eMeSinE(DerivativeStructure) | | 0% | | 0% | 2 | 2 | 10 | 10 | 1 | 1 |
getEccentricAnomaly(DerivativeStructure) | | 88% | | 75% | 3 | 7 | 4 | 30 | 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) | | 80% | | 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 | 8 | 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 |