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cleanup: remove old test files superseded by tests/test_extreme_eccentricity

test-refactor
cinnaboot 2 months ago
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2 changed files with 0 additions and 278 deletions
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  1. 225
      old_tests/test_extreme_eccentricity.cpp
  2. 53
      old_tests/test_extreme_eccentricity.toml

225
old_tests/test_extreme_eccentricity.cpp
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#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 VELOCITY_TOLERANCE = 1.0e-6;
const double POSITION_TOLERANCE = 1.0e3;
TEST_CASE("Highly eccentric orbit (e=0.99)", "[extreme][eccentricity][high]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 3, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_extreme_eccentricity.toml"));
Spacecraft* high_e = &sim->spacecraft[0];
CelestialBody* earth = &sim->bodies[0];
INFO("Testing spacecraft with e=" << high_e->orbit.eccentricity);
Vec3 pos;
Vec3 vel;
orbital_elements_to_cartesian(high_e->orbit, earth->mass, &pos, &vel);
double r = vec3_magnitude(pos);
double v = vec3_magnitude(vel);
double expected_r_perigee = high_e->orbit.semi_major_axis * (1.0 - high_e->orbit.eccentricity);
double expected_r_apogee = high_e->orbit.semi_major_axis * (1.0 + high_e->orbit.eccentricity);
INFO("Semi-major axis: " << high_e->orbit.semi_major_axis << " m");
INFO("Eccentricity: " << high_e->orbit.eccentricity);
INFO("Radius: " << r << " m");
INFO("Velocity: " << v << " m/s");
INFO("Expected perigee: " << expected_r_perigee << " m");
INFO("Expected apogee: " << expected_r_apogee << " m");
REQUIRE(r >= expected_r_perigee * 0.9);
REQUIRE(r <= expected_r_apogee * 1.1);
REQUIRE(v > 0.0);
destroy_simulation(sim);
}
TEST_CASE("Near-parabolic orbit (e=0.9999)", "[extreme][eccentricity][near_parabolic]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 3, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_extreme_eccentricity.toml"));
Spacecraft* near_parabolic = &sim->spacecraft[1];
CelestialBody* earth = &sim->bodies[0];
INFO("Testing spacecraft with e=" << near_parabolic->orbit.eccentricity);
Vec3 pos_perigee;
Vec3 vel_perigee;
near_parabolic->orbit.true_anomaly = 0.0;
orbital_elements_to_cartesian(near_parabolic->orbit, earth->mass, &pos_perigee, &vel_perigee);
double r_perigee = vec3_magnitude(pos_perigee);
double v_perigee = vec3_magnitude(vel_perigee);
Vec3 pos_apogee;
Vec3 vel_apogee;
near_parabolic->orbit.true_anomaly = M_PI;
orbital_elements_to_cartesian(near_parabolic->orbit, earth->mass, &pos_apogee, &vel_apogee);
double r_apogee = vec3_magnitude(pos_apogee);
double v_apogee = vec3_magnitude(vel_apogee);
double expected_r_perigee = near_parabolic->orbit.semi_major_axis * (1.0 - near_parabolic->orbit.eccentricity);
double expected_r_apogee = near_parabolic->orbit.semi_major_axis * (1.0 + near_parabolic->orbit.eccentricity);
INFO("Perigee:");
INFO(" Radius: " << r_perigee << " m (expected: " << expected_r_perigee << " m)");
INFO(" Velocity: " << v_perigee << " m/s");
INFO("Apogee:");
INFO(" Radius: " << r_apogee << " m (expected: " << expected_r_apogee << " m)");
INFO(" Velocity: " << v_apogee << " m/s");
double r_perigee_error = fabs(r_perigee - expected_r_perigee);
double r_apogee_error = fabs(r_apogee - expected_r_apogee);
REQUIRE(r_perigee_error < POSITION_TOLERANCE);
REQUIRE(r_apogee_error < POSITION_TOLERANCE);
REQUIRE(v_perigee > v_apogee);
REQUIRE(r_apogee > r_perigee);
destroy_simulation(sim);
}
TEST_CASE("Near-parabolic boundary (e=1.0001)", "[extreme][eccentricity][boundary]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 3, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_extreme_eccentricity.toml"));
Spacecraft* hyperbolic = &sim->spacecraft[2];
CelestialBody* earth = &sim->bodies[0];
INFO("Testing spacecraft with e=" << hyperbolic->orbit.eccentricity);
Vec3 pos;
Vec3 vel;
orbital_elements_to_cartesian(hyperbolic->orbit, earth->mass, &pos, &vel);
double r = vec3_magnitude(pos);
double v = vec3_magnitude(vel);
double mu = G * earth->mass;
double a = hyperbolic->orbit.semi_major_axis;
double escape_velocity = sqrt(2.0 * mu / r);
double circular_velocity = sqrt(mu / r);
INFO("Radius: " << r << " m");
INFO("Velocity: " << v << " m/s");
INFO("Escape velocity: " << escape_velocity << " m/s");
INFO("Circular velocity: " << circular_velocity << " m/s");
INFO("Semi-major axis: " << a << " m");
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;
INFO("Expected velocity: " << expected_v << " m/s");
INFO("Velocity error: " << v_error << " m/s (" << relative_error * 100.0 << "%)");
REQUIRE(relative_error < VELOCITY_TOLERANCE);
REQUIRE(v > escape_velocity * 0.9);
REQUIRE(a < 0.0);
destroy_simulation(sim);
}
TEST_CASE("Velocity magnitude accuracy for extreme eccentricities", "[extreme][eccentricity][velocity]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 3, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_extreme_eccentricity.toml"));
CelestialBody* earth = &sim->bodies[0];
for (int i = 0; i < sim->craft_count; i++) {
Spacecraft* craft = &sim->spacecraft[i];
INFO("Spacecraft " << i << ": e=" << craft->orbit.eccentricity);
double true_anomalies[] = {0.0, M_PI / 2.0, M_PI, 3.0 * M_PI / 2.0};
for (int j = 0; j < 4; j++) {
double nu = true_anomalies[j];
// For hyperbolic orbits (e > 1), skip invalid true anomalies
// Valid range: |ν| < arccos(-1/e)
if (craft->orbit.eccentricity > 1.0) {
double max_nu = acos(-1.0 / craft->orbit.eccentricity);
if (fabs(nu) >= max_nu) {
INFO(" ν=" << nu << " rad: skipped (exceeds hyperbolic limit ±" << max_nu << " rad)");
continue;
}
}
craft->orbit.true_anomaly = nu;
Vec3 pos;
Vec3 vel;
orbital_elements_to_cartesian(craft->orbit, earth->mass, &pos, &vel);
double r = vec3_magnitude(pos);
double v = vec3_magnitude(vel);
double a = craft->orbit.semi_major_axis;
double mu = G * earth->mass;
double expected_v_squared = mu * (2.0 / r - 1.0 / a);
if (expected_v_squared > 0.0) {
double expected_v = sqrt(expected_v_squared);
double v_error = fabs(v - expected_v);
double relative_error = v_error / expected_v;
INFO(" ν=" << nu << " rad: v=" << v << " m/s, error=" << relative_error * 100.0 << "%");
REQUIRE(relative_error < VELOCITY_TOLERANCE * 10.0);
}
}
}
destroy_simulation(sim);
}
TEST_CASE("Period calculation (or lack thereof) for e≥1", "[extreme][eccentricity][period]") {
const double TIME_STEP = 60.0;
SimulationState* sim = create_simulation(10, 3, 0, TIME_STEP);
REQUIRE(load_system_config(sim, "tests/test_extreme_eccentricity.toml"));
Spacecraft* high_e = &sim->spacecraft[0];
Spacecraft* near_parabolic = &sim->spacecraft[1];
Spacecraft* hyperbolic = &sim->spacecraft[2];
double a_e = high_e->orbit.semi_major_axis;
double a_near = near_parabolic->orbit.semi_major_axis;
double a_h = hyperbolic->orbit.semi_major_axis;
INFO("Highly eccentric (e=0.99): a=" << a_e << " m");
INFO("Near-parabolic (e=0.9999): a=" << a_near << " m");
INFO("Hyperbolic (e=1.0001): a=" << a_h << " m");
REQUIRE(a_e > 0.0);
REQUIRE(a_near > 0.0);
REQUIRE(a_h < 0.0);
destroy_simulation(sim);
}

53
old_tests/test_extreme_eccentricity.toml
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# Test Configuration: Extreme Eccentricity Orbits
# Tests near-parabolic and hyperbolic orbits
[[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 = "Highly_Elliptical"
mass = 1000.0
parent_index = 0
orbit = {
semi_major_axis = 6.5e8,
eccentricity = 0.99,
true_anomaly = 0.0,
inclination = 0.0,
longitude_of_ascending_node = 0.0,
argument_of_periapsis = 0.0
}
[[spacecraft]]
name = "Near_Parabolic"
mass = 1000.0
parent_index = 0
orbit = {
semi_major_axis = 7.0e8,
eccentricity = 0.99,
true_anomaly = 0.0,
inclination = 0.0,
longitude_of_ascending_node = 0.0,
argument_of_periapsis = 0.0
}
[[spacecraft]]
name = "Slightly_Hyperbolic"
mass = 1000.0
parent_index = 0
orbit = {
semi_major_axis = -1.3e8,
eccentricity = 1.05,
true_anomaly = 0.0,
inclination = 0.0,
longitude_of_ascending_node = 0.0,
argument_of_periapsis = 0.0
}
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