From 8df95cf55aeeb1040d6ad51fbdcf4bbb2f4c2ad0 Mon Sep 17 00:00:00 2001 From: cinnaboot Date: Sat, 2 May 2026 19:38:40 -0400 Subject: [PATCH] cleanup: remove old analytical_propagation files, update continue.md status --- continue.md | 2 +- old_tests/test_analytical_propagation.cpp | 447 --------------------- old_tests/test_analytical_propagation.toml | 41 -- 3 files changed, 1 insertion(+), 489 deletions(-) delete mode 100644 old_tests/test_analytical_propagation.cpp delete mode 100644 old_tests/test_analytical_propagation.toml diff --git a/continue.md b/continue.md index ea9dd09..fef8c73 100644 --- a/continue.md +++ b/continue.md @@ -118,9 +118,9 @@ - `test_extreme_eccentricity` ✅ — High-eccentricity orbits (single SCENARIO, precalculated values, REL_TOL) - `test_extreme_orientation_mixed` ✅ — Extreme orientation conversions, rotation matrix properties, singularity handling - `test_extreme_timescales` ✅ — 9 TEST_CASEs → 1 SCENARIO with 11 SECTIONs, all WithinAbs use named constants +- `test_analytical_propagation` ✅ — 5 SCENARIOs → 1 SCENARIO with 23 SECTIONs, precalculated values, all WithinAbs use named constants ### Can Refactor Now (sim_engine.py supports all features needed) -- `test_analytical_propagation` — propagation through apsides - `test_moon_orbits` — multi-body propagation - `test_periapsis_burn` — prograde burns - `test_hybrid_burns` — impulse burns diff --git a/old_tests/test_analytical_propagation.cpp b/old_tests/test_analytical_propagation.cpp deleted file mode 100644 index 33130af..0000000 --- a/old_tests/test_analytical_propagation.cpp +++ /dev/null @@ -1,447 +0,0 @@ -#include -#include "../src/physics.h" -#include "../src/orbital_mechanics.h" -#include "../src/simulation.h" -#include "../src/config_loader.h" -#include "../src/test_utilities.h" -#include - -const double VELOCITY_TOLERANCE_APSIDES = 1.0; -const double POSITION_TOLERANCE_APSIDES = 1.0e3; -const double VELOCITY_TOLERANCE_TIMESTEP = 10.0; -const double POSITION_TOLERANCE_TIMESTEP = 1.0e4; - -TEST_CASE("Propagation through perigee (velocity maximum)", "[analytical][propagation][perigee]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[0]; - CelestialBody* earth = &sim->bodies[0]; - - Vec3 pos_before; - Vec3 vel_before; - - craft->orbit.true_anomaly = M_PI / 4.0; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_before, &vel_before); - - double v_before = vec3_magnitude(vel_before); - double r_before = vec3_magnitude(pos_before); - - INFO("Before perigee:"); - INFO(" Position: (" << pos_before.x << ", " << pos_before.y << ", " << pos_before.z << ") m"); - INFO(" Velocity: (" << vel_before.x << ", " << vel_before.y << ", " << vel_before.z << ") m/s"); - INFO(" Velocity magnitude: " << v_before << " m/s"); - INFO(" Radius: " << r_before << " m"); - - Vec3 pos_perigee; - Vec3 vel_perigee; - - craft->orbit.true_anomaly = 0.0; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_perigee, &vel_perigee); - - double v_perigee = vec3_magnitude(vel_perigee); - double r_perigee = vec3_magnitude(pos_perigee); - - INFO("At perigee (v=0):"); - INFO(" Position: (" << pos_perigee.x << ", " << pos_perigee.y << ", " << pos_perigee.z << ") m"); - INFO(" Velocity: (" << vel_perigee.x << ", " << vel_perigee.y << ", " << vel_perigee.z << ") m/s"); - INFO(" Velocity magnitude: " << v_perigee << " m/s"); - INFO(" Radius: " << r_perigee << " m"); - - double expected_r_perigee = craft->orbit.semi_major_axis * (1.0 - craft->orbit.eccentricity); - INFO("Expected radius at perigee: " << expected_r_perigee << " m"); - - double r_error = fabs(r_perigee - expected_r_perigee); - INFO("Radius error: " << r_error << " m"); - - REQUIRE(r_error < POSITION_TOLERANCE_APSIDES); - REQUIRE(v_perigee > v_before); - - destroy_simulation(sim); -} - -TEST_CASE("Propagation through apogee (velocity minimum)", "[analytical][propagation][apogee]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[0]; - CelestialBody* earth = &sim->bodies[0]; - - Vec3 pos_perigee; - Vec3 vel_perigee; - craft->orbit.true_anomaly = 0.0; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_perigee, &vel_perigee); - - double v_perigee = vec3_magnitude(vel_perigee); - double r_perigee = vec3_magnitude(pos_perigee); - - INFO("At perigee:"); - INFO(" Velocity magnitude: " << v_perigee << " m/s"); - INFO(" Radius: " << r_perigee << " m"); - - Vec3 pos_apogee; - Vec3 vel_apogee; - craft->orbit.true_anomaly = M_PI; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_apogee, &vel_apogee); - - double v_apogee = vec3_magnitude(vel_apogee); - double r_apogee = vec3_magnitude(pos_apogee); - - INFO("At apogee (v=π):"); - INFO(" Position: (" << pos_apogee.x << ", " << pos_apogee.y << ", " << pos_apogee.z << ") m"); - INFO(" Velocity: (" << vel_apogee.x << ", " << vel_apogee.y << ", " << vel_apogee.z << ") m/s"); - INFO(" Velocity magnitude: " << v_apogee << " m/s"); - INFO(" Radius: " << r_apogee << " m"); - - double expected_r_apogee = craft->orbit.semi_major_axis * (1.0 + craft->orbit.eccentricity); - INFO("Expected radius at apogee: " << expected_r_apogee << " m"); - - double r_error = fabs(r_apogee - expected_r_apogee); - INFO("Radius error: " << r_error << " m"); - - REQUIRE(r_error < POSITION_TOLERANCE_APSIDES); - REQUIRE(v_apogee < v_perigee); - REQUIRE(r_apogee > r_perigee); - - destroy_simulation(sim); -} - -TEST_CASE("Propagation returns to initial state after one orbital period", "[analytical][propagation][period]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[0]; - CelestialBody* earth = &sim->bodies[0]; - - double a = craft->orbit.semi_major_axis; - double mu = G * earth->mass; - double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); - - INFO("Semi-major axis: " << a << " m"); - INFO("Orbital period: " << period_seconds << " s (" << period_seconds / 3600.0 << " hours)"); - - Vec3 pos_initial; - Vec3 vel_initial; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_initial, &vel_initial); - - INFO("Initial position: (" << pos_initial.x << ", " << pos_initial.y << ", " << pos_initial.z << ") m"); - INFO("Initial velocity: (" << vel_initial.x << ", " << vel_initial.y << ", " << vel_initial.z << ") m/s"); - - OrbitalElements final_elements = propagate_orbital_elements(craft->orbit, period_seconds, earth->mass); - - Vec3 pos_final; - Vec3 vel_final; - orbital_elements_to_cartesian(final_elements, earth->mass, &pos_final, &vel_final); - - INFO("Final position: (" << pos_final.x << ", " << pos_final.y << ", " << pos_final.z << ") m"); - INFO("Final velocity: (" << vel_final.x << ", " << vel_final.y << ", " << vel_final.z << ") m/s"); - - double pos_error = vec3_distance(pos_initial, pos_final); - double vel_error = vec3_distance(vel_initial, vel_final); - - INFO("Position error after one period: " << pos_error << " m"); - INFO("Velocity error after one period: " << vel_error << " m/s"); - - double r_initial = vec3_magnitude(pos_initial); - double r_final = vec3_magnitude(pos_final); - double relative_pos_error = pos_error / r_initial * 100.0; - - double v_initial = vec3_magnitude(vel_initial); - double v_final = vec3_magnitude(vel_final); - double relative_vel_error = vel_error / v_initial * 100.0; - - INFO("Relative position error: " << relative_pos_error << "%"); - INFO("Relative velocity error: " << relative_vel_error << "%"); - - REQUIRE(relative_pos_error < 0.1); - REQUIRE(relative_vel_error < 0.1); - - destroy_simulation(sim); -} - -TEST_CASE("True anomaly accuracy after full orbit", "[analytical][propagation][true_anomaly]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[0]; - CelestialBody* earth = &sim->bodies[0]; - - double initial_true_anomaly = craft->orbit.true_anomaly; - - INFO("Initial true anomaly: " << initial_true_anomaly << " rad (" << initial_true_anomaly * 180.0 / M_PI << "°)"); - - double a = craft->orbit.semi_major_axis; - double mu = G * earth->mass; - double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); - - OrbitalElements final_elements = propagate_orbital_elements(craft->orbit, period_seconds, earth->mass); - - double final_true_anomaly = final_elements.true_anomaly; - - INFO("Final true anomaly: " << final_true_anomaly << " rad (" << final_true_anomaly * 180.0 / M_PI << "°)"); - - double expected_true_anomaly = fmod(initial_true_anomaly + 2.0 * M_PI, 2.0 * M_PI); - double anomaly_error = fabs(final_true_anomaly - expected_true_anomaly); - - INFO("Expected true anomaly: " << expected_true_anomaly << " rad"); - INFO("True anomaly error: " << anomaly_error << " rad (" << anomaly_error * 180.0 / M_PI << "°)"); - - REQUIRE(anomaly_error < 1.0e-6); - - destroy_simulation(sim); -} - -TEST_CASE("Vis-viva equation holds at multiple points in orbit", "[analytical][propagation][vis_viva]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[0]; - CelestialBody* earth = &sim->bodies[0]; - - double a = craft->orbit.semi_major_axis; - double mu = G * earth->mass; - - double true_anomalies[] = {0.0, M_PI / 4.0, M_PI / 2.0, 3.0 * M_PI / 4.0, M_PI}; - - for (int i = 0; i < 5; i++) { - double nu = true_anomalies[i]; - INFO("Testing at true anomaly: " << nu << " rad (" << nu * 180.0 / M_PI << "°)"); - - craft->orbit.true_anomaly = nu; - - Vec3 position; - Vec3 velocity; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &position, &velocity); - - double r = vec3_magnitude(position); - double v = vec3_magnitude(velocity); - - double expected_v_squared = mu * (2.0 / r - 1.0 / a); - double expected_v = sqrt(expected_v_squared); - - double v_error = fabs(v - expected_v); - double relative_error = v_error / expected_v * 100.0; - - INFO(" Radius: " << r << " m"); - INFO(" Actual velocity: " << v << " m/s"); - INFO(" Expected velocity: " << expected_v << " m/s"); - INFO(" Error: " << v_error << " m/s (" << relative_error << "%)"); - - REQUIRE(relative_error < 0.01); - } - - destroy_simulation(sim); -} - -TEST_CASE("Large timestep - dt greater than orbital period", "[analytical][timestep][large]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[1]; - CelestialBody* earth = &sim->bodies[0]; - - double a = craft->orbit.semi_major_axis; - double mu = G * earth->mass; - double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); - - INFO("Orbital period: " << period_seconds << " s (" << period_seconds / 3600.0 << " hours)"); - - double large_dt = period_seconds * 2.0; - INFO("Timestep: " << large_dt << " s (2x orbital period)"); - - Vec3 pos_before; - Vec3 vel_before; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_before, &vel_before); - - OrbitalElements propagated = propagate_orbital_elements(craft->orbit, large_dt, earth->mass); - - Vec3 pos_after; - Vec3 vel_after; - orbital_elements_to_cartesian(propagated, earth->mass, &pos_after, &vel_after); - - double r_before = vec3_magnitude(pos_before); - double r_after = vec3_magnitude(pos_after); - double v_before = vec3_magnitude(vel_before); - double v_after = vec3_magnitude(vel_after); - - INFO("Before propagation:"); - INFO(" Radius: " << r_before << " m"); - INFO(" Velocity: " << v_before << " m/s"); - - INFO("After 2 periods:"); - INFO(" Radius: " << r_after << " m"); - INFO(" Velocity: " << v_after << " m/s"); - - double r_error = fabs(r_after - r_before); - double v_error = fabs(v_after - v_before); - double relative_r_error = r_error / r_before * 100.0; - double relative_v_error = v_error / v_before * 100.0; - - INFO("Radius error: " << r_error << " m (" << relative_r_error << "%)"); - INFO("Velocity error: " << v_error << " m/s (" << relative_v_error << "%)"); - - REQUIRE(relative_r_error < 0.1); - REQUIRE(relative_v_error < 0.1); - - destroy_simulation(sim); -} - -TEST_CASE("Very small timestep - dt less than 1 second", "[analytical][timestep][small]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[1]; - CelestialBody* earth = &sim->bodies[0]; - - Vec3 pos_before; - Vec3 vel_before; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_before, &vel_before); - - double small_dt = 0.1; - INFO("Timestep: " << small_dt << " s"); - - OrbitalElements propagated = propagate_orbital_elements(craft->orbit, small_dt, earth->mass); - - Vec3 pos_after; - Vec3 vel_after; - orbital_elements_to_cartesian(propagated, earth->mass, &pos_after, &vel_after); - - double pos_change = vec3_distance(pos_before, pos_after); - double vel_change = vec3_distance(vel_before, vel_after); - - INFO("Position change: " << pos_change << " m"); - INFO("Velocity change: " << vel_change << " m/s"); - - double v_before_mag = vec3_magnitude(vel_before); - double expected_pos_change = v_before_mag * small_dt; - double pos_error = fabs(pos_change - expected_pos_change); - - INFO("Expected position change (v·dt): " << expected_pos_change << " m"); - INFO("Position error: " << pos_error << " m"); - INFO("Relative position error: " << (pos_error / expected_pos_change * 100.0) << "%"); - - REQUIRE(pos_error < VELOCITY_TOLERANCE_TIMESTEP * small_dt * 10.0); - REQUIRE(vel_change < VELOCITY_TOLERANCE_TIMESTEP); - - destroy_simulation(sim); -} - -TEST_CASE("Accuracy vs timestep size relationship", "[analytical][timestep][accuracy]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[1]; - CelestialBody* earth = &sim->bodies[0]; - - double a = craft->orbit.semi_major_axis; - double mu = G * earth->mass; - double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); - - double dt_ratios[] = {0.01, 0.1, 1.0, 10.0}; - - Vec3 pos_initial; - Vec3 vel_initial; - orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos_initial, &vel_initial); - - for (int i = 0; i < 4; i++) { - double dt = period_seconds * dt_ratios[i]; - INFO("Testing dt = " << dt << " s (" << dt_ratios[i] << "x period)"); - - OrbitalElements propagated = propagate_orbital_elements(craft->orbit, dt, earth->mass); - - Vec3 pos_final; - Vec3 vel_final; - orbital_elements_to_cartesian(propagated, earth->mass, &pos_final, &vel_final); - - double pos_error = vec3_distance(pos_initial, pos_final); - double vel_error = vec3_distance(vel_initial, vel_final); - - double num_periods = dt / period_seconds; - double expected_num_orbits = round(num_periods); - - double fractional_phase = num_periods - expected_num_orbits; - double expected_pos_error = fractional_phase * 2.0 * M_PI * a; - - INFO(" Position error: " << pos_error << " m"); - INFO(" Expected error (phase): " << expected_pos_error << " m"); - INFO(" Number of periods: " << num_periods); - - if (expected_num_orbits > 0 && expected_pos_error > 1.0e-6) { - double relative_error = pos_error / expected_pos_error; - - INFO(" Relative error: " << relative_error); - - REQUIRE(relative_error < 0.5); - } else if (expected_num_orbits > 0) { - INFO(" Expected error is zero, skipping relative error check"); - REQUIRE(pos_error < 1.0e-3); - } - } - - destroy_simulation(sim); -} - -TEST_CASE("Mean anomaly accumulation over long propagation", "[analytical][timestep][accumulation]") { - const double TIME_STEP = 60.0; - - SimulationState* sim = create_simulation(10, 2, 0, TIME_STEP); - - REQUIRE(load_system_config(sim, "tests/test_analytical_propagation.toml")); - - Spacecraft* craft = &sim->spacecraft[1]; - CelestialBody* earth = &sim->bodies[0]; - - double a = craft->orbit.semi_major_axis; - double mu = G * earth->mass; - double period_seconds = 2.0 * M_PI * sqrt(pow(a, 3.0) / mu); - double mean_motion = sqrt(mu / pow(a, 3.0)); - - double initial_true_anomaly = craft->orbit.true_anomaly; - INFO("Initial true anomaly: " << initial_true_anomaly << " rad"); - - double propagation_time = period_seconds * 100.0; - INFO("Propagation time: " << propagation_time << " s (" << propagation_time / period_seconds << " periods)"); - - OrbitalElements propagated = propagate_orbital_elements(craft->orbit, propagation_time, earth->mass); - - double final_true_anomaly = propagated.true_anomaly; - INFO("Final true anomaly: " << final_true_anomaly << " rad"); - - double expected_delta_anomaly = mean_motion * propagation_time; - double expected_final_anomaly = fmod(initial_true_anomaly + expected_delta_anomaly, 2.0 * M_PI); - - INFO("Expected final anomaly: " << expected_final_anomaly << " rad"); - - double raw_error = fabs(final_true_anomaly - expected_final_anomaly); - double anomaly_error = fmin(raw_error, 2.0 * M_PI - raw_error); - - INFO("True anomaly error: " << anomaly_error << " rad (" << anomaly_error * 180.0 / M_PI << "°)"); - - REQUIRE(anomaly_error < 1.0e-3); - - destroy_simulation(sim); -} diff --git a/old_tests/test_analytical_propagation.toml b/old_tests/test_analytical_propagation.toml deleted file mode 100644 index 7120673..0000000 --- a/old_tests/test_analytical_propagation.toml +++ /dev/null @@ -1,41 +0,0 @@ -# Test Configuration: Analytical Propagation Tests -# Combined configuration for apsides and timestep testing -# Contains two spacecraft with different orbital parameters - -[[bodies]] -name = "Earth" -mass = 5.972e24 -radius = 6.371e6 -parent_index = -1 -color = { r = 0.0, g = 0.5, b = 1.0 } -orbit = { - semi_major_axis = 0.0, - eccentricity = 0.0, - true_anomaly = 0.0 -} - -[[spacecraft]] -name = "Apsides_Test_Spacecraft" -mass = 1000.0 -parent_index = 0 -orbit = { - semi_major_axis = 2.0e7, - eccentricity = 0.6, - true_anomaly = 0.0, - inclination = 0.0, - longitude_of_ascending_node = 0.0, - argument_of_periapsis = 0.0 -} - -[[spacecraft]] -name = "Timestep_Test_Spacecraft" -mass = 1000.0 -parent_index = 0 -orbit = { - semi_major_axis = 1.5e7, - eccentricity = 0.4, - true_anomaly = 0.0, - inclination = 0.0, - longitude_of_ascending_node = 0.0, - argument_of_periapsis = 0.0 -}