Info

engines = major weight reduction

• Low orbit deployment and check-out

• Can have reusable transfer vehicle

• Failure recovery possible

transfer to GEO

• High radiation exposure

• Needs autonomous transfer lor cost efficiency

final orbit, as illustrated in Fig 7-9C. In this case the total efficiency will approximate that of a two-burn Hohmann transfer, because all of the energy is being provided near perigee or apogee, as it is for the Hohmann transfer. With a low-thrust chemical transfer, we can deploy and check out a satellite in low-Earth orbit where we can still recover it before transferring it to a high-energy orbit where we cannot. Low-thrust transfer provides substantially lower acceleration and, therefore, a more benign environment Also, we are more likely to be able to recover a satellite if the propulsion system fails. The principal disadvantage of low-thrust chemical transfer is that it is a very nontraditional approach. Wertz, Mullikin, and Brodsky [1988] and Wertz [2001] describe low-thrust chemical transfer further.

Another type of low-thrust transfer uses electric propulsion, with extremely low acceleration levels—at levels of 0.001 g or less [Cornelisse, Schôyer, and Wakker, 1979]. Transfer therefore will take several months, even when the motors are thrusting continuously. Consequendy, as Fig. 7-9D shows, electric propulsion transfer requires spiralling out, with increased total AV(see Table 7-9). We need far less total propellant because of electric propulsion's high ls„. Electric propulsion transfer greatly reduces the total on-orbit mass and, therefore, tne launch cost. However, much of the weight savings is lost due to the very large power system required. In addition, the slow transfer will keep the satellite longer in the Van Allen belts, where radiation will degrade the solar array and reduce mission life.

Flybys or gravity-assist trajectories can save much energy in orbit transfers. Because they must employ a swing-by of some celestial object, however, missions near Earth do not ordinarily use them. Gravity-assist missions can use the Earth, but the satellite must first recede to a relatively high altitude and then come back near the Earth.* For a more extended discussion of gravity-assist missions, see Kaufman, Newman, and Chromey [1966], or Wertz [2001]. Meissinger et al. [1997, 1998] and Farquhar and Dunham [1998] have separately proposed interesting techniques for using a different orbit injection process to substantially increase the mass available (and, therefore, reduce the launch cost) for high-energy interplanetary transfers.

FireSat Transfer Orbit. We assume that FireSat will be launched into a 150-km, circular parking orbit at the proper inclination and need to determine how to get to the operational orbit of 700 km. For now, we assume some type of orbit transfer. When the spacecraft weight becomes better known and a range of launch vehicles selected, another trade will be done to determine whether it is more economical for the launch vehicle to put FireSat directly into its operational orbit.

The FireSat orbit transfer AV from Table 7-3 is a modest 309 m/s. It is not worth the added cost, solar array weight, or complexity for electric propulsion transfer. There is no reason for a high-energy transfer. We are left to select between a Hohmann transfer and a low-thrust chemical transfer. The Hohmann transfer is the traditional approach.

Low-thrust chemical transfer provides a more benign transfer environment and the potential for low-orbit deployment and checkout so that satellite recovery would be a possibility. The propulsion system would be lighter weight and require less control authority. We may be able to do the orbit transfer using just the mission orbit control modes and hardware which could completely eliminate a whole set of components and control logic.

For FireSat we will make a preliminary selection of low-thrust chemical transfer. This is non-traditional, but probably substantially lower cost and lower risk. Later in the mission design, the launch vehicle may eliminate this transfer orbit entirely.

7.5.2 Parking and Space-Referenced Orbits

In parking or space-referenced orbits, the position of the spacecraft relative to the Earth is unimportant except for blockage of communications or fields of view. Here the goal is simply to be in space to observe celestial objects, sample the environment, or use the vacuum or low-gravity of space. These orbits are used, for example, for space manufacturing facilities, celestial observatories such as Space Telescope and Chandra X-Ray Observatory, or for testing various space applications and processes. Because we are not concerned with our orientation relative to the Earth, we select such orbits to use minimum energy while maintaining the orbit altitude, and possibly, to gain an unobstructed view of space. For example, Sun-synchronous orbits may be

'Using the Earth for a gravity assist was first proposed by Meissinger [1970], appropriate for maintaining a constant Sun angle with respect to a satellite instrument Another example is the parking or storage orbit: a low-Earth orbit high enough to reduce atmospheric drag, but low enough to be easy to reach. We may store satellites (referred to as on-orbit spares) in these orbits for later transfer to other altitudes.

An interesting class of orbits which have been used for environmental monitoring and proposed for space manufacturing are libration point orbits or Lagrange orbits, named after the 18th century mathematician and astronomer, Joseph Lagrange. The Lagrange points for two celestial bodies in mutual revolution, such as the Earth and Moon or Earth and Sun, are five points such that an object placed at one of them will remain there indefinitely. We can place satellites in "orbit" around the Lagrange points with relatively small amounts of energy required to maintain these orbits. (For more details, see Wertz [2001].)

7.6 Constellation Design

In designing a constellation, we apply all of the criteria for designing a single-satellite orbit. Thus, we need to consider whether each satellite is launchable, survivable, and properly in view of ground stations or relay satellites. We also need to consider the number of satellites, their relative positions, and how these positions change with time, both in the course of an orbit and over the lifetime of the constellation.

Specifying a constellation by defining all of the orbit elements for each satellite is complex, inconvenient, and overwhelming in its range of options. A reasonable way to begin is by looking at constellations with all satellites in circular orbits at a common altitude and inclination, as discussed in Sec. 7.4. This means that the period, angular velocity, and node rotation rate will be the same for all of the satellites. This leads to a series of trades on altitude, inclination, and constellation pattern involving principally the number of satellites, coverage, launch cost, and the environment (primarily drag and radiation). We then examine the potential of elliptical orbits and the addition of an equatorial ring. The principal parameters that will need to be defined are listed in Table 7-10. After exploring the consequences of some of the choices, we will summarize the orbit design process in Sec. 7.6.2. A more detailed discussion is given in Wertz [2001].

No absolute rules exist. A constellation of satellites in randomly spaced low-Earth orbits is a serious possibility for a survivable communications system. The Soviet Union has used a constellation of satellites in highly eccentric Molniya orbits for decades. Various other missions may find satellite clusters useful. One of the most interesting characteristics of the low-Earth orbit communications constellations is that the constellation builders have invested billions of dollars and arrived at distinctly different solutions. For example, a higher altitude means fewer satellites, but a much more severe radiation environment (as discussed in Sec. 8.1), such that the cost of each satellite will be higher and the life potentially shorter. Similarly, elliptical orbits allow an additional degree of freedom which allows the constellation to be optimized for multiple factors, but requires a more complex satellite operating over a range of altitudes and velocities and passing through heavy radiation regimes. (See, for example, Draim [1985].) Because the constellation's size and structure strongly affect a system's cost and performance, we must carefully assess alternate designs and document the reasons for final choices. It is this list of reasons that allows the constellation design process to continue.

TABLE 7-10. Principal Factors to be Defined During Constellation Design. See Sec. 7.6.2 for a summary of the constellation design process.

Factor

Effect

Selection Criteria

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