Category Soviet Robots in the Solar System

As early as 1959 the Soviets had a plan for manned circumlunar flights. When the Americans decided in mid-1961 to go to the Moon, Korolev was already designing the Soyuz spacecraft for these missions. It was the same three-module arrangement with w’hich we are all familiar, with a support module containing all the resources required for power, propulsion, communication, navigation and consumables for the cosmonauts, a descent module to carry them aloft and to return them to Earth, and a compartment to provide more room for the cosmonauts on long flights. After the Vostok and Voskhod manned capsules, this system was introduced and remains the reliable Russian system still in use today.

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Figure 5.4 Soyuz 7K-L1 ‘Zond’ circumlunar spacecraft (from Space Travel Encyclope­dia).

The Soyuz 7K-L1 was a version of the 7K-LOK lunar orbital spacecraft modified to perform a circuit! lunar mission. Although the three-stage R-7 used for Soyuz flights in Earth orbit was replaced by the more powerful four-stage Proton, mass limitations meant that the 7K-L1 did not have the ‘orbital’ module and was designed to carry only two cosmonauts. The idea was to fly circumlunar missions with two astronauts using the 7K-L1 as a precursor to performing a lunar landing using the 7K-LOK version of the Soyuz (which would have an orbital module) and the LK lunar lander, all launched by the massive N-l rocket. To prepare for the manned circumlunar missions, several automated flights of the 7K-L1 were conducted, the first two in Earth orbit and then nine others over the years 1967-1970 either to lunar distance or performing actual circumlunar flights. Zond 4 reached lunar distance before returning to Earth, but in a direction away from the Moon in order to simplify navigation, and Zond 5 to 8 each made circumlunar flights. Zond 4 self – destructed on re-entry, Zond 5 had significant but non-fatal problems with on board systems, and Zond 6 crashed on landing only a few weeks before the Apollo 8 mission. Although Zond 7 and Zond 8 were complete successes, the Soviets never used the system for a manned circumlunar flight.

After the failures of the 1M Mars missions in October 1960 and the 1VA Venus missions in February 1961, Korolev resolved to develop an improved, second

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 8,

© Springer Science+Business Media, LLC 2011

Launch date

1961

23 Aug Ranger 1 lunar mission test

18 Nov Ranger 2 lunar mission test

1962

26 Jan Ranger 3 lunar hard lander

23 Apr Ranger 4 lunar hard lander

22 Jul Mariner 1 Venus flyby

25 Aug Venera entry probe

27 Aug Mariner 2 Venus flyby

1 Sep Venera entry probe

12 Sep Venera flyby

18 Oct Ranger 5 lunar hard lander

24 Oct Mars flyby

1 Nov Mars 1 flyby

4 Nov Mars entry probe generation planetary spacecraft. In the spring of 1961 lie directed that a new multi­mission spacecraft be designed that could be configured for either flyby or entry probe missions at either Mars or Venus. This new scries was the first modular interplanetary spacecraft, with a standardized multipurpose ‘orbital’ module (in the US vernacular this was a carrier vehicle) to guide the spacecraft to either planet, and a separate module to carry a science payload tailored to the planet and mission type. Two standard types of science module were provided, the first a pressurized vessel to accommodate instruments for studying the planet during a flyby, and the second an entry vehicle for atmospheric probe or lander missions. For the latter, the entry vehicle was detached at arrival and the carrier vehicle discarded and left to burn up in the atmosphere. This was a major improvement over the 1VA design, where the probe was retained and simply expected to survive the destruction of the spacecraft on entry.

The communications, attitude control, thermal control, entry, and propulsion systems were much improved over the 1M and 1VA spacecraft. This new generation set the design precedent for all Molniya-launched planetary missions until the more capable Proton launcher was introduced. The initial design, designated 2MV, lasted only for the 1962 Mars and Venus opportunities. Six were built and launched, three for Venus and three for Mars. Korolev also upgraded the 8K78 launcher to lift these heavier spacecraft by improving the strap-on booster engines and lengthening both the interstage between the third and fourth stages and the aerodynamic shroud. Only one 2MV survived its launch vehicle, Mars 1. After the 1962 campaign, the design was upgraded to produce the 3MV.

For the 1962 launch opportunity for Venus the Soviets prepared two 2MV-1 entry probe spacecraft and one 2MV-2 flyby spacecraft. The mission of the entry probes was to penetrate belowr the veil of clouds, survive landing, and return data profiling

2MV-1 No.3 [Sputnik 19]

Venus Atmosphere,’Surface Probe

USSR ОКБ-1

Molniya

August 25, 1962 at 02:18.45 UT (Baikonur)

Failed to leave Earth orbit, fourth stage failure,

2MV-1 No.4 [Sputnik 20]

Venus Atmosphere.’Surface Probe

USSR ОКБ-1

Molniya

September E 1962 at 02:12:30 UT (Baikonur)

Failed to leave Earth orbit, fourth stage failure,

2MV-2 No. l [Sputnik 21]

Venus Flyby USSR ОКБ-1 Molniya

September 12, 1962 at 00:59:13 UT (Baikonur)

Failed to leave Earth orbit, third and fourth stage failures.

the temperature, pressure, density, and composition of the atmosphere, and then the composition of the surface. The flyby spacecraft would photograph the planet using an upgraded version of the camera originally intended to fly on the 1M spacecraft. Frustratingly, all three spacecraft would be lost to failures of the launcher’s fourth stage.

Spacecraft:

The 2MV spacecraft was 1.1 meters in diameter and 3.3 meters long in total, and measured 4 meters across the solar panels with the thermal radiators deployed. It was divided into two attached parts. The main spacecraft, known as the ’orbital’ (or carrier) module, was 2.7 meters long including the 60 an long propulsion system at one end. The carrier s pressurized compartment contained the flight system avionics and scientific instrumentation. The propulsion system used cold gas jets for attitude control and the KDU-414 gimbaled engine that delivered a thrust of 2 kN and was capable of more than one midcourse correction, for a total firing time of 40 seconds. Attached to the other end of the carrier module was either a 60 cm long pressurized flyby instrument module or a detachable 90 cm diameter spherical entry probe.

In addition, significant changes were made to the communication system. A 1.7 meter parabolic antenna and radio system transmitting at either 5, 8 or 32 cm was provided for high rate communications during the interplanetary cruise and for data transmission on arrival at the target. A separate omnidirectional antenna and meter

band transmitter was added to supplement the decimeter high gain directional unit. Commands were received at 39 cm (768.96 MHz) using semi-directional antennas attached to the thermal radiators, which were also used for transmission at 32 cm (922.776 MIIz) in the vicinity of Earth and at a reduced rate at longer range in the event of an emergency. A backup 1.6 meter band radio was provided for operations

near Earth at 115 and 18.3.6 MHz using whip antennas mounted on top of the solar panels. For flyby missions, the camera in the instrument module had its own 5 cm band impulse transmission system. The high gain antenna was fixed, pointing in the opposite direction to the solar arrays. To use it, the spacecraft was required to adopt Earth-pointing attitude. An onboard tape recorder was provided to store data while Sun pointing and to replay it when the high gain antenna had locked on. The power supply system consisted of 2.6 square meters of solar cells that supplied 2.6 kW to a 42 amp-hour NiCd battery array.

The attitude control system was upgraded by providing an Earth sensor for high gain antenna pointing, instead of using a radio bearing. And based on the Venera 1 experience, the Sun, Earth, and star sensors were repositioned inside the controlled environment of the carrier module, looking out through a quartz window dome. New’ and more reliable sequencers were used, with element-by-element redundancy. As with Venera 1, several orientation modes were provided. While cruising, ihe spacecraft was to maintain a low-precision З-axis Sun pointing mode in order to keep the solar panels illuminated. To avoid losing control of the spacecraft as with Venera 1, it was decided never again to turn off the receivers during the cruise phase of a planetary mission. For high gain transmission sessions, the spacecraft would terminate solar panel Sun pointing and reorient itself to high gain antenna Harth pointing using Sun and Earth optical sensors in conjunction with gyroscopes for precise attitude eontrol. For mid course maneuvers, the required engine orientation relative to the Sun and the star Canopus was controlled by the gyroscope system. The gyro stabilization system also provided feedback to adjust the angle of the engine during the burn and terminated the burn when integrating aeeelerometers detected the specified velocity change. The orientation of the spacecraft during its planetary encounter would be controlled using optical sensors associated with the imaging system in the instrument module.

Thermal control was improved by abandoning the motorized shutters in favor of a binary gas-liquid thermal control system that had two liquid hemispherical radiators mounted on the ends of the solar panels. Separate heating and cooling lines carrying different liquids were coupled by heat exchangers to the dry nitrogen circulating in the interior. The spacecraft were also covered with metal foil and insulating blankets of fiberglass cloth that do not show7 in the available photographs.

The carrier module had instruments for cruise science, measurements in the near­vicinity of the planet and, on entry missions, in the ionosphere prior to destruction. For entry missions, the carrier module deployed the probe by a command from Earth that triggered a timer just prior to entering the atmosphere. Pyrotechnic charges w ere fired to release the restraining straps and the entry probe w as ejected by a spring-like mechanism. The entry system was a 90 cm diameter sphere protected by an ablative aeroshell material. In addition to the science instruments, the entry probe contained a three-stage parachute system, silver-zinc batteries, and a decimeter band radio with a semi-directional antenna for direct transmission to Earth. Based on the best guess of surface conditions at the time, the probes were designed to survive pressures up to 5.0 bar and temperatures up to 77 C. The Venus and Mars probes were almost identical, but those for Venus had thicker shells and smaller parachutes and w hereas the Mars probes were cooled by air circulation the Venus probes were cooled using a passive ammonia-based system. Unlike Venera 1, the new probes were chemically sterilized by being soaked in an atmosphere of 60% ethylene oxide and 40% methyl bromide in order to prevent biological contamination of the surface of their target on landing.

Launch mass: 1.097 kg (probe version)

— 890 kg (flyby version)

Probe mass: ^305 kg

Payload:

Carrier spacecraft:

1. Magnetometer to measure the magnetic field

2. Scintillation counters to detect radiation belts and cosmic rays

3. Gas discharge Geiger counters

4. Cherenkov detector

5. Ion traps for electrons, ions and low energy protons

6. Radio to detect cosmic waves in the 150 to 1,500 meter band

7. Micromctcoroid detector

This list is for the Mars 1 payload, and it is assumed here that the carrier modules of all the 2MV series were similarly instrumented. The magnetometer was mounted on a 2.4 meter boom, and ribbon antennas were extended for the cosmic wave radio detector. Starting with these 2M V spacecraft, piezoelectric micromctcoroid detectors with a total area of 1.5 square meters were attached to the rear of the solar panels.

Descent! landing capsule:

1. Temperature, pressure and density sensors

2. Chemical gas analyzer

3. Gamma-ray detector system to measure radiation from the surface

4. Mercury level wave motion detector

The chemical gas analyzer consisted of simple chemical test cells, precursors for the proper chemical lest instruments that would be flown on later missions. Platinum wire resistance thermometers were utilized, and the density gauge was an ionization chamber for measurements in the upper atmosphere where the pressure was less than 10 millibars.

Vlyhy instrument module:

Instrumentation was probably the same as the Mars flyby module with the exception of the infrared spectrometer, which for Venus was designed to study the atmosphere instead of the surface.

1. Facsimile imaging system to photograph the surface

2. Ultraviolet spectrometer in the camera system for ozone detection

3. Infrared spectrometer to study the thermal balance of the atmosphere

The imaging system was complex and heavy. It weighed 32 kg and was mounted inside the pressurized instrument module, peering out through portholes on the end. It focused 35 mm and 750 mm lenses on 70 mm film with a capacity of 112 images, alternately shot with square frames and 3 x 1 rectangular frames. Individual frames could be scanned or rescanned at 1,440, 720, or 68 lines and stored on wire tape for later transmission. An ultraviolet spectrograph projected its spectrum onto the film alongside the images. The imaging system had a dedicated 5 an (6 GHz) impulse transmitter housed inside the instrument module. This transmitter would issue short 25 kW pulses with an average power output of 50 W. The transmission rate was 90 pixels/second, requiring about 6 hours to transmit a high resolution image of 1.440 x 1,440 pixels. The pixels were probably encoded as analog pulse position rather than as binary values. The infrared spectrometer was on the exterior of the instrument module and bore-sighted with the camera.

Mission description:

All three missions were lost to fourth-stage failures after successful insertion into parking orbit. On the 2MV-1 No.3 mission, only three of four ullage eontrol solid rocket motors on this stage fired, causing it to somersault after 3 seconds. The main engine did ignite, but because of the tumbling motion it burned for only 45 of the planned 240 seeonds. Several pieees were left in orbit. On the 2M V-l No.4 mission, a stuck valve blocked the fuel line and the fourth stage failed to reignite.

The 2MV-2 No. l Venus flyby spacecraft was lost due to a violent shutdown of the third stage. An engine in the third stage exploded at shutdown because the LOX valve did not close, continuing to feed LOX into the combustion chamber. The third stage broke up into seven pieces. The fourth stage eontinued into parking orbit, but the tumbling imparted to it by the destruction of the third stage induced cavitation in the oxidizer pump which caused the engine to shut down less than a second after it was reignited for the escape burn.

Results:

None.

Campaign objectives:

Since their origins in 1960 the Soviet Mars and Venus programs had been strongly intertwined, using slightly different versions of the same spacecraft. When NPO – Lavochkin took over the planetary program it set out to transform OKB-l’s 3MV-3 design into a 1,000 kg spacecraft to be launched on an upgraded Molniya-M at the 1967 flight opportunity to Mars. But this approach was soon abandoned. The Mars program had been a disaster. Seven attempts in the period 1960 through to 1964 had failed, including one test mission. Then the Zond 2 Mars flyby spaeeeraft created an embarrassment by failing as Mariner 4, launched by the US at almost the same time, went on to make a successful flyby in July 1965. In that same month Zond 3, after operating successfully at the Moon, failed its Mars deep space test flight objectives. Aware that the US was turning away from Venus in favor of Mars, starting with dual flybys planned in 1969 and with orbiters and landers to follow, perhaps as early as 1973. the Soviets decided to perfect a Mars lander that would outdo the American flyby missions.

The entry vehicle for the 3MV Mars spacecraft had been designed in the early 1960s on the presumption that the atmospheric pressure at the surface was between 80 and 300 millibars. The Mariner 4 flyby in July 1965 showed it to be a mere 4 to 7 millibars. The design of the 3MV entry probe was therefore fatally flawed. A new technique would be required to perform entry, descent and landing in such a rarefied atmosphere. In October 1965 NPO-Lavochkin abandoned the 3MV for Mars, but retained it for Venus because it was suitable for that dense atmosphere. The Soviets skipped the 1967 Mars launch opportunity to develop a more capable spacecraft for the 1969 opportunity.

The powerful Proton launch vehicle made its debut in 1965. It doubled the mass that could be delivered to low Earth orbit compared to the three-stage Molniya. and when augmented by the Block D fourth stage (as the Proton-K) it facilitated a whole new generation of heavier, more capable and complex lunar and planetary spacecraft than the Molniya-launched 3MV. Capable of dispatching over 4 metric tons onto an interplanetary trajectory, the Proton-K became the standard launcher for lunar and Mars missions after 1966, and for Venus missions after 1972.

The engineering requirements for new Mars and Venus missions during the time period 1969 73 were defined in March 1966 by the head of NPO-Lavochkin, Georgi Bab akin:

1. Use of the Proton-K to achieve parking orbit and escape onto an interplanetary trajectory

2. IJsc of a "universal” multi-purpose, modular on board propulsion system for trajectory correction while coasting and then insertion into an orbit around the target with a pericenter about 2,000 km and apocenter not exceeding

40.0 km

3. Use of descent-from-flyby and descent-from-orbit mission designs for soft landers to place instruments on the surface

4. Use of the main spacecraft as either a flyby vehicle or an orbiter to relay information from the lander at about 100 bits/s to the Earth

5. IJsc of a telemetry system capable of transmission from the main spacecraft of about 4,000 bits/s.

It was decided that in addition to trajectory correction maneuvers, entry vehicle targeting and planetary orbit insertion and trim maneuvers, the universal propulsion system should also participate in establishing the desired interplanetary trajectory by firing after the spent Block D stage was jettisoned.

These requirements were not applied to Venus until the successful Venera type of the 3MV had fulfilled all of the objectives for that planet in 1972, but they were applied immediately to Mars for the 1969 opportunity. Also, it was decided that for the initial Mars mission the descent module would be an atmospheric probe to obtain the data required for designing a landing system for that rarefied atmosphere. Another key objective was to improve the ephemeris for Mars for future missions. The science objectives for Mars missions using this new spacecraft system were: [1] [2] [3] [4] [5] [6] [7] [8] [9] opportunity that was only 33 months away, an incredibly short period of time in which to try to develop a spacecraft of such an unprecedented complexity. And by devoting part of this time to modifying the 3MV to score a success at Venus in 1967 they left themselves with only 20 months to develop the new spacecraft. Then problems with the design left them with only 13 months. Given the intense pressure to outdo the US at Mars, the risks taken were enormous.

The workload was intense during the last years of the 1960s as the Soviets tried to compete with Apollo. NPO-Lavochkin was overloaded developing the Luna rover and sample return missions, continuing to milk the successful Venus missions, and making a valiant effort on M-69. This was a brand new spacecraft like none built before, and the rushed development showed. Nothing went smoothly. The spacecraft suffered from the same development problems as OKB-l’s early rushed designs and engineers were not terribly optimistic about its chances. The winter of 1968-69 was exceedingly harsh, pipes burst and heating systems failed, creating near-impossible working conditions. Control and telemetry systems were plagued with troubles and the design of the spacecraft actually prevented easy access for servicing. The entry probe had to be deleted very late in the process due to insufficient time and system mass growth, and was replaced by a compartment for additional orbital instruments.

The Soviets were to fail in their first attempt with this new spacecraft in 1969, but the engineering and science requirements for the M-69 program set a precedent for all of the Mars mission designs that were to follow7. At that time almost nothing was known of these missions in the West, and 30 years would elapse before they were described in any detail.

Spacecraft:

The initial design:

As Babakins engineers worked with their OKB-1 colleagues in 1966-67 to prepare a 3MV spacecraft for what would become the successful Venera 4 mission, others at NPO-Lavochkin were working on a new7 spacecraft for the Luna series that would be launched by the Proton-K instead of the Molniya. Unlike the previous 2MV, 3MV and Luna series spacecraft where the avionics compartment was the main structural element, this time a quartet of spherical propellant tanks connected together in the shape of a square using cylindrical inter-tank sections became the element on w7hich everything else was mounted.

Given the short period of time available for the development of a Proton – launched Mars spacecraft, it was decided to exploit this work. The initial M-69 design had the entry probe attached to the tank assembly where the lunar rover w7ould otherwise be carried, and the remaining systems attached to the underside’. The two solar panels were spread out from opposite sides of the square, and the antenna and engine were opposite each other on the remaining sides. This design could meet the schedule, but was not easily reconfigured and failed to satisfy some of the requirements. Also, the designers struggled with a number of engineering

problems in trying to adapt a lunar spacecraft for Mars exploration. The main issues centered on the fundamental tank design, and ultimately it was abandoned, forcing a total redesign 13 months before the launch date.

The final design:

The new design used a single large spherical tank at the center of the spacecraft as the main structural element. The tank had an internal baffle to separate the UDMH fuel from the nitrogen tetroxide oxidizer. The Isayev engine was attached to the base of the tank. A cylindrical interstage with a pressurized container for electronics was attached to the top of the tank, and the entry vehicle was installed above that. Two hermetically sealed cylindrical modules were attached on opposite sides of the tank, one for communication, navigation systems and optical orientation sensors, and the other for science instruments including the cameras. There were also science sensors attached to the outside of ihe spacecraft.

The antenna system, including both a large high gain and small conical antennas, was affixed to the cylindrical interstage. The two 3.5 square meter solar panels were mounted outboard of the instrument modules. The panels were supplemented with a NiCad. battery that delivered power at 12 amps with a 110 amp-hour capacity. Both passive insulation and active thermal control vrere employed. The active system operated in the pressurized compartments and consisted of a ventilation and air circulation system to route air between two radiators, one exposed to sunlight and the other to shadow. The thermal control radiators were inboard of the solar panels, between the modules across the main tank. The avionics of the M-69 spacecraft were

much improved over the 3MV series. It was the first Soviet planetary spacecraft to carry a computer. .An advanced data processing system weighing only II kg was provided that could program the instruments and acquire, process and compress the data from both engineering and science systems for transmission to Earth.

A new telemetry system was provided that consisted of a transponder-receiver for

non-imaging data and an impulse transmitter for images, a 2.8 meter parabolic high gain directional antenna and a trio of low gain semi-directional conical antennas for decimeter and centimeter bands. The arrangement of the conical antennas was such that when the solar panels were pointed at the Sun, they would be pointing at Earth. The transponder-receiver had two transmitters and three receivers in the decimeter band at 790 to 940 MHz with 100 W of power, and facilitated Doppler tracking at a transmitted data rate of 128 bits/s with 500 data channels. These transmitters and receivers could use either the conical antennas or the high gam. One receiver was always on and connected to one of the conical antennas for continuous reception. The remaining receivers and the transmitters were cycled through these antennas by timers in order to ensure the reliability of the system. As part of the payload, a new film camera system with facsimile processing was developed. The imaging system had a 5 cm impulse 50 W transmitter for a data rate of 6 kbits/s using short pulses at 25 kW.

For the attitude control system, new Sun and star sensor systems and new nitrogen gas micro-engines were developed. There were two Sun sensors, two star sensors, two Earth sensors, and two Mars sensors. Nine helium-pressurized tanks provided nitrogen gas stored in ten separate tanks to eight attitude control thrusters,
two each for pitch and yaw and the other four for roll. The nitrogen lank pressure of 350 bar was regulated to 6 bar for maneuvering and 2 bar for attitude maintenance. During cruise and routine operations the vehicle used one set of sensors to maintain itself in a rough attitude that faced the solar panels towards the Sun. For high gain antenna operations, midcourse maneuvers, and orbital mapping, it used a more accurate set of sensors for precise З-axis stabilization. Both optical sensors and gyroscope control were provided for the altitude control system.

The entry system was a prototype of that which would be used in 1971, and was to have been deployed w hile 2 days from Mars. But it was ultimately deleted from the 1969 mission due to mass growth of the spacecraft and insufficient time to test the parachute descent system in balloon drops. The entry probe w as designed around a large spherical tank with three attached pressurized compartments. No other details are available.

Launch mass:

Or hi ter mass:

Probe mass:

Payload:

Or biter:

1. Facsimile imaging system (FTU)

2. Infrared Fourier spectrometer (UTV1) for atmosphere and surface studies

3. Infrared radiometer (RA69) for surface temperature

4. Ultraviolet spectrometer (USZ) for reflected radiation

5. Water vapor detector (I VI)

6. Mass spectrometer for ionosphere composition and hydrogen, helium detection (UMR2M)

7. Multi-channel gamma-ray spectrometer (GSZ)

8. Low – energy ion spectrometer (RIB803)

9. Charged particle deteetor (KM69) for solar electrons and protons

10. Magnetometer

11. Micrometeoroid detector

12. Low frequency radiation detector

13. Cosmic ray and radiation belt detector

14. X-ray radiometer

15. Gamma-ray burst detector

Total mass: 85 kg.

The new FTU was an advanced film facsimile imaging system consisting of three cameras, each with red, green and blue color filters. The image format was 1,024 x 1,024 pixels. One camera had a 35 mm lens, a second had a 50 mm lens and a field of view of 1,500 x 1,500 km, and the third had a 250 mm lens and a field of view of 100 x 100 km with a best resolution of 200 to 500 meters. The film was processed on

board, encoded digitally and supplied to the impulse transmitter. The film was to be chemically activated upon arrival at Mars in order to avoid damage by radiation in cruise. Each camera had sufficient film for 160 images.

Atmosphere probe (deleted):

1. Pressure sensors

2. Temperature sensors

3. Accelerometers for atmospheric density

4. Chemical gas analyzer

Total mass: 15 kg.

Mission Description:

The plan was to use the first three stages of the Proton and the Block D upper stage to achieve parking orbit. After one orbit, the Block D would be reignited for the first part of the escape sequence under the control of the spacecraft. After burnout of the Block D and separation, the spacecraft would fire its main engine for the final boost onto the interplanetary trajectory. This would be the first time that this new scheme was used, adding more risk to an already challenging project. The spacecraft engine w ould also be used for two trajectory corrections during the 6 month cruise to Mars, one 40 days out from Earth and the other 10 to 15 days prior to arrival. The fourth burn of the engine would be made at the closest point of approach to Mars in order to enter a 1.700 x 34,000 km orbit inclined at 40 degrees to the equator with a period of 24 hours. No immediate trim burns were planned, despite the expectation that the errors would be considerable. After some photography and other science from this initial orbit over several weeks, the periapsis would be lowered to about 600 km for an additional 3 months of imaging and data collection. At that point the mission was expected to be concluded.

Unfortunately, neither spacecraft even reached Earth orbit. М-69Л was lost to a third-stage explosion when a rotor bearing malfunction caused a turbopump to fail and catch fire. The engine shut down at the 438 second mark and the stage exploded. M-69B was lost when one of the six first stage engines exploded just at launch. The vehicle continued to climb on the five remaining engines until the 25 second mark, at which time it tipped over to the horizontal at an altitude of 1 km. The remaining engines shut down and 41 seconds into the flight the vehicle fell to the ground 3km from the pad and exploded. Remarkably, the second stage landed intact.

The failure of the Soviets to exploit the 1969 opportunity for Mars passed largely unnoticed in the West, mainly because the two attempted launehes failed so early in flight. But the Protons may have saved the Soviets from the larger embarrassment of another Mars mission failing due to the spacecraft being rushed too hard through its design and development. As one of its designers remarked, M-69 was an example of how’ not to build a spacecraft.

Results:

None.

The Proton was experiencing its worst period in development at this time, with a very high failure rate. It was responsible for the loss of many spacecraft including a large number of lunar missions. The failure of the M-69 launches was a bitter pill for the spacecraft team to swallow after all the difficult and frantic work that had gone into the preparation. To rub salt into the wound, soon thereafter the US achieved the Apollo 11 lunar landing and the successful Mariner 6 and 7 flybys of Mars.

TIMELINE: 1977-1978

With nothing left to accomplish on the Moon, and having abandoned Mars for the immediate future, Soviet scientists and engineers focused their robotie exploration solely on Venus. In 1978 they launched a second pair of spacecraft which were near duplicates ol" Venera 9 and 10. Because the energetics for this opportunity were less favorable, it was not practicable to send an orbiter/1 under and instead the lander was to be delivered by a spacecraft that would perform a flyby and relay to Earth the data from the entry system and lander. Although both of the Venera 11 and 12 landers touched down, they suffered a number of problems and in particular were unable to provide imagery.

The US also sent spacecraft to Venus in 1978, but these were very much smaller. The Pioneer 12 Venus orbiter was an outstanding success, reporting information on the upper atmosphere for many years. Pioneer 13 adopted a collision course and deployed one large and three small entry probes, all of which successfully returned atmospheric data during their descent.

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 15,

© Springer Science+Business Media, LLC 2011

OKB-1

The founding space exploration enterprise in the Soviet Union was Experimental Design Bureau No. l (OKB-1). It had its beginnings in the Scientific Research Institute No.88 (N11-88). A new design section. Department No.3, was set up by the government in May 1946 for the dozens of engineers who had just returned from over a year of investigating the German rocket industry. Sergey Korolev headed the department as Chief Designer. It comprised almost 150 engineers and technicians, and its task, Stalin stated, was to build a Soviet version of the V-2. After succeeding with the R-l. and proceeding to design new rockets of its ow n, the department was restructured into a larger design bureau OKB-1 in the early 1950s and then separated from NII-88 in 1956. OKB-1 built the first Soviet ballistic missile to carry a nuclear warhead, the intermediate range R-5M, and the first submarine-launched ballistic missile, the R-11FM. Korolev’s proposal to build the first intercontinental ballistic missile, the R-7, was approved by the government in 1954. The first successful test of the missile wra$ carried out in August 1957 and on October 4, 1957 it was used to launch Sputnik. The R-7 has been modified, augmented and upgraded in various forms to become the most prolific and successful space launch vehicle in history.

While building for the military. Korolev’s real passion was for space exploration. OKB-1 would eventually lose the military rocket business to rivals, but it achieved great success in space exploration, along with frustrating failure, before Korolev’s death in 1966. After Sputnik. Korolev and OKB-1 pursued more ambitious goals robotic flights to the Moon and planets, and manned flights into Earth orbit. OK B-1 built the first spacecraft to impact the Moon, Luna 2, the first to photograph the far side of the Moon, Luna 3. and the first interplanetary spacecraft intended for Mars and Venus, but the failure rate was terrific. From 1958 through 1965, only four of 21 robotic flights to the Moon were successful (Luna 1, 2 and 3, and Zond 3); none of eleven attempts at Venus and none of the seven attempts at Mars w ere successful. On the other hand, OKB-1 had a singularly excellent record in manned spaceflight, launching the first man into space in 1961, the first woman into space in 1963. the first multi-person spacecraft in 1964, and the first spacewalker in 1965.

There were other design bureaus critical to the space program in the mid-1960s. Valentin Glushko’s OK B-456 w as the premier developer of rocket engines. Glushko

supplied engines for Korolev’s early rockets as well as other military rocket builders such as Chclomey. Chclomcy’s OKB-52 built the Proton rocket which became the staple heavy launcher for Soviet lunar and planetary spacecraft. In 1964 the Soviet Union made the late decision to compete with the IJS and send cosmonauts to the Moon. Korolev, Glushko and Chelomey each presented plans to the government for building the necessary rockets and spacecraft. After considerable wrangling. OKB-1 won on the basis of its head start in the manned program and long-standing work on the design of a Moon rocket. Chelomey did save his Proton rocket from the military scrapheap for the precursor manned circumlunar flights, but OKB-1 was to provide the final upper stage and the spacecraft.

During the battle for control of the manned lunar program, tvhile still conducting both manned and robotic flight programs, succeeding with one and struggling with the other. Korolev realized that OKB-1 had taken on too much. It was essentially responsible for the entire Soviet space effort including communications satellites, reconnaissance satellites, robotic and manned space exploration programs. OKB-1 had to offload something in order to relieve the pressure on his organization, so in March 1965 Korolev reluctantly transferred the robotic program to NPO – Lavochkin. Keldysh played a significant role in this decision. If any comparison to the US could be made at this point, it would be that the USSR had two NASAs one for manned missions (OKB-1) and another for robotic missions (NPO – Lavochkin). This is not a perfect comparison, however, since neither had full control of its own funding or its suppliers; that came from MOM.

After Korolev died in January 1966. OKB-1 was renamed the Central Design Bureau of Experimental Machine Building (TsKBEM) and his deputy Vasily Mishin took over. But unlike Korolev. Mishin was not a charismatic and politically savvy leader and he immediately ran into trouble. He introduced Korolev’s three-person Soyuz spacecraft into service for the first time in April 1967 with tragic results, killing the test pilot Vladimir Komarov when the parachute failed to deploy properly as he returned to Earth. He then presided over the repeated failure of the N-l rocket, which would have launched the Soviet Union’s challenge to Apollo. In 1974 he was replaced by Glushko, who merged the organization with his OKB-456, and then with Chclomcy’s OKB-52, to form the giant NPO-Encrgiya. This organization went on to produce the Energiya heavy lift rocket, the Buran space shuttle, and the Salyut and Mir space stations. Now known as the S. P. Korolev Rocket and Space Corporation Energiya (RRK Energiya) it dominates the Russian manned space flight enterprise, having operated the Mir space station for almost 15 years, supplied the Zvezda habitat module for the International Space Station, and a decade of flights of the Soyuz and Progress spacecraft to service the ISS.

There were essentially three general design series of Russian planetary spacecraft. None of them resembled their American counterparts because, unlike the latter, the Russian spacecraft required pressurized containers for most of their electronics. The Venus and Mars flights in 1960-61 used the first generation spacecraft, which were simple pressurized canisters with attached solar panels and high gain antennas. Their payloads were specific to the target planet, but in general the spacecraft were the same. Of the four launched, only Venera 1 was successfully dispatched and it failed early in its cruise through interplanetary space.

The second generation introduced the first modular spacecraft, with a pressurized carrier that had the propulsion system at one end and a module for the payload at the other. They were individually outfitted for missions to Mars or Venus, with either an entry probe or a flyby module for remote sensing. (This same modular approach was adopted for the second generation Ye-6 lunar spacecraft series.) There were two sub-types of this spacecraft, 2MV and 3MV. Six 2MV spacecraft were launched in 1962, three for Venus and three for Mars, but only one, Mars 1. survived its launch vehicle. The flight of Mars 1 was plagued with problems and it succumbed half way to its target, but the lessons learned were applied in developing the 3MV. Seventeen 3MV spacecraft were launched between 1963 and 1972, five of which. Venera 4 to 8. achieved their planetary objectives. One of the Mars types, Zond 3. did achieve significant results by imaging the far side of the Moon as it departed and subsequently testing the communications system by transmitting the pictures from deep space.

The third generation planetary spacecraft were a major design change, enabled by the powerful Proton launcher with the Block D upper stage. These much larger and

figure 5.5 Representative Soviet planetary spacecraft to scale: first generation Venera 1 (upper left); second generation Venera 4 to 8 (upper right); and third generation Venera 9 to 14 at lower left; and Mars 2. 3, 6 and 7 at lower right (from Space Travel Encyclopedia).

more complex spacecraft were meant to provide planetary orbiters and soft-landers, starting with Mars in 1969 and Venus in 1975. Of twenty-two launched, Venera 9 to 16 and Vega 1 and 2 ran up a string of straight successes at Venus. The other twelve experienced a more difficult challenge at Mars, where only five can be deemed even partial successes, Mars 2 and 3, Mars 5 and 6, and Phobos 2. The Phobos missions of 1988 and the Mars-96 spacecraft were derivatives of this class, but with upgrades sufficiently significant for them perhaps to be regarded as another generation.

In normal flight, Russian spacecraft were flown in uniaxial orientation in which their static solar panels were oriented constantly towards the Sun and the craft spun at 6 revolutions per hour on the axis perpendicular to the plane of the solar panels. The command uplink was at 768.6 MHz through semi-directional conically-shaped spiral antennas which were also used for low-rate data transmission. Because these antennas generate funnel-shaped radiation patterns, several were placed around the spacecraft pointing at the Sun, and at any point in the mission the one with the best funnel angle for Earth was used. For high data rate transmissions, a parabolic high – gain antenna was affixed to the spacecraft. This had to be aimed directly at Earth by disabling the uniaxial control mode, reorienting the spacecraft appropriately, and switching to the three-axis orientation control mode. Circularly polarized decimeter (~920 MHz) and centimeter (~5.8 GHz) band transmitters shared the dish antenna.

In 2MV and 3MV missions, planetary probes and landers were designed for direct transmission to Earth by small spiral antennas with pear-shaped radiation patterns.

The heavier Proton-launched Mars and Venera landers were designed to relay their transmission through flyhy or orbiter spacecraft using large meter band (186 MHz) helical antennas mounted on the rear of the solar panels. The Mars 3 class of entry vehicle carried small wire antennas on the entry stage and another set on the lander. The Venera 9 class of entry vehicle had another large helical antenna installed on top of the lander. Data from the Mars and Venus entry systems was stored for later transmission, but in the case of the Venus landers it was also relayed in real-time as a precaution. The entry system data link operated at 72,000 bits/s for Mars and at 6,144 bits/s for Venus.

Campaign objectives:

After the three Venus launches failed in late August and early September, Korolev’s team scrambled to prepare for three more launches to Mars in late October and early November. Many measures were taken to enhance the reliability of the fourth stage. There was some pressure to abandon the Mars attempts until the problems with this stage were solved, but Korolev blazed ahead.

The 1962 Mars campaign consisted of two flyby missions and one entry probe. The objectives of the entry probe were to obtain in-situ data on the composition and structure of the atmosphere, and data on surface composition. The objectives of the flyby missions were to examine the interplanetary environment between Earth and Mars, to photograph that planet in several colors, to search for a planetary magnetic field and radiation belt, to search for ozone in the atmosphere, and to search for organic compounds on the surface. A comprehensive payload was prepared for each spacecraft, but apart from the camera and a magnetometer most of the payload was deleted when it was decided instead to install instrumentation to monitor the fourth stage to find out why it was suffering so many failures. These missions then became primarily engineering test flights of the 8K78 fourth stage, with Mars as a secondary objective.

Although the fourth stage failed again on two of the launches, the second of three worked and provided the Soviets with their first spacecraft to Mars. Unfortunately, as in the case of Venera 1 it was immediately clear that Mars 1 had attitude control problems. The inability to perform a midcourse maneuver ruled out the desired close flyby of Mars. On the other hand, communications with Mars 1 were maintained for almost 5 months before it fell silent about half w ay to its target.

Spacecraft:

The 2M V Mars spacecraft were virtually identical to the versions described in detail above for the 1962 Venus missions. Although we have no description of the 300 kg entry probe of the 2M-3 No. l spacecraft we know7 it w7as not designed as a lander but as a simple spherical entry system containing a parachute, radio, and instruments intended for measurements during descent. Surviving impact must have been more a hope than a goal. In fact, since the designers had no idea just how thin the Martian atmosphere is. the entry probe would have crashed into the surface before any useful data could have been returned.

The Mars 1 spacecraft is depicted in Figure 8.4 in a stand. Above the stand is the pressurized compartment containing the scientific instruments for the flyby. Next is the ‘orbital’ compartment. The large port in the front is the star sensor, and to the right of that is the Sun sensor. The gas bottles for the attitude control system are on
the waist separating the two compartments. Topping the spacecraft is the propulsion system. The parabolic high gain antenna is fixed pointing in the opposite direction to the solar panels, and there are hemispherical radiators mounted on the ends of the panels.

Launch mass: 893.5 kg (Mars 1)

1.097 kg (probe version)

305 kg ‘

Payload:

Many of the instruments developed for the 2MV Mars spacecraft were removed in order to accommodate systems to monitor the fourth stage of the launcher. There is no information on how many were actually removed, but the magnetometer and the flyby imaging system are known to have been carried by Mars 1.

The original set of instruments is given in this list.

Carrier spacecraft:

1. Magnetometer to measure the magnetic field

2. Scintillation counters to detect radiation belts and cosmic rays

3. Gas discharge Geiger counters

4. Cherenkov detector

5. Ion traps for electrons, ions and low-energy protons.

6. Radio to detect cosmic waves in the 150 to 1,500 meter band

7. Micrometeoroid detector

Descent I landing capsule:

1. Temperature, pressure and density sensors

2. Chemical gas analyzer

3. Gamma-ray detector system to measure radiation from the surface

4. Mercury level movement detector

Flyby instrument module:

1. Facsimile imaging system to photograph the surface

2. Ultraviolet spectrometer in the camera system for ozone detection

3. Infrared spectrometer to search for organic compounds

These instruments were identical to those built for the Venus mission, except that the Mars infrared spectrometer operated in the 3 to 4 micron C-H band to search for organic compounds and vegetation on the surface of Mars.

Mission description:

Two of the three missions were lost to the new and as yet unreliable fourth stage. The 2MV-4 No.3 Mars flyby was launched on October 24, 1962, but failed to leave parking orbit when the fourth stage turbo pump failed after 17 seconds due either to a foreign particle in the assembly or to the pump overheating after a lubricant leak. The fourth stage and spacecraft broke into five large pieces that re-entered over the course of the next few’ days. The US Ballistic Missile Early Warning System radar in

Alaska, which was at a state of high alert in the midst of the Cuban missile crisis, dc tec ted the debris after launch and was initially concerned that it might represent a Soviet nuclear ICBM attack, but rapid analysis of the debris pattern put this fear to rest.

The rocket carrying the second spacecraft was rolled out to the pad the next day. October 25, at the peak of the missile crisis. Shortly thereafter the firing range was ordered to battle readiness, which required the preparation for launch of the two R-7 combat missiles. One of these was stationed at the launch site where the Mars rocket stood. Stored in a corner of the Assembly and Test Building, it was uncovered and the launch team switched from supporting the Mars launch to preparing the missile. Fortunately, when the order to stand down came on October 27 the Mars rocket had not yet been removed from the launch pad. The 2MV-4 No.4 flyby spacecraft was successfully launched on the optimum date of the window, November 1, and became the first spacecraft to be sent towards Mars. The mission was named Mars 1. Just as in the case of Venera 1, a serious problem was discovered immediately after launch. The pressure in one of the two nitrogen gas containers was dropping rapidly because of a leaking valve. Later analysis showed that manufacturing had allowed debris to foul one of the valves. The outgassing caused the spacecraft to tumble out of control. When the tank drained after several days, ground controllers managed to use the gas in the remaining tank to halt the tumbling, restore the spacecraft to the desired Sun pointing attitude and spin it at 6 revolutions per hour so that the batteries would be continuously recharged from the solar panels. But by then most of the dry nitrogen for the cold gas jets of the attitude control system and for pressurizing the engine was expended. The backup gyro system used for attitude control was not designed for continuous use. Stuck in the backup Sun pointing spin mode, the spacecraft was unable to point its high gain antenna at the Earth or to make a midcourse correction. The Earth link was maintained through the UHF system and the medium-gain semi­directional antennas. Contact was established every 2 days for the first 6 weeks, and then every 5 days thereafter. On March 2. 1963, the signal strength began to decline and communications were lost on March 21, probably due to a final breakdown of the attitude control system at the unprecedented range of 106,760,000 km. The silent spacecraft would have passed Mars at a distance of about 193,000 km on June 19, 1963; the intended flyby distance was between 1.000 and 10.000 km.

The third spacecraft to be launched, 2MV-3 No. l, was stranded when the fourth stage failed to reignite properly. Vibrations in the core stage caused by cavitation in its oxidizer lines had dislodged a fuse and igniter in the fourth stage. Its engine was commanded to shut down after 33 seconds. The Americans detected five pieces of debris whose origins were unclear. The spacecraft is believed to have re-entered on January 19, 1963.

Of the six 2M V spacecraft launched between August and November 1962, four were lost to failures of the fourth stage, one was lost to a failure of both the third and fourth stages. The other one was launched successfully and named Mars 1, but failed in transit. No more 2MV spacecraft were built. The design wras improved to produce the 3MV spacecraft for the next series of Mars and Venus missions in 1964 1965.

In the US, the orbital remains of the 1962 Venus and Mars spacecraft, including

Mars 1, were designated as Sputniks 19 to 24 in order of launch. All the spacecraft stranded in parking orbit re-entered within days.

Results:

No information was obtained on Mars. However, Mars 1 did acquire data during its cruise before it fell silent. The radiation zones around Earth were detected, and the distribution and flux of particles were measured. Л third zone at 80,000 km was detected. The solar wind and magnetic fields were measured in interplanetary space to a farther distance than Venera 1. Л solar wind storm was measured on November 30, 1962. ‘flic intensity of cosmic rays had almost doubled since 1959 due to a less active Sun. The micrometeoroid collision rate decreased with distance from Earth and showed intermittent increases as meteoroid showers were traversed. The Taurid meteor shower w as encountered twice at ranges from 6,000 to 40.000 km, and again at distances from 20 to 40 million km. with a strike rate of one every 2 minutes on average.

In late 1968 and early 1969 it became apparent to the Soviet Union that American astronauts might very well reach the Moon before Russian cosmonauts. Anxious to ensure that a Soviet mission was first to return lunar soil to Earth, NPO-Lavochkin hurriedly modified the Ye-8 spacecraft for a sample return mission. The lunar rover variant of this spacecraft was well advanced in design by the end of 1968. and could readily be modified simply by replacing the payload of the lander. Even although it would have scientific merit, the sample return mission had a far greater significance than being just another task for the Ye-8. The robotic sample return mission became the means to upstage Apollo by returning a sample to Earth before the Americans could do so. The fact that these complex spacecraft could be designed and built so readily, and ultimately work so well, is amazing in hindsight. It would seem to be a Russian characteristic to ”just do if to dismiss the hardship, use whatever you have at hand, and fix things up on the fly during and after build.

The modification of the Ye-8 lunar rover spacecraft to the Ye-8-5 for the sample return mission faced daunting problems, not the least of which were mass limitations on the return vehicle, lifting off from the Moon and navigating back to Earth. It was originally believed that the return vehicle would require the same complex avionics as any interplanetary spacecraft, to enable its position to be determined and to make midcourse correction maneuvers. The avionics necessary to meet these requirements far exceeded the available mass. However. D. Ye. Okhotsimskiy, a scientist at the Institute of Applied Mathematics, found a small set of flight trajectories for launches from the surface of the Moon that did not require midcourse corrections. In essence, the large gravitational influence of the Earth at lunar distance could, under certain conditions, assure an Earth return. These trajectories were limited to specific points on the Moon, varying within a general locus with the time of year, and required the lander to set dow n within 10 km of its target and the lunar liftoff for a direct ascent to occur at a precise moment. Accurate knowledge of the lunar gravitation field was also required, but this information had already been determined by the Luna 10, 11. 12 and 14 orbitcr missions.

These passive return trajectories simplified the ascent vehicle enormously. Only a single burn of the ascent vehicle was required. No active navigation was necessary, and no midcourse maneuvers were required. The only problem with a passive return was the very large error ellipse on arrival at Earth, which would make recovering the small capsule impraclically difficult. This problem was solved by using a low-mass meter wave radio beacon on the ascent vehicle so that radio tracking would be able to determine its actual trajectory, supplemented by optical observations from Earth during the latter half of its flight. In addition, the return capsule would have its own radio beacon to assist in recovery operations.

Even with these ingenious solutions, the engineers could not trim the design mass of the Ye-8-5 below 5,880 kg. At that time the most that the Proton-К could send to the Moon was 5,550 kg. However, Babakin managed to cajole the Proton maker into providing sufficient additional mass capability to launch his sample return spacecraft to the Moon. This was accomplished without major changes to the launch vehicle.

Spacecraft;

Lander stage:

The lander stage was essentially the same as designed for the rover mission and its mission profile through to lunar landing was identical. The only differences were the attachment of a surface sampling system and, for the first eight spacecraft, a pair of television cameras for stereo imaging of the sampling site and floodlights for night landings. The rover and ramps were replaced by a toroidal pressurized compartment which held the instruments and avionics for surface operations. The ascent stage was mounted on top, with the entire lander and toroidal compartment acting as its launch pad.

The ascent stage was powered by a silver-zinc 14 amp-hour battery, and the return capsule by a 4.8 amp-hour battery. Lander communications were provided at 922 and 768 MHz, with backups at 115 and 183 MHz. The ascent stage communicated at 101.965 and 183.537 MIIz. The return capsule had beacons at 121.5 and 114.167 MHz for radio tracking.

The sampling system for the Ye-8-5 consisted of an upright 90 cm long boom arm capable of two degrees of freedom, with a drill at its end for surface sampling. Three movements were required to place the drill on the surface through a 100 degree arc of swing, and then another three to transfer the sample to the ascent stage. From the stowed position it first swung itself vertical, then rotated in azimuth to line up on the selected sample site before swinging down onto the surface. A movement in azimuth with the head on the ground might be used to clear a small area to improve drilling. This sequence was reversed to transfer the sample to the return capsule of the ascent stage. Mounted at the end of the boom was a cylindrical container 90 mm diameter and 290 mm long for a hollow rotary/percussion drill. The drill bit had a diameter of 26 mm and was 417 mm long. Its cutter was a crown with sharp teeth. The drill was equipped with different coring mechanisms for hard coring and for loose coring. At a speed of

500 rpm, it required 30 minutes to fill the entire core length of 38 cm. The drill was both insulated and hermetically sealed, and to enable the mechanism to be lubricated using oil vapor it was not opened until just before use. Some parts used a lubricant designed to reduce friction in a vacuum. A standby motor was provided as a contingency to overcome obstacles encountered during drilling. The whole device weighed 13.6 kg.

An improved drill system tvas provided for the Ye-8-5M version, which had a rail mounted deployment mechanism. This drill was capable of penetrating to a depth of

2.5 meters and preserving the stratigraphy, but it could not be articulated to select a sampling site. It used an elevator mechanism rather than the articulated boom arm to transfer the sample to the return capsule.

Ascent stage:

The ascent stage was a smaller, vertically mounted open structure composed of a pressurized cylindrical avionics compartment above three spherical propellant tanks and the rocket engine. This was the same engine as used on the lander, but was not throttled. Four vernier engines were attached outboard of the propellant tanks. There were perpendicular antennas mounted radially at 90 degree intervals near the top of the avionics compartment. The spherical return capsule was held in place on top by deployable straps. Including the return capsule, the ascent stage was 2 meters tall. It weighed 245 kg dry and 520 kg with propellant. The KRD-61 Isayev engine burned nitric acid and UDMH and produced a thrust of 18.8 kN for 53 seconds to impart a velocity of 2.6 to 2.7 km/s, which was enough to escape from the Moon on a direct ascent trajectory.

Return capsule:

The return capsule was a 50 cm sphere covered with ablative material for entry at a speed of about 11 km/s and a peak deceleration load of 315 G. It had three internal sections. The upper section contained the parachutes (a 1.5 square meter drogue and a 10 square meter main) and beacon antennas, the middle section contained the lunar sample, and the base had the heavy equipment including batteries and transmitters.

on entry.

Luna 15 launch mass: 5,667 kg

Luna 16 launch mass: 5,727 kg

Luna 18 launch mass: 5,750 kg

Luna 20 launch mass: 5,750 kg

Lima 23 launch mass: 5,795 kg

Luna 24 launch mass: 5,795 kg

4,800 kg (Luna 24)

1,880 kg

520 kg (515 kg for Luna 23 and 24) 35 kg (34 kg for Luna 23 and 24)

Payload:

1. Stereo panoramic imaging system with lamps (deleted on Luna 23 and 24)

2. Remote arm for sample collection (improved drill on Luna 23 and 24)

3. Radiation detectors

4. Temperature sensor inside capsule

The stereo imaging system had two 300 x 6,000 panoramic scan cameras of the type used on the earlier Yc-6 landers and Lunokhod rovers. Mounted on the lander just below the level of the ascent stage on the same side as the sampling system, they were spaced 50 cm apart, angled at 50 degrees to the vertical, and gave a field of view of 30 degrees. The orientation of the lander was determined by measuring the position of Earth in a panoramic image. Stereo images were taken of the surface between the
two landing legs to select the position to be sampled. They also imaged sampling and drilling operations, For the Luna 23 and 24 sample ret urn missions the earner as and lamps were deleted.

Mission description:

Only six of the eleven spacecraft in this series were launched successfully. Of these, three succeeded in returning lunar samples to Earth.

The first attempt

The first launch (Ye-8-5 No.402) was attempted on June 14, 1969. one month prior to the Apollo 11 launch date, but the Block D failed to ignite for its first burn and the payload re-entered over the Pacific Ocean.

Luna 15

The second spacecraft in this series was successfully launched on July 13, 1969, just 3 days before Apollo 11, and the Soviets announced that Luna 15 was to land on the Moon on July 19, one day ahead of the Americans, with the objective of returning something to Earth. At 10:00 UT on July 17 it entered a 240 x 870 km lunar orbit inclined at 126 degrees. ‘This orbit was much higher than intended, so the next day it was trimmed to 94 x 220 km. Another trim a day later yielded an orbit 85 x 221 km. Ideally the orbit should have been near-circular at about 100 km, but the Soviets had underestimated the effect of the lunar in a scons and they were also suffering attitude control problems. Meanwhile. Apollo 11 had arrived and entered an equatorial orbit The drama was palpable. In Russia its nature was clear, but in America the ultimate purpose of Luna 15 was mysterious and opinions ranged from the suspicious to the sublime to the ridiculous. Apollo 8 astronaut Frank Borman, just back from a visit to the Soviet Union, appealed for information and the Academy of Sciences supplied orbit data, operational frequencies, and assurances that Luna 15 would not endanger the Apollo 11 mission.

On July 20. after several more orbit changes, Luna 15 began its descent sequence during its 39th orbit by lowering its perilune to 16 km above the landing site in the Sea of Crises. The intention was to land just 2 hours before Apollo 11 landed further west in the Sea of Tranquility. But when controllers saw the radar data from the first perilune pass they became concerned. The one and only target appeared uneven and potentially dangerous. It must have been with the utmost reluctance and dismay that the decision was taken to postpone the landing to test the radar and perform further observations. As a result of this delay, not only was there now’ no chance of landing ahead of the Americans, the nature of the return trajectory would make it impossible to get a sample back first. All of this w as unknown to an anxious world, wondering wrhat stunt Luna 15 was going to pull in order to upstage Apollo 11. Eighteen hours later, on its 52nd orbit, Luna 15 w*as commanded to land at 15:46:43 UT on July 21. after Armstrong and Aidrin had already walked on the Moon. The descent maneuver failed and. for reasons still to be explained, the transmission ceased 4 minutes after the de-orbit burn started. It erashed at 17 N 60 H, about 800 km east of Tranquility Base. Jodrell Bank flashed notification to the Americans that Luna 15 had impacted at a velocity of 480 m/s just as Apollo 11 ‘s lunar module was preparing to leave the Moon. The Soviets reported that Luna 15 had ’"reached the lunar surface in a preset area" but remained silent on its true mission and there was no propaganda victory.

Campaign objectives:

The 1978 Venus campaign objectives were to repeat the resounding successes of the Venera 9 and 10 landers with new instruments to analyze both the atmosphere and surface. The 1976-77 opportunity was skipped in order to build the new7 apparatus, and a launch in 1978 dictated a much higher arrival velocity at Venus than in 1975. The larger propellant load required for the longer orbit insertion burn was unable to be accommodated together with the mass of the new instruments, so the carrier was downgraded to a flyby role. A positive outcome wras that a flyby spacecraft would remain in view7 of the lander for longer in order to relay data from the surface. Both of the previous landers had still been operating when their orbiters flew7 below their horizons. The new7 lander investigations featured a high resolution color camera and an experiment to drill into the surface. The descent investigations included new7 experiments to study the chemical composition of the atmosphere, the nature of the clouds, and any electrical activity in the atmosphere. The flyby instrumentation was reduced in order to maximize the mass available for the descent and surface science.

Spacecraft:

Although assigned only to a flyby role the Venera 11 and 12 spacecraft were almost identical to their or biter predecessors, but the lander relay was increased to 3 к bits/s per channel. After releasing the entry system 2 days prior to arriving at the planet, Venera 11 (and all later flyby spacecraft) made a deflection maneuver to establish a flyby w7hich would enable it to relay to Earth for longer than w^as possible using an
orbitcr. The spacecraft were identical and the landers had the same configuration as their recent predecessors but the floodlights were deleted and the camera lens cap was redesigned. The complexity and mass of the parachute system was reduced to accommodate more instruments. Only a single supersonic braking parachute was used instead of a sequence of two, and only one main parachute was used instead of a system of three. Some of the instruments were modified and new ones were added, in some cases being installed on the shock absorbing impact ring. All landers from now through to Vega 2 carried a technology experiment consisting of a set of small solar cells arranged around the lander ring.

4,450 kg (Venera 11) 4,461 kg (Venera 12) 2,127 kg 1,600 kg

731 kg

Payload:

Flyby spacecraft:

1. Extreme-ultraviolet (30 to 166 nm) spectrometer (France)

2. Magnetometer

3. Plasma spectrometer

4. Solar wind detectors

5. High energy particle detectors

6. KONUS gamma-ray burst detector

7. SNEG gamma-ray burst detectors (France-USSR)

KONUS was an interplanetary cruise experiment to try and identify the source of mysterious astronomical gamma-ray bursts by having the two spacecraft coordinate with the Prognoz satellite in Earth orbit to triangulate on individual bursts. SNEG was an instrument complementary to KONUS, built in cooperation with the French. And a new French-built extreme-ultraviolet spectrometer covered the spectral lines of atomic hydrogen, helium, oxygen and other elements that it was thought might be present in the exosphere of Venus. The solar wind detector was a hemispherical proton telescope, and the high energy particle experiments used four semiconductor counters, two gas-discharge counters and four scintilla­tion counters.

Lander:

Entry and descent:

1. Scanning spectrophotometer (0.43 to 1.17 microns)

2. Mass spectrometer Гог atmospheric composition

3. Gas chromatograph for atmospheric composition

4. Nephelometer for aerosols of about 1 micron in size

5. X-ray fluorescence spectrometer for elemental composition of aerosols

6. Accelerometers for atmospheric structure from 105 to 70 km

7. Temperature and pressure sensors for 50 km to the surface

8. GROZA radio sensor at 8 to 95 kHz for electrical and acoustic activity

9. Doppler experiment for wind and turbulence

The experiments for determining the light scattering properties of the atmosphere were modified for a higher spectral resolution. The angular scattering nephelometers were deleted because they had satisfactorily measured the particle size and refractive index when carried on the Venera 9 and 10 landers. Instead, only the back scattering nephelometers w ere retained to examine the spatial uniformity of the cloud layers in different regions of the planet. The scanning spectrophotometer was improved for a higher spectral resolution. Every 10 seconds it measured radiation coming from the zenith using a ramp interference filter at a resolution of about 20 nm continuously over the range 430 to IT70 nm, and the angular distribution of radiation (a full 360 degrees) in the vertical plane in the bands 0.4 to 0.6, 0.6 to 0.8, 0.8 to 1.3. and 1.1 to

1.6 microns using a rotating prism mounted on the aerobrake. The temperatures and pressures were measured by a suite of four thermometers and three barometers, flic monopole radio-frequency mass spectrometer of Venera 9 and 10 w as replaced by a Bennett radio-frequency design and the inlet system modified to prevent it becoming clogged by cloud particles which might then contaminate the atmospheric readings. The microscopic leak admitting the atmosphere to the instrument w7as replaced by a piezoelectric valve that would open a relatively large hole for a very short time in order to admit a pulse of atmosphere into a long sample tube which w? ould trap cloud particles. In addition the instrument was not to be operated until the lander wras at about 25 km. well below the aerosols. The apparatus w^as pumped down between atmospheric readings to purge the sample. Two other new atmospheric composition experiments were included. The gas chromatograph used neon to carry atmospheric samples through columns of porous materials and a Penning ionization detector. It had one column 2 meters long that was optimized for water, carbon dioxide and the compounds hydrogen sulfide, carbonyl sulfide, and sulfur dioxide; a second column 2.5 meters long for the volatile gases helium, molecular hydrogen, argon, molecular oxygen, molecular nitrogen, krypton, methane and carbon monoxide; and a third column just 1 meter long specifically for argon. An x-ray fluorescence spectrometer used gamma rays to excite the emission of x-rays from cloud particles collected on a cellulose acetate filter by drawing atmospheric gas through the instrument, thereby measuring the elemental composition of the aerosols.

The GROZA experiment comprised an acoustic detector and an electromagnetic wave detector using loop antennas with four narrow-band receivers at 10, 18, 36 and 80 kHz and a wide-band receiver over the range 8 to 95 kHz. This was to start at an altitude of 60 km and operate down to and on the surfaec. The electromagnetic wave detector was to register radio bursts from lightning and the acoustic signals could be interpreted in terms of thunder, wind speed past the lander during the descent and, while on the ground, perhaps even seismic quakes.

Surface:

1. Panoramic two-camera color imaging system

2. Soil drill with x-ray fluorescence spectrometer analysis system

3. Rotating conical soil penetrometer (PrOP-V)

The panoramic camera system had been improved by adding clear, red, green and blue filters for three-color imaging, and by increasing the image quality from 128 x 512 pixels at 6 bit encoding to 252 x 1,024 pixels at 9 bit encoding and 1 bit parity. It was capable of resolving detail as fine as 4 or 5 mm at a range of 1.5 meters. The transmission bandwidth had been increased by a factor of twelve, one reason for this being the Kvant-D upgrade to the Soviet communications facilities; in particular the introduction of 70 meter antennas at Yevpatoria and Ussuriisk. The increase in the transmission rate from the surface of Venus from 256 bits, s to 3,000 bits/s enabled a color panorama to be sent in 14 minutes, as against 30 minutes for a lower resolution black-and-white panorama previously.

The gamma-rav soil analysis instrument inside the Venera 8, 9 and 10 landers was replaced by a superior instrument. A drill mounted on the shock-absorbing impact ring w as to core a sample of the surface and then pass it through a series of pressure – reduction stages to the x-ray fluorescence spectrometer carried inside the lander. The penetrometer was on a deployable arm and reported its results on a dial that was to be read by the cameras.

1ission description:

Venera 11 lander:

Venera 11 was launched on September 9, 1978, and made midcourse corrections on September 16 and December 17. After the entry capsule was released on December 23 the spacecraft made the deflection burn in order to perform a flyby of the planet at the desired relay communications altitude, and on December 25 the entry system hit the atmosphere at 11.2 km/s. After a 1 hour descent the lander touched down at a speed of 7 to 8 m/s on the day-side at 14 S 299 E. It w as 03:24 UT, 11:10 Venus solar time, and the solar zenith angle was 17 degrees. The lander transmitted from the surface for 95 minutes before the relay spacecraft flew over the horizon after 110 minutes, so none of the transmission was lost.

figure 15.3 Venera 11 and Venera 12 encounter design, showing entry capsule targeting followed by flyby vehicle deflection and lander relay communications.

Venera 12 lander:

Venera 12 was launched on September 14, 1978, pursued a faster trajectory with midcourse corrections on September 21 and December 14, and arrived ahead of its partner. It released its entry system on December 19. This entered the atmosphere on December 21 at a velocity of 11.2 km/s. The parachute was jettisoned at 49 km and after a 1 hour descent the lander touched down at about 8 m/s on the day-side at 7°S 294°E. It was 03:30 UT, 11:16 Venus solar lime, and the solar zenith angle was 20 degrees. Unlike Venera 11, it kicked up a cloud of dust that took about 25 seconds to settle. Both landers encountered an unexplained anomaly at an altitude of 25 km, where instrument readings went off-scale and there was an electrical discharge from the vehicle. This lander transmitted from the surface for 110 minutes until the flyby spacecraft passed below the horizon. It is therefore not known when it finally ceased to function.

Venera 11 and 12 flyby spacecraft:

After the deflection maneuver, each spacecraft Hew by Venus at a range of about

35,0 km and relayed the data from its lander to Earth throughout the descent and then during the period of surface activity. The last reports from the flyby spacecraft were in January 1980 for Venera 11 and March 1980 for Venera 12.

Results:

Venera 11 and 12 decent measurements:

The landers inferred atmospheric density from accelerometer data over the altitude range 100 to 65 km and then directly measured atmospheric temperature and pressure from 61 km down to the surface. Opacity was measured from 64 km to the surface, the chemical composition of aerosols from 64 km to 49 km, aerosol scattering from 51 km down to the surface, and thunderstorm activity from 60 km down to the surface. The gas chromatograph analyzed nine atmospheric samples from 42 km to the surface. The new mass spectrometer measured atmospheric composition from 23 km to 1 km. Wind velocities were measured from about 23 km down to the surface, and altitude profiles of horizontal wind speed and direction were obtained from Doppler data.

The spectrophotometer produced the first realistic water vapor profile, identifying water vapor as the second most important greenhouse gas in the atmosphere (after carbon dioxide). The contemporary analysis indicated a profile that decreased from 200 ppm at the cloud base to 20 ppm at the surface, but a re-analysis many years later obtained a better fit for the Venera 11,12 and 14 spectrophotometer data using a constant mixing ratio for water vapor of about 30 ppm from 50 km to the surface. The mass spectrometers on these missions reported values as high as 0.5% at 44 km and 0.1% at 24 km; these are much larger mixing ratios for water vapor than were obtained from the spectrophotometer and other remote spectral measurements from Hanh, and are considered suspect.

The mass spectrometer results from Venera 11 and 12 obtained by the analysis of 176 complete spectra of 22 samples were reported as:

carbon dioxide

molecular nitrogen

argon

neon

krypton with isotopic ratios as follows:

0. 0112 ± 0.0002 1.19 ± 0.07 0.197 ± 0.002

The gas chromatograph made eight measurements betw een 42 km and the surface with the following results:

2.5 + 0.3 %

25 to 100 ppm

40+10 ppm

130 + 35 ppm

28 + 7 ppm (low altitudes)

less than 20 ppm

krypton

hydrogen sulfide carbonyl sulfide

The x-ray fluorescence spectrometer on Venera 12 measured cloud particles from 64 km to 49 km and was then overcome by the high temporal tires. It missed sulfur (< 0.1 mg/m3). but found chlorine (0.43 + 0.06 mg/nri) in cloud particulates. The chlorine was suspected to be a non-volatile compound such as aluminum chloride at the time, but it was not specifically identified. The large amount of chlorine relative to sulfur was incompatible with the theory that the clouds were composed of sulfuric acid droplets, but these anomalous data were corrected by the Venera 14 mission.

Both Venera 11 and 12 detected a large number of electromagnetic pulses in the descent from 32 to 2 km similar to those produced by distant lightning Hashes on Harth. The activity was more intense on Venera 11 than Venera 12, and diminished in intensity towards the surface. No such pulses were detected by Venera 11 after it touched down, but one large burst was noted by Venera 12 while on the surface. The microphones were saturated by aerodynamic noise during the descent and detected no thunder on the surface, but they did pick up sounds issued by the instruments and surface activities.

As with Venera 9 and 10. the light scattering data indicated clouds with a base at an altitude of 47 km, and a much lower loading of aerosols below that. Venera 11 and 12 found the atmosphere to be generally free of aerosols below ^30 km. The nephelometer on Venera 11 measured cloud particles throughout the descent and its results confirmed the uniformity of the cloud layers as reported by Venera 9 and 10. The base cloud layer was located between 51 and 48 km, with a mist below that. The nephelomeler on Venera 12 did nol function correctly. It was confirmed that only about 3 to 6% of the sunlight reaches the surface. Intense Rayleigh scattering in the dense atmosphere gives poor visibility. Above several kilometers altitude the surface must be invisible. At ground level the horizon will be visible, but the detail of the landscape must fade quickly into an orange haze. The Sun is not visible as a disk, merely a uniformly lit hazy sky.

Venera 11 and 12 surface measurements:

The temperature at the Venera 11 landing site was 458 + 5 C and the pressure was 91 + 2 bar. There was no surface imaging because the lens covers would not open. These had been redesigned after the problems w ith one camera on each of Venera 9 and 10, but with disastrous results. The transmit ted pictures were uniformly black. The soil drill collected a sample, but it was not properly delivered to the instrument container and no soil analysis was accomplished.

The temperature at the Venera 12 site was 468 + 5 C and the pressure was 92 + 2 bar. The fact that this suffered exactly the same camera and soil analysis experiment failures as its partner implied a systematic design flaw. Vibrations while descending broke the sample transfer system on the drill and no soil analysis was possible. The soil penetrometers also failed on both landers.

The surface experiments were an almost total failure on both landers. It is possible
that they suffered rough landings which damaged the instruments that were mounted on the impact ring. The lack of results was very disappointing, hut in typical Soviet fashion this spurred the engineers on to succeed at the next flight opportunity.

Venera 11 and 12 flyby spacecraft:

The ultraviolet spectrometer detected Lyman-alpha emissions from hydrogen atoms and 584 angstrom (He-I) emissions from helium atoms. These provided exospheric temperatures and number densities. Time profiles for 143 gamma-ray bursts were obtained by Venera 11 and 12 and the results triangulated with an identical detector on Prognoz 7 in harth orbit. On Kebruary 13 and March 17, 1980, Venera 12 used its extreme-ultraviolet spectrometer to observe Comet Bradfield.

The Scientific Production Organization NPO-Lavochkin was originally founded in 1937 as the Lavochkin Aircraft Design Bureau, OKB-301, named for its Chief Designer. Lavochkin produced a number distinguished fighter aircraft during WW-11

and then surface-to-air missile designs after the war, producing the first operational system for the defense of Moscow. In 1953 the SAM business was transferred to a new design bureau and OKB-301 pursued ramjet intercontinental cruise missiles as a hedge against problems with ICBM development. But the successful introduction of ICBMs in the late 1950s left OKB-301 without work. Semyon Lavochkin died in 1960 and the organization transferred in 1962 to Chelonrey’s OKB-52. The factory was closed, but reopened in 1965 as NPO-Lavochkin under the steady and capable leadership of Georgi Nikolayevich Babakin specifically to take responsibility for the robotic lunar and planetary spacecraft programs transferred from OKB-1.

The new NPO-Lavochkin realized immediate success using its inheritance from OKB-1 augmented by a history of great skill and experience in aviation technology. Luna 9 soft landed on the Moon’s surface in January 1966, and before the year was over there were three successful lunar orbiters and a second soft lander. The first successful Venus entry probe, Venera 4, followed in 1967. NPO-Lavochkin went on to continue this highly successful series of spacecraft at Venus, a successful series of lunar orbiters, rovers and sample return missions, and the singularly complex and successful Vega missions which delivered landers and balloons to Venus enroute to a flyby of Halley’s comet. LTnfortunately, NPO-Lavochkin had no success at Mars: their campaigns in 1969, 1971, and 1973 were riddled with failures, and worse was to come in 1988 and 1996. Their astronomy missions have met with better success, in particular the Granat and Astron space observatories. Today NPO-Lavochkin is the single engineering center for the production of robotic scientific spacecraft.