Category Soviet Robots in the Solar System

Lunar and interplanetary missions required facilities for tracking spacecraft on their journeys through deep space and to communicate with them for navigation, control and data acquisition. Korolev chose Ycvpatoria in 1957 as the site for these facilities because it offered a southerly latitude near the plane of planetary orbits. It was also conveniently close to the Black Sea and Crimean resorts.

Known as the Center for Long Range Space Communications (TsDUC), its first facility was a 22 meter antenna built in 1958 for lunar missions. The first phase of construction for planetary missions was ready in 1960. Korolev built his receiving stations by scavenging old naval parts – using the hull of a scrapped submarine, a revolving turret from an old battleship, and a railway bridge on top of the turret to hold the antenna array. Each array consisted of eight antennas in two rows of four all of which moved in unison. There were two sites, one to the north for the receivers and the other to the south for the transmitters. The receiving station had two such antenna arrays, each using eight 15.8 meter dishes. They operated in the meter band at 183.6 MHz, in the decimeter band at 922.763 and 928.429 MHz (32 cm), and in the centimeter band at 3.7 GHz (8 cm) and 5.8 GHz (5 cm). The transmitter station had one array of eight 8 meter dishes. This ‘Plutori transmitter was rated at 120 kW and operated at 768.6 MHz (39 cm). A ground-link microwave station was set up for transmitting data to a second station at Simferopol and then on to other locations in the USSR. The TsDUC facilities went online September 27, 1960, only a day before

Figure 3.4 Receiving station at north facility (left) and transmitting station at south facility (right), Yevpatoria.

Figure 3.5 Cosmonaut Yuri Gagarin, king of the Soviet tracking ship fleet.

the optimal launch date for Mars, although the first Mars launch did not occur until October 10. Between 1963 and 1968 Yevpaioria and Simferopol each received a 32 meter “Saturn’ dish, and five others were installed at Baikonur in Kazakhstan, Sary Shagan in Balkash, Shclkovo near Moscow, and Ycniscicsk in Siberia. In 1979 a 70 meter ‘Kvanf dish was built at Yevpatoria. TsDUC now also has a 64 meter antenna at Bear Lake near Moscow, and a 70 meter dish in Ussurusk near Vladivostok. All deep space missions were operated from Yevpatoria until a new control facility was opened in Moscow in 1974.

The USSR did not have a worldwide network of tracking stations like NASA’s Deep Space Network, with serious consequences for deep space mission operations. Critical operations such as planetary encounters had to be planned for times when the spacecraft could communicate. And since signals could not be picked up when a spacecraft was below the horizon, spacecraft were designed to transmit only when Yevpatoria had a line of sight. This system required carefully controlled timing of spacecraft operations and reorientation of the spacecraft for high-gain operations. To provide a measure of relief from the limitations on spacecraft operations imposed by a single ground station, the Soviets deployed tracking ships in the world’s oceans. These ships also tracked missile tests, covered manned space missions, and tracked interplanetary missions making the second firing of their upper stage to escape Earth orbit into interplanetary space. The ships were not a wholly satisfactory solution for deep space tracking, as only small dishes could be mounted on the ships and weather conditions could severely hamper operations. The first ships deployed in 1960 were the Illchevsk, Krasnodar and Dolinsk. In 1965/6 the Illchevsk and Krasnodar were replaced by the Ristna and Bezhitsa. A third generation consisting of the Borovichi* Kegostrov, Morzhovets and Neve! were deployed in 1967. These were all converted merchant ships of about 6Л00 tons displacement with crews of 36. In May 1967 the first purpose-built traeking ship was introduced, the Cosmonaut Vladimir Komarov (17.000 tons). The Cosmonaut Yuri Gagarin (45,000 tons) and Academician Sergey Korolev (21,250 tons) joined the fleet in 1970. In addition, a number of smaller traeking vessels were deployed: the Cosmonaut Pavel Belyayev, Cosmonaut G corgi Dobrovolskiy Cosmonaut Viktor Patsayev and Cosmonaut Vladislav Voikow

The Mars program was the first to make use of the powerful Proton launcher. The 1969 launch opportunity was a particularly favorable one and the large increase in available spacecraft mass offered the opportunity to attempt a soft landing mission, but it was decided on this initial campaign to send a spacecraft that would release an entry probe from orbit around the planet. If successful, it would provide the first in- situ measurements of the Martian atmosphere.

The first design for the Mars-69 spacecraft took advantage of all the work that had been done for the new lunar landing vehicle, the Ye-8. This design ultimately turned out to be impractical for Mars and a complete redesign produced an in-line modular configuration similar to, but much larger and more robust, than the preceding 3MV spacecraft.

The core of the Mars-69 spacecraft was a spherical propellant tank which had the engine beneath and a cylindrical section above, and held the solar panels, antennas and thermal control system. The navigation system and orbital instrument modules were mounted on opposite sides of the tank. The entry probe was installed above the cylinder. The avionics were a significant improvement on the 3MV spacecraft. Due to insufficient time for testing and significant growth in spacecraft mass, the entry vehicle had to be deleted from the 1969 campaign. Two identical orbiter spacecraft were launched in late March and early April, 1969, and unfortunately both were lost to launch vehicle failures.

Campaign objectives:

By the end of 1962, the Soviets had made five attempts at Venus. Only one mission. Venera 1, survived its launch vehicle and succeeded in reaching interplanetary space but the spacecraft failed early in transit. All six spacecraft in the second-generation 2MV Mars/Venus series, including the three Venus missions, fell victim to launch failures. Adding to the frustrations, the Americans had a successful flyby of Venus in 1962 with their Mariner 2. Undaunted, the Soviets improved the 2M V as the 3M V for the 1964 opportunities to Venus and Mars.

The initial flight tests of the new spacecraft were lost to launch vehicle failures, the first in November 1963 on a test to Mars distance and the second in February 1964 on a test to Venus distance. Despite of these losses, the Soviets proceeded with their 1964 program for Venus and Mars.

Spacecraft launched
First spacecraft:	3MV-1A No.4A
Mission Type:	Venus Spacecraft Test
Country; Builder:	USSR OKB-1
fjnmch Vehicle:	Molniya-M
Launch Date ‘: I ime:	February 19, 1964 at 05:47:40 UT (Baikonur)
Outcome:	Launch failure, third stage exploded.
Second spacecraft:	3MV-1 No.5 (Cosmos 27)
Mission Type:	Venus Atmosphere .’Surface Probe
Countryi Builder:	USSR OKB-1
Launch Vehicle:	Molniya-M
Launch Date! Time:	March 27, 1964 at 03:24:42 UT (Baikonur)
Outcome:	Failed to leave Harth orbit, fourth stage failure.
Third spacecraft:	Zond 1 (3MV-1 No.4)
Mission Type:	Venus Atmosphere; Surface Probe
Country і Builder:	USSR OKB-1
launch Vehicle:	Molniya-M
Launch Date: Time:	April 2, 1964 at 02:42:40 UT (Baikonur)
Mission End:	May 25, 1964
Encounter Date/Time:	July 19, 1964
Outcome:	Failed in transit, pressurization lost.

Launches to Venus were attempted on March 27 and April 2, straddling the ideal date. The first spacecraft was stranded in parking orbit but the second was deployed on a trajectory for Venus. Just as the launch vehicle was plagued by a troublesome fourth stage, the spacecraft had troublesome avionics systems. Immediately named Zond 1 when it became evident 3MV-1 No.4 would not be able to fulfill its mission at Venus, the spacecraft fell silent after less than 2 months in flight. Even if it had succeeded in deploying its entry probe, the decent capsule would not have survived to the surface, it was designed to survive only to 1TC and withstand pressures up to 5 bar. At the time there were two opposing theories for conditions on the surface of Venus. The high brightness temperatures measured by terrestrial radio telescopes, and confirmed by Mariner 2 during its flyby in 1962, could be interpreted either as a surface as hot as 400°C or as a hot ionosphere and cool surface. The easier design path was for the popular vision of a planet like Earth with a cool surface and maybe even an ocean. When the 1964 Venus launch window opened and Zond 1 set off, the controversy was not yet settled, although the weight of evidence was leaning to a hot surface. Radio observations of the planet from Earth later in 1964 would discredit the cool surface theory, but by then it was too late to redesign the Venus probes for the 1965 launches. The hot surface theory was firmly established after the flight of Venera 4 in 1967, which coincided with the highly successful Manner 5 llyby.

Spacecraft:

Currier spacecraft:

The 3MV Venus spacecraft were almost identical to their 3MV Mars counterparts, although the solar panels were less densely populated with solar cells. On the probe versions of the 2MV and 3MV, the entry vehicle was to be deployed just prior to the spacecraft entering the atmosphere and burning up. Instruments for interplanetary science and measurements in the near vicinity of the planet prior to destruction were carried on the main spacecraft. Communications from the probe would be direct to the Earth.

Figure 9,8 The 3MV-1 Venus probe spacecraft (courtesy Energiya Corp).

Launch mass: 800 kg (3MV-1A No.4A)

Launch mass: 948 kg (Cosmos 27 and Zond 1)

Entry system mass: 290 kg

Entry vehicle:

The 3MV entry probes were intended to obtain data during the descent through the atmosphere, survive impact, and return data from the surface. In the case of Venus the dense atmosphere that meant the probes would impact slowly enough to have a fair chance of surviving and operating for a short period of time on the surface. The 3MV probes were similar to the 2MV ones, being 90 cm in diameter and containing parachutes, batteries, sequencers, and two redundant 32 cm transmit­ters each with an antenna for direct communications to Earth, in addition to science instruments.

Figure 9.9 Zond 1.

Payload:

The payload for the test flight to Venus distance was probably similar to that of the lost test flight to Mars distance. The payloads for the missions to Venus launched on March 27 and April 2. 1964. were identical.

Zond l carrier spacecraft:

1. Radiation detector

2. Charged particle detector

3. Gas discharge and scintillation cosmic ray and gamma-ray detectors

4. Ion traps

5. Magnetometer

6. Micrometeoroid detector

7. Lyman-alpha atomic hydrogen detector

Zond 1 descent I landing capsule:

1. Temperature, pressure and density sensors

2. Atmospheric composition, acidity and electrical conductivity experiments

3. Gamma-ray surface composition detector and cosmic ray detector

4. Visible air glow photometer

5. Mercury level experiment

The atmospheric structure experiment consisted of two platinum wire resistance thermometers with ranges of -60 C to 460 C and 0 to 330 C, an aneroid barometer with a range of 0.13 to 6.9 bar. and a beta ray ionization chamber densitometer that was integrated with the thermometers and had a range of 0.0005 to 0.015 g/cc with a 5% error. The atmospheric composition, acidity and electrical experiments consisted of a set of gas analyzer cartridges wnth chemical and electrical tests for various gases including carbon dioxide, nitrogen, oxygen, and water vapor. The photometer was to search for airglow during the night landing. It was sensitive over the range 0.001 to 10,000 lux, and included the mercury level experiment to measure wave motion in a putative ocean. The anti-coincidence gas discharge and scintillation counter cosmic ray and gamma-ray detector was primarily to measure the surface composition of radioactive elements potassium, thorium and uranium from gamma-ray emissions on the surface, but it was also to be used during the interplanetary cruise to measure primary cosmic rays.

A micro-organism detector w as planned for the 3MV Venus and Mars landing capsules, but was never included in the payloads.

Mission description:

The first lest launch of this series failed when the third stage exploded. LOX leaking through a valve froze a fuel line which later broke. The loss of 3MV-1A No.4A so close to the imminent Venus launch window must have been disheartening, but the Soviets continued with preparations to launch the other two spacecraft.

The first attempt to launch the 3MV-1 No.5 spacecraft on March 1 was postponed owing to problems with the launcher during pre-launch tests. The second attempt on March 27 using the same vehicle failed when an electrical fault caused the fourth stage to lose attitude control and the engine did not restart for the escape burn. It was designated Cosmos 27 by the Soviets. This loss did have a very valuable result. For the first time a flight recorder had been added to the fourth stage telemetry system, and on its second pass the downlinked telemetry indicated a failure that was able to be traced to a generic problem in the I-100 control system circuitry. It required only 20 minutes of re-soldering to fix this for subsequent flights.

The third Venus spacecraft was dispatched successfully on April 2, 1964, but its initial trajectory was inaccurate and a midcourse maneuver was made the next day at a range of 564,000 km from Earth. It was the first midcourse maneuver carried out by a Soviet planetary spacecraft. Venera 1 and Mars 1 had both had this capability, but neither had been able to exercise it. However, Zond 1 w as in serious trouble. Л leak in the pressurized avionics section was detected right after launch due to a erack in the weld seam of the quartz dome for the Sun and star navigational sensors. The location of the leak w as determined from analysis of how’ the escaping gas perturbed the spacecraft. After a week, the transmitters and other electronics failed when they were switched on as the pressure fell to about 5 millibars, which permitted coronal discharges to short out power lines. The ion engines also failed their test, operating erratically. Owing to sensible backup design, communications w^ere able to continue using the entry system, and a second midcourse maneuver w7as made on May 14 at a distance of more than 13 million km from Earth. This resulted in a trajectory that would fly by the target at 100,000 km. In fact, the initial trajectory w’as probably so wide of the mark that even if the spacecraft had been fully functional, it would not have been able to adopt a collision course. Due to the pressure leak the Soviets did not reveal Venus as the target, merely announcing that the mission w7as a deep space engineering test, and named it Zond 1 rather than Venera 2. The leak was fatal, and on May 25 thermal control w? as lost and communications failed. The inert spacecraft passed Venus on July 19.

Results:

Zond 1 did return data on interplanetary plasma, including cosmic ray and Lyman – alpha measurements from the avionics module and proton measurements from the cosmic ray instrument in the lander capsule, but much of the data returned appears to have been lost.

The Venera 4, 5 and 6 missions were major successes for the Soviets, sending back detailed information while descending on their parachutes, as they were designed to, but they did not survive all the way to the surface. The controversy about the surface conditions in the wake of the Venera 4 and Mariner 5 missions was not resolved in time to permit modification of the Venera 5 and 6 probes to survive the temperature and pressure at the surface. It was decided to launch them anyway in order to obtain more information on atmospheric conditions, which they did well.

Spacecraft launched
First spacecraft:	Venera 7 (3V No.630)
Mission Type:	Venus Atmosphere/Surface Probe
Country і Builder:	USSR NPO-Lavochkin
Launch Vehicle:	Molniya-M
Launch Date ‘: 7 ‘ime:	August 17, 1970 at 05:38:22 UT (Baikonur)
Em ounier Da te l Time:	December 15, 1970
Outcome:	Successful, transmitted from surface.
Second spacecraft:	Cosmos 359 (3V No.631)
Mission Type:	Venus Atmosphere/Surface Probe
Country і Builder:	USSR/NPO-Lavochkin
Launch Vehicle:	Molniya-M
Launch Date: Time:	August 22, 1970 at 05:06:09 UT (Baikonur)
Outcome:	Failed to depart Karth orbit.

For the 1970 Venus opportunity the Soviets were determined to reach the surface. They now had unequivocal scientific proof that the pressure at the surface was about 100 bar and the temperature exceeded 450 C. But after years of controversy among scientists, the engineers were wary and they designed the new capsule to withstand 180 bar and a tempera lure of 540 C for 90 minutes. The additional mass required a significant reduction in the number of science instruments.

One mission was dispatched successfully but the other was a launcher failure. The entry capsule of Venera 7 was the first to reach the surface in an operational state and became the first successful planetary lander. Over the decade 1960 70 the Soviets had seventeen unsuccessful attempts at Venus missions, seven of which were intended to reach the surface. Now’ persistence had paid of!’. At the same time, their first Lunokhod rover was traversing the surface of the Moon.

Spacecraft:

Carrier spacecraft:

The carrier vehicle was unchanged from Venera 4, 5, and 6, but fewer instruments were carried to accommodate the larger mass of the entry system.

Entry vehicle:

The descent capsule was significantly modified as described above. In particular the entry system was made more egg-shaped to accommodate extra thermal insulation and a new’ shock absorber. A new’ spherical pressure vessel was used instead of the previous flat-capped hemisphere. The pressure vessel w as made of titanium, and to minimize weak points it had the fewest possible number of feed-throughs and welds. The temperature and pressure sensors w ere on the exterior of the shell, under the top hatch. The design was verified in a new test chamber at 150 bar and 540"C.

The goal was to reach the surface as fast as possible without losing the capsule in order to maximize its lifetime on the surface. A single parachute was used. On first opening, it was reefed by a cord wrapped around the shroud lines to limit its area to 1.8 square meters. Being smaller than the main parachute of previous missions, this would produce a faster descent. But the reefing cord was designed to melt at 200°C, deep in the atmosphere, opening the parachute to its full 2,5 square meters in order to achieve a soft landing. The parachute was built to survive the high temperatures at the surface. The capsule was pre-coolcd to -8°C prior to being released by the carrier spacecraft to maximize its survival time. The total design lifetime of the capsule was 90 minutes.

Launch mass: 1,180 kg (entry capsule 490 kg)

Figure 12.1 Venera 7 spacecraft.

Payload:

Carrier spacecraft:

1. Solar wind charged particle detector

Descentjlanding capsule:

1. Temperature, pressure and density sensors

2. Radio altimeter

3. Gamma-ray spectrometer

4. Doppler wind experiment

Due to the additional mass required to accommodate the very high pressure limit, the spacecraft had only the solar wind charged-particle detector and it was necessary to delete the atmospheric composition experiment and airglow photometer front the descent capsule. The aneroid barometer could measure pressures of 0.5 to 150 bar, and the resistance thermometer had a range of 25 to 540°C. Density was measured by an accelerometer during the entry phase. A gamma-ray instrument was added for measurement of surface rock type.

Mission description:

Venera 7 was dispatched successfully on August 17, 1970, and conducted midcourse maneuvers on October 2 and November 17. The second launch on August 22 failed

Figure 12.2 Venera 7 descent capsule.

to depart low Earth orbit when the fourth stage misfired. As a result of a sequencer problem and a power system failure the engine ignited late and shut down after only 25 seconds. This vehicle was designated Cosmos 359 by the Soviets and re-entered on November 6. Venera 7 initiated its planetary encounter activities on December 12 when the capsule batteries were charged by the solar panels. The internal equipment compartment was activated and chilled down to -8°C. The capsule was released at 04:58:44 UT on December 15. It struck the atmosphere at an altitude of 135 km at

11.5 km/s. By the time it was down to 54 km it had been slowed to 200 m/s, and a pressure reading of 0.7 bar triggered the deployment of the parachute just above the cloud layer. The capsule transmitted for 35 minutes during its descent in darkness. It survived the impact at 05:34:10 UT and a weak signal was received for another 23 minutes. The landing site was at 5°S 351 °E, where ir was 4:42 Venus solar time and the solar zenith angle was 117 degrees.

■ Parachute Deployed Figure 12.4 Doppler frequency plot for the Venera 7 descent capsule (from Don Mitchell).

This success was not immediately obvious, and it was originally thought that the capsule had failed to reach the surface. The 35 minute descent to the surface turned out to be a wild ride. After the first 13 minutes the reefing cord melted away and the parachute opened fully, just as it was meant to do. Six minutes later the parachute ripped, and over the next several minutes the descent rate increased and the capsule oscillated wildly as the rip extended. A few minutes before reaching the surface the parachute failed and the capsule fell freely. All of this was evident from the Doppler shift in the transmitter carrier frequency. The capsule hit the surface at 16.5 m/s, the signal disappeared into noise, and it was concluded that the capsule must have been destroyed. There was no immediate announcement of the success of Venera 7. After the New Year holiday, an expert in signal processing re-ran the data tapes and in all the noise found the barely perceptible signal from the capsule on surface. The signal strength had reduced to 3% at impact, returned to full strength for one second, then dropped back to 3% for the next 23 minutes before terminating. The capsule would seem to have bounced on impact and come to rest tilted at about 50 degrees to the vertical, aiming the radiation pattern of its antenna well off Earth and resulting in a very low received power. The team members were elated by this discovery. It may not have been a graceful touchdown, but it was another first for the Soviet planetary program.

The Venera 7 carrier spacecraft returned measurements on the upper atmosphere and ionosphere prior to breaking up in the atmosphere.

Results:

The Venera 7 probe measured the temperature of the atmosphere from an altitude of 55 km to the surface, but a commutator failure resulted in no pressure or altimetry information being transmitted. Initially il was thought the temperature measure­ments returned were internal but fortunately they did turn out to be atmospheric, and when combined with Doppler data and with thermodynamic and aerodynamic modeling it was possible to construct a profile of temperature and pressure down to

the surface. Altitude profiles of; horizontal wind speed and direction were also obtained from the Doppler data and aerodynamic modeling. There was a fast wind at high altitudes in the same retrograde direction as the axial rotation. This con finned astronomical evidence from ultraviolet cloud flow that the upper atmosphere was ‘super-rotating*. Wind speeds of less than 2.5 m s were measured at the surface. The probe temperature sensor oscillated between binary readings of 457 C and 474 C on the surface. The computed surface pressure was 92 bar. Doppler data at the moment of touchdown, plus the fact that the capsule survived the high speed impact, implied that the surface was harder than sand but no harder than pumice. No surface composition measurements were returned because of the stuck commutator.

Campaign objectives:

For the Soviets this campaign combined a Venus flyby/entry mission with a flyby of Comet Halley, and it was their first (and thus far only) multiple-target mission. After releasing their entry systems at Venus in June 1985 the two flyby spacecraft were to be re-targeted by the gravity-assist of their encounter with the planet onto a course to intercept Comet Halley in March 1986.

In addition to a lander, the entry system carried an atmospheric balloon. The idea to float a balloon in the atmosphere of Venus grcwr from French-Soviet cooperation initiated after the successful Venera 4 mission in 1967. France and the Soviet Union had come to a rapprochement of sorts in the Cold War, opening a breach in the Iron Curtain by establishing cooperation in space science. In 1974 Dr. Jacques Blamont of CNES and Boris PeLrov, Chairman of the Intercosmos Council, began to discuss a joint mission consisting of an entry probe to deliver a large French balloon into the atmosphere of Venus, and a Soviet orbiter to provide the communications relay. By 1977 a date had been tentatively set for a 1984 launch of the ‘Venera-84’ mission to mark the bicentennial of the Montgolfier brothers’ invention of the hot-air balloon, and the division of work had been established. Jacques Blamont and Mikhail Marov were named as science co-chairs for the mission. The French would supply the two 10 meter diameter balloons with their 50 kg gondolas, including transponders for very long baseline interferometry (VLB!) tracking, and the Soviets would supply the spacecraft, entry systems, and the remaining mission support. But events changed these plans.

In the late 1970s the world’s space science community was beginning to plan for the eagerly awaited apparition of Comet Halley in 1986. The IJS offered to carry a French ultraviolet instrument on one of its spacecraft. When the US withdrew from this effort in 1979 the Soviets offered to fly the French instrument on Venera-84 to enable it to observe the comet from Venus orbit – which would be a more favorable vantage point because although the comet would approach that planet no closer than 40 million km. that was much closer than it would approach Harth. In the process of investigating how to improve observations of Halley from Venus the Soviets found that it would be possible to utilize a gravity-assist during a flyby of Venus to set up an encounter with Halley. The science value of a mission to both Venus and Halley as argued by Jacques Blamont intrigued Roald Sagdeev, Director of 1K1, who set out to have it supersede the Venera-84 mission. The new project w as called ‘Vega’ as a Russian contraction of’Venera’ and ’Galley’, with the name of the comet using a G’ because there is no H’ in the Cyrillic alphabet. Valery Barsukov. Director of the Vernadsky Institute, w as far more interested in Venus than he w as in the comet, but Sagdeev sold the mission to him by including a lander, albeit at the cost of reducing the size of the balloon package to enable both to fit inside the standard entry system. Three years of intensive development of the Venera-84 mission, including partially manufactured hardware, was lost. When the furious French declined to participate further, the small balloon became a Russian project. Nevertheless, Sagdeev managed to coax the French into providing several instruments for the lander and balloon, as well as two key remote sensing instruments for the Halley encounter. And by taking advantage of their bridging position between the East and the West, the French were able to gain the participation of the Deep Space Netw ork in the VLB1 network that would measure the dynamics of the balloons as they drifted in the atmosphere of Venus. For the first time, therefore, the arch rival Americans became a participant in a Soviet planetary mission, albeit by providing tracking resources. The University of Chicago supplied an instrument to investigate dust particles during the Halley flyby, but this was arranged through the science community as a private venture rather than at government level and the principal investigator had to assure the US military that he was using only commercial parts from his local Radio Shack store! He dismissed the military’s concerns with. "Let them [the Soviets] copy this, it will set them back years.’

Sagdeev, by enthusiasm, energy, and personal effort, instituted the new project as a broadly international venture by off ering 120 kg on the spacecraft for instruments originating from countries outside the USSR. This extensive internationalization was unprecedented for the historically closed Soviet space program. And internally the perestroika initiative enabled him to overcome resistance by the Soviet bureaucracy.

But the final credit must go to Chief Designer Vyacheslav Kovtunenko and the NPO-Lavochkin scientists and engineers wTo, by building the most comprehensive and successful deep space mission in their history, created a legacy for Soviet lunar and planetary exploration.

Vega 1 (5VK No. 901)

Venus Flyby/Lander/Balloon and Halley Flyby USSR, NPO-Lavochkin ’ ‘ ‘

Proton-K

December 15, 1984 at 09:16:24 UT (Baikonur) June 11, 1985 March 6, 1986 Successful.

Vega 2 (5VK No.902)

Venus Flyby/Lander/Balloon and Halley Flyby USSR, NPO-Lavochkin ’ ’ ‘

Proton-K

December 21, 1984 at 09:13:52 UT (Baikonur) June 15, 1985 March 9, 1986 Successful.

The Vega missions became an integral part of the International Halley Mission (IHM) organized initially by the European Space Agency to coordinate operations and data analysis for the various Halley missions being planned by Europe, Japan, the IJS and the Soviet Union. An Interagency Consultative Group consisting of high level representatives of the space agencies overseeing the IIIM provided cover for US participation in the midst of the Cold War, effectively ci re um venting the absence of a formal agreement between the US and USSR. Ironically spacecraft were sent to Halley by all these nations except the US, whose formal involvement was ultimately limited to providing tracking and science support.

With lander, balloon, and flyby components the Vega missions were both very ambitious, and by involving a host of international interfaces including a large array of international instruments were extraordinarily complex. The nations participating included Austria, Bulgaria, Czechoslovakia, East Germany, France, Hungary. West Germany, Poland, and the United States. The Hungarians built part of the navigation system and the Czechs supplied the optical system for the automated scan platform. Foreign investigators were allowed into the country to participate fully and actively in the project from beginning to end; not passively as previously by delivering their completed instruments in advance and waiting at home to find out what happened to them. Team meetings were held in the USSR and foreign contributors were allowed into Soviet facilities for development, testing and integration activities. This style of cooperation w ith the USSR w7as unprecedented. An organization called Intercosmos had existed since the 1960s for coordination of cooperation in space research mainly among Eastern Bloc nations and with France, but this was the first time the activity assumed such a large scale and included Western nations to such a degree.

The 1984 launches gave the Soviets enormous influence in the international space
community. With such a bold move to internationalization, leadership in planetary exploration passed to the USSR. After the busy era of Mariner, Pioneer. Viking and Voyager launches in the 1970s, the US launch rate had fallen precipitously to zero in the 1980s. The USSR continued to reap a harvest from its Venera series, and began its transition from a closed program to an open program far more international than any flight project in the US. The Soviets now issued open calls for participation in its science missions. US science missions would not become more international than ”participation by invitation onlv*

The Vega missions were highly successful in meeting all their science objectives, and a major achievement for the Soviet robotic lunar and planetary program. They concluded the run of ten consecutive highly successful heavy-class Venera missions that started with Venera 9 in 1975. and they were the final Soviet missions to Venus after twenty-nine launch attempts since 1961. During this 24 year period only three of sixteen windows for Venus were not used. Nineteen of the twenty-nine launches sent spacecraft on trajectories to Venus, of w hich fifteen successfully delivered three entry probes, ten landers, tw’o balloons, and four or biters. The Soviet scientists and engineers participating in the Vega missions would have dismissed as ridiculous the prospect of there being only two more campaigns in the Soviet planetary exploration program, both of w’hich would be embarrassing failures.

Spacecraft:

Flyby spacecraft:

The flyby spacecraft was nearly identical to Venera 9 to 14 but used the larger solar panels of Venera 15 and 16 to handle the pow er demand and w^as loaded w ith 590 kg of propellant instead of the usual 245 kg. It w’as protected from hypervelocity comet dust impacts by an aluminum shield consisting of an outer multi-layer sheet of 100 micrometers thickness mounted at a standoff distance of 20 to 30 cm.

A data rate of 65 к bits/s was provided for the comet encounter, but a slower mode would be used in the cruise phase. Approximately half of the spacecraft w as devoted to the Halley science instruments and half to the Venus entry system. In making the flyby of Venus in the manner required to set up the Halley encounter, the spacecraft would relay to Earth the transmission from the lander during its descent and surface operations as previously. However, the balloon would transmit its telemetry directly to Earth.

The spacecraft was fitted with an 82 kg articulated scan platform that could rotate from -147 to +126 degrees in azimuth and from -60 to +20 degrees in declination for a pointing accuracy of 5 minutes of arc and a stability of 1 minute of arc per second. Its automated tracking would enable instruments to be continuously pointed at the nucleus of the comet during the rapid flyby while the spacecraft held an orientation that permitted its high gain antenna to point at Earth for real-time transmission. The pointing was controlled either by an eight-element photometer or by using the wide angle camera, and gyroscopic attitude control was provided as a

Figure 18.1 Vega spacecraft (courtesy NPO-Lavochkin). Scan platform folded on left, parabolic antenna on the right, toroidal instrument compartment on the bottom with external instruments.

Figure 18.2 Museum model Vega spacecraft without insulation and dust shields. Front side at right shows solar panels, parabolic antenna, and navigation instruments. Back side at left show’s camera platform hanging down below toroidal instrument section, radiator panels and black disks where helical lander relay antennas were mounted.

precaution against comet dust upsetting the optical sensors. The scan platform carried the narrow and wide angle cameras, an infrared sounder, and a three – channel spectrometer. All other experiments were body-mounted except for two magnetometer sensors and various plasma probes and plasma wave analyzers which were mounted on a 5 meter boom. The total science payload for Halley weighed 130 kg.

Figure 18.3 Vega 1 folded and ready to launch. Note scan platform, insulation and metal shielding.

Entry system:

The entry system was virtually identical to the recent Venera missions, consisting of an insulated sphere 2.4 meters in diameter whose upper and a lower hemispheres were joined non-hermetically. In this case, however, the lander was installed in the lower half and the balloon in the upper half.

Figure 18.4 Entry capsule cross-section (by James Garry): 1. Antenna; 2. Balloon compartment; 3. Helium inflation tank; 4. Lander aerodynamic stabilizer; 5. Gas chromatograph; 6. Spectrophotometer; 7. Entry heat shield; 8. Thermal insulation; 9. Oscillation damper; 10. Battery; 11. Stabilizing vanes; 12. Crushable impact torus; 13. Drill and sample collector; 14. Coolant delivery piping: 15. Balloon aerobrake; 16. Science instrument bay; 17 Parachute.

Lander:

The Vega landers were almost identical to the Venera 13 and 14 landers with some aerodynamic modifications for increased stability while free falling. These included spoke-like blades interior to the landing ring to reduce spinning and a thin collar-like sleeve installed beneath the disk of the aerobrake to minimize the turbulence which would be induced by the externally mounted instruments.

Figure 18.5 (left) show’s the sleeve and the blades. Tn view on the landing ring are the two white hygrometer compartments, the temperature and pressure unit offset to its right, and also the drill. Figure 18.5 (right) shows the large shiny cylindrical gas chromatograph on the ring to the left, the horizontal drill vacuum reservoir, and the penetrometer and the hydrometers on the far right. The impact velocity of 8 m/s was to be cushioned by the shock absorbers that support the main spherical pressurized compartment.

Figure 18.5 Venera 13 and Venera 14 landers during tests at Lavochkin

Strap to balloon

Semi-Directional / Antenna

Straps

Radio Transmitter and Electronics

Temperature Sensors

Photometer Pressure Sensor Lithium Batten" „ Nephelometer Windows

r-3 ——— lb кґ 4 ,—^

Anemometer

Figure 18.6 Gondola diagram (from Don Mitchell) and testing on a short tether.

Balloon:

The balloons were a new component, and were to be carried in and deployed by the upper hemisphere of the entry system. The super-pressure helium aerostat with its attached gondola w as designed to float in the middle layer of cloud at an altitude of 54 km, where the temperature was a mild 32’C and the pressure was 5.35 millibars.

Each radio-transparent balloon had a mass of 11.7 kg and when inflated it wfas 3.4 meters in diameter and held 19.4 cubic meters of helium that had a mass of 2 kg. A 13 meter long tether suspended the 7.0 kg gondola (including 1.6 kg for the tether). The entire system weighed a little over 20.7 kg. The rate of helium diffusion w’as sufficiently low to sustain pressure for about 5 days.

The 1.2 meter long 14 cm wide gondola contained a transmitter with a stabilized oscillator for Doppler tracking, a conical antenna, a vertical anemometer, sensors for ambient temperature and pressure, a light photometer, a nephelometer. a eontrol and ballast system, and sixteen lithium batteries for 300 wratt-hours of power. The 1 kg battery package w as designed for 46 to 52 hours of life. To simplify the task for the network of radio telescopes wdiich would track the balloons, the 4.5 W transmitter operated in the 18 cm astronomical band at 1.6679 GHz. It transmitted direct to Earth via the conical antenna at either 1 or 4 к bits, s. Except for the lightning counter which was sampled every 10 minutes and the photometer twice every 30 minutes, all the other instruments were sampled once every 75 seconds. The data w as stored on a 1,024 bit memory. A 5.5 minute burst of data was sent to Earth every 30 minutes, alternating between two transmission inodes in a predetermined sequence. In the first mode, 852 bits of data collected from the instruments were transmitted in a 270 second burst preceded and followed by 30 seconds of carrier for VLB1 velocity measurements. In the alternative 330 second mode, only two tones were transmitted for VLBI position and velocity.

The balloon system had to be folded up during cruise and entry, survive the forces of deployment, and then withstand the corrosive atmosphere of sulfuric acid aerosol. The envelopes were made using a woven teflon and cloth matrix, the gondola was covered with a white paint resistant to sulfuric acid, and the tethers were made of a type of nylon. Timing, as determined by pressure sensors, was critical to successful deployment: if the envelope were inflated at too high an altitude it would burst in the low pressure; if it were inflated at too low’ an altitude it w ould not gain the necessary buoyancy, would penetrate too deep and be destroyed by the high temperature. The inflation system had 2 kg of helium, and altitude control would be by the release of ballast.

The balloon system was carried in the upper hemisphere of the entry system, in a toroidal canister that surrounded the helical antenna of the lander. In addition to the folded balloon and gondola, this canister contained a 35 square meter parachute and the spheres of pressurized helium to inflate the balloon. The deployment began at an altitude of 64 km by separating the hemispheres while on the drogue parachute. This released the lower hemisphere containing the lander. Separation deployed a braking parachute for the lander, which then performed its own deployment sequence as on previous missions. The upper hemisphere then released the toroidal balloon package

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Scientific instruments

Navigation sensors

Scan platforms

Oust shield

Scientific instruments

Figure 18.8 The Vega spacecraft configured for Halley encounter (courtesy NPO – Lavochkm).

at 62 km, deploying its parachute in the process. At 57 km the package deployed the balloon system. At 55 km the inflation system was activated. By the 53 km level the envelope had inflated and the package on the parachute released the balloon system. At 50 km the balloon system released its ballast, deployed the boom that carried the temperature sensors and anemometer, and then rose to 54 km to travel wherever the prevailing wind took it. Since the temperatures in this altitude range w’ere benign there was no requirement for thermal control.

On Earth, a global distribution of twenty international antennas consisting of two networks was ready to perform Doppler tracking and receive the scientific data from the balloons – one at a time, as they were to arrive at the planel several days apart. One network was led by 1K1 and used six Soviet antennas including a new 70 meter dish that was built for the Vega missions. The second network was led by CNES and used the three 64 meter antennas of the Deep Space Network in the US, Australia and Spain, and astronomical antennas in Brazil, Canada, England, Germany, Puerto

Rico, South Africa, and Sweden. Doppler tracking by each antenna gave the range and velocity along the Earth-Vcnus line, but lateral motion of the balloons required interferometry that combined phase information from antennas located far apart and linked electronically. In addition, the network simultaneously tracked carrier wave signals provided by one or other of the two flyby spacecraft to provide a third leg to greatly increase the precision of distance and velocity measurements in a differential interferometry technique developed by the IJS for the Apollo lunar missions.

Vega spacecraft system mass

Payload:

flyby spacecraft:

Mounted on the scan platform:

1. TV imaging system (TVS, USSR-France-Hungary)

2. Three-channel (ultraviolet, visible and near-infrared) spectrometer (IKS. France-USSR-Bulgaria)

3. Infrared spectrometer (IKS, France)

Body mounted:

1. Dust mass spectrometer (PUMA, FRG-USSR-France)

2. Dust particle counter (SP-1)

3. Dust particle counter (SP-2)

4. Dust particle detector (DUCMA, USA)

5. Dust particle detector (FOTON)

6. Neutral gas mass spectrometer (ING, FRG)

7. Plasma spectrometer (PLASMAG)

8. Energetic particle analyzer (TUNDE-M. Hungary-USSR-FRG-ESA)

9. Energetic particles (MSU-TASPD)

10. Magnetometer (MISCHA, Austria)

11. Low frequency wave and plasma analyzer (APV-N, USSR-Poland-Czecho – slovakia)

12. High frequency wave and plasma analyzer (APV-V, USSR-France-ESA)

Three instruments for remote sensing of Halley were mounted on the ASP-G scan platform; the 32 kg TVS camera, the 14 kg TKS three-channel spectrometer, and the

18 kg IKS far-infrared spectrometer. The camera was Russian and the spectrometers were provided by France. The far-infrared spectrometer was cryogenically cooled by a Joule-Thompson cryostat and operated in the range 2.5 to 12.0 microns. The three – channel instrument operated in the ultraviolet 120 to 290 nm, visible 275 to 715 nm, and near-infrared 950 to 1,200 nm. The flyby range at Ilailey was deliberately large to avoid damaging to the spacecraft, so to obtain the desired view of the nucleus the camera required a narrow angle optical system capable of a resolution of 150 meters at a range of 10,000 km. The computers for the science instruments were enhanced by using Western electronics. Hoivever. because CCD technology was restricted the Soviets had to develop their ота 512 x 512 device for the camera. The optics were built by the French, and comprised a 150 mm f/3 wide angle lens that was limited to the red, and a 1,200 mm f/6.5 narrow angle lens with six filters from the visible to infrared. The Hungarians were responsible for the camera electronics with assistance from the Soviets.

The five instruments to study the dust issued by the nucleus of the comet were the

19 kg PUMA dust particle impact mass spectrometer to measure the composition of

A

individual dust particles, the 2 kg SP-1, 4 kg SP-2, and 3 kg DUCMA dust particle counters to determine the flux and mass distribution of dust particle in different si/e ranges, and the FOTON dust particle detector that was installed to measure the large particles that punched through the standoff shield. In-situ measurements by the 7 kg INC neutral gas mass spectrometer would analyse gas in the space through which the spacecraft was traveling. The composition and energy spectrum of ions would be determined by the 9 kg PLASMAG plasma spectrometer, and the flux and energy of ions would be measured by the 5 kg TUNDE-M energetic particle analyzer. It also had the 4 kg MISCHA magnetometer, two plasma w*ave analyzers, the 5 kg APV-N for ion flux and frequencies below 1 kHz, and the 3 kg APV-V for plasma density, temperature and frequencies in the range 0 to 300 kHz.

Including the scan platform and its supporting structure, the instrument payload of the flyby spacecraft was 253 kg.

Lander:

Entry and descent:

1. Temperature, pressure and wind sensors (МЕТЕО, USSR-France)

2. Hydrometer for water vapor concentrations ( VM-4)

3. Ultraviolet spectrometer for atmospheric SO2 and sulfur measurements (ISAV-S)

4. Optical nephelometer-seatterometer for aerosol size and properties (ISAV-A)

5. Particle si/e spectrometer for aerosols (LSA)

6. X-ray fluorescence spectrometer for aerosol elemental analysis (IFP)

7. Gas chromatograph for aerosol chemical analysis (SIGMA-3)

8. Mass spectrometer for aerosol chemical analysis (MALAKIIIT-V. USSR – France)

9. Doppler experiment for wind and turbulence

The deseent instruments focused on aerosols in particular. There were two particle size instruments for measuring the physical properties of aerosols, two instruments for aerosol chemistry, and one instrument for an elemental analysis of the aerosols. These five instruments had externally mounted components with limited insulation from the ambient temperature and pressure, but since the aerosols were confined to the upper atmosphere they were required to function only above 35 km. The aerosols were carried into the instruments by inlet tubes. Some instruments analyzed the light scattered by the aerosol partieles in these tubes to determine their si/e. The ISAV-A instrument also included a nephelometer to determine the cloud density by shining a beam of light through a window’ in the pressure vessel and measuring the light returned through this window. It shared electronics with the ultraviolet spectro­meter.

The gas chromatograph instrument was specifically designed for Vega to measure sulfuric acid aerosol by trapping the droplets in a carbon saturated filter that reacted with sulfuric acid to produce sulfur dioxide and carbon dioxide.

The x-ray spectrometer was a significant improvement on the ones carried by the

Venera 13 and 14 landers. It distinguished grain size using laser imaging. The mass spectrometer sampling system used an aerodynamic inertial separator to segregate grains into small and large sizes on two separate filters. These were then vaporized and analyzed in the mass spectrometer.

The ultraviolet spectrometer was an active experiment, particularly effective for a descent in darkness. It had an ultraviolet lamp and a 1.7 meter path length absorption cell into which the atmosphere was admitted in order to measure the absorption at 512 points between 230 and 400 nm. The objective was to determine the nature of the mysterious ultraviolet absorber’ deduced from remote sensing measurements. The spectrometer was inside the lander, but there was a pipe through the hull to allow the atmosphere into the instrument. It was operated from 62.5 km down to the surface.

The temperature and pressure instruments were similar to those of the Venera 13 and 14 landers but revised for greater accuracy. They comprised two platinum wire thermometers and three pressure sensors covering the ranges 0 to 2, 0 to 20 and 2 to 110 bar. The hydrometer was also improved.

Surface:

1. Drill and surface sampler (SSCA)

2. X-ray fluorescence speciromcter (BDRP)

3. Gamma-ray spectrometer (GS-15STsV)

4. Dynamic penetrometer (PrOP-V)

As both the gravity-assist to deflect the flyby trajectory for Halley and the mission of the balloons required the Vega entries to occur on the night-side, the landers were not given cameras or optical instruments, and those instruments they did carry w’ere similar to those utilized previously. I he Vega landers were focused mainly on solving mysteries about the atmosphere and rectifying problems with instruments on previous missions that were caused by the hostile atmosphere. The gamma-ray soil spectrometer had been deleted after Venera 9 and 10 in favor of the combined drill and x-ray fluorescence spectrometer; this time they were all carried. And since there were no imagers the penetrometer was upgraded to provide an electrical readout.

Lander instrument mass 117 kg.

Balloon:

1. Temperature and pressure sensors (‘USSR-France)

2. Vertical wind velocity anemometer

3. Nephelometer for density and particle size of local aerosols (USA)

4. Light level photometer and lightning detector

5. Stable oscillator for VLBI measurements

A boom w as deployed from the side of the gondola to expose sensors. One w as a propeller anemometer. It measured vertical winds as fast as 2.0 m/s. The horizontal winds were measured by VLBI analysis of radio tracking. The ambient temperature

was measured by two thin-film resistance thermometers with a range of 0 to 7Cf’C and an accuracy of 0.5°C mounted at separate positions on the boom. Pressure was measured by a vibrating quartz beam sensor with a range of 0.2 to 1.5 bar and an accuracy of 0.25 millibar. The photometer consisted of a silicon PIN diode sensitive in the 400 to 1,100 nm range with a 60 degree field of view at the nadir. It was also designed to detect lightning by counting short bursts of abnormally bright intensity. The ncphelomcter was a simple backscatter instrument similar to those of previous missions.

Mission description:

Flyby spacecraft:

In keeping with the international nature of the project. Westerners were allowed to visit Baikonur and view the launches of Vega 1 on December 15. 1984, and Vega 2 on December 21. This was also the first time that Soviet television showed a Proton launch. And although the US routinely tracked Soviet spacecraft, this w’as the first time that this was done officially. The announcement that an American instrument was onboard prompted a small furor in the US. One of the booms for the plasma wave experiment initially failed to deploy on each spacecraft, but these both sprung out after the first midcourse maneuver.

Vega 1 arrived at Venus in early June 1985, only weeks after Venera 16 had been switched off. The spacecraft released their entry systems 2 days out from the planet,

on June 9 for Vega 1 and on June 14 for Vega 2. The points at which they were to enter the atmosphere were on the night-side in order to enable the spacecraft to head for Halley and to maximize the cruising lifetime of the balloons before they suffered solar heating. After releasing its entry system Vega 1 maneuvered to pass the planet at a range of 39,000 km for the gravitv-assist to llallev and to relay the data from its lander. Vega 2 did likewise at a range of 24.500 km. Each spacecraft turned off-Sun to receive the transmission from its lander at a rate of 33)72 hits/s in the meter band and to relay it to Earth in the centimeter and decimeter bands. They did not conduct any science observations at Venus. On finishing the relay, each spacecraft resumed cruise operations. The gravity-assist of the flyby did most of the work in deflecting the path of each spacecraft toward Halley, but maneuvers were needed to refine the final approach.

Vega 1 flew past the nucleus of the comet at a range of 8.890 km on March 6. 1986. and Vega 2 did likewise at a range of 8,030 km on March 9. Both made highly successful scientific measurements. Two Japanese spacecraft had been observing the comet at extreme distance and Europe’s Giotto was scheduled to arrive on March 13 for a daring close flyby at a range of only 500 km. By combining tracking data with imaging, the Vega spacecraft gave a more precise position for Halley in space than was possible using terrestrial telescopes. This was used to improve the accuracy of Giotto’s terminal maneuvers, both to reduce the targeting error in order to obtain the intended observations and to reduce the potential risk to that spacecraft. Both Vega spacecraft flew through the tail of the comet and were pummeled by small grains impacting at 80 km s. The shields installed on one side of each vehicle protected it from damage. The solar panels suffered both dust impacts and electrical discharges induced by the comet plasma. Vega 1 lost 40% of its power supply and Vega 2 lost 80%. After a circuit around the Sun, both spacecraft passed through the tail again in 1987. providing further data. Vega 1 ran out of attitude control gas on January 30 of that year and then on March 24 contact with Vega 2 was discontinued.

Entry system:

The Vega 1 capsule entered the night-side atmosphere at 01:59:49 UT on June 11. 1985, at a speed of 10.75 kin/s and at an angle of 17.5 degrees. The Vega 2 capsule entered at 01:59:30 UT on June 15, at 10.80 km/s and at an angle of 18.13 degrees. The pilot parachutes were deployed at an altitude of 65 km. Eleven seconds later, at

64.5 km, the capsules split into hemispheres and the pilot parachutes drew the upper hemispheres containing the balloon systems away, in the process deploying the main parachutes of the landers in the lower hemispheres. Hour seconds later, at 64.2 km. the landers shed their hemispheres. Having slowly descended to 47 km. each lander released its parachute in order to free-fall to the surface. The new aerodynamic drag devices successfully reduced both vibration and spin, thereby increasing the stability of the descending landers.

Meanwhile, the balloon packages were released from their hemispheres at 62 km. in the process deploying the pilot parachute of each balloon package. At 57 km the

main parachute was deployed. At 55 km the inflation of the envelope was initiated. With the balloon fully inflated, the main parachute was released at 53 km. At 50 km the inflation system and ballast was released and the balloon system rose to 54 km in the middle of the cloud layers with its gondola deployed to make measurements.

Landers:

The Vega 1 lander settled at 7.11 N 177.48 E. just north of eastern Aphrodite Terra and 0.6 + 0.1 km below the planetary mean radius. It was 03:02:54 IJT on 11 June. 0:24 local time, and the solar zenith angle was 169.3 degrees. The measured surface temperature was 467’C and the pressure was 97 bar. The transmission was curtailed 20 minutes after landing in order to conserve energy on the flyby spacecraft, which was not facing its solar panels at the Sun. and to ensure readiness for the subsequent Halley trajectory maneuver.

At an altitude of 17 km the Vega 1 lander experienced electrical spikes and the Doppler tracking data showed violent upward excursions. This shock triggered the accelerometer that was to indicate contact with the ground, causing a premature start to the surface activity sequence, including deployment and operation of the drill and x-ray spectrometer. As the x-ray soil analysis instrument had failed its pre­launch tests and been flown regardless, this may not have mattered. Venera 11 to 14 and the four IJS Pioneer probes also experienced electrical anomalies in the altitude region 12 to 18 km, but Venera 9 and 10 and Vega 2 did not. The cause of these anomalies remains unknown.

The Vega 2 lander touched down at 7.52°S 179.4 E, 1,300 to 1,500 km southeast of Vega 1 and 0.1 +0.1 km above the planetary mean radius. It was 03:00:50 UT on 15 June, 1:01 local time, and the solar zenith angle was 164.5 degrees. The surface temperature was 462°C and the pressure was 90 bar. The transmission was truncated 22 minutes after landing to preserve energy on the flyby spacecraft. There were no anomalies during the descent and the surface operations were performed nominally.

Balloons:

The Vega balloons were both successfully deployed at the anti-solar point (i. e. local midnight) and drifted with the wind at an altitude of about 53 km where the pressure was about 0.5 bar, right in the middle of the three cloud layers. They were carried longitudinally by zonal winds through the night-side atmosphere for 30 hours before crossing the dawn terminator. No latitude measurements could be made, and it was assumed that the balloons remained at a constant latitude. 8°K in the case of Vega 1 and 7 S for Vega 2. Each balloon transmitted for 46.5 hours until its batteries were exhausted. Loss of signal occurred in the early morning hours on Venus after having traveled some 10.000 km. about one-third the way around the planet. The balloons continued silently into the day-side where they would eventually have succumbed to solar heating and burst their envelopes.

Results at Venus:

Landers on descent:

A telemetry problem prevented Vega 1 temperatures from being transmitted during the descent, but the Vega 2 data indicated the presence of a sharp thermal inversion that reached a minimum temperature of -20°C at an altitude of 62 km. The optical spectrometers operated between 63 and 30 km and reported an atmospheric structure similar to that seen by earlier landers and confirming a three layer cloud deck. But on this mission, as for Venera 8, no sharp lower cloud boundary was observed. Aerosol particle size measurements were taken down to 47 km, and were in general agreement with earlier Soviet results and the data from the Pioneer entry probes and confirmed that there were at least two layers of differing particle sizes. The measurements from Vega 1 and 2 were highly consistent, indicating the cloud layers to be very similar at their entry points except in the uppermost layer where Vega 2 found less dense aerosols than Vega 1. The smallest ‘mode Г particles were speculated to be aluminum and/or ferric chloride. About 80% of the larger ‘mode T particles were shown to be spherical with a refractive index of 1.4, a characteristic consistent with sulfuric acid, wdiile the remaining 20% had a refractive index of 1.7, suggestive of solid sulfur, The highest particle counts wrere in the altitude range 58 to 50 km. The Vega instruments were insensitive to the largest ‘mode 3’ particles reported by Pioneer probes.

The Vega 1 and 2 gas chromatographs and the Vega 1 mass spectrometer were the first to make an in-situ detection of sulfuric acid, confirming remote sensing results and yielding a density for the altitude range 63 to 48 km of about 1 milligram of sulfuric acid per cubic meter. The Vega 1 mass spectrometer heavy particle sample contained sulfur trioxide (sulfuric acid anhydride) and chlorine. Unfortunately, the Vega 2 mass spectrometer failed. The x-ray fluorescence spectrometer on Vega 2 detected sulfur (—1.5 mg/m3), chlorine ( — 1.5 mg/m3), and iron (0.2 + 0.1 mg/’mJ). It also made the first detection of phosphorus (—6 mg/m3), this possibly in the form of phosphoric acid, and explaining the persistence of a small amount of aerosol in the sub-cloud region with a base at 33 km. Iron was also reported by the x-ray

Figure 18.11 Chlorine, sulfur and phosphorus profiles from the descent x-ray aerosol analyzer (from Don Mitchell).

analysis, perhaps as ferric chloride in the aerosols. The Vega 1 x-ray fluorescence instrument failed. The ultraviolet speetrometcrs gave vertical profiles for sulfur dioxide mixing ratios with upper region abundances in general agreement with remote sensing and other sources, and generally deer easing towards /его at the surface. The possibility of elemental sulfur vapor was also noted. Solar ultraviolet was completely absorbed below an altitude of 10 km. although this was probably due to aerosols coating the instrument. The hydrometer reported a water vapor abundance of 0.15% at high altitudes (60 to 55 km) decreasing by a factor of ten at lower altitudes (30 to 25 km). The fact that this large abundance is inconsistent with other measures may indicate that the instrument was confused by other atmospheric constituents. The water vapor profile on Venus remains poorly determined.

Landers on the surface:

The Vega 1 lander conducted a gamma-ray soil analysis but the drill had failed and so no x-ray soil analysis could be performed. The Vega 2 gamma-ray spectrometer, drill, and x-ray fluorescence experiments all worked well.

X-ray fluorescence results from Vega 2 (as oxides):

silicon	47%
titanium	0.2%
aluminum	16%
iron	8.5%
manganese	0.14%
magnesium	11%
calcium	7.3%
potassium	0.1%
sulfur	4.7%
chlorine	<0.3%

These analyses showed rocks poor in iron and magnesium but rich in silicon and aluminum, indicating a composition similar to lunar highland rocks. The fairly high sulfur abundance may be an indicator of older rocks.

Gamma-г а г results:

The potassium, uranium and thorium values were very similar to Venera 9 and 10. in contrast to the Venera 8 results that showed significantly higher concentrations of all three elements.

Balloons:

Even although this was the first attempt at deploying a planetary aerostat, both of the balloons succeeded. They made the first measurements of the horizontal structure of the atmosphere to complement the many vertical profiles from descent probes. The temperature in the Vega 1 air mass was a constant 40°C. It was about 6°C cooler for the Vega 2 balloon. The atmosphere was more turbulent than expected. At times the balloons precipitously plunged in downdrafts of 1 to 3 m/s by hundreds of meters, sometimes several kilometers. The Vega 1 balloon encountered heavy turbulence at the start of its run and then again towards its end. Shortly after sunrise, passing over the Aphrodite Terra highlands, the Vega 2 balloon plunged more than 3 km to a pressure level of 0.9 bar, very close to the lower limit of its buoyant zone, before it rebounded.

atm	‘		km
0.6	–	——-	53
0.7	*	V 1 „	52
o.8	–	__ t___ i____ 1___ I___ i____ l____ I___ i___	5i

Ю 20 30 40 hOlirS

Figure 18.12 Flight profile of the Vega 1 balloon (from Don Mitchell).

The nephelometer on the Vega 1 balloon was hard to interpret due to calibration problems but generally seemed to agree with particle data from the nephelometers on the descent probes, showing the middle cloud in which the balloon drifted to be horizontally homogeneous with no clear regions. Unfortunately, the Vega 2 balloon nephelometer failed. In their cruise to the dawn terminator, the photometers noted some variation in light levels that may have been due to variations in the underlying clouds, and although there were some light flashes there was no strong evidence for lightning. Vega 1 crossed the terminator into daylight 34 hours into the flight, and its photometer registered dawn 2 hours prior to sunrise. The Vega 2 photometer did not
function correctly, but indicated dawn 3 hours before the terminator crossing. The anemometers reported downdrafts of 1 m s. The VLB! Doppler measurements found horizontal winds of up to 240 km/hr, made the first in-situ observations of the ‘super-rotation’ of the atmosphere at this altitude, and made measurements of atmospheric turbulence.

Results at Halley:

The results of the Vega Halley encounter were more than just scientific, they were also cultural and political. The project would be the first to image the nucleus of the world-famous Comet Halley. For the first time, a Soviet mission and its purpose was made known well in advance. The portion of the Vega mission at Venus went barely noticed outside scientific circles, but the whole world was waiting in expectation for the spectacle of the Halley encounter and the Soviets were well aware that this was unlike any other space mission they had ever conducted.

The Vega 1 spacecraft closed in on the comet at the blazing speed of 79.2 km/s in early March 1986. It performed a final trajectory correction on February 10. Its scan platform locked onto the comet on February 14 and began tracking. Far encounter images on March 4 and 5 demonstrated the camera’s performance. On March 6, the day of close encounter, the world’s press was present in the IKI control room for the first time, disturbing the usual professional calm with a bustling jumble of people eager to experience a Soviet mission event as it happened, including US television and media with both Roald Sagdeev and Carl Sagan providing commentary. Sagan as commentator for a Soviet spacecraft encounter in real-time was clear evidence that perestroika had become reality. Vega 1 switched to high rate telemetry 2 hours before closest approach and took over 500 images during the 3 hour encounter. The raw images looked overexposed and fuzzy. It was hard to pick out the nucleus from the obvious dust jets. But the IKI press room was filled with awe and applause. The images and other data streamed in for another 2 days.

Vega 2 closed in 3 days later at 76.8 km s. It did not require a final correction but 30 minutes before the encounter on March 9 it gave its controllers a scare when the computer guidance system failed. However, the spacecraft quickly switched over to the backup system and the observations began as planned. By the time the encounter was over on March 11 the spacecraft had provided over 700 images.

The images of Halley revealed a potato-shaped nucleus 14 x 7 km with a very dark albedo of 4%. a rotation rate of 53 hours, and at least five dust jets that could be counted on its sunward side. The environmental sensors on board the two spacecraft made pioneering measurements of the plasma fields in the vicinity of the comet, and defined the interaction of the solar w ind with the out-flowing eometary gases. Some of the constituents of the gas were identified and measured. The size and flux of dust particles varied enormously as the spacecraft flew through and in between the jets of dust and gas. A number of instruments were lost during the encounter, and the solar panels were extensively damaged by impacts and the electrical discharges that were induced by the cometary plasma.

Figure 18.14 Vega 2 image of Halley (processing by Ted Slryk).

The infrared spectrometer on Vega 2 failed due to a leak in the cryogenic system. The Vega 1 infrared spectrometer was sent an erroneous command which put it into calibration mode during the 30 minutes at closest approach, which was unfortunate, but it did report data taken at greater distances. The C-TT band of hydrocarbons was detected. The fact that the temperature of the nucleus was 300 to 400K meant that it had an insulating layer at its surface. The dust and gas were jetting through fissures in this crust opened by the heating of volatiles contained within. The three-channel spectrometer on Vega 1 was crippled by an electrical fault, and despite its partner on Vega 2 losing the ultraviolet channel this was able to detect water, carbon dioxide, the hydroxyl radical and the cyano radical, various other products of hydrocarbon photolysis, ammonia and other organic materials in the coma. It was concluded that the principal components of the gas were water containing carbon monoxide and carbon dioxide molecules, as well as photo-produced radicals and atomic hydrogen, oxygen and carbon.

Analysis of the dust in the jets revealed grains in the submicron-lo-micron size range of compositions varying from metallic to siliceous to carbonaceous. The dust mass spectrometers returned results showing three families of materials: one very similar Lo the carbonaceous chondrile meteorites which are thought to be the most primitive of Solar System material, another enriched in carbon and nitrogen, and the third enriched with water and carbon dioxide ice.

Thus ended one of the most daring, innovative, complex and successful missions in the history of robotic space exploration to that time. It established the USSR as the leader in the field; a distinction that was sadly short-lived and later forgotten.

EARLY SOVIET ROCKET DEVELOPMENT

The enabling technological step towards lunar and planetary space flight was the development of the military intercontinental ballistic missile (ICBM). From this, it is only a small incremental step to the development of a rocket capable of launching Earth-orbiting satellites, and then only another small step to one capable of sending spacecraft on trajectories to the Moon and beyond. The developers of ICBMs in both the US and USSR dreamed about space flight from the very beginning, and always in the back of their minds knew that the weapons on which they were working could ultimately be used for space exploration. This was as true for Sergey Korolev in the Soviet Union as for Wernher von Braun both in wartime Germany and later in the US. Each rapidly adapted their large rockets for flights to Earth orbit and beyond. The launch of Sputnik and the first Soviet launches to the Moon were made during the initial months of testing the R-7, the Soviet Union’s first ICBM. Subsequently, various versions of the R-7 became standard launchers for both military and civilian Soviet space missions. The ‘space race’ in the 1960s between these two nations was essentially defined by the development of ever more powerful rockets on both sides. The first intercontinental rockets developed in the US were the Atlas and Titan, and both were used in the civilian program for manned and robotic missions. However, the giant Soviet N-l and American Saturn V rockets were developed to land men on the Moon, and hence were far larger than required for military applications. Military rockets were modified by both nations to send spacecraft to the Moon and planets by adding upper stages for the extra boost required to achieve interplanetary velocities. Without these military rockets and the development of their associated upper stages, there would have been no access to space for interplanetary missions.

The history of rocketry in Russia can be traced back to their use by the military in the 13th Century – the same time that rockets made their appearance as a weapon in western Europe. A Rocket Enterprise was founded in Moscow’ in the 1780s. and in 1817 the Russian engineer Alexander Zasyadko wrote a manual on the production of

W. T. Huntress and M. Y. Marov, Soviet Robots in the Solar System: Mission Technologies and Discoveries, Springer Praxis Books 1, DOl 10.1007/978-1-4419-7898-1_4,

© Springer Science 4-Business Media, LLC 2011

Figure 4.1 Early GIRD rocket and team in the 1930s.

rockets and their use for artillery bombardment. By the beginning of WW-I, Russia had developed the artillery rocket into a significant weapon with a range of almost 10 km. This development gained momentum after the Russian revolution in 1917, as the newly established Soviet Union became an industrial state with a large military force. The establishment of the Gas Dynamics Laboratory in Leningrad in 1928 for development of military missiles marked the beginning of the later powerful Soviet military rocket design bureaus.

The first consideration of the rocket for use other than as a military weapon was by the Russian visionary Konstantin Tsiolkovskiy, whose book ‘The Exploration of the World’s Space with Jet-Propulsion Instrument’ wfas published in 1903; the same year as the Wright brothers’ first powered flight. Tsiolkovskiy, a schoolteacher, laid the theoretical foundation for space flight and interplanetary space travel using the rocket. In the 1930s, his work led a number of enthusiasts to found an organization called the Group of Research in Jet Propulsion (GIRD) whose first project was to construct a rocket-powered airplane. Sergey Korolev, the famed ‘Chief Designer’ of the Soviet space program in the 1960s, was a founding member. The government

The Cold War race to build an I CBM 33

began to sponsor the organization in 1932. and the group launched both a hybrid engine rocket and a liquid-fueled rocket in 1933. They were merged with the Gas Dynamics Laboratory in September 1933 as the Jet Propulsion Scientific Research Institute (RMI).

Progress was slow and resources very limited for these amateur rocketry pioneers in the 1930s. At that time, no government was interested in supporting a program to develop peaceful exploration of space. Military applications were the only hope for obtaining state budgetary support, and this happened first and most successfully in Germany during WW-II.

The energy requirements for a Mars flight were larger in 1971 than in 1969. This, and several engineering problems with the multiple instrument modules used in the Mars-

Figure 5.11 Mars 3 spacecraft (courtesy NPO-Lavochkin).

69 design, prompted yet another redesign. In the new version, the propulsion system at the base of the spacecraft formed the main structural element, and a single instrument module was mounted at the base of the cylindrical fuel and oxidizer tank system, forming a torus around the engine. As before, the solar panels, antennas, and thermal control system were attached to the side of the propellant tanks. New digital electronics were provided based on the avionics for the final stage of the N-l rocket. Advantage was taken of this heritage to save mass by removing the control system of the Proton Block D and allowing the spacecraft to manage the upper stage engine operations.

The higher energy requirements of the 1971 launch opportunity did not allow the orbiter to carry the entry vehicle into Martian orbit, so it would have to be deployed prior to orbit insertion. The higher atmospheric entry velocities and the decision to perform a soft landing, demanded a new entry vehicle design with a larger aerobrake possessing a shallower cone angle. The parachute would have to open at supersonic velocities, which was unprecedented. The final entry vehicle design was a modular stack consisting of the aerobrake at the forward end, the egg-shaped lander nested in the aerobrake, the toroidal parachute container on top of the lander, and a propulsion assembly at the rear of the entry vehicle. For the cruise, the entry vehicle was carried on top of the orbiter.

Lacking a sufficiently precise Mars ephemeris to provide accurate targeting of the entry systems prior to launch, it was decided to send an advance spacecraft to enter orbit around the planet and provide the navigational data necessary for the following two orbiter/lander missions to target and deploy their landers inbound to the planet. Unfortunately, the launch of the orbiter failed in May 1971 due to a stored command error. This accident had two very negative effects, the first being that the American Mariner 9 spacecraft would become the first to orbit Mars, and the second being that the two orbiter/landers would have to rely on a back-up, real­time and less accurate optical targeting technique. The launches were successful, and Mars 2 and 3 were on their way. The Mars 2 lander crashed w’hcn the back-up targeting system failed. On December 2, 1971, the Mars 3 entry system succeeded and its lander became the first to touch down on Mars. Unfortunately, the lander transmitted for only 20 seconds before failing and returned no useful data. Both parent spacecraft successfully achieved orbit.

The 1973 Mars launch opportunity was even less energetically favorable, making orbiter/lander combinations impractical. The lander would have to be deployed by a flyby vehicle. Four spacecraft were launched in July and August 1973, two orbiters and two flyby/landcrs. The spacecraft were essentially the same as in 1971, but the 1973 spacecraft were plagued by electronics problems due to manufacturing changes in a transistor used throughout the system. The engine on Mars 4 failed to ignite and the orbiter sailed past the planet. The Mars 5 orbiter succeeded, but failed after only about one month in orbit. The Mars 6 carrier had telemetry difficulties throughout its cruise, but managed to deploy its lander. The entry vehicle performed properly and transmitted the first in-situ atmospheric data, but no signal was ever received from the lander after it was dropped in close proximity to the surface. Mars 7 failed to put its lander on a proper trajectory, causing it to miss the planet.

Campaign objectives:

Nineteen months after their frustrating third campaign to Venus, the Soviets were ready with three more spacecraft for the late 1965 launch window. They had tried to reach this planet at every opportunity since February 1961, but after one test launch and seven launches they had nothing to show for it. Only two of the seven spacecraft survived their launch vehicles, and both of these failed in flight rather quickly. But the engineers reckoned they had fixed the problems that crippled Zond 1 and were encouraged by the success of Zond 3 at the Moon and its long interplanetary flight, so they prepared for the second 3MV Venus campaign with confident expectation.

Several 3MV spacecraft were left over from the November 1964 Mars campaign when only one had been launched during the window, flying as Zond 2. Another had been launched in July 1965 as Zond 3 for a test to Mars distance. Three 3MV Mars spacecraft, one configured with an entry probe (3MV-3 No. l) and the other two for flyby observations (3MV-4 No.4 and No.6), were modified for the Venus window in 1965. Their original target. Mars, accounts for their anomalous ‘tail numbers’. Only

Spacecraft launched
First spacecraft:	Venera 2 (3MV-4 No.4)
Mission Type:	Venus Flyby
Country j Builder:	USSR/OKB-1
Launch Vehicle:	Molniya-M
Launch Date: Time:	November 12, 1965 at 05:02:00 UT (Baikonur)
Mission End:	February 10, 1966
Encoun ter Dale і 7 ‘іme:	February 27, 1966
Outcome:	Failed in transit, communications lost.
Second spacecraft:	Venera 3 (3MV-3 No. l)
Mission ‘type:	Venus Atmosphere Surface Probe
Country і Builder:	USSR/OKB-1
Launch Vehicle:	Molniya-M
Launch Date; Time:	November 16, 1965 at 04:19:00 UT (Baikonur)
Mission End:	February 16, 1966
Encoun ter Dale і 7 ime:	March 1, 1966
Outcome:	Failed in transit, communications lost.
Third spacecraft:	3MV-4 No.6 (Cosmos 96)
Mission Type:	Venus Flyby
Country і Builder::	USSR/OKB-1
Launch Vehicle:	Molniya-M
Launch Date: Time:	November 23. 1965 at 03:22:00 UT (Baikonur)
Outcome:	Failed to depart Farth orbit.

two were successfully dispatched. Venera 2 and 3 flew to the vicinity of their target and became the first truly successful interplanetary cruises since Korolev had begun launching planetary spacecraft in 1960. The long interplanetary cruise provided new confidence in the spacecraft, but the fact that they failed at or near their target made them agonizing disappointments. There was a fourth spacecraft, probably with an entry probe, but this was unable to be launched before the window closed.

Venera 2 and 3 were also the last planetary spacecraft to be built and launched by OKB-1 because in late 1965 Korolev had transferred responsibility for robotic lunar and planetary missions to NPO-Lavochkm. The next Venera spacecraft for the 1967 window would be built and launched under the leadership of Gcorgi Babakin.

Spacecraft:

The Venera 2 and 3 spacecraft were basically the same as Zond 2 and 3 but modified for the new target. The Venera 3 entry probe was essentially the same as that carried by Zond 1. By the time the mission was launched, there was strong evidence that the surface of Venus was hot, possibly 400°C. Although the surface pressure was not yet well determined, it was apparent that conditions were beyond the limits to which the 3MV probe was designed (77aC and 5 bar). As it was too late to make changes, Venera 3 was launched in full knowledge that its probe would provide only data on the atmosphere and would not survive the full descent to the surface.

Figure 9.10 Venera 2 (left) and Venera 3 (right).

963 kg (Venera 2)

958 kg (Venera 3)

~ 950 kg (Cosmos 96) 337 kg

Payload:

Venera 2 carrier spacecraft:

1. Lyman-alpha and oxygen spectrometer

2. Triaxial fluxgatc magnetometer

3. Micrometeoroid detector

4. Charged particle detectors

5. Cosmic ray gas discharge and solid state detectors

6. Cosmic radio emission receivers for 20 to 2,200 kl-Iz

7. Decimeter band, radio solar plasma detector

The cosmic ray detectors now consisted of the gas discharge counters and silicon solid-state detectors. The decimeter band radiometer dish antenna was mounted on the ring between the avionics and instrument compartments.

Venera 2 flyby instrument module:

1. Facsimile imaging system

2. Ultraviolet spectrometer at 285 to 355 nm in the imaging system

3. Ultraviolet spectrometer for ozone at 190 to 275 nm

4. Infrared spectrometer at 7 to 20 and 14 to 38 microns

The camera system and ultraviolet spectrometers were identical to those carried by Zond 2 and 3. The camera was provided with a 200 mm lens. The Venus infrared spectrometer was similar to that of Mars 1 but designed to measure thermal radiation from the atmosphere and clouds. It covered two ranges in 150 increments each, the first using an InAn window7 and the second a LiF mirror. The instrument had a mass of 13 to 15 kg, was 50 cm in size, and was mounted outside the instrument module, coaxial with the imaging system, and included a visible photometer for reference. It could also make a spatial scan of the planet at the two fixed wavelengths of 9.5 and 18.5 microns.

Venera.? carrier spacecraft:

1. Lyman-alpha and atomic oxygen photometers

2. Triaxial fluxgate magnetometer

3. Charged particle detectors

4. Cosmic ray gas discharge and solid state detectors

5. Decimeter band radio solar plasma detector

The cosmic ray instrument had an additional gas discharge counter on Venera 3. and both the micrometeoroid detector and the radio emission receivers were deleted.

Venera.? descent I landing capsule:

1. Temperature, pressure and density sensors

2. Atmospheric composition, acidity and electrical conductivity experiments

3. Gamma-ray surface composition detector and cosmic ray detector

4. Visible airglow photometer

5. Mercury level motion experiment

The probe instruments were spares from the 1964 campaign. The photometer was included again since Venera 3 was to be a night-time landing. As with all of the ЗМ V missions, the probe also carried pennants of the Soviet Union.

Mission description:

The Venera 2 flyby spacecraft was successfully launched on November 12, 1965 .It was intended to fly in front of the sunlit hemisphere of Venus and photograph it at a range of less than 40,000 km. The initial trajectory w as so precise that no midcourse maneuver was required. The thermal system did not function well and the spacecraft began to overheat as it neared its target, causing problems with the communications system. An improper coating of the radiation domes was suggested as the cause. By February 10. which proved to be the final interrogation session, the temperature was considerably increased, the quality of communications was seriously degraded, and the command from Harth to initiate flyby observations was not acknowledged. After the flyby Venera 2 failed to respond to commands to download the flyby data, and on March 4 it was declared lost. It may very well have achieved its mission and been unable to transmit its results to Harth. The closest point of approach to the planet was at 02:52 UT on February 27, 1966, at a distance of 23.950 km.

Venera 3 was dispatched towards Venus on November 16, 1965. It performed satisfactorily during cruise and a midcourse correction on December 26 put it on an impact trajectory 800 km from the bull’s-eye. How7ever. the communications system failed on February 16, just seventeen days prior to arrival. The spacecraft may have released its entry probe automatically at 06:56 UT on March 1. 1966, but there was no telemetry from the capsule. Even so. the probe became the first human artefact to reach another planet, near the terminator on the night side somewhere between 20°S and 20 N and between 60 E and 80 E.

The post-mission investigation into the loss of Venera 2 and 3 revealed problems with the thermal control system in both spacecraft which had caused components in the communications system to overheat and fail.

The third spacecraft, 3MV-4 No.6, was launched on November 23. A broken fuel line caused one of the engine chambers in the third stage to explode shortly prior to stage shutdown, with the result that the fourth stage inherited an unstable attitude. It managed to achieve orbit, but the tumbling prevented it from restarting its engine for the escape maneuver. Written off as Cosmos 96, it re-entered on December 9.

A fourth spacecraft (probably 3MV-3 No.2) w as to be launched at the very end of the w indow on November 26, 1965. but was scrubbed w7hen a problem was found in the launch vehicle during pre-flight checks. The launch was abandoned because the vehicle could not be recycled before the window closed.

These were the last robotic interplanetary spacecraft launched by OKB-1. Out of a total of 39 launch attempts in a period of a little more than seven years, only Luna 2, Luna 3, and Zond 3 fulfilled their missions. Twenty lunar launch attempts gave eight successful launches, with only three spacecraft being fully successful. Hlevcn Venus launch attempts gave four successful launches, but unfortunately no spacecraft were successful. Out of six Mars launch attempts only two succeeded, but both spacecraft failed. Two ЗМ V test launches also failed.

Results:

The 1965 campaign produced no data from Venus. Some results were published on micrometeoroids, the interplanetary magnetic field, cosmic rays, low energy charged particles, solar wind plasma fluxes and their energy spectra.

Campaign objectives:

The Soviets had a strong desire to follow their original long-term plan for the 1971 campaign and build a new entry vehicle containing a soft lander, but the M-69 losses meant NPO-Lavochkin lacked both the detailed ephemeris for the planet and the atmospheric data which was required to design a soft lander. One option was to repeat the atmospheric probe mission with a hard lander in 1971 to obtain this data, and postpone the soft lander to 1973. But the 1973 opportunity would require more energy, and so would require separate rather than combined launches for an orbitcr and lander. This would mean launching at least four vehicles, two orbiters and two flyby spacecraft carrying landers, and redesigning the entry vehicle to accommodate entry from the initial approach rather than from orbit. This scenario was deemed too expensive at the time, but it is exactly w? hat the Soviets ended up doing in 1973. An alternative w as to get the data from the IJS. The Mariner 4, 6 and 7 flyby missions in 1965 and 1969 had studied the atmosphere and estimates of the surface pressure had been published, but the crucial ephemeris had not been published and the Americans w7ere unwilling to supply it to the Soviets since the antagonism of the Cold War was rife at the time.

Ultimately, the Soviets settled upon a clever but risky approach to implementing a soft lander which facilitated the launch of combined orbiter, landers in 1971 without requiring pre-launch data on the planet’s ephemeris. This involved sending another spacecraft ahead of the two orbiter/landers to enter orbit around Mars and serve as a radio beacon that the other spacecraft would use to achieve the desired navigational accuracy. On this orbiter the mass which would normally have been allocated to the entry system facilitated the larger propellant load required to achieve a high energy, fast trajectory and increased the scientific payload. Optical tracking during approach and radio tracking in orbit would enable the ephemeris to be derived in sufficient time for trajectory corrections to be sent to the orbiter landers. Once in orbit, the leading spacecraft would act as a radio beacon to assist the entry vehicles navigate

Spacecraft launched
First spacecraft:	M-71S (М-71 Ко. 170 and Cosmos 419)
Mission Type:	Mars Orbiter
Country; Builder:	USSR/NPO-Lavochkin
Launch Vehicle:	Proton-K
Launch Date ‘: 7 ime:	May 10, 1971 at 16:58:42 UT (Baikonur)
Outcome:	Stranded in orbit, fourth stage failed to reignite.
Second spacecraft:	Mars 2 (M-71 Nod 71)
Mission Type:	Mars Orbit er/Lander
Country і Builder:	USSR/NPO-Lavochkin
Launch Vehicle:	Prolon-K
Launch Date; Time:	May 19. 1971 at 16:22:44 UT (Baikonur)
Encounter Date/ Time:	November 27, 1971
Mission End:	August 22, 1972
Outcome:	Or biter successful, lander crashed
Third spacecraft:	Mars 3 (M-71 No. 172)
Mission Type:	Mars Orbit er/ Lander
Country і Builder:	USSR/NPO-Lavochkin
Launch Vehicle:	Proton-K
Launch Date: Time:	May 28. 1971 at 15:26:30 UT (Baikonur)
Encounter Date/ Time:	December 2, 1971
Mission End:	August 22, 1972
Outcome:	Or biter successful, lander failed on the surface

their approach following release by their carriers. The Americans were planning to send two Mariner spacecraft to enter orbit around Mars at this launch opportunity. Sending a spacecraft on ahead offered the Soviets the propaganda advantage of being first to insert a spacecraft into orbit around the planet.

The scientific objectives of all the Soviet orbiters were to image the surface of the planet and its clouds, study the topography, composition and physical properties of the surface, measure properties of the atmosphere, make temperature measurements, and study the solar wind and interplanetary and planetary magnetic fields. The two carrier vehicles were also to relay back to Larth the transmissions from their landers. The entry system was to make atmospheric measurements during entry and deliver the lander to the surface. The objectives of the lander were to return images from the surface, obtain data on meteorological conditions and atmospheric composition, and deploy a small rover that would measure the mechanical and chemical properties of the soil.

These Soviet missions and the US Mariner 9 or biter in 1971 had the potential to transcend the pervasive competition between the two space faring powers with the first cooperation by a telephone ;hot line’ that was set up between the Jet Propulsion Laboratory in Pasadena and the Soviet space center in Yevpatoriya, Crimea, for the exchange of results.

Spacecraft:

Orbiters:

Designated M-71S (S for Sputnik, or orbiter), the lead orbiter would require much larger tanks than the M-69 spacecraft to enable it to fly the higher energy trajectory required to arrive at Mars ahead of the orb ter landers. In conjunction with a number of engineering problems with the multiple instrument modules of the M-69 design, this prompted the Soviets once again to redesign the entire spacecraft. Instead of the propellant tank being the main structural element, this function was assigned to the KTDU-425A propulsion system. The fuel and oxidizer tanks formed a 3 meter long cylinder on top of the propulsion system. The avionics and science instruments were in a hermetically sealed module at the base of the cylinder, forming a toroid around the propulsion system. The gimbaled. engine nozzle attached at the base of the tank protruded through the center of the instrument module. Instruments could be reached during testing simply by detaching the lower half of the toroidal cover.

Figure 12.5 Mars-71 S orbiter spacecraft.

Two 2.3 x 1.4 meter solar arrays extended from opposite sides of the cylindrical tank. Attached to the solar arrays were cold gas attitude control jets, an antenna for relaying the lander’s transmission, and the magnetometer booms. A parabolic high – gain antenna 2.5 meters in diameter was mounted on the side to support redundant transmitters for 5 and 32.5 cm (5.8 GHz and 928.4 MHz). Three omnidirectional spiral antennas were installed near the high-gain antenna. The thermal control radiators and tanks of attitude control propellant were on the side of the cylinder. Navigational optics were on the outside of the instrument module – a pair of star sensors pointing downward in terms of the vehicle’s structure, three Sun sensors in a vertical stack, all pointing radially out, an Earth sensor that was aligned with the parabolic antenna, and a Mars sensor aimed horizontally off to one side.

M-71S launch mass: 4,549 kg (dry mass 2,164 kg)

The orbilcr/’landcrs to follow the M-71S were designated M-71P (P for Posadka, or lander). They had shorter tanks with less propellant, the mass being used for the entry system carried on top of the tank, but otherwise they were almost identical to the M-71S and they were almost identical to each other. With its lander the M-71P was 4.1 meters high with a base diameter of 2 meters. The span across the deployed solar panels was 5.9 meters. They incorporated a new digital guidance and control computer based on the prototype for the Block D stage of the N-l rocket. 1’his was capable of significantly greater navigational accuracy, but with a mass of 167 kg and a power rating of 800 W it was rather demanding. The extra mass was compensated by deleting the control system from the Block D and instead using the spacecraft to control the stack. This is an interface design that would never have been considered m the US.

Figure 12.6 Mars 3 spacecraft.

Figure 12.7 Mars-71 orbiter/landcr spacecraft: 1. Lander; 2. Parabolic antenna; 3. Attitude control jets; 4. Spiral antenna; 5. Mars sensor; 6. Star sensor; 7. Star sensor; 8. Propulsion system; 9. Instrument compartment; 10. Attitude control gas tanks; 11. Thermal radiators; 12. Earth sensor; 13. Solar panels; 14. Magnetometer; 15. ‘STEREO" experiment antenna.

Mars 2 and 3 launch mass: 3,440 kg (orbiter; dry mass 2,265 kg)

1,210 kg (entry vehicle)

635 kg (lander system on descent)

358 kg (lander)

4,650 kg (total)

The 1971 spacecraft were much easier to work on in testing operations, and were more readily modified for various planetary missions by changing instruments in the module, attaching various modules to the top of the tank, and changing the length of the tank itself. The 1971 design formed the basis for all subsequent Mars spacecraft, and all Venera spacecraft beginning with Venera 9 through the Vega spacecraft, and for astrophysics spacecraft in Earth orbit.

Entry system:

A new entry system was required to slow the spacecraft rapidly in the thin Martian atmosphere for a soft landing. The steep cone angle of the entry vehicle designed for the (unflown) 1969 atmospheric probe would not be adequate. For a soft landing in 1971 a much larger entry shell 3.2 meters in diameter and with an open vertex angle of 120 degrees was devised to maximize the altitude at which the parachute opened. Furthermore, the parachute would have to open at a supersonic velocity of Mach 3.5, a feat that had never been done before. This engineering and test challenge w’as met by a program of drop tests using balloons at an altitude of 35 km and meteorological rockets at 130 km. Due to the lack of data on the Martian atmosphere, the aerobrake for the M-71 system was designed for an uncontrolled ballistic descent instead of the controlled descent to be used by the Viking entry vehicles that the Americans were designing.

The entry system comprised four stacked assemblies: the aerobrake at the forw ard end. the egg-shaped lander nested in the aerobrake, the toroidal parachute container above the lander, and the propulsion assembly at the rear with the latter including a structural ring. The stack was held together by four crossbars linking the rim of the aerobrake to the ring at the rear. Unlike US designs, there w? as no monolithic back shell. The role of the solid rocket in the center of the propulsion ring assembly was to separate the entry system from the orbiter after release and to transfer from the flyby trajectory to the desired entry trajectory. The carrier would remain on the flyby trajectory until firing its own engine for orbit insertion. For attitude control, tanks mounted on the interior of the propulsion ring assembly provided nitrogen to the cold gas micro-engines located on the crossbars near the rim. Small solid rocket micro-engines were affixed to the aerobrake rim in order to spin the vehicle prior to entry and to de-spin it following entry in readiness for deploying the parachute. The vehicle was actively З-axis controlled from its release to the spin-up for entry, passively aerodynamically controlled during entry, and passively controlled for parachute descent. The toroidal section holding the parachutes, deployment devices, and terminal rocket engines w’as attached to the lander. The aerobrake w^as connected to the parachute container by metal bands on the underside. The avionics to control the sequence of entry, descent and landing were contained in a small

Figure 12.8 Mars 3 entry system diagram: 1. Main solid rocket; 2. Avionics; 3. Main parachute; 4. Lander surface station; S. Aeroshell; 6. Altimeter antenna; 7. Parachute container; 8. Relay antennas; 9. Drogue parachute pyro.

Figure 12.9 Mars-71 entry system.

cylinder attached to the underside of the toroid, which was itself designed to separate into two halves. A solid rocket device with four small nozzles affixed to the side of the upper half dragged the 13 square meter drogue parachute from the toroid. The upper half of the toroid was separated and carried away by the drogue, which in turn pulled out the 140 square meter main parachute whose lines were connected to the bottom half. The solid terminal rockets were deployed in a container part way up the

Figure 12,10 Mars-71 lander diagram: 1. Radar altimeters: 2. Shock absorber; 3. Telemetry units; 4. Automatic radio system; 5. Antennae; 6. Radio; 7. Radio system units; 8. Science instrument module; 9. Imaging system; 10. Petal locking pin; 11. Instrument deployment system; 12. Science sensors; 13. Internal thermal insulation; 14. External thermal insulation; 15. Petal deployment mechanisms; 16. Petals; 17. Aeroshell cap displacement balloon; 18. Aeroshell cap; 19. Aeroshell cap shock absorber; 20. Gas cartridge for displacement balloon; 21. Control system; 22. Batteries; 23. Pressure sensor.

Figure 12.11 Mars-71 engineering lander in test bed. and a sectioned model (insert).

shroud lines. The radar altimeter was mounted inside the lander at the bottom of the instrument compartment.

Lander:

The lander was an egg-shaped capsule 1.2 meters in diameter across the middle that was entirely covered with a 20 cm thick layer of foam. The foam was in two pieces, one an aeroshcll cover in the form of an ejectable cap over the larger top portion of the lander capsule and fitting onto a small skirt encircling the bottom of the capsule; and the other a lens-shape which was permanently mounted on the bottom, under the encircling skirt, in order to absorb the shock of landing. The foam aeroshell cap was ejected after landing by inflating a balloon to allow the petals to open, in the process righting the lander and exposing its internal instruments. Two camera ports and four deployable elastic aerials protruded from the top of the sphere for communicating with the orbiter. The tethered rover was mounted on a deployable

Figure 12.12 Lander diagrams showing surface deployments and impact shock absorber (from Ball ct al.).

arm. The lander was powered with batteries that would be charged by the orbiter prior to separation. Temperature control was by thermal insulation covering the exposed portions and a system of radiators. It was designed to survive the chill of the Martian night.

The entire lander capsule weighed 358 kg and was sterilized prior to launch by germicidal lamps to prevent contamination of the Martian environment. It was tested using catapults and rated for horizontal speeds of 28.5 m/s. vertical speeds of 12 m/s and impacts of 180 G. Figure 12.10 shows it with petals closed and encapsulated in its foam aeroshell cap and impact shock absorber.

Entry у descent and landing:

Rather than having the inbound carrier spacecraft target the atmospheric entry point, release a passive entry system, and then perform a deflection maneuver to reach the position where it would perform orbit insertion, the Soviet mission design targeted the carrier at its insertion point and required a more complex entry system that had a propulsion system with which to maneuver for the requisite atmospheric entry point and angle of attack. The difference in entry strategies for Venus and M ars was due to the nature of; their atmospheres. The atmosphere of Venus is so thick that a simple spherical shell with an offset center of mass for attitude alignment is readily able to reduce the entry velocity to subsonic far above the surface. The atmosphere of Mars is rarefied and requires a large conical aeroshell to slow the velocity rapidly enough and high enough in the atmosphere for parachutes and terminal rockets to be able to cancel the residual velocity prior to surface contact. The Martian atmosphere levied stringent requirements on the entry angle: if it were too steep then the vehicle would reach the surface before the various velocity reduction steps could be completed; too shallow, and the vehicle would skip out of the atmosphere. Furthermore the conical shield had to be properly oriented relative to the incoming velocity vector and spin stabilized to hold this orientation. The requirement to deliver the vehicle on a precise trajectory and entry angle despite the lack of an accurate ephemeris for Mars, drove the designers to enable the carrier to autonomously undertake optical navigation as it closed in on the planet and release the entry system just hours prior to entry. The Venera carriers released their entry systems 2 days before entry and followed them into the atmosphere and destroyed themselves. But for the 1971 Mars missions the carrier was to enter orbit. To have maneuvered the entire spacecraft to the trajectory for atmospheric entry, released the entry system, and then performed a deflection maneuver so near the planet would have required a prohibitive amount of propellant. The tradeoff in mass therefore favored the orbiter by complicating the entry system with a maneuvering engine and active З-axis attitude control capability.

Figures 12.13 to 12.17 illustrate the approach, separation, trajectory correction, entry, descent and landing sequences for the Mars-71 entry system. All events after the entry system separates from the orbiter occur automatically, without command from Farth. The entry mission begins with the pyrotechnic separation of the entry system from the orbiter at a distance from Mars of about 46,000 km. At this time the

Figure 12.13 Mars-71 approach and targeting sequence: 1. First optical navigation measurement at ~ 70,000 km range to update orbiter and entry vehicle trajectory parameters; 2. Trajectory correction maneuver (the Lhird since leaving Earth) to target the orbiter, with a velocity change of less than 100 m/s changing the periapsis from ~2.350 – F 1,000 km to 1,500 + 200 km; 3. Entry vehicle separation about 6 hours before entry; 4. Entry vehicle trajectory correction maneuver to target entry vehicle. Entry angle accuracy ™ 5 deg, velocity change ~ 100 m/s, propulsion system ejected post maneuver; 5. Entry vehicle reorientation to entry attitude and spin-up; 6. Second optical navigation measurement at ~20,000 km range to update orbit insertion parameters; 7. Mars orbit insertion maneuver. Velocity change ~ 1,190 m/s, orbital period accuracy ~2 hrs.

Figure 12.14 Mars-71 entry sequence: 1. Entry system separation 6 hours from entry; 2. Solid rocket ignition to retarget from flyby to entry trajectory; 3. Separation of the propulsion system and spin-up; 4. Spin-down after peak deceleration; 5. Aero braking.

Figure 12.15 Mars-71 pilot parachute braking sequence: 1. Accelerometer initiates descent program timer at I = 0, auxiliary parachute cover is severed and extraction rocket is ignited: 2. Drogue parachute and cover is extracted from its container; 3. Drogue parachute shroud line is extracted from the container and tension huilt up in suspension lines: 4. Drogue parachute is released from the extraction mechanism and opened at t =0.7 sec; 5. Top half of the toroidal main parachute cover is severed and drawn away; 6. Main parachute is extracted with shroud lines attached to the bottom half of the toroidal compartment; 7. Main parachute is deployed, hut reefed by a ripcord to prevent overload. Descent science instruments activated at t =3.1 sec.

entry system is under З-axis attitude control. After 900 seconds (by now’ hopefully a safe distance from the orbiter) the main solid rocket is fired to provide an impulse of 120 m/s and adopt the required entry trajectory. 100 seconds later, the vehicle rotates to the proper entry attitude. After another 50 seconds, a set of solid micro-engines on the aerobrake rim are ignited, each delivering 0.5 kN for 0.3 second to spin up the vehicle to 10 rpm. Then the propulsion ring assembly is jettisoned, taking with it the attitude control system and the mounting bars. The spin-stabilized vehicle coasts to its target.

The vehicle enters the atmosphere at about 5.8 km/s. When the load drops to 2 G after peak deceleration, spin stabilization is no longer required and the second set of solid micro-engines on the aerobrake rim arc fired to de-spin the vehicle. After about 100 seconds, at a preset G equivalent to about Mach 3.5, an accelerometer triggers the start of the descent program timer at t = 0 and deploys the 13 square meter drogue parachute. The toroidal section is bisected at t = 2.1 seconds and its Lop half is pulled away by the drogue, drawing out the main parachute. The drogue is then released. The 140 square meter main parachute is reefed to prevent over stressing it

f igure 12.16 Mars-71 main parachute descent sequence: 1. Ripcord cut at 12.1 sec to fully open the main parachute; 2. Heat shield separated at t = 14 see. At t~ 19 sec the high altitude radar altimeter is activated; 3. At t—25 see, pyros are fired to release the terminal rocket; 4. The main parachute extracts the rocket on a new set of shroud lines. At l = 27 see the low altitude radar is activated; 5. After 30 to 200 seconds on the parachute, at a height of 16 to 30 meters the low altitude radar turns off the descent science instruments and ignites the terminal landing rockets; 6. The parachute is carried away by another rocket and the lander is dropped; 7. The lander free falls to the surface.

at such a high speed. The descent science instruments are activated at t = 3.1 seconds. At t = 12.1 seconds, after the speed has become subsonic, the reef lines arc cut and the canopy opens fully. The aerobrake is jettisoned at t= 14 seconds.

The high altitude radar is activated at t= 19 seconds and a descent rate of about 65 m/s. At t — 25 seconds the lower shroud lines are withdrawn from the toroid with the terminal solid rocket system at their top, and at t^27 seconds the low altitude radar is activated. After 30 to 200 seconds on the parachute, at a height of 16 to 30 meters the radar triggers the landing sequence in which, in rapid succession, a second timer is initiated, the descent science instruments are turned off, the lander terminal

Figure 12.17 Mars-71 landing sequence: 1. The terminal rockets are ignited and another rocket carries the parachute away; 2. The lander is dropped and comes to rest on the surface; 3. The displacement balloon inflates to separate the top cover of the lander (at right); 4. Petals open on the upper hemisphere to stabilize the lander, the antennas and booms are deployed, and the science package is activated.

solid rocket is ignited to deliver 56 kN for 1.1 seconds, and the parachute is carried away by a second rocket that fires for 1 second and delivers a thrust of 9 kN. After terminal rocket firing, the lander is released to fall to the surface and two small rockets on the side of the terminal rocket container deliver a horizontal impulse of 1 kN for 4 seconds in order to prevent it from falling onto the lander. Meanwhile, the lander should impact at a vertical velocity no greater than 12 in.’s.

Fifteen seconds after the lander makes physical contact with the surface, a timer commands the ejection of the foam cap covering the petals and initiates the lander’s sequence. This deploys the lour petals, antennas, and booms, and starts to transmit to the main spacecraft at a rate of 72 kbits/s on two independent VHF channels. This communication session lasting about 20 minutes has to occur before the spacecraft makes its insertion maneuver. It includes a panoramic image of 500 x 6,000 pixels. The lander is then powered down, as it will be between all communications sessions. The sessions are initiated by timer and may be as short as 1 minute depending on the location of the site, the nature of the terrain, and the mutual orbiter/lander positions. The lander was designed to operate for several local days.

The entire descent sequence was tested by fifteen M-100B sounding rocket flights using scale models dropped from 130 km.

Payloads:

M-71S orbitev:

The scientific payloads of the orbiters were almost identical, except that the M-71S and Mars 3 spacecraft both had the French STEREO instrument to measure solar outbursts. This was the first time a Soviet spacecraft carried a Western instrument. However, the Soviets still guarded their secrecy and the French simply handed over the equipment "’at the border”. They were not involved in its integration and testing. In fact, they were not shown any drawings, and were not told where and on which spacecraft the instruments would be mounted. The loss of the M-71S orbiter left this experiment with only one instrument, compromising the stereoscopic aspect of the project.

Mars 2 and 3 orbiters:

Most of the orbiter scientific instruments were mounted in the hermetically sealed instrument module, and were generally intended to be operated for 30 minutes near each periapsis. Others were externally mounted or had externally mounted sensors for in-sit u investigation of the space environment:

F FPU dual camera facsimile imaging system

2. Infrared radiometer (8 to 40 microns) for measurement of surface temperatures

3. Infrared narrow-band 1.38 micron photometer for measurement of water vapor content in the atmosphere

4. Infrared spectrometer in the 2.06 micron absorption band of carbon dioxide to measure atmosphere optical thickness and as an indicator of surface topography

5. Ultraviolet photometer with filters in the intervals 1,050 to 1Л80, 1,050 to 1,340 and 1,225 to 1.340 angstroms to detect atomic hydrogen, oxygen, and argon

6. L у man-alpha photometer (French-Soviet) for measurement of upper atmo­sphere hydrogen

7. Six channel visible photometer in range 0.35 to 0.7 microns for measurement of color and albedo of the surface and atmosphere

8. Microwave radiometer (3.4 cm) for measurement of dielectric constant and subsurface temperatures to depths of 25 to 50 cm

9. Radio science investigation to determine atmospheric structure (temperalure and density profiles)

10. Cosmic ray charged particle detector consisting of a Cherenkov counter, four gas discharge detectors and seven silicon solid-state detectors

11. Solar wind plasma sensors (8) for measurement of speed, temperature and composition in the 30 eV to 10 keV energy range

12. Boom mounted three-axis iluxgate magnetometer

13. STEREO instrument on M-71S and Mars 3 to measure solar radiation outbursts at 169 MHz in conjunction with Earth-based receivers (French- Soviet).

The Mars 2 and 3 photo-television imaging system was an improvement over the M-69 system, and consisted of two bore-sighted film cameras, one with a 52 mm wide angle lens and several color filters and the other with a 350 mm narrow angle lens and an orange filter. At the planned periapsis altitude, surface resolutions of 100 to 1,000 meters were expected. There was film for 480 images, most of which were pre-programmed for the first 40 days of the orbital mission.

The science instruments on the Mars 3 orbiter weighed a total of 89.2 kg.

Mars 2 and 3 entry systems:

A radio altimeter attached to the toroid provided data during the descent. The lander payload had a mass of 16 kg and consisted of:

1. Accelerometer for atmospheric density during entry

2. Temperature and pressure sensors for descent and landing

3. Radio altimeter for providing altitudes on descent

4. Mass spectrometer for atmospheric composition on descent and landing

5. Atmospheric density and wind velocity on the surface

6. Two panoramic television cameras for stereo view ing of the surface

7. X-ray spectrometer for soil composition deployed to the surface from a petal

8. PrOP-M walking robot deployed to the surface from this same petal with onboard gamma-ray densitometer and conical penetrometer.

The cameras were similar to those of the Luna 9 lander with a single photometer and a scanning mirror that tilted to scan vertically and rotated to scan horizontally, returning a single brightness value for each scan position. A full panorama spanned 500 x 6,000 pixels. The mass spectrometer w as an early form of the Bennett radio – Irequency instrument being developed for Venera 9 and 10. There was no telemetry during the descent. All data obtained during this time was stored for transmission in the communication session programmed for immediately after touchdown.

The 4.5 kg PrOP-M rover was a box 250 x 250 x 40 mm with a small protrusion rising from the center of its upper surface. The body w*as supported by two skis, one projecting down from each side. By moving the skis in alternating fashion the rover w*as able to ‘walk and by moving them in opposite directions it could turn. There were obstacle-sensing bars at the front, and it was programmed to reverse in order to circumnavigate an obstacle. The rover was to be deployed by a 6-joint manipulator arm and moved into the field of view’ of the cameras. It w as tethered by a 15 meter long cable for direct communication w ith the lander, and was to pause at intervals of

Figure 12.18 PrOP-M ‘Marsokhodnik’ rover.

1.5 meters to make measurements. It carried a dynamic penetrometer and a gamma – ray densitometer, and its tracks were to be photographed to investigate the physical properties of the surface.

Mission description:

M-71S:

The Soviets must have breathed a sigh of relief on May 8, 1971, when the launch of Mariner 8 failed. Their plan was for the M-71S spacecraft to arrive at Mars and enter orbit before the two US spacecraft arrived, and their chances of achieving this had just improved. The M-71S orbiter was launched two days later, on May 10, but the failure of the Block D to reignite due to an ignition timing error – “a most gross and unforgivable mistake” – left the spacecraft stranded in parking orbit. The timer was intended to have been set to reignite the engine 1.5 hours after the Block D achieved orbit, but the 8-bit code was erroneously specified as 150 hours by the programmer who input the command with the bits in reverse order. The coupled spacecraft and stage was named Cosmos 419 by the Soviets to hide its purpose. It re-entered 2 days later.

This failure not only cost the Soviets a chance to be the first to orbit Mars, it also threatened the success of the Mars 71 campaign because it meant there would be no radio beacon orbiting the planet to assist in refining Ihc trajectories of the spacecraft carrying landers. The French were not informed of the loss of their first STEREO instrument. The Soviets would have to resort to the backup method of correcting the trajectory, which was less accurate and much more risky. Lacking an accurate Mars ephemeris to calculate a pre-determined release point and how to orientate the entry system in relation to Mars, each approaching spacecraft would have to use on board optical sensors to determine its position relative to the planet and then calculate for itself the release point, the trajectory correction required to reach this point, and the orientation that the entry system must adopt in readiness for atmospheric entry. This autonomous procedure using an optical navigation instrument had been developed as a back up contingency, but the M-71S failure made it the only option. It was bold, very complex, highly sophisticated, and far ahead of its time. Several decades would pass before American mission designers adopted automated optical navigation: had they known so at the time they would have been aghast at its use for Mars 2 and 3.

The mission plan for the Mars-71 campaign allowed as many as three midcourse correction maneuvers, but nominally used only two, the first soon after leaving Earth and the second on approaching Mars. Another correction now became essential, and was dedicated to the autonomous entry system targeting procedure. The first step, at about 70,000 km from Mars, would be to make the optical navigation observations required to correctly target the entry system. After a new vector had been calculated and the course corrected, the entry system would be released to pursue its standard procedures. The main spacecraft would then undertake a second optical observation about 20,000 km from Mars in order to identify any change required for the orbit insertion maneuver. All of these operations were to be performed autonomously.

Mars 2:

After a successful launch on May 19, 1971, the first trajectory correction maneuver was conducted on June 5. Almost simultaneously on June 25 communications with both Mars 2 and Mars 3 in the primary decimeter band were lost, evidently owing to problems with the transmitters. After working for a brief period the decimeter back­up transmitter also failed on Mars 2. It proved impossible to activate its centimeter band telemetry system. The primary decimeter transmitter remained unreliable, but conditions were identified in which the back-up transmitter could be made to work. The loss of the centimeter band system was never understood, but it worked reliably on subsequent missions. There were no further incidents and on November 21, with 6 days remaining to arrival. Mars 2 performed an optical navigation sequence and 7 hours later made its second trajectory correction. The third maneuver, to target the entry system, was made on November 27 but it proved to be fatally imprecise. After being released 4.5 hours before the main spacecraft was to perform its orbit insertion maneuver, the entry system ran through its standard procedures. The orbiter made a trim burn, then the 1.19 km/s insertion maneuver and settled into a 1,380 x 24.940 km orbit inclined at 48.9 degrees. The problem with the third targeting maneuver resulted in a low er apoapsis than intended, with a period of 18 instead of 24 hours.

Meanw7hile, having entered the Martian atmosphere at a velocity of approxi­mately 6.0 km/s at a steeper angle than planned, the descent system malfunctioned and the lander hit the surface before it could deploy its parachute. It fell at 44.2°S 313.2 W. delivering a coat of arms of the USSR. Post-flight analysis show ed that the computer codes were not sufficiently developed owing to lack of development time to address all situations, including that faced by Mars 2 in which the trajectory prior to the third correction was fairly close to that desired and the ensuing procedure over-corrected and produced an overly steep entry angle.

Mars 3:

Mars 3 was launched on May 28, 1971, and performed its first midcourse correction on June 8. The primary decimeter band transmitter failed on 25 June, but the back­up functioned. The cruise was uneventful, and on November 14 the spacecraft made a second midcourse maneuver. On approaching Mars on December 2 it executed the autonomous final targeting. At 09:14 UT. some 4 hours 35 minutes prior to orbital insertion, the spacecraft cut loose the entry system. Fifteen minutes later, the entry system performed its separation maneuver and adopted the required orientation. At 13:47 UT it entered the Martian atmosphere at 5.7 km/s at an entry angle of less than 10 degrees. The drogue parachute was deployed. This drew out the main parachute, which remained reefed until the speed became subsonic and the canopy could fully open. The heat shield was jettisoned and the low altitude radar was activated. At a height of 20 to 30 meters, falling at 60 to 110 m/s, the parachute was discarded and a small rocket lifted it away from the lander. Simultaneously, the lander fired its own retro-rockets. After a descent lasting a little over 3 minutes, Mars 3 touched down at 13:50:35 UT at a speed of 20.7 m/s. The landing site was at 44.9CS 158.0CW. in the planned area.

The foam cover was immediately ejected and the four petals opened. At 13:52:05 UT, 90 seconds after landing, the capsule began to transmit to its parent. However, after 20 seconds the transmission ceased and no further signals were received. It was several hours before the main spacecraft, which had to devote its attention to making the orbit insertion maneuver, was able to replay to Earth the transmission that it had recorded from the lander. The partial image returned by the lander is uninterpretable, being essentially noise. The only real information was an imaging calibration signal. The cause of this loss of signal may have been related to the planet-wide dust storm that was raging at the time. This would also explain the bland image lighting. It has been suggested that the transmitter failed due to coronal discharge in the dusty, low-pressure atmosphere. In any event, because the data collected during the descent was stored on board the lander for transmission in that first communication session this was lost as well.

Meanwhile a computer programming error caused the Mars 3 orbiter to cut short the insertion burn and it ended up in a 1,530 x 190.000 km orbit that had a period of 12.79 days instead of 25 hours. As a result there w ere only seven opportunities for periapsis observations during its limited operating life. As in the case of Mars 2, the inclination of the orbit was 49 degrees.

In the 4 month interval between December 1971 and March 1972 the two orbiters transmitted a large amount of science data. Mars 2 had the better orbit for planetary observations but, still suffering communications problems, its telemetry w as of poor quality and almost all of the planetary data were lost except radio occultations as the spacecraft crossed the planetary limb. The telemetry system on Mars 3 was working properly, although its impulse transmitter wras malfunctioning. Its orbit was ill-suited for planetary observations but Mars 3 was able to return useful planetary data. After the science observations finished in March, both orbiters continued to operate until contact was lost almost simultaneously in July 1972 when their attitude control gas ran out. The missions were announced to have been completed on August 22, 1972. by which time Mars 2 had made 362 о Г its shorter than intended orbits and Mars 3 only 20 of its exceedingly long orbits.

These spacecraft were highly sophisticated engineering marvels. They were the first of a new generation of large, complex spacecraft designed for comprehensive and bold investigation of our planetary neighbors. Their success on this initial outing led to a whole new generation of spacecraft for exploring the planets and conducting astrophysical investigations.

Results:

Orbiters:

Imagery

The Mars 2 and 3 orbiters suffered from a combination of circumstances. First, the telemetry systems had some problems. Very’ little telemetry at all was received from Mars 2. The Mars 3 impulse transmitter failed, and only lower resolution 250-line images were returned using the PCM decimeter band transmitter. Then there was the dust storm that began in October and had fully engulfed the planet by the time the spacecraft arrived. Third, the imaging sequences were pre-programmed, and with all but the very tallest mountain summits obscured imaging was impractical. Lastly the cameras had been set at the wrong exposure. And once the ampoules containing the chemicals to process the film were opened, the time available for photography was limited. Nearly all of the Mars 3 imagery was returned in four batches. The first two batches taken on 10 and 12 December 1971 showed very little detail due to the dust storm. Due to control system problems the next two batches were postponed to 28 February and 12 March 1972, by which time the dust stonn had abated. A total of 60 pictures were returned, including color images of volcanoes whose summits rose as high as 22 km and depressions as deep as 1.2 km, but the image quality was rather poor.

Only one picture was released during the mission, a relatively featureless view of the whole planet taken from the apoapsis of Mars 3’s extremely eccentric orbit. The imaging results of the Soviet missions paled in comparison to the 7,000 pictures that Mariner 9 provided, showing about 70% of the planet in unprecedented detail. The flood of orbital data from the American spacecraft revealed a much more interesting Mars than the dry, cratered. Moon-like perception created by the Mariner 4, 6 and 7 flybys. The canyons, dry river beds, flood plains and volcanoes imaged by Mariner 9 hinted at a much wetter past and raised the prospect of there being substirfaee w ater and maybe even life. The accomplishments of the Mars 2 and 3 orbiters were lost in the glare of Mariner 9. and the Soviets could only think about what might have been had they been blessed with a little more luck.

Dust storm

The dust storm abated in late January 1972 allowing the orbiting cameras a view of

the surface, but it was many months before the very light particles of dust settled out of the atmosphere. Dust clouds were found to extend to altitudes of 10 km, but were not evenly distributed around the planet. Dust particle sizes were determined, and small micron-sized dust grains were found as high as 7 km in the atmosphere during the storm. Bright ultraviolet clouds indicated the presence of even smaller particles at higher altitudes. During the dust storm the water vapor content of the atmosphere was very low, on the order of a few preei pi table mierons. After the storm the water vapor content increased to 20 microns, with greater humidity at the equator than in the northern polar region. The dust diverted a significant amount of sunlight, and the surface temperatures rose by about 25′ C after the atmosphere had cleared.

TIMELINE: 1986-1988

Bolstered with confidence as a result of the extremely successful Vega missions and leading the internationalization of robotic planetary exploration after the Americans had sidelined themselves, the Soviets decided to make another attempt at the Red Planet in 1988. As approved in 1976 by Mstislav Keldysh after the demise of the very ambitious rover and sample return proposals, this Lime the focus would be on the moon Phobos. The spacecraft would enter Martian orbit and after several weeks of orbital phasing during which it would study the planet, it would make a very slow – pass just 50 meters over the surface of Phobos to deposit two landers and undertake not only passive remote sensing by imagers and spectrometers but also active remote sensing with radar, ion beams and laser beams. In addition to the new power hungry active remote sensing instruments, the massive spacecraft would be equipped with a variety of other scientific instruments. Once again the Soviets invited the world’s scientific community to provide investigations for the mission, and this time even American instruments were accommodated.

The Phobos project was a model for international cooperation, but in the end also turned out to be a lesson in the international dissonance caused when such a mission fails. Phobos 1 and Phobos 2 were successfully launched in July 1988, but Phobos 1 was lost early in its interplanetary cruise owing to an elementary operational error. Phobos 2 reached Martian orbit and in just a few weeks conducted enough first-class observations of the planet to make up for all the flawed Soviet missions in the past, but then, just days prior to the close encounter with Phobos, the spacecraft failed to respond to a scheduled communications session and was lost.

W. T. Huntress and M. Y. Marov, Soviet Robots in the Solar System: Mission Technologies and Discoveries, Springer Praxis Hooks 1, DOl 10.1007/978-1-4419-7898-1 19,

© Springer Science+Business Media, LLC 2011

Launch date

1986

No missions

1987

No missions

1988

7 JuJ Phobos 1 Mars orbitcr Lost enroute

12 Jul Phobos 2 Mars orbitcr Failed in orbit before Phobos encounter