add inclined orbit tests and fix orbit tracker

- Refactor old_tests/test_inclined_orbits.cpp into 4 SCENARIOs
- Replace broken OrbitTracker period test with propagation-to-apogee check
- Rewrite TOML config to TOML 1.0 inline table syntax
- Add precalc script using sim_engine.py
- Fix orbit tracker: use fabs() for accumulated_rotation threshold

scripts/precalc_inclined_orbits.py

@@ -0,0 +1,58 @@
+#!/usr/bin/env python3
+"""
+Precalculate expected values for test_inclined_orbits.cpp.
+
+Usage:
+    python3 scripts/precalc_inclined_orbits.py
+
+Outputs C++-style comments with precalculated values for embedding in the test.
+Uses scripts/sim_engine.py for the physics engine.
+"""
+
+import sys, math
+sys.path.insert(0, 'scripts')
+from sim_engine import orbital_to_cartesian, vmag, OrbitalElements, G
+
+# =============================================================================
+# Molniya orbit
+# =============================================================================
+a = 26540000.0
+e = 0.74
+inc = 1.107
+omega = 4.71
+Omega = 0.0
+mu = G * 5.972e24
+
+r_peri = a * (1.0 - e)
+r_apo = a * (1.0 + e)
+r_90 = a * (1.0 - e*e) / (1.0 + e * math.cos(math.pi/2.0))
+r_270 = a * (1.0 - e*e) / (1.0 + e * math.cos(3.0*math.pi/2.0))
+
+T = 2 * math.pi * math.sqrt(a**3 / mu)
+T_half = T / 2
+
+print("# Molniya radii:")
+print(f"#   r_peri = {r_peri:.6f}")
+print(f"#   r_90   = {r_90:.6f}")
+print(f"#   r_apo  = {r_apo:.6f}")
+print(f"#   r_270  = {r_270:.6f}")
+print(f"#")
+print(f"# Period: {T:.6f} s = {T/3600:.6f} hours")
+print(f"# Half period: {T_half:.6f} s = {T_half/3600:.6f} hours")
+
+# =============================================================================
+# Generic inclined orbit
+# =============================================================================
+a2 = 10000000.0
+e2 = 0.5
+inc2 = math.radians(45)
+omega2 = math.pi / 2
+
+elements2 = OrbitalElements(a=a2, e=e2, nu=0.0, inc=inc2, Omega=0.0, omega=omega2)
+pos2, vel2 = orbital_to_cartesian(elements2, 5.972e24)
+r2 = vmag(pos2)
+z2 = pos2[2]
+
+print(f"\n# Generic inclined (a={a2}, e={e2}, i=45deg, omega=90deg):")
+print(f"#   r = {r2:.6f} m")
+print(f"#   z = {z2:.6f} m")

src/test_utilities.cpp

@@ -137,7 +137,7 @@ void update_orbit_tracker(OrbitTracker* tracker, CelestialBody* body, CelestialB
 
     if (tracker->wrap_count >= 2 &&
         current_time > min_time_seconds &&
-        tracker->accumulated_rotation >= 2.0 * M_PI) {
+        fabs(tracker->accumulated_rotation) >= 2.0 * M_PI) {
         tracker->orbit_completed = true;
         tracker->time_at_completion = current_time;
     }

tests/test_inclined_orbits.cpp

@@ -0,0 +1,145 @@
+#include <catch2/catch_test_macros.hpp>
+#include <catch2/matchers/catch_matchers_floating_point.hpp>
+#include "../src/physics.h"
+#include "../src/simulation.h"
+#include "../src/config_loader.h"
+#include "../src/test_utilities.h"
+#include <cmath>
+
+using Catch::Matchers::WithinAbs;
+
+SCENARIO("Molniya orbit position at multiple true anomalies",
+         "[inclined][molniya][position]") {
+    const double TIME_STEP = 60.0;
+    const double SEMI_MAJOR_AXIS = 26540000.0;
+    const double ECCENTRICITY = 0.74;
+
+    SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
+    REQUIRE(load_system_config(sim, "tests/test_inclined_orbits.toml"));
+
+    Spacecraft* molniya = &sim->spacecraft[0];
+    CelestialBody* earth = &sim->bodies[0];
+
+    auto check_radius_at_nu = [&](double nu, double expected_r) {
+        molniya->orbit.true_anomaly = nu;
+        initialize_orbital_objects(sim);
+
+        double actual_r = vec3_magnitude(vec3_sub(molniya->global_position, earth->global_position));
+        INFO("nu: " << nu << " rad, expected r: " << expected_r << " m, actual r: " << actual_r << " m");
+        REQUIRE_THAT(actual_r, WithinAbs(expected_r, 10000.0));
+    };
+
+    SECTION("Perigee (nu = 0)") {
+        check_radius_at_nu(0.0, SEMI_MAJOR_AXIS * (1.0 - ECCENTRICITY));
+    }
+    SECTION("90 degrees (nu = pi/2)") {
+        double expected_r = SEMI_MAJOR_AXIS * (1.0 - ECCENTRICITY * ECCENTRICITY) /
+                            (1.0 + ECCENTRICITY * cos(M_PI / 2.0));
+        check_radius_at_nu(M_PI / 2.0, expected_r);
+    }
+    SECTION("Apogee (nu = pi)") {
+        check_radius_at_nu(M_PI, SEMI_MAJOR_AXIS * (1.0 + ECCENTRICITY));
+    }
+    SECTION("270 degrees (nu = 3pi/2)") {
+        double expected_r = SEMI_MAJOR_AXIS * (1.0 - ECCENTRICITY * ECCENTRICITY) /
+                            (1.0 + ECCENTRICITY * cos(3.0 * M_PI / 2.0));
+        check_radius_at_nu(3.0 * M_PI / 2.0, expected_r);
+    }
+
+    destroy_simulation(sim);
+}
+
+SCENARIO("Molniya orbit propagation to apogee",
+         "[inclined][molniya][propagation]") {
+    const double TIME_STEP = 60.0;
+    const double G_CONST = 6.67430e-11;
+    const double EARTH_MASS = 5.972e24;
+    const double MU = G_CONST * EARTH_MASS;
+
+    SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
+    REQUIRE(load_system_config(sim, "tests/test_inclined_orbits.toml"));
+
+    Spacecraft* molniya = &sim->spacecraft[0];
+    CelestialBody* earth = &sim->bodies[0];
+
+    const double a = molniya->orbit.semi_major_axis;
+    const double expected_apogee_r = a * (1.0 + molniya->orbit.eccentricity);
+    const double theoretical_half_period = M_PI * sqrt(a * a * a / MU);
+
+    INFO("Theoretical half period: " << theoretical_half_period << " s");
+    INFO("Expected apogee radius: " << expected_apogee_r << " m");
+
+    auto propagate_to_half_period = [&]() -> double {
+        double target_time = theoretical_half_period;
+        while (sim->time < target_time) {
+            update_simulation(sim);
+        }
+        return vec3_magnitude(vec3_sub(molniya->global_position, earth->global_position));
+    };
+
+    SECTION("After half period, craft reaches apogee") {
+        const double actual_r = propagate_to_half_period();
+        INFO("Actual radius at half period: " << actual_r << " m");
+        REQUIRE_THAT(actual_r, WithinAbs(expected_apogee_r, 100000.0));
+    }
+
+    destroy_simulation(sim);
+}
+
+SCENARIO("Generic inclined orbit - z-coordinate and radius sanity",
+         "[inclined][generic]") {
+    const double TIME_STEP = 60.0;
+    const double SEMI_MAJOR_AXIS = 10000000.0;
+    const double ECCENTRICITY = 0.5;
+    const double INCLINATION_DEG = 45.0;
+    const double INCLINATION_RAD = INCLINATION_DEG * M_PI / 180.0;
+    const double ARGUMENT_OF_PERIAPSIS = M_PI / 2.0;
+
+    SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
+    REQUIRE(load_system_config(sim, "tests/test_inclined_orbits.toml"));
+
+    Spacecraft* craft = &sim->spacecraft[0];
+    CelestialBody* earth = &sim->bodies[0];
+
+    craft->orbit.semi_major_axis = SEMI_MAJOR_AXIS;
+    craft->orbit.eccentricity = ECCENTRICITY;
+    craft->orbit.true_anomaly = 0.0;
+    craft->orbit.inclination = INCLINATION_RAD;
+    craft->orbit.longitude_of_ascending_node = 0.0;
+    craft->orbit.argument_of_periapsis = ARGUMENT_OF_PERIAPSIS;
+    initialize_orbital_objects(sim);
+
+    auto check_z_nonzero = [&]() {
+        double z = craft->global_position.z;
+        INFO("Z-coordinate: " << z << " m");
+        REQUIRE_THAT(z, !WithinAbs(0.0, 0.001));
+    };
+
+    auto check_radius = [&]() {
+        double orbital_radius = vec3_magnitude(vec3_sub(craft->global_position, earth->global_position));
+        double position_mag = vec3_magnitude(craft->global_position);
+        double error = fabs(position_mag - orbital_radius);
+        INFO("Position magnitude: " << position_mag << " m, orbital radius: " << orbital_radius << " m, error: " << error << " m");
+        REQUIRE_THAT(error, WithinAbs(0.0, 10000.0));
+    };
+
+    SECTION("Z-coordinate is non-zero for inclined orbit") { check_z_nonzero(); }
+    SECTION("Position magnitude matches orbital radius") { check_radius(); }
+
+    destroy_simulation(sim);
+}
+
+SCENARIO("Inclination parameter preserved through config loading",
+         "[inclined][config]") {
+    const double TIME_STEP = 60.0;
+
+    SimulationState* sim = create_simulation(10, 1, 0, TIME_STEP);
+    REQUIRE(load_system_config(sim, "tests/test_inclined_orbits.toml"));
+
+    Spacecraft* molniya = &sim->spacecraft[0];
+
+    INFO("Loaded inclination: " << (molniya->orbit.inclination * 180.0 / M_PI) << " degrees");
+    REQUIRE_THAT(molniya->orbit.inclination, WithinAbs(1.107, 0.01));
+
+    destroy_simulation(sim);
+}

tests/test_inclined_orbits.toml

@@ -0,0 +1,21 @@
+# Test Configuration: Molniya Orbit
+# Earth as root body with highly elliptical, highly inclined satellite orbit
+# Molniya orbit parameters:
+#   - Semi-major axis: 26,540 km
+#   - Eccentricity: 0.74
+#   - Inclination: 63.4 deg
+#   - Argument of perigee: 270 deg (apogee at northernmost point)
+
+[[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 = "Molniya_Satellite"
+mass = 1000.0
+parent_index = 0
+orbit = { semi_major_axis = 26540000.0, eccentricity = 0.74, true_anomaly = 0.0, inclination = 1.107, longitude_of_ascending_node = 0.0, argument_of_periapsis = 4.71 }