1. GN/MAE1551 Basic Astronautics Overview MAE 155A G. Nacouzi

2. GN/MAE1552 Astronautics Fundamentals Definition Introduction Example Astronautics Project Overview Spacecraft (SC) Major Subsystems Orbital Mechanics Sample Orbits Interplanetary Missions Launch Systems Space Environment

3. GN/MAE1553 ME 155A/Astronautics Course Objectives Provide students with fundamental understanding of aerospace engineering (space application) –Basic knowledge of astronautics to help in decision in future career choices and provide sufficient skills to work in aero. industry –Provide basic skills needed to complete astronautics project in 155B

4. GN/MAE1554 Definition Astronautics: The science of designing, building and operating vehicles outside the atmosphere (manned and unmanned) –Aerodynamic forces are minimal at lower orbits and insignificant at higher orbits –Must consider different environment due to exo-atmospheric conditions –Will include the design of launch vehicles for spacecraft orbital insertion

5. GN/MAE1555 Introduction First successful spacecraft (SC) was launched in 1957 by the Soviet Union (Sputnik 1). Since then, we have made great advances in the use and exploration of space and have launched thousands of SC In addition to space exploration, SC have become a critical part of our scientific, commercial, civilian and military infrastructure (think Hubble telescope, communications, Radio, TV, weather, navigation and spy satellites) All indications are that the use of satellites will continue to expand and play a more critical part in our lives

6. GN/MAE1556 Introduction The current nominal cost to launch a SC into a high orbit is about $50,000 per kg –Ongoing research and development underway to lower launch costs => Make Space more affordable –Minimizing SC mass is important! –Maximizing lifespan of SC is also important –High reliability and survivability are required Therefore a good understanding of the SC environment and its effects on the SC are critical for success Need to properly design the SC & launch vehicle to attain proper balance between weight, performance, reliability and cost

7. GN/MAE1557 Launched on Space Shuttle in 1999 Highly Sensitive X Ray Telescope, Spectrometer & other devices. 14m x 20 m, 4784 kg, Largest payload launched by Space Shuttle 1.2 m Diameter Mirror, 7Angstrom roughness.

8. GN/MAE1558 Some Chandra Facts Highly Elliptical orbit –Apogee: 133,000 km; Perigee: 16,000 km=> Allows Chandra to spend most of its time (55 of 64 hrs) above Van Allen to allow uninterrupted observation Light from some quasars detected by Chandra has traveled about 12 billion years Can observe Xrays from particles up to the last second before falling into a black hole! Electrical power required to operate Chandra SC and instruments is about 2KW (as much as a hairdryer!) Ref: NASA

9. GN/MAE1559 National Polar Orbiting Operational Environmental Satellite System: Next Generation Environmental Monitoring Satellite System First Spacecraft Launch around 2011

10. GN/MAE15510 Astronautics Project Overview A project usually starts by first analyzing the customer mission requirements and developing an initial design concept –Mission requirements: interpretation & analysis of needs, with customer feedback, requirements definition & flowdown to lower levels –Technical considerations include purpose, cost, schedule, technology, risk and priority –Initial Design synthesis: develop a reasonable and achievable conceptual system design (architecture). Trade studies and other tools are used to support this effort. Presented to customer for feedback –Develop final system design & verify that system meets customer requirements

11. GN/MAE15511 Mission Objectives User Reqts Political Constraints Financial Constraints Mission Reqts PerformanceCoverage Reliability CostLife Ground Segment Ground Station Data Processing Launch Vehicle Volume Envt Stages SC System Reqts OrbitPower Config.Mass Ops Subsystem Reqts - Payload ThermalPowerCommunications StructureElectronicsACSPropulsion Mission Objectives & Requirements

12. GN/MAE15512 Initial Design Synthesis: Mission Concept Definition

13. GN/MAE15513 Major SC Subsystems Power (Peak, average) Attitude Determination & Control System (ADCS) => Pointing accuracy (include guidance and control) Propulsion (Attitude control, station keeping, orbital maneuvers, disposal) Structures (SC, deployment mechanisms, appendages) Communications (Timeliness, rates…) Thermal control (Temp limits) Command & Data Handling (C&DH)

14. GN/MAE15514 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

15. GN/MAE15515 Overview: Orbital Mechanics Study of S/C (Spacecraft) motion influenced principally by gravity. Also considers perturbing forces, e.g., external forces, 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 laws of planetary motion –Includes elliptical orbits of planets –Also developed Kepler’s eqtn which relates position & time of orbiting bodies

16. GN/MAE15516 Overview: Kepler Laws Kepler's three laws of planetary motion can be described as follows:  The Law of Ellipses: The path of the planets about the sun are elliptical in shape, with the center of the sun being located at one focus.  The Law of Equal Areas: An imaginary line drawn from the center of the sun to the center of the planet will sweep out equal areas in equal intervals of time.  The Law of Harmonies: The ratio of the squares of the periods of any two planets is equal to the ratio of the cubes of their average distances from the sun.

17. GN/MAE15517 Introduction: Orbital Mechanics Motion of satellite is influenced by the gravity field of multiple bodies, however, Two Body assumption will be used. 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 higher term of gravity field) –Gravity and centrifugal forces are principal forces

18. GN/MAE15518 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: F g = M m G/R 2 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

19. GN/MAE15519 Circular Orbits Equations Circular orbit solution offers insight into understanding of orbital mechanics and are easily derived Consider: F g = M m G/R 2 & F c = m V 2 /R (centrifugal F) V is solved for (by equating F g and F c ) to get: V=  (MG/R) =  (  /R) Period* is then: T=2  R/V => T = 2  (R 3 /  ) FcFc FgFg V R * Period = time it takes SC to rotate once wrt earth

20. GN/MAE15520 Some Orbit Types... Several types of orbits, some common ones: –Low Earth Orbit (LEO), R < 2000 km –Mid Earth Orbit (MEO), 2000< R < 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 (wrt to equator) = 90 degree –Molniya ~ Highly eccentric orbit with 12 hr period (developed by Soviet Union to optimize coverage of Northern hemisphere)

21. GN/MAE15521 Some Orbit Periods Orbits can be designed to have periods ranging from about an hour to 24 hours (stationary) and above

22. GN/MAE15522 Orbital Inclination Inclination is the angle between the orbital plane and the equatorial plane

23. GN/MAE15523 Basic Orbit Description SC in Elliptical orbits spend more time in the apogee ‘side’ of the orbit

24. GN/MAE15524 Spacecraft Earth Coverage: Effects of Orbits Area coverage of satellite increases for higher orbits. Note that resolution capability of an electro-optical (EO) sensor decreases as the orbit altitude is increased. High resolution (EO systems) need to be in lower altitude to maximize performance Nadir point

25. GN/MAE15525 Sample Orbits LEO at 0 & 45 degree inclination Elliptical, e~0.46, I~65deg Ground trace from i= 45 deg 45 deg latitude

26. GN/MAE15526 Sample GEO Orbit Figure ‘8’ trace due to inclination, zero inclination has no motion of nadir point (or satellite sub station) Nadir for GEO (equatorial, i=0) remain fixed over point 3 GEO satellites provide almost complete global coverage

27. GN/MAE15527 Interplanetary Missions A heliocentric coordinate system, i.e. Sun centered, is usually used for interplanetary trajectory analysis The SC is injected into a hyperbolic escape trajectory at a large distance from Earth For interplanetary travel, trajectory design is a highly complex task requiring the use of advanced concepts for trajectory optimization

28. GN/MAE15528 Interplanetary Missions Interplanetary trajectories –A simplified approach, patch conics method, provides a tool for preliminary analysis Considers the influence of a single body at a time SC is assumed to follow Keplerian motion while in the Sphere of Influence (SOI) of the body (SOI of Earth ~ 1E6 km) When SC leaves one SOI, it is captured by the SOI of the next body, and the trajectories are ‘patched’ together –Gravity Assist is a method used to rotate SC velocity vector and increase its magnitude wrt Sun SC is on a hyperbolic trajectory relative to a planet, its gravity field and motion are used to modify SC trajectory

29. GN/MAE15529 Atmospheric Entry As SC approaches planet with atmosphere, it experiences a large heat flux caused by the high velocity SC interacting with atmospheric gases The heat flux increases as the SC enters the continuum regime of the atmosphere –Rarefied (very low density) environment at high altitude causes relatively low heating rates – As the SC penetrates the atmosphere and reaches the continuum regime, substantial heat transfer from the flow to the SC occurs

30. GN/MAE15530 Launch Systems Typical launch system involves a ground based multi-stage rocket (Air launches are also used for small payloads, e.g., Pegasus) Rocket motors use liquid or solid fuel or a combination of both (hybrid) Boeing Delta IV Launch

31. GN/MAE15531 Liquid Rocket Engine Fuel and oxidizer are stored in different tanks and mixed in a combustion chamber to produce high pressure gas High Isp and engine can be throttled Requires liquid handling equipment

32. GN/MAE15532 Solid Rocket Motors Fuel and oxidizer are formed together into a solid propellant using a binder. No liquid handling equipment is needed Easier to handle, simpler and safer than liquid engine Not easy to throttle, lower Isp (efficiency)

33. GN/MAE15533 Launch Systems Simple ideal rocket equation is easily derived using a force balance approach to yield the incremental velocity delivered by a rocket motor: DV = Isp g 0 ln (M 0 /M) - DV g - DV d ; where, Isp ~ specific impulse (Thrust/mass flux); g 0 ~ gravity const. M 0 ~ Initial mass = M struct +M fuel +M payload, M ~ Final mass DV~ Velocity loss due to gravity; DV d ~ Velocity loss due to drag

34. GN/MAE15534 Launch Systems The required incremental velocity delivered by the launch system is a function of the orbit into which the SC needs to be injected. For circular orbits the circular velocity ranges from about 3 to 8 km/s plus the velocity losses due to drag, gravity and any other contributors Launch System required DV: Required DV = Orbital V + DV drag + DV gravity loss + other where, Orbital V ~ 3-8 km/s DV drag ~ losses due to drag (~ 0.2-0.4 km/s) DV gravity~ losses due to gravity (up to 2 km/s) Other ~ losses due to trajectory, thrust vector control (up to 2 km/s)

35. GN/MAE15535 Launch Systems In order to avoid carrying the total structure mass during the whole flight, multi-stage rockets are used, e.g., –Pegasus is a 3 stage air launch vehicle with total mass of 19000 kg, provides a DV of 8.1 km/s to a 455 kg payload. –An equivalent single stage rocket would provide the same payload with a DV of 5.4 km/s => For Pegasus, Multi stage improves performance by about 50%

36. GN/MAE15536 Orbital Velocity Requirements

37. GN/MAE15537 Space Environment Artist rendition of Ulysses

38. GN/MAE15538 Space Environment NOAA maintains space weather website for monitoring solar activity including Sun flares and solar wind conditions

39. GN/MAE15539 Space Environment Microgravity => Positive control of fluids is needed. Feeding mechanisms, pressurized tanks or bladders are needed since fluids ‘float’ in space Vacuum => Air or gases trapped in equipment exposed to space cause a pressure force and/or outgasing and potential contamination Solar => Solar heating can cause SC components or surface to exceed allowable temperature. Sun intensity in space ~ 1370 W/m 2. Solar flares can affect the operations of electronic components. Solar wind can affect the geomagnetic field around Earth

40. GN/MAE15540 Space Environment Radiation Envt: Radiation belts contain energetic protons and electrons trapped in the Earth magnetic field Plasma/Geomagnetic: Ionized gas in upper atmosphere. Plasma has an equal # of + & - charged particles. Rarefied Atm. => Drag, Erosion  Meteroids/Orbital Debris => Impact Hazards

41. GN/MAE15541 Space Environment Example: Radiation Radiation environment requires shielding of SC components Inner belt (Van Allen) has a maximum density (protons) at about 3000 km Outer belt max density is around 17000 km (e-) Solar wind distorts toroidal shape of belts

42. GN/MAE15542 Summary & Conclusions Development of an aerospace system involves understanding and analyzing mission requirements and synthesizing an effective design which integrates the different components of the system from launch and mission operations to deactivation and disposal Given the high costs associated with the development and construction of a SC and the costs of launch vehicles, understanding the SC operating environment and performance capability are critical to a successful system

43. GN/MAE15543 General References C.D. Brown “Elements of Spacecraft Design”, AIAA Education Series M.D. Griffin & J.R. French “Space Vehicle Design”, AIAA Education Series P. Fortescue et al. “ Spacecraft Systems Engineering” Wiley 2003 J. R. Wertz & W.J. Larson “ Space Mission Analysis and Design”, Microcosm 1999 J.J. Sellers et al. “Understanding Space” McGraw Hill 2004 NASA Website