -85 to LRO.
let lro_frame = Frame::from_ephem_j2000(-85);
// To build the trajectory we need to provide a spacecraft template.
let sc_template = Spacecraft::builder()
.dry_mass_kg(1018.0) // Launch masses
.fuel_mass_kg(900.0)
.srp(SrpConfig {
// SRP configuration is arbitrary, but we will be estimating it anyway.
area_m2: 3.9 * 2.7,
cr: 0.96,
})
.orbit(Orbit::zero(MOON_J2000)) // Setting a zero orbit here because it's just a template
.build();
// Now we can build the trajectory from the BSP file.
// We'll arbitrarily set the tracking arc to 48 hours with a one minute time step.
let traj_as_flown = Traj::from_bsp(
lro_frame,
MOON_J2000,
almanac.clone(),
sc_template,
5.seconds(),
Some(Epoch::from_str("2024-01-01 00:00:00 UTC")?),
Some(Epoch::from_str("2024-01-02 00:00:00 UTC")?),
Aberration::LT,
Some("LRO".to_string()),
)?;
println!("{traj_as_flown}");
// ====================== //
// === MODEL MATCHING === //
// ====================== //
// 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 moon_pa_frame = IAU_MOON_FRAME;
let sph_harmonics = Harmonics::from_stor(
almanac.frame_from_uid(moon_pa_frame)?,
HarmonicsMem::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![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}");
// Now we can build the propagator.
let setup = Propagator::default_dp78(dynamics.clone());
// For reference, let's build the trajectory with Nyx's models from that LRO state.
let (sim_final, traj_as_sim) = setup
.with(*traj_as_flown.first(), almanac.clone())
.until_epoch_with_traj(traj_as_flown.last().epoch())?;
println!("SIM INIT: {:x}", traj_as_flown.first());
println!("SIM FINAL: {sim_final:x}");
// Compute RIC difference between SIM and LRO ephem
let sim_lro_delta = sim_final
.orbit
.ric_difference(&traj_as_flown.last().orbit)?;
println!("{traj_as_sim}");
println!(
"SIM v LRO - RIC Position (m): {:.3}",
sim_lro_delta.radius_km * 1e3
);
println!(
"SIM v LRO - RIC Velocity (m/s): {:.3}",
sim_lro_delta.velocity_km_s * 1e3
);
traj_as_sim.ric_diff_to_parquet(
&traj_as_flown,
"./04_lro_sim_truth_error.parquet",
ExportCfg::default(),
)?;
// ==================== //
// === OD SIMULATOR === //
// ==================== //
// After quite some time trying to exactly match the model, we still end up with an oscillatory difference on the order of 150 meters between the propagated state
// and the truth LRO state.
// Therefore, we will actually run an estimation from a dispersed LRO state.
// The sc_seed is the true LRO state from the BSP.
let sc_seed = *traj_as_flown.first();
// 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: PathBuf = [
env!("CARGO_MANIFEST_DIR"),
"examples",
"04_lro_od",
"dsn-network.yaml",
]
.iter()
.collect();
let devices = GroundStation::load_many(ground_station_file)?;
// 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 trkconfg_yaml: PathBuf = [
env!("CARGO_MANIFEST_DIR"),
"examples",
"04_lro_od",
"tracking-cfg.yaml",
]
.iter()
.collect();
let configs: BTreeMap<String, TrkConfig> = TrkConfig::load_named(trkconfg_yaml)?;
// Build the tracking arc simulation to generate a "standard measurement".
let mut trk = TrackingArcSim::<Spacecraft, RangeDoppler, _>::new(
devices,
traj_as_flown.clone(),
configs,
)?;
trk.build_schedule(almanac.clone())?;
let arc = trk.generate_measurements(almanac.clone())?;
// Save the simulated tracking data
arc.to_parquet_simple("./04_lro_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(sc_seed)
.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{sc_seed:x}\n{initial_estimate}");
let kf = KF::new(
// Increase the initial covariance to account for larger deviation.
initial_estimate,
// Until https://github.com/nyx-space/nyx/issues/351, we need to specify the SNC in the acceleration of the Moon J2000 frame.
SNC3::from_diagonal(10 * Unit::Minute, &[1e-11, 1e-11, 1e-11]),
);
// We'll set up the OD process to reject measurements whose residuals are mover than 4 sigmas away from what we expect.
let mut odp = ODProcess::ckf(
setup.with(initial_estimate.state().with_stm(), almanac.clone()),
kf,
Some(ResidRejectCrit::default()),
almanac.clone(),
);
odp.process_arc::<GroundStation>(&arc)?;
let ric_err = traj_as_flown
.at(odp.estimates.last().unwrap().epoch())?
.orbit
.ric_difference(&odp.estimates.last().unwrap().orbital_state())?;
println!("== RIC at end ==");
println!("RIC Position (m): {}", ric_err.radius_km * 1e3);
println!("RIC Velocity (m/s): {}", ric_err.velocity_km_s * 1e3);
odp.to_parquet("./04_lro_od_results.parquet", ExportCfg::default())?;
// In our case, we have the truth trajectory from NASA.
// So we can compute the RIC state difference between the real LRO ephem and what we've just estimated.
// Export the OD trajectory first.
let od_trajectory = odp.to_traj()?;
// Build the RIC difference.
od_trajectory.ric_diff_to_parquet(
&traj_as_flown,
"./04_lro_od_truth_error.parquet",
ExportCfg::default(),
)?;
Ok(())
}
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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(())
}
Sourcepub fn from_models(
orbital_dyn: OrbitalDynamics,
force_models: Vec<Arc<dyn ForceModel>>,
) -> Self
pub fn from_models( orbital_dyn: OrbitalDynamics, force_models: Vec<Arc<dyn ForceModel>>, ) -> Self
Initialize new spacecraft dynamics with a vector of force models.
Sourcepub fn guidance_achieved(
&self,
state: &Spacecraft,
) -> Result<bool, GuidanceError>
pub fn guidance_achieved( &self, state: &Spacecraft, ) -> Result<bool, GuidanceError>
A shortcut to spacecraft.guid_law if a guidance law is defined for these dynamics
Sourcepub fn with_guidance_law(&self, guid_law: Arc<dyn GuidanceLaw>) -> Self
pub fn with_guidance_law(&self, guid_law: Arc<dyn GuidanceLaw>) -> Self
Clone these spacecraft dynamics and update the control to the one provided.
Examples found in repository?
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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_from_uid(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)
.dry_mass_kg(1000.0) // 1000 kg of dry mass
.fuel_mass_kg(1000.0) // 1000 kg of fuel, totalling 2.0 tons
.srp(SrpConfig::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::SMA, 42_164.0, 5.0), // 5 km
Objective::within_tolerance(StateParameter::Eccentricity, 0.001, 5e-5),
Objective::within_tolerance(StateParameter::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 = Harmonics::from_stor(
almanac.frame_from_uid(IAU_EARTH_FRAME)?,
HarmonicsMem::from_cof(&jgm3_meta.uri, 8, 8, true)?,
);
orbital_dyn.accel_models.push(harmonics);
let srp_dyn = SolarPressure::default(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 = MultivariateNormal::new(
sc,
vec![StateDispersion::zero_mean(StateParameter::SMA, 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);
// For all of the resulting trajectories, we'll want to compute the percentage of penumbra and umbra.
rslts.to_parquet(
"03_geo_sk.parquet",
Some(vec![
&EclipseLocator::cislunar(almanac.clone()).to_penumbra_event()
]),
ExportCfg::default(),
almanac,
)?;
Ok(())
}
More examples
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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_from_uid(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)
.dry_mass_kg(1000.0) // 1000 kg of dry mass
.fuel_mass_kg(1000.0) // 1000 kg of fuel, totalling 2.0 tons
.srp(SrpConfig::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::SMA, 42_165.0, 20.0),
Objective::within_tolerance(StateParameter::Eccentricity, 0.001, 5e-5),
Objective::within_tolerance(StateParameter::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 = Harmonics::from_stor(
almanac.frame_from_uid(IAU_EARTH_FRAME)?,
HarmonicsMem::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(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!("{:x}", orbit);
// 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 fuel_usage = sc.fuel_mass_kg - final_state.fuel_mass_kg;
println!("{:x}", final_state.orbit);
println!("fuel usage: {:.3} kg", fuel_usage);
// Finally, export the results for analysis, including the penumbra percentage throughout the orbit raise.
traj.to_parquet(
"./03_geo_raise.parquet",
Some(vec![
&EclipseLocator::cislunar(almanac.clone()).to_penumbra_event()
]),
ExportCfg::default(),
almanac,
)?;
for status_line in ruggiero_ctrl.status(&final_state) {
println!("{status_line}");
}
ruggiero_ctrl
.achieved(&final_state)
.expect("objective not achieved");
Ok(())
}
Trait Implementations§
Source§impl Clone for SpacecraftDynamics
impl Clone for SpacecraftDynamics
Source§fn clone(&self) -> SpacecraftDynamics
fn clone(&self) -> SpacecraftDynamics
1.0.0 · Source§fn clone_from(&mut self, source: &Self)
fn clone_from(&mut self, source: &Self)
source
. Read moreSource§impl Display for SpacecraftDynamics
impl Display for SpacecraftDynamics
Source§impl Dynamics for SpacecraftDynamics
impl Dynamics for SpacecraftDynamics
Source§type HyperdualSize = Const<9>
type HyperdualSize = Const<9>
type StateType = Spacecraft
Source§fn finally(
&self,
next_state: Self::StateType,
almanac: Arc<Almanac>,
) -> Result<Self::StateType, DynamicsError>
fn finally( &self, next_state: Self::StateType, almanac: Arc<Almanac>, ) -> Result<Self::StateType, DynamicsError>
Source§fn eom(
&self,
delta_t_s: f64,
state: &OVector<f64, Const<90>>,
ctx: &Self::StateType,
almanac: Arc<Almanac>,
) -> Result<OVector<f64, Const<90>>, DynamicsError>
fn eom( &self, delta_t_s: f64, state: &OVector<f64, Const<90>>, ctx: &Self::StateType, almanac: Arc<Almanac>, ) -> Result<OVector<f64, Const<90>>, DynamicsError>
Source§fn dual_eom(
&self,
delta_t_s: f64,
ctx: &Self::StateType,
almanac: Arc<Almanac>,
) -> Result<(OVector<f64, Const<9>>, OMatrix<f64, Const<9>, Const<9>>), DynamicsError>
fn dual_eom( &self, delta_t_s: f64, ctx: &Self::StateType, almanac: Arc<Almanac>, ) -> Result<(OVector<f64, Const<9>>, OMatrix<f64, Const<9>, Const<9>>), DynamicsError>
Auto Trait Implementations§
impl Freeze for SpacecraftDynamics
impl !RefUnwindSafe for SpacecraftDynamics
impl Send for SpacecraftDynamics
impl Sync for SpacecraftDynamics
impl Unpin for SpacecraftDynamics
impl !UnwindSafe for SpacecraftDynamics
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Source§impl<T> BorrowMut<T> for Twhere
T: ?Sized,
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T: ?Sized,
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fn borrow_mut(&mut self) -> &mut T
Source§impl<T> CloneToUninit for Twhere
T: Clone,
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T: Clone,
§impl<T> Instrument for T
impl<T> Instrument for T
§fn instrument(self, span: Span) -> Instrumented<Self>
fn instrument(self, span: Span) -> Instrumented<Self>
§fn in_current_span(self) -> Instrumented<Self>
fn in_current_span(self) -> Instrumented<Self>
Source§impl<T> IntoEither for T
impl<T> IntoEither for T
Source§fn into_either(self, into_left: bool) -> Either<Self, Self>
fn into_either(self, into_left: bool) -> Either<Self, Self>
self
into a Left
variant of Either<Self, Self>
if into_left
is true
.
Converts self
into a Right
variant of Either<Self, Self>
otherwise. Read moreSource§fn into_either_with<F>(self, into_left: F) -> Either<Self, Self>
fn into_either_with<F>(self, into_left: F) -> Either<Self, Self>
self
into a Left
variant of Either<Self, Self>
if into_left(&self)
returns true
.
Converts self
into a Right
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otherwise. Read more§impl<T> Pointable for T
impl<T> Pointable for T
§impl<SS, SP> SupersetOf<SS> for SPwhere
SS: SubsetOf<SP>,
impl<SS, SP> SupersetOf<SS> for SPwhere
SS: SubsetOf<SP>,
§fn to_subset(&self) -> Option<SS>
fn to_subset(&self) -> Option<SS>
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from the equivalent element of its
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is actually part of its subset T
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fn to_subset_unchecked(&self) -> SS
self.to_subset
but without any property checks. Always succeeds.§fn from_subset(element: &SS) -> SP
fn from_subset(element: &SS) -> SP
self
to the equivalent element of its superset.