Category Soviet Robots in the Solar System

Campaign objectives:

With the survival of the Venera 8 capsule on the surface of Venus in 1972, the 3MV spacecraft had reached the limit of its capability and the Soviets were ready for the next step. They now had enough data on the atmosphere of Venus and conditions at the surface to design a very capable lander that included sophisticated imaging and surface science instruments. The challenge was to enable this apparatus to operate in such harsh conditions. Also, the new heavy Proton-launched Mars spacecraft had proved itself in 1971 with the Mars 2 and 3 missions. Both orbiters were successful, and the Mars 3 lander was successfully delivered to the surface. This orbiter served as the basis for designing the Venus spacecraft. However, the entry vehicle had to be completely redesigned. For the first time since their initial launch to Venus in 1961, the Soviets skipped a Venus opportunity in October 1973 while they w’orked on their new spacecraft.

The main difference between the heavy Proton-launched spacecraft for Mars and for Venus was the entry system. A vehicle to enter the rarefied atmosphere of Mars needs a broad conical aerobrake lor rapid deceleration in the upper atmosphere and large robust parachutes to slow to a safe speed before reaching the surface. The thick atmosphere of Venus, on the other hand, is much more forgiving and permits the use
of a simpler entry system. The new system for Venus was a hollow spherical vessel that contained the heavy lander and its parachute system. The previous probes had revealed the atmosphere to be so thick that to reach the surface in a reasonable time the new system w as designed to jettison its parachute high in the atmosphere and let the lander fall using an aerobrake. and since the free-fall velocity at the surface was slow enough for the impact to be survivable there w as no requirement for a terminal retro-rocket.

There were minor modifications to the spacecraft, including changes in the size and position of the solar panels, the thermal systems, and increased reliability. One key change was to replace the dircet-to-Harth communications system of the lander with a system to relay via the or biter, which greatly improved the data rate from the lander. The mission plan for Venus differed from that for Mars. The orbiter lander w’as first targeted at the atmospheric entry point, rather than at the orbital insertion point. Several days from Venus the passive entry system w’as released. Immediately afterwards, the spacecraft made a deflection maneuver for the orbital insertion point. The timing was such that by the time the lander started to transmit, the spacecraft would have just completed its orbit insertion burn and have a line of sight to relay the transmission.

The principal scientific goal of the lander w^as to obtain the first panoramic image on the surface of Venus. This determined the minimum time that the lander must be able to operate on the surface and also the data rate of the relay through the orbi ter. These new* capabilities, and the large mass available on the lander, meant a number of instruments that had never been flown before could be carried Гог descent science. These included instruments to measure the vertical structure of aerosols w’ithin and under clouds, the vertical and spectral distribution of solar flux penetrating through

the clouds for several look angles, chemical and isotopic analysis of the atmosphere, and direct measurements of the winds at the surface. To undertake the first science from orbit, the spacecraft had experiments to report upon the plasma environment around the planet and its atmospheric structure, upper cloud layers, and outgoing thermal radiation.

Early on, consideration had been given to using a flyby spacecraft to support the lander mission but NPO-Lavochkin and IK I argued hard for the orbiter in order to obtain the additional and original science that only an orbiter could provide. And, of course, being first to place a spacecraft in orbit around the planet would be another significant first achievement in space exploration.

Spacecraft:

As the first of another generation of spacecraft, Venera 9 and 10 were five times heavier than their predecessors and were launched by the more powerful four-stage Proton-K. This launcher was introduced in 1969 for the Ye-8 lunar missions and then used for the M-69, M-71 and M-73 campaigns. These new spacecraft consisted of an orbiter with the entry system strapped on top, inside of w hich w as the lander. The new entry system and lander were extensively tested in wind tunnels and by airdrops.

The two spacecraft for this opportunity were essentially identical, but Venera 10 was slightly heavier and required a larger fuel load for its longer orbit insertion burn.

kg (Venera 9) 5,033 kg (Venera 10) kg (Venera 9) 1,159 kg (Venera 10) kg (Venera 9) 2,314 kg (Venera 10) kg

The orbiter was based on the M-71 spacecraft. The IJDMH and nitrogen tetroxide tanks formed a cylindrical body. At 110 cm in diameter, this was narrower than the 180 cm Mars version and it was also 1 meter shorter. Below’ was the KTDU-425A restartable rocket engine which could be throttled between 9,856 and 18,890 N for a total of 560 seconds. The avionics and science instruments were in a pressurized toroidal module with a diameter of 2.35 meters that was attached at the base of the cylinder, with the gimbaled engine nozzle protruding through the donut. Navigation optics attached to the exterior of the instrument module included a number of solar sensors mounted in a linear cluster, bordered either side by duplicate down-looking telescopic Canopus sensors. The Earth sensor was installed in such a way as to point in the same direction as the parabolic high gain antenna. With the entry system on top, the spacecraft was 2.8 meters tall.

Tw o 1.25 x 2.1 meter solar arrays extended from opposite sides of the cylinder with an overall span of 6.7 meters. These supported cold gas attitude control jets, the

magnetometer booms, and a relay antenna for the entry system and lander. Also on the side of the cylinder were thermal control gas radiators and tanks which contained nitrogen at 350 bar for the attitude control system. During the interplanetary cruise, louvers on the entry shell provided passive thermal control of the entry system. Communications from the entry system and lander were received by the orbiter and relayed to Earth. There was a 1.6 meter diameter parabolic high gain antenna on the side of the cylinder for communicating with Earth on decimeter and centimeter bands. Six omnidirectional helical antennas were attached near the parabolic antenna, four for Earth and two for the lander. The command uplink was by the helical antennas at 769 MHz. There was a 16 megabyte magnetic tape system for data storage. The downlink to Earth was phase modulated PCM at 3 kbits/s via the

high gain antenna, or in an emergency at a much slower rate using the helical antennas. Data from the lander was retransmitted through the parabolic antenna to Earth in real-time and also stored on tape for later backup transmission. The spacecraft computers were similar to those carried by the M-71 missions.

As it did for Mars, this design served as the basis for all Venus spacecraft starting in 1975, and the Proton-K became the singular planetary launch vehicle in the Soviet inventory.

Entry system:

The sophisticated new lander was contained within a 2.4 meter diameter spherical entry system, and was deployed after the rate of descent through the atmosphere had been reduced to subsonic. The entry vessel was a simple sphere covered by ablative material that consisted of asbestos composite over honeycomb, and it was stabilized during entry by placing the center of mass towards leading side. The entry angle was shallower than for the 3MV capsules, reducing the peak load from around 450 G to a more modest 150 to 180 G. After entry, the sphere would split into hemispheres, releasing the lander and its parachute system.

Lander:

The lander was 2 meters in height, which was much larger than the 3MV capsules, and capable of carrying more scientific instruments. Previous Venera probes were limited to a transmission rate of 1 bit/s by their direct communications link to Earth. The lander was battery powered and transmitted to the orbiter through two VHF channels at 256 bits/s for relay to Earth using the high gain antenna.

The lander was basically a hermetically sealed titanium spherical pressure vessel 80 cm in diameter, containing most of the instrumentation and electronics. This was affixed to a ring-shaped landing cushion by a set of shock absorbers. Above was a disk-shaped aero brake 2.1 meters diameter, to slow the rate of descent during free – fall. This disk also acted as a reflector for the cylindrically wound omnidirectional antenna above. Inside this 80 cm diameter 40 cm tall cylinder were the parachutes and some of the descent instruments. The sphere consisted of several sections bolted together with gold wire seals. It was surrounded with a 12 cm layer of honeycomb insulating material and a thin surface of titanium. The inside of the sphere was also lined with a polyurethane foam insulating material. The thermal design was similar to the earlier landers. The lander was pre-chilled to -10°C by cold air from the main spacecraft via two pipes through the entry vessel. A lithium nitrate trihydrate phase – change material which melted at 33°C absorbed heal that penetrated the insulation, and a gas circulation system distributed it evenly. These measures, and the life of the batteries, allowed operation for about an hour after landing.

Entry, descent and, landing:

The midcourse maneuvers were to target the spacecraft for the entry point. The entry system w’ould be released 2 days from Venus to make a ballistic approach and enter the atmosphere at 10.7 km/s at an angle in the range 18 to 23 degrees. Six seconds later it would experience the peak deceleration of 170 G. After 20 seconds, having slowed to 250 m/s and under a 2 G load at an altitude of about 65 km. the small pilot

parachute would deploy with a ripcord to draw out the 2.8 meter drogue parachute. The spherical shell would then split into hemispheres and the drogue would pull the upper hemisphere and attached lander away from the lower hemisphere, at the same time deploying a second braking parachute. After 11 seconds, by now at an altitude of 60 to 62 km and descending at 50 m/s, the upper hemisphere would release the lander, in the process extracting the three 4.3 meter diameter main parachutes from the cylindrical section on top of the lander.

Once on its main parachutes, the lander would activate its instruments. It would spend about 20 minutes descending through the cloud layer at a rate of about 50 m/s. On reaching an altitude of 50 km, the lander would shed its parachutes and spend the next 55 minutes free falling, with the drag of the disk-shaped aerobrake slowing its rate of descent as it penetrated the thicker air close to the surface. This strategy was selected to minimize heat inflow during descent, and hence extend the lifetime of the lander on the surface. The terminal velocity on reaching the surface would be 7 m/’s, and the impact would be cushioned by the compressible, metal annular landing ring.

Payloads:

Orbiter:

1. Panoramic ultraviolet cloud cameras, 345 to 380 nm and 355 to 445 nm

2. Cloud infrared spectrometer (1.6 to 2.8 microns)

3. Cloud thermal infrared radiometer (8 to 28 microns)

4. Cloud ultraviolet imaging spectrometer (352 and 345 nm) [USSR-France]

5. Cloud photopolarimeters (335 to 800 nm)

6. Lyman-alpha H/D photometer

7. Airglow spectrometer (300 to 800 nm).

8. Triaxial magnetometer

9. Plasma electrostatic analyzer

10. Charged particle traps

11. Cherenkov energetic particle detectors

12. Centimeter and decimeter radio occultation experiment

13. Bistatic 32 cm radar mapping experiment

The two cloud cameras were the same as the linear scanning photometer cameras of the Mars 4 and 5 orbiters. providing cross-track scanning of 30 degrees and using the motion of the spacecraft to scan along the orbital track. The Venus cameras used violet and ultraviolet filters, scanned 500 cycles/line at 2 lines/second. Images were usually transmitted at 256 pixels/line with 6 bits/pixel. Panoramas were typically

6,0 pixels in length. For a periapsis at about 5,000 km the resolution at the cloud tops was on the order of 6 to 30 km.

Between them, the spectrometers and photometers could make measurements of the clouds throughout the ultraviolet, visible and infrared parts of the spectrum. The photopolarimeters were an improved form of those of the M-71 and M-73 missions, with design help from the French. The cloud infrared spectrometer used a circular – ramp interference filter and made high resolution spatial scans across the planet. The thermal infrared instrument used two horn radiometers for bands at 8 to 13 microns and 18 to 28 microns, both of which were relatively transparent in an atmosphere of carbon dioxide. The French-built cloud ultraviolet imaging spectrometer measured spatial profiles across the planet at two wavelengths with a resolution of 16 seconds of arc. The particle detectors included low energy electron, proton and alpha particle sensors, three semiconductor counters, two gas discharge counters, and a Cherenkov detector.

Lander:

Entry and descent:

1. Broadband photometer with three visible and two infrared channels for radiation flux

2. Narrow-band infrared photometer with three channels near 0.8 microns for radiation flux ratios in water, carbon dioxide and background bands

3. Back scatter and multi-angle nephelomctcrs at 0.92 microns for light scattering between the altitudes of 63 and 18 km

4. Pressure and temperature measurements from 62 km to the surface

5. Accelerometers for atmospheric structure between 110 and 76 km

6. Mass spectrometer for atmospheric composition from 63 to 34 km altitude

7. Doppler experiment for wind and turbulence

In comparison to Venera 8, the new photometers were greatly improved and more complex. Both up welling and down welling integrated radiation was measured in the range 0.440 to 1.160 microns using green, yellow, red. IR1. and IR2 glass filters for five wavelength hands with widths of 0.1 to 0.3 microns. This was complemented by an near-infrared photometer operating in three channels, one centered on the carbon dioxide band at 0.78 microns, another on the water band at 0.82 microns, and a third background channel at 0.80 microns, with each band being only 0.005 microns wide. Both back scattering and angular scattering nephelometers were carried. These were new, measuring how7 the atmosphere scattered light from a pulsed light source. This information could be used to infer the size distribution, refractive index, and density of cloud droplets. The sensors for the photometers and nephelometer w ere mounted in the external environment, had their own thermal protection, and were linked by fiber optics to the instruments inside. The mass spectrometer was a radio-frequency monopole unit with a pressure regulator designed for input pressures of 0.1 to 10 bar. The Doppler experiment was facilitated by an ultra-stable master oscillator for the transmitter.

Surface:

1. Panoramic imaging system, two cameras with floodlights

2. Surface wind rotary anemometer

3. Gamma-ray spectrometer (uranium, potassium, and thorium) for surface rocks

4. Gamma-ray densitometer

The scanning photometer imaging camera w as similar to that carried by the M-71 landers and had a mass of 5.8 kg. There were tw o, in sealed insulated containers on either side of the lander just beneath the disk of the aero brake to give a vantage point 90 cm off the ground. The rotational axis of the mirror system was tilted from the landers vertical by 50 degrees in order that the center of the image was the surface directly in front of the camera at a distance of 1.5 to 2 meters, with the field of view extending 90 degrees to either side in order to include a small section of the horizon. The cameras peered through 1 cm thick cylindrical quartz windows using a lens to compensate for refraction and provide a total angular field of 40 x 180 degrees. Each 128 x 512 panorama consisted of a 115 x 512 image. w7ith the first 13 bits of each line containing a calibration pattern. Each measurement consisted of a 6 bit picture element and 1 bit for parity checking. The quality of the imagery w7as limited by the projected 30 minute surface lifetime and the transmission rate of 256 bits s. To send a panorama at 3.5 seconds per line would require 30 minutes. The panoramas w7ere to be transmitted simultaneously on separate VHF channels. Because scientist were worried after Venera 8 that the illumination at the surface would be very weak, each camera was provided with a 10,000 lux floodlight system with tw7o lamps in order to ensure that there would be sufficient light to acquire an image.

The deployable densitometer had a cesium-137 irradiation source and detectors to measure the gamma rays reflected back by the environment. During the descent this measured atmospheric scattering. Immediately after landing it deployed a 4 x 36 cm

‘paint roller’ on the surface in order to measure soil scattering. In addition, a sodium iodide gamma-ray spectrometer similar to that of Venera 8 was carried inside the sphere for measurement of potassium, uranium and thorium soil abundances. Two anemometers were mounted on the upper side of the aerobrake disk.

Mission description:

Venera 9 orbitev:

Venera 9 was launched on June 8, 1975. It maneuvered on June 16 and October 15 to align its trajectory with the desired entry point in the atmosphere of Venus. After releasing its entry system on October 20, it performed a 247.3 m/s deflection burn to head for the orbit insertion point, where on October 22 it made a 922.7 m/s burn and inserted itself into an orbit with a period of 48.30 hours. The relay of the transmission from the entry system and lander followed immediately thereafter. This was the first spacecraft to enter orbit around the planet. The initial orbit was 1.500 x 111,700 km inclined at 34.17 degrees. This was changed to 1,300 x 112,200 km and finally to 1,547 x 112,144 km at 34.15 degrees. The orbiter conducted 3 months of scientific observations which were terminated by the failure of its transmitter.

Venera 9 lander:

The entry system penetrated the atmosphere of Venus at a speed of 10.7 к m/s and at an angle of 20.5 degrees. At 05:13 UT on October 22 the lander touched down at a speed of 7 to 8 m/s on the day-side of the planet at 31.01 N 291.64 E, where it was 13:12 Venus solar time and the solar zenith angle was 33 degrees. The site was on a slope of 15 to 20 degrees and the lander was tilted a further 10 to 15 degrees by the uneven, rocky surface. It immediately began its surface activities, relaying its data to Earth via the orbiter until that flew out of range 53 minutes later, by which time the temperatures inside the lander had risen to 60 C.

Venera 10 orbiter:

After launch on June 14. 1975, Venera 10 flew7 almost the same route as Venera 9. making trajectory corrections on June 21 and October 18, releasing its entry system on October 23 and then making a 242.2 m s deflection burn. On October 25 it made a 976.5 m/s insertion burn. Its initial orbit was 1,500 x 114.000 km inclined at 29.50 degrees with a period of 49.38 hours. This was later changed to 1,651 x 113,923 km at 29.10 degrees. After relaying the transmission from the entry system and lander, the orbiter began its scientific observations. It succumbed 3 months later to the same problem as disabled its partner.

Venera 10 lander:

The entry system penetrated the atmosphere of Venus at an angle of 22.5 degrees at 01:02 UT on October 25. The lander touched down at 02:17 UT at a speed of about 8 m/s at 15.42 N 291.5UE, about 2,200 km from w here Venera 9 landed. It was on the day-side, at 13:42 Venus solar time with a solar zenith angle of 27 degrees. The surface was fairly level but the lander was perched on a rocky mass that tilted it at about 8 degrees. The lander was still transmitting when the orbiter flew out of range, curtailing the relay operation after 65 minutes.

Results:

Venera 9 lander:

Entry and decent science

The Venera 9 lander inferred atmospheric density from accelerometer data between the altitudes 110 and 76 km. It directly measured atmospheric temperature, pressure, composition and light levels from 62 km to the surface. Light scattering data from the nephelometer. together with the photometer data, indicated clouds with a base at an alii Hide of 49 to 48 km and a lower loading of aerosols extending down to about 25 + 5 km. The clouds were similar to a light fog. with much smaller drop size than normal for Earth and a visibility of several kilometers. Distinct layers were detected at altitudes of 60 to 57 km, 57 to 52 km and 52 to 49 km. The refractivity index was measured to be as high as 1.46; much higher than for water ice and consistent with sulfuric acid droplets. The cloud particles all scattered light but there was absorption in the blue and this, along with heavy Rayleigh scattering, led to increasing orange color with depth. The atmosphere below 25 km appeared to be free of aerosols. Red light was found to reach further towards the surface than blue, shifting the spectrum towards longer w avelengths, producing both orange colored skies and, by reflection, an orange tinted surface. Doppler data provided altitude profiles of horizontal wind speed and direction.

The detailed chemical composition measurements attempted by the first landers of this type gave poor results. The mass spectrometers did not function properly due to inappropriate cleaning procedures prior to launch and apparent clogging of the inlet system by cloud particles. The mixing ratio of molecular nitrogen to carbon dioxide was determined. Argon was deteeted in the atmosphere. A large ratio of argon-36 to argon-40 was measured and confirmed by later missions, but went unreported owing to mistrust of the instrument. Near-infrared photometer results for the water vapor mixing ratio proved to be spurious when spectral measurements were conducted on later missions.

Surface science

The photometers detected dust raised by the landing, but this quickly settled. The surface conditions were 455 + S C and 85 + 3 bar, and there w? as a light w? ind of 0.4 to 0.7 m/s.

Only one 180 degree panorama was taken because the cover for the other camera failed to deploy. This back-and-white image was the first picture from the surface of another planet. It showed a level landscape with a variety of flat, apparently young angular rocks without much erosion. Л portion of the image extended to the horizon, and there w’as no indication of dust in the atmosphere. The illumination was similar to Earth mid-latitudes on a cloudy summer day, and the light scattering did not east shadows. The floodlights for the camera were not triggered on. They were eliminated from subsequent missions. The visibility was a pleasant surprise to the scientists who, after reviewing Venera 8, had predicted a dark, murky and dusty atmosphere in which only the near field would be available for inspection.

The indistinctness and apparent nearness of the horizon in all of the images from the Venera landers were due to the high refractivity of the dense atmosphere which made Venus appear to be a small-diameter spherical body with a horizon much less than 1 km away. The phenomenon is similar to a terrestrial mirage and is probably a function of the observer’s height above the ground.

The gamma-ray composition analysis of the surface material measured potassium, uranium and thorium abundances more typical of terrestrial basalt than meteorites. The fact that the surface rocks differed from primitive meteorites in a way consistent with trends observed in terrestrial rocks indicated that Venus must have been thermal differentiated into a core, mantle and crust. The reflectivity of the surface in five wavelengths was consistent with material of a basaltic composition. The penetrometer indicated a rock density of 2.7 to 2.9 g/’cc.

Venera 10 lander:

Entry and descent science

The Venera 10 lander inferred atmospheric density from accelerometer data between the altitudes 110 and 63 km. ft directly measured atmospheric temperature, pressure, composition and light levels from 62 km to the surface, and structure, microphysical properties and composition of the clouds. The three distinct cloud layers observed by Venera 9 were confirmed. Doppler data profiled the horizontal wind speed and direction during the descent, and then an anemometer measured wind velocity on the surface. As the results were generally in agreement with Venera 9, some conclusions could be drawn on atmospheric convective stability and turbulence. The profile of temperature and pressure showed 33 bar and 158"’C at 42 km, 37 bar and 363°C at 15 km, and 91+3 bar and 464 + 5°C at the surface.

Surface science

As in the case of Venera 9, one of the camera covers failed to deploy and this lander also only provided a single 180 degree black-and-white panoramic image. It showed a surface that was smoother with large, more eroded pancake rocks interspersed with lava or other weathered rocks. The horizon was visible and there was no evidence of dust in the atmosphere. As with Venera 9, the photometers detected some dust raised on touchdown which quickly settled. A surface albedo of 0.06 was derived from the imaging and photometers on both landers.

The surface winds were light at 0.8 to 1.3 m/s. The gamma-ray results and surface reflectivity were suggestive of a basaltic composition. Apparently both landers came down on young volcanic shield structures with lavas close in composition to the tholeiitie basalts that emerge from oceanic spreading ridges on Earth. The penetrometer indicated a surface density of 2.7 to 2.9 g/cc, just as at the Venera 9 site. The surface of Venus appeared to be harder than the Moon or Mars.

Venera 9 and 10 orbitersi

The panoramic cameras returned 1,200 km long images taken using several different filters to distinguish cloud structures and some surface features, although the latter were poorly defined. The results included imagery of the clouds in the ultraviolet, infrared radiometry, photometry, spectrometry of both day and night sides, photo- polarimetry, radio occultation and plasma data. Orbital data suggested a cloud base at an altitude of 30 to 35 km with three distinct layers. The orbiters obtained data on the clouds above 64 km, which is the altitude at which the descent data started. The day-time temperature of the upper cloud was -35°C. warming by about 10 degrees at night. The night-time atmosphere was found to glow in the visible spectral range in bands that later investigation established to be a molecular oxygen band system that is not excited in the Earth’s atmosphere owing to its lower concentration of carbon dioxide.

The airglow spectrometer on the Venera 9 orbiter found optical evidence of night-side lightning, but Venera 10 did not. Reflection spectra of clouds in the infrared at 1.7 to 2.8 microns measured the aerosol scale height near their upper boundary, and infrared wide band radiometry in the range 8 to 28 microns prompted the conclusion that outgoing radiation is systematically stronger on the night-side than the day-side.

The dual frequency radio occult ations at wavelengths of 8 and 32 cm gave a set of temperature and pressure profiles for altitudes in the range 40 to 80 km that revealed details of the night-side ionosphere and the existence of a large diurnal variation of ionospheric electron density. The bistatic radar experiment mapped fifty-five strips of the surface 100 to 200 km wide by 400 to 1,200 km long. Early analysis provided onc-dimensional terrain profiles at a resolution of 20 to 80 km. Later processing of the Venera 10 data produced a two-dimensional local topography for live regions at a resolution of 5 to 20 km.

Measurements of the scattering of solar Lyman-alpha radiation by the hydrogen corona that surrounds Venus, including its line width, gave an estimate of 450°C for

the temperature of the atmosphere at the exobase. Many features of how the solar wind interacts with the ionosphere were measured. No intrinsic planetary magnetic field was detected. Nonetheless, the interaction of the solar wind with the ionosphere created a magnetic plasma tail.

As the first spacecraft to enter Venus orbit, Venera 9 and 10 were able to provide the first long-term survey of the Venusian atmosphere with a comprehensive battery of scientific instrumentation. Their landers performed marvelously, returning the first pictures of the surface of the planet. These missions began what became an unbroken string of successes running to Venera 16 in 1983 and ending with the two Vega missions that made flybys in 1985.

PARTY, GOVERNMENT AND MILITARY

In the Soviet Union there were three separate organizations that ran the country – the Communist Party, the government and the military. The Party was in overall charge through its Central Committee and the executive Politburo. The Party business was managed by a Secretariat that included a Secretary of Defense Industry and Space, and a Department of Sciences. The Soviet Academy of Sciences, while claiming to be independent, was implicitly part of the Department of Sciences and hence a Party organization. It ran the Inter-Department Scientific and Technical Council on Space Exploration (MNTS KI; Mezliduvedomstvennyi Nauchno- Tckhnichcskii Soviet po Kosmicheskim Isslcdovaniyam) which was nominally responsible for specifying the national policy and strategy for the space program.

The Soviet government comprised the Council of Ministers and its executive the Presidium. The Presidium had a Commission for Military Industry (VPK; Voenno- Promyshlennaya Komissiya) which included the various ministers controlling ihe defense industries. The Ministry of General Machine Building (MOM; Ministerstvo Obshchego Mashinostroenija) controlled the planning and budget for the Soviet space program. MOM was the closest equivalent to NASA in the USSR but it had a much wider remit, including the design and production of rockets and space systems for the military. It established, controlled and funded the various design bureaus (OKBs) that developed the rocket and space systems, and the scientific research institutes (Nils) that provided the science and technical support required for space projects. There was no separation of civilian and military space programs. MOM was the focal point of the powerful Soviet military industrial complex. It controlled a massive industrial system, providing funding enormous in scale, and operating in complete secrecy.

How this dual system functioned between Party and government, sitting atop a large and convoluted system of industry, design bureaus and research institutes, is somewhat mystical. It was complicated even more by the third force in the Soviet system, the military. The Armed Service’s Strategic Missile Forces was the

W. T. Huntress and M. Y. Marov, Soviet Robots in the Solar System: Mission Technologies 23

and Discoveries, Springer Praxis Books 1, DOl 10.1007/978-1-4419-7898-1_3,

© Springer Science+Business Media, LLC 2011 heavyweight competing for funding from MOM. and the military controlled the launch pads, the launch ranges, and the і гаек in g stations. In reality, the personal power and political influence of the Chief Designers of the design bureaus such as Korolev, Glushko. Chelomey. Yangel and Babakin were most influential for planning and execution, especially in the early years. These powerful men competed mightily with one another for influence and funding, sometimes bitterly, as between Korolev and Glushko or between Korolev and Chelomey.

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.

wwamttw* ■mv w:*5i’

f-WW fjr. "A’.WJ

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.