#### Transcript Orbital Mechanics and Design

Engineering 176 Orbital Design Mr. Ken Ramsley [email protected] (508) 881- 5361 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 Illustrated Useful Orbits and Maneuvers to Get There Interplanetary Space and Beyond 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 Image: Courtesy of tychobrahe.com Tycho Brahe's Uraniborg Observatory and 90° Star Sighting Quadrant 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 Day 40 Day 50 Day 30 Day 60 Day 20 Day 70 Day 80 Day 90 Day 10 Day 100 A line between an orbiting body and primary body sweeps out equal areas in equal intervals of time. Day 110 Day 120 Day 0 Kepler’s Third Law This defines the relationship of Orbital Period & Average Radius for any two bodies in orbit. EXAMPLE: R2 Earth R1 For a given body, the orbital period and average distance for the second orbiting body is: P1 P2 P = 1 Year R = 1 AU Mars P = 1.88 Years R = 1.52 AU P 2 = R3 P = Orbital Period R = Average Radius Vernal Equinox – The Celestial Baseline First some astronomy… June 21st When the Sun passes over the equator moving south to north. Sun Epoch 2000 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). December 22nd Vernal Equinox (March 20th) Defines a fixed vector in space through the center of the Earth to a known celestial coordinate point. Conic Sections (shape) Eccentricity • • • • e=0 e<1 e=1 e>1 -- circle -- ellipse -- parabola -- hyperbola e < 1 Orbit is ‘closed’ – recurring path (elliptical) e > 1 Not an orbit – passing trajectory (hyperbolic) Keplerian Elements e e, a, and v (3 of 6) 120° 150° 90° Eccentricity (0.0 to 1.0) Apogee 180° v True anomaly (angle) a Perigee 0° Semi-major axis (nm or km) e=0.8 vrs e=0.0 Apo/Peri gee – Earth Apo/Peri lune – Moon Apo/Peri helion – Sun Apo/Peri apsis – non-specific e a v defines ellipse shape defines ellipse size defines satellite angle from perigee Inclination i Intersection of the equatorial and orbital planes (4th Keplerian Element) Inclination (above) (angle) i (below) Ascending Node Equatorial Plane ( defined by Earth’s equator ) Sample inclinations Ascending Node is where a satellite crosses the equatorial plane moving south to north 0° -- Geostationary 52° -- ISS 98° -- Mapping Right Ascension [1] 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 plane Perigee Direction ω = angle from ascending node to perigee on the orbital plane ω Ω Ascending Node [1] Vernal Equinox Right Ascension is the astronomical term for celestial (star) longitude. 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 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 w = 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 Sun Drifts 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. (above) Ascending Node (below) Nodal Regression can be very useful. 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). 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 Changes Burn must take place where the initial and target planes intersect. Even a small amount of plane change requires lots of ΔV θ 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 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 1. launch to ‘GTO’ 3. Hohmann circularizing burn 3. Second Hohmann burn circularizes at GEO ‘Super GTO’ GEO Target Orbit 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 1. Launch to ‘Super GTO’ 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. Rendezvous Launch 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’ February 2009 Iriduim / Cosmos collision created > 1,000 items > 10cm diameter 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. Deep Space Cassini – Saturn orbit insertion using good ‘ol fashion rocket power. Using Lagrange Points to ‘stay put’ Halo Orbits (stability from motion) AeroBraking Earth, 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 Assignments for April 2 Reading on Orbits: SMAD ch 6 – scan 5 and 7 TLOM ch 3 and 4 – scan 5 and 17 HOMEWORK: Design minimum two, preferably three orbits your mission could use. Create a trade table to compare orbit designs. For the selected orbits: Trade criteria should include: Describe it (orbital elements) How will you get there? How will you stay there? Estimate perturbations Orbit suitability for mission. Cost to get there – and stay there. Space environment (e.g., radiation). Engineering 176 Orbits