| eMeSinE(UnivariateDerivative2) |   | 60% | | 50% | 1 | 2 | 4 | 10 | 0 | 1 |
| getEccentricAnomaly(UnivariateDerivative2) | | 98% | | 91% | 1 | 7 | 1 | 30 | 0 | 1 |
| propagateInEcef(AbsoluteDate) | | 100% | | n/a | 0 | 1 | 0 | 34 | 0 | 1 |
| GNSSPropagator(GNSSOrbitalElements, Frame, Frame, AttitudeProvider, double) | | 100% | | n/a | 0 | 1 | 0 | 10 | 0 | 1 |
| getTk(AbsoluteDate) | | 100% | | 100% | 0 | 3 | 0 | 7 | 0 | 1 |
| getTrueAnomaly(UnivariateDerivative2) | | 100% | | n/a | 0 | 1 | 0 | 3 | 0 | 1 |
| propagateOrbit(AbsoluteDate) | | 100% | | n/a | 0 | 1 | 0 | 3 | 0 | 1 |
| static {...} | | 100% | | n/a | 0 | 1 | 0 | 6 | 0 | 1 |
| resetInitialState(SpacecraftState) | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| resetIntermediateState(SpacecraftState, boolean) | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| getMU() | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| getECI() | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| getECEF() | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| getOrbitalElements() | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| getFrame() | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |
| getMass(AbsoluteDate) | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |