A BETTER SPACECRAFT: A SECOND TRY AT VENUS: 1962 Campaign objectives

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,

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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

Figure 8.1 The 2MV flyby spacecraft (courtesy Energiya Corp): 1. Pressurized orbital module; 2. Pressurized imaging module; 3. Propulsion system; 4. Solar panels; 5. Thermal control radiators; 6. High gain parabolic antenna; 7. Low gain omnidirectional antennas; 8. Low gain omnidirectional antenna; 9. Meter-band antenna; 10. Emergency omni antenna; 11. Camera and planet sensor port; 12. Science instruments; 13. Meter band antenna; 14. Sun and star tracker; IS. Emergency radio system; 16. Continuous sun sensor; 17. Earth tracking antenna; 18. Attitude control nozzles; 19. Attitude control nitrogen tanks; 20. Attitude sensor light baffle; 21. Coarse sun tracker; 22. Sun tracker.

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

Figure 8.2 The 2MV probe spacecraft (courtesy Energiya Corp): 1. Orbital module; 2. Entry capsule: 3. Propulsion system; 4. Solar panels; 5. Thermal control radiators; fi. High gain antenna; 7. Medium gain antennas; 8. Entry capsule test antenna; 9. Meter band transmit antenna; 10. Meter band receiver antenna; 11. Magnetometer and boom antenna; 12. Low gain antennas; 13. Earth sensor; 14. Science instruments; 15. Sun/Star sensor; 16. Emergency radio system; 17. Sun sensor; 18. Attitude control nozzles; 19. Nitrogen tanks; 20. Sun sensor.

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.