Category Soviet Robots in the Solar System

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

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-

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

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.

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

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.

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.

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

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

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

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

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

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

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

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

At the end of WW-II the US and USSR each captured German rocket scientists. V-2 rockets, and rocket development equipment. The captured engineers and technology enabled both nations to vastly accelerate their own rocket development work. The V-2 was flown many times in Russia and America. This famous rocket became the springboard for initiating a race between the two post-war superpowers to be first to build an ICBM capable of dropping a nuclear warhead on the other side. Building rocket defenses to counteract US strategic bombers, in the 1950s the USSR initially appeared to have the edge in this competition, wrhich led in turn to the perception in America of a "missile gap’. ‘This term was applied in several ways. Technically, it meant a gap in the range and ‘throw weight’ of a missile, but in some cases it was simply a measure of how many operational weapons each side had. and the Americans had an exaggerated view of the number of missiles actually pointed at them from inside the Soviet Union a misperception that delighted the Soviet government.

Despite starting at the same point in the late 1940s with captured V-2 rockets, the Soviet and American development programs took different paths towards an ICBM. In the early 1950s. the IJS had substantial advantages both in electronics technology and the ability to construct smaller, high-yield weapons. The fact that Soviet atomic devices were much heavier led to more pow erful rockets than those required by the Americans. The Soviets led in rocket mobility and deployment, in part because they assembled their rockets horizontally in production line fashion, and rolled them fully assembled on a railcar to the launching facility. The Americans built their rockets in sections and assembled them slowly on the pad by stacking them vertically one stage at a time.

The Soviet program culminated in the versatile R-7 two-stage rocket, which had its first successful test in August 1957. The kerosene and liquid oxygen propellants imposed a lengthy loading procedure prior to launch. The R-7 reached operational capability, but only five were ever deployed because by then the Soviets had learned howr to build smaller warheads and were developing a more suitable missile. It was rapidly replaced by smaller rockets that could be placed in hardened silos and loaded with storable propellants. The R-7 survived to serve as a space launcher because of its large "throw weight’, the mass that it could launch, and its versatility to use upper

stages for various military and civilian missions. It was called the Vostok’ launcher after the Gagarin flight, and became the base vehicle for Soviet lunar and planetary missions until superseded by the larger Proton vehicle in the 1970s. It is still in use today as the core vehicle for the Soyuz family of Russian launch vehicles.

The Venus objectives of the 3MV series were fulfilled when Venera 7 and 8 both survived landing and provided data on surface conditions. This led to the decision to design a new spacecraft for more extensive operations on the surface of this planet. For the first time since the initial launch to Venus in 1961, a launch opportunity was skipped in 1973 to devote the time to developing a new, heavier, more complex and capable Venus orbiter/lander system based on the Proton-launched Mars spacecraft. The success of the American Mariner 9 orbiter in 1971, and the anticipated superior capability of that nation’s Viking landers to be launched in 1975, led to the decision to focus the mote expensive Proton-launched missions on Venus rather than Mars in the immediate future.

The new Venera orbiter was nearly identical to the Mars orbiter, with changes in solar panel size and thermal design. But the entry vehicle was significantly different. The thicker, deeper atmosphere of Venus allowed for a simpler entry and landing system consisting of a large, hollow, spherical entry vessel containing the lander and parachute system. Pairs of spacecraft were launched on three launch windows, and all were successful, the Venera 9 and 10 orbiter /landers in 1975, the Venera 11 and 12 flyby/landers in 1978. and the Venera 13 and 14 flyby/landers in 1981. Not only did the Venera 9 lander provide the first imagery and composition measurements from the surface, the parent spacecraft was the first successful Venus orbiter. Flight energy requirements in 1978 and 1981 did not allow for orbiters. In 1983 the lander module was replaced with an imaging radar, and the Venera 15 and 16 orbiters were successful in providing the first radar imagery of the surface of the planet.

Nearly simultaneously with the 1983 radar mission, another flyby/lander mission was being prepared in a French partnership to deploy a large balloon which would be equipped with a comprehensive science payload and drift around the planet in the cloud deck. But when it was recognized that the flyby spacecraft could be retargeted to Halley’s comet after releasing the entry vehicle for Venus, the flyby spacecraft payload was redesigned for Ilalley, the balloon significantly descoped. a lander added, and the launch date adjusted to provide encounters with both Venus and Ilalley. Renamed Vega 1 and Vega 2, all aspects of these missions were carried out very successfully at both targets.

TIMELINE: JAN 1966-NOV 1968

The race to put humans on the Moon heated up in 1966. Both the Soviets and Americans stepped up the pace of their robotic lunar missions in support of eventual manned missions, with each successfully sending landers and orbiters. In September 1967 the Soviets began automated tests of a version of the Soyuz spacecraft that was intended to fly cosmonauts on circumlunar missions. In September 1968 the Zond 5 mission flew this spacecraft around the Moon and returned to Earth with a biological payload and high quality photographs of Earth from deep space. Its success spurred the American decision to send Apollo 8 into lunar orbit in the hope that astronauts would beat cosmonauts to the vicinity of the Moon. The Soviets lost their chance in November, when Zond 6 flew a repeat of the circumlunar mission and crashed on its return. This left the way clear for Apollo 8 in December, whose success nullified the propaganda value to be gained from sending cosmonauts on a circumlunar mission.

On their first launch of 1966 the Soviets were finally successful with their Luna soft-lander series. Luna 9 became the first lander on the Moon on February 3. It was followed in December by a second lander. Luna 13. Immediately after the success of Luna 9 an orbital module was hastily assembled to replace the lander module, and the spacecraft converted to support the call by the manned program for information on the lunar gravity field and surface properties. With the Luna 10 mission launched on March 31, 1966, the Soviet Union became the first nation to put a spacecraft into orbit around the Moon. Additional orbiters were sent, and by the close of 1968 this series was concluded with Luna 14.

On May 30,1966, America succeeded with its first attempt at a lunar soft landing. Surveyor 1 was a more sophisticated lander than Luna 9, as were the spacecraft of its Lunar Orbiter series, the first of which was successfully inserted into orbit around the Moon in August 1966. The purpose of the US landers and orbiters was precisely

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 10,

© Springer Science+Business Media, LLC 2011

the same as their Soviet counterparts, to determine the properties of the Moon and to identify candidate landing sites for manned missions. All five Lunar Orbiters were successful, as were five of the seven Surveyor landers.

In the midst of all this lunar activity, the Soviets had their first fully successful planetary mission, launched as Venera 4 on June 12, 1967. A second launch on June 17 was unsuccessful. Venera 4 entered the atmosphere of Venus on October 18 and transmitted atmospheric data while descending by parachute until it fell silent at an altitude that was still far above the surface. The next day America made its second successful flyby mission of Venus, with Mariner 5.

Both the nations ignored the Mars launch opportunity in 1967. The US could not afford as many launches as the Soviets and focused on Venus in 1967. The Soviets, frustrated by their six failures at Mars, by their continuing difficulties with the 3MV spacecraft, and by the Mariner 4 flyby in 1965. decided to forego the 1967 window and create a more complex Mars spacecraft that would be dispatched by the Proton launcher on a mission to land on the planet. In contrast, the tantalizing performance of Venera 2 and 3 ensured that this ЗМ V line would continue for Venus, and after the success of Venera 4 it became evident that a landing on the surface of that planet was achievable.

Temperature and pressure profiles of the upper atmosphere were obtained from the frequent Mars 2 radio occuRations, as was data showing there to be a neutral upper atmosphere made almost entirely of carbon dioxide with about 2% atomie oxygen. A night-side airglow was detected 200 km beyond the terminator. The base of the ionosphere was at 80 to 110 km. From 100 to 800 km, the carbon dioxide became increasingly dissoeiated into atomie oxygen and carbon monoxide. And the Lyman – alpha experiment detected an atomic hydrogen corona out to 20,000 km. There was less atomic hydrogen than Mariner 6 and 7 had found in 1969. presumably because the hydrogen atoms derived from w ater dissociation and during the dust storm there was less water present in the atmosphere. Charged particles were measured in the ionosphere, and the bow shock that defines the interaction of the solar wind with the planetary ionosphere was detected.

Lower atmosphere

The data obtained included temporal and spatial changes of temperature in the low er atmosphere, and waiter vapor concentrations 5.000 times weaker than in the Earth’s atmosphere. Images of the limb showed the layered structure of the atmosphere and its extent out to 200 km. Clouds were observed in the low er atmosphere composed of sub-micron particles at altitudes as high as 40 km. The composition was reported to be 90% carbon dioxide, 0.027% molecular nitrogen, 0.02% molecular oxygen.

0. 016% argon, and water vapor variable in the range 10 to 20 precipitable microns.

Campaign objectives:

The contrast between the Soviet space program in the mid-1970s and the mid-1980s was stark. In the mid-1970s the Soviets were in the deep shadow of the Americans. They had convincingly lost the race to land a man on the Moon. The Mars lleets of 1971 and 1973 were failures, whereas the Mariner 9 and Viking missions were total successes. In the 1970s the Americans launched inward lo Mercury and outward to Jupiter, Saturn, Uranus, and Neptune. The Soviets were confined Lo Venus. But in just 10 years all this was turned on its head. By the mid-1980s the US launch rate lor planetary missions had fallen to zero and their ambitious plans for Venus, Mars and Jupiter orbiters were troubled with delays and funding problems. By failing to fund a spacecraft for the Halley armada, the US had yielded the leadership it had worked so hard to achieve to Europe and Russia.

While the American program was struggling in the 1980s, the Soviet program was flourishing. Since 1975 they had achieved a string of unbroken successes at Venus with ever more complex spacecraft. They had taken the international lead from the US by opening the Vega missions to international participation on an unprecedented scale. Their Proton-launched spacecraft could carry a large, comprehensive payload of sophisticated and complex instruments. With landers and balloons at Venus and a gravity-assist to Halley the Vega mission had demonstrated boldness, ambition and success on a grand scale that silenced Western criticism of the quality of the Soviet program. The USSR was launching a space station. Mir, launching the largest rocket ever built, Energiya, developing a competitor to the US Space Shuttle, Buran, and launching almost 100 spacecraft per year. The National Geographic magazine gave a profile on the massive Soviet space industry and its successes, and Time magazine ran a cover story ‘Surging ahead Soviets overtake the US as the No. l spacefaring nation’. By the mid-1980s, therefore, Soviet confidence and optimism were running high.

After the Vega missions there was a consensus that the Venera series had run its course, having accomplished about all that it could. Of course there were ideas for long – lived landers and aerostats, but these were for the future. As the Vega missions were underway, the ‘Martians’ at IK I were insistent that it was now time to revive interest in Mars. The early proposals for a Mars sample return mission had fallen by the wayside, with even the plan designed to use the Proton launeher being discarded in 1977. Since then, however, the ‘Martians’ had been investigating more practical missions and had revived interest in a proposal to investigate Phobos. the larger of the planet’s two small moons. Strategic planning at IK I was still to seek to outdo the Americans, and since the US had not considered Phobos as a target there was scope for another Soviet showplace. They would not be repeating or competing with the US. After the Viking landers did not find any strong evidence for life on the surface of Mars the Americans had all but abandoned the planet. In addition, the continuing development of the Mars 2 to 7. Venera 9 to 16. and Vega 1 and 2 lines of heavy Proton-launched spacecraft provided the technology to undertake very capable Mars missions on a scale that the US was no longer able to fund. The lack of immediately compelling Venus missions, the lack of US interest in Mars, and the availability of proven technology, were compelling reasons to mount another campaign at Mars. So Soviet engineers decided to develop a new-gcncraiion planetary spacecraft based on the Venera-Vega heritage and to switch their attention from Venus to Mars. It was expected that the new spacecraft would serve as the baseline for the next 20 years of Soviet planetary exploration. The three Vega spacecraft spares were refurbished for astronomy missions in Earth orbit, two of which flew as Astron in 1983 and Granat in 1989.

The concept of a mission to Phobos actually preceded the Vega missions, but the latter had priority because their launches were dictated by the apparition of Comet Halley. The intention to send spacecraft to Phobos was first announced in November 1984, a month before the launch of the Vega missions. At that time the schedule was to launch in 1986. but this was slipped to 1988. Several mission scenarios had been considered including a Phobos landing, an outpost on Phobos for remote sensing of Mars, and a Phobos landing with sample return. The mission design selected was no less audacious, but less risky to arrange. The plan was to hover a large spacecraft about 50 meters above Phobos and conduct both passive and active remote sensing using lasers and particle guns. Later the mission expanded to include deploying two small landers for Phobos, one built by IKI and the other by the Vernadsky Institute. This campaign was to be the first in a Mars-focused exploration program which the Soviets expected to be every bit as successful as their exploration of Venus.

The objectives of the Phobos missions were to:

1. Conduct studies of the interplanetary environment

2. Perform observations of the Sun

3. Characterize the plasma environment in the Martian vicinity

4. Conduct surface and atmospheric studies of Mars

5. Study the surface composition and environment of the moon Phobos.

Having reaped major scientific and political results by the internationalization of the Vega missions, this policy was repeated for the Phobos campaign. This time the involvement of the US and Western countries was even more extensive. Many of the investigations and instruments were supplied by European countries; not only Soviet Bloc but also Western Europe. Given its history of cooperation in Soviet missions starting with the 1971 Mars missions. France was the primary Western contributor. The fact that the IJS supplied one instrument, a number of science co-investigators, and tracking support, was a major breakthrough initiated by the Soviet Union as the Cold War thawed. As a result, each spacecraft had a record number of instruments with which to perform a comprehensive scientific investigation of Mars and Phobos. In fact, each spacecraft carried twenty-four experiments provided by the USSR, the European Space Agency, and thirteen other countries including the IJS. The Phobos missions constituted an aggressive, innovative, and very impressive scientific attack on both Mars and Phobos, especially when compared to the modest orbiter that the IJS was then planning to send to Mars in 1990.

Spacecraft:

Orbiter:

The spacecraft was similar to the Vega design but represented a new generation in technology, and was the first major design change in Soviet planetary missions since Mars 2 and 3. It was the heaviest planetary spacecraft yet devised, with a total mass over 6,200 kg including 3,600 kg for the separable ADU propulsion system. The scientific payload capacity was an incredible 500 kg. It was constructed around a pressurized toroidal compartment for electronics, with four spherical outrigger tanks containing hydrazine monopropellant for the onboard propulsion system. The solar panels were mounted with their flat planes parallel to the toroid, perpendicular to the earlier Mars and Venus spacecraft. Attached to the outrigger tanks were twenty-four

engines rated at 50 N of thrust and four engines rated at 10 N for maneuvering, and a do/cn thrusters rated at 0.5 N for attitude control; and there were additional thrusters on the body and the solar panels. Two basic attitude control modes were provided: 3- axis control, and ‘drift mode’ control with the solar panels facing the Sun and the vehicle spin stabilized. The attitude control avionics consisted of Sun sensors, star sensors, gyroscope and accelerometer platforms, and a triply-redundant computer.

f or the Phobos mission this basic spacecraft had a cylindrical pressurized module mounted above the toroid. This module contained spacecraft control avionics, radio systems, and science experiments. It had a 2-axis 1.65 meter diameter parabolic high gain antenna on top. A 30 megabit data recorder was provided as a buffer. With the upgraded ground systems, the new 50 W transmitter was capable of 65 to 131 kbits/s from Mars. The old Venera control system was abandoned and a new dual-processor computer that incorporated a 4.8 gigabyte memory developed jointly with Hungary was to perform the complex maneuvers to match orbits with the target moon and make the flyby during which the spacecraft would undertake its closest observations and release its landers. Payload modules were mounted both inside and outside the cylinder. Unlike Vega, there was no scan platform; to make planetary observations the spacecraft had to be reoriented away from Earth-pointing. While cruising it would employ the drift mode in which it spun slowly with its solar panels pointed at the Sun.

Another innovation was the Autonomous Propulsion System (ADU) for the final escape burn, midcourse corrections, and orbit insertion. It was to be discarded after completing the major orbital maneuvers. This large assembly was carried below the toroidal section. It comprised eight tanks, four 1.02 meters in diameter and four 0.73 meters in diameter, nested around a single К I DIJ-425A restartable engine that could be throttled between 9.8 and 18.6 kN, burning UDMH and nitrogen tetroxide. It had a total burn time of 560 seconds. The ADU weighed 600 kg dry and was capable of carrying 3,000 kg of propellants. It w^as an extensive redesign of the main propulsion stage of the old Ye-8 lunar spacecraft. Part of the strategy was that the ADU would serve as a fifth stage of the launcher to achieve the desired interplanetary trajectory, and then perform the most energy intensive maneuvers at the target planet. It would eventually become the ‘Fregaf stage and be extensively used to augment the Soyuz launcher.

The Phobos mission strategy was to place the spacecraft into an initial elliptical near-equatorial orbit that maneuvers over several weeks would eireularize very close to that of Phobos. The ADU w ould be jettisoned, and from this intermediate orbit the moon would be approached several times at low velocity over a period of w eeks at distances varying from several hundred kilometers to 35 km. Observations of the moon would be used to refine knowledge of its orbital parameters sufficiently for a maneuver to be calculated which the onboard propulsion system would perform to produce the desired close flyby. The radar would be activated at a range of 2 km and the spacecraft w ould be controlled using the radar to perform a 20 minute pass at an altitude of about 50 meters at a relative velocity of 2 to 5 m/s. Considering the large topographic variation, this low’ altitude flyby would need a very robust and complex automated control system. Two types of lander w ere to be deployed, one of which
would be stationary and the other capable of‘hopping’ around in the weak gravity. Two active remote sensing experiments would be attempted using laser and ion guns in a manner that would enable the spacecraft to fly through the material evaporated off the surface for analysis with mass spectrometers. An active radar system would map the regolith material down to 2 meters depth. Passive remote sensing included both imaging and spectrometry. After the low pass over Phobos, the spacecraft was to adjust its orbit and spend its remaining time observing Mars. If the Phobos flyby were a success, there was the possibility of maneuvering the spacecraft to make a similar flyby of Dcinios.

Launch mass:

AD U propulsion system (Arbiter wet mass:

Instrument payload:

Mobile lander:

The PrOP-F hopper (‘Frog’) resembled a small, flattened ball consisting of a 50 cm diameter hemisphere on top of a semi-cylindrical base. It would be ejected from the side of the spacecraft at the time of closest approach, and fall to the surface in the moon’s weak gravity – some 2,000 times weaker than that of Earth. It was designed

to survive contact at a horizontal speed of.3 m/s and a vertical speed of 0.45 m/s. A damper truss would reduce the time taken to settle on the surface after initial impact. This would be ejected after settling, and a set of four long levers (‘whiskers’) on the base, two of which could be rotated, would deploy in order to orient the ball flat-side down. After about 20 minutes of sensing and the transmission of results, the hopper would flex its levers to propel itself to another location about 20 meters away. Each such hop would peak at a height of about 20 meters. The levers would right the ball after each landing. Ten hops were planned. The hopper was powered by a 20 amp – hour battery, and a 0.3 W transmitter would send to the spacecraft at 224 bits/s. Operations would begin immediately upon landing, and run for the 4 hour battery lifetime, by which time the spacecraft would have traveled about 300 km from the moon. Due to mass limitations the hopper was assigned only to Phobos 2.

Mobile lander mass: 50 kg

Payload: 7 kg

Stationary lander (DAS):

The stationary lander was carried on top of the spacecraft and was to be deployed at 2.2 m/s by a pair of arms. Once free, it would use cold gas thrusters to spin itself for stability at about 2 radians per second and also to propel itself towards the surface. It was designed to survive contact at a vertical speed of 4 m/s and a horizontal speed of at most 2 m/s. Contact probes underneath the lander were to ignite hold-down solid rockets, and simultaneously fire a harpoon on the underside down into the surface. The harpoon was tethered and nesting motors would wind up the tension to hold the lander firmly in place on the surface. After allowing the dust to settle for 10 minutes, the lander would extend three legs to raise its instrument platform 80 cm above the surface while maintaining the tension in the harpoon tether, and deploy and point its solar panels and antenna. Because Phobos rotates relative to the Sun, the orientation of the solar panels would be controlled by Sun sensors.

The lander was to undertake science investigations for 3 months, communicating directly with Earth by a transmitter and receiver at 1,672 MHz using antennas on the

instrument platform. It was realized that there might be shadowing problems on the solar panels, but there was insufficient time to develop mitigation options even using a secondary battery. Also, the Soviets and French developed different algorithms for data compression and because of insufficient capacity in the computer to implement both, the landers on the two spacecraft had different algorithms. The dual-processor computer was supplied by Hungary under the supervision of IKI. At a data rate that varied from 4 to 16 bits/s. three or four communication sessions would be needed to transmit a single image frame.

Stationary lander mass: 67 kg

Payload: 20.6 kg

The number of science instruments and the complexity of the spacecraft and its operations was unprecedented. A special MORION central interface was provided to handle the complexities. In the end, the mass budget was exceeded and some of the instruments had to be deleted from each spacecraft. The PrOP-F hopper was deleted from Phobos 1, as were the TERMOSKAN and ISM infrared instruments. Phobos 2 lost the RLK radar, the TEREK solar telescope, and the 1PNM neutron detector. The complexity of the mission was daunting, both to develop with all of its international interfaces and to operate with all of its instruments competing for operational time including spacecraft targeting and data transmission.

Payload:

Ovhiter:

Phobos active remote sensing:

1. Laser mass spectrometer for elemental surface composition (LIMA-D. USSR-Bulgaria-Finland-FRC-DDR-Czechoslovakia)

2. Ion gun mass spectrometer for elemental surface composition (DION, USSR-Austria-Finland-France)

3. Radar system for subsurface structure and mapping (RLK, USSR), Phobos 1 only

The 80 kg laser mass spectrometer was to conduct active remote sensing. It would lire 150 laser pulses, each of 10 nanoseconds, to evaporate material in the uppermost 1 mm of the surface of Phobos. The mass spectrometer would analyze the ions in the resulting plasma cloud to provide an elemental analysis for ranges up to 100 meters. The 24 kg ion gun mass spectrometer was to lire krypton ions at the moon and then measure the ions scattered back from its surface. Between them the two instruments were to measure about 100 sites on Phobos. The 41 kg radar system was to operate after the landers had been deployed and after the remote sensing at closest approach was done. When the spacecraft had opened the range to 2 km, the radar would map the surface of the moon and sound its subsurface to a depth of 2 meters.

Photos and Mars passive remote sensing:

1. CCD camera and spectrometer for surface mapping at three wavelengths (VSK, USSR-Bulgaria-GDR)

2. Thermal infrared radiometer and ultraviolet-visible spectrometer for surface temperature, thermal inertia, stratospheric temperature and aerosol char­acteristics (KRFM, USSR-France)

3. Near-infrared mapping spectrometer for mineralogy and atmospheric structure (ISM, IJSSR-France); Phobos 2 only

4. Thermal infrared mapping radiometer for surface temperature mapping

(TERMOSKAN, USSR); Phobos 2 only "

5. Gamma-ray spectrometer for surface radioactive element content (CS-14, USSR)

6. Neutron spectrometer to search for water in surface layers (IPN M-3, USSR); Phobos 1 only

7. Solar oeeultation ultraviolet and near-infrared spectrometer for distribution of minor constituents and aerosols (AUGUSTE, USSR-France)

The 52 kg VSK imaging system comprised a spectrometer and three cameras with 288 x 505 pixel CCD arrays. There was a narrow angle camera with a clear filter across the range 400 to 1,100 nm, a wide angle camera with a blue-green filter (400 to 600 nm), and a wide angle camera with a near-infrared filter (800 to 1,100 nm). The solid-state memory supplied by East Germany could store over 1,000 images. Bulgaria provided the electronics and full assembly and test with the help of France, Finland and the US. The KRFM multi-wavelength instrument would measure the reflectivity and thermal properties of the regolith, optical properties of atmospheric aerosols, and temperature of the Martian stratosphere using ultraviolet, visible and thermal infrared wavelengths. The ISM near-infrared mapping spectrometer would obtain spectra in a single pixel of the surface, providing data on mineralogy and, for Mars, the column depth of carbon dioxide w hich w as related to the elevation of the surface. The pixel was scanned across-track, and along-track scan was by spacecraft motion. It would provide a 1,600 km ship map of the planet’s surface at a resolution of 5 km during early Mars operations and a resolution of 30 km later in the orbit for the Phobos encounter.

The 28 kg TERMOSKAN infrared multispectral imager was a new line-scanning photometer camera w ith better detectors than on the Venera 9 and Mars 5 missions, one of them cryogenic ally cooled with liquid nitrogen from a Stirling refrigerator for thermal wavelengths between 8.0 and 12.5 microns. The other detector was for the red and near-infrared between 600 and 950 nm. TERMOSKAN recorded thermal emission (essentially the temperature) of the surface in 512 x 3,100 pixel panoramas at a resolution of about 2 km. Only 384 of the 512 pixels were image data, the rest provided calibration information. The image width w as about 650 km and the length was on the order of 1,600 km at a resolution of about 1.8 km/pixel. Images could be presented as surface temperature, thermal inertia, and texture. The 18 kg AUGUSTE experiment used a combination of two spectrometers and an interferometer to obtain vertical profiles for o/onc, carbon dioxide, w-ater, and oxygen by observing the limb of the planet at orbital sunrise and sunset and measuring atmospheric absorption in the solar spectrum. The IPNM gas scintilla lion neutron detector would identify areas on Mars (or indeed Phobos) that had hydrogen atoms in the regolith that were almost certainly due to water. T his eould provide evidence of habitable areas on Mars. The CS-14 gamma-ray spectrometer was mounted 3 meters away from the spacecraft on the edge of one of the solar panels and was to measure the elemental composition of the surfaces of both Mars and Phobos.

Solar wind and the plasma environment of Mars:

1. Plasma scanning analyzer for ion composition and direction, electron distribution, magnetosphere structure and dynamics (ASPERA, Sweden – IJSSR-Finland)

2. Plasma wave analyzer for plasma density and frequency spectrum of plasma waves (APV-F. ESA-Poland-Czechoslovakia-IJSSR)

3. Flux gate magnetometer for Martian magnetic field (FCMM, USSR-GDR)

4. Flux gate magnetometer for Martian magnetic field (MAGMA. USSR – Austria)

5. Electrostatic analyzer for energy and angular distribution of ions and electrons (HARP, Austria-Hungary)

6. Electrostatic and magnetic analyzer for direction and velocity of protons, alpha-particles and heavy ions (TAUS. Austria-Hungary)

7. Energy, mass, and charge spectrometer for ion composition, energy

distribution and plasma structure (SOVIKOMS, USSR-Austria-IIungary – FRG) "

8. Low energy telescope for solar wind and cosmic rays (LET, USSR-IIungary-

ESA-FRG) ‘

9. Energetic charged particle spectrometer for low-energy cosmic rays (SLED. USSR-IIungary-Ireland-FRG)

Of the two llux gate magnetometers, MAGMA was derived from the instruments used by the Venera and Vega missions and FGMM was a new one from a IJSSR – Germany collaboration. Both were mounted on a 3.5 meter boom, MAGMA at the tip and FGMM a meter from the tip. The APV-F plasma wave instrument included a dipole antenna and Langmuir probe for electron fluxes, and two 10 cm spheres at a separation of 1.45 meters to measure electromagnetic waves and plasma instabilities. The ASPERA instrument had two spectrometers on a scanning platform to measure plasma properties around the entire spacecraft. LET could measure the flux, energy spectrum, and composition of the solar wand and cosmic rays from atomic hydrogen up to iron. It was to complement measurements from a similar instrument on ESA’s Ulysses mission, but that launch w? a$ delayed when the Space Shuttle w? as grounded after the Challenger accident. TAIJS was to measure the energy and distribution of ions in the environment of Mars. HARP would measure electrons and ions from eight different directions.

Solar physics and astrophysics:

1. Solar telescope to observe the solar corona in x-ray and visible light (TEREK, l J S S R – Czech oslo v a к і a); Phobos 1 only

2. Solar high precision photometer for solar oscillations (IFIR, Switzerland – Francc – ES A -1J SS R)

3. Ultraviolet photometer for solar extreme-ultraviolet monitoring (SUFR. USSR)

4. Solar x-ray and gamma-ray analyzer (RF-15, IJSSR-C/echoslovakia)

5. Gamma-ray burst monitor for high energy solar and galactic bursts. 100 keV to 10 MeV (VGS APEX, USSR-France)

6. Gamma-ray burst monitor for low energy solar and galactic bursts, 3 keV to 1 MeV (LILAS, USSR-France)

The 36 kg TEREK solar telescope had three sets of optics, one a coronagraph to view the corona in the visible and the others the whole Sun in different x-ray bands using CCD detectors. The plan was to make observations in conjunction with Earth – based telescopes in order to compose a 360 degree view of the Sun. A three-channel photometer was to precisely measure solar irradiance in order to detect oscillations. The SUFR photometer was to monitor the ultraviolet flux of solar irradiation. The RF-15 instrument was similar to those of the geostationary meteorological satellites operated by America, and would be able to view different hemispheres of the Sun simultaneously. The two gamma-ray burst monitors were on the tip of one of the solar panels. The high energy instrument would have some utility for measuring the composition of the surface of Mars.

DAS small stationary Phobos lander:

1. CCD camera for surface imaging and microstructure (France)

2. Alpha, proton and x-ray spectrometer for surface elemental composition (FRC)

3. Harpoon anchor penetrometer with accelerometer and temperature sensor

4. Seismometer for internal activity and structure

5. Sun angle position sensor for determination of libration (France)

6. VLB! celestial mechanics experiment for orbital motion (IJSA-IJSSR – France)

The French were major partners in providing instruments and operational support for the DAS. They supplied the CCD camera and the optical sensor for tracking the Sun in order to determine the libration motions of the moon, and were participants in the VLBI experiment. The harpoon was instrumented with a temperature sensor and accelerometer to serve as a penetrometer to determine surface properties during the anchoring operation. The seismometer on the first DAS to land on the moon would have the opportunity to record the arrival of its partner. Its sensitivity was sufficient to detect the hopping activities of the PrOP-F.

PrOP-F small mobile Phobos lander, Phobos 2 only:

1. X-ray fluorescence spectrometer for surface elemental composition

2. Magnetic susceptibility and electric resistance probes for surfaee properties

3. Dynamic penetrometer for surface mechanical properties

4. Temperature sensors for measurement of surface layers

5. Radiometer for surface thermal flux

6. Magnetometer for surface magnetic field and permeability

7. Gravimeter (pendulum) to determining gravity field during descent

8. Accelerometer for surface properties

Mission description:

Phobos 1

Phobos 1 was launched on July 7, 1988. Another precedent was set when the launch was attended by the press, the international group of scientists participating in the mission, and even a delegation of US military. The rocket was also adorned with advertising for Italian and Austrian steel companies! The spacecraft used its ADU to achieve interplanetary injection for Mars, and again on July 16 for its first midcourse maneuver. However, on September 2 the spacecraft failed to respond to a scheduled communications session. Attempts to re-establish contact in September and October were unsuccessful and Phobos 1 was abandoned on November 3. Prior to its loss, it had returned data from its solar physics, plasma, and cosmic radiation instruments.

An investigation followed immediately. The communications failure w7as traced to a software upload made on August 29. An error was discovered in a command that w7a$ intended to turn on the gamma-ray spectrometer. An omitted hyphen created an unintended command to deactivate the attitude thrusters. The loss of Sun-pointing left the vehicle free to tumble and it depleted its batteries. The erroneous command w7as part of a test program that w as encoded in software PROMs which had not been removed and replaced prior to launch owing to time pressure. This was extremely humiliating. For a spacecraft this complex, expensive, and international in scope it w7as unimaginable that there would not be sufficient operational checks in place to prevent something as simple as a human coding error. Adding to the embarrassment w7as a battle that summer between the Moscow7 and Yevpatoria control centers over responsibility. Moscow’ had been given responsibility, and Yevpatoria was to cheek everything for transmission. When Moscow provided this command on August 29 the Yevpatoria checking equipment w as out of order and so the command w as sent to the spacecraft unchecked. Compounding the problem was that the spacecraft had not been programmed to undertake its own checks and reject fatal commands. In the midst of the operations team’s fear of reprisal and anxiety over the loss of Phobos 1 and additional problems on Phobos 2, no one was shot as might have been the case in earlier times but the Yevpatoria commander lost his job.

Photos 2

Phobos 2 was dispatched towards Mars on July 12, 1988, and performed midcourse maneuvers on July 21 and January 23, 1989; the latier occurring six days from Mars and on the same day as its silent partner passed the planet. During the interplanetary cruise Phobos 2 experienced serious problems. Its primary transmitter failed, and it continued on a less powerful backup transmitter that reduced the data rate. Also, one of three independent attitude control processors in the onboard computer failed and a second was occasionally giving spurious results. The three-fold redundancy of the computer system required two of the three processors to function properly. If two failed, the remaining functional processor could be outvoted by the failed ones! This was a serious design Haw, and it would ultimately decide the fate of the mission. In spite of its mishaps, controllers had been able to operate the spacecraft nominally. The solar telescope instrument had pointing problems but nevertheless returned a fair amount of good data. The gamma-ray instruments detected hundreds of bursts and measured their fine structure. The various other solar plasma, solar physics, and astrophysics instruments all operated well.

The spacecraft fired its ADU near Mars at 12:55 UT on January 29, 1989, and successfully entered into orbit. This initial orbit was 876 x 80,170 km inclined at an angle of 0.87 degrees to the equator with a period of 77.91 hours. Observations of the plasma environment were made during this time. A burn on February 12 raised the periapsis to 6,400 km and increased the period to 86.5 hours. There was some anxiety when the spacecraft temporarily fell silent on February 14. The apoapsis was gradually reduced until the final ADU maneuver on February 18 almost circularized the orbit near 6,270 km, several hundred kilometers above the orbit of Phobos. This maneuver also reduced the inclination to 0.5 degree and trimmed the period to 7.66 hours, a few minutes longer than that of its target. The ADU was then jettisoned. All future maneuvers were to be made by the onboard propulsion system. Observations of both Mars and Phobos were made from this orbit while the final maneuvers were being planned to produce the Phobos encounter in early April.

High resolution images were taken during two relatively close passes of Phobos on February 23 at 860 km range and on February 28 at 320 km range. These enabled the orbit of the moon to be refined to an accuracy of 5 km. On March 7, the plane of the orbit was aligned to precisely 0 degrees. The orbit was trimmed twice more on March 15 and March 21 into a 5,692 x 6,276 km path that was nearly synchronous with the moon, with the separation varying periodically between 200 and 600 km. A third close pass at 191 km was made on March 25 with all passive remote sensing instruments operating to identify landing points for the two DAS landers. Sufficient data having been gained to design the maneuver that would produce a low pass over the surface of the moon, the encounter was scheduled for April 9.

In the meantime Phobos 2 was suffering additional degradation. Both the backup transmitter and another attitude control processor were experiencing malfunctions. Pictures and thermal images were taken of Phobos on March 26. The following day additional navigation images were returned during communications sessions at 8:25 and 12:59 UT. Each such session required the spacecraft to turn towards Phobos for

imaging and then to turn the high gain antenna back to Earth for transmission. At the next scheduled communication session (15:58) the spacecraft failed to respond. A weak signal, perhaps from the omnidirectional antenna, was detected between 17:51 and 18:03 but no telemetry was received. Analysis of the signals indicated that the vehicle had lost attitude control and started to tumble. Deprived of solar power, the spacecraft would die after 5 hours when its batteries drained. Efforts to re-establish contact were futile and the mission was officially declared lost on April 1 5.

A board of inquiry was established on March 31. The cause of the failure was attributed to the omission of fail-safe software capable of automatically dealing with onboard emergencies, most notably to orient the spacecraft to the Sun in the event of a critical power shortage. The most likely immediate cause was failure of the second attitude control processor. Other design failures cited included the failure to enable the surviving good processor to outvote two failed ones, and the failure to undertake command uplink checks. It was evident that the systems and software for this new UMVL spacecraft had not been developed to maturity prior to launch. In the view of some, a contributing factor was that IKI scientists had not been able to participate at the top level of project management with the engineers at NPO-Lavochkin who were charged with building the spacecraft, as they had in past missions. The Ministry of General Machine Building had deleted the ‘supervising science team" and assigned management of the mission exclusively to the manufacturer. It is interesting to note that in the US. scientists were generally excluded from critical project management decisions and often still are today Гог major missions.

These lost spacecraft were the first after an unbroken string of successes for the Soviet program starting with Venera 9 in 1975 and culminating with the fabulously successful Vega investigation of Comet Halley in 1986. It was a shock, and set off an acrimonious debate of blame between Soviet scientists and engineers in the full light of international attention. The team of international scientists was summoned to Moscow in May 1989 for a post-mortem in which the Deputy Director of Lavochkin delivered a fog of excuses unrelated to the spacecraft that fooled no one and angered everyone. He could not resort to the old Soviet habit of concealment, not admitting a fault with the system. However his colleagues and the scientists at IKI. including its Director, Sagdeev. were more forthcoming about the faults with the systems on the spacecraft and the reasons for their failures. The questions from the audience were angry and pointed, something to which the Soviets were wholly unaccustomed. The dejected community of scientists that departed Moscow left behind a demoralized Soviet project team.

The loss of these complex spacecraft, especially Phobos 2 shortly prior to the culmination of its mission, was a staggering disappointment for the Soviet side and a great loss to the international community of planetary scientists eager for the results. After two decades of near secrecy in pursuing their planetary exploration program, the Soviets had opened it up for the Vega campaign, complete with international participation and coordination with missions by other nations. The Vega experience was a watershed for the Soviet program and brought positive results and widespread recognition. Built on the same implementation model, the Phobos campaign had the potential to propel the Soviets to an insurmountable lead in international planetary exploration. The abrupt and humiliating end of the missions brought a major crisis in confidence both internally and externally and presaged a swift decline in the Soviet planetary exploration program as the Soviet Empire came to an end in 1991.

Results:

Despite its premature loss, Phobos 2 provided a significant scientific return both on Mars and on Phobos. The mission failed to meet its primary objectives at Phobos, but during its 2 month lifetime as an orbitcr the spacecraft returned more data than all previous Soviet Mars missions combined. Furthermore, this data was of a quality never before obtained for the planet.

Thirty-seven images of Phobos were obtained with a coverage and resolution that complemented the Mariner 9 and Viking results. About 80% of the surface of the moon was documented. Thermal inertia values for the surface were determined from the thermal emission spectrum, highlighting some inhomogeneities. The reflectance from the ultraviolet through the near-infrared (i. e. 0.3 to 0.6 and 0.8 to 3.2 microns) at a spatial resolution of 1 km appeared to indicate a more carbonaceous chondrile composition than a water-rich chondrile, and also indicated strong inhomogeneities on the surface. Perturbations of the spacecraft’s orbit by Phobos provided a value for the mass of the moon which, together with a volume derived from imagery, gave a density in the range 1.85 to 2.05 g/cc. This was a low value even for a primitive volatile-rich meteorite, and suggested that Phobos might be more porous, or contain more ice, than expected. The temperature of the sunlit surface was 27°C. There were enticing indications from the magnetometer during close passes that the moon might possess a weak magnetic field.

Mars:

Phobos 2 had an array of instruments for multispectral investigations. Limb-to-limb photometric profiles were obtained using a prism spectrometer in the visible at about 25 km resolution, with matching profiles in the thermal infrared (5 to 60 microns) at about 50 km resolution. The data from these two instruments provided information on clouds and aerosols in the atmosphere and temperature profiles at altitudes in the 10 to 30 km range. Information was gained on optical depths, the sizes of particles, and the vertical distribution of aerosols. The occultation spectrometer saw a day-to­day variability in the atmospheric ozone profile, and a large altitude variation in the water vapor mixing ratio. Water vapor vertical profiles between 20 and 60 km were measured for the first time. The vapor content of the atmosphere was only 0.005%. Mars appeared to be losing its atmosphere at a rate of 2 to 5 kg per second, which

was fairly significant given the low atmospheric mass; this was equivalent to losing a global ocean to a depth of 1 to 2 meters.

Imaging of the surface of Mars was conducted at thermal wavelengths (8 to 12 microns) with about 2 km resolution using pixel-by-pixel coincident visible-infrared (0.5 to 0.95 micron) images, and the near-infrared mapping spectrometer provided multispectral images at spatial resolutions between 5 km at the low periapsis of the initial orbit and 30 km at the higher altitude of the circular orbit. The thermal data covered most of the equatorial region. The thermal imagery was less sensitive to the atmospheric haze, and showed the same surface features as in the visible range but with sharper contrast, emphasizing the pervasiveness of atmospheric haze on Mars. The measured surface temperatures ranged from -93°C to + 30′ C. Surface thermal inertia from the passage of the shadow of Phobos indicated that there must be a good insulating material down to a depth of 50 microns and poorer insulation below. The near-infrared mapping spectrometer provided data for most of the major geological formations on Mars except the polar regions. Data on the mineralogy of the surface and its local variations revealed a more pronounced variability than was apparent in Earth-based studies. Two bands were of particular interest, the 3.1 micron band of hydrated minerals and the 2 micron band of carbon dioxide, and some contour maps were derived from carbon dioxide measurements.

Gamma-ray spectroscopy conducted during the first four low periapsis passages in the initial orbit provided data on bulk elemental abundances that were consistent with results from Mars 5 and also the x-ray fluorescence measurements made by the Viking landers. Due to its extensive instrumentation, Phobos 2 was able to make a detailed survey of the plasma environment around Mars and how this interacted with the solar wind. No permanent intrinsic planetary magnetic field was measured, even at the low periapsis of the initial orbit. The SLED detector showed levels of radiation in orbit around Mars to be less than limits that would be dangerous to humans.