cinnaboot
/
claudes_game
1
0
Fork 0
Code Issues Pull Requests Projects Releases Wiki Activity
Browse Source

refactor: remove old_cartesian_to_elements_basic from old_tests/

test-refactor
cinnaboot 2 months ago
parent
9fd1e29947
commit
db5c781392
2 changed files with 0 additions and 217 deletions
Whitespace
Show all changes Ignore whitespace when comparing lines Ignore changes in amount of whitespace Ignore changes in whitespace at EOL
Split View
Diff Options
Show Stats Download Patch File Download Diff File
  1. 190
      old_tests/test_cartesian_to_elements_basic.cpp
  2. 27
      old_tests/test_cartesian_to_elements_basic.toml

190
old_tests/test_cartesian_to_elements_basic.cpp
Unescape View File

@ -1,190 +0,0 @@
#include <catch2/catch_test_macros.hpp>
#include "../src/physics.h"
#include "../src/orbital_mechanics.h"
#include "../src/simulation.h"
#include "../src/config_loader.h"
#include <cmath>
const double POSITION_TOLERANCE = 1.0e6;
const double VELOCITY_TOLERANCE = 10.0;
const double ELEMENT_TOLERANCE = 1.0e-6;
TEST_CASE("Round-trip conversion: orbital elements → state vectors → orbital elements", "[cartesian][elements][roundtrip]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_cartesian_to_elements_basic.toml"));
Spacecraft* craft = &sim->spacecraft[0];
OrbitalElements original_elements = craft->orbit;
Vec3 position_from_elements;
Vec3 velocity_from_elements;
orbital_elements_to_cartesian(original_elements, sim->bodies[0].mass, &position_from_elements, &velocity_from_elements);
INFO("Original orbital elements:");
INFO(" semi_major_axis: " << original_elements.semi_major_axis << " m");
INFO(" eccentricity: " << original_elements.eccentricity);
INFO(" true_anomaly: " << original_elements.true_anomaly << " rad");
INFO(" inclination: " << original_elements.inclination << " rad");
INFO(" longitude_of_ascending_node: " << original_elements.longitude_of_ascending_node << " rad");
INFO(" argument_of_periapsis: " << original_elements.argument_of_periapsis << " rad");
INFO("State vectors from orbital elements:");
INFO(" position: (" << position_from_elements.x << ", " << position_from_elements.y << ", " << position_from_elements.z << ") m");
INFO(" velocity: (" << velocity_from_elements.x << ", " << velocity_from_elements.y << ", " << velocity_from_elements.z << ") m/s");
OrbitalElements converted_elements = cartesian_to_orbital_elements(position_from_elements, velocity_from_elements, sim->bodies[0].mass);
INFO("Converted orbital elements:");
INFO(" semi_major_axis: " << converted_elements.semi_major_axis << " m");
INFO(" eccentricity: " << converted_elements.eccentricity);
INFO(" true_anomaly: " << converted_elements.true_anomaly << " rad");
INFO(" inclination: " << converted_elements.inclination << " rad");
INFO(" longitude_of_ascending_node: " << converted_elements.longitude_of_ascending_node << " rad");
INFO(" argument_of_periapsis: " << converted_elements.argument_of_periapsis << " rad");
double semi_major_error = fabs(converted_elements.semi_major_axis - original_elements.semi_major_axis);
double eccentricity_error = fabs(converted_elements.eccentricity - original_elements.eccentricity);
double inclination_error = fabs(converted_elements.inclination - original_elements.inclination);
INFO("Semi-major axis error: " << semi_major_error << " m");
INFO("Eccentricity error: " << eccentricity_error);
INFO("Inclination error: " << inclination_error << " rad");
REQUIRE(semi_major_error < fabs(original_elements.semi_major_axis) * ELEMENT_TOLERANCE);
REQUIRE(eccentricity_error < ELEMENT_TOLERANCE);
REQUIRE(inclination_error < ELEMENT_TOLERANCE);
destroy_simulation(sim);
}
TEST_CASE("Position magnitude preservation through conversion", "[cartesian][elements][position]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_cartesian_to_elements_basic.toml"));
Spacecraft* craft = &sim->spacecraft[0];
Vec3 position_1;
Vec3 velocity_1;
orbital_elements_to_cartesian(craft->orbit, sim->bodies[0].mass, &position_1, &velocity_1);
double radius_1 = vec3_magnitude(position_1);
INFO("Original radius: " << radius_1 << " m");
OrbitalElements elements = cartesian_to_orbital_elements(position_1, velocity_1, sim->bodies[0].mass);
Vec3 position_2;
Vec3 velocity_2;
orbital_elements_to_cartesian(elements, sim->bodies[0].mass, &position_2, &velocity_2);
double radius_2 = vec3_magnitude(position_2);
INFO("Reconstructed radius: " << radius_2 << " m");
double radius_error = fabs(radius_2 - radius_1);
INFO("Radius error: " << radius_error << " m");
REQUIRE(radius_error < POSITION_TOLERANCE);
destroy_simulation(sim);
}
TEST_CASE("Velocity magnitude preservation through conversion", "[cartesian][elements][velocity]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_cartesian_to_elements_basic.toml"));
Spacecraft* craft = &sim->spacecraft[0];
Vec3 position_1;
Vec3 velocity_1;
orbital_elements_to_cartesian(craft->orbit, sim->bodies[0].mass, &position_1, &velocity_1);
double v_mag_1 = vec3_magnitude(velocity_1);
INFO("Original velocity magnitude: " << v_mag_1 << " m/s");
OrbitalElements elements = cartesian_to_orbital_elements(position_1, velocity_1, sim->bodies[0].mass);
Vec3 position_2;
Vec3 velocity_2;
orbital_elements_to_cartesian(elements, sim->bodies[0].mass, &position_2, &velocity_2);
double v_mag_2 = vec3_magnitude(velocity_2);
INFO("Reconstructed velocity magnitude: " << v_mag_2 << " m/s");
double velocity_error = fabs(v_mag_2 - v_mag_1);
INFO("Velocity error: " << velocity_error << " m/s");
REQUIRE(velocity_error < VELOCITY_TOLERANCE);
destroy_simulation(sim);
}
TEST_CASE("Semi-major axis accuracy", "[cartesian][elements][semi_major]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_cartesian_to_elements_basic.toml"));
Spacecraft* craft = &sim->spacecraft[0];
double expected_a = craft->orbit.semi_major_axis;
Vec3 position;
Vec3 velocity;
orbital_elements_to_cartesian(craft->orbit, sim->bodies[0].mass, &position, &velocity);
OrbitalElements elements = cartesian_to_orbital_elements(position, velocity, sim->bodies[0].mass);
double actual_a = elements.semi_major_axis;
double a_error = fabs(actual_a - expected_a);
double relative_error = a_error / fabs(expected_a);
INFO("Expected semi-major axis: " << expected_a << " m");
INFO("Actual semi-major axis: " << actual_a << " m");
INFO("Absolute error: " << a_error << " m");
INFO("Relative error: " << relative_error * 100.0 << "%");
REQUIRE(relative_error < ELEMENT_TOLERANCE);
destroy_simulation(sim);
}
TEST_CASE("Eccentricity accuracy", "[cartesian][elements][eccentricity]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_cartesian_to_elements_basic.toml"));
Spacecraft* craft = &sim->spacecraft[0];
double expected_e = craft->orbit.eccentricity;
Vec3 position;
Vec3 velocity;
orbital_elements_to_cartesian(craft->orbit, sim->bodies[0].mass, &position, &velocity);
OrbitalElements elements = cartesian_to_orbital_elements(position, velocity, sim->bodies[0].mass);
double actual_e = elements.eccentricity;
double e_error = fabs(actual_e - expected_e);
INFO("Expected eccentricity: " << expected_e);
INFO("Actual eccentricity: " << actual_e);
INFO("Absolute error: " << e_error);
REQUIRE(e_error < ELEMENT_TOLERANCE);
destroy_simulation(sim);
}

27
old_tests/test_cartesian_to_elements_basic.toml
Unescape View File

@ -1,27 +0,0 @@
# Test Configuration: Basic Elliptical Orbit
# Moderate eccentricity, zero inclination for testing Cartesian ↔ orbital elements conversion
[[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 = "Test_Spacecraft"
mass = 1000.0
parent_index = 0
orbit = {
semi_major_axis = 1.5e7,
eccentricity = 0.5,
true_anomaly = 0.0,
inclination = 0.0,
longitude_of_ascending_node = 0.0,
argument_of_periapsis = 0.0
}
Write Preview
Loading…
Cancel
Save