GN/MAE155A 1 Orbital Mechanics Overview MAE 155A Dr. George Nacouzi
Jan 26, 2016
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Orbital Mechanics Overview
MAE 155A
Dr. George Nacouzi
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James Webb Space Telescope, Launch Date 2011
Primary mirror: 6.5-meter aperture Orbit: 930,000 miles from Earth , Mission lifetime: 5 years (10-year goal)Telescope Operating temperature: ~45 Kelvin Weight: Approximately 6600kg
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Overview: Orbital Mechanics
• Study of S/C (Spacecraft) motion influenced principally by gravity. Also considers perturbing forces, e.g., external pressures, on-board mass expulsions (e.g, thrust)
• Roots date back to 15th century (& earlier), e.g., Sir Isaac Newton, Copernicus, Galileo & Kepler
• In early 1600s, Kepler presented his 3 laws of planetary motion– Includes elliptical orbits of planets
– Also developed Kepler’s eqtn which relates position & time of orbiting bodies
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Overview: S/C Mission Design
• Involves the design of orbits/constellations for meeting Mission Objectives, e.g., area coverage
• Constellation design includes: number of S/C, number of orbital planes, inclination, phasing, as well as orbital parameters such as apogee, eccentricity and other key parameters
• Orbital mechanics provides the tools needed to develop the appropriate S/C constellations to meet the mission objectives
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Introduction: Orbital Mechanics• Motion of satellite is influenced by the gravity field of multiple
bodies, however, Two body assumption is usually sufficient. Earth orbiting satellite Two Body approach:
– Central body is earth, assume it has only gravitational influence on S/C, assume M >> m (M, m ~ mass of earth & S/C)
• Gravity effects of secondary bodies including sun, moon and other planets in solar system are ignored
– Solution assumes bodies are spherically symmetric, point sources (Earth oblateness can be important and is accounted for in J2 term of gravity field)
– Only gravity and centrifugal forces are present
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Two Body Motion (or Keplerian Motion)
• Closed form solution for 2 body exists, no explicit soltn exists for N >2, numerical approach needed
• Gravitational field on body is given by:Fg = M m G/R2 where,
M~ Mass of central body; m~ Mass of Satellite
G~ Universal gravity constant
R~ distance between centers of bodies
For a S/C in Low Earth Orbit (LEO), the gravity forces are:
Earth: 0.9 g Sun: 6E-4 g Moon: 3E-6 g Jupiter: 3E-8 g
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General Two Body Motion Equations
2t
rd
d
2 r
R 0 GMwhere
Solution is in form of conical section, i.e., circle, ellipse, parabola & hyperbola.2
trd
d
2 r
R 0
V
2
2
R
KE + PE, PE = 0 at R= ∞ ∞
V 2R
a
a~ semi major axis of ellipse
H = R x V = R V cos (), where H~ angular momentum & ~ flight path angle (between V & local horizontal)
& r ~Position vector
V
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General Two Body Motion Trajectories
Central Body
Circle, a=r
Ellipse, a> 0
Hyperbola, a< 0
Parabola, a =
a
• Parabolic orbits provide minimum escape velocity• Hyperbolic orbits used for interplanetary travel
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General Solution to Orbital Equation
• Velocity is given by:
• Eccentricity: e = c/a where, c = [Ra - Rp]/2
Ra~ Radius of Apoapsis, Rp~ Radius of Periapsis
• e is also obtained from the angular momentum H as:
e = [1 - (H2/a)]; and H = R V cos ()
V 2R
a
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Circular Orbits Equations
• Circular orbit solution offers insight into understanding of orbital mechanics and are easily derived
• Consider: Fg = M m G/R2 & Fc = m V2 /R (centrifugal F)
V is solved for to get:
V= (MG/R) = (/R)
• Period is then: T=2R/V => T = 2(R3/)
Fc
Fg
V
R
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Elliptical Orbit Geometry & Nomenclature
Periapsis
ApoapsisLine of Apsides
R
a c
V
Rpb
• Line of Apsides connects Apoapsis, central body & Periapsis• Apogee~ Apoapsis; Perigee~ Periapsis (earth nomenclature)
S/C position defined by R & , is called true anomalyR = [Rp (1+e)]/[1+ e cos()]
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Elliptical Orbit Definition
• Orbit is defined using the 6 classical orbital elements including:– Eccentricity, semi-
major axis, true anomaly and inclination, where
• Inclination, i, is the angle between orbit plane and equatorial plane
i
Other 2 parameters are: • Argument of Periapsis (). Ascending Node: Pt where S/C crosses equatorial plane South to North • Longitude of Ascending Node ()~Angle from Vernal Equinox (vector from center of earth to sun on first day of spring) and ascending node
Vernal Equinox
AscendingNode
Periapsis
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Some Orbit Types...
• Extensive number of orbit types, some common ones:– Low Earth Orbit (LEO), Ra < 2000 km
– Mid Earth Orbit (MEO), 2000< Ra < 30000 km
– Highly Elliptical Orbit (HEO)
– Geosynchronous (GEO) Orbit (circular): Period = time it takes earth to rotate once wrt stars, R = 42164 km
– Polar orbit => inclination = 90 degree
– Molniya ~ Highly eccentric orbit with 12 hr period (developed by Soviet Union to optimize coverage of Northern hemisphere)
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Sample Orbits
LEO at 0 & 45 degree inclination
Elliptical, e~0.46, I~65deg
Ground tracefrom i= 45 deg
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Sample GEO Orbit
Figure ‘8’ trace due to inclination, zero inclination in nomotion of nadir point(or satellite sub station)
• Nadir for GEO (equatorial, i=0) remain fixed over point• 3 GEO satellites provide almostcomplete global coverage
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Orbital Maneuvers Discussion
• Orbital Maneuver– S/C uses thrust to change orbital parameters, i.e., radius, e,
inclination or longitude of ascending node
– In-Plane Orbit Change• Adjust velocity to convert a conic orbit into a different conic orbit.
Orbit radius or eccentricity can be changed by adjusting velocity
• Hohmann transfer: Efficient approach to transfer between 2 Non-intersecting orbits. Consider a transfer between 2 circular orbits. Let Ri~ radius of initial orbit, Rf ~ radius of final orbit. Design transfer ellipse such that: Rp (periapsis of transfer orbit) = Ri (Initial R) Ra (apoapsis of transfer orbit) = Rf (Final R)
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Hohmann Transfer Description
DV1
DV2
TransferEllipse
Final Orbit
Initial Orbit
Rp = RiRa = RfDV1 = Vp - ViDV2 = Va - Vf
Note:( )p = transfer periapsis( )a = transfer apoapsis
RpRa
RiRf
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General In-Plane Orbital Transfers...
• Change initial orbit velocity Vi to an intersecting coplanar orbit with velocity Vf DV2 = Vi2 + Vf2 - 2 Vi Vf cos (a)
Initial orbit
Final orbit
a Vf
DV
Vi
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Other Orbital Transfers...
• Bielliptical Tranfer– When the transfer is from an initial orbit to a final orbit
that has a much larger radius, a bielliptical transfer may be more efficient
• Involves three impulses (vs. 2 in Hohmann)
• Plane Changes– Can involve a change in inclination, longitude of
ascending nodes or both
– Plane changes are very expensive (energy wise) and are therefore avoided if possible
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Basics of Rocket Equation
F = Ve dm/dt
F = M dV/dt
Thrust
dV/dtVe ~ Exhaust Vel.m ~ propellant massF = Thrust = Force
M ~S/C MassV ~ S/C Velocitygc~ gravitational constant
S/C
M dV/dt = Ve dm/dt = - Ve dM/dt
=> DV = Ve ln (Mi/Mf) where, Mi ~ Initial Mass; Mf~ Final Mass
Isp = Thrust/(gc dm/dt) => Ve = Isp x gc
Calculate mass of propellant needed for rocket to provide a velocitygain (DV)
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Basics of Rocket Equation (cont’d)M dV/dt = Ve dm/dt = - Ve dM/dt
=> DV = Ve ln (Mi/Mf) where, Mi ~ Initial Mass; Mf~ Final Mass
Isp = Thrust/(gc dm/dt) => Ve = Isp x gc
Substituting we get:Mi/Mf = exp (DV/ (gc Isp))but Mp = Mf - Mi =>Mp = Mi[1-exp(-DV/ gc Isp)]
Where, DV ~ Delta Velocity, Mp ~ Mass of Propellant
Mass of propellant calculated from Delta Velocity and propellant Isp.For Launch Vehicles: Isp ~ 260 - 300 sec for solid propellant Isp ~ 300 - 500 sec for liquid bipropellant
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