NAVAL SPACE

3-2 The solar wind travels at one million mph, carrying various particles from the sun. The interaction of the solar wind with the Earth’s magnetic field produces a cavity in the interplanetary medium known as the Earth’s Magnetosphere (See Figure 3-2). If the solar wind did not exist, we would have a decreasing dipole magnetic field extending into space from Earth indefinitely. However, the incoming solar gas compresses the field on the sunlit side of the Earth and sweeps magnetic field lines from near the poles back into a long tail. When the solar wind reaches Earth and interacts with its magnetic field, electric currents are formed that travel along magnetic field lines. Much like a magnet, the current is forced along the magnetic field lines down to the Polar Regions and excites the nitrogen and oxygen molecules in the ionosphere. This electrical energy is converted to light creating what is called the aurora, which is seen as a yellow-green light in the sky. It appears in many forms such as arcs with rays, bands, pulsation surfaces, and draperies. The aurora usually occurs in the northern latitudes of 65° to 70° and is known as the Aurora Borealis or Northern Lights. It also occurs in the Southern Hemisphere, where it is called the Aurora Australis. SOLAR WIND EARTH’S ORBIT _ _ _ HIGH SPEED STREAM IN THE SOLAR WIND SOLAR SECTOR BOUNDARY + EARTH’S MAGNETOSPHERE + + ++ + + + + + ++ + + _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ + + + + INTERPLANETARY MAGNETIC FIELD LINES + _ + + _ Figure 3-1. Solar Wind This excited state of atmospheric molecules degrades radar performance in the auroral zones, including ballistic missile warning radar. It can also adversely affect satellites at altitudes to 600 miles, to include polar orbiting satellites.

3-3 Figure 3-2. General Configuration of the Magnetosphere In addition to the solar wind, the sun can also emit an explosive burst of electrically charged particles called a “solar flare.” Generally, the stronger the solar flare, the greater the intensity of a particle stream, and the more severe the impact of the event on space systems operating in that environment. The main cause of solar activity is the solar flare which occurs within a relatively small region of the sun’s atmosphere. Flares are characterized by the stronger than normal X-ray, ultraviolet, optical and/or radio wave emissions. All of these wavelengths travel at the speed of light and reach the Earth in about eight minutes. Impacts are almost entirely limited to the daylight hemisphere since the rays do not penetrate or bend around the Earth’s surface. Normally, effects tend to last from only a few minutes up to about an hour or two although some lasting up to two days have been recorded.

3-4 Flares usually occur in the vicinity of “sunspots” or their pre-cursors, bright active regions called plage. Sunspots are transient dark spots on the surface of the sun. They appear as dark spots because they are cooler than the rest of the sun’s atmosphere. Individual sunspots have lifetimes that range from a few hours to several months, with most dissipating within several days of their appearance. The energy released by a flare is the energy stored in the intense, complex magnetic fields, which produce the sunspots. Some of the space systems that can be impacted by this radiation include communications satellites, navigation satellites, and radar. If the sun is in the field of view of a receiver and a burst is at the right frequency and intensity, Radio Frequency Interference (RFI) may occur. These electromagnetic impacts are almost entirely limited to the Earth’s sunlit hemisphere and occur simultaneously with the solar flare that caused them (See Figure 3-3). RADIO BURST EFFECTS RADAR INTERFERENCE SUN RADIO BURST SATCOM INTERFERENCE Figure 3-3. Radio Burst Effects Solar activity is cyclic in nature, following a 11-year cycle which is called the Solar Cycle (See Figure 3-4). Generally there is a 4-year rise to a solar maximum, followed by a gradual 7-year decline to solar minimum.

3-5 Also related to this phenomena are geomagnetic storms. Geomagnetic storms are worldwide events that normally occur on Earth a day or two after a large solar flare erupts on the sun, and have an occurrence frequency that is directly related to the 11-year solar cycle. Geomagnetic storms are created when a “gust” of the solar wind compresses the Earth’s magnetic field and are recorded at geomagnetic observatories as a sudden change in the intensity of the local magnetic field. Since geomagnetic storms greatly hamper communications, it is fortunate that their effects usually die out within a few days. THE SOLAR CYCLE SOLAR MINIMUM SOLAR MAXIMUM SOLAR CYCLES 19-23 19 20 21 22 Figure 3-4. The Solar Cycle Also, because the sun rotates, the emissions experienced in the near-Earth environment can vary as the sun’s active regions rotate around to the other side of the sun. A full rotation is completed about every 27 days, so active regions that “disappear” from our perspective on Earth, may appear a couple of weeks later as they move back into view. ELECTRICALLY CHARGED PARTICLES Both low and high earth-orbiting spacecraft and satellites are subject to a number of environmental radiation hazards, such as direct physical damage and/or electrical upsets caused by charged particles (See Figure 3-5).

3-6 PARTICLE RADIATION ENERGETIC PARTICLES ELECTROMAGNETIC RADIATION SOLAR WIND MAGNETOSPHERE POLAR CUSP THE “SOLAR WIND” IS A CONTINUOUS OUTFLOW OF ENERGETIC CHARGED PARTICLES (PROTONS & ELECTRONS). Figure 3-5. Particle Radiation High Energy Particles These are primarily protons and electrons, but occasionally cosmic rays can reach the Earth within fifteen minutes to a few hours after the occurrence of a strong solar flare. The major impact of these particles is over the polar caps, where the protons have ready access to low altitudes through the funnel-like cusps that are created by the Earth’s magnetic field lines that terminate at the North and South poles. These impacts can last from a few hours to several days depending on the intensity of the flare. Potential impacts include satellite disorientation, physical damage to spacecraft, false sensor readings, navigation errors, and absorption of HF radio signals. Very high-energy protons or ions are capable of penetrating completely through a satellite. As they pass through, they will ionize particles deep inside the satellite. In fact, a single proton or cosmic ray can, by itself deposit enough charge to cause an electrical upset (circuit switch, spurious command or memory change or loss) or serious physical damage to on-board computers or other components. Hence these occurrences are called “single event upsets” (SEU) and are depicted in Figure 3-6. SEUs are very random, almost unpredictable events. They can occur at any time during the 11-year Solar Cycle. In fact, SEUs are actually most common near solar minimum, when the interplanetary magnetic field emanating from the sun is

3-7 weak and unable to provide the Earth much shielding from cosmic rays originating outside the Solar System. SINGLE EVENT UPSETS • CAUSE - HIGH ENERGY PROTON OR COSMIC RAY • EFFECT - CAN PENETRATE COMPLETELY THROUGH SATELLITE AND IONIZE MATERIAL DEEP INSIDE • RESULT - SINGLE PARTICLE CAN CAUSE PHYSICAL DAMAGE AND/OR DEPOSIT ENOUGH CHARGE TO CAUSE AN ELECTRICAL UPSET (CIRCUIT SWITCH, FALSE COMMAND, OR MEMORY CHANGE/LOSS) OR PHYSICAL DAMAGE MAGNETOSPHERE INTERPLANETARY MAGNETIC FIELD Figure 3-6. Single Event Upsets Additionally, charged particles may be trapped in the Van Allen Radiation Belts. The Outer and Inner Van Allen Radiation Belts are two concentric, donut-shaped regions of stable, trapped charged particles that exist because the geomagnetic field near the Earth is strong and field lines are closed. The Inner Belt has a maximum proton density approximately 5,000 km above the Earth’s surface and contains mostly high-energy protons produced by cosmic ray collisions with the Earth’s upper atmosphere. The Outer Belt has a maximum proton density at an altitude ranging from 16,000 to 20,000 km and contains low to medium energy electrons and protons whose source is the influx of particles from the magneto-tail during geomagnetic storms. These radiation belts can have serious impact on satellite operations. Communications satellites in “Geosynchronous” orbit (35,782 km or 22,235 statute miles altitude) suffer whenever the Van Allen belt moves inward or outward. Satellites in a semi-synchronous orbit such as GPS satellites suffer from a variable, high-density particle environment. Both orbits are particularly vulnerable to the directed motion of charged particles

3-8 that occurs during geomagnetic storms. Space vehicles in low circular orbit (125-130 miles) however, receive an insignificant amount of radiation from the Van Allen belts (See Figure 3-7). Van Allen Radiation Belts GEOSYNCHRONOUS ORBIT SEMI-SYNCHRONOUS ORBIT (NAVSTAR GPS) OUTER VAN ALLEN BELT INNER VAN ALLEN BELT Figure 3-7. Van Allen Radiation Belts An additional problem occurs when satellites rely on electro-optical sensors to maintain their orientation in space. These sensors lock onto certain patterns in the background stars and use them to achieve precise pointing accuracy. Such star sensors are vulnerable to cosmic rays and high-energy protons, which can produce flashes of light as they impact a sensor. The bright spot produced on the sensor may be falsely interpreted as a star. When computer software fails to find this false star in its star catalogue or incorrectly identifies it, the satellite can lose attitude lock with respect to the Earth. Directional communications antennae, sensors, and solar cell panels would then fail to see their intended targets. The result may be loss of communications with the satellite, loss of satellite power and, in extreme cases, loss of the satellite due to drained batteries (gradual star sensor degradation can also occur under constant radiation exposure). Disorientation occurs primarily when solar activity is high and on geosynchronous or polar- orbiting satellites (See Figure 3-8).

3-9 SATELLITE DISORIENTATION PARTICLE STREAM (PROTONS) ? FALSE STAR Figure 3-8. Satellite Disorientation from False Star Images Low to Medium Energy Particles Streams of lower energy particles may arrive at the Earth about two to three days after a flare. These particles can occur at any time due to other non-flare solar activity. The radiated particles cause geomagnetic and ionospheric storms that can last from hours to several days. Most common impacts include spacecraft electrical charging, increased drag on low orbiting satellites, radar interference, space tracking errors, and radio wave propagation anomalies. These impacts are most frequently experienced in the night-side hemisphere of the Earth. The intense ionospheric irregularities found in the Earth’s auroral zones are also a cause of “scintillation” at high geomagnetic latitudes. Scintillation is the rapid, random variation in signal amplitude, phase and/or polarization caused by small scale irregularities in the electron density along a signal’s path. The result is signal fading and data drop-outs on satellite command uplinks, data down-links or communications signals.

3-10 Scintillation tends to be a localized impact. Only if the signal path penetrates the ionospheric region where these small scale electron density irregularities are occurring will an impact be felt. Low altitude, nighttime links with geosynchronous communications satellites are particularly vulnerable to intermittent signal loss due to scintillation. GPS satellites, which are located at semisynchronous altitude, are vulnerable to ionospheric scintillation. In particular, scintillation can cause a GPS receiver to lose signal lock with a particular satellite which may result in a potentially less accurate position fix. Unlike other solar phenomena, there is no fielded network of ionospheric sensors capable of detecting real-time scintillation occurrences. Presently, space environmental forecasters are heavily dependent on their known association with other environmental phenomena. Another source for space object positioning errors is that of atmospheric drag (See Figure 3-9) on low orbiting objects (less than 1,000-km altitude). Energy deposited in the Earth’s upper atmosphere by charged particle bombardment heats the atmosphere, causing it to expand outward over a period of time. This produces more frictional drag on a satellite than expected and decreases its altitude while increasing its speed. Consequently, the satellite will be some distance below and ahead of its expected position when a ground radar or optical telescope attempts to locate it. Conversely, just the opposite conditions result when exceptionally calm solar and/or geomagnetic conditions cause less atmospheric drag than predicted and the object is higher and behind its expected location. ATMOSPHERIC DRAG - ORBIT CHANGES EXPECTED POSITION ACTUAL POSITION Figure 3-9. Atmospheric Drag on Satellites.

3-11 The consequences of atmospheric drag include: • Inaccurate satellite locations which can hinder rapid acquisition of SATCOM links for commanding or data transmission; • Costly orbit maintenance maneuvers may become necessary; • De-orbit predictions may become unreliable. An additional problem resulting from charged particle bombardment during a geomagnetic storm or proton event is potential damage to a launch vehicle or satellite. For example, an electric charge can be deposited on or inside the spacecraft. The electrostatic charge deposited may be discharged without serious impact by on-board electrical activity such as vehicle commanding. However, on occasion, this discharge has damaged payload circuitry. SUMMARY Despite engineering efforts, satellites are still susceptible to solar events (See Figure 3- 10). In fact, with newer microelectronics and their lower operating voltages, it will actually be easier to cause electrical upsets than on older, simpler vehicles. Furthermore, with the perceived lessening of the man-made nuclear threat, there has been a trend to build new satellites with less nuclear radiation hardening. This previous hardening had also protected the satellites from space environmental radiation hazards. SOLAR EMISSIONS & IMPACTS SPACECRAFT CHARGING & DRAG SPACETRACK ERRORS LAUNCH TRAJECTORY ERRORS RADAR INTERFERENCE RADIO PROPAGATION ANOMALIES POWER BLACKOUTS ELECTROMAGNETIC RADIATION ARRIVAL: IMMEDIATELY DURATION: 1-2 HOURS HIGH ENERGY PARTICLES ARRIVAL: 15 MIN TO FEW HOURS DURATION: DAYS LOW-MEDIUM ENERGY PARTICLES ARRIVAL: 2-3 DAYS DURATION: DAYS X-RAYS, EUV, RADIO BURSTS PROTON EVENTS GEOMAGNETIC STORMS SATCOM INTERFERENCE RADAR INTERFERENCE SHORTWAVE RADIO FADES SATELLITE DISORIENTATION FALSE SENSOR READINGS SPACECRAFT DAMAGE LAUNCH PAYLOAD FAILURE HIGH ALTITUDE AIRCRAFT RADIATION SHORTWAVE RADIO FADES Figure 3-10. Solar Emissions and Impacts

3-12 EFFECTS ON SPACECRAFT AND MATERIALS During the design process, engineers analyze the hazards and risks that result directly from the effects of the natural environment, and attempt to minimize their impact on operations. The following provides an overview of the primary hazards found in the space environment, and highlight the factors that restrict the tactical use of satellites. Upper Atmosphere Density Variations Density is the number of molecules per unit volume. Density of the atmosphere varies as a result of the balance between the gravitational force on molecules of different masses and the thermal energy of these molecules. A good approximation of this variation is that up to 100 miles altitude, air density decreases by a factor of 10 every 10 miles. Above 100 miles, the decrease is more exponential as the lighter elements (such as hydrogen and helium), become more predominant. Our atmosphere actually extends for thousands of miles above the Earth’s surface, but in ever decreasing densities. Space, of course, provides an excellent vacuum, far better than that obtainable in any Earth laboratory. While this is an important advantage for many purposes, it can also cause serious problems. Designers and operators must be careful about the problem of outgassing of volatiles from spacecraft hardware exposed to the vacuum. Escaping volatiles can condense on cooler external surfaces, or sometimes change the chemical properties of the substance from which they escaped. This condition can be a serious problem for optical instrumentation, where the deposition of even a very thin layer of outgassed material on lenses or mirrors can have adverse effects. Also, total outgassing takes a considerable amount of time (for some materials, upwards of 90 days has been recorded), so under normal conditions it should not be expected that a spacecraft would provide tactical information immediately following launch and orbit insertion. Another problem associated with operating in a hard vacuum is that joints between spacecraft components in a vacuum tend to "cold weld" to each other preventing freedom of motion. In Earth’s atmosphere, an extremely thin layer of air adheres to mostly all surfaces, acting as a lubricant between materials. Without this "natural" lubricant, substitutes must be developed and utilized. Common lubricants used routinely on Earth often boil away in space, so special substances must be employed in their place. Some metals are stronger in the hard vacuum of space. If a crack forms in a metallic surface on the Earth where air surrounds the metal, air molecules immediately enter the crack and a chemical reaction with the metal occurs. In some instances the reaction causes a wedging action to take place and the crack is enlarged, thus weakening the metal. In a hard vacuum, this chemical reaction does not happen.

3-13 Long-term variations in the extreme ultraviolet (EUV) and soft X-ray emissions from the sun change the amount of upper atmospheric heating, which can affect the drag on LEO satellites. On a shorter time scale, plasma injections during geomagnetic disturbances are also an important source of upper atmospheric density variations. Although restricted to high latitudes, the atmospheric response to these storms can alter the orbits of polar-orbiting satellites, causing tracking and positioning errors. Thermal Variations There are two ways in which the temperature of an object or particle can be classified: • That which we can feel and measure with a thermometer because of its density (usually measured in degrees) • That which is inherent within a particle due to its energy level. Although space is relatively cold when measured with a thermometer, the various particles and plasma constituents may contain high equivalent temperatures in the thousands of degrees. Spacecraft have limits of heat and cold which they can withstand for a specific period of time. In space, the temperature of an unprotected object will rise rapidly on the sunlit side, while simultaneously dropping to very low temperatures on the shaded side. Various types of materials are employed which reflect sunlight and insulate a spacecraft to maintain acceptable temperatures. Additionally, operational equipment generates heat that must be dissipated or dumped by some means. The only way to do this in a vacuum is by radiating the excess heat into space, which is a much less efficient process than doing so by convection. Radiation Space radiation hazards can emanate from several sources: • Solar flares • Energetic solar particles • Trapped particles • Artificial events (such as a nuclear detonation in space) • High energy galacticions Space radiation consists of protons (p+), electrons (e-), neutrons (n), photons and HZE. Due to interactions in the solar system with radiation sources, radiation varies as a function of time and location. Although HZE has only an estimated 1-2% influence, it contributes 50% of the total radiation dosage received.

3-14 Proton radiation particles can travel at the speed of light and hence can reach Earth and LEO in 8 minutes. However, longer times are usually experienced because (1) the particles diffuse out from the sun; (2) electrically charged particles spiral around the magnetic fields (See Figure 3-11) instead of traveling linearly; (3) particles are disturbed by shock waves. Figure 3-11. Typical path of charged particles in a magnetic field. The Earth is protected from space radiation by both its strong geomagnetic field and the depth of its atmosphere. At some locations, such as the poles, the magnetic protection is minimum yet the atmospheric protection is sufficient. Satellites in LEO and low inclination are generally shielded by the magnetic field, not by the atmosphere. However, due to the structure of the magnetic field and ionic trapping effects, there are areas where the magnetic field collects particles rather than acts as a shield. On hardware, especially solid state electronic devices operating in space, SCR can generate enough electrons to change the state of a circuit element, or produce a "bit flip" (changes the electronic state of a 1 to a 0 in computer memory), resulting in software errors or permanent damage. Radiation can also cause interference such as noise, or impact damage leading to degraded operation or destruction of the equipment. Also, as we have seen, the near-Earth space environment is not a total vacuum. This means that the spacecraft is constantly being bombarded by atmospheric particles moving at relatively fast velocities. Some of these particles are the monatomic oxygen atoms (O) created by photo-dissociation. When these atoms strike a satellite, being a highly reactive element, they tend to combine with the spacecraft materials and create potentially damaging corrosion. The effects of this corrosion are quite small, but could build up over long periods of time in space, especially on the exposed surfaces in the direction of flight.

3-15 Sensors used in spacecraft are usually very sensitive to particular wavelengths of energy. Direct exposure to solar radiation may result in unwanted signal reception and possible sensor damage. If the purpose of a system is to pick up the weak radiation of a far away planet or star, and the sensor happens to inadvertently look in the direction of the sun (or a reflection of some of its energy), the desired signals may be obscured in the sun’s radiation and the sensor itself may be damaged. Additionally, the sensors themselves must be protected from the general effects of the space environment described above, or performance may be adversely affected. Spacecraft Charging The term spacecraft charging refers to the variation in the electrostatic potential of a spacecraft surface with respect to the surrounding plasma, where tens of thousands of volts can develop between the two. This can occur in LEO, as well as in deep space and can result in structural and electrical equipment damage. Although the build-up of static charge may affect certain spacecraft sensors, the real danger lies in any possible resulting discharge or arcing because structural damage is a real possibility. Even weak discharges may bring about spurious electronic switching, the breakdown of thermal coatings, and solar cell and optical sensor degradation. The major sources of charging that change a spacecraft’s potential are as follows: • Electrons and ions from the surrounding thermal plasma • High energy particles • Secondaries produced as a result of external particle impacts • Photoelectrons • Secondaries from internal particle sources, such as radioisotope thermal generators (RTGs) Which of these sources is most important depends on the region of space involved. In near Earth space, thermal electron and photoelectron fluxes predominate. These two sources work against each other in that captured thermal electrons will charge a spacecraft negatively while ejected photoelectrons will leave behind a positively charged spacecraft. These effects are more pronounced for spacecraft having odd shaped exteriors with protuberances, depressions, holes, etc. Sometimes the surface potential varies considerably from one part of a spacecraft surface to another. Because there are so many spacecraft at geosynchronous altitude, spacecraft charging has been studied there in particular. Space vehicles at this altitude are susceptible to plasma injection events that accompany geomagnetic disturbances and substorms. Such events may occur several times per day even on quiet days, and may produce a ten-fold increase in ion densities and a thousand-fold increase in electron densities at geosynchronous orbit. Such charging is most likely near local midnight (spacecraft time) and not during daytime. Because of

3-16 the anomalous behavior of some military spacecraft at geosynchronous orbit, the USAF carried out a research program to investigate the phenomenon of spacecraft charging at high altitude (SCATHA). The results demonstrated the concept of active discharging by a plasma gun to reduce spacecraft potentials safely when necessary, and provided a computer code for calculating satellite surface potentials. SCATHA developed a method for modeling active charging/discharging and created an atlas of the geosynchronous environment. MACROSCOPIC BODIES Macroscopic bodies range in size from small micrometeoroid particles to man-made objects, large meteoroids, and satellite debris. When compared to natural space debris, man-made debris usually travels at slower orbital velocities. Natural space bodies, on the other hand, can enter and cross through the path of the Earth and thus have considerably higher relative velocities and a correspondingly higher destructive impact potential. Meteoroids and Micrometeoroids In addition to the ionized particles of the solar wind and magnetic fields, interplanetary space also contains a considerable amount of solid matter, most of which is in the form of small particles called interplanetary dust. The entire assortment of all the solid pieces in the interplanetary space is called the meteoritic complex. The term meteoroid refers to a particle while it is moving in space. When a meteoroid enters the atmosphere and begins to glow, it is called a meteor. If the same particle survives the journey through the atmosphere and hits the Earth, the remnant is called a meteorite. A few meteorites are large (the largest ever found weighed about 50 tons and there is crater evidence for even larger ones), but most are very small and are called micrometeorites. Meteoroids move with speeds between about 30,000 mph to 160,000 mph. At these speeds the impact of a large meteoroid on a satellite would be catastrophic. Impacts between micrometeoroids and a satellite would not necessarily be catastrophic, but could erode the satellite’s surface, a potentially serious hazard to optical surfaces or to surfaces used for thermal control or solar power generation. Because of the conjectural nature of the process, meteoroids were frequently blamed when early satellites ceased to function. A number of satellites, including Explorers 16 and 23 and the three satellites in the Pegasus series, have been used to study the problem. The present conclusion is that the probability of a satellite being hit catastrophically by a large meteoroid is very small, but the probability of its being hit by many small micrometeoroids is quite large.

3-17 HARDNESS AND SURVIVABILITY REQUIREMENTS Survivability is the ability of a space system to perform its intended function after being exposed to a stressing environment created by an enemy or hostile agent. Hardness is an attribute defining the environmental stress level which a space system can survive. A military space system or commercial satellite must be survivable if we will need its services in times of high stress, such as a nuclear war. To do this, we must understand what may cause the system to malfunction and design it to protect against failures. Survivability requirements include identifying the environments and their intensities and, in most cases, designing the space system so it will continue to perform its intended function for a certain time after exposure. Commercial or scientific s atellites usually do not need to be survivable, but planners must be aware that an unhardened satellite may stop operating after even very distant nuclear explosions. A slight hardening of satellites can make them much more survivable. It is important to consider survivability from the outset of mission design. For example, if the satellite can function within a range of orbit altitudes, the highest of these is both the hardest to attack and the most expensive to reach. We should consider the system’s survival in each of its life cycles phases, including concept definition, engineering design and development, and operations in orbit. Historically, we have not hardened launch systems because of cost and weight, as well as undefined need. The main threats against space systems are nuclear weapons, including directed energy designs such as X-ray lasers; ground – and space-based radio frequency (microwave) weapons; homing kinetic energy weapons; and beam weapons using neutral atomic particles. We may use several approaches to make a system survivable, with hardening of the satellite as a key element. THE NUCLEAR WEAPONS ENVIRONMENT AND ITS EFFECT ON SPACE SYSTEMS Nuclear weapons pose the most severe threat to spacecraft or space systems. The yield, or explosive power, and accuracy of delivery are such that if a nuclear weapon directly attacks a spacecraft, ground station, or any other node of a space system, the node will be destroyed. Nuclear weapon yields can range from a few tons to many megatons of TNT equivalent (one kiloton of TNT is defined to be 10 12 calories). Future nuclear exchanges could use yields of a few hundred kilotons to a few megatons, depending on the purpose of the specific attack and the weapon’s delivery accuracy. Accurate delivery of low yields will achieve the desired kill probability, whereas less accurate delivery requires higher yields. Approximately 80 percent of the energy from a nuclear weapon detonated in space appears in the form of X-rays. Other important effects include small amounts of gamma rays and neutrons, as well as small fractions in residual radioactivity and kinetic energy of bomb debris.

3-18 SPACE ENVIRONMENTAL SUPPORT The 55th Space Weather Squadron (55SWXS) 55SWXS is DoD’s only space environmental analysis and forecasting facility. The squadron is a 24-hour support operation providing tailored space environmental products and services to DoD and national program customers. The 55SWXS headquarters is at Schreiver AFB, Colorado and operates several Geographically Separated Units (GSUs) to monitor the sun. Known as the Solar Electro-Optical Network (SEON), it is the only network in the world dedicated to observing the sun at optical and radio wavelengths in real time. Mission 55SWXS provides space environmental support for worldwide operations The squadron gathers and processes space environmental data from ground and space-based sensor networks, analyzes and models the space environment, forecasts solar and space environmental phenomena, and provides alerts, warnings and assessments for operational impacts to Air Force and other DoD agencies. Support to customers can be provided at the unclassified, collateral and Sensitive Compartmented Information (SCI) levels. Systems supported include satellite vehicle and payload operations, ground and satellite-based communications, navigation, surveillance, and weapon system radar, as well as high-altitude reconnaissance aircraft and the Space Shuttle. 55SWXS products fall into one of four categories: • Parameter Obervations: The 55SWXS monitors solar activity through the data received from SEON, other ground-based ionospheric sounder networks, and satellite-based sensors. Critical parameters from this data are used to optimize tailored environmental models used in specifying satellite locations and enhancing HF and satellite communication links, as well as radar and satellite tracking correction and calibration. • Analysis: Near-Real Time and Post Analysis. This category gives system operators, engineers, and decision makers expert analyses of the role the space environment plays in system anomalies. This provides quicker resolution of anomalies, reducing system down-time, and saving time searching for other causes. • Forecasts: 24-Hours, days, months and years. All portions of the radio spectrum are subject to variability in the ionosphere. The 55SWXS provides predications of critical parameters for optimizing HF and satellite communication operations and planning, satellite drag predication, and radar and satellite signal correction.

3-19 • Warnings: Navigation systems are influenced by energetic proton flux into the polar caps as well as geomagnetic activity. Also, energetic protons pose a significant health hazard to high-altitude reconnaissance aircraft pilots and astronauts operating in the space environment. Satellite systems in certain orbits can perform anomalously or be damaged during solar flare induced particle storms. The squadron provides situational awareness products, potential systems effects, and aids in system anomaly resolution in support of radar, satellite vehicle, and payload operations. MAN-MADE DEBRIS In the short time that man has been able to place objects in space, he has already created a serious situation with the pollution of the environment consisting of the debris created by launch vehicles and spacecraft operations. At present U.S. Space Command has counted approximately 10,000 artificial space objects in near Earth vicinity, where the probability of a spacecraft collision with one of these objects is greater than that of a collision with a natural meteoroid. SUMMARY The sun shapes the characteristics of terrestrial space, causing phenomena that can significantly affect operation of Earth-orbiting satellites. In some situations, these phenomena can be predicted and negative effects may be avoided. However, present technology does not allow us to circumvent all negative effects caused by the space environment. Until we better understand the space environment we will be at the mercy of this harsh operating medium.

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4-1 CHAPTER 4 ORBITAL MECHANICS Note: A film entitled "Spaceflight: The Application of Orbital Mechanics," jointly produced by the Naval Space Command and the National Aeronautics and Space Administration (NASA), is available to supplement the following text on Orbital Mechanics. This 35-minute film uses state-of-the-art computer graphics to clearly portray the fundamentals of orbital mechanics in a simple, understandable manner. The film is available in all formats (3/4", VHS, Beta) and can be acquired through the Fleet Audiovisual Libraries. It is also available at no cost through the Defense Audiovisual Information System (Product Identification Number 804859DN). INTRODUCTION In order to speak and understand the “space language” being introduced into Naval operations, a basic knowledge of the fundamentals of orbital mechanics is essential. Consequently, this chapter is designed to provide you with a layman’s understanding of the field of Celestial or Orbital Mechanics – a basic explanation of the way objects move in orbit around the earth. A short section on the history of celestial mechanics is included to give you an appreciation of the permanency and complexity of this, the most basic science governing our operations in space. The study of trajectories and the orbits of vehicles in space is not a new science. However, the application of the concepts of celestial mechanics to man-made vehicles is less than a century old. To differentiate, celestial mechanics is mainly concerned with the determination of trajectories and orbits of the stars and planets in space. Orbital mechanics uses the same procedures and techniques, but applies them toward the computation of orbits and trajectories of man-made objects in outer space. (Astronomers say “celestial mechanics,” people in the space operations business say “orbital mechanics.”) EARLY ASTRONOMY To fully understand the concepts and laws associated with orbital mechanics, it is necessary to go back to the very beginning of civilization. It is impossible to state with certainty when the earliest quantitative observations of the heavenly bodies were made; however, it is known that many early civilizations recognized the pattern and regularity of the motions of the stars and the planets. Some made an effort to track and predict celestial events. In particular, the invention of the calendar required an elementary knowledge of astronomy.

4-2 The Chinese had a working calendar at least as early as the fourteenth century B.C. They also maintained accurate records of comets, meteor showers, meteorites, and other phenomena. Egyptian astronomers were able to roughly predict the flooding of the Nile River each year near the time when the star Sirius could be seen rising in the dawn sky just before the sun. The Bronze age people in Northwestern Europe built many monuments such as Stonehenge, which was certainly used as a crude calendar. One of the earliest known attempts to describe the location of the planets, Sun, Moon, and their motions with respect to each other was by Pythagoras. He assumed that the Earth was stationary at the center of a system comprised of the seven known moving objects visible to the unaided eye−the Sun, Moon, Mercury, Venus, Mars, Jupiter, and Saturn. Pythagoras theorized that all revolved about the Earth in complex spherical orbits. This was known as the geocentric theory. Aristotle (384-322 B.C.), one of the most famous of Greek philosophers, understood such phenomena as the phases of the Moon and eclipses. Aristotle considered the theory that the apparent daily motion of the sky could be explained by a hypothesis of the rotation of either the Earth or of the celestial sphere. Perhaps most importantly though, Aristotle considered the possibility that the Earth revolves around the Sun, instead of the more popular idea that the Earth was the center of the universe. Aristarchus of Samos (310-230 B.C.), another famous Greek astronomer, was the first to profess a belief in the heliocentric theory−that the Earth moves about the Sun. He also believed that the stars must be extremely distant to account for the fact that their apparent positions in the sky remain the same all year long. However, because Aristarchus’ ideas were too revolutionary, they were rejected, and the geocentric theory continued to be accepted for centuries. About A.D. 140, Claudius Ptolemy amplified the geocentric theory by developing an elaborate geometrical representation of the solar system, which predicted the apparent motion of the planets with considerable accuracy. According to his theory, the planets revolve about imaginary planets, which in turn revolve around the Earth. Ptolemy’s hypothesis, with some later modifications, was accepted as absolute authority throughout the Middle Ages. MODERN ASTRONOMY It was not until some 1800 years after Aristarchus had first proposed the heliocentric theory, that a Polish monk named Nicolaus Copernicus (1473-1543) published a book in defense of the heliocentric system. Copernicus postulated that the Earth was one of the six (then known) planets that revolve around the Sun. Starting nearest the Sun, he ordered the planets Mercury, Venus, Earth, Mars, Jupiter, and Saturn. He further deduced that the closer a planet is to the Sun, the greater its orbital speed. Thus, the retrograde (or, backward) motion of Mars, Jupiter, and Saturn was easily explained. He also calculated the approximate scale of the

4-3 solar system. Contrary to what has been a popular belief, Copernicus did not prove that the Earth revolves about the Sun. A Danish astronomer, Tycho Brahe (1546-1601), established an astronomical observatory in 1576, where for 20 years he carried out the most complete and accurate astronomical observations made up to that time. In 1600 Brahe was joined by a young German mathematician named Johannes Kepler (1571-1630). Brahe put Kepler, an early convert to the heliocentric theory, to work on finding a satisfactory theory of planetary motion. LAWS OF MOTION Kepler’s most detailed study was of Mars, for which Brahe’s observational data was the most extensive. For 10 years he attempted to fit combinations of circular motion to the observed motion of Mars. Finally, Kepler tried to represent the orbit of Mars with an oval and soon discovered that the orbit could be fitted by a curve known as an ellipse. KEPLER’S LAWS OF PLANETARY MOTION Kepler found that Mars has an orbit that is an ellipse, with the Sun at one focus (the other focus of the ellipse is empty, an unoccupied point in space). Kepler generalized that what is true for Mars must be true for the other planets as well. This generalization has become known as Kepler’s First Law of Planetary Motion. Kepler’s First Law of Planetary Motion (Law of Ellipses) Each planet moves in an elliptical orbit with the Sun a t one focus and the other focus empty. This law also applies to man-made objects orbiting the Earth (i.e., satellites). The orbit of each satellite is an ellipse with the center of the Earth at one focus and the other focus unoccupied. This law simply implies that the purpose of the occupied focus (i.e., the Earth) is to provide gravitational attraction to the satellite to keep the satellite in its elliptical orbit (see Figure 4-1). Figure 4-1. Kepler’s 1st Law applied to an Earth satellite.

4-4 Kepler’s Second Law of Planetary Motion (Law of Areas) The line joining a planet to the Sun sweeps over equal areas in equal time intervals. As applied to an Earth-orbiting satellite, the line joining it to the Earth sweeps over equal areas in equal periods of time (see Figure 4-2). This law implies that the speed of a satellite changes depending on its distance from its gravitational attraction source, the center of the Earth. A satellite’s speed is greatest at the point in the orbit closest to the Earth, and is slowest at the point farthest from the Earth. If a satellite is traveling in a circular orbit, then the speed of the satellite is constant. Figure 4-2. Kepler’s 2nd Law. It is important to understand that the orbit followed by a satellite is not dependent on its mass. A large, heavy satellite could be in the same orbit with a small, light satellite with each sweeping out equal areas in equal periods of time. Kepler’s Third Law of Planetary Motion (Law of Harmonics) For any planet, the square of its period of revolution about the Sun is directly proportional to the cube of its mean distance from the Sun. When applied to earth satellites, this

4-5 law explains that the farther a satellite is from the Earth, the longer it will take to complete its orbit, the greater the distance it will travel, and the slower its average speed.

4-6 This law is expressed as follows: p2/d3 = K where p is the period d is the distance K is a constant with the same value for all satellites GALILEO GALILEI An Italian mathematician, Galileo Galilei (1564-1642), contributed greatly to the understanding of the behavior of objects at rest or in motion. In addition to being the first person to develop and use a telescope, he accumulated a great deal of evidence in support of the heliocentric theory. In 1632, he published the "Dialogue of the Two Great World Systems," which examined all arguments for and against the heliocentric theory. NEWTON’S LAWS OF MOTION AND UNIVERSAL GRAVITATION While declaring his three laws of planetary motion, Kepler deduced that a force from the Sun pulled on the planets, but he did not determine the mathematical nature of this force. Galileo discovered some of the basic laws governing the behavior of physical objects. Sir Isaac Newton (1643-1727), regarded as the father of classical mechanics, drew upon the work of both Kepler and Galileo to formulate the Law of Universal Gravitation and the three Laws of Motion. While Kepler’s laws provided a conceptual model of orbital motion, Newton’s laws provided the foundation for the mathematical description of orbits and why a satellite remains in orbit. Newton’s First Law of Motion (Law of Inertia) A body in motion will keep moving at the same speed and in the same direction unless acted upon by an external force. This explains why a satellite moves in a curved path around the Earth, because the Earth’s gravitational pull acts as an external force on it. Newton’s Second Law of Motion (Law of Momentum) If the sum of forces acting on an object is not zero, the object will have an acceleration proportional to the magnitude and in the direction of the net force. This provides some of the explanation for changes in a satellite’s orbit due to external forces other than the Earth’s gravity, and the need to adjust a satellite’s orbit.

4-7 Newton’s Third Law of Motion (Law of Action & Reaction) For every action, there is an equal and opposite reaction. This law explains how a satellite gets into orbit. Newton’s Law of Universal Gravitation Any two objects in the universe attract each other with a force directly proportional to the product of their masses, and inversely proportional to the square of the distance between them. Mathematically, this law is expressed as follows: F = G*(m 1*m2)/d2 where F is the force acting between the bodies G is the universal constant of gravitation m is the mass of a body, and d is the distance between the bodies Simply stated, the more massive the objects, or the closer they are together, the greater the gravitational pull between them. ORBITAL PARAMETERS To better understand the operational application of satellites, you should become familiar with the fundamental terms and relationships associated with satellite orbits. ORBITAL INSERTION To illustrate how a satellite is placed into orbit, refer to Figure 4-3. Imagine someone standing atop a tall mountain and throwing a ball horizontally. Not considering atmospheric friction, the Earth’s rotation, or any other force except gravity, the path of the ball will curve downward due to gravity and it will strike the Earth (Path A). Now, if that person fires a gun horizontally, the bullet will obviously go farther than the ball, but it will still be pulled down to the Earth (Path B). However, if a projectile is fired horizontally at a very high speed, approximately 17,500 miles per hour, the curvature of its path due to gravity will match the curvature of the Earth below it. The projectile will then continue to "fall" around the Earth just as fast as the Earth curves away from the projectile, and become an Earth-orbiting satellite (Path C).

4-8 Figure 4-3. Orbital insertion. For an object to achieve orbit, sufficient energy must be provided so that the path does not intersect the surface of the Earth. However, the object cannot be given too much energy or it will escape the effects of the Earth’s gravity. The following three conditions are required to place an object in orbit: • The object must be put above most of the Earth’s atmosphere (approximately 60 miles high) to negate the effects of atmospheric friction • The object must be imparted with a speed of approximately 17,500 miles per hour within about 200 miles of the Earth’s surface • The object’s speed must be in a direction parallel to the surface of the Earth. POSSIBLE SATELLITE ORBITS Suppose that an object is boosted to an altitude of approximately 200 miles above the Earth’s surface, then turned so that it is horizontal to the Earth, and finally provided with a forward horizontal motion. The object will enter into an orbit the size and shape of which depends on the exact direction and speed of the object at "burnout" (when thrusting terminates). If the object is moving horizontally to the earth, the possible types of orbits it can enter are depicted in Figure 4-4.

4-9 Figure 4-4. Possible satellite orbits. If the object’s burnout velocity is slightly greater than that required for a perfect circular orbit, the resulting orbit will be elliptical with the center of the Earth at one focus. The point farthest from the center of the Earth will be the point of burnout. If the burnout velocity is substantially below the circular-orbit velocity requirement, the object will strike the Earth’s surface. If the burnout velocity is just slightly below the circular-orbit velocity, the object may barely clear the Earth’s surface, but atmospheric drag would soon cause it to slow down below the required orbital velocity and it will strike the Earth. If the burnout velocity is exactly the circular-orbit velocity requirement, a circular orbit results (in practical situations, this is nearly impossible to achieve). Burnout velocities equal to or greater than the escape velocity from the Earth’s gravitational field result in parabolic or hyperbolic orbital paths. At these velocities, satellites will escape the gravitational pull of the Earth and never return. ORBITAL TERMS AND ELEMENTS When an object’s burnout velocity results in an elliptical orbit, a point on the orbit farthest from the center of the Earth is referred to as apogee (apogee – “away”). The point closest to the center of the Earth will be halfway around the orbit and is called perigee (see Figure 4-5). Apogee and perigee are always 1/2 revolution apart. For a circular orbit, apogee and perigee altitude are equal.

4-10 Figure 4-5. Orbital parameters. Orbital Elements Six parameters, often referred to as the "orbital element set," establish the size, shape, and orientation of an orbit in space as well as the location of a satellite in its orbit. These parameters, some of which are depicted in Figure 4-6, are as follows: • Semi-major axis • Eccentricity • Inclination • Right ascension of the ascending node • Argument of perigee • Time

4-11 Figure 4-6. Orbital elements. The semi-major axis is simply o ne-half the diameter of an ellipse. The length of the semi-major axis is used to define the size of an orbit. From this, the orbital period (the time it takes a satellite to complete one orbital revolution) can be calculated. Eccentricity defines the shape of an ellipse. For all ellipses, the value of eccentricity lies between zero and one. The larger the value, the more elliptical the orbit. Circular orbits have an eccentricity equal to zero. A satellite orbit with an eccentricity equal to or greater than one will escape the Earth’s gravitational field (parabolic or hyperbolic orbit). The semi-major axis and the eccentricity of the orbit are determined before launch to design the orbit for that particular satellite’s mission. Inclination is used to orient the orbital plane with respect to the Earth. It is the angular measurement made at the ascending node (the point at which a satellite crosses the equatorial plane going from south to north) from the equatorial plane to the orbital plane (see Figure 4-6). Inclination orients the orbital plane to the equatorial plane, and determines the northernmost and

4-12 southernmost latitudes covered by the satellite’s ground track. Inclination also defines another manner in which orbits are described: • Prograde: inclination > 0° but < 90° • Retrograde: inclination > 90° but < 180° • Polar: inclination equal to 90° • Equatorial: inclination equal to 0 ° or 180° The final element used to define the size and shape of the orbit and the orbital plane is Right Ascension of the Ascending Node. This is the angular measurement in the equatorial plane from the Vernal Equinox (that point determined by a line extending from the center of the Sun through the center of the Earth to the celestial equator at the start of spring) eastward to the ascending node. The remaining two orbital elements are used to position the orbital plane. The Argument of Perigee orients the orbit within the orbital plane. It is the angular measurement from the ascending node along the orbital path to the point of perigee. Once the perigee point has been determined, the point of apogee is 180° away. Finally, to determine where a satellite is in its orbit at a specific time, Epoch Time or True Anomaly is used. Epoch Time is an arbitrary time used as a starting point. Normally, the time when a satellite is at the ascending node is used as epoch time. True Anomaly is the angular measurement from the point of perigee along the orbital path to the location of the satellite at the epoch time. Although period is not categorized as an orbital element, it is very useful in satellite operations. Typically, the period of a satellite is the time it takes to travel from one ascending node point to the following ascending node point (referred to as the "nodal period"). The minimum period a satellite can have and still be in a stable, non-decaying orbit is approximately 87.5 minutes. If there is any less time, the object will eventually reenter the Earth’s atmosphere. REFERENCE SYSTEMS Since observational data for a satellite may be derived by any one of a wide variety of types and locations of sensors, an understanding of the common reference systems used is essential. Each reference system is designed for a particular use, fulfilling at least one of the following purposes: • to locate a specific object in space

4-13 • to discriminate between two or more objects in space • to describe motion of an object in space. Reference systems may be classified into two basic types, either inertial or non-inertial. An inertial reference system is one in which direction does not change with time. It is used in locating the position of a satellite to establish the orbital plane and the satellite’s orbit within that plane. A non-inertial reference system does change direction with time. The four reference systems typically used in space operations are summarized in Table 4-1 below: Table 4-1. Summary of Common Reference Systems COORDINATE SYSTEM ORIGIN USED TO LOCATE Geographic (Non-Inertial) Earth Center Earth Radius Tracking Station Topocentric (Non-Inertial) Tracking Station Slant Range Locate satellite Ref: Tracking Station Geocentric (Inertial) Earth Center Geocentric Radius Locate satellite in space Orbit (Inertial) Earth Center Celestial Radius Locate satellite in its orbit LAUNCH AND ORBITAL MANEUVERING Many factors enter into the mission planning and subsequent design of a satellite. One of the most restrictive facets of this planning is the launch environment (i.e., booster, launch site, launch "window"). LAUNCH AZIMUTH AND INCLINATION Launch azimuth is the direction from the launch site in which a booster is launched. The relationship between the location of the launch site and the launch azimuth to the resulting inclination of a satellite’s initial orbit is very specific. The minimum orbital inclination of a satellite is equal to the latitude of the launch site, and is achieved with a launch azimuth of due East. For launches with any azimuth other than due East, the orbital inclination will always be greater than the latitude of the launch site. Practically speaking, a satellite launched on an azimuth between 0° and 180 ° will have an inclination between 0° and 90 °, or a prograde orbit. Satellites launched on azimuths between 180° and 360° will have inclinations between 90° and 180°, or a retrograde orbit.

4-14 Most U.S. launches take place from one of two complexes: Kennedy Space Center, Cape Canaveral, Florida, or Vandenberg Air Force Base in California. If a satellite is launched with the Space Shuttle from Kennedy Space Center (which is located at a latitude of 28.5° North) on a due East azimuth, its orbital inclination will be 28.5° and the limits of its ground track will be between 28.5° North and 28.5° South latitude. Therefore, the limits of the ground track equal the launch inclination. Launch Site Limitations Many safety factors must be considered when launching from a particular site. Of course, one of the safety considerations is not to launch over populated areas. Consequently, there are limitations to the launch azimuths from each site. Because of safety considerations, the maximum practical inclination from a Kennedy Space Center launch is 57°. However, after an initial orbit is established, the inclination of an orbit can be changed by out-of-plane maneuvers (described later in this section), but this is costly in terms of on-orbit propellant. To obtain an orbit with an inclination greater than 57°, U.S. satellites are launched from Vandenberg AFB, CA. A significant advantage of launching from Vandenberg is the capability to economically achieve polar orbits, with ground tracks covering all latitudes from the North Pole to the South Pole. Another limitation induced by the location of a launch site is the affect of the Earth’s rotational velocity. The speed of the Earth changes with latitude, ranging from zero mph at the poles to about 1037 mph at the equator. Since the Earth rotates from West to East, all points on its surface have an eastward velocity, with the greatest eastward velocity occurring at the equator. The farther a launch site is from the equator, the less the Earth’s rotational velocity imparts energy to the booster. This results in requiring more fuel to get a satellite into orbit, or a tradeoff with the satellite’s payload weight to conserve on-orbit fuel. Launching from an equatorial site offers a significant advantage, and is an important consideration since many satellites operate in low-inclination orbits. Launch Window Satellite launches take place within a specified time interval referred to as the "launch window." Some of the factors affecting the launch window are: • Launch and orbit lighting conditions • Sun angles • Payload orbit requirements • Satellite system phasing

4-15 • Tracking and communications requirements • Collision avoidance with other space objects. During the winter months, the available launch window for lighting conditions alone can be as little as 3 hours per day. When combined with other factors, launch windows become very constrained. ORBITAL MANEUVERING It is rare that a satellite is launched directly into its final orbit. A satellite will typically need to change its orbit at least once to be able to perform its mission. For example, the Space Shuttle may deploy a communications satellite designed for placement in a geosynchronous orbit. The Shuttle itself cannot reach geosynchronous altitude, so a small booster attached to the satellite has to be pointed in the proper direction at the right time and place, and thrusted for a precise length of time. Then, once it gets to geosynchronous altitude, it will have to thrust again to stay in its desired orbit. Mission Considerations After a satellite has been on orbit for a period of time, it is often necessary to change its orbit by thrusting. Firing a spacecraft’s thrusters results in a Delta "V"−a change in the satellite’s velocity. The amount of fuel used during a burn depends on the desired velocity change and the mass of the satellite. Since the amount of fuel carried is limited, fuel consumption is one of the primary considerations in satellite mission planning, and is critical to mission life. On orbit, a satellite can thrust in any direction. Thrusting along the flight path, forward or backward, is the most common. Forward thrusting increases a satellite’s velocity and is known as a "prograde burn." With prograde burns, the orbital path of the satellite will be raised at all points except the burn point. Thrusting opposite to the direction of the orbit, which slows a satellite down, is called a "retrograde burn." For retrograde burns, the orbit will be lowered at all points except the burn point. Maneuvers Most maneuvers may be performed any place in the orbit; however, there are certain points that minimize energy requirements for a particular type of maneuver: • The most economical method for obtaining a change in orbital period is by applying thrust at perigee or apogee

4-16 • The most economical place to perform an inclination change is at either the ascending or descending node

• The most economical place to perform a change in right ascension of the ascending node is at either of the tw o midway points between the ascending and descending nodes

• The change in inclination or right ascension is easiest if it is performed at apogee

• The most economical place to perform a maneuver to change perigee height is at apogee.

In-Plane Maneuvers

When a satellite needs to change its altitude, period, or eccen tricity, additional energy is required. This energy requirement is true whether or not the element's value is being increased or decreased. The type of tran sfer used to meet this energy requirement depends on the satellite's mission, and the amount of fuel available.

The most energy efficient in-plane ma neuver is known as the Hohmann Transfer (see Figure 4-7). The Hohmann Transfer is a two-impulse maneuver between two co- planar orbits.

Courtesy of Damon, Thomas D. (2001) Introduction to Space: The Science of Spaceflight, Third Edition. Krieger Publishing Company. Figure 4-7. Hohmann transfer.

In accomplishing a Hohmann Transfer, two applications of thrust are required. Each thrust changes the speed of the satellite and places it in a new orbit. If an increase in altitude is

4-17 desired, the point of departure becomes the perigee of the tran sfer orbit and the point of injection into the higher circular orbit becomes the apogee of the transfer orbit. The use of the Hohmann Transfer minimizes the velocity change required, with the advantage of using minimum fuel. The disadvantage of the H ohmann Transfer is th at it takes longer than most other transfers. The Hohmann Transfer is used extensively with interplanetary satellites, because it does not require a large maneuver engine or fuel tanks.

Another way to accomplish an altitude change is referred to as the Fast Transfer (see Figure 4-8). This transfer is most useful when time is a critical factor. In the Fast Transfer, the transfer orbit crosses the final orbit at an angle. The burn is performed in two increments but requires substantially more fuel than the Hohmann Transfer. It is, however, the quickest way to get to the final required orbit. The Fast Transfer is typically used by surveillance satellites to reposition over new target areas.

Courtesy of Damon, Thomas D. (2001) Introduction to Space: The Science of Spaceflight, Third Edition. Krieger Publishing Company. Figure 4-8. Fast transfer.

Out-of-Plane Maneuvers

To change inclination or right ascension, an out-of-plane maneuver is required. The inclination maneuver changes only inclination and maintains right ascension. This plane change requires one burn at either the ascending node or descending node of the original orbit. The amount of inclination change depends on the length of the burn.

The right ascension maneuver changes only right ascension without changing inclination. This plane change requires one burn anywhere in the original orbit except at the ascending or

4-18 descending nodes. The maximum effect is achieved when the burn is performed midway between the nodes. ORBIT TYPES AND APPLICATIONS The laws of nature force the orbits of all satellites to lie in planes that pass through the center of the Earth. A satellite’s ground track is formed by the intersection of the surface of the Earth and a line between the Earth’s center and the satellite. As the satellite moves in its orbit, this intersection traces out a path on the ground below it. GROUND TRACK If the Earth did not rotate, a satellite would retrace the same ground on each revolution (see Figure 4-9). Notice that the maximum latitude North and South of the equator over which the satellite passes is equal to the satellite’s inclination. Figure 4-9. Example of ground track (non-rotating Earth). The orbital plane of a satellite remains fixed in space as the Earth rotates under it. The effect of this rotation is to displace the ground track westward on each successive revolution of the satellite by the number of degrees the Earth turns during one orbital period (see Figure 4-

4-19 10). This is referred to as nodal regression. The resulting ground tracks provide operators with an indication of the position of their satellite in accomplishing its mission.

4-20 Figure 4-10. -Example of ground track (rotating Earth). As mentioned above, a satellite in either a circular or an elliptical orbit will trace out a path over the Earth between the limits of latitude as determined by the inclination angle. A satellite in a circular orbit will spend equal amounts of time North and South of the equator. However, a satellite in an elliptical orbit will remain North or South of the equator for unequal periods of time. Only satellites that are in circular orbits travel along their ground track at a constant velocity. When an orbit is inclined to the equator, the component of satellite velocity in the direction due East or due West varies continuously throughout the orbit. More specifically, a satellite in a nearly circular orbit moves slower in an easterly or westerly direction at the equator and faster when it is at its most northerly or southerly points. The relative speeds of satellites in elliptical orbits vary even more. FIELD OF VIEW The field of view of a satellite is defined as the area of the Earth’s surface that is in view from the satellite at any given time. Satellites in high orbits have greater fields of view than those in lower orbits. For example, a satellite at an altitude of 800 nmi has a circular field of view with a diameter of about 4,100 nmi. A satellite at 200 nmi has a circular field of view with a diameter of about 2,000 nmi.

4-21 TYPES OF ORBITS The various types of orbits are the result of different satellite missions, and the desire for satellites to perform their missions over different areas of the Earth’s surface (i.e., have different ground tracks). Low Earth Orbit A satellite is considered to be in a low earth orbit (LEO) at altitudes between approximately 150 and 800 miles above the Earth’s surface. At an altitude of approximately 150 miles, a satellite’s period will be about 90 minutes. The Space Shuttle and some scientific satellites are typically placed in low inclination, low earth circular orbits. Polar Orbit In contrast to a low inclination orbit and its latitude limitations, a polar orbit passes over the entire surface of the Earth. A polar orbit has an inclination of 90° and is usually circular (see Figure 4-11). Due to the ability to pass over the entire surface of the earth throughout the course of several days, the polar orbit is used extensively by imagery satellites. Figure 4-11. Polar orbit. Geosynchronous Orbit A satellite placed in orbit with an average altitude of approximately 19,300 nautical miles ( nm) will have an average angular velocity exactly equal to that of the Earth’s. Stated more simply, the satellite would have a period approximately equal to one day. This means that it would take as long for the satellite to complete one revolution around the Earth, as it takes for the earth to rotate once about its axis. Such an orbit is called a geosynchronous orbit.

4-22 If a geosynchronous orbit with an inclination of 0° were perfectly circular, the satellite would appear to remain stationary in space above the same point on the Earth’s surface. This is referred to as a geostationary orbit (see Figure 4-12).This orbit is predominantly used by relay satellites to provide a continuous communications capability among ground stations within their very broad field of view. Some surveillance and warning satellites also use the geostationary orbit. The geostationary field of view is constant, covering nearly one-third of the Earth’s surface with latitude limitations of approximately 70° North and South of the equator. Effective satellite communications from a geostationary orbit is not possible at either pole. Figure 4-12. Field of View from a Geostationary Orbit. Elliptical Orbit To obtain satellite communications capability in the northern or southern latitudes, a highly eccentric elliptical orbit, commonly referred to as a Molniya orbit is used (see Figure 4-

4-23 13). The Molniya orbit has an apogee nearly equivalent to the geosynchronous altitude and an inclination of approximately 63° to 64°. A satellite in this type of highly eccentric elliptical orbit slows down at apogee in the Northern Hemisphere (providing longer duration over its greatest field of view) and whips through perigee (smallest field of view) in the Southern Hemisphere. This provides communications in the Northern Hemisphere for nearly 75 percent of the satellite’s orbital period. If the apogee/perigee points were shifted 180°, this orbit would cover the Southern Hemisphere. Figure 4-13. Molniya orbit. Semi-Synchronous Orbit An average orbit with an altitude of approximately 10,800 nmi results in a period of about 12 hours and is referred to as a semi-synchronous orbit. The purpose of placing satellites in this type of orbit is to allow a user to receive signals from more than one satellite at any time. Navigation and certain kinds of area communications (Molniya) satellites use the semi- synchronous orbit. Sun-synchronous Orbit The sun-synchronous orbit takes advantage of the precession of the orbital plane caused by the Earth not being a perfect sphere. All sun-synchronous orbits are highly inclined (normally with inclinations between 95° and 105 °) retrograde orbits that precess eastward around the Earth’s polar axis at the rate on one revolution per year. Since the Earth-Sun line also revolves eastward at the rate of one revolution per year, the orbital plane will maintain a

4-24 constant orientation relative to the Earth-Sun line (see Figure 4-14). If the satellite’s orbital period is then synchronized with the rotation of the Earth, it will pass over the same point on the Earth’s surface at the same local time at a regular interval. Figure 4-14. Sun-synchronous orbit. A sun-synchronous satellite ensures that a constant sun angle and uniform lighting exists for the same field of view from revolution to revolution. For certain mission requirements, a noon/midnight sun-synchronous orbit can be selected that would provide good photography for about one-half of every revolution. Most meteorological and earth resources LEO satellites are placed in sun-synchronous orbits, imaging the entire Earth on a regular schedule. ORBITAL PERTURBATIONS AND DECAY There are other forces besides the Earth’s gravity (referred to as perturbations) which act on an orbiting satellite. Although much smaller, these other forces are of sufficient magnitude to cause orbital decay, and result in significant changes in a satellite’s ground track over time if not recognized and corrected for periodically. Changes in a satellite’s ground track can result in reduced mission effectiveness. To neutralize the effects of natural perturbations, a satellite may perform "station-keeping" thruster burns.

4-25 ORBITAL PERTURBATIONS The Earth is not a perfect sphere. The North Pole region is more pointed than the flatter South Pole region, and a bulge exists at the equator. This bulge, referred to as the Earth’s oblateness, causes the equator to be slightly elliptical. Earth’s Asymmetry These asymmetrical conditions have an influence on the orbital parameters of satellites in low-and-medium- altitude orbits. One effect of the Earth’s asymmetry is positive, in that the nodal regression of an orbit can be timed to make an orbit sun-synchronous without using precious fuel to maneuver for that purpose. Atmospheric Drag The Earth’s atmosphere does not suddenly end, it gradually tapers off into interplanetary space. Generally speaking, atmospheric drag circularizes and decreases the apogee of a low earth orbit. Any drag caused by air resistance can normally be disregarded above 300 nmi. Occasionally, however, a period of extreme solar activity may occur that will heat the atmosphere and cause it to expand. Third Body Effects Newton’s Law of Universal Gravitation indicates that there is a force of attraction (gravity) between all objects and bodies in the universe. The gravitational pull of "third bodies" such as the Moon and the Sun can affect the orbits of geosynchronous and deep space satellites. Radiation As discussed previously, the Sun is continually expelling matter in the form of ionized gas. The particles in this gas (mostly electrons and protons) that penetrate interplanetary space move with high velocities. These particles, commonly referred to as the solar wind, exert a pressure on satellites, particularly those with large area-to-mass ratios. This pressure induces a restraining force on satellites, and is only present on the daylight (or, windward) side of the Earth. This causes irregular perturbations of a satellite’s orbit. Electrons and protons can become trapped when encountering the Earth’s magnetic field. They oscillate back and forth along the lines of magnetic force, and since the magnetic field completely encircles the Earth, the trapped particles completely encircle the Earth. This region of trapped particles, mentioned earlier in Chapter 3, is known as the Van Allen Radiation Belt.

4-26 The belt has an inner and outer portion. The inner Van Allen belt starts at an altitude of approximately 250 to 750 miles, depending on the latitude. It extends to about 6,200 miles where it begins to overlap the outer belt. This inner belt extends from 45° North latitude to about 45° South latitude. The outer Van Allen belt begins around 6,200 miles and extends to an altitude that varies from 37,000 to 52,000 miles. The upper boundary is dependent on the activity of the Sun. Electromagnetic Forces The Earth’s magnetic field extends far into space. As a satellite orbits the Earth, it is traveling through this magnetic field. The electronic components of a satellite also produce a magnetic field that consequently reacts with the Earth’s magnetic field. Additionally, the ions and electrons in the Earth’s magnetic field collide with the satellite, causing a negative charge on the satellite’s surface. The negative charge is larger on the day-side of an orbit than on the night side. The interaction of these induced fields with the Earth’s magnetic field causes a magnetic drag to act on a satellite. This drag can cause charging or torquing of a satellite. ORBITAL DECAY AND DEORBIT For satellites that pass close to the Earth (low orbit or highly elliptical orbits), we can arrange for the satellite to re-enter, or let it re-enter by itself. Deliberate re-entry of a satellite with the purpose of recovering the vehicle intact is deorbiting. We usually do this to recover something of value: people, experiments, film, or the vehicle itself. The natural process of spacecraft (or any debris: rocket body, payload, or piece) eventually re-entering Earth’s atmosphere is decay. In some situations the satellites are in such stable orbits that natural perturbations won’t do the disposal job for us. In these situations, we plan to remove the satellite from the desirable orbit. To return a satellite to Earth (or low Earth orbit), it would take just as much energy as it did to place it in orbit. Obviously it is impractical to return old satellites to Earth from a high orbit. We usually just boost the satellite into a slightly higher orbit to get it out of the way, and there it will sit for thousands of years to come. SUMMARY The overall purpose of understanding the principles of orbital mechanics is to recognize the elements that affect the design and planning of a satellite’s mission. Many mission planning factors and constraints, such as on-orbit fuel, satellite weight, and launch site considerations, were mentioned in this chapter. However, one factor essential to the planning of a satellite’s mission remains to be discussed, and that is the selection of a satellite booster, or launch vehicle.

5-1 CHAPTER 5 LAUNCH AND RECOVERY SYSTEMS INTRODUCTION This chapter provides an overview of the systems and operations associated with the launch and recovery of spacecraft. The discussion includes a description of the U.S. launch and orbit transfer vehicles used to place payloads into orbit, an overview of launch and recovery operations, a description of the three major U.S. launch sites, a summary of the major ground processing activities that contribute to a successful launch and recovery, and a look at some future launch systems. It should be noted that the information contained in this chapter represents only a snapshot of the current U.S. launch vehicle inventory. As you might expect when dealing with a dynamic industry, this inventory will change as technologies emerge, evolve, and are replaced or removed. BACKGROUND During the early years of the space program, both the National Aeronautics and Space Administration (NASA) and the Department of Defense ( DoD) developed launch vehicles to satisfy their own specific requirements. Consequently, there has been a wide variety of launch vehicles used to support the space program. DoD developed the Atlas, Delta, and Titan series of launch vehicles from ballistic missile technology. Meanwhile, NASA developed the Scout and Saturn. Used to send Apollo crews to the Moon, both are now out of production. All these launch vehicles, which can only be used once, are called expendable launch vehicles (ELVs) and each has a different capability to put satellites in orbit. From the least to the most capable in terms of lift, they are Pegasus, Taurus, Delta, Atlas, and Titan. The Pegasus is a recent addition to the inventory of expendable launchers and is aircraft-launched. The Taurus is a ground-launched, mobile vehicle. Currently, there is no one vehicle that can launch all satellites into all the orbits required for the various missions. NASA and DoD select the vehicle best suited for their particular mission based on the size, weight, and desired orbit of the payload. In 1972, President Nixon approved NASA’s plan to create a reusable launch vehicle called the Space Shuttle, and directed that it become the primary U.S. launch vehicle, replacing all ELVs except the Scout. This made both NASA and DoD dependent on a single launch vehicle for access to space but, in theory, the resulting high launch rate for the Shuttle would significantly reduce the cost per flight. The Shuttle was first launched in 1981, and was declared operational in 1982. The phase-out of ELVs began. However, in 1984, concerned about having “assured access to space,” DoD successfully argued that it needed a “complimentary” ELV as a backup to the Shuttle and initiated what became known as the “Titan IV Program.” Production lines for the Delta and Atlas launch vehicles were closed down in anticipation that only the Shuttle and Titan IVs would be used by the end of the 1980s.