Presentation on theme: "Engineering 176 Orbital Design"— Presentation transcript:
1 Engineering 176 Orbital Design Mr. Ken Ramsley (508)
2 Class Topics 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 IllustratedUseful Orbits and Maneuvers to Get ThereInterplanetary Space and BeyondEN176 Orbital Design
3 Claudius Ptolemaeus (AD 83 – c.168) The AncientsAristotle (384 BC – 322 BC)Claudius Ptolemaeus (AD 83 – c.168)
4 Nicolaus Copernicus (1473 - 1543) Copernicus and TychoNicolaus Copernicus ( )Tycho Brahe ( )
5 The Copernicus Solar System Image: Courtesy of tychobrahe.comTycho Brahe's Uraniborg Observatory and 90° Star Sighting Quadrant
6 Kepler and Galileo Johannes Kepler (1571 - 1630) Galileo Galilei ( )
7 Joseph Louis Lagrange (1736-1813) Newton and LaGrangeIsaac Newton ( )Joseph Louis Lagrange ( )
9 Geodesics: The Science and Art of 4D Curved Space Trajectories. All objects in motion conserve momentum through a balance ofGravity Potential and Velocity Vector(think rollercoaster)Geodesics: The Science and Art of 4D Curved Space Trajectories.
10 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 RollSatellite Location – Where are we in this orbit?
11 Kepler’s First LawEvery orbit is an ellipse with the Sun (main body) located at one foci.
12 Kepler’s Second Law Day 40 Day 30 Day 50 Day 60 Day 20 Day 70 Day 80 A line between an orbiting body and primary body sweeps out equal areas in equal intervals of time.Day 120
13 Kepler’s Third LawThis 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:EXAMPLE:EarthP = 1 YearR = 1 AUMarsP = 1.88 YearsR = 1.52 AUR2R1P1P2P2 = R3P = Orbital PeriodR = Average Radius
14 Vernal Equinox – The Celestial Baseline First some astronomy…June 21stWhen 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.SunEpoch 2000The Vernal Equinox drifts ~0.014° / year. Orbits are therefore calculated for a specified date and time, (most often Jan 1, 2000, 2050 or today).December 22nd
15 Conic Sections (shape) Eccentricity e=0 -- circlee<1 -- ellipsee=1 -- parabolae>1 -- hyperbolae < 1 Orbit is ‘closed’ – recurring path (elliptical) e > 1 Not an orbit – passing trajectory (hyperbolic)
16 Keplerian Elements e, a, and v (3 of 6) 120°150°90°Eccentricity (0.0 to 1.0)vTrue anomaly (angle)aApogee 180°Perigee 0°Semi-major axis (nm or km)e=0.8 vrs e=0.0e defines ellipse shape a defines ellipse size v defines satellite angle from perigeeApo/Peri gee – Earth Apo/Peri lune – Moon Apo/Peri helion – Sun Apo/Peri apsis – non-specific
17 Inclination i (4th Keplerian Element) Intersection of the equatorial and orbital planesiInclination (angle)(above)(below)Ascending NodeEquatorial Plane ( defined by Earth’s equator )Sample inclinations 0° -- Geostationary ° -- ISS ° -- MappingAscending Node is where a satellite crosses the equatorial plane moving south to north
18 Right Ascension  of the ascending node Ω and Argument of perigee ω (5th and 6th Elements) Ω = angle from vernal equinox to ascending node on the equatorial planePerigee Directionω = angle from ascending node to perigee on the orbital planeωΩAscending Node Right Ascension is the astronomical term for celestial (star) longitude.Vernal Equinox
19 The Six Keplerian Elements a = 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 timei = Inclination The angle between the orbital and equatorial planesΩ = Right Ascension (longitude) of the ascending node The angle from the Vernal Equinox vector to the ascending node on the equatorial planew = Argument of perigee The angle measured between the ascending node and perigeeShape, Size, Orientation, and Satellite Location.
20 Sample Keplerian Elements (ISS) TWO LINE MEAN ELEMENT SET - ISSU 98067ASatellite: ISSCatalog Number: 25544Epoch time: = yrday.fracdayElement set: 900Inclination: degRA of ascending node: degEccentricity:Arg of perigee: degMean anomaly: degMean motion: rev/day (semi-major axis derivable from this)Decay rate: E-04 rev/day^2Epoch rev: 2917Checksum: 315
21 State Vectors NonKeplerian Coordinate System Cartesian x, y, z, and 3D velocity
22 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).
23 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 = m/s7 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  plus another 1500 m/s lost to drag during early portion of flight.
24 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) = m/s (97% of ground ∆V, 94% of energy) ∆V (∆Height) = m/s (98% of ground ∆V, 96% of energy)∆V (total, with airplane) = 8486 m/s km/s drag loss = m/s ∆V (total, from ground) = 8715 m/s km/s drag loss = m/sTotal 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.
25 Ground TracksGround tracks drift westward as the Earth rotates below an orbit.Each orbit type has a signature ground tract.
26 More Astronomy Facts The Sun The Earth… Drifts east in the sky ~1° per day Rises hours later each day.(because the earth is orbiting)The Earth…Rotates 360° in hours(Celestial or “Sidereal” Day)Rotates ~361° in hours(Noon to Noon or “Solar” Day)Satellites orbits are aligned to the Sidereal day – not the solar day
27 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.
28 ‘LEO’ < ~1,000km (Satellite Telephones, ISS) ‘MEO’ = ~1,000km to 36,000km (GPS) ‘GEO’ = 36,000km (CommSats, HDTV) ‘Deep Space’ > ~GEOLEO is most common, shortest life. MEO difficult due to radiation belts. Most GEO orbit perturbation is latitude drift due to Sun and Moon.
29 Nodal Regression can be very useful. Orbital planes rotate eastward over time.(above)Ascending Node(below)Nodal Regression can be very useful.
30 Sun-Synchronous Orbits Relies 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).
31 Molniya - 12hr Period‘Long loitering’ high latitude apogee. Once used used for early warning by both USA and USSR
32 ‘Tundra’ Orbit - 24hr Period Higher apogee than Molniya. For dwelling over a specific upper latitude (Used only by Sirius)
33 GPS Constellation ~ 20200km alt. GPS: Six orbits with six equally-spaced satellites occupying each orbit.
34 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.
35 Orbital Plane Changes θ Burn must take place where the initial and target planes intersect.Even a small amount of plane change requires lots of ΔVLess ΔV required at higher altitudes (e.g., slower orbital velocities).Often combined with Hohmann transfer or rendezvous maneuver.θSimple 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
36 Fast Transfer OrbitRequires 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 HohmannTypically an upper stage restart where excess fuel is often available.
37 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 plane1. launch to ‘GTO’3. Hohmann circularizing burn
38 ‘Super GTO’ 3. Second Hohmann burn circularizes at GEO GEO Target OrbitInitial 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.1. Launch to ‘Super GTO’2. Plane change plus initial Hohmann burn
39 Low Thrust Orbit Transfer A series of plane and altitude changes.Continuous electric engine propulsion.PROs: Lower mass propulsion system. Same system used for orbital maintenance CONs: Weeks or even months to reach final orbit. Van Allen Radiation belts.
40 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
41 Orbital Debris a.k.a., ‘Space Junk’ February 2009 Iriduim / Cosmos collision created > 1,000 items > 10cm diameterCurrently > 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.
42 Cassini – Saturn orbit insertion using good ‘ol fashion rocket power. Deep SpaceCassini – Saturn orbit insertion using good ‘ol fashion rocket power.
51 Complex Orbital Trajectories Galileo (Jupiter)Cassini (Saturn)
52 Designing Deep Space Missions …yes, there are software tools for this
53 Assignments for April 2 Reading on Orbits: HOMEWORK: SMAD ch 6 – scan 5 and 7TLOM ch 3 and 4 – scan 5 and 17HOMEWORK: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 perturbationsCreate 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).Engineering 176 Orbits