The Space Congress® Proceedings 1968 (5th) The Challenge of the 1970's
Apr 1st, 8:00 AM
Experiment Payloads for Manned Encounter Missions to Mars and Venus
W. B. Thompson Belle omm 3 Inc. Washington, D. C.
J. E. Volonte Belle omm 3 Inc. Washington, D. C.
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Scholarly Commons Citation Thompson, W. B. and Volonte, J. E., "Experiment Payloads for Manned Encounter Missions to Mars and Venus" (1968). The Space Congress® Proceedings. 1. https://commons.erau.edu/space-congress-proceedings/proceedings-1968-5th/session-10/1
This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. EXPERIMENT PAYLOADS FOR MANNED ENCOUNTER MISSIONS TO MARS AND VENUS
W. B. Thompson J. E. Volonte Belle omm 3 Inc. Washington, D. C. Summary Trajectory opportunities have been Mariner flyby probes through possible manned identified for free return manned flyby, Mars landings in the 1980 f s are being or encounter, missions to Mars and Venus. studied. It appears that a planetary Using Saturn V launch vehicle technology program covering that spectrum of mis and assuming the development of a manned sions could achieve many of the scienti planetary spacecraft with two year capa fic, technological and national prestige bility, missions to these planets with objectives associated with one of the experiment payloads of 50,000 Ibs are major goals of our national space program— possible. the exploration of the solar system. Selecting as a design reference Assuming that manned planetary explo mission a triple planet (Venus-Mars-Venus) ration in the late 1970 T s is a possibility, flyby with a 1977 Earth launch date, a this paper postulates a planetary program possible experiment program is outlined concept and illustrates the roles fulfilled which employs unmanned probes to explore by unmanned precursory and manned encounter Mars and Venus during the planetary en missions. In particular, the possible con counter phase. To complement this a pro tribution of the experiments payload of the gram of space science and astronomy latter class of missions to our knowledge experiments is carried out during the of the solar system is described. remaining portion of the mission. 2.0 Program Concept A precursory unmanned program of orbital reconnaissance missions with small 2.1 Exploration Objectives atmospheric and survivable surface impacter probes is assumed for both planets. Based One of the major goals of our national on this the prime objective of the manned space program is the exploration of the encounter mission at Mars is surface sample solar system. General objectives which may return for life detection experiments. be cited in the pursuit of this goal are Samples from three different selected areas the advancement of science and technology, could be recovered during the Mars encoun with implications bearing on our national ter phase of the mission. Pour types of image. probes are considered for Venus. A mete orological balloon probe deploys a distri The scientific objectives, as stated bution of weather balloons to record by the Space Science Board of the National atmospheric data. A companion orbiter Academy of Sciences, (D are summarized as serves as a balloon tracking and data relay follows. station. Also considered are slow descent, non-survivable impacter probes which might A. The origin and evolution of the take TV pictures of the surface from below Earth, sun, and planets the cloud layer and survivable impacting lander probes to investigate surface pro Pertinent questions relate to the perties . source of the material and mechanism of formation of the visible objects of Several en route experiments have been the solar system, the time scale of the identified which take particular advantage major events which have occurred and of the trajectory of the design reference are occurring in the solar system, and mission. These include optical observa the physical processes responsible for tions of Zodiacal light, several known the principal energy release of the sun, asteroids, Mercury, and the moons of Mars. Radio observations of Jupiter and the sun B. The origin and evolution of life made in conjunction with an earth-based station would also be of interest. * Problems include the examination of what constitutes life, the search for 1.0 Introduction recognizable life elsewhere in the solar system, the possibility of living The U.S. program of planetary explora systems based on other than hydrogen- tion through space flight missions is still carbon chemistry, and an examination of in its early stages. Missions ranging in the likely conditions necessary for the technical complexity from today's unmanned origin of primitive life.
10.4-1 C. The dynamic processes that shape Utilization of the existing capa man's terrestrial environment bility for development of unmanned probes, instrumentation, and data One facet of this objective in processing. volves the examination of other bodies Acquisition of engineering design of the solar system which are either input data on the planets and inter quite different from Earth or, if planetary space for application to similar, are at different stages of future systems. geologic evolution, to stimulate in Definition of the requirements on creased understanding of the evolution technology for the exploration of and' present physical state of the the entire solar system. Earth itself. A second facet is the application of our knowledge of the Enchancement of National Prestige terrestrial environment to the explana tion of observed properties of plane Manned planetary encounter mission. tary atmospheres, surfaces, and Mars surface sample return. interiors . Unmanned spacecraft rendezvous with an asteroid. The technological objectives, although Unmanned spacecraft rendezvous with not necessarily independent of the scienti a comet. fic objectives, are focused on stimulating Manned planetary landing. a wide spectrum of scientific and engineer ing disciplines in the nation and providing Objects of Investigation the capability to continue manned and un manned exploration of space. The selection of Mars and Venus as the prime areas of investigation in the early Prestige objectives focus on the poten phases of planetary exploration is a re tial enhancement of national power and posi flection of the fact that they are not only tion which can be accrued by demonstrating highly interesting bodies which will pro technological leadership through being first vide new data bearing on the scientific in important new accomplishments. objectives of solar system exploration, but they are also the near-Earth planets with Particular scientific, technological, higher likelihood of yielding early results. and prestige objectives which may be pursued A listing of the priorities of solar system through a planetary program are summarized exploration, exclusive of the sun and the as follows : Earth, which reflects a consensus within the scientific.community,^ 1 ' is: Scientific 1. Mars 5 Venus Search for extraterrestrial life in 2. Moon the recovered Mars surface sample 3. Major planets and in situ on the Mars surface. 4* Comets and asteroids Mapping and reconnaissance of Mars 5. Mercury and Venus to understand the current 6.. Pluto physical state of the planets and 7. Interplanetary dust their history. Measurement of the atmospheric pro 2 ,2 Flight Opportunities perties of Mars and Venus, especially the dynamics of the Venus atmosphere. Table I outlines possible unmanned fly- Optical observations of Mercury, by- missions to Mars and Venus in terms of light scattered from interplanetary the Earth departure energy requirements, CU, dust, the satellites of Mars, and velocities at which cannot be and the hyperbolic excess selected asteroids planetary encounter, ¥«,. Of particular made from. Earth orbit . fact that the V^ values all of Jupiter and interest is the Radio observations within the relatively low range of the sun which cannot be made from lie Earth orbit* 2.4-5.6 km/sec. of Increase in our understanding The parameters of representative manned Mars and Venus to a level where full Mars and Venus encounter missions are set use can be made of the exploration II, They are all low-energy succeeding genera forth in Table potential of the trajectories of a free-return type, i.e., tion of manned planetary orbiting only minor trajectory corrections are neces and landing missions. sary to achieve Earth entry following the initial planetary injection maneuver in Earth orbit. The missions include single planet exploration, as well as dual and Utilization of the existing techno flights* • logical capability for manned triple planet 'exploration of space,
10.4-2 2.3 Precursory Program Early orbiter probes are desirable to provide reconnaissance data useful in the It is expected that the planetary pro targeting of later probes such as soft gram of the next decade would pave the way landers. Furthermore, as a comparison of for continued achievement of our national Tables 1 and II indicates, the passage space objectives through manned planetary velocities at both Mars and Venus are in orbiting and landing missions. The time general less when the probe is launched scales for carrying out the more advanced directly from Earth, thereby reducing the missions are difficult to forecast, but a propulsion requirements for planetary orbit reasonable planning assumption is that inj ectlon, their accomplishment would be feasible in the 1980 T s. This timing suggests that an Small atmospheric probes in the 50-200 initial manned planetary encounter mission Ib class could be delivered to different be conducted in the late 1970 T s and that regions of the atmospheres of both planets it be preceded by (1) a manned Earth or by being deployed from a parent arbiter bital program aimed specifically at the spacecraft either before or after the or development of a capability for long bital Injection maneuver. Data on the atmo duration flight and (2) an unmanned pro spheric structure provided by these probes gram to provide early data on the planets would be used in the mission planning for through the mechanism of Earth-launched lander probes targeted to specific surface probes. sites and would provide preliminary mete orological data, This assumption is one of several that could be made at this time. Although Impacting lander probe technology for partially based on optimism toward what Venus has already been tested by the might be done, it was chosen to permit the Russians (Venus *J spacecraft). Atmospheric selection of an illustrative manned encoun braking without retropropulsion Is suffi ter mission as the basis for a conceptual cient, and the technology Is available for design of an experiment payload. The sys short term surface missions at the high tem, mission and program details described temperatures which are believed to exist, should be considered as illustrative of Such a probe would probably communicate feasibility rather than as a plan for the through Its parent orbiter. future. JPL has studied short term (^1 day) Achievement of a long duration space survivable impact probes for Mars which craft would undoubtedly require a major new could be delivered by a Mariner flyby start in the manned space flight program. spacecraft . ^ 2 ' Having a gross weight of It is envisioned as having a versatile, 350 Ibs, this probe would land about 13 flexible capability built around a "stand Ibs of scientific instruments and trans ardized" module which can provide at least mit 600 bits of data directly to Earth. two years of life support Independent of Such a probe should also be deliverable external sources. The following precursory from an orbital spacecraft. By instrumen needs of the manned planetary program ting the entry shell of this type of probe, should be provided in the manned Earth or the function of the atmospheric probe could bital program: be performed on the same mission. This is the mode assumed in Figure 1. Determination of crew capability Soft lander precursory probes have Maintenance and repair been studied for Mars, but without satis Experiment operation and data factory solution. Candidate concepts are analysis the Saturn V Voyager and the Titan III/ Centaur lander. The latter concept Is Development of technology adopted for the purposes of Figure 1, with future study required to determine the Long-duration subsystems exact nature of the lander. An appropri Experiment subsystems, particularly ate resolution of this matter may be that a large optical telescope the Titan Ill/Centaur probe Include a second generation Mars orbiter spacecraft Development of operational techniques (the first generation being, perhaps, a Mariner derivative) and a large survivable Orbital assembly and launch impact probe, with soft landing probes Experiments probe deployment and left to the manned program, control. Soft lander probes for Venus were not Figure 1 outlines a possible program considered because it was felt that there which would use existing capabilities for Is insufficient data on the surface environ unmanned exploration and provide a sound ment to support the design of such probes* scientific base on which to proceed with the manned encounter mission. The four A parallel program of flyby missions- to basic types of probes considered are or- c ome t s an d as t er o i d s I n t h e 1 at e 1 9 7 0 * s wo u Id biters, atmospheric probes, survivable capitalize on Mariner spacecraft technology surface impacters, and soft landers. and require only the At las/Centaur launch
10,4-3 vehicle. Another program of flyby missions There are also non-ballistic flights with to Jupiter and the major planets in the lower initial velocity requirements. The middle and late 1970 f s would serve as the 1977 flight of this class which has been focal point of new technology in the un studied requires an impulsive maneuver of manned program which would carry over into approximately 0.2 km/sec at the, first the 1980*3 when manned systems are expected Venus encounter. to provide the major means of exploring Mars and Venus. 3.0 Space Vehicle Configuration and Mission Profile 2.4 Mission Selection 3-1 Space Vehicle Configuration Evaluation of the relative merits of the manned missions outlined in Table II To provide a basis for sizing the ex reveals that the 1977 triple planet encoun periment payload to be carried on the de ter is the most attractive for purposes of sign reference mission, an estimate of the this analysis. In addition to being suffi system capability necessary to meet perfor ciently late to permit a suitable unmanned mance requirements within the constraints precursory program and to afford a reason of the mission has been made, using the able period for development and test of the results of analyses conducted within NASA long duration spacecraft, that mission and Bellcomm and by other contractors. appears to win most of the trade offs in Studies have shown that a crew size of be system requirements. tween four and six will be needed to con duct this class of mission. Exclusive of Discounting the one year 1975 Venus experiments payload, the spacecraft weight encounter, the 1977 mission competes favor for a four-man crew would be in the range ably with regard to mission duration. In of 145-150,000 pounds, and increasing the particular, It is superior to the 1978 and crew to six would entail a weight penalty 1981 triple planet flights, each of which of 20-25?. The use of Saturn V technology requires in excess of two years. This reflecting improvements possible in the mission also has the lowest Earth entry next decade should permit injection of a velocity of all the missions. It should total payload weight of 240,000 pounds into be noted, though, that all the Earth entry the Interplanetary trajectory. velocities shown are in a range (12.0-14.9 km/sec) that can be accommodated by Apollo 3.2 Mission Profile techniques and technology. The space environment through which the flight would Spacecraft injection takes place late pass Is also favorable since the spacecraft in January, 1977. The first Venus encoun would not enter the asteroid belt and would ter occurs on the l49th day of the mission, not come closer to the sun than Venus. the Mars encounter on the 345th day, and the second Venus encounter on jthe 574th While the encounter geometry and para day. Earth entry takes place in January, meters are short of being ideal, they are 1979 , just short of two years after Earth no worse than, and In some cases better departure. The geometry of tljie spacecraft than, those of the other triple planet solar orbit is shown in Pigur^ 2. flights. "The combination of encounter ve locities appears to be the best; periapsis The mission is a single-Impulse flight velocities of approximately 12.0 km/sec at which makes hyperbolic encounters with each Venus are comparable to those of other planet and provides a free return to Earth missions, and the low velocity (5.6 km/sec) for entry at a velocity of 12.0 km/sec. at Mars not only affords greater time in Midcourse corrections are made as necessary the vicinity of the planet, but also re during the interplanetary phases. duces the &V requirements on experiments probe delivery and retrieval systems. The The two Venus encounters have a number periapsis altitudes are not as low as the of similar features. The angles of inclin 300-500 km range of the single planet ation to the Venus orbit plane are 80,4° missions; however, they are generally more and 80.5°, respectively, and both flights suitable to reconnaissance of the planet pass over the south polar region. On the from, the flyby vehicle than those of the first encounter periapsis occurs on the other triple planet flights. Also, during sunlit side of the planet at an altitude of each planetary approach better than half 680 km, and for the second, on the dark side of the planetary area projected on the at 700 km. The Mars encounter has a periap "plane-of-the-sky" is illuminated by the sis altitude of 3960 km on the dark side of * the planet, with the flyby plane inclined 29,7° to the Mars equatorial plane. The principal drawback of the mission is its high Earth departure velocity. In 4,0 E xperiment Program for a Tripie P1anet the family of 1977 free return opportuni Encounter Mission ties there are other missions with signi ficantly lower injection energy require 4 .1 Mi s s I o n Objective _s ments. These, however, generally require orbital plane changes en route to the The experiment program for the triple planets and have Earth entry velocities planet encounter mission can be divided that outside the range cited above, into three distinct categories, namely,
10,4-4 experiments performed at Mars, experiments The immediate bio-analysis of part of performed at Venus, and en route experiments. a sample of Mars In a laboratory on board Each of these categories has a specific set the manned spacecraft by a well trained of objectives. analyst, coupled with a more detailed analysis of the remainder of the sample in Mars earth-based laboratories, appears to be the . most promising approach to the search for The available experimental data on Mars life on Mars. There is no guarantee that suggests an environment which would be rela this approach would lead to a successful tively hospitable to soft landing spacecraft. identification of life - or proof of its The atmosphere is thin compared to the absence. However, this approach does have Earth 1 s, containing a small amount of haze the advantage of permitting scientists to which does not seem to preclude good photo exhaust their Imaginations and reflect what graphy of the surface. It is capable of they learn in the definition of additional providing a substantial amount of aerobrak- experiments to perform on the samples. On ing, thereby decreasing the total thrust the other hand, the approach of placing a required from a descent propulsion system. package of experiments on the surface of Our current understanding of this atmo Mars by an unmanned landing probe would sphere appears sufficient to successfully only produce definitive results if a posi design entry shells and retropropulsion tive identification and characterization systems for unmanned spacecraft, although occurred. additional data from atmospheric entry probes would lead to a higher confidence in Forward contamination of Mars is an these designs. obvious problem if the search for life is pursued. While adequate sterilization can The Mars surface has been photographed be approached for an unmanned probe, it to a resolution of several kilometers and can never be a guarantee for a manned sys appears "moon-like. Tt In the absence of tem. The risk of an accident on the sur strong physical evidence to the contrary we face which would expose an unsterilized can assume that this surface is landable in astronaut could bring the search for life the same sense that the lunar surface Is to a rapid conclusion. It therefore seems landable to a Surveyor or Apollo spacecraft. reasonable to try to solve the question of Here again, higher resolution imagery of life on Mars before attempting a manned the surface Is desirable before attempting landing. Sample return, we feel, offers to land any probes. the best chance for meeting this goal. One objective of the manned encounter There is also the related question of program should be to provide planning data back-contamination of the Earth. In the for the next major step, a manned Mars land flyby sample return mode to be discussed ing. The function of this data should first .in Section 4.3, a sterilized probe retrieves be to confirm that manned landing Is a worth the sample, which is subsequently trans while goal, and then to provide environmental ferred behind a biologic barrier for fur data which will affect the descent, surface, ther handling by the astronauts. Before and ascent strategy and hardware design of this spacecraft reenters the Earth's atmo the landing mission. sphere, it must be determined by analysis that there is no risk due to known patho A second objective is to extend the - genic organisms. Furthermore, if the ques scientific investigation of Mars beyond the tion of pathogenic organisms were to exist point reached in the precursory program. at the time of a proposed manned landing, Consistent with the plan outlined in Section the sum total of the constraints this might 2.3 , it is assumed that the manned encounter place on the surface systems and operational mission will be preceded by a program of or plan could be sufficient to compromise the bital reconnaissance, In situ measurements landing mission. of the atmosphere, and measurements at the planet surface. Venus A common focus for both of these objec The physical picture of Venus Is quite tives is the return of a sample from the d1ffe rent from that of Mars, • The recent surface of Mars. The search for extrater Russian probe has confirmed earlier -specu restrial life is one of the most challenging lation as to the high surface pressure, scientific problems of the planetary program, revealing It to be some 15-22 Earth atmo and it seems that sample return should be spheres. This probe recorded the surface the next step after accomplishing a soft temperature on the dark side to be approxi landing by unmanned probes. The difficulty mately 280°C, and temperatures on the sunlit in defining a finite set of experiments to side are probably higher.. However, except detect life and to adequately characterize for the surface temperature, which has also it, once having been detected, suggests been estimated- from the radio emission sample return as a reasonable course of brightness, and an average surface dielec action. tric constant which has been measured by
10,4-5 earth-based radar, physical properties of A Ranger-type TV system could provide this the surface are largely unknown. Aspects data, with several of these probes being of the surface environment of interest to deployed to different parts of the sunlit the spacecraft designer are whether it is side of the planet on a single flyby. solid or liquid, the topography, and the nature of the near surface winds. The And thirdly, it seems worthwhile to Russian Venus 4 probe apparently did not entend the exploration of the surface, survive the landing, and one can only albeit with short-term probes, to examine speculate as to the nature of the surface the range of surface properties at places which caused its demise. such as the sub- and anti-solar points and the poles. The visible picture of Venus as seen from Earth is that of a planet covered by There are at least two possible probe clouds. Photographs of the Earth from concepts which should be considered for a space lead us to speculate that the Venus follow-on program. One is a buoyant labor cloud cover may be intermittent with atory station designed to float in the regions through which the surface might atmosphere and make detailed composition be viewed. Unlike Mars, however, Mariner and thermodynamic measurements, observe the photography has not been attempted for surface, and possibly even search for life Venus, so its visible appearance must be in the more temperature regions of the atmo based on rather crude evidence. sphere. Such a probe would probably have the capacity for a several hundred pound A gross characterization of Venus, science payload, at the same time providing then, is that it is a cloud-covered planet a controlled environment in which the ex with a dense atmosphere and a hot surface. periments could be operated. The weather Also, it is roughly the size of the Earth balloon experiment is a necessary precursor with a Venus day being equal to about 120 to this probe in gathering data on wind Earth days. This sort of picture is not circulation patterns, atmospheric turbulence, readily conducive to thoughts of a manned and temperature and pressure profiles. Venus landing. Instead, the reaction is one of probing the environment more A seco
10,4-6 Meteorological Balloon Probe System The accelerometers will measure the gross effects of wind gusts and turbulence The primary objectives of the meteoro on the balloon motion. To reduce power logical balloon system are: (1) to deter dissipation and the associated thermal mine the gross atmospheric circulation control requirement the acceleration data patterns by tracking balloons from an over would be collected for three minutes in flying orbiter, and (2) to gain a general every fifteen. During these periods a understanding of existing atmospheric con record would be made of the number of oc ditions from the balloon sensor payloads. casions upon which the vector sum of the three accelerations exceeded a preset The balloons are delivered to the threshold, and also of the integrated atmosphere by two separate probes 3 each length of time for which this occurred. containing six balloons 3 to be deployed at The complete record would, in addition, altitudes of 45, 40, 30, 25, 10 and 5 km. contain the maximum acceleration experi A shallow entry path angle of about 12 de enced in each of the three directions dur grees is selected as a reasonable compromise ing the data collection period, It is between the total heat load, maximum heating expected that knowledge gained from the rate., and maximum deceleration experienced analysis of the acceleration data will be by the probe. applied to the design of larger and more complex buoyant Venus probes, The two target areas for the first pas sage are shown in Figure 5- These areas Suitable temperature measurements can nominally lie near the terminator about be made with ceramic thermistor devices in. 12 degrees from the limb, separated by-156 the form of beads and thin wires.. As each degrees of central angle. Since for track balloon floats in the atmosphere at a known ing purposes the targets must lie close to constant density, measurement of tempera the orbital plane, a near polar orbit re ture allows the ambient pressure to be quirement for the tracking orbiter is calculated. However, since the redundancy established. Although somewhat arbitrary of a pressure measurement requires only a at this time, the targeting strategy small weight penalty, it is considered adopted allows the acquisition of atmospher worthwhile to make this measurement as a ic information from two widely separated means of checking the temperature record. regions of the planet. Precursory data Within the range of pressures that will be from the reconnaissance orbiters and atmo encountered up to an altitude of 4 5 tan the spheric probes should provide better guide performance of conventional aneroid gauges lines on where to target the meteorological will be quite adequate. Simultaneous pres balloon probes. sure and temperature measurements would be taken every 15 minutes. The six altitudes at which balloons are deployed were chosen to provide a means The sferics detector carried by each of investigating the low, intermediate and balloon will enable observations to foe high altitude domains of the atmosphere. made of the general electrical activity of Deployment at altitudes' of 5 and 10 kin is the atmosphere. The detector, consisting suggested to cover the near surface region of a, small whip antenna and a broad-band of the atmosphere. Specification of 25 radio receiver, will operate continuousIf and 30 km intermediate altitudes covers the and will count' the number of discrete elec domain in which a buoyant laboratory station trical discharges which produce sufficient might be deployed. These balloons would electromagnetic radiation for detection. probably operate inside the clouds. The two From the analysis of this data, a general high altitude balloons at 40 and 45 km are picture of the electrical activity of the expected to yield data on the conditions Venus atmosphere will be obtained. It above the clouds, probably in the region of should be possible to detect any correla the tropopause. Earth-based observations tions between, such, activity and position on indicate that very high wind velocities the planet. parallel to the equator may be encountered at these altitudes. The balloons were de Direct spectroscopic observations have signed using a model atmosphere based on shown that water vapor exists above the data from both the Soviet Venus 4 mission clouds covering Venus, and it is possible and brightness temperature data obtained that the clouds themselves are partly com from earth-based observations. Nominal posed of ice crystals . 1t 1. s 9 t here fore , balloon lifetime is one month. considered desirable to equip the balloons with humidity de t ec tor s* Such a detector» Figures 6 and 7 indicate the balloon consisting of an electrically conducting probe conceptual design, while Figures 8 chemical film, registers changes in humid- and 9 show aspects of the mission profile* *ity by its change In resistivity at 15 ml n ut e i n t e r va Is » The payload subsystem carried by each of the meteorological balloons will include The balloon tracking relay a pressure gauge, three mutually orthogonal orbiter is deployed several days before accelerometers, a sferics detector, and a f ly b y ve hi e 1 e pe r i ap s is, Irame di at e ly aft e T thermistor for ambient temperature measure separation from the manned vehicle the or ment. The pressure and temperature gauges biter performs a. preprogrammed velocity will be designed to operate within ranges correction to adjust its perlapsis passage appropriate to the altitude of deployment, altitude to **QQO km,, At periapsis it Is In addition, each balloon will be equipped deboosted via a two stage retro system with a humidity gauge, 10,4-7 into a circular orbit with about a 3 hour period. Figure 10 shows the orbiter space visibility conditions would influence the craft design concept. experiment selection and mission plan for the floating laboratory station, which The balloon tracking and data communi might be a second generation atmospheric cations system employs a square array probe. antenna with a 27° half angle. The orbital tracking geometry-is shown in Figure 11. In addition to TV this probe carries The orbiter measures the angle and range to a complement of atmospheric experiments each balloon within its field of view on t'o measure temperature, pressure, composi every orbit. It also interrogates each tion, light level and altitude. All ex balloon for the atmospheric data It has In periments cease to function when the probe on-board storage. The position of the impacts the surface. Prior knowledge of orbiter relative to the planet Is estab probe descent times gained from the precur lished by tracking from Earth. The latitude sory program will allow accurate timing so and longitude of each balloon can be deter that the photographic readout from the mined to an uncertainty of about 100 km, sinker probe occurs when the flyby vehicle This is not a serious error since the objec is In the optimal range position. tive is to determine atmospheric circulation patterns on a planetary scale. Balloon Pictures are acquired every 10 seconds altitude will be calculated by comparing the from an altitude of about 30 km. Initial on-board pressure and temperature measure pictures have a resolution of approximately ments with atmospheric profile data collec 120 meters and a coverage of 500 km 2 . A ted by the photo sinker, lander, and maximum transmission range of 60,000 km Is atmospheric probes. adequate for this experiment. Each balloon is programmed to transmit The design concept for this probe Is some 800 bits per orbit when interrogated, shown In Figure 12 and its mission profile Consequently, the orbiter can accumulate as In Figure 13. Probe power is supplied by many as 9,600 bits per orbit. This data Is' batteries, and thermal control is by insu relayed to Earth through an omnidirectional lation with an internal coolant such as command data link at a 5 bit/second rate, ice . requiring 120 watts of Input power. The solar array power supply was sized for 265 This probe can be targeted for atmo watts continuous. Attitude control consists spheric entry angles ranging from vertical of horizon scanners and a cold gas system incidence to 25° off vertical. This latter to keep the tracking antenna pointed along constraint is the half angle of the commu the gravity gradient, nications antenna which is aligned with the spacecraft spin axis. The only other tar The propulsion system was sized for a geting requirement Is Impact on the sunlit total AV of 6.25 km/sec (approximately 0,25 side. Precursory data Is expected to con km/sec to change the periapsis altitude from tribute to the targeting strategy. that of the nominal flyby path; and for mid- course maneuvers, with the remaining 6 kin/sec Yerms Lander Probe for circularlzation and plane changes). The total AV Is divided equally between two The lander probe Is designed to Impact stages, with- stage 1 being used both for the Venus surface and survive for one hour. periapsis altitude adjustment and deboost. During this time It collects data on the Table III summarizes the subsystem weight surface environment,, including a panoramic breakdown for the orbiter probe. television scan, and transmits the data directly to the manned flyby vehicle. The Photo^Sinker^robje objective of this probe is to find out whether surface operations on Venus are This probe Is essentially an atmospheric feasible and to indicate which areas are drag probe instrumented to take TV pictures more suitable for landing. This requires- below the Venus cloud layer. If the pre the use of several lander probes. The cursory program establishes that the surface large payload capacity of the manned vehi is not visible from orbit, but there Is cle makes it possible to investigate eight sufficient light below the clouds for surface locations on a single mission. illumination (photometer measurement), this type of probe might provide the only large Figures 14 and 15 Illustrate the probe area coverage of the surface in the visible mission profile and conceptual design, re range* The utility of the encounter mission spectively. The lander probe Is designed Is that it can deliver several'probes on a for entry at angles ranging from, vertical single passage to provide a good statistical incidence to about 78° off vertical, which sample. If the surface visibility is not is the skip out angle. It can also take uniformly good for photography,, multiple TV pictures in the dark 3 as It carries its probes will provide a better chance of loca own Illumination system, ting a single good area* If the visibility is uniformly good over the sunlit side of As this probe Is the subject of a the planet, multiple probes will provide a companion paper(3) .for this conference,,, it statistical sample of terrain over different will not be discussed further here* regions of the planet. Knowledge of the
10.4-8 On Board Experiments used to achieve low velocity landing through the use of range and Doppler radar and an A variety of remote sensor experiments autopilot system similar to that of Surveyor., might be usefully carried out on board the After the velocity has been reduced to about manned vehicle during Venus encounter. In 1000 fps the aeroshell is separated from the particular, a one meter diffraction-limited landing vehicle and the landing legs are telescope which is proposed for en route extended. Touchdown occurs at a vertical astronomy (see Section 4.5) could be used velocity of about 5 fps. for photography and spectral measurements of Venus during the approach and periapsis The landing points of the three MSSR passage phases. Some of the most compre probes were chosen so that they would all hensive imagery of the surface might be be rotated Into the flyby plane for launch obtained from a radar imagery experiment about 6 minutes prior to manned vehicle conducted at periapsis passage. periapsis passage, as shown in Figure 20. The variation in arrival time allows flex 4.3 Mars Passage ibility in the probe landing location. Figure 21 shows the early arrival times The geometry of the Mars passage is required to properly position the MSSR In shown in Figures 16 and 17. Only one type the flyby plane at launch as a function of of probe for Investigating Mars is included landing latitude. in the probe complement of this mission. The Mars Surface Sample Recovery (MSSR) Shortly after landing the probes be probe collects samples of the Martian sur gin programmed pre-launch surface opera face material and transports them via a tions * The equipment doors are unfolded rendezvous rocket to the manned flyby vehi- as shown In Figure 2:2, and the high gain cle 3 returns full color photographs taken antenna is deployed. Sample collection is on the surface , and emplaces long term ex begun using a roek drill to collect sub periments on the surface. Three MSSR surface material and an aerosol filter to probes are used to increase the likelihood collect samples of material suspended in of successful sample return and hopefully the atmosphere. As soon as e ommun1cat I ons to return samples from different areas. are established with the manned vehicle, a panoramic television picture is transmitted Figure 18 shows the significant events to enable the astronauts In the flyby vehi in the MSSR mission profile. Prior to de cle to select the most interesting areas ployment /the three MSSR probes are encased for surface sample collection, Since it is in individual sterilization canisters and necessary to collect surface samples at stored in the probe compartment of the flyby least 100 feet from, the landed MSSR, probe vehicle. Deployment occurs about 5 days to minimize contamination by rocket exhaust before flyby vehicle periapsis passage (M-5 ga s e s, th e s urfac e samp1e acquls111on de- days) and consists of ejecting-the probes vices are propelled, radially outward by from the flyby vehicle and separating the mortars remotely aimed in the favorable sterilization canisters in such a way as to di re c 11. ons . Drag buck e t s or va c uum c le aner prevent contamination of the probes. t y p e s amp le r s , or a c omb in at I, on o f the s e , are used to collect samples of the surface The injection propulsion maneuver, material. The entire sample acquisition which is started as soon as deployment is procedure and other interactions of the complete, causes the probes to intercept probe with the surface, such as footpad the planet from 2 to 4 hours prior to flyby penetration and rocket crater ing., are vehicle periapsis passage. Following two photographed in color. The film is placed, midcourse maneuvers and jettisoning of the ' in the rendezvous rocket along with the rocket , the probes enter the atmosphere at surface samples and the probe is readied a height of 220 km and an entry angle of for launch, . 19° below the local horizontal. These entry conditions constrain the touchdown On. command from the manned vehicle the points to a locus of about 11° behind the antennas and cameras on the MSSR are then limb of Mars as viewed from the approaching; retract e d. The c on I. c a 1 s t rue t ural she 11s manned vehicle. Figure 19 shows the ap which surround and support the rendezvous- proximate landing geometry. Since the rocket during the entry and landing are planet rotates approximately 14° per hour,, folded outward, over the equipment doors. it takes almost an hour after landing for This frees the rocket for launch and pro the MSSR probes to establish line-of-sight tects the equipment from damage by the communication's with the manned vehicle, rocket exhaust. On further commands from the manned vehicle the three rendezvous A 60° half-angle sphere-cone shaped rockets are launched, nearly simultaneously* aeroshell equipped with an ablative heat Inertial guidance is used to control the shield protects the probe during entry, .rockets as they accelerate-to the speed of The relatively low ballistic coefficient of the flyby vehicle. The terminal maneuvers about 0.7 slugs/ft causes the velocity to are made using the small attitude control be reduced to about 2000 fps at an altitude rocket s on the rendezvous vehIcle* Rendez— of 20,000 feet. At this point, landing vous and docking accomplished using rockets are ignited to slow the vehicle command guidance under optical and radar further. obs e r va t 1 on from the manned vehi cle* Dock- ing and transfer of the payload from the Gimbaled, variable thrust landing three rendezvous rockets to the flyby vehi rockets and attitude control rockets are cle are done in a manner which prevents 10,4-9 contamination of the payload and back- Zodiacal Light contamination of the flyby vehicle. Zodiacal light is the faint scatter Post-launch surface operations begin ing of sunlight from interplanetary dust soon after launch of the rendezvous rockets and gas in the ecliptic plane back to and continue for the life of the probe - Earth. The distribution of this dust in nominally two years. A 180-pound geophysics space cannot be resolved solely with earth- laboratory containing nine experiments in based measurements. Of particular concern cluding a television camera is included in is the question of whether this dust is the landed payload. An additional 100 uniformly distributed around the sun, or pounds of landed payload for exobiology ex whether there is a local concentration in periments or more geophysics is provided in the vicinity of the Earth, giving rise to the weight estimate shown in Table IV, the enhanced scatter from the anti-solar These experiments are supplied power from direction known as gegenschein (see Figure the MSSR power subsystem and transmit data 25). Observations from the manned vehicle over the probe communication subsystem, en route to and from the planets could pro When the increasing distance to the depart vide useful data in determining the distri ing flyby vehicle reduces the transmission bution of this dust. Using a small hand rate below that of a direct Mars-Earth link* held camera with a wide field of view (at the MSSR antenna acquires the Earth and least 10°), a record of the changing inten communicates directly for the remainder of sity, and possibly polarization, of its mission. Zodiacal light in various directions as the spacecraft moves away from Earth should So far as on-board experiments are provide key data for the analysis of this concerned, optical photography and spectro- problem. Simultaneous measurements of scopy using the one meter telescope may be light intensity from Earth orbit would be acquired during the approach and departure required to compensate for any possible phases of the mission, but infrared and time variations in intensity. radar imagery would have to suffice at perlapsis passage , since this is a darkside Asteroid Observations flyby of Mars . Many hundreds of asteroids with sizes 4 . 4 Sec pnd Ve n us P as s age estimated to range from a fraction of a kilometer to several hundred kilometers are The geometry of the second Venus pas in orbit about the sun 3 mostly between the sage is shown in Figures 23 and 24. The orbits of Mars and Jupiter. Little is objective is to repeat the type of experi known of the physical properties of these ment program carried out on the first pas objects. During the 1977 flyby mission the sage in the different planet geometry. manned spacecraft comes reasonably close to Figure 5 illustrates possible target areas four of these asteroids, as shown below.* for the meteorological balloon probes. A large portion of the data collected on the . Asteroid Encounter Distance (AU) Date first passage will have been analyzed to provide clearer direction to the experiment Icarus ,048 5-11-77 and .probe mission planning. ' In this stra-, Aethra .38? 12-5-77 tegy the second .generation class of surface Icarus ,670 8-5-78 or atmospheric probes could be carried on Prisma .532 4-14-78 some later mission. Alinda ,124 4-25-78 By the time of this passage all of the Using a one meter diffraction-limited tele earlier balloon probes would have exceeded scope, the size of asteroids with diameters their design lifetime. To insure a success greater than 160*R kilometers, where R is ful mission for the new meteorological the distance to the asteroid In AU, could balloon probes, a separate tracking and re be measured. Knowing the diameter, the lay orbiter would be provided for the second asteroid albedo could be determined. Only passage experiment program. However, if the four asteroid albedos are known to date, first orbiter were found to be still opera making it difficult to determine whether tional,, the second could be used to provide there are several classes of asteroids additional spatial coverage for the new differing In composition and structure. balloon probes. The location of this orbit (still circular at 4000 km altitude) would Another useful observation which could be selected after examination of the first be made during the encounter mission would passage balloon tracking data, be a measure of asteroid reflectance and polarization as a function of phase angle* 4, 5 E For most asteroids, phase angle data ob tained from Earth Is limited to a maximum A number of particularly rewarding ex of 27°; this could be increased to 40° or periments which take advantage of the orbi 50° by observing from the flyby spacecraft tal geometry of the flyby mission are near its aphelion position. suggested. In general these experiments cannot be done equally well in Earth orbit, although in some cases similar measure ments made simultaneously in Earth orbit *Data Is courtesy of Dr. D. P* Bender of would enhance their total value, North American-Rockwell Inc. 10,4-10 Stellar and Galactic Astronomy greater than the point 90° away on the equator, this difference having been main A large aperture diffraction-limited tained for perhaps a considerable fraction telescope may be used to advantage during of the planet's history,. This may have the flyby mission to obtain long exposure led to permanent changes In the surface photographs to search for faint stars and properties, as a function of longitude, galaxies. Using a one meter diffraction- which may be observed optically or in the limited telescope able to point within infrared* about 0.04 arc seconds of a given direc tion for exposure periods of 10 to 40 Unfortunately Mercury is a very dif hours 3 an improvement in the detection of ficult planet to observe from Earth, As faint sources of several stellar magni it never gets more than 28° from the sun, tudes should be possible compared with the optical observations from the Earth's present detection limit of the 200 inch surface are handicapped either by daylight Mt. Palomar reflector. By proper selec viewing for small zenith angles, or by tion of the orbit inclination, comparable extreme atmospheric distortion for night- exposure periods could be obtained for time viewing. (At times when the sun is selected regions of the sky from Earth below the horizon and Mercury is still in orbital spacecraft. However, in this mode sight, it must be viewed through the of operation thermal problems may be in equivalent of several Earth atmospheres.) troduced by the passage from sunlight to darkness on each orbit. There could also Observations from Earth orbit would be an advantage to the flyby mode if it is provide a considerable improvement, as determined that the Zodiacal light back there is no atmospheric distortion. Arbi ground is reduced by moving the telescope trarily assuming that the telescope line- away from the Earth. of sight is kept at least 10° from the Earth-sun line (to minimize the scatter of A sky survey should be made to count sunlight down the telescope barrel), the faint stars in our own galaxy and the minimum distance to Mercury is about 0.5** average number of galaxies as a function AU, although at this time the phase angle of distance (determined by the red shift). (sun-Mercury-observer) is nearly 160°. Once the sky survey is recorded on film, For a phase angle of 90° the minimum dis the sources of interest would then be tance Is about 0,87 AU. Figures 26 and. 27 calibrated photoelectri.cally by counting Illustrate the observation geometry of photons from the source region and com Mercury from the flyby spacecraft during paring with the neighboring sky back the two passages of Yenus on the 1977 en ground. counter mission, and Table ¥ tabulates the pertinent data. . So that this data may be A possible extension of the faint comp.ared with the Earth, orbital observa source detection experiment is the measure tions, an optical res o1u11on fIgure of ment of the spectra of these sources. One me r 11 Is indicated'in the far r 1 gfa t c o lurah „ advantage in making these measurements in With the same telescope In Earth orbit, space is the absence of an airglow spectrum this figure of merit would be 1.58 when of the night sky which provides an effective Mercury is closest to Earth, with a minimum noise background at Earth. Spectral data value of .about 0,87. In other words,' there could be used to correlate velocities of are periods on the flyby mission when, the distant galaxies (red shift measurement) best linear resolution obtainable at the • with apparent magnitudes (as a measure of surface of 'Mercury would be approximately distance) to further test Hubble T s law. a factor of two better than could be Also stellar spectra in the near "infrared achieved from Earth orbit, i.e,, a figure and ultraviolet are of interest in determin of merit as low as G.W, ing stellar structure and composition. Infrared observations would also be Another advantage of the flyby-based valuable, and again approximately a factor telescope is the possibility of uninterrup of two improvement in linear resolution ted observation of variable sources. This could be achieved.' In particular, on J.D. is especially true for quasar intensity 2443276.5 Mercury Is at a minimum range of fluctuations, where several days of unin 0.3 AU with nearly the full disk in dark- •' terrupted observation at several wavelengths ness. 'The comparable range from Earth may shed new light on the structure of these orbit Is 0.5-4 AU. obj ects. The Natural Satellites of Mars Ob s ervat ions __of Mercury , The diameters of Phobos and Delinos As it is the closest planet to the sun, have been estimated at 19 and 10 kilome— information on Mercury is important to our t e r s , re s p e c 11 ve ly , b as e d on a me as ure d understanding of the solar system. In addi b ri gh t nes s and an ass nined aIbedo ^ B In ce tion, the unusual relationship between its they cannot be resolved optically from period of revolution and its spin period Earth* Phobos is the only observed leads to a substantial amount of uneven satellite in the solar system whose orbi heating of the surface. It has been esti tal period is shorter than, the spin, period mated that the total energy absorbed at the of its primary:,, Its large angular momen perihelion subsolar point is 2.5 times tum suggests that it is a captured
10.4*11 asteroid. High resolution photographs of program. Consistent with the scientific Phobos could be compared with similar photos objectives and probe deployment strategy of known asteroids to search for similari outlined earlier, the following prelimi ties . nary payload selection is made. The geometry of the nominal encounter First Venus Passage Weight during the 1977 flyby mission is shown in Figure 28. In this configuration both satel (2) Meteorological Balloon Probes 3280 Ibs lites would be seen from the dark side. (6 Balloons each) However, by delaying the encounter time by (1) Balloon Tracking and Data approximately one hour, both satellites would Relay Orbiter 5750 " have moved into extremely favorable viewing (4) Venus Lander Probes 3600 " positions. Types of observations which (4) Photo Sinker Probes 1532 " could be made include gross optical features such as shape, diameter, and albedo, and Total 14,162 Ibs possibly the cooling rate if the satellite is eclipsed by Mars. It should also be Mars Passage noted that there may be additional satel- ites of Mars which have not been detected. (3) Mars Surface Sample Recovery Probes 14,151 Ibs Decameter Radio Emission from Jupiter (Each establishing a remote geophysics and biologic Decameter radiation (5-40 Mcps) in station in addition to return short, intense bursts which seem to be con ing surface material) fined to cones of about 10° half-angle has been observed coming from Jupiter. More Second Passage specifically, it has been observed that the satellite lo seems to affect both the prob Same as First Passage 14,162 Ibs ability that emission will be received at Earth and the spectral character of the Assuming a 50,000 Ib experiment payload emission above about 20 Mcps. Apparently capability, this leaves about 7,500 Ibs for the emission phenomenon depends on the experiments on board the flyby vehicle. favorable positioning of both lo and the Major items here include the one meter dif Jovian longitude as seen from Earth. Obser fraction-limited telescope, the biological vations of the emission from two different laboratory, and a microwave mapping radar positions would allow a determination of its system. changing character as the lo-Jovian longi tude geometry changes. The geometry of the 5.2 Mission Return - Scientific 1977 encounter mission provides the oppor tunity to carry out such an experiment. In Origin and Evolution of the Earth, Sun, and particular, at the time of Mars passage the Planets spacecraft-Jupiter-Earth angle is approxi mately 20°. The radiometer equipment used Inferences as to the origin of the for this experiment could have several Earth and planets are based on extrapola alternate uses, such as simultaneous deca tion from knowledge of their present meter observations of the sun from the flyby physical and chemical states. Important spacecraft and the Earth. properties of the various solar system elements which will be investigated by the Objects of Opportunity flyby experiment program include the appear ance of Mars, its moons, Venus, Mercury, and The sudden appearances of comets or the asteroids; composition of the Mars asteroids near us In the solar system, and surface and the atmospheres of Mars and novae or supernovae in our own or other Venus; possible distribution and density of galaxies, provides a good secondary reason interplanetary dust; and the state of the for having an observatory available to look Mars interior, as determined by surface and In any direction on short notice. In addi orbital geophysics measurements. Knowledge tion, it would be interesting to return of these parameters, in particular the photography of the Earth, Mars and Venus similarities and differences between the taken from distances ranging from a few new bodies under investigation and the thousand miles up to an Astronomical Unit Earth, provides important constraints which to compare their appearances at different the various hypotheses of planetary origin distances. and evolution must satisfy. Astronomical observations of other stars at different 5.0 Payload Selection and Mission Return stages of development will aid in under standing the origin and history of the sun. 5.1 Payload Selection The Origin and Evolution of Life Study has shown that experiment pay- loads oh the order of 50,000 Ibs can be Samples returned from several areas on effectively utilized on a manned planetary the Mars surface will be investigated for encounter mission. This payload weight is life forms; these could be current or compatible with estimates of allowed pay- fossil life. The results of this investi load assuming a Saturn V technology based gation, carried out both on board the
10.4-12 manned spacecraft and, subsequently, in an In the case of Venus the complexity earth-based laboratory, should establish of the environment, combined with our lack the existence or absence of life on Mars. of sufficient data to describe it, sug Supplementary measurements to establish the gest that the role of the technological present, and perhaps past, chemical and return will be to determine just what physical aspects of the biologic environ shape the subsequent course of manned ment (chemical composition, temperature, exploration should actually take. In etc.) should shed new light on the mecha the event that manned landing is ruled nism of life development on Earth and the out for some time to come, the data possibilities for similar (or dissimilar) gathered during the encounter mission development elsewhere in the solar system.. will probably be used to determine the design and performance parameters for Dynamic Processes that Shape Man's second generation probes which might Terrestrial Environment work in conjunction with a manned Venus orbiter mission. One of the primary forces affecting our everyday lives is the weather, which Technological return bearing on the can be described as the interaction be overall mission may be considered in tween solar radiation and the Earth's terms of two very general but interrela atmosphere, with perturbing effects due to ted areas: first, an evaluation of the the Earth's surface (continents and performance of both man and the space oceans). Understanding features of this craft (including all subsystems) as a weather phenomenon, perhaps to the point measure of the technological require of accurate prediction and limited con ments to continue and extend our space trol, would certainly be a considerable exploration capabilities; and second, an achievement. While much experimental and evaluation of the manned encounter (flyby) theoretical work remains to be carried out mission as a competitive mode of planetary on our own weather system, it is probable exploration. Specific points related to that the study of the rudimentary aspects the first area cited above include the of a radically different weather system, physiological and pyschological effects of as quite likely exists on both Mars and long term exposure to zero gravity and Venus, will provide new insight into the confinement when no longer in Earth orbit working of our own atmosphere. Obvious and the performance of the spacecraft with parameters of importance that are varied regard to general reliability and repair- by studying Mars and Venus are solar ability effectiveness. Regarding the energy flux, mass and composition of the second area, the importance of evaluating atmosphere, rotation rate of the planet, encounter (flyby) as a mode of exploration and physical nature of the surface. can be appreciated when it Is realized that as our solar system exploration program The investigation of another planet develops, there will be an increasing num at a different stage of geologic evolu ber of objects which man will wish to tion, with different degrees of both investigate closely without ever committing internal and external activity, serves himself to an orbit or landing mode. to isolate and emphasize the role various factors play in their complex interaction Acknowledgement on Earth. For example if it were found that Venus had a molten core but no mag Much of the material in this paper is netic field, this might well provide a based on studies being carried out by counter example to the dynamo-type magne others within Bellcomm. The authors wish tic field generation we believe takes to acknowledge the contributions of these place within the Earth. Such a situation individuals. could call for a radical reappraisal of current theories. The results would, in References general, be applicable to all the planets. 1. Space Research: Directions for the 5.3 Mission Return - Technological Future, Space Science Board, NAS, NRC, December, 1965 The technological application of data derived from the Mars encounter mission has 2. A Possible 1971 Mariner Mars Landed two related functions: first, to verify Experiment, JPL Plight Projects that manned JVlars landing is a feasible goal, Document 601-1, July 24, 1967 and secondly, to determine the physical aspects of the planet environment which will .3• A Venus Lander Probe for Manned affect the planning and performance of such Flyby Missions, P. L. Chandeysson, a mission. The successful performance of Bellcomm, Inc. the MSSR probe should demonstrate manned Mars landing feasibility. Environmental data col lected by the precursory unmanned program, coupled with the encounter mission soft land- er experiments and sample return, should pro vide an adequate characterization to proceed with the follow-on program.
10.4-13 TABLE I UNMANNED PLANETARY FLYBY MISSION PARAMETERS
EARTH DEPARTURE PLANET ENCOUNTER MISSION C 3 (km2/sec2) v, (km/sec)
1971 MARS (I) 8.2 - 9.6 2.8 - 2.9 1972 VENUS (I) 13.3 - 29.1 3.8 - 5.6 1973 MARS (I) 15.0 - 21.5 2.4 - 3.0 1973 VENUS (I) 11.0 - 21.6 3.0 - 1.1 1973 VENUS (I I) 7.7 - 9.0 1.5 1975 MARS (I) 21.0 -26.0 2.1-3.6 1975 MARS (II) 13.6 - 19.3 2.1 - 2.9 1975 VENUS (I) 7.0 - 12.5 3.0 - 3.1 1975 VENUS (II) 6.8 - 12.I 3.0 - 3.7 1977 MARS (I) 20.0 - 29.0 2.1 - 3.6 1977 MARS (II) II.0 - 13.6 2.1 - 2.6 1977 VENUS (I) 7.9 - 9.1 1.I - 1.6
NOTES: (I) IN THE MISSION COLUMN, "I" AND "II" IDENTIFY TYPE I AND TYPE II TRAJECTORIES, RESPECTIVELY. (2) THE DATA ARE BASED ON PERIAPSIS ALTITUDES OF 1000 km AND 2000 km AT MARS AND VENUS, RESPECTIVELY.
10.4-14 TABLE II- MANNED PLANETARY ENCOUNTER MISSION PARAMETERS
Planet Encounter Earth Earth Aphelion/ Duration Departure Entry Mission Perihelion Periapsis Periapsis (days) C3 Velocity V» Planet Illumination (AU) Planet Velocity Altitude (krr,2/sec2) (km/sec) (km/sec) On Approach* (km/sec) (km) 1975 Mars 648 34.9 14.9 2.3/1.0 Mars 8.6 10.0 300 Sunlit Encounter
1975 Venus 368 10.5 13.5 1.2/0.7 Venus 4.6 10.9 500 Sunlit Encounter
1977 Triple 716 40.8 12.0 Venus 6.7 11.8 680 Sunlit Planet 1.6/0.7 Mars 4.4 5.6 3960 Sunlit Venus 7.1 12.0 700 Sunlit
1978 Dual- 645 27.8 13.7 1.7/0.5 Venus 10.5 14.1 1170 Sunlit Planet Mars 5.4 7.3 200 Dark
800 1978 Triple 26.3 13.1 1.6/0.6 Venus 10.3 13.7 1750 Sunlit Planet Mars 4.9 5.9 4670 Dark Venus 6.5 11.1 1910 Sunlit
1979 Mari 686 14.4 2.4/0.9 Mars 10.6 12.0 300 Sunlit Incounter 32.4 1981 Triple 790 45.4 12.5 1.6/0.5 Venus 6.0 8.8 9270 Sunlit Planet Mars 7.6 8.4 2990 Dark Venus 10,9 13.3 5060 Dark
li a h "Sunlit 1 if half or the Is by the TABLE III
BALLOON TRACKING AND DATA RELAY ORBITER SUBSYSTEM WEIGHT BREAKDOWN
ORBITAL SPACECRAFT 510 LBS POWER SUBSYSTEM 180 LBS. COMMUNICATION SUBSYSTEM 150 ATTITUDE CONTROL 80 STRUCTURE 65 EQUIPMENT 35
PROPULSION STAGE II 1200 LBS
PROPELLANT (Lb nP 325SEC) 1065 INERT WEIGHT 135
PROPULSION STAGE I MM LBS PROPELLANT (l sp 325 SEC) 3590 INERT WEIGHT 450
5750 LBS
10.4-16 TABLE IV
MSSR PROBE SUBSYSTEM WEIGHTS
RENDEZVOUS ROCKET 910
SAMPLE ACQUISITION SUBSYSTEM 88 GEOPHYSICS EXPERIMENTS 180 ADDITIONAL EXPERIMENTS 100 FLIGHT CONTROL SENSORS AND ELECTRONICS 60 RADARS 43 COMMUNICATIONS INCLUDING ANTENNAS 70 DATA HANDLING 30 POWER 243 ATTITUDE CONTROL PROPULSION 50 CABLING 80 SPACEFRAME 760 LANDING PROPULSION 900 AEROSHELL 578 INJECTION PROPULSION 380 STERILIZATION CANISTER 405
GROSS WT. 4717 LBS
10.4-17 J.D. J.D.
DATE
1977
'FIGURE
1977
RESOLUTION
TRIPLE
TRIPLE
2443252. 2443692,
OF
3732.5
3724, 3700, 3708, 3300. 3716. 3308.5 3284.
3268. 3276.
3260. 3292.
MERIT
OBTAINABLE
PLANET
PLANET
=
SEPARATION ENCOUNTER
ENCOUNTER
ON
LATITUDE
THE
(DEGREES)
MERCURY
-5.3 -6.7 -0.6
-3.2
-6.9 -4.9 -5.4
-6.7 -0.6 -3.2 +5.2 +2.3
+2.3
+5.2
TRAJECTORY
TRAJECTORY
DISTANCE
ILLUMINATED
OF
OBSERVATIONAL
MARS
EARTH VENUS
VENUS EARTH
1977
DIVIDED
-
-
MARS
EARTH
PORTION
ARRIVAL
TRIPLE
ARRIVAL-
ARRIVAL- ARRIVAL-
LAUNCH
TO
BY
TO
DISTANCE
TABLE
GEOMETRY
SEPARAT
VENUS
COS
OF
-
- PLANET
VENUS
J.D.
THE .55
;5 .3 .4 .35
.35
.7 .3
.5 .4 .6
.5
.4
.5
(
V
v
LEG
LEG
SURFACE. -
ION
(AU)
2443166.
ENCOUNTER
FOR
y),
3315.
3511. 3739.
3882.
MERCURY
WHICH
00
00
99 66
15
SPACECRAFT
ANGLE
(DEGREES)
IS
MERCURY-
SUN-
119 141
PROPORTIONAL 145 163
128
168
106
160
142 138
168
122
109
92
(u-)
TO
RESOLUTION
THE
OF
OPTICAL
BEST
1.21 1.43
1.04
1.47
1.91
.63
.64
.53
.52 .60 .44
MERIT*
.65
.53
.59
FIGURE
LINEAR
o
00 5
PROBE PROBE + +
PROBE PROBE
PROBE PROBE
SWINGBY SWINGBY
LANDER LANDER
LANDER LANDER
ATMOSPHERIC ATMOSPHERIC
VENUS
+ +
+ +
11/CENTAUR
II PROGRAM
I I
I I IMPACT IMPACT
III
MISSION MISSION
TITAN TITAN
VENUS VENUS ORBITER ORBITER
VENUS/MERCURY VENUS/MERCURY
TITAN TITAN
ORBITER ORBITER
TITAN TITAN
V
ENCOUNTER ENCOUNTER
SATURN SATURN
PLANET PLANET
PROBE PROBE
TRIPLE TRIPLE
PRECURSORY PRECURSORY UNMANNED PLANETARY
- -
PROBE PROBE
LANDER LANDER PROBE LANDER LANDER
I I
MANNED MANNED
FLYBYS FLYBYS
LANDER LANDER
IMPACT IMPACT
IMPACT IMPACT
FIGURE FIGURE
+ + I
+ +
+ +
MI/CENTAUR
III
II II
MARS
MARINER MARINER ATLAS/CENTAUR
TITAN TITAN
ORBITER ORBITER
ORBITER ORBITER (2) (2)
TITAN TITAN
ORBITER ORBITER
TITAN TITAN
1977
1976
1975
1974
1973
1972
1971
1970 1969
LU LU
>-
««
o
> to TRAJECTORY SYMBOLS A SPACECRAFT $ EARTH 9 VENUS OMARS 3 MERCURY
270
350 0° \*°
POSITION DAYS INTO MISSION DATE 1 0 JAN. 23, 1977 2 74 3 149 (VENUS ENCOUNTER) JUNE 21, 1977 4 245 5 345 (MARS ENCOUNTER) JAN. 3, 1978 6 459 7 574 (VENUS ENCOUNTER) AUG. 20, 1978 8 645 9 716 (EARTH ENTRY) JAN. 9, 1979
FIGURE 2 - 1977 TRIPLE PLANET MISSION SOLAR ORBIT GEOMETRY
10.4-20 ANTI-SOLAR ; i POI NT DEPARTURE 141 EQUATOR; ASYMPTOlF
FIGURE 3 - FIRST VENUS ENCOUNTER - PROJECTION IN THE FLYBY PLANE
10.4-21 SUS-SOLAR POINT
VIEW ON ARRIVAL
ANT I-EARTH POINT
lllf
VIEW ON DEPARTURE
DEPARTURE FIGURE ¥ - FIRST VENUS ENCOUNTER - VIEW ON ARRIVAL AND
10.4-22 GROUND
PATTERN
REPRESENTATIVE ANTENNA
SATELLITE
POLE
SOUTH
RELAY
ENCOUNTER
OF
X
VENUS
ORBIT
SECOND
ORBIT
INITIAL
ALTITUDE-\
KM
AND
CIRCULAR
4,000
AREAS
TARGET
GROUND
PATTERN
REPRESENTATIVE ANTENNA
BALLOON PROBE
ENCOUNTER
VENUS
m
METEOROLOGICAL
-
TARGETS
6
FIRST
NORTH
3
®m
ORBIT
ALTITUDE |||||AREA|
FIGURE
illllPOLE.il
KM
TERMINATORS^®:;
CIRCULAR
4,000
O
¥ GENERAL
WEIGHT
3. 2.
TOTAL BALLOON
1. STERILIZATION SUPPORT PROBE PROBE ENTRY PROPULSION
0 KM 40 KM 45
25 30
10
5
WEIGHT
30°
BALLISTIC
SUMMARY:
KM KM M ALT. KM KM
PROBE CHARACTERISTICS:PROBE
WT. WT. SYS. BALLOON
SPHERE
ALT. ALT. SYSTEM
ALT. ALT. ALT.
STRUCTURE
AT AT
AT
WT.
SYS.
ENTRY
SEP.
SEPARATION COEFFICIENT:
SYS. CAN.
-
WEIGHTS:
(LBS.)
AND
CONE
WT.
WT. WT.
ENTRY
=
1,640
SHELL
.50
(LOW
1,065 1,640
IBS
353 222
ALT.)
FIGURE
<
A
- 6
<
.90
METEOROLOGICAL
(HIGH
ALT.)
BALLOON
PROBE
i
CM
d 4
SUPER SUPER
INSULATION
EVACUATED EVACUATED
DETAIL
PACKAGE PACKAGE
PAYLOAD PAYLOAD
BALLOONS
15.7
14.6
11.4
10.6
10.8
12.4
P(IN)
8.5
7.1
7.0
9.7
15.0
21.8
d(FT)
METEOROLOGICAL METEOROLOGICAL
WEAVE
WEAVE
DEPLOYED DEPLOYED
MATERIAL
MYLAR
MYLAR
KAPTON
KAPTON
- -
BALLOON
7 7
STEEL STEEL
WALL WALL
PACKAGE
)
(LBS (LBS
PAYLOAD PAYLOAD
m
10 44
m
WT, WT,
109
122
(KM)
s ALT, ALT,
10
25
45
30
10
o JSM JETTISON
WT
t
-i*
1,065
HRS. PROPULSION
IBS
SYSTEM-
^xlO
ENTRY:
II"
ALT
WT
<
1
t
*
EPA 7xI0 -3
1,065
V
HRS
<
5
SxiO LBS
FT
13
FIGURE
1
*
FT/SEC
8
-
METEOROLOGICAL
MIDCOURSE
M
WT
t
-2.5
150 1,132
CORRECTION:
FT/SEC
DAYS
BALLOON
IBS
PROBE:
SEPARATION: ETSN STERILIZATION JETTISON
INJECTION
WT AV
WT
WT
SEPARATION
t t
t
-5 -5
-5
1,640
1,287 1,250
1,157
DAYS DAYS
DAYS
MANEUVER:
LBS
LBS FT/SEC
LBS
TO
ENTRY
CAN:
MISSION
PROFILE
o
4
CN
co
3
3
DIAM'/VT)
NOTE NOTE
NOTE NOTE
15.0
13.5 13,9 14.3
IS
AND AND
SEE SEE
SEE SEE
"MYLAR" "MYLAR"
)
2
"q" "q"
2 2
15
17 23
24
20
23
(LB/FT
HIQH HIQH
q q
LB/FT
A A
OF
IALLOON*
.75 .75
SINCE SINCE
q q
40 95
30
120
WEAVE" WEAVE"
350
280
(FT/SEC)
INFLATION INFLATION
Y T
PROBE
AND AND
IS IS
"STEEL "STEEL
MAINTAIN MAINTAIN
A A
)
2
TO TO
OF OF
FLAT",
CHUTE CHUTE
SIZED SIZED
,60 .70
,50 ,80
,88 .90
"SOLID "SOLID
(SLUGS/FT
, ,
- -
ARE ARE
A A
I I
IS IS
INFLATION INFLATION
D
INFLATION,
M/C
STAilLIZATION STAilLIZATION
(KM)
OF OF A A
T
"KAPTON" "KAPTON"
^ ^
TYPI TYPI
ONLY ONLY
S
10
25 30
140
45
ALTITUDE ALTITUDE
B B
1. 1.
2. 2.
|
DEPLOYMENT DEPLOYMENT
IN
o fi ACS JET NOZZLE CLUSTER
YLINDRICAL SOLAR CELL ARRAY - 100 FT 2 BALLOON TRACKING ANTENNA 55" x 55" SQUARE ARRAY
FIGURE 10 - BALLOON TRACKING AND DATA RELAY ORBITER
10.4-28 ALTITUDE = 4,000 KM ORBIT RADIUS =10,050 KM PERIOD = 3.08 HRS. ORBITAL VELOCITY = 5.65 KM/SEC
TRACKING ANTENNA APERTURE
FIGURE II - BALLOON TRACKING AND DATA RELAY ORBITER TRACKING GEOMETRY
10.4-29 WEIGHT BREAKDOWN
LANGMUIR PROBE^ SCIENTIFIC INSTRUMENTATION 79 LBS COMMUNICATIONS 33 LBS PROPULSION 26 LBS FREE MOLECULE PRESSURE STRUCTURE 158 LBS PROBE COOLANT 12 LBS STERILIZATION CANISTER 75 LBS HEAT SHIELD TOTAL 383 LBS
VIEWING WINDOW THERMAL INSULATION
SCIENCE PAYLOAD
COMMUNICATIONS AND POWER SUPPLY
ANTENNA
FINS
ANTENNA
FUEL TANK ENGINE
FIGURE 12 - OUTLINE OF PHOTO SINKER PROBE
10.4-30 ATMOSPHERIC - ENTRY (TIME = 0)
HEAT FINS SHIELD
PROBE
MAXIMUM HEATING AND DECELERATION (TIME = 10 SEC)
150 KM
SHIELD \ /DISCARDED
IMPACT WITH SURFACE TIME = 2000 SEC)
FIGURE 13 - PHOTO SINKER MISSION PROFILE
10.4-31
WEIGHT WEIGHT
INJECTION INJECTION
(MID
780#
FIGURE FIGURE
14- 14-
SPACEFLIGHT
VENUS VENUS
LANDER LANDER
PROBE PROBE
MISSION MISSION
WEIGHT WEIGHT
ENTRY ENTRY
PROFILE
370#
EJECTION
WEIGHT WEIGHT
LANDING LANDING
150#
4
CO CM CM X"STERILIZATION CANISTER
PROPULSION SUBSYSTEM ANTENNA
SOLAR CELLS
ELECTRONICS COMPARTMENT
SEPARATION BOLTS
SUSPENSION/EJECTION MECHANISM
HEAT SHIELD
FIGURE 15 - VENUS LANDER PROBE GENERAL ARRANGEMENT
10.4-33 ENCOUNTER TRAJECTORY
£ANTI-SOLAR PER I APS IS POINT ANT I-EARTH POINT
FIGURE 16 - MARS ENCOUNTER - PROJECTION IN THE FLYBY PLANE
10.4-34 ANT I-EARTH POINT SUB-SOLAR POINT
VIEW ON ARRIVAL
+ + SUB-EARTH POINT
VIEW ON DEPARTURE
FIGURE 17 - MARS ENCOUNTER - VIEW ON ARRIVAL AND DEPARTURE
10.4-35 I
TIME
WEIGHT
DISTANCE
A
INJECTION
V
M-5
430
4,512
FPS
DAYS
2x10
LBS.
FLYBY
KM
VEHICLE
MARS
TRAJECTORY
ENTRY
WEIGHT
DISTANCE
^V
18,000
4,132
220
FPS
KM
LBS,
VELOCITY
ALTITUDE
WEIGHT
IUE 18 FIGURE
3,554
3
1,000
KM
LBS.
-
TIME
WEIGHT
LANDING
PAYLOAD
FPS
HSSR
M-2
2,789
WEIGHT
MISSION
TO
4
LBS,
HR
1,278
PROFILE
TIME
WEIGHT
LAUNCH
LBS.
M
910
-6
MIN
LBS,
SECOND
IGNITION
WEIGHT
A
M
12,500
STAGE
204
TIME
FPS
IBS.
IGNITION
M
-0.5
MIN
Jfe
WEIGHT
BURNOUT
AM
RENDEZVOUS
10,000
76
TIME
FPS
IBS,
M
+
5
MIN
8 LANDING
LIMB OF MARS VIEWED FROM MANNED VEHICLE
LOCUS OF POSSIBLE LANDING POINTS TO MANNED VEHICLE
FIGURE 19 - MSSR ENTRY AND LANDING GEOMETRY
10.4-37 LOCUS OF POSSIBLE LANDING ! POI NTS
COPLANAR LAUNCH SITES FLYBY PLANE
SELECTED L LANDING SITES
FIGURE 20 - MOTION OF LANDED MSSR PROBES DURING STAY TIME
10.4-38 3 -
2 _
10° 20° 30°
LANDING LATITUDE (DEGREES SOUTH)
FIGURE 21 - MSSR TARGETING
10.4-39
STAGE STAGE
2 2
FIGURE FIGURE
TANK
22 22
- -
?AYLOAD ?AYLOAD
MSSR MSSR
PROBE PROBE
COMPARTMENT
(LANDED (LANDED
CONFIGURATION)
EQUIPMENT EQUIPMENT
CONICAL CONICAL
& &
DEPLOY DEPLOY
SYSTEMS SYSTEMS
DOORS DOORS
PROVIDE PROVIDE
TO TO
CONTAIN CONTAIN
STRUCTURAL STRUCTURAL LOCATED LOCATED
CLEAR ASCENT ASCENT CLEAR
BLAST BLAST
LANDED LANDED
PROTECTION PROTECTION
IN IN
EQUIP. EQUIP.
SHELLS SHELLS
PAYLOAD
VEHICLE VEHICLE
COMPT
FOR FOR
o
3 o* PER I APS IS
ENCOUNTER TRAJECTORY
FIGURE 23 - SECOND VENUS ENCOUNTER - PROJECTION IN THE FLYBY PLANE
10.4-41 SOUTH POLE VIEW ON ARRIVAL
SUB-EARTH' POINT
VIEW ON DEPARTURE
FIGURE 21 - SECOND VENUS ENCOUNTER - VIEW ON ARRIVAL AND DEPARTURE
10.4-42 The Zodiacal Light (The drawing is in the plane of the ecliptic)
Scattered Sunlight
'. i *v V£i/!/•>».''• ',**•- ••.:r..:.^.« !?; Interplanetary Dust ^ Field of Viev ••••^,v.^X^^-- " ^^£S£ff J
Earth
Sun
Light Intensity of the Sky Background in Space vs. 8 (Schematic)
Solar Corona Zodiacal Light
-C O) O) O
FIGURE 25 -
10.4-43 SPACECRAFT TRAJECTORY
VERNAL EQUINOX 3308.5 J.D.-3252.5
= VENUS ENCOUNTER
FIGURE 26 - SUN-MERCURY-SPACECRAFT GEOMETRY ON FIRST VENUS PASSAGE
10.4-44 SPACECRAFT TRAJECTORY
J.D.-3692.5
SUN
3732.5
= VENUS ENCOUNTER
VERNAL EQUINOX
FIGURE 27 - SUN-MERCURY-SPACECRAFT GEOMETRY ON SECOND VENUS PASSAGE
10.4-45
TO TO
28.7° 28.7°
OF OF
(DIRECTION (DIRECTION
SUN
DIAGRAM)
BELOW BELOW
ACTUALLY ACTUALLY
PLANE PLANE
FIGURE FIGURE
28 28
- -
POSITIONS POSITIONS
OF OF
PHOBOS PHOBOS
AND AND
DEIMOS DEIMOS
ON ON
THE THE
1977 1977
TRIPLE
d
5 to