| eMeSinE(DerivativeStructure) |  | 0% | | 0% | 2 | 2 | 10 | 10 | 1 | 1 |
| getEccentricAnomaly(DerivativeStructure) |  | 90% | | 80% | 2 | 6 | 3 | 29 | 0 | 1 |
| resetInitialState(SpacecraftState) | | 0% | | n/a | 1 | 1 | 1 | 1 | 1 | 1 |
| resetIntermediateState(SpacecraftState, boolean) | | 0% | | n/a | 1 | 1 | 1 | 1 | 1 | 1 |
| getTk(AbsoluteDate) | | 81% | | 75% | 1 | 3 | 1 | 6 | 0 | 1 |
| getFrame() | | 0% | | n/a | 1 | 1 | 1 | 1 | 1 | 1 |
| getMU() | | 0% | | n/a | 1 | 1 | 1 | 1 | 1 | 1 |
| propagateInEcef(AbsoluteDate) | | 100% | | n/a | 0 | 1 | 0 | 26 | 0 | 1 |
| GPSPropagator(GPSPropagator.Builder) | | 100% | | n/a | 0 | 1 | 0 | 7 | 0 | 1 |
| getTrueAnomaly(DerivativeStructure) | | 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 |
| getGPSOrbitalElements() | | 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 |
| getMass(AbsoluteDate) | | 100% | | n/a | 0 | 1 | 0 | 1 | 0 | 1 |