A1.5-HOUR LAUNCH WINDOW on November 2, 1973 (November 3, on the East Coast), provided the best science data return possibility for the mission. Although there were opportunities in other years, the 1973 opportunity offered one of the lowest launch energies to swing by Venus and subsequently encounter Mercury. When the project was formally initiated in December 1969, four years were available to plan and implement this complex new interplanetary mission: the first use of gravity-assist and the first two-planet mission to be undertaken by the National Aeronautics and Space Administration.
Not only was the mission under tight cost constraints that demanded use of new management philosophies, but also some significant changes had to be made to the earlier Mariner spacecraft to meet the special requirements of the Venus-Mercury mission.
Scientists, too, were constrained in their experiments-the rule was to achieve maximum science for minimum new development. Since there were options on the flyby path at Mercury encounter, conflicts in the demands of scientists arose from science opportunities offered by these different modes. The mode finally selected, passage on the night side, provided good conditions for nonimaging science return but was the worst situation for imaging science. To meet these constraints the TV imaging system had to be redesigned, and there were demands for real-time return of data at satisfactory error rates. For example, while detailed analysis had shown that TV imaging could accept relatively high rates of bit errors (about 1 in 50) and still produce high-resolution pictures of suitable quality, the other science experiments had to be assured of very low bit error rates (about I in 10,000). To constrain the TV system to such low bit error rates would have considerably reduced the number of TV images and made it impossible to produce a full-disc, high-resolution mosaic of Mercury during the short period available for TV imaging during the night-side pass. The solution to this conflict was the implementation of a two-channel, independently commendable data stream, using a new form of carrier modulation devised for this purpose by JPL's telecommunications engineers.
Several activities connected with the design of the spacecraft produced conflicts of requirements to meet the objectives of the mission. Quick and satisfactory resolution of these conflicts was a continuing challenge to the management of the program, which had to meet the tight schedule and yet still keep costs within the budget limitations.
 Upgraded Spacecraft Capability
A decision to increase the size of the nitrogen gas tank, thereby increasing the amount of reaction control gas from the 2.45 kg (5.4 lb) used by Mariner Mars 1971 to 3.62 kg (8.0 lb) for Mariner Venus/Mercury was made early, principally to accommodate predicted worst-case effects of solar pressure on the appendages and to allow for total depletion of one of the redundant halves of the gas system by a valve failure early in the mission. Without this change, there would not have been sufficient reaction control gas-a margin of 0.91 kg (2 lb) was provided-to allow the extended mission, which added so much to the science coverage at Mercury by three encounters with the planet.
Originally the spacecraft was to be capable of a maneuvering velocity change of 56 m/sec (184 ft/ sec). By using an improved propulsion unit, incorporating a larger tank used on the Pioneer 10 and 11 spacecraft, the final capability of the spacecraft was more than doubled, to 122 m/sec (401 ft/sec). Again, this made possible the multiple encounters with Mercury.
The solar panels, originally specified as being of a fixed tilt of 60 deg, were, in the final spacecraft design, capable of an independent variable tilt, from 0 to 76 deg, on command from the ground. This capability was eventually used for " solar sailing " which, as it turned out, made the extended mission possible when trouble developed in the gyro system. The original two-position, low-gain antenna was upgraded to provide three positions on command, allowing communication with Earth following first encounter. Moreover, the final articulation pointing system of Mariner Venus/Mercury was much improved over the original scan system based on the Mariner Mars 1971 design. High-gain antenna articulation was increased from one to two degrees of freedom to permit transmitting to Earth at all times instead of just at Venus and the first Mercury encounter. All elements controlled by the subsystem-the scan platform, high-gain antenna, and solar panels-could be pointed to within 0.35 deg and their positions reported back to the flight data system for telemetry to Earth within 0.1 deg in the position mode. In the incremental mode the subsystem positioned the scan platform within 0.075 deg.
Sail vs V-Tilt Solar Panels
On Mariner spacecraft the thousands of solar cells that convert sunlight into electrical energy were mounted on the face of flat rectangular panels extending like wings from the spacecraft. Since the new spacecraft had to travel from the orbit of the Earth to that of Mercury, its solar cell energy-gathering system had to accommodate to the change of nearly S times in the amount of solar radiation that would be received. Early studies by JPL and Boeing concluded that the best way to keep the solar panels at the right temperature of about 100°C (212°F), while still providing a fairly constant power output from them into the spacecraft electrical system and also meeting the weight constraints, would be to....
....articulate or tilt the solar panels as the spacecraft approached the Sun. Furthermore, after considering two ways of tilting the panels, a V-tilt was chosen (Fig. 4-1). However, as the project developed, it was discovered that V-tilt might lead to unacceptable thermal input to the bus and the instruments mounted on the platform, so a study was started to compare V-tilt and rotatable ("sail") configuration.
The study concluded that even though the structures and mechanisms needed for a sail configuration (Fig. 4-2) were more complicated, weighed more, and would cost more than those for a V-tilt, scan platform temperature control would be simpler, and the solar panel temperatures would also be lower at high angles of tilt. In addition, this design permitted the mounting of the roll/yaw cold gas jets at the ends of the sails, providing added leverage for their thrust. Accordingly, the sail configuration was recommended by Boeing in August 1971, and was accepted by Project management at JPL.
Protecting the Rocket Engine From the Sun
Another problem arising from the close approach to the Sun was protecting the maneuvering rocket engine from direct solar radiation and preventing the conduction of solar heat from the engine into the octagonal equipment compartment. Earlier studies had concluded that the preferred direction of the rocket nozzle should be toward the Sun during the cruise phase of Mariner operations.
The basic configuration proposed by Boeing at the time of contract award made use of a thermal door over the rocket nozzle to minimize the effects of solar radiation. But this door imposed a reliability problem if it should fail in either an open or a closed position. If the door remained open it would allow the propulsion subsystem to overheat. If, on the other hand, the door failed to open and remained closed it would be blown off the first time the engine fired. Again the spacecraft would be unprotected and would have to survive the resultant solar heating. Since in both failure modes spacecraft survival without the door would be necessary, engineers took a new look at the true effectiveness of the rocket engine door to ascertain whether or not it could be eliminated from the spacecraft design.
The main parts of the propulsion system that had to be safeguarded from solar overheating were the propellant tank, the valve that controlled the flow of propellant to the rocket thrust chamber, and the thrust vector control actuator. Engineers at Boeing used a computer thermal model of the propulsion unit and found that with suitable modifications the engine could survive solar heating without a door.
The modified design ensured that radiation entering the sunward-facing nozzle would be radiated into space from the thrust plate and from  the inner surface of the nozzle. The plan was that surfaces of the thrust plate and the thrust vector control actuator would be coated with a paint that emitted heat readily, while the outer nozzle surface and the inner surface of the thrust vector control actuator support were polished to reflect radiant heat from the inner nozzle. Finally, the propellant tanks and lines were protected by multiple layers of insulation.
Although the engineering analysis indicated that it was theoretically safe to eliminate the door, engineers decided that the consequences of such a decision were so far-reaching that a practical test had to be made. Accordingly, high priority was given to a verification test in which actual materials were tried out under simulated conditions. The theoretical analysis was vindicated, and the tests confirmed that the engine thermal door with its attendant problems of reliability could be eliminated from the flight spacecraft.
Further Protection From the Sun
In addition to solar panels and propulsion system, many other components of the spacecraft needed protection from solar radiation. At the beginning of the program, considerable doubt existed that Kapton, a commonly employed heat-protection material, would survive in the anticipated environment. Kapton had been suggested as a replacement to the Teflon used on earlier Mariner spacecraft when engineering tests showed that Teflon failed at the intensities of solar radiation expected at Mercury. However, Kapton was found to become brittle with long exposure to temperatures above 354°C (670°F) and also in the environment of ultraviolet light and protons expected sunward of Earth's orbit.
The scientists at Boeing quickly subjected several alternative materials to exhaustive tests at the Boeing Radiation Effects Laboratory. But it was discovered that the tests were inconclusive because of contamination of the material by an unidentified substance and because the flux of protons was not large enough to simulate the flux at the distance of Mercury. There was no time to conduct a further test series, so a working group of materials experts from JPL and Boeing sought acceptable alternatives to Kapton. After much investigation and long working hours the group....
....was able to recommend that several alternatives were indeed available: stainless steel cloth backed by aluminized Kapton; metallic foil of titanium, stainless steel, or aluminum; Teflon-coated glass fiber cloth, known as beta cloth, aluminized on one side and backed by aluminized Kapton; clear-anodized polished aluminum, known as Alzak; and optical solar reflectors. Project management selected the beta cloth and the clear-anodized aluminum for more detailed study.
Checks were made to see if the outgasing from the beta cloth and its subsequent loss of weight resulted in deposits on neighboring surfaces and what, if any, changes took place to the cloth, such as discoloration and reaction with other spacecraft materials. The results were that Teflon-coated glass cloth could be expected to survive the environment at Mercury encounter. Analysis of the beta cloth suitability for the sunshade showed that even if the cloth discolored and turned black, the sunshade would still function adequately.
 The working group therefore selected a foldable beta cloth sunshade and suggested that all sunlit thermal blankets used on the spacecraft should have an outer layer of beta cloth also.
Two flight sunshades and a backup unit (Fig. 4-3) were therefore designed and fabricated. The use of foldable designs resulted in the project's being able to make good use of existing hardware that had demonstrated reliability for deploying sunshades in space, but because of the increased weight of the beta cloth, some modifications were needed to strengthen deployment springs and the deployment assembly generally.
Making Sure the Spacecraft Obeyed Commands
Whereas future interplanetary spacecraft will have a redundant central computer and sequencer that provides alternative paths, Mariner Venus/ Mercury had only one computer and sequencer. A backup command capability had to be provided through the flight command unit that would not leave the articulation and pointing system susceptible to command errors. This system pointed the high-gain antenna and controlled the scan control subsystem for directing both the TV and ultraviolet instruments at the target planet. The impact of an incorrect command on the scan platform was unacceptable since it would take too long to detect back on Earth and issue a correcting command.
After considerable debate and test activity on different designs, it was decided to use both a position mode and an incremental mode design. Thus an initial position in a typical scan sequence would be commanded by position commands, and then the following frames of a photomosaic would be acquired by incremental updates. These joint modes obviated problems of storage of sufficient position commands for a complete mosaic sequence, while at the same time they ensured that should a command be in error the scan platform would be returned to a correct position for a subsequent sequence with loss of only part of the encounter sequence, and not all of it. Also, most important, the incremental command mode was required to allow fine stepping with sufficient picture-to-picture and slit-to-slit precision for mosaicking and proper UV scanning.
Fine Tuning for Encounter
Once the spacecraft was placed on a trajectory to Venus, and when this trajectory had been accurately determined by tracking, controllers had to fine-tune the trajectory to get a more precise Venus encounter that would lead to the minimum use of propellant to fly by Mercury at the correct time, distance, and orbital inclination. Such trajectory correction maneuvers (four were planned but only three were actually needed during the flight to the two planets, five more for the subsequent returns to Mercury) relied upon the spacecraft's being oriented accurately in space by command and then provided with a given thrust for a known period of time in a definite direction. The spacecraft carried a main propulsion system for this purpose.
The Mariner Venus/Mercury main propulsion system was a modified Mariner Mars 1969 design, but by the spring of 1971 additional velocity change requirements were imposed on this propulsion system by mission planners. More propellant had to be carried aboard the spacecraft, and a larger storage tank was needed. The one used on Pioneer 10 and 11 was chosen. By 1973, Boeing, which contracted with JPL to produce the new propellant subsystem, had delivered a subsystem mockup and, just over a month later, a development test model of the new subsystem. The choice of a "blowdown" design, wherein the driving gas pressure on the propellant is allowed to diminish from firing to firing, was a major departure from earlier Mariners and represented a first for this class of spacecraft. Another change was made later when the Skylab program experienced difficulties with thrusters, which had oxidizer valves and valve inlet fittings identical to those on Mariner Venus/Mercury. The time was August 1973, only three months before scheduled launch of Mariner. Propulsion engineers reworked the subsystem and replaced valve inlet fittings. Propellant loading was completed with only six days left before the scheduled time for installation in the spacecraft.
 Test, Test, and Test Again
Interplanetary spacecraft must be reliable. Once the spacecraft has been launched into space, a failure cannot be repaired. The spacecraft must either carry a redundant part to replace a failed part, or controllers must devise ways to complete the mission by working around the failed part. Because many parts of a spacecraft are critical- their failure could be catastrophic-the whole process of designing and building and launching a spacecraft is accompanied by test upon test upon test.
For example, the high-gain antenna required considerable development to meet the performance requirements of Mariner Venus/Mercury, especially for the distant Mercury encounter. Every microwatt of transmitted power was required at Mercury to get the many pictures back to Earth during the short period of the encounter. When a new combined S- and X-band feed was installed on the antenna dish it was found that the expected gain from the antenna was below that required for the 1220-mm (48-in.) diameter antenna. After considerable effort to increase the gain, the decision had to be made to increase the diameter of the reflector to 1370 mm (54 in.) to obtain the required effective radiated power.
Similarly, development testing of the articulation and pointing system took place during July and August of 1972 to confirm that the backlash and stiffness of each movable unit were acceptable, to make sure that sufficient reserve of power was available to move the various actuators, and to check on the positional accuracy with which the scan platforms could be moved. Checks were also made to ascertain that these various movable elements could be unlatched from their stowed position and would reliably lock up in their operating positions.
During these tests some problems were encountered in locking up the dish of the high-gain antenna, resulting in the requirement that the lockup pin be changed. Later, these same development tests (Fig. 4-4) were repeated on the two flight spacecraft, and all functions were found to be satisfactory.
A most critical item to the success of the mission was the solar panel system, since if these panels did not move from the stowed to the operating position, the spacecraft would be starved of electrical energy and could not operate.
Viscous boost dampers had to be developed for the solar panels as well as for the magnetometer boom to prevent the two systems from banging together during the launch. The objectives of this development program were to verify that the damping force would meet the requirements for the spacecraft over the operational range of vibration frequencies, and to develop assembly techniques for the magnetometer boom damper. During the development program, both dampers were subjected to small- and large-amplitude vibration testing at many frequencies. These tests were successful and showed that the dampers would safeguard the solar panels and the magnetometer boom.
A development test fixture was used to verify that the release mechanism exerted sufficient force to unlatch the solar panels and the other deployable structures and that there were adequate forces to move all these systems into flight position at the required rate. The test fixture was also used to test the structural strength of the solar panels and other assemblies to make sure that they would be capable of withstanding the...
 ....forces exerted during launch. Reliable operation was demonstrated after several modifications had been completed. Since the appendages mounted on booms had to deploy in the weightless or zero-g condition of flight through space, other tests were performed to simulate the zero-g condition. As a result of these tests it was necessary to reduce the rate at which the low-gain antenna deployed, to modify the high-gain antenna's latching mechanism by adding a kickoff spring to the boom release, and to develop a special tool for the assembly of the restraint mechanism for the plasma science experiment.
As the program progressed, extensive testing took place with the developmental test model of the spacecraft, which was subjected to high- and low-frequency vibrations and to acoustical inputs ("noise") of various frequencies (Fig. 4-5). All the structures and mechanisms tested were either flight hardware or flightlike hardware, except for the high-gain antenna reflector and some other minor components which were simulated. Thus it was shown that the spacecraft and all its subsystems could withstand the vibrational stresses of the large rocket engines of the launch vehicle. Test results agreed closely with the calculated effects of vibration. No failures or excessive deformations were observed during any development test model vibration testing, nor during the test firing of pyrotechnic charges on  the spacecraft to release structures from launch stowage to the cruise positions. The subsystems of the Mariner Venus/Mercury spacecraft had been certified with a clean bill of health to ride in the launch vehicle to interplanetary speeds.
The spacecraft was also tested rigorously to determine its ability to survive the thermal stress of the inner Solar System (Fig. 4-6), the degradation of external surfaces by the environment of space, and hot-firing of the rocket engine. The high-gain antenna was also tested for its ability to perform when exposed to the intensity of nearly five suns' radiation (Fig. 4-7).
Successful deployment of the sunshade was critical to the success of the mission to Mercury. If it failed to open and shield the spacecraft, the intense solar heat would damage the electronic and scientific equipment. To establish confidence in the concept used for deployment of the sunshade, a development test (Fig. 4-8) faced the sunshade upward and downward to deploy both against and with the force of gravity. The sunshade's deployment was also tested with broken lanyards, nonsymmetrical solar panel deployment, and broken deployment springs. The sunshade passed all these tests, showing that it would deploy under the most severe conditions of a combination of possible failures.
The final series of tests began when the flight spacecraft was subjected to rigorous solar thermal vacuum conditions in the space simulator at JPL (Fig. 4-9), where testing simulated the harsh solar environment that the spacecraft must expect as it speeds into the inner Solar System. The spacecraft was again tested to make sure that its performance met all specifications, followed by further tests to ensure that it was compatible with the Deep Space Network and the Mission Operations System. These tests at JPL occupied most of July 1973, and the spacecraft came through them with flying colors. There were, as expected, some minor glitches and some differences in the actual temperature experienced at equivalents of 1, 2, and 4.8 suns, but these were within tolerances, and the spacecraft was considered ready for launch.
Meanwhile, at the Eastern Test Range, the backup spacecraft had arrived from Boeing on August 4 and went through a series of tests to verify the adequacy of the test procedures and all the spacecraft-related equipment that would later be needed for the flight spacecraft on its arrival at the range.
As with most complex technical programs, many problems beset the engineers and scientists as they developed the new spacecraft. And, as always, the challenge was to identify and resolve the problem as quickly and as inexpensively as possible.
Once, for example, during testing at the Boeing facility, the radio frequency subsystem lost uplink lock while operating near threshold. Another similar unit lost lock while operating during solar vacuum testing in the test chamber at JPL. Both units had to be returned to the subcontractor, where the trouble was corrected by the installation of a filter in the wiring.
Another problem was identified early during evaluation tests of the data storage subsystem. The tape transport failed to start consistently from the left-hand end of the tape-parking window, and engineers thought that this might be due to wear that was caused by low humidity of the magnetic tape itself. To overcome the difficulty, the unit was preconditioned prior to shipment to the launch site. A humidified mixture of argon and helium was used within the tape transport while the tape was being run continuously. The final relative humidity of both the magnetic tape and the internal atmosphere of the tape transport was established at the required level. Actual usage of the system in flight exceeded that proposed  before the flight. Mission playback of both high and low-rate recorded data was very good, with low bit error rates.
In December of 1972, the spacecraft battery was accidently connected to the attitude control electronics and several electronic modules and harnesses were damaged. Not only had the damaged modules to be replaced but nearby components had to be checked thoroughly to make sure that they, too, had not suffered stress from the accident. Further tests called for replacement of more modules, and the rebuilding caused a delay in the spacecraft systems test which could only be regained by a specially devised " catchup " test in which the attitude control electronics were powered-up continuously for several days and all the functions of the subsystem were exercised faster than normally. In this way the tight schedule for delivery of the spacecraft to meet the launch window was not affected.
During design and development of the magnetometer boom, it was discovered that intrinsic frequency characteristics would likely interact adversely with the spacecraft during maneuvers, causing dangerous vibrations. As a result, the engineering group at Goddard Space Flight Center had to redesign the sensor canisters, brackets and cable supports. The weight reduction shifted the characteristic frequency away from the dangerous region. Subsequent testing during the summer of 1974 at the Boeing Company's facility called for the zero-gravity environment to be simulated. A special compensation string and support was developed for this. Unfortunately, during one of the deployment tests, the string broke and the boom was dropped. A critical interface bracket and inboard hinge of the two-section boom was damaged. A rather frantic rebuilding effort had to take place at Goddard Space Flight Center to provide another qualified and tested boom for flight. Much weekend and overtime work was demanded and trips to international scientific meetings in August and September had to be cancelled for the principal investigator but the deadline was met.
Meeting the Milestones
It is a long hard road (Fig. 4-10) from contract award in response to a written proposal to the shipment of a finished spacecraft to mate with its launch vehicle at the Kennedy Space Center,...
 ....Florida. It is a road punctuated by milestone events which must be reached at certain times in order to meet the launch window, predetermined by the inexorable movements of planets in their orbits.
The planets will not wait on human failure. Men have to work and produce and be ready with their space machines on time, or the whole effort is in vain and the opportunity for an interplanetary mission is lost. Often the opportunity is not repeated for decades, sometimes centuries. A special breed of men and a special type of human endeavor are required to meet the requirements of space missions.
Management of a planetary mission requires the discipline of control of each major effort in the program. It must define accurately the events, activities, and resources necessary to reach objectives on time and within budgets. Responsibility for all tasks has to be clearly defined.
Major milestones in a program master schedule provide key dates from which detailed schedules for work units are derived. With Mariner Venus/ Mercury, many formal progress reviews took place, keyed to program events that earlier had been identified as major indicators of progress significant in previous planetary efforts. Free communications on program matters speeded management decisions when corrective actions were needed to keep the program on schedule.
Considerable emphasis was placed on early identification and reporting of problems. Special technical sessions followed each regular monthly program review and identified problems needing solutions. In June 1972, Boeing instituted a weekly log prepared by the work unit personnel in each area of activity and summarized in a weekly problem report. When the test phase of the program was reached, daily meetings were held with test and operations people. But in all this activity the basic premise was that the success of the project depended upon men and women, not upon management systems. Dedicated people were supported by good communications to top management.
Following selection of the experiments to be flown on Mariner Venus/Mercury, the Project Science Steering Group, consisting of principal investigators and science team leaders, was constituted. The emphasis of the Science Steering Group was upon interaction of the scientists with the project. One example was the design of the telemetry system with regard to the allocation of the rate at which information would be transmitted from spacecraft to Earth. In negotiating with the experimenters early during the project, the scientific experiments and their interfaces with the spacecraft flight data subsystem were discussed.
In one meeting of the Science Steering Group at JPL, the principal investigators got together with the project staff and allocated the two planned bit rates of 2450 and 490 among the various experiments. This was the time at which the need for a lower bit rate was identified because the higher rate could not be accommodated during the extended mission on the 26-m (85-ft) antenna net. The range of rates needed was from 1050 bits/sec for magnetometer experiments to only 33 bits/sec for the infrared radiometer. There were some conflicting requirements aired at the meeting, but these were all resolved through mutual understandings and discussions.
Another example was the quick recognition that not only did bit rates differ for the different experiments but also the quality of data needed for the TV experiment was quite different from the nonimaging science. Whereas the latter could accept only very low bit rate errors, the photo scientists could obtain images of usable quality at high bit error rates, thereby obtaining more pictures at higher transmission rates.
The imaging experimenters also needed the flight data system to have the capability to command a quick change from high to low transmission rates or vice versa depending on inspection of the quality of the incoming pictures in real-time. The result was that it was decided to decouple the imaging science data from the nonimaging science data by the use of separate high and low data rate channels. Dr. Stan Butman at JPL designed a special modulation scheme, termed interplex modulation, which permitted decoupling of two data transmission rates. The allocation of power in two subcarriers was changed so that a cross modulation channel could be used to carry the lower data rate subcarrier.
 Another major interaction between scientists and engineers was in regard to the placement of instruments on the spacecraft to provide suitable fields of view necessary for the science experiments. Several months after the start of spacecraft design at Boeing, the locations of all instruments were changed in order to better satisfy requirements of scientists for the fields of view of their instruments. The Canopus tracker's 30-deg offset resulted from this rearrangement.
Thus, in the early phases of spacecraft design, there was much interaction among JPL divisions and project staff, the Boeing Company, and the principal investigators, separately and as a science group-this interaction being aimed toward designing a spacecraft that would really do the exploratory job assigned to it and do the job well. Costs of the science experiments were also rigorously controlled.
There was a very strong interaction between the project and the principal investigators in defining the data records and detailed planning on how these records would be obtained during the mission. Originally, it was planned to use the real-time data from the stations and to generate the master data record from this real-time data and a log tape called the system data record. When this plan was examined more closely with the principal investigators, it was ascertained that the amount of data recovery would, from a percentage point of view, be very high: on the order of 95%. But with the help of people who had done systems analysis of data returned from the Pioneer 10 spacecraft to Jupiter, it was found that the way in which the errors were distributed in the real-time link was such that there would be uninterrupted error-free data for brief periods of only a few minutes. At other times, the data would contain errors. So the design of the data system was completely changed-the original data record as recorded at the DSN station was to be flown to JPL and merged with the system data record obtained in real-time over the ground data links to produce a relatively error-free master data record from which the scientific data would be supplied to the experimenters. During the encounters, these error-free data were available to the investigators within a few hours. During the cruise mode, where time was not so important, the data were made available within 1 to 2 weeks.
All principal investigators met the schedules for delivery of their science instruments, although there were some anxious times. A low-energy telescope was added to the particle experiment to extend the lower bound of charged particle measurements and to permit low-energy protons to be detected in the presence of low-energy electrons. Changes were made to the plasma science experiment, which originally proposed a body-fixed biaxial detector to look in an antisolar direction. This instrument was deleted at the time of selection to reduce cost. Later, at the suggestion of the plasma experimenter, a single-axis instrument was added to the scan package at no additional cost. This was fortunate because the main instrument on the scanning platform failed because of a stuck protective door, and the only plasma data were obtained with the add-on unit.
A third addition following initial selection of science instruments was the wide-angle filter that was added to the television optics in order to allow the search for structure in Venus's visible clouds. The path by Venus was from the planet's dark side, so the only good phase angle views between spacecraft, planet, and Sun suitable for cloud analysis would be around closest approach. But at closest approach the high-resolution cameras of Mariner would show only a very small area of clouds, much smaller than the scale of the searched-for features, and thus a wide-angle capability was required.
Preparing for Launch
Before launch, during the summer of 1973, an almost true-to-life launch was simulated and carried off at the Mission Control and Computing Center at JPL (Fig. 4-1 1). Members of the....
 ....mission operations system team spent hours of intense concentration as the various teams went through the exercise of a mock launch and initial collection of data from the simulated flight. This exercise tried out all the ground systems needed to support the spacecraft on its long mission.
Meanwhile, the Mariner Venus/Mercury spacecraft was shipped from the Boeing plant on Friday, June 30, 1973, to JPL, where it went through exhaustive tests in the solar simulator, as discussed above. The spacecraft left the Laboratory in early August in carefully packed sections aboard a convoy of specially equipped vans en route to Florida by road. It arrived at the Air Force Eastern Test Range at Cape Kennedy on August I 1, 1973, and was placed in Building AO's spacecraft checkout area for final verification tests. In providing launch operations, the John F. Kennedy Space Center handles scheduling of test milestones and review of data to assure that the launch vehicle has met all of its test requirements and is ready for launch.
Atlas/Centaur 34 was erected on Complex 36's Pad B in July 1973. The flight spacecraft was moved into the Explosive Safe Facility on September 25 for installation of ordnance and loading of its hypergolic propellant. It was encapsulated for mating with its launch vehicle in mid-October.
A flight events demonstration test took place successfully during the third week of October to assure that the space vehicle was electrically ready for final launch preparations. The test included running the computer and programmer through post-launch events and monitoring the data to assure correct response to all signals when the umbilical was ejected.
About 10 days before the planned launch, the spacecraft was mated with Atlas/Centaur and further electrical tests were conducted (Fig. 4-12). The Composite Electrical Readiness Test for the overall space vehicle took place a few days prior to launch to verify the ability of the launch vehicle-spacecraft combination to go through post-liftoff events. Range support elements participated along with the spacecraft and launch vehicle just as during a launch.
The launch (T)-1 day functional test involved final preparation in getting vehicle and support ready for launch, preparing ground support equipment, completing readiness procedures, and...
 ....installing ordnance on the launch vehicle ( Fig. 4-13). The countdown was picked up at T-600 min. All systems were checked against readiness procedures, establishing the integrity of the vehicle and ground support equipment interface prior to removing the tower at T-120 min. Cryogenic propellants of liquid oxygen and liquid hydrogen began to flow into the launch vehicle's tanks at T-100 min, culminating in complete vehicle readiness at T-1 min. The terminal count began with monitoring all systems and topping off the venting propellant and purge systems. At T-10 sec, the automatic release sequence was initiated and the space vehicle was cleared for liftoff.