THE FIRST LANDER ON MARS: 1971

Campaign objectives:

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

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

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

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

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

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

Spacecraft:

Orbiters:

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

Figure 12.5 Mars-71 S orbiter spacecraft.

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

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

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

Figure 12.6 Mars 3 spacecraft.

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

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

1,210 kg (entry vehicle)

635 kg (lander system on descent)

358 kg (lander)

4,650 kg (total)

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

Entry system:

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

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

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

Figure 12.9 Mars-71 entry system.

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

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

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

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

Lander:

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

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

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

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

Entry у descent and landing:

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Payloads:

M-71S orbitev:

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

Mars 2 and 3 orbiters:

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

F FPU dual camera facsimile imaging system

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

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

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

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

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

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

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

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

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

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

12. Boom mounted three-axis iluxgate magnetometer

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

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

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

Mars 2 and 3 entry systems:

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

1. Accelerometer for atmospheric density during entry

2. Temperature and pressure sensors for descent and landing

3. Radio altimeter for providing altitudes on descent

4. Mass spectrometer for atmospheric composition on descent and landing

5. Atmospheric density and wind velocity on the surface

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

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

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

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

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

Figure 12.18 PrOP-M ‘Marsokhodnik’ rover.

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

Mission description:

M-71S:

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

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

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

Mars 2:

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

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

Mars 3:

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

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

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

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

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

Results:

Orbiters:

Imagery

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

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

Dust storm

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

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