In the gravitational two-body problem, the specific orbital energy (or vis-viva energy) of two orbiting bodies is the constant sum of their mutual potential energy () and their total kinetic energy (), divided by the reduced mass. According to the orbital energy conservation equation (also referred to as vis-viva equation), it does not vary with time:
It is expressed in J/kg = m2?s-2 or MJ/kg = km2?s-2. For an elliptic orbit the specific orbital energy is the negative of the additional energy required to accelerate a mass of one kilogram to escape velocity (parabolic orbit). For a hyperbolic orbit, it is equal to the excess energy compared to that of a parabolic orbit. In this case the specific orbital energy is also referred to as characteristic energy.
For a parabolic orbit this equation simplifies to
For a hyperbolic trajectory this specific orbital energy is either given by
or the same as for an ellipse, depending on the convention for the sign of a.
In this case the specific orbital energy is also referred to as characteristic energy (or ) and is equal to the excess specific energy compared to that for a parabolic orbit.
It is relevant for interplanetary missions.
For an elliptic orbit the rate of change of the specific orbital energy with respect to a change in the semi-major axis is
In the case of circular orbits, this rate is one half of the gravitation at the orbit. This corresponds to the fact that for such orbits the total energy is one half of the potential energy, because the kinetic energy is minus one half of the potential energy.
If the central body has radius R, then the additional specific energy of an elliptic orbit compared to being stationary at the surface is
The International Space Station has an orbital period of 91.74 minutes (5504s), hence the semi-major axis is 6,738km.
The energy is -29.6MJ/kg: the potential energy is -59.2MJ/kg, and the kinetic energy 29.6MJ/kg. Compare with the potential energy at the surface, which is -62.6MJ/kg. The extra potential energy is 3.4MJ/kg, the total extra energy is 33.0MJ/kg. The average speed is 7.7km/s, the net delta-v to reach this orbit is 8.1km/s (the actual delta-v is typically 1.5-2.0km/s more for atmospheric drag and gravity drag).
The increase per meter would be 4.4J/kg; this rate corresponds to one half of the local gravity of 8.8m/s2.
For an altitude of 100km (radius is 6471km):
The energy is -30.8MJ/kg: the potential energy is -61.6MJ/kg, and the kinetic energy 30.8MJ/kg. Compare with the potential energy at the surface, which is -62.6MJ/kg. The extra potential energy is 1.0MJ/kg, the total extra energy is 31.8MJ/kg.
The increase per meter would be 4.8J/kg; this rate corresponds to one half of the local gravity of 9.5m/s2. The speed is 7.8km/s, the net delta-v to reach this orbit is 8.0km/s.
Taking into account the rotation of the Earth, the delta-v is up to 0.46km/s less (starting at the equator and going east) or more (if going west).
For Voyager 1, with respect to the Sun:
Thus the hyperbolic excess velocity (the theoretical orbital velocity at infinity) is given by
However, Voyager 1 does not have enough velocity to leave the Milky Way. The computed speed applies far away from the Sun, but at such a position that the potential energy with respect to the Milky Way as a whole has changed negligibly, and only if there is no strong interaction with celestial bodies other than the Sun.
Then the time-rate of change of the specific energy of the rocket is : an amount for the kinetic energy and an amount for the potential energy.
The change of the specific energy of the rocket per unit change of delta-v is
which is |v| times the cosine of the angle between v and a.
Thus, when applying delta-v to increase specific orbital energy, this is done most efficiently if a is applied in the direction of v, and when |v| is large. If the angle between v and g is obtuse, for example in a launch and in a transfer to a higher orbit, this means applying the delta-v as early as possible and at full capacity. See also gravity drag. When passing by a celestial body it means applying thrust when nearest to the body. When gradually making an elliptic orbit larger, it means applying thrust each time when near the periapsis.
When applying delta-v to decrease specific orbital energy, this is done most efficiently if a is applied in the direction opposite to that of v, and again when |v| is large. If the angle between v and g is acute, for example in a landing (on a celestial body without atmosphere) and in a transfer to a circular orbit around a celestial body when arriving from outside, this means applying the delta-v as late as possible. When passing by a planet it means applying thrust when nearest to the planet. When gradually making an elliptic orbit smaller, it means applying thrust each time when near the periapsis.
If a is in the direction of v:
the Earth's surface
|Speed||Orbital period||Specific orbital energy|
|Earth's own rotation at surface (for comparison-- not an orbit)||6,378km||0km||465.1m/s (1,674km/h or 1,040mph)||23h 56min 4.09sec||-62.6MJ/kg|
|Orbiting at Earth's surface (equator) theoretical||6,378km||0km||7.9km/s (28,440km/h or 17,672mph)||1h 24min 18sec||-31.2MJ/kg|
|Low Earth orbit||6,600-8,400km||200-2,000km||
||1h 29min - 2h 8min||-29.8MJ/kg|
|Molniya orbit||6,900-46,300km||500-39,900km||1.5-10.0km/s (5,400-36,000km/h or 3,335-22,370mph) respectively||11h 58min||-4.7MJ/kg|
|Geostationary||42,000km||35,786km||3.1km/s (11,600km/h or 6,935mph)||23h 56min 4.09sec||-4.6MJ/kg|
|Orbit of the Moon||363,000-406,000km||357,000-399,000km||0.97-1.08km/s (3,492-3,888km/h or 2,170-2,416mph) respectively||27.27days||-0.5MJ/kg|