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CHAPTER 16 LAUNCH OPPORTUNITIES AND SEASONAL ACTIVITY ON MARS CARL SAGAN and J. W. HAUGHEY The orbits of Earth and Mars are, within 1°51', coplanar. The sidereal period of revolution of the Earth is 365.257 days; that of Mars, 686.980 days. Thus, the synodic period, the time for a given configuration Sun- Earth-Mars to be repeated, is 779.94 days, or approximately 26 months. This is also the period between successive oppositions of Mars, that is, successive configurations in which the Sun, the Earth and Mars are in a straight line, with Earth in the middle. The time of closest approach of Mars to Earth during a synodic period occurs within a few days of oppo- sition. For this reason, most astronomical observations of Mars are made near opposition. Not all oppositions are, however, equally favorable, because the orbit of Mars is highly eccentric (e = 0.0933). Figure 1 illustrates the relative orbits of Mars and Earth. The eccentricity of the Martian orbit is discernible. The figure shows the dates of all opposi- tions between 1960 and 1975; and, for each opposition, the distance to Mars in miles and the diameter of Mars in seconds of arc. The ensemble of relative planetary configurations at opposition in Figure 1 constitute a Metonic cycle, a period of 15 or 17 years during which the configurations repeat themselves. The oppositions of closest approach are called "favorable" oppositions; in the present Metonic cycle, they occur in 1969 and 1971. Such favorable oppositions will not recur until 1984. For a given vehicle system, a spacecraft launched to Mars at certain 283

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284 APPROACHES AND REMOTE OBSERVATIONS Figure 1. Diagram of the orbits of Mars and \ Earth. * times during the synodic period requires much less energy than a launch at other times. The time of launch is a function of the desired payload, the desired flight time, and the three-dimensional configuration of Mars and Earth. In a given synodic period there is one time interval during which the energy required for a given planetary mission is minimized. This is called a launch window, or a planetary opportunity. Roughly speaking, Martian opportunities occur about three months before Martian oppositions. Since characteristic flight times are about 200 days, space- craft launched to Mars during the launch window leave the Earth before, and arrive at Mars after, opposition. An "opportunity" is denned by the existence of a launch vehicle with adequate thrust to place a spacecraft on a Martian transfer trajectory during the time interval of interest. The mission may be a fly-by, an orbiter, or a lander. In a fly-by, the spacecraft spends a brief period of time (of the order of a few hours) within close range of the planet. An orbiter may spend a much greater period of time at close range, depending upon the eccentricity of the orbit and the useful lifetime of the orbiter. A lander is a spacecraft that is designed to be deposited on the planetary surface (preferably at a selected site), and in most cases, is designed to give useful information about conditions at and near the surface for some period of time after landing. The simplest planetary mission is a fly-by. A planetary fly-by trajectory comprises, essentially, two segments: powered flight and ballistic flight. During powered flight, the necessary velocity vector is imparted to the vehicle. The powered flight trajectory desired is a function of the characteristics of the launch vehicle, the engineering constraints (payload, guidance, tracking, and telemetry), and the geo- metric constraints (launch site location, launch azimuth, and the relative configuration of the Earth and Mars). In the powered flight segment,

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Launch Opportunities and Seasonal Activity on Mars 285 either the direct ascent or the parking orbit mode may be adopted. In the direct ascent mode, the launch vehicle fires in one time interval only; at the end of the powered phase, the spacecraft must have the required velocity. In the parking mode, the spacecraft is first established in an approximately circular parking orbit about the Earth. After this has been done, one of the booster stages must be fired in order to leave the parking orbit for the planet. In general, the parking orbit mode is more efficient than the direct ascent mode, and allows a larger payload to be carried. The planetary transfer trajectories, after either direct ascent or parking modes, can be divided into two categories, Type I and Type II. If the path of a spacecraft traverses less than 180° of heliocentric longitude, it is said to be on a Type I trajectory. Should the range of heliocentric longitude traversed by a spacecraft lie between 180° and 360°, it is on a Type II trajectory. Spacecraft in Type I trajectories generally require more energy, but take less time to reach their destinations than do space- craft on Type II trajectories. Thus, it is possible for a spacecraft launched on a Type II trajectory to arrive at Mars after a spacecraft launched later, but on a Type I trajectory. Use of Type I and Type II trajectories is determined in part by the non-coplanarity of the orbits of Earth and Mars, and by the fact that it is difficult to launch two space vehicles from the same launch pad in a short period of time. In Figure 2, the required launch or injection energy per unit mass of spacecraft, C3, is shown for the 30-day launch windows in each opportunity between 1969 and 1979. The larger the value of C3, the smaller the payload, and the shorter the launch period during which a given launch vehicle may be used effec- tively. We note that by far the most favorable opportunities in the next INJECTION ENERGY PER UNIT MASS C»,(Km/S«c)1 KUICH WIDOW a KM OWI mi w mi mi ins HIT mi orroiiumiT Figure 2. Injection energy C3 required to trans- fer unit mass from Earth to Mars at six opportunities between 1969 and 1979. The parameter C3 is, by con- vention, twice the kinetic energy per unit mass.

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286 APPROACHES AND REMOTE OBSERVATIONS INJECTION ENERGY PER UNIT MASS Cs,( Km/Sec)2 40 30 20 IO 64 65 66 67 68 69 70 71 72 73 74 73 76 77 CALENDAR YEARS Figure 3. Variation of injection energy C3 with time of launch, for transfer of unit mass from Earth to Mars. The parameter C3 is, by convention, twice the kinetic energy per unit mass. decade and a half occur in 1969 and 1971. The injection energy per unit mass required in the middle 1970's is approximately twice that re- quired in 1969 and 1971. Not until 1984 will a launch opportunity as favorable as those of 1969 and 1971 recur. There are also two restrictions on utilization of planetary opportunities that are functions of the launch azimuth. They arise from limitations due to the declination of the outbound velocity vector, and due to range safety considerations. In general, if the declination of the outbound velocity vector is greater than the launch site latitude, a range of launch azimuths, symmetrical about due East, are excluded for a given launch site. If the outbound velocity vector declination is less than or equal to the launch site latitude, it is then possible to launch at all azimuths within the range safety limits. The range safety requirement restricts the use of launch azimuths that would cause overflight of populated areas, and unnecessarily jeopardize the lives and property of the inhabitants. A synopsis of the energies, C3, launch azimuths, and flight times for Martian opportunities within 30-day launch periods between 1969 and 1977 is shown in Table 1. From Figure 3, it is clear that great gains in cost, fuel, and mission reliability can be achieved by synchronizing the launch schedule with the firing windows. A given spacecraft, launched along a minimum-energy

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288 APPROACHES AND REMOTE OBSERVATIONS trajectory, will, of course, arrive at Mars at a known time during the Martian year. The astronomical observations suggest that not all arrival times are of equal interest. Much of the visual and polarimetric evidence that is suggestive of life on Mars has a seasonal character. Markedly periodic is the wave of darkening, a progressive albedo decline of the Martian dark areas (but not the bright areas) starting in local springtime from the edge of the vaporizing polar ice cap, and moving towards and across the equator. The wave of darkening has been observed photometrically by Focas [1959], and his measurements are shown in Figure 4. Here, we have plotted Martian latitude against heliocentric longitude. Following the astronomical convention, south is at the top. The heliocentric longitude is shown related to the Martian season and the day of the Martian year. Differences in the nature of the darkening between adjacent dark areas at the same latitude have been suppressed. The shading shows the high- contrast third, the medium-contrast third, and the low-contrast third of the wave of darkening. The contrasts are with respect to the neighboring bright areas, deserts of seasonally constant albedo. If, as has been re- peatedly suggested, the Martian wave of darkening is related to the spring- time metabolic activities of Martian organisms, then a spacecraft intended for biological exploration of Mars should be operative during the high- contrast third of the darkening wave, whether the vehicle is an orbiter or a lander. The southern wave of darkening is much more striking than the northern; therefore, particular concentration on the southern wave is desirable. The latitudes of particular interest on visual, photometric, polarimetric, and spectrometric grounds are within 30° of the equator and at the edge of the receding polar ice cap. In a study of the Voyager concept of Mars landing missions, Swan and Sagan [1965] have computed the arrival window of Mars-bound space- craft launched along approximately minimum-energy trajectories. Each space vehicle was intended to land two 1700-pound packages on the Martian surface. The accessible latitude ranges shown depend upon the design specifications of the system, its experimental objectives, and the use of the Saturn IB booster. But the position and width of the arrival windows apply approximately to minimum-energy trajectories for any space vehicle system. There are no restrictions in longitude. The dashed arrows are 180 days in length, and represent hypothetical lander life- times. Arrival windows were computed for the entire Metonic cycle from 1969 to 1982. We see that only in 1969 will a minimum-energy trajec- tory lead to arrival at an appropriate time for investigation of the southern hemisphere darkening wave. The 1973 opportunity is favorable for

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Launch Opportunities and Seasonal Activity on Mars 289 SOUTHERN SEASONS SPRING 270 MO 130 SUMMER AUTUMN HELIOCENTRIC LONGITUDE.*) I5O WINTER EARTH OATS NORTHERN SEASONS WINTER SPRING SUMMER AUTUMN HIGH CONTRAST THIRD J •• MEDIUM CONTRAST THIRD WAVE OF DARKENING . LOW CONTRAST THIRD \ .4-tM Figure 4. Martian wave of darkening and Voyager lander footprints. See text for details. observation of the northern wave of darkening; the 1971 and 1975 mini- mum-energy arrival windows may permit some observation of the high- contrast third of the wave of darkening in the southern and the northern hemispheres, respectively; but they are not advantageously placed. The 1977, 1979, and 1982 opportunities are, from this point of view, particu- larly unfavorable. However, payload can always be exchanged for a high-energy Earth- to-Mars transfer orbit. With the Saturn IB, for example, the time of the 1971 arrival window can be changed until it overlaps that of the 1969 minimum-energy window, if we are willing to forego one of the two pro- posed 1700-pound landers. For the 1973 and later opportunities, how- ever, large reductions in payload must be made if the arrival windows are to be timed to allow observation to begin just before the southern or the northern waves of darkening, even if the Saturn IB booster is used. This argues for utilization of the 1969 and 1971 opportunities for biological exploration of Mars. However, if the Saturn V were available for Mars missions by the

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Launch Opportunities and Seasonal Activity on Mars 291 middle 1970's, it appears that a significant application of the extra launch capability could be made; high-energy transfer orbits could be selected, and arrival windows opportunely timed for investigations of the wave of darkening. A preliminary analysis of the Saturn V's capabilities has been made by Dr. Paul R. Swan, of the AVCO Corporation, as a courtesy to this study. His computations disregarded launch azimuth restrictions. Swan finds that generally, 120-day to 150-day maneuverability in helio- centric longitude could be achieved by use of the Saturn V. This means, roughly, that 3000 pounds of payload could be placed on Mars in the 1973 through 1982 opportunities. Launches in 1975, 1977, and 1979 would result in arrival just before or in the midst of the high-contrast third of the northern wave of darkening; the 1982 launch, just before the high-contrast third of the southern wave of darkening and the 1973 launch would allow observation of either wave of darkening. A further study of the Saturn V's capabilities has been made by Dr. F. G. Beuf, of the General Electric Company, as a courtesy to the study. The General Electric study involves a different spacecraft and is more detailed than the AVCO Saturn V study. The overall conclusions are generally in agreement, except in the case of the 1975 opportunity. The General Electric results are contained in Table 2. We see that quite sizable payloads could be landed in the midst of the wave of darkening throughout the 1970's, if a Saturn V booster is used. It is possible that more favorable arrival dates may be obtained by use of a different kind of transfer trajectory in which the spacecraft is launched first towards Venus, where it makes a close fly-by with a hairpin trajec- tory [Seiff, 1964]. This imparts momentum to the spacecraft in a way that could not be achieved with the vehicle's own fuel supply. The con- cept is the same as the "gravitational machine" of Dyson [1963]. The synodic period of Venus is 584 days. Therefore, if such Venus fly-by trajectories can be exploited, a much larger range of launch opportunities and arrival windows becomes available. The practicality of such a trajec- tory, with its high accelerations, long transit times, and range of distances from the Sun, remains to be explored. REFERENCES Dyson, F. J. (1963), In: Interstellar Communication, A. G. W. Cameron, ed., Benjamin, New York. Focas, J. H. (1959), Compt. Rend. 248:924. Seiff, A. (1964), private communication. Swan, P. R., and C. Sagan (1965), /. Spacecraft and Rockets 2:18.