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| {{More footnotes|date=February 2010}}
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| [[File:Orbital motion.gif|thumb|200px|A satellite orbiting the earth has a [[Speed#Tangential speed|tangential velocity]] and an inward [[acceleration]].]]
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| {{Astrodynamics}}
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| '''Orbital mechanics''' or '''astrodynamics''' is the application of [[ballistics]] and [[celestial mechanics]] to the practical problems concerning the motion of [[rocket]]s and other [[spacecraft]]. The motion of these objects is usually calculated from [[Newton's laws of motion]] and [[Newton's law of universal gravitation]]. It is a core discipline within space mission design and control. Celestial mechanics treats more broadly the orbital dynamics of systems under the influence of [[gravity]], including both spacecraft and natural astronomical bodies such as star systems, [[planet]]s, [[moon]]s, and [[comet]]s. Orbital mechanics focuses on spacecraft [[trajectory|trajectories]], including [[orbital maneuver]]s, orbit plane changes, and interplanetary transfers, and is used by mission planners to predict the results of [[Spacecraft propulsion|propulsive maneuvers]]. [[General relativity]] is a more exact theory than Newton's laws for calculating orbits, and is sometimes necessary for greater accuracy or in high-gravity situations (such as orbits close to the Sun).
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| ==History==
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| Until the rise of [[Spaceflight|space travel]] in the twentieth century, there was little distinction between orbital and celestial mechanics. The fundamental techniques, such as those used to solve the [[Keplerian problem]] (determining position as a function of time), are therefore the same in both fields. Furthermore, the history of the fields is almost entirely shared.
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| [[Johannes Kepler]] was the first to successfully model planetary orbits to a high degree of accuracy, publishing [[Kepler's laws of planetary motion|his laws]] in 1605. [[Isaac Newton]] published more general laws of celestial motion in his 1687 book, ''[[Philosophiæ Naturalis Principia Mathematica]]''.
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| {{Expand section|date=August 2008}}
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| ==Practical techniques==
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| {{Further2|[[List of orbits]]}}
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| ===Rules of thumb===
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| The following rules of thumb are useful for situations approximated by [[classical mechanics]] under the standard assumptions of astrodynamics. The specific example discussed is of a satellite orbiting a planet, but the rules of thumb could also apply to other situations, such as orbits of small bodies around a star such as the Sun.
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| * [[Kepler's laws of planetary motion]], which can be mathematically derived from Newton's laws, hold strictly only in describing the motion of two gravitating bodies, in the absence of non-gravitational forces, or approximately when the gravity of a single massive body like the Sun dominates other effects:
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| ** Orbits are [[ellipse|elliptical]], with the heavier body at one [[Focus (geometry)|focus]] of the ellipse. Special cases of this are circular orbits (a circle being simply an ellipse of zero [[eccentricity (mathematics)|eccentricity]]) with the planet at the center, and [[parabola|parabolic]] orbits (which are ellipses with eccentricity of exactly 1, which is simply an infinitely long ellipse) with the planet at the focus.
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| ** A line drawn from the planet to the satellite sweeps out ''equal areas in equal times'' no matter which portion of the orbit is measured.
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| ** The square of a satellite's orbital period is proportional to the cube of its average distance from the planet.
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| * Without applying [[thrust]] (such as firing a rocket engine), the height and shape of the satellite's orbit won't change, and it will maintain the same orientation with respect to the fixed stars.
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| * A satellite in a low orbit (or low part of an elliptical orbit) moves more quickly with respect to the surface of the planet than a satellite in a higher orbit (or a high part of an elliptical orbit), due to the stronger gravitational attraction closer to the planet.
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| * If thrust is applied at only one point in the satellite's orbit, it will return to that same point on each subsequent orbit, though the rest of its path will change. Thus to move from one circular orbit to another, at least two brief applications of thrust are needed.
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| * From a circular orbit, thrust in a direction which slows the satellite down will create an elliptical orbit with a lower [[periapse]] (lowest orbital point) at 180 degrees away from the firing point. If thrust is applied to speed the satellite, it will create an elliptical orbit with a higher [[apoapse]] 180 degrees away from the firing point.
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| The consequences of the rules of orbital mechanics are sometimes counter-intuitive. For example, if two spacecraft are in the same circular orbit and wish to [[Space rendezvous|dock]], unless they are very close, the trailing craft cannot simply fire its engines to go faster. This will change the shape of its orbit, causing it to gain altitude and miss its target. One approach is to thrust [[retrograde]], or opposite to the direction of motion, and then thrust again to re-circularize the orbit at a lower altitude. Because lower orbits are faster than higher orbits, the trailing craft will begin to catch up. A third firing at the right time will put the trailing craft in an elliptical orbit which will intersect the path of the leading craft, approaching from below.
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| To the degree that the standard assumptions of astrodynamics do not hold, actual trajectories will vary from those calculated. For example, simple [[atmospheric drag]] is another complicating factor for objects in [[Earth orbit]]. These rules of thumb are decidedly inaccurate when describing two or more bodies of similar mass, such as a [[binary star system]] (see [[n-body problem]]). ([[Celestial mechanics]] uses more general rules applicable to a wider variety of situations.) The differences between [[classical mechanics]] and [[general relativity]] can also become important for large objects like planets.
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| ==Laws of astrodynamics==
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| {{see also|Laplace–Runge–Lenz vector}}
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| The fundamental laws of astrodynamics are [[Newton's law of universal gravitation]] and [[Newton's laws of motion]], while the fundamental mathematical tool is his differential [[calculus]].
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| Every orbit and trajectory outside atmospheres is in principle reversible, i.e., in the space-time function the time is reversed. The velocities are reversed and the accelerations are the same, including those due to rocket bursts. Thus if a rocket burst is in the direction of the velocity, in the reversed case it is opposite to the velocity. Of course in the case of rocket bursts there is no full reversal of events, both ways the same delta-v is used and the same [[mass ratio]] applies.
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| Standard assumptions in astrodynamics include non-interference from outside bodies, negligible mass for one of the bodies, and negligible other forces (such as from the solar wind, atmospheric drag, etc.). More accurate calculations can be made without these simplifying assumptions, but they are more complicated. The increased accuracy often does not make enough of a difference in the calculation to be worthwhile.
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| [[Kepler's laws of planetary motion]] may be derived from Newton's laws, when it is assumed that the orbiting body is subject only to the gravitational force of the central attractor. When an engine thrust or propulsive force is present, Newton's laws still apply, but Kepler's laws are invalidated. When the thrust stops, the resulting orbit will be different but will once again be described by Kepler's laws. The three laws are:
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| # The [[orbit]] of every [[planet]] is an [[ellipse]] with the sun at one of the [[Focus (geometry)|foci]].
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| # A [[line (mathematics)|line]] joining a planet and the sun sweeps out equal areas during equal intervals of time.
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| # The [[square (algebra)|square]]s of the [[orbital period]]s of planets are directly [[Proportionality (mathematics)|proportional]] to the [[cube (arithmetic)|cube]]s of the [[semi-major axis]] of the orbits.
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| ===Escape velocity===
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| The formula for [[escape velocity]] is easily derived as follows. The [[specific energy]] (energy per unit [[mass]]) of any space vehicle is composed of two components, the specific [[potential energy]] and the specific [[kinetic energy]]. The specific potential energy associated with a planet of [[mass]] ''M'' is given by
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| <center><math>- \frac{G M}{r} \,</math></center>
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| while the [[specific kinetic energy]] of an object is given by
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| <center><math>\frac{v^2}{2} \,</math></center>
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| Since [[conservation of energy|energy is conserved]], the total [[specific orbital energy]]
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| <center><math>\frac{v^2}{2} - \frac{G M}{r} \,</math></center>
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| does not depend on the distance, <math>r</math>, from the center of the central body to the space vehicle in question. Therefore, the object can reach infinite <math>r</math> only if this quantity is nonnegative, which implies
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| <center><math>v\geq\sqrt{\frac{2 G M}{r}}</math></center>
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| The escape velocity from the Earth's surface is about 11 km/s, but that is insufficient to send the body an infinite distance because of the gravitational pull of the Sun. To escape the Solar System from a location at a distance from the Sun equal to the distance Sun–Earth, but not close to the Earth, requires around 42 km/s velocity, but there will be "part credit" for the Earth's orbital velocity for spacecraft launched from Earth, if their further acceleration (due to the propulsion system) carries them in the same direction as Earth travels in its orbit.
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| ===Formulae for free orbits===
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| Orbits are [[conic section]]s, so, naturally, the formulas for the distance of a body for a given angle corresponds to the formula for that curve in [[polar coordinates]], which is:
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| : <math>r = \frac{ p }{1 + e \cos \theta}</math>
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| : <math>\mu= G(m_1+m_2)\,</math>
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| : <math>p=h^2/\mu\,</math>
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| where μ is called the [[gravitational parameter]], G is the [[gravitational constant]], m<sub>1</sub> and m<sub>2</sub> are the masses of objects 1 and 2, and h is the [[specific angular momentum]] of object 2 with respect to object 1. The parameter ''θ'' is known as the [[true anomaly]], ''p'' is the [[conic section|semi-latus rectum]], while ''e'' is the [[orbital eccentricity]], all obtainable from the various forms of the six independent [[orbital elements]].
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| ===Circular orbits===
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| All bounded orbits where the gravity of a central body dominates are elliptical in nature. A special case of this is the circular orbit, which is an ellipse of zero eccentricity. The formula for the velocity of a body in a circular orbit at distance ''r'' from the center of gravity of mass ''M'' is
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| :<math>\ v = \sqrt{\frac{GM} {r}\ }</math>
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| where <math>G</math> is the [[gravitational constant]], equal to
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| : 6.673 84 × 10<sup>−11</sup> m<sup>3</sup>/(kg·s<sup>2</sup>)
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| To properly use this formula, the units must be consistent; for example, ''M'' must be in kilograms, and ''r'' must be in meters. The answer will be in meters per second.
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| The quantity ''GM'' is often termed the [[standard gravitational parameter]], which has a different value for every planet or moon in the [[Solar System]].
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| Once the circular orbital velocity is known, the [[escape velocity]] is easily found by multiplying by the square root of 2:
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| :<math>\ v = \sqrt 2\sqrt{\frac {GM} {r}\ } = \sqrt{\frac {2GM} {r}\ }.</math>
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| ===Elliptical orbits===
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| If 0<e<1, then the denominator of the equation of free orbits varies with the true anomaly θ, but remains positive, never becoming zero. Therefore, the relative position vector remains bounded, having its smallest magnitude at periapsis r<sub>p</sub> which is given by:
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| :<math>r_p=\frac{p}{1+e}</math>
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| The maximum value r is reached when θ = 180. This point is called the apoapsis, and its radial coordinate, denoted r<sub>a</sub>, is
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| :<math>r_a=\frac{p}{1-e}</math>
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| Let 2a be the distance measured along the apse line from periapsis P to apoapsis A, as illustrated in the equation below:
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| :<math>2a=r_p+r_a</math>
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| Substituting the equations above, we get:
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| :<math>a=\frac{p}{1-e^2}</math>
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| a is the semimajor axis of the ellipse. Solving for r we get:
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| :<math>r=\frac{a(1-e^2)}{1+e\cos\theta}</math>
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| ==== Orbital period ====
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| Under standard assumptions the [[orbital period]] (<math>T\,\!</math>) of a body traveling along an elliptic orbit can be computed as:
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| :<math>T=2\pi\sqrt{a^3\over{\mu}}</math>
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| where:
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| *<math>\mu\,</math> is [[standard gravitational parameter]],
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| *<math>a\,\!</math> is length of [[semi-major axis]].
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| Conclusions:
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| *The orbital period is equal to that for a [[circular orbit]] with the orbit radius equal to the [[semi-major axis]] (<math>a\,\!</math>),
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| *For a given semi-major axis the orbital period does not depend on the eccentricity (See also: [[Kepler's laws of planetary motion#Third law|Kepler's third law]]).
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| ====Velocity====
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| Under standard assumptions the [[orbital speed]] (<math>v\,</math>) of a body traveling along an '''elliptic orbit''' can be computed from the [[Vis-viva equation]] as:
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| :<math>v=\sqrt{\mu\left({2\over{r}}-{1\over{a}}\right)}</math>
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| where:
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| *<math>\mu\,</math> is the [[standard gravitational parameter]],
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| *<math>r\,</math> is the distance between the orbiting bodies.
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| *<math>a\,\!</math> is the length of the [[semi-major axis]].
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| The velocity equation for a [[hyperbolic trajectory]] has either + <math>{1\over{a}}</math>, or it is the same with the convention that in that case ''a'' is negative.
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| ====Energy====
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| Under standard assumptions, [[specific orbital energy]] (<math>\epsilon\,</math>) of elliptic orbit is negative and the orbital energy conservation equation (the [[Vis-viva equation]]) for this orbit can take the form:
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| :<math>{v^2\over{2}}-{\mu\over{r}}=-{\mu\over{2a}}=\epsilon<0</math>
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| where:
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| *<math>v\,</math> is the speed of the orbiting body,
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| *<math>r\,</math> is the distance of the orbiting body from the center of mass of the [[central body]],
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| *<math>a\,</math> is the [[semi-major axis]],
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| *<math>\mu\,</math> is the [[standard gravitational parameter]].
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| Conclusions:
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| *For a given semi-major axis the specific orbital energy is independent of the eccentricity.
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| Using the [[virial theorem]] we find:
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| *the time-average of the specific potential energy is equal to 2ε
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| **the time-average of ''r''<sup>−1</sup> is ''a''<sup>−1</sup>
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| *the time-average of the specific kinetic energy is equal to -ε
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| ===Parabolic orbits===
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| If the eccentricity equals 1, then the orbit equation becomes:
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| :<math>r={{h^2}\over{\mu}}{{1}\over{1+\cos\theta}}</math>
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| where:
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| *<math>r\,</math> is the radial distance of the orbiting body from the mass center of the [[central body]],
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| *<math>h\,</math> is [[specific angular momentum]] of the [[orbiting body]],
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| *<math>\theta\,</math> is the [[true anomaly]] of the orbiting body,
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| *<math>\mu\,</math> is the [[standard gravitational parameter]].
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| As the true anomaly θ approaches 180°, the denominator approaches zero, so that ''r'' tends towards infinity. Hence, the energy of the trajectory for which ''e''=1 is zero, and is given by:
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| :<math>\epsilon={v^2\over2}-{\mu\over{r}}=0</math>
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| where:
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| *<math>v\,</math> is the speed of the orbiting body.
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| In other words, the speed anywhere on a parabolic path is:
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| :<math>v=\sqrt{2\mu\over{r}}</math>
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| ===Hyperbolic orbits===
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| If ''e''>1, the orbit formula,
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| :<math>r={{h^2}\over{\mu}}{{1}\over{1+e\cos\theta}}</math>
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| describes the geometry of the hyperbolic orbit. The system consists of two symmetric curves. the orbiting body occupies one of them. The other one is its empty mathematical image. Clearly, the denominator of the equation above goes to zero when cos''θ'' = -1/''e''. we denote this value of true anomaly
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| :<center>''θ''<sub>∞</sub> = cos<sup>-1</sup>(-1/''e'')<center>
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| since the radial distance approaches infinity as the true anomaly approaches ''θ''<sub>∞</sub>. ''θ''<sub>∞</sub> is known as the ''true anomaly of the asymptote''. Observe that ''θ''<sub>∞</sub> lies between 90° and 180°. From the trig identity sin<sup>2</sup>''θ''+cos<sup>2</sup>''θ''=1 it follows that:
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| :<center>sin''θ''<sub>∞</sub> = (e<sup>2</sup>-1)<sup>1/2</sup>/e
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| ====Energy====
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| Under standard assumptions, [[specific orbital energy]] (<math>\epsilon\,</math>) of a [[hyperbolic trajectory]] is greater than zero and the [[orbital energy conservation equation]] for this kind of trajectory takes form:
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| :<math>\epsilon={v^2\over2}-{\mu\over{r}}={\mu\over{-2a}}</math>
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| where:
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| *<math>v\,</math> is the [[orbital speed|orbital velocity]] of orbiting body,
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| *<math>r\,</math> is the radial distance of orbiting body from [[central body]],
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| *<math>a\,</math> is the negative [[semi-major axis]],
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| *<math>\mu\,</math> is [[standard gravitational parameter]].
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| ====Hyperbolic excess velocity====
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| {{see also|Characteristic energy}}
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| Under standard assumptions the body traveling along hyperbolic trajectory will attain in infinity an [[Kinetic energy|orbital velocity]] called hyperbolic excess velocity (<math>v_\infty\,\!</math>) that can be computed as:
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| :<math>v_\infty=\sqrt{\mu\over{-a}}\,\!</math>
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| where:
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| *<math>\mu\,\!</math> is [[standard gravitational parameter]],
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| *<math>a\,\!</math> is the negative [[semi-major axis]] of [[orbit]]'s [[hyperbola]].
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| The hyperbolic excess velocity is related to the [[specific orbital energy]] or characteristic energy by
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| :<math>2\epsilon=C_3=v_{\infty}^2\,\!</math>
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| ==Calculating trajectories==
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| ===Kepler's equation===
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| One approach to calculating orbits (mainly used historically) is to use [[Kepler's equation]]:
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| :<math> M = E - \epsilon \cdot \sin E </math>.
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| where ''M'' is the [[mean anomaly]], ''E'' is the [[eccentric anomaly]], and <math>\displaystyle \epsilon </math> is the [[Eccentricity (mathematics)|eccentricity]].
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| With Kepler's formula, finding the time-of-flight to reach an angle ([[true anomaly]]) of <math>\theta</math> from [[periapsis]] is broken into two steps:
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| # Compute the eccentric anomaly <math>E</math> from true anomaly <math>\theta</math>
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| # Compute the time-of-flight <math>t</math> from the eccentric anomaly <math>E</math>
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| Finding the eccentric anomaly at a given time ([[Kepler's Equation#Inverse problem|the inverse problem]]) is more difficult. Kepler's equation is [[transcendental function|transcendental]] in <math>E</math>, meaning it cannot be solved for <math>E</math> [[algebraic function|algebraically]]. Kepler's equation can be solved for <math>E</math> [[analytic function|analytically]] by inversion.
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| A solution of Kepler's equation, valid for all real values of <math> \textstyle \epsilon </math> is:
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| <math>
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| E =
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| \begin{cases}
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| \displaystyle \sum_{n=1}^{\infty}
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| {\frac{M^{\frac{n}{3}}}{n!}} \lim_{\theta \to 0} \left(
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| \frac{\mathrm{d}^{\,n-1}}{\mathrm{d}\theta^{\,n-1}} \left(
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| \frac{\theta}{ \sqrt[3]{\theta - \sin(\theta)} } ^n \right)
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| \right)
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| , & \epsilon = 1 \\
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| \displaystyle \sum_{n=1}^{\infty}
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| { \frac{ M^n }{ n! } }
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| \lim_{\theta \to 0} \left(
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| \frac{\mathrm{d}^{\,n-1}}{\mathrm{d}\theta^{\,n-1}} \left(
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| \frac{ \theta }{ \theta - \epsilon \cdot \sin(\theta)} ^n \right)
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| \right)
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| , & \epsilon \ne 1
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| \end{cases}
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| </math>
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| Evaluating this yields:
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| <math>
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| E =
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| \begin{cases} \displaystyle
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| x + \frac{1}{60} x^3 + \frac{1}{1400}x^5 + \frac{1}{25200}x^7 + \frac{43}{17248000}x^9 + \frac{ 1213}{7207200000 }x^{11} +
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| \frac{151439}{12713500800000 }x^{13} \cdots \ | \ x = ( 6 M )^\frac{1}{3}
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| , & \epsilon = 1 \\
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| \\
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| \displaystyle
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| \frac{1}{1-\epsilon} M
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| - \frac{\epsilon}{( 1-\epsilon)^4 } \frac{M^3}{3!}
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| + \frac{(9 \epsilon^2 + \epsilon)}{(1-\epsilon)^7 } \frac{M^5}{5!}
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| - \frac{(225 \epsilon^3 + 54 \epsilon^2 + \epsilon ) }{(1-\epsilon)^{10} } \frac{M^7}{7!}
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| + \frac{ (11025\epsilon^4 + 4131 \epsilon^3 + 243 \epsilon^2 + \epsilon ) }{(1-\epsilon)^{13} } \frac{M^9}{9!} \cdots
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| , & \epsilon \ne 1
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| \end{cases}
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| </math>
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| <br />
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| Alternatively, Kepler's Equation can be solved numerically. First one must guess a value of <math>E</math> and solve for time-of-flight; then adjust <math>E</math> as necessary to bring the computed time-of-flight closer to the desired value until the required precision is achieved. Usually, [[Newton's method]] is used to achieve relatively fast convergence.
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| The main difficulty with this approach is that it can take prohibitively long to converge for the extreme elliptical orbits. For near-parabolic orbits, eccentricity <math>\epsilon</math> is nearly 1, and plugging <math>e = 1</math> into the formula for mean anomaly, <math>E - \sin E</math>, we find ourselves subtracting two nearly-equal values, and accuracy suffers. For near-circular orbits, it is hard to find the periapsis in the first place (and truly circular orbits have no periapsis at all). Furthermore, the equation was derived on the assumption of an elliptical orbit, and so it does not hold for parabolic or hyperbolic orbits. These difficulties are what led to the development of the [[universal variable formulation]], described below.
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| ===Conic orbits===
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| For simple procedures, such as computing the [[delta-v]] for coplanar transfer ellipses, traditional approaches{{Clarify|date=February 2009}} are fairly effective. Others, such as time-of-flight are far more complicated, especially for near-circular and hyperbolic orbits.
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| ===The patched conic approximation===
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| {{Main|Patched conic approximation}}
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| The Hohmann transfer orbit alone is a poor approximation for interplanetary trajectories because it neglects the planets' own gravity. Planetary gravity dominates the behaviour of the spacecraft in the vicinity of a planet and in most cases Hohmann severely overestimates delta-v, and produces highly inaccurate prescriptions for burn timings.
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| A relatively simple way to get a [[Orders of approximation|first-order approximation]] of delta-v is based on the 'Patched Conic Approximation' technique. One must choose the one dominant gravitating body in each region of space through which the trajectory will pass, and to model only that body's effects in that region. For instance, on a trajectory from the Earth to Mars, one would begin by considering only the Earth's gravity until the trajectory reaches a distance where the Earth's gravity no longer dominates that of the Sun. The spacecraft would be given [[escape velocity]] to send it on its way to interplanetary space. Next, one would consider only the Sun's gravity until the trajectory reaches the neighbourhood of Mars. During this stage, the transfer orbit model is appropriate. Finally, only Mars's gravity is considered during the final portion of the trajectory where Mars's gravity dominates the spacecraft's behaviour. The spacecraft would approach Mars on a hyperbolic orbit, and a final retrograde burn would slow the spacecraft enough to be captured by Mars.
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| The size of the "neighborhoods" (or [[sphere of influence (astrodynamics)|spheres of influence]]) vary with radius <math>r_{SOI}</math>:
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| : <math>r_{SOI} = a_p\left(\frac{m_p}{m_s}\right)^{2/5}</math>
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| where <math>a_p</math> is the [[semimajor axis]] of the planet's orbit relative to the [[Sun]]; <math>m_p</math> and <math>m_s</math> are the [[mass]]es of the planet and Sun, respectively.
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| This simplification is sufficient to compute rough estimates of fuel requirements, and rough time-of-flight estimates, but it is not generally accurate enough to guide a spacecraft to its destination. For that, numerical methods are required.
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| <!--
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| ====Derivation of formulas====
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| For a spacecraft to travel from one planet to another, the patched conic approximation breaks the trip into three phases:
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| #Escape from the origin planet
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| #Interplanetary transfer
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| #Capture by the destination planet
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| We begin by considering phase 2. In this phase, the gravity of the planets is neglected, and only the Sun's gravity is considered. Therefore, this phase can be treated as a standard Hohmann transfer orbit.
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| The [[specific orbital energy]] of an orbit is <math>-\mu/2a</math>, where <math>a</math> is the [[semimajor axis]] and <math>\mu</math> is the [[standard gravitational parameter]] of the gravitating body. For phase 2, we are interested in the following orbits around the Sun: | |
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| {| border=1
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| |-
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| ! Orbit !! Major axis !! Total specific orbital energy
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| |-
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| | At origin || <math> 2r_0 \,</math> || <math>-\frac{\mu}{2r_0}</math>
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| |-
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| | Transfer || <math> r_0 + r_1 \,</math> || <math>-\frac{\mu}{r_0+r_1}</math>
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| |-
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| | At destination || <math> 2r_1 \,</math> || <math>-\frac{\mu}{2r_1}</math>
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| |}
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| *<math>-\mu/2r_0</math> for the origin planet
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| *<math>-\mu/2r_1</math> for the destination planet
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| *<math>-\mu/(r_0+r_1)</math> for the Hohmann transfer orbit
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| We must first get from the origin planet's orbit to the Hohmann transfer orbit, which requires a delta-v sufficient to make up the difference in orbital energy between the two orbits involves a change in specific kinetic energy of
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| A useful concept is the speed of a circular orbit speed at a given distance from a given body, which we will refer to as <math>C_b(r)</math>, where <math>b</math> is the body and <math>r</math> is the distance. (We may omit <math>b</math> and <math>r</math> for clarity.) This speed is usually easy to determine, simplifies the formulas, and tells us a lot about the gravitational conditions of a location in space. The [[specific orbital energy]] for a body in a circular orbit is <math>-\frac{1}{2C}</math>, and the [[specific kinetic energy]] is <math>\frac{1}{2C}</math>.
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| -->
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| | |
| ===The universal variable formulation===
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| To address computational shortcomings of traditional approaches for solving the 2-body problem, the [[universal variable formulation]] was developed. It works equally well for the circular, elliptical, parabolic, and hyperbolic cases, the differential equations converging well when integrated for any orbit. It also generalizes well to problems incorporating perturbation theory.
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| <!-- TODO: Describe and possibly derive the universal variable formulation. Talk about the f and g functions, and the X variable, and maybe outline an algorithm, maybe in wikicode. Try not to just lift this from Fundamentals of Astrophysics text. -->
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| ===Perturbations===
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| The universal variable formulation works well with the variation of parameters technique, except now, instead of the six Keplerian orbital elements, we use a different set of orbital elements: namely, the satellite's initial position and velocity vectors <math>x_0</math> and <math>v_0</math> at a given epoch <math>t = 0</math>. In a two-body simulation, these elements are sufficient to compute the satellite's position and velocity at any time in the future, using the universal variable formulation. Conversely, at any moment in the satellite's orbit, we can measure its position and velocity, and then use the universal variable approach to determine what its initial position and velocity ''would have been'' at the epoch. In perfect two-body motion, these orbital elements would be invariant (just like the Keplerian elements would be).
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| However, perturbations cause the orbital elements to change over time. Hence, we write the position element as <math>x_0(t)</math> and the velocity element as <math>v_0(t)</math>, indicating that they vary with time. The technique to compute the effect of perturbations becomes one of finding expressions, either exact or approximate, for the functions <math>x_0(t)</math> and <math>v_0(t)</math>.<!-- TODO: Explain it more -->
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| The following are some effects which make real orbits differ from the simple models based on a spherical earth. Most of them can be handled on short timescales (perhaps less than a few thousand orbits) by perturbation theory because they are small relative to the corresponding two-body effects.
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| * Equatorial bulges cause [[precession]] of the node and the perigee
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| * [[Tesseral harmonic]]s<ref>{{cite web|url= http://mathworld.wolfram.com/TesseralHarmonic.html|title= Tesseral Harmonic}}</ref> of the gravity field introduce additional perturbations
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| * Lunar and solar gravity perturbations alter the orbits
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| * Atmospheric drag reduces the semi-major axis unless make-up thrust is used
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| Over very long timescales (perhaps millions of orbits), even small perturbations can dominate, and the behaviour can become [[Chaos theory|chaotic]]. On the other hand, the various perturbations can be orchestrated by clever astrodynamicists to assist with orbit maintenance tasks, such as [[Orbital station-keeping|station-keeping]], [[ground track]] maintenance or adjustment, or phasing of perigee to cover selected targets at low altitude.
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| ==Orbital maneuver==
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| {{main|Orbital maneuver}}
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| In [[spaceflight]], an '''orbital maneuver''' is the use of [[spacecraft propulsion|propulsion]] systems to change the [[orbit]] of a [[spacecraft]]. For spacecraft far from Earth—for example those in orbits around the Sun—an orbital maneuver is called a ''deep-space maneuver (DSM)''.{{citation needed (lead)|date=September 2011}}
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| ===Orbital transfer===
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| [[File:Orbital Hohmann Transfer.svg|thumb|250px|A Hohmann transfer from a low circular orbit to a higher circular orbit.]]
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| [[File:Bi-elliptic transfer.svg|thumb|250px|A bi-elliptic transfer from a low circular starting orbit (dark blue), to a higher circular orbit (red).]]
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| [[File:Orbital Two-Impulse Transfer.svg|thumb|250px|A two-impulse transfer from a low circular orbit to a higher circular orbit.]]
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| [[File:Orbital General Transfer.svg|thumb|250px|A general transfer from a low circular orbit to a higher circular orbit.]]
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| [[Transfer orbit]]s are usually elliptical orbits that allow spacecraft to move from one (usually substantially circular) orbit to another. Usually they require a burn at the start, a burn at the end, and sometimes one or more burns in the middle.
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| *The [[Hohmann transfer orbit]] requires a minimal [[delta-v]].
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| *A [[Bi-elliptic transfer]] can require less energy than the Hohmann transfer, if the ratio of orbits is 11.94 or greater,<ref>{{Cite book | last = Vallado | first = David Anthony | title = Fundamentals of Astrodynamics and Applications | page = 317 | publisher = Springer | year = 2001 | isbn = 0-7923-6903-3 | url = http://books.google.com/books?id=PJLlWzMBKjkC&printsec}}</ref> but comes at the cost of increased trip time over the Hohmann transfer.
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| *Faster transfers may use any orbit that intersects both the original and destination orbits, at the cost of higher delta-v.
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| For the case of orbital transfer between non-coplanar orbits, the change-of-plane thrust must be made at the point where the orbital planes intersect (the "node").
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| ===Gravity assist and the Oberth effect===
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| In a [[gravity assist]], a spacecraft swings by a planet and leaves in a different direction, at a different speed. This is useful to speed or slow a spacecraft instead of carrying more fuel.
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| This maneuver can be approximated by an [[elastic collision]] at large distances, though the flyby does not involve any physical contact. Due to Newton's Third Law (equal and opposite reaction), any momentum gained by a spacecraft must be lost by the planet, or vice versa. However, because the planet is much, much more massive than the spacecraft, the effect on the planet's orbit is negligible.
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| The [[Oberth effect]] can be employed, particularly during a gravity assist operation. This effect is that use of a propulsion system works better at high speeds, and hence course changes are best done when close to a gravitating body; this can multiply the effective [[delta-v]].
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| | |
| === Interplanetary Transport Network and fuzzy orbits ===
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| {{Main|Interplanetary Transport Network}}
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| {{See also|Low energy transfers}}
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| It is now possible to use computers to search for routes using the nonlinearities in the gravity of the planets and moons of the Solar System. For example, it is possible to plot an orbit from high earth orbit to Mars, passing close to one of the Earth's [[Trojan point]]s. Collectively referred to as the [[Interplanetary Transport Network]], these highly perturbative, even chaotic, orbital trajectories in principle need no fuel beyond that needed to reach the Lagrange point (in practice keeping to the trajectory requires some course corrections). The biggest problem with them is they can be exceedingly slow, taking many years to arrive. In addition launch windows can be very far apart.
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| They have, however, been employed on projects such as [[Genesis (spacecraft)|Genesis]]. This spacecraft visited the Earth-Sun Lagrange {{L1}} point and returned using very little propellant.
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| == See also ==
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| {{Portal| Spaceflight }}
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| * [[Kepler orbit]]
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| * [[Spacecraft propulsion]]
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| * [[Tsiolkovsky rocket equation]]
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| * [[Aerodynamics]]
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| * [[Astrophysics]]
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| * [[Celestial mechanics]]
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| * [[Universal variable formulation]]
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| * [[Chaos theory]]
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| * [[Lagrangian point]]
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| * [[N-body problem]]
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| * [[Orbit]]
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| * [[Orders of magnitude (speed)]]
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| * [[Roche limit]]
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| * [[Canonical units]]
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| * [[Aerospace Engineering]]
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| * [[Mechanical Engineering]]
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| | |
| ==References==
| |
| | |
| {{Reflist}}
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| * {{cite book|author=Curtis, Howard D.,|title=Orbital Mechanics for Engineering Students, 2e|publisher=Elsevier|location=New York|year=2009|isbn=978-0-12-374778-5}}
| |
| | |
| * {{cite book
| |
| |last= Bate
| |
| |first= Roger R.
| |
| |coauthors= Mueller, Donald D., and White, Jerry E.
| |
| |title=Fundamentals of Astrodynamics
| |
| |year=1971
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| |publisher=Dover Publications
| |
| |location=New York
| |
| |isbn=0-486-60061-0
| |
| |pages=
| |
| }}
| |
| | |
| *{{cite book
| |
| |last=Sellers
| |
| |first=Jerry J.
| |
| |coauthors= Astore, William J., Giffen, Robert B., Larson, Wiley J.
| |
| |editor=Kirkpatrick, Douglas H.
| |
| |title= Understanding Space: An Introduction to Astronautics
| |
| |edition=2
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| |year=2004
| |
| |publisher=McGraw Hill
| |
| |isbn=0-07-242468-0
| |
| |pages= 228}}
| |
| | |
| * {{cite web | url = http://space.au.af.mil/primer/orbital_mechanics.pdf | title = Air University Space Primer, Chapter 8 - Orbital Mechanics | publisher = USAF}}
| |
| | |
| ==External links==
| |
| * [http://www.braeunig.us/space/orbmech.htm ORBITAL MECHANICS] (Rocket and Space Technology)
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| * [http://jat.sourceforge.net Java Astrodynamics Toolkit]
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| | |
| ==Further reading==
| |
| | |
| Many of the options, procedures, and supporting theory are covered in standard works such as:
| |
| * {{cite book|author= Bate, R.R., Mueller, D.D., White, J.E.,|title=Fundamentals of Astrodynamics|publisher=Dover Publications, New York|year=1971|isbn=978-0-486-60061-1}}
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| * {{cite book|author=Vallado, D. A.|title=Fundamentals of Astrodynamics and Applications, 2nd Edition|publisher=Springer|year=2001|isbn=978-0-7923-6903-5}}
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| * {{cite book|author=Battin, R.H.|title=An Introduction to the Mathematics and Methods of Astrodynamics|publisher = American Institute of Aeronautics & Ast, Washington, D.C.|year=1999|isbn=978-1-56347-342-5}}
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| * {{cite book|author=Chobotov, V.A. (ed.)|title=Orbital Mechanics, 3rd Edition|publisher=American Institute of Aeronautics & Ast, Washington, D.C.|year=2002|isbn=978-1-56347-537-5}}
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| * {{cite book|author=Herrick, S.|title=Astrodynamics: Orbit Determination, Space Navigation, Celestial Mechanics, Volume 1 |publisher=Van Nostrand Reinhold, London|year=1971|isbn=978-0-442-03370-5}}
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| * {{cite book|author=Herrick, S.|title=Astrodynamics: Orbit Correction, Perturbation Theory, Integration, Volume 2 |publisher=Van Nostrand Reinhold, London|year=1972|isbn=978-0-442-03371-2}}
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| * {{cite book|author=Kaplan, M.H.|title=Modern Spacecraft Dynamics and Controls|publisher=Wiley, New York|year=1976|isbn=978-0-471-45703-9}}
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| * {{cite book|author=Tom Logsdon|title=Orbital Mechanics|publisher=Wiley-Interscience, New York|year=1997|isbn=978-0-471-14636-0}}
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| * {{cite book|author=John E. Prussing and Bruce A. Conway|title=Orbital Mechanics|publisher=Oxford University Press, New York|year=1993|isbn=978-0-19-507834-3}}
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| * {{cite book|author=M.J. Sidi|title=Spacecraft Dynamics and Control|publisher=Cambridge University Press, New York|year=2000|isbn=978-0-521-78780-2}}
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| * {{cite book|author=W.E. Wiesel|title=Spaceflight Dynamics, 2nd edition|publisher=McGraw-Hill, New York|year=1996|isbn=978-0-07-070110-6}}
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| * {{cite book|author=J.P. Vinti|title=Orbital and Celestial Mechanics|publisher=American Institute of Aeronautics & Ast, Reston, Virginia|year=1998|isbn=978-1-56347-256-5}}
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| * {{cite book|author=P. Gurfil|title=Modern Astrodynamics|publisher=Butterworth-Heinemann|year=2006|isbn=978-0-12-373562-1}}
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| {{Astronomy subfields}}
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| {{Orbits}}
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| {{Spaceflight}}
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| {{DEFAULTSORT:Orbital Mechanics}}
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| [[Category:Astrodynamics|*]]
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