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CHAPTER 17 SPACE VEHICLES FOR PLANETARY MISSIONS ELLIOTT C. LEVINTHAL INTRODUCTION AND SOME DEFINITIONS The launch vehicles contain the booster engine or engines arranged in one or more stages. These provide the thrust necessary to launch the spacecraft containing the payload on the desired trajectory. The launch vehicles associated with the possible unmanned planetary missions of immediate interest all use chemical propellants, as distinct from nuclear propellants, and are known as Earth launch vehicles. Orbit launch ve- hicles are assembled in Earth orbit from payloads placed there by Earth launch vehicles. If several vehicles are required to place this payload in orbit, the procedure would be known as a multi-Earth launch. This prac- tice is encountered in studies of manned interplanetary missions where, because of the limitations in thrust, the vehicles are launched from an initial orbit about the Earth. For a given vehicle system, a spacecraft launched to Mars at certain times during the synodic period requires much less energy than it would at other times. The selection of a time for launch is affected by the de- sired payload, the desired flight time, the three-dimensional configuration of Mars, Earth and the Sun, the particular trajectory desired, including considerations of arrival time during the synodic period, approach velocity of the vehicle relative to the planet, the type of mission to be accomplished 292

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Space Vehicles for Planetary Missions 293 on arrival, and the energy that can be provided by the launch vehicle. In a given synodic period there is a particular interval of time during which the energy required for a given planetary mission is minimized. This is called a launch window or a planetary opportunity. The mechanics of interplanetary flight is a complex subject. To study the feasibility of a particular mission, however, one can often make use of simplifying assumptions. The problem can be reduced to a two-body problem to the extent that the trajectory can be considered to traverse different regions of space, starting from the departure planetocentric space, crossing interplanetary or heliocentric space and arriving at the target planetocentric space. Each of these regions can be described in terms of a central force field, in which the vehicle's motion is essentially that of a comet with constant orbital elements. The transfer orbit is that orbit connecting the departure planet and target planet. To the extent to which orbits are Keplerian (i.e., non-powered) and lie in heliocentric space, their orbital elements are invariant. However, as they enter different regions of space, they are perturbed and elements change. Rocket thrust, either impulsive or continuous, can be used to effect an orbit change. If transfer orbits are tangential to their departure and target orbits, they require very low and, in most cases, the minimum transfer energy. These orbits generally require a comparatively long transfer time and are known as Hohmann transfer orbits. For such an orbit, the Martian opportunities occur about three months before Martian opposition. Since characteristic flight times are about 200 days, spacecraft launched to Mars during the launch window leave Earth before, and arrive at Mars after, opposition. At such a time in the synodic period, the Earth to Mars distance is in- creasing, and this is especially significant in connection with the trans- mission of data back to the Earth. Fast transfer ellipses are transfer orbits that intersect one terminal orbit or both. They provide the means for reducing the Earth to Mars distance at encounter. A fast transfer orbit might reduce this distance by a half, thus providing four times the data transmission for the same transmitter and receiver. The process of transfer from a helicentric orbit to a planetocentric orbit is known as capture. The reverse process is known as escape. For circumplanetary probes that are captured by the central force field of the target planets, fast orbits are particularly expensive in energy. This is because the relative velocity between the spacecraft and the planet at encounter is large. While this is important, no matter what the type of mission may be, it is more important for orbiter missions, for the relative velocity between the spacecraft and the planet has a direct bearing upon the amount of retro-propulsion that must be carried in order to put the spacecraft into orbit about the planet. Changes in the three-dimensional

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294 APPROACHES AND REMOTE OBSERVATIONS configuration of Mars and Earth during the Metonic cycle, can cause the weight of the retro-propulsion system for otherwise similar missions to vary by a factor of three. The approach velocity is also important for landing missions because this changes the velocity at which the capsule would enter the Martian atmosphere and, hence, the amount of heat- shielding material with which it must be equipped. Trajectories are often classified by the angle traversed with respect to the Sun, from Earth launch to Mars encounter. Those sweeping out less than 180° are known as Type 1 trajectories, and those between 180° and 360° are known as Type 2 trajectories. Figures 1 through 4 show typical trajectories from Earth to Mars for the opportunities in 1969, 1971, 1973 and 1975. Launching requirements for planetary missions are often described in terms of the over-all minimum velocity required for surface launching to accomplish the particular mission. This defines an ideal velocity that will give the correct potential and kinetic energy with respect to the Earth's surface for the desired transfer orbit. For example, the circular velocity in a 300 nautical-mile-high Earth orbit .is 24,900 feet per second, and the corresponding ideal veloclty is 27,000 feet per second. The difference of 2100 feet per second is accounted for as potential energy. To achieve a trajectory that would yield hyperbolic passage to Mars, additional energy is required. This is often expressed in terms of a velocity known as the hyperbolic excess velocity or injection velocity. It is the velocity of the vehicle with respect to the Earth when the vehicle-Earth distance is large. Thus, the ideal velocity for a Mars mission is obtained by adding the ideal velocity, corresponding to the initial 300-mile Earth orbit (27,000 feet per second), and the hyperbolic excess velocity of approximately 11,600 feet per second, giving a total of 38,600 feet per LAUNCH ENCOUNTER (280 WYS FROM LAUNCH) ENCOUNTER+3MOS ENCOUNTERV6MOS MARSOIST FROW EARTH IMILLION Km> MARSDIST. FROM SUN IAUI m> Ltt ai LX •LI in mi Lt) Figure 1. Earth to Mars, February 24, 1969

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Space Vehicles for Planetary Missions 295 LAUNCH ENCOUNTER (182 DAYS FROM LAUNCH) ENCOUNTER-HMOS. ENCOUNTER I 6 MOG MARSOIST. FROM EARTH MIU.KM M) MARS DIST FROM SIM jAu HI L« Mi Ml L44 L» K3 L«6 Figure 2. Earth to Mars, June 3, 1971 LAUNCH ENCOUNTER ()66 DATS FROM LAUNCH) ENCOUNTER 13 MOS. ENCOUNTER + 6MOS. MARS 01 ST. FROM EARTH (MILLION Km) MARSDIST. FROM SUN IAUI I03L3 1919 141 Lit •bi Mi 1.U 1.M Figure 3. Earth to Mars, August 7, 1973 LAUNCH ENCOUNTERdeaDATS FROM LAUNCH) ENCOUNTER + JMOS. ENCOUNTER+eMOS. MARSDIST. FROMtABTH IMILLIONKm1 MARS DIST FROM SUN OCR for page 292
296 APPROACHES AND REMOTE OBSERVATIONS second. In Figure 5, the launching requirements for several different planetary missions are shown in terms of ideal velocity. The evidence of extremely low Martian atmospheric densities has im- posed additional constraints on missions that involve landing. Assuming ballistic entry and parachutes for descent, the low densities have a direct effect on mission capability. Figure 6 shows the percentage of entered weight remaining at impact as a function of the Martian surface pressure. Curves for different impact velocities are shown. These curves are meant simply to show typical examples of trends. They indicate a significant reduction in the weight available for instruments, power and communica- tions as the Martian atmospheric pressure is reduced. In designing a spacecraft one must take into account the uncertainties in our knowledge of the atmospheric density. Assuming that the uncertainty is 10-40 milli- bars, the capsule must be designed to cover this total range. If the surface pressure is 10 millibars, a large penalty in capability is not paid in design- ing for 10-40 millibars. If the surface pressure is much above 10 millibars, a large penalty is paid in designing for pressures as low as 10 millibars. Another effect of the low surface pressure is on the physical size of the entry capsule. Figure 7 shows the diameter of the capsule as a function of its weight at entry into the Martian atmosphere. Four curves are shown for different surface pressures. The physical size of the capsule is very large for the 11 millibar atmosphere. This has a direct effect on the selec- tion of the launch vehicle because some rockets have small shrouds, whereas Figure 5. Velocities required for planetary missions and corresponding durations of flight. (Courtesy of the National Aeronautics and Space Administra- tion) g «... JIMt OF fUGHT [DAYS!

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Space Vehicles for Planetary Missions 297 100 160 Worit COM ballistic entry with "free-fall" terminal velocity -^ 2 Stage Parachute System 10 20 60 TO Figure 6. 30 40 50 SURFACE PRESSURE (MB) Effect of surface atmospheric pressure on the fraction of the weight of a Martian lander that could be al- located to instruments, power, communications, etc. Atmospheric retardation is assumed'and the terminal velocities indicated. (Courtesy of the National Aero- nautics and Space Administration) others allow larger diameter spacecraft to be carried. In some cases, it may be necessary to provide the capsule with extensible flaps in order to achieve the effective diameter required by the conditions of atmospheric entry. Figure 7. Effect of Martian atmospheric pres- sure on size and weight of a lander. (Courtesy of the National Aero- nautics and Space Administration) 1000 !000 3000 4000 MOO LANDER WEIGHT AT ENTRY, POUNDS

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298 APPROACHES AND REMOTE OBSERVATIONS LAUNCH VEHICLES Several potentially suitable launch vehicles are available or under devel- opment; they are shown in Figure 8. The Atlas-Agena has a Mars or Venus capability of about 500-600 Ibs. It was used for the Mariner flight to Venus in 1962 and for the 1964 Mars flight (Mariner IV). The approxi- mate cost of this rocket system, including launching costs, is $8 million. (The costs given here are approximate and do not include development costs or the costs of the spacecraft.) The Atlas-Centaur is under develop- ment and is scheduled to be available by 1966; it will have a 1200-1500 Ib planetary transfer capability and will cost about $11 million per launch. The next largest launch vehicle with potentialities in the planetary program Figure 8. United States launch vehicles capable of planetary missions. (Courtesy of the National Aeronautics and Space Administration)

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Space Vehicles for Planetary Missions 299 is the Titan IIIC; it is being developed under contract from the United States Air Force. It could be available for planetary exploration in 1966 at a unit cost for launch of between $12-20 million, depending upon the level of production. The Titan IIIC would be able to inject approxi- mately 3000 Ibs on a Mars transfer trajectory. The largest launch vehicle now under development with planetary transfer capability is the Saturn V. It is being developed under the direction of the Marshall Space Flight Center and would be able to inject spacecraft weighing up to 60,000 Ibs into a transfer trajectory to Mars. The availability of the Saturn V for planetary exploration would depend on the progress of the manned lunar landing program, for which it was designed. It could, perhaps, be avail- able for planetary exploration in 1969. The cost per launch is estimated at approximately $100 million. In addition to these rocket systems, other configurations are being studied. For example, if fluorine were added to the liquid oxygen used in the Atlas-Centaur, it would increase its payload capacity to 2000 Ibs and slightly increase the launch cost. A Titan II with a Centaur upper stage would be able to transport about 2500 Ib to Mars at a cost of $11 million per launch. Extensive studies have been made of the capabilities that would result from the addition of a third stage such as the Centaur, to the Saturn IB. It is proposed to use a rocket system of this kind in the Voyager program; it would be able to transport 6000 to 7000 Ib to Mars at an estimated launch cost of $25 million. The costs for developing the three- stage Saturn IB have been estimated at between $100-300 million. All planetary missions require the use of a rocket system capable of an ideal velocity greater than about 40,000 feet per second; the launch vehicles mentioned above that meet this requirement are listed in Table 1. UNMANNED MARS MISSIONS Even with all the limitations that have been described, the number of different missions that could be flown is very large. This is particularly true when one is considering the launch vehicles with the larger payload capabilities. While generalizations about their characteristics are possible, they must be applied with great caution. Determination of the feasibility of any particular mission requires detailed study. For example, more than 1000 man-months was expended on a study of a few of the very many possible Saturn IB planetary missions. In these studies the scientific pay- load varied from 8-12% of the capsule weight; while this ratio is fairly common, it may not apply in all such cases. For orbiter missions, approxi- mately hah* of the weight injected on a transfer trajectory to the planet

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300 APPROACHES AND REMOTE OBSERVATIONS •5! £ i Q E I 3 K I a as Q g° 88- X o \O 18 So I II « 0\ —-1 ^ Q WJ o 0 00 « o rI 8 a c 0) 00 3 3 c« c« •a 6000- in M t> oo 4 ^ I

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Space Vehicles for Planetary Missions 301 must be reserved for the propulsion required to place the payload in orbit. Thus, the effective payload for orbiters is about one-half of fly-bys on otherwise similar missions. About one-half the payload may be landed if it is first put into orbit, rather than landed directly from a ballistic trajectory. The nature of the search for extraterrestrial life is such that it is not possible to estimate quantitatively the potential yield from any particular mission. One can only approach this problem by considering individual experiments, such as the potential spectral and spatial resolution for a given orbiting infrared spectrometer and, with this information, attempt to assign a value to the inferences that might be drawn from the results. A highly oversimplified criterion of the yield of a mission is the number of bits of information that the scientific payload can transmit back to Earth. The significance of this criterion is again dependent upon the objective of the mission. For example, the payload that a Saturn V could put into orbit about Mars could, for lifetimes in orbit greater than a certain value, transmit more data for the duration of its mission than could the maximum payload of the same launch vehicle if put into a fly-by trajectory past the planet. It might turn out, however, that because of its larger payload, the fly-by mission might take a few pictures of a limited area of the planet's surface at a higher resolution than would be achieved by the orbiter. This is not to suggest that a fly-by is a better mission for video reconnaissance than an orbiter, but simply to emphasize the care that is required when using a single criterion for the yield of a mission. Prof. Bruce Murray of the California Institute of Technology has suggested a trajectory that is intermediate between that of an orbiter and a fly-by: some energy is used at encounter to change the fly-by hyperbolic trajectory to another hyperbola that brings the spacecraft back towards the Earth. This might be called a bi-hyperbolic trajectory. It has some interesting properties with respect to reconnaissance data. For a given risk of collision, the fly-by can approach closer to the surface. This gives a linear gain in resolution for the same payload. The time in orbit, which has a linear relationship to total bits transmitted, must be compared to the inverse R2 dependence of transmitted bandwidth or the bit-rate for a given transmitter power, R being the Earth- to-spacecraft distance during transmission of data. Thus the situation can be reversed and the non-orbiting mission could give more total bits and higher resolution (closer approach). It should be pointed out, however, that the duration of the mission becomes large if much advantage is to be taken of the procedure outlined above. Some of the missions that have been planned or studied for each of the possible launch vehicles are reviewed below.

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302 APPROACHES AND REMOTE OBSERVATIONS ' IM GM JWIENM IlwnD HOUIIOI sau Ptmua VIM Figure 9. The Mariner IV spacecraft, shown here in diagram, was launched with an Atlas Agena rocket on Novem- ber 28, 1964, and completed a successful flight to within 6,000 miles of Mars in July 1965. With solar panels extended, the spacecraft spans about 22V6 ft; its gross weight is 575 Ib, including about 60 Ib of instruments, as follows: television camera, plasma probe, ionization chamber, trapped radiation detec- tor, cosmic-ray telescope, vector magnetometer, and cosmic dust detector. (Courtesy of the Jet Propulsion Laboratory and the National Aeronautics and Space Administration.) Fly-By Missions All the launch vehicles mentioned above have the capability of carrying out fly-by missions. The 1964 Mars Mariner was such a mission utilizing the Atlas-Agena. The payload included a television system designed to obtain about twenty pictures of the Martian surface with 5 km resolution, together with instruments for observing micrometeoroids, charged particle fluxes and the magnetic field. The characteristics of this spacecraft are illustrated in Figure 9. This represents the minimum mission studied and essentially exhausts the capability of the Atlas-Agena. Fly-by missions have generally been considered worthwhile only in con- nection with launch vehicles of moderate capability. Larger rockets may find some application in this fashion if the suggestion of Murray is followed (see above). A spacecraft contemplated for the Saturn IB, with Centaur upper stage, comprises a 2000 Ib "bus" that could serve either as a fly-by or, with the addition of a retro-rocket, as an orbiter. This bus could carry 200 to 300 Ib of scientific instruments with 600 watts of power, a 10 foot diameter microwave antenna and could transmit at a 2000 bits per second rate. With data storage, it could provide a total transmission of 10* bits, if used as a fly-by.

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314 APPROACHES AND REMOTE OBSERVATIONS m >n 0 **! cs VO in o § *j O\ in » CN >/-) PJ <"1 oo ^ . 3 n "S ct /•> ON in fI VO d ri C VO 8 M fN •^ VO in ^ W •o > 1 1 e D £ £ £ £ « « « 3 13 oo 1 M i-l ^ CQ iU 1 *-J H 3 CU -^ — a - & Scientific Payload Lander Gross Weight ^ Base Diameter (DB) ef Scientific P/L Nomin; Communication Nom Antenna Dish Diamel Gross Payload i i § •5 a ft I I

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Space Vehicles for Planetary Missions 315 Figure 10. Scientific payload power versus weight for Saturn V landers woo SCIEKTriCWYLOAO)LBI Figure 11. Scientific payload bit rate versus weight for Saturn V landers 250 500 RANGE - 1.4 AU 2500 5000 I I 200 400 IOOO 4OOO 10000 SCIENTIFIC PAYLOAD WEIGHT (LB) the initial assumptions. For example, assumptions concerning guidance accuracies and method of retardation lead to a maximum in the curve of gross payload weight as a function of lander gross weight, so that if the entry angle were less than 23°, the total payload on Mars with all

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316 APPROACHES AND REMOTE OBSERVATIONS 1000 Figure 12. Communication subsystem data rate versus subsystem power for Saturn V lander Figure 13. Communication subsystem data rate versus subsystem weight for Saturn V lander 200 900 «OO 50° SUBSYSTEM POWER (WATTS) 4- 200 5OO 400 SOO SUBSYSTEM WEIGHT (LB)

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Space Vehicles for Planetary Missions 317 Figure 14. Comparative launch capabilities other assumptions remaining the same could be increased by about a factor of 3. An alternate approach would be to eject the landers from orbit about Mars. Ordinarily this would be considered inefficient due to the energy required to place the payload in orbit. However, entry from orbit would allow a tighter entry corridor with the same guidance accuracies and a significant reduction in the entry velocities. Both these effects allow significantly higher ballistic coefficients and thus a much higher payload weight for a given total lander weight. .This more than doubles the possible scientific payloads landed per mission. Figure 14 shows the comparative capability of the Saturn V launch vehicle as a function of the "vis-viva energy", or C3, the injection energy of the Earth escape hyperbolic trajec- tory. This points up the fact that the use of this launch vehicle opens up a new regime of mission opportunities and problems. The potential of the Saturn V for the biological exploration of Mars can perhaps be exploited even further. According to H. H. Koelle, Director of the Future Projects Office, MSFC, 7200 Ib of the 50,000 Ib a Saturn V could deliver to a Mars transfer trajectory would be required to return a 600 Ib capsule from Mars to Earth on a minimum energy mission of about' 950 days duration. This simply shows that such a mission is possible with the Saturn V and, for that matter, even with the three-stage Saturn IB. A detailed study would be required to establish the technical feasibility of such a mission. Because of the hazards of back contamination, one might want to impose additional requirements on such a study, such as the de- livery of the return capsule to a scientific station on the Moon or to an Earth-orbiting space station. On the Moon, man could study at first hand the interaction of terrestrial and Martian biota without the dangers of con- tamination of either Earth or Mars. Concern with planetary contamination could prevent the use of orbiting laboratories for this purpose. Table 12 gives a comparison of the estimates of cost, weight, data trans- mitted and lifetime for many of the planetary missions discussed above.

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318 APPROACHES AND REMOTE OBSERVATIONS V S N E J s s 8 o 'i! pp in E E E E J % 1 - a 2 Sa| f £ £"£ o *> o^ 8 a fiu 1 ll| Isi S •a .•" -O 0 C "£ C u a 6 o o o o 1 J .g efl C w tf l s-* c B 2^1 <<% o H) ri WO £w^- 0,^« Is li •a i CO r*- Cd ON c 1 I c m o 88 88 5^- r- _ a 5 > C P> . — — 2 ** ca s-^ 5 ^fsi a cu 0.0- ''-/) U- 3 cj -2 g •"« 5 ,/-: C/3 s~' IB o U 5 O O 8 H CS ^ 0 & c« D, « >n ^ iiuoCi 01 u ft 00 D & ATLAS-CENTA With S-6 Stage With Centaur SATURN IB SATURN V pacecraft Orbiter o0 Lander 2-1 .§ ?s 1 J 89 p

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319 COMMENTS The excitement of the search for extraterrestrial life, which extends well beyond the domain of science, need not be reiterated here. The strategies suggested for implementation of this search arise in a context of engineering constraints—mainly those of weight, power and communications. When it was felt that there were opportunities for one- to ten-pound instruments and communication rates of a bit per second in the early 1960's, and that a significant relaxation of these constraints would not occur for approxi- mately a decade, the strategic debate focused on the selection of the "best" "life detector". One argued that growth was a "riskier guess" than metabo- lism and that a single step functional test for a ubiquitous enzyme had more generally than a multi-step metabolic degradation. The morphologists, on either a macroscopic or microscopic scale, could not sharpen their criteria to justify the utilization of the total communication channel required for one picture of even a modest combination of field of view and spatial resolution. This debate has not been fruitless, for it has served as a reminder that the unambiguous recognition and characterization of life requires a diversity of experimental investigations. An examination of the instrumentation developments that have been completed or are under way gives the conviction that a sufficient range of specific experimental techniques can be exploited with an instrument payload of thousands of pounds and perhaps even several hundred. The requirement for thousands of pounds becomes necessary when one makes the essential inclusion of the non-specific technique of visual observation on any scale. Its singular value is its non-specificity. The full benefit of this value is achieved only by the transmission of an enormous amount of data. Ten minutes of microscopic observation represent more than 109 bits of data presentation. It is the interaction of these data with the collec- tive store of scientific intelligence represented by the observers viewing the images at Goldstone that gives power to the specific instruments situated in a laboratory on Mars. A Saturn V lander makes possible video trans- mission at high rates during a good deal of the mission. It is continually necessary to face the task of evaluating whether or not a mission is "possible". What is "possible" is determined by many kinds of constraints in addition to those imposed directly by technological mat- ters. There are financial considerations having to do with total costs, and the phasing of costs with the motion of the planets and fiscal budgets. There are manpower problems and political considerations that concern the scientific community, the Government, Congress, and the country at large.

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320 APPROACHES AND REMOTE OBSERVATIONS To avoid the danger of circular reasoning with regard to what is "pos- sible," it is important that scientists and engineers separate and identify the various considerations that affect their judgments about the possibility or desirability of a mission. It has been generally true so far, that the constraints of size and weight have been so overriding on planetary missions that there have been unre- lenting pressures to push miniaturization to "state-of-the-art" limits and to maximize the number of different scientific objectives a mission could achieve. This has led to a certain comparability in some of the parameters of different scientific payloads, such as cost per pound, development costs and development time per pound, the dependence of cost on bits trans- mitted and dependence of reliability on weight or size. Statements or im- plications that these are generally invariant parameters of space missions and that they cannot be radically changed should be examined quite care- fully. It is possible that with larger launch vehicles, good engineering and scientific management of the opportunities presented by large payloads could lead to more favorable levels of performance, economy and yield than would be expected by extrapolation from past experience with smaller rockets and smaller payloads. Similar precautions are required with regard to "learning" or "success" curves of previous missions or vehicles. It has been suggested that data on failures as a function of experience show that one arrives best at a successful large payload by means of a series of mod- est steps. A more careful examination should be made to be sure that the contrary is not true. In attempting a radical departure in size and capa- bility for a planetary mission the number of failures that must be expected before achieving success may be an insensitive function of the number of launch experiences with different missions of much smaller scale. If this is true, and a large mission is thought to be necessary in any case, the greatest effectiveness and economy may be achieved by starting in the first place with a large mission. These remarks suggest a compromise: the mission strategy is based on the use of a large launch vehicle, spacecraft, power source, and data capability but the initial scientific objectives are kept simple in order to maximize reliability. The mission could then have a large potential for growth in scientific yield. The scientific yield or objective is then the parameter that changes with time or experience, rather than the "basic" mission hardware. Such a procedure could lead to a different and improved dependence of scientific yield on cost. Acknowledgements In addition to the utilization of material from the referenced reports, material was also abstracted from memoranda prepared by Mr. J. W.

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Space Vehicles for Planetary Missions 321 Haughey, NASA representative to the Exobiology Summer Study; the talk to the Exobiology Summer Study, June 17, 1964, by Mr. Donald P. Hearth of NASA Headquarters; and material prepared by Professor Von Eshleman of Stanford University on planetary entry probes. BIBLIOGRAPHY Atlas-Centaur 1. General Electric Re-Entry Systems Department; Mariner B Entry Vehicle, Nov. 26, 1963; submitted to Jet Propulsion Laboratory. 2. Goddard Space Flight Center; Experiments from a Small Probe which Enters the Atmosphere of Mars; NASA Technical Note D-1899, Dec. 1963. 3. Ames Research Center; Use of Entry Vehicle Responses to Obtain Measure- ment of the Structure and Composition of the Mars Atmosphere. 4. General Electric; Mariner B Entry Vehicle—Vol. 1: Technical Study; Nov. 26, 1963. 5. General Electric; Mariner Mars 66 Capsule, Vol. 1: Technical Proposal; Jan. 17, 1964. 6. Lockheed; Proposal for Preliminary Mars Atmospheric Probe Design; Dec. 18, 1963. 7. Goddard Space Flight Center; Measurement of Upper Atmosphere Structure by Means of the Pitot-Static Tube; J. E. Ainsworth; NASA Technical Note D-670, Feb. 1961. 8. Jet Propulsion Laboratory; Mariner Mars 1969 Orbiter Feasibility Study; EDP-250,Nov. 16, 1964. 9. Jet Propulsion Laboratory; Mariner Mars 1969 Lander Feasibility Study; EDP-261,Dec. 28, 1964. Three-Stage Saturn IB 1. Jet Propulsion Laboratory; Study of Mars and Venus Orbiter Missions launched by the Three-Stage Saturn CIB vehicle; EPD-139, Vol. 3, Dec. 31,1963. 2. Research and Advance Development Division; AVCO Corporation, Wilm- ington, Mass.; Voyager Design Studies; AVCO RAD-TR-63-34, Oct. 15. 1963; prepared under contract NASw697. 3. General Electric Missile and Space Division; Voyager Design Study; Docu- ment 63SD801; Oct. 15, 1963; prepared under contract NAS W-696. 4. Lockheed Missiles and Space Co., Sunnyvale, Calif.; Voyager Program Study; LMSC-5-53-63-4; Oct. 4, 1963. Titan IIIC Launch Vehicle 1. General Electric Spacecraft Department; Voyager Spacecraft Systems Study (Phase I - Titan IIIC Launch Vehicle) Document 64-SD933; Aug. 4, 1964; prepared under contract 950847 for Jet Propulsion Laboratory.

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322 APPROACHES AND REMOTE OBSERVATIONS Saturn V 1. General Electric Re-Entry Systems Department; A Study of Advanced Voyager/Beagle for Saturn V; GE-RSD70036; March, 1964. 2. General Electric Space Department; Voyager Spacecraft System Study (Phase II-Saturn V Launch Vehicle) Final Report, Volume 1 Summary. Document No. 64SD4376; Dec. 9, 1964; prepared under contract 950847 for Jet Propulsion Laboratory.

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PART VII MARTIAN LANDINGS: UNMANNED

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