engineering 176 orbital design mr. ken ramsley [email protected] (508) 881- 5361...
TRANSCRIPT
When Orbits Were Perfect (and politically dangerous)
Einstein’s Geodesics (the art and science of motion)
Kepler’s Three Laws (based on Tycho’s meticulous data)
Orbital Elements Defined and Illustrated Useful Orbits and Maneuvers to Get
There Interplanetary Space and Beyond
Class Topics
EN176 Orbital Design
The Ancients
Aristotle (384 BC – 322 BC) Claudius Ptolemaeus (AD 83 – c.168)
Copernicus and Tycho
Nicolaus Copernicus (1473 - 1543) Tycho Brahe (1546 - 1601)
The Copernicus Solar System
Tycho Brahe's Uraniborg Observatory and 90° Star Sighting Quadrant
Image: Courtesy of tychobrahe.com
Kepler and Galileo
Johannes Kepler (1571 - 1630) Galileo Galilei (1564 - 1642)
Newton and LaGrange
Isaac Newton (1643 - 1727) Joseph Louis Lagrange (1736-1813)
Einstein
Geodesics: The Science and Art of 4D Curved Space
Trajectories.
All objects in motion conserve
momentum through a balance of
Gravity Potential and Velocity Vector(think rollercoaster)
Defining Simple 2-Body Orbits
This is all we need to know…
• Shape – More like a circle, or stretched out?• Size – Mostly nearby, or farther into space?• Orbital Plane Orientation – Pitch, Yaw, and Roll • Satellite Location – Where are we in this orbit?
Kepler’s First Law
Every orbit is an ellipse with the Sun (main body) located at one foci.
Kepler’s Second Law
A line between an orbiting body and primary body sweeps out equal areas in equal intervals of time.
Day 0
Day 10
Day 20
Day 30Day 40
Day 50
Day 60
Day 70
Day 80
Day 90
Day 100
Day 110Day 120
Kepler’s Third Law
P2 = R3
P1 P2
R2
R1
EXAMPLE:
Earth
P = 1 Year
R = 1 AU
Mars
P = 1.88 Years
R = 1.52 AU
This defines the relationship of Orbital Period & Average Radius for any two bodies in orbit.
For a given body, the orbital period and average distance for the second orbiting body is:
P = Orbital PeriodR = Average Radius
Vernal Equinox – The Celestial Baseline
First some astronomy…
When the Sun passes over the equator moving south to north.
Vernal Equinox (March 20th)
Defines a fixed vector in space through the center of the Earth to a known celestial coordinate point.
June 21st
December 22nd
Sun
The Vernal Equinox drifts ~0.014° / year. Orbits are therefore calculated for a
specified date and time, (most often Jan 1, 2000, 2050 or today).
Epoch 2000
Conic Sections (shape) Eccentricity
• e=0 -- circle• e<1 -- ellipse• e=1 -- parabola• e>1 -- hyperbola
e < 1 Orbit is ‘closed’ – recurring path (elliptical) e > 1 Not an orbit – passing trajectory (hyperbolic)
Keplerian Elements e, a, and v (3 of 6)
Perigee 0°
Apogee 180°
e defines ellipse shape
a defines ellipse size
v defines satellite angle from perigee
Semi-major axis (nm or km)
True anomaly (angle)
Eccentricity (0.0 to 1.0)
Apo/Peri gee – Earth Apo/Peri lune – Moon Apo/Peri helion – Sun Apo/Peri apsis – non-specific
90°120°
a
ev
150°
e=0.8 vrs e=0.0
Inclination i (4th Keplerian Element)
Inclination (angle)
Equatorial Plane ( defined by Earth’s equator )
Intersection of the equatorial and orbital planes
(below)
(above)
Sample inclinations 0° -- Geostationary 52° -- ISS 98° -- Mapping
Ascending Node
Ascending Node is where a satellite crosses the equatorial
plane moving south to north
i
Right Ascension [1] of the ascending node Ω and Argument of perigee ω (5th
and 6th Elements)
Vernal Equinox
Perigee Direction
Ωω
Ω = angle from vernal equinox to ascending node on the equatorial plane
ω = angle from ascending node to perigee on the orbital plane
[1] Right Ascension is the astronomical term for celestial (star) longitude.
Ascending Node
The Six Keplerian Elementsa = Semi-major axis (usually
in kilometers or nautical miles)
e = Eccentricity (of the elliptical orbit)
v = True anomaly The angle between perigee and satellite in the orbital plane at a specific time
i = Inclination The angle between the orbital and equatorial planes
Ω = Right Ascension (longitude) of the ascending node The angle from the Vernal Equinox vector to the ascending node on the equatorial plane
= Argument of perigee The angle measured between the ascending node and perigee
Shape, Size, Orientation, and Satellite
Location.
Sample Keplerian Elements (ISS)
TWO LINE MEAN ELEMENT SET - ISS
1 25544U 98067A 09061.52440963 .00010596 00000-0 82463-4 0 9009
2 25544 51.6398 133.2909 0009235 79.9705 280.2498 15.71202711 29176
Satellite: ISS
Catalog Number: 25544
Epoch time: 09061.52440963 = yrday.fracday
Element set: 900
Inclination: 51.6398 deg
RA of ascending node: 133.2909 deg
Eccentricity: .0009235
Arg of perigee: 79.9705 deg
Mean anomaly: 280.2498 deg
Mean motion: 15.71202711 rev/day (semi-major axis derivable from this)
Decay rate: 1.05960E-04 rev/day^2
Epoch rev: 2917
Checksum: 315
State Vectors NonKeplerian Coordinate
System Cartesian x, y, z, and 3D velocity
Orbit determination
On Board GPS
Ground Based Radar: Distance or “Range” (kilometers).Elevation or “Altitude” (Horizon = 0°, Zenith = 90°).Azimuth (Clockwise in degrees with due north = 0°).
On board Radio Transponder Ranging:Alt-Az plus radio signal turnaround delay (like radar).
Ground Sightings: Alt-Az only (best fit from many observations).
Launch From Vertical Takeoff
• Raising your altitude from 0 to 300 km (‘standing’ jump)
– Energy = mgh = 1 kg x 9.8 m/s2 x 300,000 m ∆V = 1715 m/s
• 7 km/s lateral velocity at 300 km altitude (orbital insertion)
– ∆V (velocity)= 7000 m/s – ∆V (altitude) = 1715 m/s– ∆V (total) = 8715 m/s [1]
[1] plus another 1500 m/s lost to drag during early portion of flight.
Launch From Airplane at 200 m/s and 10 km altitude
Raise altitude from 10 to 300 km (‘flying’ jump) Energy = mgh = 1 kg x 9.8 m/s2 x 290,000 m
∆V = 1686 m/s (98% of ground based launch ∆V) (96% of ground based launch energy)
Accelerate to 7000 m/s from 200 m/s ∆V (velocity) = 6800 m/s (97% of ground ∆V, 94% of energy) ∆V (∆Height) = 1686 m/s (98% of ground ∆V, 96% of energy)
∆V (total, with airplane) = 8486 m/s + 1.3 km/s drag loss = 9800 m/s ∆V (total, from ground) = 8715 m/s + 1.5 km/s drag loss = 10200 m/s
Total Velocity savings: 4%, Total Energy savings: 8%
Downsides: Human rating required for entire system, limited launch vehicle dimension and mass, fewer propellant choices, airplane expenses.
Ground Tracks
Ground tracks drift westward as the Earth rotates below an orbit.
Each orbit type has a signature ground tract.
More Astronomy Facts
The SunDrifts east in the sky ~1° per day. Rises 0.066 hours later each day.
(because the earth is orbiting)
The Earth…Rotates 360° in 23.934 hours
(Celestial or “Sidereal” Day)Rotates ~361° in 24.000 hours
(Noon to Noon or “Solar” Day)
Satellites orbits are aligned to the Sidereal day – not the solar day
Orbital Perturbations
“All orbits evolve” Atmospheric Drag (at LEO altitudes, only)
– Worse during increased solar activity. – Insignificant above ~800km.
Nodal Regression – The Earth is an oblate spheroid. This adds extra “pull” when a satellite passes over the equator – rotating the plane of the orbit to the east.
Other Factors – Gravitational irregularities – such as Earth-axis wobbles, Moon, Sun, Jupiter gravity (tends to flatten inclination). Solar photon pressure. Insignificant for LEO – primary perturbations elsewhere.
‘LEO’ < ~1,000km (Satellite Telephones,
ISS) ‘MEO’ = ~1,000km to 36,000km (GPS) ‘GEO’ = 36,000km (CommSats,
HDTV) ‘Deep Space’ > ~GEO
LEO is most common, shortest life. MEO difficult due to radiation belts. Most GEO orbit perturbation is latitude drift due to Sun and Moon.
Nodal Regression
Orbital planes rotate eastward over time.
(below)
(above)
Ascending Node
Nodal Regression can be very useful.
Sun-Synchronous OrbitsRelies on nodal regression to shift the ascending node ~1° per day.Scans the same path under the same lighting conditions each day.
The number of orbits per 24 hours must be an even integer (usually 15).Requires a slightly retrograde orbit (I = 97.56° for a 550km / 15-orbit SSO).
Each subsequent pass is 24° farther west (if 15 orbits per day).Repeats the pattern on the 16th orbit (or fewer for higher altitude SSOs).
Used for reconnaissance (or terrain mapping – with a bit of drift).
Molniya - 12hr Period‘Long loitering’ high latitude apogee. Once used used for early warning by both USA and USSR
‘Tundra’ Orbit - 24hr Period
Higher apogee than Molniya. For dwelling over a specific upper latitude (Used only by Sirius)
GPS Constellation ~ 20200km alt.
GPS: Six orbits with six equally-spaced satellites occupying each orbit.
Hohmann Transfer Orbit
Hohmann transfer orbit intersects both orbits.
Requires co-planar initial and ending orbits.
After 180°, second burn establishes the new orbit.
Can be used to reduce or increase orbit altitudes.
By far the most common orbital maneuver.
Orbital Plane ChangesBurn must take place where the initial and target planes intersect.
Even a small amount of plane change requires lots of ΔV
Less ΔV required at higher altitudes (e.g., slower orbital velocities).
Often combined with Hohmann transfer or rendezvous maneuver.
Simple Plane Change Formula (No Hohmann component):
Plane Change ΔV = 2 x Vorbit x sin(θ/2)
Example: Orbit Velocity = 7000m/s, Target Inclination Change = 30°
Plane Change ΔV = 2 x 7000m/s x sin(30°/ 2)
Plane Change ΔV = 3623m/s
θ
Fast Transfer Orbit
Requires less time due to higher energy transfer orbit.
Also faster since transfer is complete in less 180°.
Can be used to reduce or increase orbit altitudes.
Less common than Hohmann
Typically an upper stage restart where excess fuel is often available.
Geostationary Transfer Orbit ‘GTO’
Requires plane change and circularizing burns.
Less plane changing is required when launched from near the equator.
2. Plane change where GTO plane intersects GEO plane
3. Hohmann circularizing burn
1. launch to ‘GTO’
‘Super GTO’
Initial orbit has greater apogee than standard GTO.
Plane change at much higher altitude requires far less ΔV.
PRO: Less overall ΔV from higher inclination launch sites.
CON: Takes longer to establish the final orbit.
2. Plane change plus initial Hohmann burn
GEO Target Orbit
1. Launch to ‘Super GTO’
3. Second Hohmann burn circularizes at GEO
Low Thrust Orbit Transfer
PROs: Lower mass propulsion system. Same system used for orbital maintenance. CONs: Weeks or even months to reach final orbit. Van Allen Radiation belts.
A series of plane and altitude changes. Continuous electric engine propulsion.
RendezvousLaunch when the orbital plane of the target vehicle crosses launch pad.
(Ideally) launch as the target vehicle passes straight overhead.
Smaller transfer orbits slowly overtake target (because of shorter orbit periods).
Course maneuvers designed to arrive in the same orbit at the same true anomaly.
Apollo LM and CSM
Rendezvous
Orbital Debris a.k.a., ‘Space Junk’
Currently > 19,000 items 10cm or larger. ~ 700 (4%) functioning S/C. In as few as 50 years, upper LEO and lower MEO may be unusable.
February 2009 Iriduim / Cosmos collision created > 1,000 items > 10cm diameter
Deep Space
Cassini – Saturn orbit insertion using good ‘ol fashion rocket power.
Using Lagrange Points to ‘stay put’
Halo Orbits (stability from motion)
AeroBrakingEarth, Mars, Jupiter, etc.“The poor man’s Hohmann maneuver”
The Solar System ‘Super Highway’ …designing geodesic trajectories – like tossing a message bottle into the sea at exactly the right time, direction, and
velocity.
Gravity Assist (Removing Velocity)
Gravity Assist (adding velocity)
Solar Escape
Multiple Mission
Trajectories
Complex Orbital Trajectories
Galileo (Jupiter)
Cassini (Saturn)
Designing Deep Space Missions
…yes, there are software tools for this
Engineering 176 Orbits
Assignments for April 2
Create a trade table to compare orbit designs.
Trade criteria should include: Orbit suitability for mission. Cost to get there – and stay there. Space environment (e.g., radiation).
HOMEWORK:Design minimum two, preferably three orbits your mission could use.
For the selected orbits: Describe it (orbital elements) How will you get there? How will you stay there? Estimate perturbations
Reading on Orbits:SMAD ch 6 – scan 5 and 7TLOM ch 3 and 4 – scan 5 and 17