Hohmann transfer orbit, labelled 2, from an orbit (1) to a higher orbit (3)

An example of a Hohmann transfer orbit between Earth and Mars, as used by the NASA InSight probe:
   InSight ·   Earth ·   Mars

Part of a series on
Astrodynamics
Orbital mechanics
Orbital elements Apsis Argument of periapsis Eccentricity Inclination Mean anomaly Orbital nodes Semi-major axis True anomaly
Types of two-body orbits by eccentricity Circular orbit Elliptic orbit Transfer orbit (Hohmann transfer orbit Bi-elliptic transfer orbit) Parabolic orbit Hyperbolic orbit Radial orbit Decaying orbit
Equations Dynamical friction Escape velocity Kepler's equation Kepler's laws of planetary motion Orbital period Orbital velocity Surface gravity Specific orbital energy Vis-viva equation
Celestial mechanics
Gravitational influences Barycenter Hill sphere Perturbations Sphere of influence
N-body orbits Lagrangian points (Halo orbits) Lissajous orbits Lyapunov orbits
Engineering and efficiency
Preflight engineering Mass ratio Payload fraction Propellant mass fraction Tsiolkovsky rocket equation
Efficiency measures Gravity assist Oberth effect

In astronautics, the Hohmann transfer orbit (/ˈhoʊmən/) is an orbital maneuver used to transfer a spacecraft between two orbits of different altitudes around a central body. Examples would be used for travel between low Earth orbit and the Moon, or another solar planet or asteroid. In the idealized case, the initial and target orbits are both circular and coplanar. The maneuver is accomplished by placing the craft into an elliptical transfer orbit that is tangential to both the initial and target orbits. The maneuver uses two impulsive engine burns: the first establishes the transfer orbit, and the second adjusts the orbit to match the target.

The Hohmann maneuver often uses the lowest possible amount of impulse (which consumes a proportional amount of delta-v, and hence propellant) to accomplish the transfer, but requires a relatively longer travel time than higher-impulse transfers. In some cases where one orbit is much larger than the other, a bi-elliptic transfer can use even less impulse, at the cost of even greater travel time.

The maneuver was named after Walter Hohmann, the German scientist who published a description of it in his 1925 book Die Erreichbarkeit der Himmelskörper (The Attainability of Celestial Bodies).^[1] Hohmann was influenced in part by the German science fiction author Kurd Lasswitz and his 1897 book Two Planets.

When used for traveling between celestial bodies, a Hohmann transfer orbit requires that the starting and destination points be at particular locations in their orbits relative to each other. Space missions using a Hohmann transfer must wait for this required alignment to occur, which opens a launch window. For a mission between Earth and Mars, for example, these launch windows occur every 26 months. A Hohmann transfer orbit also determines a fixed time required to travel between the starting and destination points; for an Earth-Mars journey this travel time is about 9 months. When transfer is performed between orbits close to celestial bodies with significant gravitation, much less delta-v is usually required, as the Oberth effect may be employed for the burns.

They are also often used for these situations, but low-energy transfers which take into account the thrust limitations of real engines, and take advantage of the gravity wells of both planets can be more fuel efficient.^[2][3][4]

Example

The diagram shows a Hohmann transfer orbit to bring a spacecraft from a lower circular orbit into a higher one. It is an elliptic orbit that is tangential both to the lower circular orbit the spacecraft is to leave (cyan, labeled 1 on diagram) and the higher circular orbit that it is to reach (red, labeled 3 on diagram). The transfer orbit (yellow, labeled 2 on diagram) is initiated by firing the spacecraft's engine to add energy and raise the apogee. When the spacecraft reaches apogee, a second engine firing adds energy to raise the perigee, putting the spacecraft in the larger circular orbit.

Due to the reversibility of orbits, a similar Hohmann transfer orbit can be used to bring a spacecraft from a higher orbit into a lower one; in this case, the spacecraft's engine is fired in the opposite direction to its current path, slowing the spacecraft and lowering its perigee to that of the elliptical transfer orbit. The engine is then fired again at the lower distance to slow the spacecraft into the lower circular orbit. The Hohmann transfer orbit is based on two instantaneous velocity changes. Extra fuel is required to compensate for the fact that the bursts take time; this is minimized by using high-thrust engines to minimize the duration of the bursts. For transfers in Earth orbit, the two burns are labelled the perigee burn and the apogee burn (or apogee kick^[5]); more generally, they are labelled periapsis and apoapsis burns. Alternately, the second burn to circularize the orbit may be referred to as a circularization burn.

Calculation

For a small body orbiting another much larger body, such as a satellite orbiting Earth, the total energy of the smaller body is the sum of its kinetic energy and potential energy, and this total energy also equals half the potential at the average distance ${\displaystyle a}$ (the semi-major axis):

$${\displaystyle E={\frac {mv^{2}}{2}}-{\frac {GMm}{r}}={\frac {-GMm}{2a}}.}$$

Solving this equation for velocity results in the vis-viva equation,

$${\displaystyle v^{2}=\mu \left({\frac {2}{r}}-{\frac {1}{a}}\right),}$$

where:

• ${\displaystyle v}$ is the speed of an orbiting body,
• ${\displaystyle \mu =GM}$ is the standard gravitational parameter of the primary body, assuming ${\displaystyle M+m}$ is not significantly bigger than ${\displaystyle M}$ (which makes ${\displaystyle v_{M}\ll v}$), (for Earth, this is μ~3.986E14 m³ s⁻²)
• ${\displaystyle r}$ is the distance of the orbiting body from the primary focus,
• ${\displaystyle a}$ is the semi-major axis of the body's orbit.

Therefore, the delta-v (Δv) required for the Hohmann transfer can be computed as follows, under the assumption of instantaneous impulses:

$${\displaystyle \Delta v_{1}={\sqrt {\frac {\mu }{r_{1}}}}\left({\sqrt {\frac {2r_{2}}{r_{1}+r_{2}}}}-1\right),}$$

to enter the elliptical orbit at ${\displaystyle r=r_{1}}$ from the ${\displaystyle r_{1}}$ circular orbit, where ${\displaystyle r_{2}}$ is the aphelion of the resulting elliptical orbit, and

$${\displaystyle \Delta v_{2}={\sqrt {\frac {\mu }{r_{2}}}}\left(1-{\sqrt {\frac {2r_{1}}{r_{1}+r_{2}}}}\right),}$$

to leave the elliptical orbit at ${\displaystyle r=r_{2}}$ to the ${\displaystyle r_{2}}$ circular orbit, where ${\displaystyle r_{1}}$ and ${\displaystyle r_{2}}$ are respectively the radii of the departure and arrival circular orbits; the smaller (greater) of ${\displaystyle r_{1}}$ and ${\displaystyle r_{2}}$ corresponds to the periapsis distance (apoapsis distance) of the Hohmann elliptical transfer orbit. Typically, ${\displaystyle \mu }$ is given in units of m³/s², as such be sure to use meters, not kilometers, for ${\displaystyle r_{1}}$ and ${\displaystyle r_{2}}$. The total ${\displaystyle \Delta v}$ is then:

$${\displaystyle \Delta v_{\text{total}}=\Delta v_{1}+\Delta v_{2}.}$$

Whether moving into a higher or lower orbit, by Kepler's third law, the time taken to transfer between the orbits is

$${\displaystyle t_{\text{H}}={\frac {1}{2}}{\sqrt {\frac {4\pi ^{2}a_{\text{H}}^{3}}{\mu }}}=\pi {\sqrt {\frac {(r_{1}+r_{2})^{3}}{8\mu }}}}$$

(one half of the orbital period for the whole ellipse), where ${\displaystyle a_{\text{H}}}$ is length of semi-major axis of the Hohmann transfer orbit.

In application to traveling from one celestial body to another it is crucial to start maneuver at the time when the two bodies are properly aligned. Considering the target angular velocity being

$${\displaystyle \omega _{2}={\sqrt {\frac {\mu }{r_{2}^{3}}}},}$$

angular alignment α (in radians) at the time of start between the source object and the target object shall be

$${\displaystyle \alpha =\pi -\omega _{2}t_{\text{H}}=\pi \left(1-{\frac {1}{2{\sqrt {2}}}}{\sqrt {\left({\frac {r_{1}}{r_{2}}}+1\right)^{3}}}\right).}$$

Example

Total energy balance during a Hohmann transfer between two circular orbits with first radius ${\displaystyle r_{p}}$ and second radius ${\displaystyle r_{a}}$

Consider a geostationary transfer orbit, beginning at r₁ = 6,678 km (altitude 300 km) and ending in a geostationary orbit with r₂ = 42,164 km (altitude 35,786 km).

In the smaller circular orbit the speed is 7.73 km/s; in the larger one, 3.07 km/s. In the elliptical orbit in between the speed varies from 10.15 km/s at the perigee to 1.61 km/s at the apogee.

Therefore the Δv for the first burn is 10.15 − 7.73 = 2.42 km/s, for the second burn 3.07 − 1.61 = 1.46 km/s, and for both together 3.88 km/s.

This is greater than the Δv required for an escape orbit: 10.93 − 7.73 = 3.20 km/s. Applying a Δv at the Low Earth orbit (LEO) of only 0.78 km/s more (3.20−2.42) would give the rocket the escape velocity, which is less than the Δv of 1.46 km/s required to circularize the geosynchronous orbit. This illustrates the Oberth effect that at large speeds the same Δv provides more specific orbital energy, and energy increase is maximized if one spends the Δv as quickly as possible, rather than spending some, being decelerated by gravity, and then spending some more to overcome the deceleration (of course, the objective of a Hohmann transfer orbit is different).

Worst case, maximum delta-v

As the example above demonstrates, the Δv required to perform a Hohmann transfer between two circular orbits is not the greatest when the destination radius is infinite. (Escape speed is √2 times orbital speed, so the Δv required to escape is √2 − 1 (41.4%) of the orbital speed.) The Δv required is greatest (53.0% of smaller orbital speed) when the radius of the larger orbit is 15.5817... times that of the smaller orbit.^[10] This number is the positive root of x³ − 15x² − 9x − 1 = 0, which is ${\textstyle 5+4\,{\sqrt {7}}\cos \left({1 \over 3}\arctan {{\sqrt {3}} \over 37}\right)}$. For higher orbit ratios the Δv required for the second burn decreases faster than the first increases.

Application to interplanetary travel

When used to move a spacecraft from orbiting one planet to orbiting another, the situation becomes somewhat more complex, but much less delta-v is required, due to the Oberth effect, than the sum of the delta-v required to escape the first planet plus the delta-v required for a Hohmann transfer to the second planet.

For example, consider a spacecraft travelling from Earth to Mars. At the beginning of its journey, the spacecraft will already have a certain velocity and kinetic energy associated with its orbit around Earth. During the burn the rocket engine applies its delta-v, but the kinetic energy increases as a square law, until it is sufficient to escape the planet's gravitational potential, and then burns more so as to gain enough energy to get into the Hohmann transfer orbit (around the Sun). Because the rocket engine is able to make use of the initial kinetic energy of the propellant, far less delta-v is required over and above that needed to reach escape velocity, and the optimum situation is when the transfer burn is made at minimum altitude (low periapsis) above the planet. The delta-v needed is only 3.6 km/s, only about 0.4 km/s more than needed to escape Earth, even though this results in the spacecraft going 2.9 km/s faster than the Earth as it heads off for Mars (see table below).

At the other end, the spacecraft will need a certain velocity to orbit Mars, which will actually be less than the velocity needed to continue orbiting the Sun in the transfer orbit, let alone attempting to orbit the Sun in a Mars-like orbit. Therefore, the spacecraft will have to decelerate in order for the gravity of Mars to capture it. This capture burn should optimally be done at low altitude to also make best use of the Oberth effect. Therefore, relatively small amounts of thrust at either end of the trip are needed to arrange the transfer compared to the free space situation.

However, with any Hohmann transfer, the alignment of the two planets in their orbits is crucial – the destination planet and the spacecraft must arrive at the same point in their respective orbits around the Sun at the same time. This requirement for alignment gives rise to the concept of launch windows.

The term lunar transfer orbit (LTO) is used for the Moon.

It is possible to apply the formula given above to calculate the Δv in km/s needed to enter a Hohmann transfer orbit to arrive at various destinations from Earth (assuming circular orbits for the planets). In this table, the column labeled "Δv to enter Hohmann orbit from Earth's orbit" gives the change from Earth's velocity to the velocity needed to get on a Hohmann ellipse whose other end will be at the desired distance from the Sun. The column labeled "v exiting LEO" gives the velocity needed (in a non-rotating frame of reference centered on the earth) when 300 km above the Earth's surface. This is obtained by adding to the specific kinetic energy the square of the speed (7.73 km/s) of this low Earth orbit (that is, the depth of Earth's gravity well at this LEO). The column "Δv from LEO" is simply the previous speed minus 7.73 km/s.

Note that in most cases, Δv from LEO is less than the Δv to enter Hohmann orbit from Earth's orbit.

To get to the Sun, it is actually not necessary to use a Δv of 24 km/s. One can use 8.8 km/s to go very far away from the Sun, then use a negligible Δv to bring the angular momentum to zero, and then fall into the Sun. This can be considered a sequence of two Hohmann transfers, one up and one down. Also, the table does not give the values that would apply when using the Moon for a gravity assist. There are also possibilities of using one planet, like Venus which is the easiest to get to, to assist getting to other planets or the Sun.

Comparison to other transfers

See also

• Bi-elliptic transfer
• Delta-v budget
• Geostationary transfer orbit
• Halo orbit
• Lissajous orbit
• List of orbits
• Orbital mechanics

Citations

General and cited sources

• Bate, R. R.; Mueller, D.D.; White, J.E. (1971). Fundamentals of Astrodynamics. New York: Dover Publications. ISBN 978-0-486-60061-1.
• Battin, R .H. (1999). An Introduction to the Mathematics and Methods of Astrodynamics. Washington, DC: American Institute of Aeronautics & Ast. ISBN 978-1-56347-342-5.
• Hohmann, Walter (1925). Die Erreichbarkeit der Himmelskörper. Munich: R. Oldenbourg Verlag. ISBN 3-486-23106-5.
• Thornton, Stephen T.; Marion, Jerry B. (2003). Classical Dynamics of Particles and Systems (5th ed.). Brooks Cole. ISBN 0-534-40896-6.
• Vallado, D. A. (2001). Fundamentals of Astrodynamics and Applications (2nd ed.). Springer. ISBN 978-0-7923-6903-5.

Further reading

• "Orbital Mechanics". Rocket and Space Technology. Robert A. Braeunig. Archived from the original on 2012-02-04. Retrieved 2005-08-17.
• "4. Interplanetary Trajectories". Basics of Spaceflight. JPL: NASA.