Browse Source
- Combined test_analytical_propagation_apsides.cpp (248 lines) and test_analytical_propagation_timesteps.cpp (209 lines) into test_analytical_propagation.cpp (422 lines, -35 lines saved) - All 10 test cases preserved with unique spacecraft names - Merged config files into test_analytical_propagation.toml - Tests pass: 240,299 assertions in 133 test casesmain
5 changed files with 257 additions and 278 deletions
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# Test Configuration: Analytical Propagation Tests |
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# Combined configuration for apsides and timestep testing |
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# Contains two spacecraft with different orbital parameters |
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[[bodies]] |
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name = "Earth" |
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mass = 5.972e24 |
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radius = 6.371e6 |
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parent_index = -1 |
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color = { r = 0.0, g = 0.5, b = 1.0 } |
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orbit = { |
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semi_major_axis = 0.0, |
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eccentricity = 0.0, |
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true_anomaly = 0.0 |
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} |
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[[spacecraft]] |
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name = "Apsides_Test_Spacecraft" |
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mass = 1000.0 |
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parent_index = 0 |
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orbit = { |
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semi_major_axis = 2.0e7, |
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eccentricity = 0.6, |
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true_anomaly = 0.0, |
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inclination = 0.0, |
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longitude_of_ascending_node = 0.0, |
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argument_of_periapsis = 0.0 |
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} |
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[[spacecraft]] |
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name = "Timestep_Test_Spacecraft" |
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mass = 1000.0 |
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parent_index = 0 |
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orbit = { |
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semi_major_axis = 1.5e7, |
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eccentricity = 0.4, |
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true_anomaly = 0.0, |
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inclination = 0.0, |
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longitude_of_ascending_node = 0.0, |
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argument_of_periapsis = 0.0 |
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} |
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# Test Configuration: Elliptical Orbit for Analytical Propagation |
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# Moderate eccentricity to test propagation through apsides |
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[[bodies]] |
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name = "Earth" |
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mass = 5.972e24 |
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radius = 6.371e6 |
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parent_index = -1 |
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color = { r = 0.0, g = 0.5, b = 1.0 } |
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orbit = { |
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semi_major_axis = 0.0, |
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eccentricity = 0.0, |
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true_anomaly = 0.0 |
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} |
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[[spacecraft]] |
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name = "Elliptical_Orbit_Spacecraft" |
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mass = 1000.0 |
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parent_index = 0 |
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orbit = { |
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semi_major_axis = 2.0e7, |
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eccentricity = 0.6, |
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true_anomaly = 0.0, |
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inclination = 0.0, |
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longitude_of_ascending_node = 0.0, |
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argument_of_periapsis = 0.0 |
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} |
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#include <catch2/catch_test_macros.hpp> |
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#include "../src/physics.h" |
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#include "../src/orbital_mechanics.h" |
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#include "../src/simulation.h" |
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#include "../src/config_loader.h" |
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#include "../src/test_utilities.h" |
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#include <cmath> |
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const double VELOCITY_TOLERANCE = 10.0; |
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const double POSITION_TOLERANCE = 1.0e4; |
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TEST_CASE("Large timestep - dt greater than orbital period", "[analytical][timestep][large]") { |
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const double TIME_STEP = 60.0; |
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SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP); |
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REQUIRE(load_system_config(sim, "tests/test_analytical_propagation_timesteps.toml")); |
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Spacecraft* craft = &sim->spacecraft[0]; |
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CelestialBody* earth = &sim->bodies[0]; |
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double a = craft->orbit.semi_major_axis; |
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double mu = G * earth->mass; |
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double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); |
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INFO("Orbital period: " << period_seconds << " s (" << period_seconds / 3600.0 << " hours)"); |
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double large_dt = period_seconds * 2.0; |
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INFO("Timestep: " << large_dt << " s (2x orbital period)"); |
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Vec3 pos_before; |
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Vec3 vel_before; |
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orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_before, &vel_before); |
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OrbitalElements propagated = propagate_orbital_elements(craft->orbit, large_dt, earth->mass); |
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Vec3 pos_after; |
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Vec3 vel_after; |
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orbital_elements_to_cartesian(propagated, earth->mass, &pos_after, &vel_after); |
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double r_before = vec3_magnitude(pos_before); |
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double r_after = vec3_magnitude(pos_after); |
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double v_before = vec3_magnitude(vel_before); |
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double v_after = vec3_magnitude(vel_after); |
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INFO("Before propagation:"); |
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INFO(" Radius: " << r_before << " m"); |
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INFO(" Velocity: " << v_before << " m/s"); |
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INFO("After 2 periods:"); |
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INFO(" Radius: " << r_after << " m"); |
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INFO(" Velocity: " << v_after << " m/s"); |
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double r_error = fabs(r_after - r_before); |
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double v_error = fabs(v_after - v_before); |
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double relative_r_error = r_error / r_before * 100.0; |
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double relative_v_error = v_error / v_before * 100.0; |
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INFO("Radius error: " << r_error << " m (" << relative_r_error << "%)"); |
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INFO("Velocity error: " << v_error << " m/s (" << relative_v_error << "%)"); |
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REQUIRE(relative_r_error < 0.1); |
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REQUIRE(relative_v_error < 0.1); |
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destroy_simulation(sim); |
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} |
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TEST_CASE("Very small timestep - dt less than 1 second", "[analytical][timestep][small]") { |
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const double TIME_STEP = 60.0; |
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SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP); |
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REQUIRE(load_system_config(sim, "tests/test_analytical_propagation_timesteps.toml")); |
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Spacecraft* craft = &sim->spacecraft[0]; |
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CelestialBody* earth = &sim->bodies[0]; |
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Vec3 pos_before; |
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Vec3 vel_before; |
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orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_before, &vel_before); |
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double small_dt = 0.1; |
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INFO("Timestep: " << small_dt << " s"); |
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OrbitalElements propagated = propagate_orbital_elements(craft->orbit, small_dt, earth->mass); |
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Vec3 pos_after; |
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Vec3 vel_after; |
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orbital_elements_to_cartesian(propagated, earth->mass, &pos_after, &vel_after); |
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double pos_change = vec3_distance(pos_before, pos_after); |
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double vel_change = vec3_distance(vel_before, vel_after); |
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INFO("Position change: " << pos_change << " m"); |
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INFO("Velocity change: " << vel_change << " m/s"); |
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double v_before_mag = vec3_magnitude(vel_before); |
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double expected_pos_change = v_before_mag * small_dt; |
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double pos_error = fabs(pos_change - expected_pos_change); |
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INFO("Expected position change (v·dt): " << expected_pos_change << " m"); |
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INFO("Position error: " << pos_error << " m"); |
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INFO("Relative position error: " << (pos_error / expected_pos_change * 100.0) << "%"); |
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REQUIRE(pos_error < VELOCITY_TOLERANCE * small_dt * 10.0); |
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REQUIRE(vel_change < VELOCITY_TOLERANCE); |
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destroy_simulation(sim); |
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} |
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TEST_CASE("Accuracy vs timestep size relationship", "[analytical][timestep][accuracy]") { |
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const double TIME_STEP = 60.0; |
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SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP); |
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REQUIRE(load_system_config(sim, "tests/test_analytical_propagation_timesteps.toml")); |
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Spacecraft* craft = &sim->spacecraft[0]; |
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CelestialBody* earth = &sim->bodies[0]; |
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double a = craft->orbit.semi_major_axis; |
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double mu = G * earth->mass; |
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double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); |
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double dt_ratios[] = {0.01, 0.1, 1.0, 10.0}; |
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Vec3 pos_initial; |
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Vec3 vel_initial; |
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orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_initial, &vel_initial); |
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for (int i = 0; i < 4; i++) { |
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double dt = period_seconds * dt_ratios[i]; |
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INFO("Testing dt = " << dt << " s (" << dt_ratios[i] << "x period)"); |
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OrbitalElements propagated = propagate_orbital_elements(craft->orbit, dt, earth->mass); |
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Vec3 pos_final; |
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Vec3 vel_final; |
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orbital_elements_to_cartesian(propagated, earth->mass, &pos_final, &vel_final); |
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double pos_error = vec3_distance(pos_initial, pos_final); |
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double vel_error = vec3_distance(vel_initial, vel_final); |
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double num_periods = dt / period_seconds; |
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double expected_num_orbits = round(num_periods); |
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double fractional_phase = num_periods - expected_num_orbits; |
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double expected_pos_error = fractional_phase * 2.0 * M_PI * a; |
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INFO(" Position error: " << pos_error << " m"); |
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INFO(" Expected error (phase): " << expected_pos_error << " m"); |
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INFO(" Number of periods: " << num_periods); |
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if (expected_num_orbits > 0 && expected_pos_error > 1.0e-6) { |
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double relative_error = pos_error / expected_pos_error; |
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INFO(" Relative error: " << relative_error); |
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REQUIRE(relative_error < 0.5); |
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} else if (expected_num_orbits > 0) { |
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INFO(" Expected error is zero, skipping relative error check"); |
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REQUIRE(pos_error < 1.0e-3); |
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} |
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} |
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destroy_simulation(sim); |
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} |
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TEST_CASE("Mean anomaly accumulation over long propagation", "[analytical][timestep][accumulation]") { |
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const double TIME_STEP = 60.0; |
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SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP); |
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REQUIRE(load_system_config(sim, "tests/test_analytical_propagation_timesteps.toml")); |
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Spacecraft* craft = &sim->spacecraft[0]; |
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CelestialBody* earth = &sim->bodies[0]; |
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double a = craft->orbit.semi_major_axis; |
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double mu = G * earth->mass; |
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double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); |
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double mean_motion = sqrt(mu / pow(a, 3.0)); |
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double initial_true_anomaly = craft->orbit.true_anomaly; |
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INFO("Initial true anomaly: " << initial_true_anomaly << " rad"); |
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double propagation_time = period_seconds * 100.0; |
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INFO("Propagation time: " << propagation_time << " s (" << propagation_time / period_seconds << " periods)"); |
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OrbitalElements propagated = propagate_orbital_elements(craft->orbit, propagation_time, earth->mass); |
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double final_true_anomaly = propagated.true_anomaly; |
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INFO("Final true anomaly: " << final_true_anomaly << " rad"); |
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double expected_delta_anomaly = mean_motion * propagation_time; |
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double expected_final_anomaly = fmod(initial_true_anomaly + expected_delta_anomaly, 2.0 * M_PI); |
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INFO("Expected final anomaly: " << expected_final_anomaly << " rad"); |
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double raw_error = fabs(final_true_anomaly - expected_final_anomaly); |
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double anomaly_error = fmin(raw_error, 2.0 * M_PI - raw_error); |
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INFO("True anomaly error: " << anomaly_error << " rad (" << anomaly_error * 180.0 / M_PI << "°)"); |
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REQUIRE(anomaly_error < 1.0e-3); |
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destroy_simulation(sim); |
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} |
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# Test Configuration: Standard Orbit for Timestep Testing |
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# Moderate eccentricity orbit for testing various timestep sizes |
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[[bodies]] |
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name = "Earth" |
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mass = 5.972e24 |
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radius = 6.371e6 |
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parent_index = -1 |
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color = { r = 0.0, g = 0.5, b = 1.0 } |
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orbit = { |
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semi_major_axis = 0.0, |
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eccentricity = 0.0, |
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true_anomaly = 0.0 |
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} |
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[[spacecraft]] |
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name = "Standard_Orbit_Spacecraft" |
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mass = 1000.0 |
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parent_index = 0 |
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orbit = { |
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semi_major_axis = 1.5e7, |
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eccentricity = 0.4, |
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true_anomaly = 0.0, |
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inclination = 0.0, |
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longitude_of_ascending_node = 0.0, |
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argument_of_periapsis = 0.0 |
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} |
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