A sophisticated lander for scientific exploration of Mars

Planc~t. Spaw Pergamon

Sci., Vol. 46. No. 6/l, pp. 717-737, 1998 c 1998 Elsevier Science Ltd

All rights reserved. Printed in Great Britain 0032p0633/98 $19.00+0.00 PII : S0032_0633(98)0000s-7

A sophisticated lander for scientific exploration of Mars : scientific objectives and implementation of the Mars-96 Small Station V. Linkin,’ A.-M. Harri,3 A. Lipatov,’ K. Belostotskaja,6 B. Derbunovich,’ A. Ekonomov,’ L. Khloustova,’ R. Kremnev,’ V. Makarov,’ B. Martinov,’ D. Nenarokov,’ M. Prostov,’ A. Pustovalov,’ G. Shustko,’ 1. Jlrvinen,’ H. Kivilinna,’ S. Korpela,8 K. Kumpulainen,8 A. Lehto,3 R. Pellinen,3 R. Pirjola,3 P. Riihell,3 A. Salminen,” W. Schmidt,3 T. Siili,3 J. Blamont,’ T. Carpentier,” A. Debus, C. T. Hua,” J.-F. Karczewski,” H. Laplace,’ P. Levacher,14 Ph. Lognonn6,” C. Malique, lo M. Menvielle,” G. Mouli,” J.-P. Pommereau,“’ K. Quotb,“’ J. Runavot,” D. Vienne,4 F. Grunthaner,” F. Kuhnke,16 G. Musmann,16 R. Rieder,” H. Wlnke,15 T. Economou,” M. Herring,‘” A. Lane” and C. P. McKay” ’ IKI-Russian Space Research Institute, Profsojuznaja 82142, Moscow 117810, Russia ’ Babakin Research and Design Center, 24-a Leningradskaya Street, 141400 Khimky Moscow Region. Russia ’ FMI-Finnish Meteorological Institute, Geophysical Research Division, P. 0. Box 503. 00101 Helsinki. Finland ‘CNES-French Space Agency. 3 avenue Edouard Belin, 41201 Toulouse Cedex, France ‘(CNES-French Space Agency. 2 Place Maurice Quentin, 75039 Paris Cedex I, France “Special Research Bureau of Moscow Power Engineering Institute. Krasnokasarmennaja 12. Moscow 1 I 1350. Russia ’ Finnyards Ltd. Electronics, Naulakatu 4.44100 Tampere, Finland ‘Space Systems Finland Inc., Keilaranta 8,02150 Espoo. Finland ” National Technical Research Center-VTT. Otakaari 10, 02150 Helsinki, Finland “‘Service d’Aeronomie, BP 4, 91471 Verrieres-le-Buisson Cedex, Paris, France ” Institute National des Sciences de I’Universe. 2. Avenue de Neptune, 92107 Saint Maur des Fosses Cedex. France ” Institute de Physique de Globe de Paris, 2 Place Jussieu, 52 Paris Cedex, France ’ ’ Laboratory of Terrestrial and Planetary Physics, Bat 502, Universitk Paris Sud. 9 1205 Orsy Cedex, France “‘LAS-Laboratory of Space Astronomy, Traverse du Siphon-12i&me Arr./B.P.X, 14476 Marseille Cedex 12. France Ii Max Planck Institute for Chemistry, Abteilung Kosmochemie, Saarstrasse 24. D-6500 Mainz, Germany “‘Technical University of Braunschweig. Institute for Geophysics and Meteorology. Mendelsohnsstrasse 4. D-4400 Braunschweig, Germany ‘-University of Chigaco. Laboratory for Astrophysics and Space Research. 933 East 56th Street. Chicago, Illinois 60637. L.S.A. ‘“NASA Jet Propulsion Laboratory, 4800 Oak Grove Drive. Pasadena, CA 91109-8099. U.S.A. I” NASA Ames Research Center, Space Science Division (Code SS), Moffett Field. CA 94035-1000, U.S.A. Received

1 March

1997; revised 1 September

1997; accepted

2 September

Abstract. A mission to Mars including two Small Stations, two Penetrators and an Orbiter was launched at Baikonur, Kazakhstan, on 16 November 1996. This was called the Mars-96 mission. The Small Stations were expected to land in September 1997 (L, approximately 178”), nominally to Amazon&Arcadia region on locations (33 N, 169.4 W) and (37.6 N, 161.9W). The fowth stage of the Mars-96 launcher malfunctioned and hence the mission was lost. However, the state of the art concept of the Small Station can be applied to future Martian lander missions. Also, from the manufactwing and performance point of view, the Mars-% Small Station could be built as such at low cost, and be fairly easily accommodated on almost any

* Correspotdetw

fo

: A-M. Harri

1997

forthcoming Martian mission. This is primarily due to the very simple interface between the Small Station and the spacecraft. The Small Station is a sophisticated piece of equipment. With the total available power of approximately 400 mW the Station successfully supports an ambitious scientific program. The Station accommodates a panoramic camera, an alpha-proton-x-ray spectrometer, a seismometer, a magnetometer, an oxidant instrument, equipment for meteorologi& observations, and sensors for atmospheric measurement duriag the descent phase, including images taken by a descent phase camera. The total mass of the Small Station with payload on the Martian surface, including the airbags, is only 32 kg. Lander observations on the surface of Mars combined with data from Orbiter instruments will shed

718

V. Linkin et al. : A sophisticated lander for scientific exploration of Mars

light on the contemporary Mars and its evolution. As in the Mars-96 mission, specific science goals co&d be exploration of the interior and surface of Mars, investigation of the structure and dynamics of the atmosphere, the role of water and other materials containing volatiles and in situ studies of the atmospheric boundary layer processes. To achieve the scientific goals of the mission the lander should carry a versatile set of instruments. The Small Station accommodates devices for atmospheric measurements, geophysical and geochemical studies of the Martian surface and interior, and cameras for descent phase and panoramic views. These instruments would be able to contribute remarkably to the process of solving some of the scientific puzzles of Mars. Q 1998 Elsevier Science Ltd. All rights reserved

1. Introduction The exploration of Mars requires landing missions as one of the tools in the pursuit of enhancing our knowledge of the planet. The landing missions were started in the earIy 70s and the first successful landers were the two Viking Landers in the mid-70s (Soffen, 1977). The next mission to land is the Mars Pathfinder due to arrive in Mars in July, 1997. This paper defines the scientific objectives of a landing mission, and draws the outline of the Mars-96 Small Station, which is a good example of a mission with ambitious scientific goals. The Mars-96 mission was started already in 1988, aiming at a launch in 1992. At that time the payload consisted of an orbiter. multiple Small Stations to be deployed on the surface of Mars, and a balloon that would float in the Martian atmosphere at a low altitude. The Orbiter, Small Stations and the balloon were planned to carry a versatile set of scientific instruments. At the very beginning of the Station development phase it was envisaged that Small Stations would be used as ballast for the balloon. Gradually the balloon would lose its buoyancy and it would drop Small Stations one at a time on the Martian surface and thus regain net lift. This concept gave the name METEGG for the Small Stations during the first development phase back in 1988. The word METEGG came from METeorological EGG describing the idea that a balloon would drop Small Stations as if it were laying eggs. On the other hand, METEGG also reflected the fact that the first Small Station concept was concentrating on meteorological observations (METEGG report, 1988). When the development work continued, it was discovered that the concept of using Small Stations as a ballast for the balloon was not feasible, when taking into account the scientific requirements. Later on also the balloon was omitted altogether from the Mars-96 mission, due to technical difficulties. Eventually, the Mars-96 mission comprised an Orbiter, two Small Stations and two Penetrators (Mars-94 overview, 1992). The main responsibility of the Mars-96 mission rested with the Russian space institutes that also provided the launch. Scientists from about 20 countries participated in the mission. The mission came across two changes of

Table 1. The orbit parameters of the Mars-96 Orbiter

Mars-96 Orbiter orbit parameters CI: 52,000 km e: 0,8755 T: 43.09 h i: 106” 0: 107” 1%’ : 152”

Hn300

km

launch window due to programmatic, technical, and financial difficulties. Finally, the unsuccessful launch took place on November 16, 1996. The rationale behind Martian exploration is primarily scientific curiosity, the neighbouring planets have always been in the focal point of mankind’s pursuit of knowledge (Kieffer et al., 1992; Martin et al., 1992). Along with the basic scientific research, the comparative planetology between Mars and the Earth is an extremely interesting topic to be investigated. The Earth and Mars have a lot in common because their rotation periods and inclination angles are almost equal. The planets are also comparable in size. Also, due to the lack of open water bodies and surface vegetation and to the cold and thin atmosphere, the history of Mars is partialIy preserved on its surface structures. By studying the present state of Mars we may be able to infer the past. and thus be able to predict the future of the Earth.

2. Mars-96 Small Station mission outline The Mars-96 mission, including four landing elements, was launched by Proton, a Russian launcher with an excellent launch record. The launch took place at Baikonur, Kazakhstan, on November 16, 1996. The cruise to Mars was to take about 10 months. The two Small Stations were planned to be separated from the spacecraft 4 to 5 days before the spacecraft was to adopt an intermediate orbit around Mars. With a small delta-v the Stations would have started their descent toward the Martian surface. After that the Mars-96 Orbiter would have made orbit manoeuvring and finally reached an orbit with the parameters presented in Table I. In addition to the scientific payload, the Orbiter was carrying telecommunications equipment to communicate with the Small Stations and the Penetrators. The data transmission from the landing elements to the Earth were planned to be relayed by the Mars-96 Orbiter and the Mars Global Surveyor (MGS), both of them carrying the same kind of data transponders. The Penetrators would have been deployed on Martian surface about one month later than the Small Stations, when the Orbiter was already on the working orbit around Mars. The landing ellipses of the Small Stations and Penetrators were estimated to be 320 km x 2200 km. The estimated co-ordinates of the landing sites are shown in Table 2. The Small Stations would have arrived at Mars on September 12, 1997. The nominal lifetime of the Stations is one terrestrial year, equivalent to about half a Martian year. The system responsibility for the Small Stations was divided between the Russian Space Research Institute (IKI), the Babakin Engineering Research Center (NPOL),

V. Linkin et cd. : A sophisticated lander for scientific exploration of Mars Table 2. The planned locations of the Mars-96 landing elements Landing element

latitude

longitude

Small Station 1 Small Station 2 Penetrator I Penetrator 2

37.6 N 33 N 36’ N 36’ N

161.9’W 169.4.W 161”W 251 w

the Finnish Meteorological Institute (FMI). and the French Space Agency (CNES). The prime responsibility rested with IKI. who provided the mechanical structures and thermal subsystems, the primary power system, the radio equipment and the descent systems. The manufacturing and the final integration work of the Small Station, as well as the decontamination of the Station body and some instruments, was performed by NPOL. FM1 was responsible for the central electronics unit ((‘EU) taking care of the data management and commanding of the Small Station, as well as the power conditioning subsystem. CNES provided the data relay onboard the Mars-96 Orbiter and the MGS. the Station batteries, and the components needed for data protection coding and image compression. CNES contributed to the methods and means to assess and control decontamination. and provided the integration facility based on a laminar flow tent to satisfy the planetary protection requirements. The scientific instruments came from various institutes in many countries. lK1 provided some sensors for meteorological observations and the panoramic camera and FM1 developed meteorological equipment. CNES was responsible for the seismometer, the atmospheric optical depth instrument and the descent-phase camera, as well as sharing the responsibility for the Magnetometer with DARA (Germany). DARA also provided the alphaproton-x-ray spectrometer with University of Chicago (U.S.A.). NASA/JPL (U.S.A.) was responsible for the instrument to measure the oxidant properties of the Martian surface material.

Small Station mass (kg) 87 33 12

After separation from S/C At Mars (with airbags) Payload mass Fig. 1 The Small Station in cruise configuration cm)

Zone ofvisibility f

(dimenxion~

in

Arrival at Mars and the delivery of the small stations

I-> Transfer orbit

3. Small Station deployment phase The Small Station launch configuration and the mass breakdown are presented in Fig. 1. The total Station mass during cruise is about 87 kg including the parachute and airbag systems for the descent. The Mars-96 flight to Mars would have taken approximately 300 days. The Station would have been jettisoned 3 to 5 days before the arrival at Mars, the separation pulse giving the proper delta-v for the Station to target its landing area on Mars, as illustrated by Fig. 2. During cruise time the Station is unpowered and the main battery discharged. The spacecraft should start to charge the Station rechargeable NiCd battery about 4 days before separation. At the time of separation the Station battery would be fully charged. Parallel to the NiCd battery there is also a 5 Ah Lithium battery that provides additional energy for the Station during the descent phase. Five and a half minutes before separation the spacecraft

Fig. 2. The scheme Martian

of the Small Station

deployment

on the

surface

switches power to the Small Station, causing the Station to perform self-check routines taking 270 s and producing some 600 bytes of status information. The Station transmits the results of the health-check test to the spacecraft through a 128 bit s-’ serial line. When the health-check data has been transmitted the Station is ready for the kick-off.

V. Linkin et ul. : A sophisticated lander for scientific exploration of Mars

720

During separation the activation switch turns on permanently, the Station will wake up and start heading for Mars. The moment of separation is also the cold start of the Station. From this point onwards until the end of the Station’s life, primary power from the RTGs is flowing to the battery keeping the central electronics powered and thus the Station functional. After the separation the Station is operating totally autonomously, there is no way to command the Station from the Earth.

Table 3. Timing of the Small Station cruise activities

During the free space flight period the Station makes some sensor calibration measurements. Also the magnetometer is activated a few times in an attempt to register some effects of the Martian magnetosphere. After having flown 4 to 5 days in space the Station will reach the Martian atmosphere, and begin to decelerate. A three-axis accelerometer onboard the Station will detect the deceleration. When the Station reaches the outskirts of the atmosphere, its velocity relative to Mars is 5750 f 30 m SC’. During the atmospheric entry phase the Stations are spin-stabilised-at the time of separation from the orbiter the stations are spun up to approximately 12 rotations per second. There is also a kind of a wing in the rear end of the Station helping ‘in stabilising the motion during the entry. During atmospheric entry the heat shield temperature rises by several hundreds of Kelvins. The outermost layer of the heat shield is made of an ablative material which will melt and partially outgass. Thus the overheating of the Station body will be prevented by the phase transformation. Hence the heat shield temperature does not rise above a certain temperature as long as there is still ablative material remaining. During this phase the temperature of the Station interior will rise less than 10 K. When triggered by the internal clock, the Station will poll the accelerometers at regular intervals, trying to detect the atmosphere after four days of space flight. There are four separate time windows for polling the accelerometer. When the first traces of atmosphere have been detected the Station will start a continuous sampling of and temperature and pressure the accelerometers, measurements will be started later on. From the accelerometer, temperature and pressure data we can determine the atmospheric density. pressure, temperature and mean molecular weight as functions of altitude by applying the following equations to each data point along the descent trajectory :

TSW

2ina, P=-

Activity

Time(s)

Battery charging of the two Stations started by Orbiter (4 days before separation) The two Stations SWITCHED ON (TSW) by the Orbiter (331 s before

TSW

separation Station self-check performed by CEU (duration 370 s), including data retrieval from Optimism and MOX CEU dumps data to the Orbiter

TSW + 270 s

(duration 60 s) TSW+330

s

TSW+331 s TSW + 900 s TSW+24:05 h TSW+68: 15 h TSW+87:

I5 h

TSW+92:

15 h

TSW+95:

15 h

TSW+NN

Cables 1 and 2 between Station and Orbiter cut Station jettisoned from the Orbiter Lander time saved at the APX memory as a backup (every 15 min) PTU-sensor housekeeping measurement Magnetometer measurements (24 h before first atm. polling window) MOX switch-ON (5 h before first atm. polling window) Accelerometer switch-ON. First search for atmosphere Accelerometer switch-OFF (if atmosphere not detected. three more polling windows to come : 100: 15ilOl :20, 116: 15/119: 15. 124: 15/125:20) Final pyrotechniques check (IO s after detection of atmosphere)

Table 4. Timing of the Small Station phase

activities

during the entry

Time(s)

Activity

TO

Decrease of atm. speed of Station 1 5485 m ss’. Station 2 5510 or 0.65 g deceleration (trailing edge) reached. Parachute bay cover ejected Aerodynamic/Heat shield ejected Parachute deployed First half of airbag inflated

TO+11 T0+11.4 TO+22 T0+22.2 T0+22.5 TO+23 TD

Second half of airbag inflated DESCAM and pressure and temperature sensors activated High frequency polling of accelerometer Touch Down, Parachute rope cut, local time initialised

C,A V;

where p is the atmospheric density,

m is the Station mass, a, is the acceleration along the path of the Station, CD is the drag coefficient, A is the cross-sectional area of the Station, pr is the Station speed relative to the atmosphere, p is the atmospheric pressure, g is the acceleration of gravity, z is the altitude from the surface of Mars, T is the

atmospheric temperature, p is the mean molecular weight, and R is the gas constant. The timing of the Station actions, starting from the battery charging down to the impact on the Martian surface (Touch-Down), is depicted in Tables 3 and 4. The deceleration of Station will reach its maximum at approximately 18 km above the Martian surface, after which deceleration starts to diminish. During descent the accelerometer values will have been integrated over time to determine the Station speed relative to the atmosphere. Decisions on when to release the heat shield and to open

V. Linkin et 111.: A sophisticated

lander for scientific exploration

the parachute will be made based on the accelerometer values and the Station airspeed. The parachute will be opened either just after the deceleration has reached the value 0.65 g on the decreasing leg of the deceleration curve. or after the integrated loss of the Station’s velocity reaches a certain value (Station 1 : 5485 m s-‘, Station 2 : 55 IO m s-l). After the first release command the parachute is only partially functional due to a rope surrounding it. A second pyrotechnic device cuts the rope after a few seconds resulting in a fully open parachute. The Station is connected to the parachute with a 130 m long rope to avoid entanglement during and after the impact on Martian surface. The airbags are inflated at an altitude of 4 to IO km above the reference surface level, depending on the atmospheric entry angle. to protect the Station from the landing shock. The moment of airbag activation is also the starting time of scientitic measurements. The Station starts to meaatmospheric temperature and pressure and sure DESCAM. the Descent Phase Camera. starts taking images of the terrain below. Accelerometers operate continuously to detect the landing shock, which nominally happens within two minutes after the parachute has been released. The atmospheric entry and descent scenario is depicted in Fig. 3.

Fig. 3. The Small Station atmospheric entry and descent phase. reference surf&e level, depending on the atmospheric entry angle

of Mars

7’1

The Station lands on Mars with a vertical speed of 20 rns-’ causing an impact shock of < 200 g with a duration of 30 to 70 ms. The horizontal velocity is estimated to be 20 m s ’ as well. The airbags will dampen the impact, but nonetheless the Station will bounce several times. Just after the first impact on Mars the rope connecting the Station to the parachute is cut. This action prevents the Station from entangling with the parachute and the rope. Also the possibility that surface winds might drag the Station along the Martian terrain is avoided. After cutting the parachute rope. the Station will wait for about 3 min. before cutting the ties of the airbags and letting them loose. This way there is time for the parachute to be blown away by wind. The sequence of the Small Station actions after the impact can be seen in Table 5. When the Station is free of the airbags it will open its four stowed petals. They will open by means of spring power and lift the Station to an upright position, as illustrated by Fig. 4. The Station will reach the working attitude regardless of its original attitude after the deflation of the airbags. The Station may fail to stand up only if the Station is pressed tightly from two sides by some rigid objects like stones. When the petals are opened, the combined antenna and sensor boom will also be lifted. The boom stands up on four flexible arms that were folded

The airbags

are inflated

at al

titude

of 4 to IO km above the


722

Table 5. Timing of the Small Station activities at the Martian surface Time(s)

Activity

TD+l

DESCAM, PT-sensors and magnetometer switched OFF MOX switched ON Airbag external connections cut off Airbag internal connections cut off Legs opened and booms on legs deployed Meteorology boom and antenna deployed DAY 1 started (first operational day) Determination of the angle between the Station X-axis and gravity vector (accelerometer) Station normal surface operation through daily cyclograms

TD+20 TD+ 120 TD+ 180 TD+845 TD + 905 TDf960 TD+961

while the Station was in stowed configuration. This technique of lifting a research vessel by means of opening petals has been used also in earlier Russian Moon missions. After the Small Station has resumed its working position, the magnetometer, alpha spectrometer and the Mars Oxidant instrument will be deployed by means of their own booms. The Small Station is now ready to start its surface mission.

Fig. 4. The Small Station lifting itself in an upright position

4. Small Station scientific objectives 4.1. General objectives and the network concept The Small Station observations combined with the data from the Orbiter instruments are expected to shed light on contemporary Mars and its evolution, by providing data on the atmosphere, surface and interior. Specific science goals are the following : A exploration of the dynamics and structure of the atmosphere, B the role of water and other materials containing volatiles, C studies of atmospheric boundary layer processes, D surface chemistry and geology, E to obtain new information of the puzzle of an intrinsic magnetic field, and F to study the interior of Mars by recording seismic activity. Scientific experiments performed at a fixed site on the Martian surface would provide information on physical and chemical parameters at that particular point. These data combined with observations made by a satellite orbiting Mars may enable more general conclusions. However, a series of simultaneous observations at several sites on the surface is needed for extensive exploration of the atmo-

V. Linkin et al. : A sophisticated lander for scientific exploration of Mars sphere and the geophysics of a planet. Hence a network of measurement stations would be required. This was also the original objective of the Mars-96 mission. To be able to effectively monitor the circulation and transportation systems predicted by theoretical models of the Martian atmosphere, a network of at least about ten stations would be preferable on the surface of the planet. The internal structure of Mars can be studied by seismological and magnetic measurements. At a single site both of them can only yield qualitative information. The minimum number of stations for a more quantitative seismological analysis is three (Lognonne et al., 1996). Magnetic data collected at several sites may help solve the basic question of the character of the Martian magnetic field : does Mars have an internal field comparable to the terrestrial geomagnetic field? Time variations of the magnetic field may be applied to the determination of the conductivity structure of the Martian ground using electromagnetic induction methods. This also requires simultaneous measurements in a network of stations, in particular as the ionospheric source field is poorly known. Thus, to make meteorological and geophysical measurements more useful and applicable, the establishment of a network of stations on the surface of Mars should be aimed at. Geochemical observations would also, of course, benefit from a network but they do not necessarily require simultaneous measurements at different sites.

123

annual mean. The CO* cycle is reviewed in greater detail in, e.g., James et al. (1992). The atmosphere of Mars is extremely dry, the seasonally and spatially averaged column abundance corresponds to an equivalent thickness of liquid water of less than 10 precipitable pm (pr pm). According to the Viking Orbiters’ Mars Atmospheric Water Detection MAWD (Farmer et al., 1977) experiments the peak value is of the order of 100 pr pm, occurring near the Northern polar cap at L, = 120”. The peak is most likely caused by sublimation of water from the residual polar cap. The Southern summertime peak is only about 15 pr pm. Despite the low absolute water content, in the low prevailing temperatures relative humidity can reach 100% leading to formation of near-surface fog and condensation clouds. Recent ground-based observations have, however, indicated that the Martian atmosphere may in fact currently be even drier than during the Viking epoch (Clancy et al.. 1992). The water cycle is reviewed in greater detail in, e.g., Jakosky and Haberle (1992). According to the Viking observations the Martian atmosphere has a background dust optical thickness larger than a few tenths at all times. The background dust load is due to the frequent dust storms ranging in spatial scale from dust devils to hemispherical and global. The local dust clouds occur every year and every season, but they coalesce and subsequently evolve to planet-encircling storms only in some years. The dust cycle is reviewed in greater detail in, e.g., Kahn ei al. (1992). The major Martian global circulation patterns are (Zurek et d., 1992) :

4.2. Exploration qf‘ Martian atmosphere l

The observations made by the Viking Landers and Orbiters provide by far the most comprehensive observational basis for our current understanding of the circulations of the Martian atmosphere. Additional observational data have since been obtained only from groundbased instruments and from the Hubble Space Telescope HST (e.g., James et al., 1994) in the visible, infra-red and microwave wavelengths (e.g., Clancy et al., 1992). Models and modelling techniques have in the meantime grown enormously in sophistication leading to a discrepancy between them and the observations. In the near future, in situ observations are expected from the Mars Pathfinder and orbital data from the Mars Global Surveyor missions. The observations-combined with the extensive body of global, mesoscale, and boundary layer modelling begun in the late 1960s-have provided us with a rich but obviously also incomplete picture of the contemporary Martian climate and weather. In many respects the picture resembles the terrestrial picture a lot, but striking differences are conspicuous as well. The Martian climate is characterised by the cycles of CO,, H20, and dust, as well as by the interplay between these cycles. The atmospheric CO? condenses onto the fall/winter polar ice cap and sublimates from the spring/summer cap leading to reduced and increased pressures at the respective cap regions. This pressure difference induces an atmospheric mass flow from the spring/summer cap towards the fall/winter cap. The mass flow manifests itself as substantial seasonal surface pressure variation-the difference between the extrema can exceed 20% of the

l

l

l

l

single Hadley cell between the summer hemisphere tropics to the winter hemisphere subtropics ; quasi-periodic baroclinic eddies in the winter hemisphere ; stationary eddies induced by the large topographical variations ; condensation/sublimation flow between the CO? polar caps ; and very strong thermal tidal winds.

Mesoscale phenomena such as summertime slope winds have also been identified in the Viking Lander data (Hess et al.. 1977) and have also been successfully modelled (e.g., Savijarvi and Siili. 1993). Surface thermal contrasts due to variations in CO? ice cover, surface albedo, and soil thermal inertia may also drive mesoscale circulation patterns (Siili, 1996 ; Siili et al., 1997). Originally as part of the Mars-96 mission the Small Stations were intended to continue the time series of Northern hemisphere observations after a long hiatus, and at regions differing from the Viking Lander sites. In view of using the stations or derivatives thereof as part of future missions, the stations have a capable instrument complement able to observe several phenomena and hence provide us with further insight thereto. The surface and near-surface air temperature measurements enable estimation of, e.g., surface heat fluxes (e.g., Tillman, lY94), surface thermal inertia (e.g., Savijarvi, 1995) and-together with the optical sensor-effects of dust on the diurnal temperature extrema. The pressure measurements can be used to monitor passages of baroclinic disturbances


724

(low- and high-pressure systems), the thermal tides, and the seasonal pressure variation due to the condensation/sublimation flow. The observations may also provide insight into the onset of dust storms (Tillman, 1988). No humidity sensors have yet been deployed onto the surface of Mars and the Mars-96 Small Stations were to perform the first in situ relative humidity measurements. Seasonal and diurnal humidity variations and hence global water transport as well as the regolith-atmosphere exchange of water, respectively, are the major topics addressed. The optical sensor (together with the panoramic camera) is dedicated to the monitoring of the dustrelated phenomena and dust characteristics. The camera can also be used to identify surface frost and near-surface fog formation. The wind measurements can identify and monitor directly boundary-layer flows such as mesoscale circulations and more indirectly global phenomena, e.g., baroclinic disturbances and thermal tides.

4.3. Geophysical

exploration

of Mars

The deep internal structure of Mars is still largely unknown. The determination of the Martian internal structure, with the knowledge of the structure of the Earth, will provide new constraints on the theory of the solar system formation (e.g. Zharkov and Trubitsyn, 1978). The structure of the Martian interior is most effectively determined by seismological experiments. The success of seismological studies depends on the level of seismicity. Due to the lack of plate tectonics, which, on the Earth, generates most quakes, the seismicity of Mars is likely to be much weaker than on Earth. It is currently estimated that the seismic activity should be somewhere between that of the Moon and the Earth. Modelling experiments show that seismic moment releases between 10” and 10” Nm per year, about two orders of magnitude more than the seismic moment released by shallow moon quakes. With such models, more than 10 events of magnitude greater than 4.5, and more than 250 events of seismic moment greater than 3.3, may be expected per year, which represents a significant potential for remote quakes detection. This is a clear justification for seismological experiments on Mars. A first attempt to address these key seismological questions was made by the Viking seismic experiment (Anderson et al., 1977). However, it did not result in convincing Mars-quake detection, basically because of too strong a wind sensitivity and too low a resolution during nonwindy conditions. The Optimism seismic experiment was expected to re-open the seismic exploration of Mars. Specific scientific objectives of the Optimism seismic experiment are (Lognonne et al., this issue) : l

l l

To determine the micro-seismic noise level in the band 2 Hza.02 Hz and to study its variation during the day ; To determine the level of seismic activity of the planet ; To make a first step towards seismic tomography of the planet by constraining the variation of seismic velocities with depth.

The frequency band of the instrument was optimised for the remote detection of quakes, taking into account the a

prioriinformation about the internal structure and attenuation of Mars (Lognonnt and Mosser, 1993). Most of the teleseismic P body wave energy is then expected to be found at periods of l-2 s, and the other phases at periods of 226 s. Performing magnetic measurements on the Martian surface will provide information on the puzzle of the magnetic field. The big question is, whether there is an intrinsic magnetic field, or is the observed magnetic field induced by the ionospheric currents. Fast changes in the data series will be an indication of the latter. The magnetometer may also detect remnant magnetisation of rocks near the Station, which could have been caused by a paleomagnetic field of Mars. The magnetic field of a planet may result from two primary sources, convection in a conductive liquid core and the interaction between the planet and solar wind plasma, and from two secondary sources, the local magnetisation of crustal and lithospheric rocks and electric currents induced in the planet by the transient magnetic fields of external origin. The different contributions to the magnetic field of a planet are presented in greater detail by Viljanen and Pirjola (1994). The study of the magnetic field of a planet would thus provide original information in various domains, as its past and present core dynamics, its lithospheric past evolution and present structure, or the electrodynamics of its ionised environment. Our present knowledge of the magnetic field of Mars was obtained through the data collected by the U.S. (Mariner-4, -9. and Viking) and Soviet (Mars-2, -3, and -5, then more recently Phobos-2) missions (Russell, 1980: Zakharov, 1992). During the descent of the Viking Landers, peak ionospheric densities of 10” cm-” have been measured at an altitude of 130 km above the planet’s surface. We have therefore no direct information about the magnetic field close to, and at the Martian surface. The Optimism (Observatoire PlaneTologique : Magnetlsme et Sismologie sur Mars) (Menvielle et al.. 1996; Kuhnke et al., 1998) aimed at getting direct information on the transient magnetic variations at the surface of Mars. The scientific objectives of the Optimism magnetic experiment, in order ofpriority, were to : (1) probe its internal structure by means of electromagnetic sounding methods; (2) characterise the time variations of the planetary magnetic field, and its spatial variations if measurements are made during the descent phase; and (3) investigate the space and time characteristics of the magnetic field sources in the planetary environment. Little is known about the behaviour of the magnetic field due to currents flowing in the ionised environment of a planet like Mars. The exact nature of the interaction between the solar wind and the high ionosphere is still unknown because of the lack of adequate observations. Whatever physically takes place, this interaction will result in a transfer of momentum from the solar wind to the plasma originating from Mars on the sunward and duskside of the planet. Qualitative considerations show that the different current systems which may exist must induce magnetic field perturbations at the surface whose maximum amplitude ranges between 10 to 20 nT, with typical variation times greater than 30 to 60 s. The morphology of the magnetic variations at the surface of Mars is expected to be intermediate between those of the IMF

V. Linkin e/ al. : A sophisticated lander for scientific exploration of Mars (Interplanetary Magnetic Field) and those resulting from the Earth environment filter (see e.g.. Menvielle rt al., 1996). In order to limit, and if possible cancel out contamination of the measurements by the Small Station, DC or low magnetic fields, the magnetometer sensor is installed outside of the Station. at the end of a hinged boom to be deployed after the landing of the Station and the opening of the petals. Because the magnetometer sensor is installed at the end of a hinged boom, its orientation should be determined during surface operations. Given the constraints on power and mass budgets, a simple way to determine the attitude of the magnetometer sensor is to have low mass and low power consumption attitude restitution devices rigidly mounted on the magnetometer sensor itself. A non-magnetic attitude restitution device made up with a Z-axis tilt meter and a sun sensor were thus developed in the frame of the Optimism magnetic experiment, and installed on the magnetometer sensor, inside the thermal housing. The determination of the tilt angles relies on measurements of the impedances between electrodes partially immersed in a conductive liquid. That of the direction of the geographic North relies on the recording of the variation in the inten5ity of the light passing through plane slits in a thick black cover mask. The tinal accuracy in the determination of the direction of the local vertical is better than half a degree of arc for temperatures in the range -40 C to + 30 C, and that in the determination of the direction of the geographic North may be as low as a tenth of arc degree, provided a large enough number of sun passes are available.

The X-ray fluorescence (XRF) instruments on board the two Viking Landers yielded the first data on the chemical composition of the Martian surface (Clark. 1982). .4lthough the two landing sites were about 6500 km apart from each other the chemical composition of the analysed soil was practically identical. Unfortunately, no rocks could be sampled within the range of the Lander’s arms. Thus we do not have any data on Martian rocks, except for the information obtained from SNC-meteorites, which are generally assumed-but not proven with absolute certainty-to represent Martian surface rocks ejected into space by large impacts. The XRF instruments on board the Viking Landers could not analyse elements lighter than magnesium. One of the most fascinating topics in Mars science is the climate history of our neighbour planet. Today, Mars is a desert planet with surface temperatures between about 250 K at noon in the equatorial regions and less than about 140 K at the South pole. Hence, water can only be expected in the form of ice, even CO, freezes out at the winter poles. However, large erosional features at the Martian surface in the form of outflow channels, rivers and lake beds seem to indicate a warm and wet climate in the early history of Mars. Such a climate is only possible if we assume a drastic green house effect, e.g. by a dense CO, atmosphere of 5 to 10 bars. It has been a question

715

whether CO2 alone could account for this effect, but the most important question is: “Where did the CO, go?“. The most widely discussed scenario assumes transformation to carbonates. So far no unequivocal detection of carbonates at the surface of Mars has been made. The Alpha-proton-X-ray spectrometer (APX) is able to analyse carbon down to about 0.1 w O/o,thus being capable of bringing new clues to this puzzling question. The APX spectrometer can also measure the absolute concentration of oxygen thus providing information on the oxidation stage of the soil. The Viking XRF instruments found about 0.7% Cl and up to 3.5% S in the Martian soil. Both elements could easily be determined by the APX. but with additional data on the total concentration of oxygen one can hope to find out if sulphur is present in form of sulphates as is widely believed-or in elementary form. Another element of high interest is potassium. Potassium concentration in SNC-meteorites ranges from 180 ppm to 1.560 ppm with a mean value of 680 ppm. However, the concentration obtained by the Russian Gamma-Ray instrument on the Phobos mission is 3000 ppm. which is higher than the upper limit of 1200 ppm given by the Viking instruments. Taking the SNC data as a baseline, this would require some concentration mechanisms in the Martian regolith. The APX spectrometer would yield additional K data and thus shed some light on this question. Although there is considerable evidence that the SNCmeteorites represent Martian surface rocks, the last doubts about this hypothesis can only be removed by comparing their elemental and/or isotope composition with high precision data of Martian rocks. With respect to isotope composition one probably has to wait for sample return. The elemental composition of the twelve known SNC-meteorites is quite variable, but certain characteristic element ratios are almost invariant and quite different from all terrestrial. lunar and meteoritic samples. The most characteristic ratio would be K!La or K/U. Unfortunately. the sensitivity of the APX spectrometer is not sufficient for the determination of La or 1J at the expected concentration levels. Not quite as characteristic but still highly valuable is the Fe:Mn ratio, which should be measurable by the APX spectrometer: the same holds for the ratios of Mg,‘Cr and N$AI. Studying the reactivity of the Martian soil and the atmosphere would enhance our understanding of the Martian geochemistry. and to some extent also exobiology. This would be accomplished by a Mars Oxidant Experiment. MOx. that would reveal the abundance of peroxides. superoxides and other reactive compounds. Both soil and the atmosphere within 10 cm from the soil will be explored. The main objective of the MOx is to study the reactivity of the Martian surface layer in the light of the oxidant hypothesis stating that one or more inorganic oxidants are responsible for the fact that, unexpectedly. the biology experiments of the Viking Landers did not find any indication of organic materials (Klein. 1978). The absence of organics in the Martian regolith may be explained by identification of the reaction mechanisms. This may also assist in predicting the depth to which one has to penetrate to be able to discover some organic compounds being relict of H period of chemical


126

and organic evolution. The MOx instrument will also study the geochemical weathering processes of the Martian surface. If any type of primitive organic compounds or even traces of the old biosphere on Mars would be discovered, it would be an issue of utmost importance. The main scientific objectives of the Small Stations do not include direct exobiological experiments, but indirectly the observations of the Small Stations may also contribute to this interesting puzzle of some type of life on Mars.

5. SmalI Station instruments 5. I . Pavload overoiew

To satisfy the scientific objectives the Small Station has a versatile set of instruments onboard. There are devices for atmospheric measurements, composition studies of the Martian surface and surface chemistry, seismic and magnetic measurements and cameras for descent phase and panoramic views. The scientific payload with responsible institutes and the scope of each instrument is presented in Table 6. Each individual instrument has its own primary scientific objectives, and the overall scientific return of Small Station is gained by combining the results of all the instruments. Also, the observations made by the Orbiter instruments will be taken into account when performing the data amalgamation. Table 7 illustrates in which manner

Table 6. Scientific instruments

the Small Station scientific payload responds to the scientific objectives. In the design of the Small Station, the opening petals have been made use of by locating instruments on three petals. The magnetometer, MOx and the APX that are mounted on the petals, also have their own opening arms taking them further apart from the vicinity of the Station. The instrument locations are depicted in Fig. 5. 5.2. Descent Phase Instrument (DPI) During the descent through the Martian atmosphere the Small Station performs temperature and pressure observations, as well as deceleration measurements that indirectly provide the atmospheric density. The DPI sensors are thus an accelerometer, and temperature and pressure sensor. The DPI pressure sensor, being of different type from the pressure sensor used at the Martian surface, is located outside the Station payload bay, but it is still under the cover of a petal. Accordingly, the pressure readings will be affected by the atmospheric turbulence generated by the Station body, as well as the delaying and damping effect of the Station covering shell. The main DPI characteristics are shown in Table 8. 5.3. Camera system (DESCAM

and PANCAM)

There are two cameras onboard the Small Station. The Descent Phase Camera (DESCAM) starts taking pictures

of the Small Station

Instrument

Responsible Institute

Description

PANCAM DESCAM

Russian Space Agency (IKI) Laboratorie Astronomie (Marseille) INSU/CNRS, IPGP, U. of Paris Tech. U. of Braunschweig Finnish Meteorological Inst./GE0 Max Planck Inst./Univ. of Chicago Service d’Aeronomie Russian Space Agency Russian Space Agency JPL/NASA

Panoramic Camera/Surface Imaging CCD/Surface Imaging during the descent phase Seismometer, Optimism system section Magnetometer sensor Atmospheric pressure and humidity Alpha-Proton-X-ray spectrometer Atmospheric transparency, aerosols, ozone Surface T and W, vertical T and P profiles Vertical density profile of the atmosphere Oxidant meter/Surface chemistry

OPTIMISM

P and RH-sensors APX Optical Sensor T,P,W-sensors Accelerometer MOx

Table 7. The response of the Small Station payload to the scientific objectives

Instrument PANCAM DESCAM Seismometer Magnetometer P and RH-sensors APX Optical sensor T,P,W-sensors Accelerometer MOx

Atmosph. Dynamics

Atmosph. Structure

Atmosph. PBL

Role of water

X

X

X

X

interior of

Surface chemistry

Geology

Magnetism

Mars

X X

X X

X

X

X

X

X X X

X

X X

X

X X

X

V. Linkin PI u/. : A sophisticated


of Mars

127 Wind, temperature and humidity sensors

Thermal

Pressure sensors inside

MOx sensor Fig. 5. The Small Station with instrument locations. The seismometer and pressure sensors are inside the electronics frame, as well as the Central Electronics Unit (CEU) and the other key system devices. The footprint of the Station body has the diameter of about 60 cm, the petals are about 35 cm long each. The tip of the sensor boom will be about 70 cm above the ground

Table 8. The DPI resolution, power figures Measurement

Dim. (electr.) Dim. (sensor) Mass Oper. power

mass, dimensions

and operational

Atm. density during descent, resolution < 0.001 g Atm. pressure. resolution 0.01 hPa Atm. temperature during descent, resolution 0.2 K 90mn~x160mmx15mm Smmx5mmx2mm(T) 200 g 150 mW

11 s after the heat shield has been ejected. It takes about 80 nested pictures of the underlying terrain producing about 1.6 MBytes of data. The resolution varies from 20 m to 1 cm. The DESCAM is switched off just after the impact. The DESCAM will be physically detached from the Station within two minutes after the touch-down. namely, the DESCAM cables will be cut before the airbags are kicked out. The panoramic camera (PANCAM) operates throughout the lifetime of the Station. The PANCAM has a CCD line-array detector with 1024 pixels. Panoramic pictures are taken by rotating the camera head by way of step motors. and taking a line array record at every step. The total panoramic view will consist of 6000 line-array records obtained when the PANCAM is revolved once around itself one step at a time. The PANCAM pictures give information on the geological structure, like stratification visible on the near-tovertical walls, traces of weathering and fluid erosion and results of volcanic activity. Also some mineralogical information may be collected by means of wavelength analysis of the images.

The light seen by the PANCAM consists of three components: direct sun light, sun light scattered by atmospheric aerosols, and sun light reflected by other surfaces. The last contribution is negligible compared with the first two components because of the low albedo (0.2-0.4) of the Martian surface, and a low solid angle of reflecting surfaces as viewed from the illuminated point. The first component mainly prevails when the atmosphere is transparent, resulting in large contrasts with deep shadows in the images. The scattered component is substantial only in the shadows but during a dust storm the second component becomes dominant, resulting in a low contrast image without shadows. Just after the Station has landed on the Martian surface, the camera memory is mostly filled by DESCAM images, about 390 kBytes of the 1.6 Mbyte camera memory is permanently reserved for PANCAM. This 390 kBytes is just enough to store one full compressed PANCAM image. The subsequent PANCAM images will be taken after more of the camera memory is vacated as the DESCAM images are transmitted. The cameras share a common data compression unit, which uses the JPEGtype discrete cosine transformation (DCT) algorithm with the compression factor of eight. The PANCAM and DESCAM principal facts are presented in Table 9. The high quality of PANCAM is underlined by comparing the performance of PANCAM with the panoramic camera of the Viking Lander, as can be seen in Table IO.

5.4. Meteorological sensors (PTUW) The Small Station meteorology package (PTUW) consists of four Barocap’ pressure sensors, two Humicap’K’ humidity sensors (Harri et al., 1995) two temperature sensors and one wind sensor. The wind, humidity and

728


Table 9. The main facts of the descent phase camera (DESCAM) and the panoramic Camera (PANCAM)

Table 11. The principal

DESCAM Measurement

Measurement

Dim. (electr.) Dim. (optics) Mass Oper. power PANCAM Measurement


Pictures during descent, 500 x 400 pix. resolution 20 m- > lcm 130mmx70mmx45mm 50 mm x 50 mm x 100 mm 500 g (electronics shared with Pancam) 1w Panoramic images of the landscape around the Small Station, a full compressed picture < 400 kbits IlOmmx llOmmx40mm 50mmx20mmx35mm 600 g (electronics shared with Descam) 1w

Dim. (electr.) Dim. (sensor)

Mass Oper. power

between the PANCAM and the Viking

Parameter

PANCAM

Viking camera

Spectral range (pm) Mode Image format Frame (pixels) Resolution Linear resolution Field depth AD conversion Sampling time Mass

0.4-0.8 SlW 66 x 362 1024 x 6000 0.06 0.5 mm 0.5 m-infinity 12 bit 0.3-3 h 0.6 kg

0.4-1.1 colour 60 x 360 1500 x 9000 0.04-O. 17 1.O mm 1.O m-infinity 6 bit 2h 7 kg

of the meteorological

Atm. pressure, Accuracy < 0.05 hPa Atm. humidity, day and night variation Atm. temperature, accuracy < 0.5 K Atm. wind. accuracy < 1 m s-’ 90mmx160mmx15mm (615mmx9mm(P), lOmmx4mm(U). 5mmx5mmx2mm(T),~70mmx40 mm (WI PU 50 g, TW 300 g PU 20 mW, TW 150 mW

Table 12. The key figures of the Optical Depth Sensor (ODS) Measurement

Table 10. Comparison panoramic camera

characteristics

sensors (PTUW)


Optical thickness. resolution I % Aerosol distribution, high altitude clouds, ozone 107mmx97mmx30mm 430 mm x 120 mm (incl. boom) 355 g (Optical box 115 g, electronics incl. cable 240 g) 180 mW (9 mA from + 15V, 3 mA from - ISV)

The main facts of the meteorological Martian atmosphere are shown in Table

sensors 11.

for the

5.5. Optical Depth Sensor (ODS)

temperature sensors are located in the end of the fourlegged sensor boom that will lift up about 70 cm above the surface. The wind sensor is in the uppermost tip of the boom, and the temperature and humidity sensors are some 5 cm lower in the end of small whiskers extending horizontally 10 cm away from the sensor boom. The four pressure sensors are mounted on the sensor electronics board located inside the Small Station. The same electronics board also accommodates the transducers for humidity sensors, as well as for the Thermocapt@ temperature sensors onboard the sensor electronics. The Thermocaps are used for temperature compensation of the pressure sensors. The temperature compensation of the humidity sensors is made by means of a platinum resistor type of temperature sensor mounted on the humidity sensors. The sensor boom orientation is perpendicular to the horizontal plane of the Small Station instrument platform. Hence, depending on the orientation of the Station, the heat flow coming from the Station will affect the temperature and humidity sensors, when the wind is calm. When designing the meteorology package, it was envisaged that such conditions can be discovered by flow simulations, and by inspecting the data from all the sensors. The effect of the heat flow can be partially subtracted from the data, resulting in a modest increase of the error margins of the temperature and humidity data.

The main objective of the Optical Depth Sensor (ODS) is to investigate the atmospheric optical thickness by analysing the sunlight spectral variations. The ODS measures the intensities of the direct sun light and the zenith scattered light. The ratio of these intensities measured in several wavelengths gives the atmospheric optical depth as well as the aerosol size distribution. In twilight the ODS observes the zenith scattered light to observe the high altitude clouds. The ODS will also give an indication of the total amount of atmospheric ozone. The instrument has four photodiode-based detectors : three narrow band detectors with interference filters giving +20 nm detection windows at 270 nm, 350 nm and 550 nm, as well as one detector with a broad band ranging from 250 nm to 750 nm. The ODS measurements are strongly linked to meteorological observations of pressure, temperature, humidity and wind (PTUW). The data from ODS and PTUW will be analysed in a coordinated fashion. The ODS main figures can be seen in Table 12.

5.6. Alpha-Proton-X-ra),

Spectrometer

(APX)

The Alpha-proton-X-ray spectrometer measures the elemental composition of the Martian surface. The method consists of recording the energy distributions of backscattered alpha particles, protons from nuclear reac-

V. Linkin ct trl. : A sophisticated

Table 13. The main characteristics Measurement

lander for scientific exploration of the APX

Dim. (electr.) Dim. (sensor)

Elemental composition, Carbon and heavier, accuracy 1% Some minor and trace elements, accuracy better than 0. I% Excitation source : alpha-particles from Curium 244. 70mmxl00mmx80mm bS2 mm x SO mm (270 mm opening boom)

Mahs

500 g

Oper. power

300 mW

of Mars

129

Table 14. The main characteristics of the Optimism including a magnetometer and a seismometer Measurements

Dim. (electr.) Dim. (sensor) Mass

Oper. power

instrument

Seismic signal of remote and regional quakes, resolution 0.5 nm at 2 s Magnetic field, resolution 0.25 nT 64 mm x 75 mm x 40 mm (Mag) 90 mm x 90 mm x 95 mm (Seismo ). (b67 mm x 90 mm (Mag) Mag-sensor head 1 I5 g. Mag-boom 180 g. Mag-electr. I04 g, Seismometer 405 g. Optimism electronics 635 g Magnetometer I85 mW. Seismometer 5

mW (own battery)

tions and characteristic X-rays, all generated by bombardment of the sample with primary alpha particles from the instrument’s source. The source and detectors are lifted by an opening boom to a distance of some 20 cm from the Station. During the cruise phase the boom, the detectors and the source are in a stowed position mounted on the inner side of one of the petals. When the petals are opened, the APX boom will be extended. The APX electronics responsible for processing and storing the data as well as communicating with the Central Electronics Unit of the Station is located inside the Station. Alpha particles and protons are measured in an energy range of 0.4 to 6 MeV, the X-rays are measured in an energy range of 1 to 10 keV. The main technical properties of APX are shown in Table 13.

5.7. Optinzim The Optimism instrument consists of a magnetometer, a seismometer, and an electronics unit controlling measurements and communicating with the CEU. The Optimism main electronics, used for both the seismometer and the magnetometer, has a mass of 635 g, including 200 g of non-rechargeable Lithium batteries. These batteries will be used during the first 100 days of the mission, when the Small Station power is used for other experiments and for data transmission. All the digital electronics is based on a 8OC3 1 micro-processor system on hybrid electronic circuit boards, and has a mass of 140 g. They include 2Mbits of data memory, and a 16 bit A/D converter. The seismometer measures the vertical acceleration by detecting the displacement of a specific seismic mass. The instrument operates in the frequency range from 0.02 Hz to 2Hz. The seismometer is mounted in the lower part of the payload bay. The mass of the seismometer is 405 g, forasizeof9x9x9cm3. The seismometer is a pendulum, supported by a leaf spring with a resonant frequency close to 2 Hz, located in a half-sphere made of titanium and under vacuum. The high gain mode of the seismometer has a resolution of about 5.10. ” rns-’ at 2 s, corresponding to a ground displacement of 5.10-“’ m. For signals with the period of 1 s the Optimism seismometer is about twice more sensitive than the Apollo short period seismometer, although at 2 s periods its performance is twice less than the Apollo long period seismometer in flat mode. The Optimism seis-

mometer is about two orders of magnitude more sensitive than the Viking seismometer. and is therefore very promising for the detection of the first Mars quake. When designing the Optimism seismometer, special attention was paid to the wind-induced noise, even if the Small Station noise level was expected to be significantly lower than that of the Viking Lander. A carbon epoxy ring was therefore delivered by the OPTIMISM team and used in the lander structure to increase the rigidity of the structure in contact with the ground and supporting the seismometer. The magnetometer is a three axis fluxgate device. the sensor being attached to one of the Station petals. The sensor is placed about 1 m away from the Station by means ofa boom with three hinges. In this way the magnetometer avoids the magnetic contamination generated by the Small Station, even if the Station was designed to be as magnetically clean as possible. The magnetometer and the boom are in stowed position under the petal during the cruise phase. The magnetometer dynamic range is kg000 nT with the resolution of 0.25 nT. Variations of the three components of the magnetic field will be measured with an interval of roughly 30 s. The main figures are shown in Table 14.

5.8. Mars Osidant

E.uperimtwt

( MOs)

The Mars Oxidant Experiment using micro mirror-based chemical sensors investigates the reactivity of Martian soil. as well as the near-surface Martian atmosphere. The MOx micro mirrors are coated by an array of chemically sensitive thin films encompassing metals, organometals. and organic dyes to produce a pattern of reflectivity changes characteristic of the species interacting with these sensing layers. The instrument includes LED light sources, optical fibre light guides, silicon micro machined fixtures. a line-array CCD detector, a micro controller-based scheduling and commanding system. and 96 separate micro mirrors. The reflectivity of the micro mirrors is monitored in real-time and it changes, when they are exposed to the Martian near-surface atmosphere and soil envisaged to be reactive. The effect of the reactive compounds can be seen in the reflectivity of the micro mirrors. Different reactive compounds can be distinguished from


730

Table 15. The Mars Oxidant Experiment (MOx) key features

Table 17. Specifications of the RTG and RHU used in the Small

Station Measurement Dim. (electr.)

Dim. (sensor) Mass Oper. power

Reactivity of soil and atmosphere CPU 70 mm x 60 mm x 32 mm, Petal box A130 mm x 120 mm x 18 mm, Battery 94 mm x 90 mm x 19 mm 52mmx70mmx33mm 1000 g 300 mW

each other due to the fact that each micro specific coating, when exposed to a reactive has its own characteristic response pattern. bers of the instrument are shown in Table

Heat power Mass Dimensions (cylinders} Ionisation rate at 1 m RTG electric power RTG output voltage

mirror with a environment, The key num15.

5.9. Ke_v System devices 5.9.1. The system devices and the instrument interfaces. The main system devices of the Small Station are the Central Electronics Unit (CEU), the primary power system and the Radio system. All the key system sections of the Station including reference to the responsible organisations and tasks of each section are presented in Table 16. The instruments are controlled by the CEU including a 80C31-based micro controller system and power conditioning electronics. CEU activates and deactivates the instruments according to a predetermined cyclogram. It also takes care of all the system operations of the Station, e.g. cruise and descent phase actions, data management and encoding, energy management, etc. Most of the interfaces between instruments and the CEU have been implemented by a serial line protocol (UART/19.2 kbits s-‘). The CEU controls directly only individual sensor devices, like optical, wind, pressure and humidity sensors, and their data is read though CEU’s parallel data bus. The connections between the instruments are illustrated by the block diagram in Fig. 6. 5.9.2. Radio nuclide units-sources for primary power and heat. The Small Station uses radio nuclide thermoelectric generators (RTG) as the primary source for electrical power, and radio nuclide heater units (RHU) to maintain the payload temperature within the specification (& 55’C). Both the RTGs and the RHUs use plutonium-

8.5 w RHU: 185 g RTG : 490 g RHU:40x60mm RTG : 85 x 120 mm 0.13 mR h-’ 220* 10 mW 15kO.2 v

238 dioxide as the radioactive power source. The key features can be seen in Table 17. Each of the two Small Stations has two RHUs and two RTGs. Accordingly, the Small Stations carry eight radioactive units, altogether. The total mass of plutonium238 carried by the two Stations is 120 g, corresponding to the radioactivity of 2080 curie. The development and manufacturing of the RHU were done in full accordance with the principles set for the use of nuclear energy sources in space approved by the UN General Assembly in resolution 47/68 of December 14, 1992, and with Russian documents regulating work with radioactive materials. The structure of the radioactive units is designed to provide radioactive reliability during the cruise phase, and in emergencies that might occur in launch or in case of reentry. The Plutonium dioxide of the radioactive units is in the form of cermet enclosed in an ampoule casing and covered by a multilayer thermal insulation. The outer structure of the unit is extremely robust with thermal resistive materials. The special ampoule with plutonium dioxide-238 has a double capsule design. The inner capsule that contains plutonium dioxide-238 is made of platinum-rhodium alloys, that are highly corrosion-proof, the outer (secure) cover is made of highly reliable tantalum-tungsten alloy. Besides, the inner ampoule has a device for scouring of radiogenic helium that comes out of radioactive decay of plutonium-238. The tightness of the ampoule with plutonium-238 dioxide was confirmed by ground tests using natural and simulative RHU models under the following emergency situations :

Table 16. The key system devices of the Small Station

Device

Responsible Institute

Objectives/Notes

CEU (Central Electronics Unit)

FM1

Transmitter/Receiver MBR (data relay at Orbiter) Primary power source

OKB/MEl CNES Babakin center

Controls all operations of Station. Data management Energy control, descopes activity in case of power shortage Includes power conditioning electronics for the payload VHF device (401/437 MHz), quadrupolar antenna Relays data to an Orbiter data storage Provides electrical power (nominally 440 mW) by using a Radioactive

Rechargeable battery Pyrotechniques

MP1 Babakin

Thermal Generator (RTG) Provides a power buffer Implements the commands to open parachute,

center

cut rope etc.


731

Fig. 6. The block diagram of the Small Station payload, and electrical interfaces. The main Station activation switch controlled by the Orbiter is inside the PCU

l

of the PROof the flame temperature from 3600 K to 400 K in 4000 s ; Spacecraft re-entry into the Earth’s atmosphere ; Explosion

and burning

of the components

TON rocket fuel system with the change l

Shock of RHU on rocks with the falling rate of up to 80 m s-’ ; l Kept under the influence of natural media including sea water in the depths of up to 10 km and also under the conditions of self-heating being partially or fully buried into the soil.

l

The ionisation rate of the RHUs and RTGs at a distance of 1 m equals 0.13 mR h-‘. Thus radioactive protection is not required, when the admissible time of work with the radioactive sources is in agreement with the radiation safety regulations. 5.9.3. The Radiosystem for data transmission. The Radiosystem comprises receiver and transmitter operation modes. The receiver is used only for establishing the radio link to the Orbiter by way of a handshaking protocol. The receiver cannot be used for commanding the Small Station from the Earth. The transmission and listening modes of the radio system are under the control of CEU. When in the listening mode, the receiver is switched on for a period of one second at 60 s intervals. If the request signal (RC-subcarrier) embedded in the Orbiter carrier wave is detected, the Station attempts to establish a transmission link by sending a pure carrier wave first, and after that a set of synchronisation words. In case no request signal from the Orbiter is detected, the Radiosystem is switched off. The Radiosystem has a quadrupolar antenna with four whisker-type legs located in the end of the sensor boom, and extending horizontally. The metallised upper part of the station body acts as a reflector for the antenna generating its emission to the upper hemisphere. The length of the antenna legs is one quarter of the wavelength,

Table 18. Main characteristics of the Radiosystem Receiving frequency Transmitting frequency Receiver threshold sensitivity Power emitted by the antenna Data transmission rate Transmitter operation power Mass Radiosystem Antenna with feeding unit Size of the radio system

437.100 MHz

401.5275 MHz better than 152 dB > l.OW 8 kbit s-’ < 0.7 A/15 V 1200g 230 g 80mmx160mmx100mm

i.e., with the transmission frequency of 437 MHz the length of the antenna is about 17 cm. The four antenna

legs are fed with a phase shift of 90” providing a field with a right circular polarisation in the azimuth plane. The CEU provides the data to be transmitted in binary form including convolution and Reed-Solomon data protection codes, as well as the Manchester modulation. The carrier and clock frequencies are generated by the Radiosystem itself. The main features of the Radiosystem are presented in Table 18.

6. Small Station surface operations surface the Small Station carries out scientific measurements and system operations according the commands given by the Central Electronics Unit (CEU). Some instruments demanding long measurement sessions are capable of working when the CEU is sleeping, but the majority of Station instruments use the data processing power of the CEU in real-time.

On the Martian

132


In the baseline situation the CEU is sleeping, but is awakened by an external circuitry at 32 s intervals. The CEU wakes up, checks out the task list called a cyclogram, and takes measures if required. If there is nothing to do, the CEU will go back to the sleep mode. On the Martian surface the Small Station is mostly in standby mode, due to the small amount of power available (nominally 440 mW). Only the vital sections of the CEU and some instruments consume small amounts of power. During the standby mode about 95% of the primary power will be available for charging the battery. The Small Station performs scientific measurements, processes the acquired data, and at regular intervals makes self-diagnosis on its health status. The resulting data would have been transmitted to the Mars-96 Orbiter, which would have relayed the data to the Earth. The data would have been received by three Russian ground stations and the DSN (NASA’s Deep Space Network). Originally, it was planned to use Mars Observer (MO) as a back-up data link for Mars-96 Small Stations. Due to the unfortunate loss of MO, the Mars Global Surveyor (MGS) was to be used instead to relay data to the Earth. Both the MGS and the Mars-96 Orbiter have similar data relays for communication with the Small Stations. When the Station wants to send data to the Earth it starts listening to the Orbiter’s request signal. When the signal is clear enough, the Station starts to transmit data at 8 kbits s-’ in packets with the length of 15 s. The Station listens to the Orbiter by switching on the receiver for 1 s at 60 s intervals. The Orbiter is commanding the Station by means of two subcarrier signals, a request signal RC and a TC-signal indicating that the quality of link is good enough for transmission. When the Station detects the RC-signal it starts to create the link by sending at first pure carrier and then synchronisation words. The Orbiter responds with a TC-signal when link is good, and the bit synchronisation has been ensured. The transmitted data will be encoded by convolution and Reed-Solomon data protection schemes to increase the reliability of the link. Information can be sent only from the Station back to the Earth through the Orbiters, there is no possibility to command the Station from the Earth. The Small Station may have had the first chance to send data to the Mars-96 Orbiter soon after landing, when the Orbiter is passing its first perihelion. This opportunity is a very short one, only a few minutes, and the probability of gaining contact is rather low. The next chance will be about one to three weeks later after the Mars-96 Orbiter has completed the orbit manoeuvring. From that moment on there will be an opportunity for transmission to the Mars-96 Orbiter every 3 to 7 days. The MGS would have been available for communication for the first time during the eight days after the Small Station has landed on Mars. After his period, the MGS starts the aerobraking phase taking a few months, starting on September 20, 1997. Eventually. when the MGS will adopt its working orbit around Mars, it will be above the Stations twice a day. The length of the communication links vary from 5 to 7 min. Accordingly, within some months after the Stations have landed on Mars, there may be two to three transmission windows per day. When communication with the Orbiters becomes avail-

able, the Station concentrates on sending at first all the data that was acquired during the descent phase. The combined data transmission opportunity provided by the Mars-96 Orbiter and the MGS is shared by the two Small Stations and two Penetrators. For the Orbiter request signal, RC, there are three subcarriers available. To avoid data jamming, the two Stations and the Penetrator in the vicinity of the Stations are allocated their own subcarrier. The subcarrier allocated to the Penetrator near to the two Stations is shared by the other Penetrator. Data jamming will not occur, because the Penetrator locations are separated from each other by approximately 90 latitudinal degrees. The Orbiters will request (RC signal) the landing elements to transmit data according to the commands sent from the Earth. The Station operations are controlled by means of daily cyclograms that contain all the actions and operations. There are altogether eight different cyclograms. Every day the CEU consults the table of cyclograms to decide which cyclogram to run. This way the measurements and instrument operations can be conveniently governed. Designing the Station measurement activities, sleeping periods and data transmission sessions requires trading off between scientific requirements and actual resources. Very seldom the design constraints, e.g. available power (440 mW for the whole Station), windows of opportunity to transmit data, and the amount of vacated data memory, support each other. As a result of the trade-off exercise, the Station cyclograms have been planned such that the scientific requirements have been taken into account as well as possible consuming only about 80f 10% of the nominal power. In case the Station battery voltage is too low, indicating shortage of power, the Station will start a special survival mode. In this mode the Station mainly sleeps and charges the battery. Normal operations will be resumed when the battery’s charge state is high again. Thus. the Station operates deterministically, with all the actions being predetermined. The survival mode is naturally an exception. It would have been easy to implement the Station with more onboard intelligence, but this would have tremendously increased the amount of testing and still would have caused a new reliability problem. Hence, to create a reliable Small Station the predeterministic way to govern the Station operations was chosen to be the most appropriate.

7. Small Station data retrieval concept 7.1. Data production The station will produce data during three functionally different periods : l

l

l

The check-out activity just before separation from the Orbiter ; Cruise and the descent phase into the Martian atmosphere ; Operations on the Martian surface.

The check-out performed prior to separation is a kind of diagnosis of the general health of the Station. Even if there

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of Mars

is no possibility for any corrective actions in case some malfunctioning is discovered, the check-out serves as a tool to understand the data. This is especially necessary, if any data anomalies are found. The amount of check-out data totals to 561 bytes, including instrument tests and general Station housekeeping data. The cruise and descent phase takes 4 to 5 days depending on the arrival window. All the data during that period is produced in the descent phase and marginally also during the polling of the atmosphere. The main source of data are the DESCAM images totalling to 16 MBytes and stored in the video memory of the camera system. Meteosensors, accelerometers and the Station housekeeping functions generate some 17 kBytes of data. Together with the check-out data still resident at the CEU, this leaves about 14 kBytes of the CEU memory vacated in the time of landing on Martian surface. The descent (and cruise) phase data is shown in Table 19. After landing on the Martian surface, the payload will produce data continuously starting from the first day. PANCAM. MOx, APX and Optimism instruments gather data in their own internal memories. The data of the meteosensors, optical sensor and the general housekeeping information of the Station will be stored in the CEU, generating about 7.5 kBytes per day (see Table 20). This means that the length of the available data buffer is for two day’s data production until the descent phase data has been transmitted, Hence it is possible that the data stored in the CEU memory during the very first days on Mars will be lost; only the data of the two latest days prior to the transmission session will be retrieved. Panoramic images and data from Optimism, MOx and APX will be retrieved starting from the first day on Mars.

Table 19. The data generated

during the descent phase

Descent phase data (SDPU)

No. of blocks

PTU Time ACC Pyro ACC

HK of atmospheric FIFO data HK data offset data

Time of heatshield Time of parachute

entry

separation opening

TP-measurements/once ACC-measurements/once T-HK-measurements/once

per s per s per 30 s

Time of airbag opening time of landing

Bytes/ block

1

2515

1 1

I

3 3

1509 16 13

Bytes 2515 7 1509 48 39

1

I

1

7

I I

600 180 20

18 13 18

10,800 2340 360

1 1

7 7

7 7

Total data at SDPU (kBytes) Act operation time before parachute opening = 180 s. Descent time after parachute opened = 600 s. Separation check-out data at SDPU = 0.55 kBytes. DESCAM images during descent = 2,000 kBytes. The overall descent phase data = 16.02 Mbits.

17.29

733

Table 20. The data generated surface

Data/sol

during the first IO days on Martian

during the first 10 sols

PTU PTU HK ODS TW Cyclogram HK data SDPU HK Battery HK DC/DC HK Current HK Total SDPU data (KBytes)

No. of blocks

Bytes/ block

Bytes

88 31 106 7

37 515 I 31 24

3256 515 217 3286 168

1 I

22 21

22

1 1

21 58

I

21 21 58 1.39

PANCAM, Optimism, MOX and APX gather data in their own memories (altogether z 1 Mbit/day). Day and night have been assumed to be of equal length.

7.2. Data transmission principle The Small Station attempts to transmit its data according to the internal cyclogram, provided that the battery charge state is sufficient. The Station starts the telecommunications session by sending synchronisation words to lock in to a transponder onboard an orbiter. If no orbiter is available, the Station continues the attempts to lock-in with predetermined schedule until the transmission has succeeded. When transmitting data the Small Station is not dependent on the knowledge about what is the type and number of orbiters, or transponders, within the reach of the Small Station transmission system. The Station is operating on its own according to its internal schedule. In the Mars-96 mission, the Mars-96 Orbiter and the MGS were planned to be used as data relays. The relay onboard the Mars-96 Orbiter was considered to be the prime equipment for the gathering of Small Station data, with MGS as a back-up. This strategy was modified, however, due to the long time (that could have been more than three weeks) needed to establish the first communication session with the Mars96 Orbiter, and the long interval (3 to 7 days) between visibility sessions of the Mars-96 orbit. The decision was made to use the MGS in the beginning of the surface mission to locate the Small Stations and to retrieve their data. The MGS could have been available during the first four 48 h orbits, with 3 to 10 min visibility sessions per orbit. The Mars-96 Orbiter would have taken the primary role as soon as sufficient amount of data of locations and orientation of the landing elements would have been obtained. The relay antenna onboard the Mars-96 Orbiter was a steerable directive antenna and the MGS had a quasiomnidirectional antenna. When taking into account the dispersion of the Mars-96 Orbit injection, the dispersion of the landing sites and the inhomogeneous beam of the Small Station’s transmission system, the probability to point the Mars-96 Orbiter antenna and to establish the first communication session was relatively small. This was the main rationale behind using the MGS as the prime

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communication relay in the beginning of the landed mission, thus taking benefit of the quasi-omnidirectional antenna of the MGS. After this decision, the MGS team accepted to make also an additional manoeuvre upon arrival to Mars, to move the periapsis of the initial orbit (a 48 h period) into the vicinity of the landing area of the Small Stations. This would have increased the safe transmission probability for the descent phase data. Thereafter the MGS would have communicated with the two Small Stations with the alternating mode, i.e. allocating a 15 s transmission window for one Station after the other, thus allowing a transmission of the descent phase data, even with the narrow window (3 to 10 min) of visibility per orbit. With the first part of the descent phase data successfully received, the health status of the Small Station can be determined, and a rough localisation of the Station can be established by analysing the Doppler shift of the transmission frequency. This information would have enabled the prediction of the time periods, when the Mars-96 Orbiter was visible to the Small Stations. Hence the Mars96 Orbiter antenna could have been directed to properly facilitate the fast retrieval of the rest of the descent phase data comprising mainly images. This procedure would have taken benefit of all the remaining capacity of the lithium batteries. During the first eight days of the Small Stations on the Martian surface, some of the descent phase data have been transmitted by means of the MGS. After this phase the main responsibility of receiving data from the Small Stations was to be with the Mars-96 Orbiter being planned to be available within three weeks after the landing of the Small Stations. After a few months of aerobraking the MGS would have been available again to provide one to two transmission opportunities per day. The capability to command the transponders onboard the MGS and the Mars-96 Orbiter from the Earth would have enabled optimisation of the data retrieval from the Stations by taking into account the Small Station cyclogram, and electrical energy status. Accordingly, the Small Station is behaving like a standalone autonomous spacecraft capable of sharing its transmission bandwidth with any number of communication systems. Hence the Small Station communication protocol provides for robust and reliable data retrieval scenarios. 7.2.1. Dataflow andprotection scheme. The data integrity was protected by using the recommended CCSDS coding with convolutional and Reed-Solomon error and detection codes. This type of error correction coding was important especially in the case, when the receiving end was the Mars-96 Orbiter planned to be used with the maximum range (5000 km) of the Small Station transmitter. Furthermore, error detection and correction algorithms were made necessary by the fact that most of the Small Station data are compressed. The protection requirement was especially important for the descent phase images, which are tightly compressed. In the receiving end the data integrity was to be guaranteed by the data relay by checking that the Viterbi decoder was locked to the transmitted signal. Thereafter the data relay was to send a data transmission authorisation to the landers by means of an additional sub-carrier.

The data received by the relay onboard one of the two Orbiters would be time-tagged and stored in the on-board memories with the inner protection code (Reed-Solomon). Eventually, the data stream was transmitted to the Earth by the Mars orbiting spacecraft and received by the Russian and US Earth-based control stations. Finally, the Small Station data recorded by the Orbiters and sent to the Earth, was planned to be reconstituted by means of the ground support equipment of the data relay and the Small Station. This procedure would generate multiple data files, one for each instrument or subsystem of the Small Station. Thereafter the data obtained at the Martian surface would be ready for analysis by the experiment teams.

8. Determining the location and orientation of the Small Station The orientation of the Small Station relative to the gravity vector can be concluded from the inclinometer readings. There is an inclinometer inside the Station body, actually as a part of the Optimism instrument. The vertical inclination can be determined with the accuracy of better than 0.5”, after averaging and processing of the data an accuracy of approximately 0.1” is achieved. The horizontal orientation will be determined by means of the data from the optical sensor, panoramic camera, and by means of an inclinometer and sun sensor onboard the magnetometer head. The horizontal orientation could be determined with the accuracy of better than 3”, based on the fact that the orientation of the magnetometer head was to be known with the accuracy of one degree. A rough determination of the landing sites of the Small Stations would have been performed by ballistic means giving a size of 320 km x 2200 km for the landing ellipse. A more detailed localisation of the Small Station was to be done by studying the Doppler shift of the carrier signal during data transmission sessions, the same type of procedure was used for the Viking Landers (Mayo et al., 1977). Data sent by the two Small Stations to the Mars96 Orbiter and the MGS was to be analysed for detecting the level of the Doppler shift. Statistical averaging would have been gained over time, when the number of telecommunications sessions would have been increased. The data streams transmitted by the Small Stations can be distinguished from each other due to the fact that both the Small Stations have their own subcarrier frequencies used in establishing the telecommunications link. The receiving end on the Martian orbit includes the knowledge of the subcarrier frequencies in the data stream. Hence the information of the data source is preserved. The localisation accuracy by using the Doppler method is dictated by the frequency stability of the Small Station transmitters and the receivers onboard the orbiters. Even with a modest transmitter stability (e.g. 10 ppm over 1 min for Small Station), the localisation with a Doppler method certainly improves the knowledge about the locations of the Small Stations by a factor of 100, compared with the first estimate given by ballistics. When the locations of the landing elements have been determined,

V. Linkin el

ol.

:

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they can be used like a network of beacons to find out in more detail the orbital parameters of the Orbiter.

9. Design requirements and drivers 9.1. Gene&

design driwrs

The main Small Station design drivers were the requirement to fit into the small allocated volume, to stay within the mass limit, to survive with a negligible amount of power, and not to exceed the data bandwidth. Finally, the Station should still satisfy the scientific requirements. Obviously the eventual design is a compromise among all of these constraints. The main Small Station resources are electrical power (nominally 440 mW), datalink capacity (1 to 5 Mbits/ week) and payload mass (12 kg). Every instrument team is naturally of the opinion that their instrument is highly important and their resource requirements should have the highest priority. This will inevitably lead to the fact that when combining the requirements of all instruments the available Station resources will be exceeded. An additional complication is due to the design constraints being intertwined. This is reflected by the fact that changing the requirements on one resource will have an effect on other resource requirements. Optimising the use of resources among the requirements of the instrument teams, and the actual resources available, was one of the major challenges in designing the Small Station. During the Small Station development phase all the scientific requirements were taken into account within the resources allocated to the Station. The Station operations are planned such that about 80% of the nominal power is consumed. The final version of the operation instructions is implemented by modifying the specific parameter lists of the operating software, enabling easy development and test work. All the design was made by using experienced personnel, and every manufactured system and subsystem was thoroughly tested. Product assurance was carried out by an independent organisation. The original plan included three Small Stations, which. besides being dictated by the network science, also served the purpose of redundancy. The idea of having a sort of redundant third Station is reflected by the requirement stating that internal doubling of the Station subsytems should be avoided to save mass and power. Finally, however, the third lander was deleted and the gain in mass was used by new instruments. The design of the Small Station was carried out keeping in mind the nominal lifetime of one terrestrial year and very harsh operating environment. The environmental conditions for flight units are largely as follows : l l l l l

l

l

Tests and decontamination before launch ; Mechanical stresses during launch ; Space flight of 10 months ; < 10 krad radiation dose for the electrical components ; Impact shocks of the order of 200 g /(30-70) ms when landing on Mars ; Martian night-time temperatures may reach as low as 150 K; and Daily temperature variation of 50 to 100 K on Mars.

of Mars

73s

In order to increase the quality of the final product, five different test models of the Station were manufactured and tested before fabricating the flight units. The test models were used to verify the design by qualification tests and also used for prototyping the electrical, thermal, and mechanical features of the subsystems. During the development phase it was decided not to implement a Station capable of making its own decisions on various matters. Rather the Station was designed such that all the operations will be largely predetermined before launch. Only in case of power loss will a special survival mode be chosen by the Station itself. Including a lot of internal intelligence in the Station would have increased the amount of testing inordinately and, at the same time. would have diminished the overall reliability. All of the Station development and tests were aimed at creating a reliable Station that would produce a maximal amount of new scientific information within the allocated resources. In theend, this philosophy resulted in the Mars96 Small Stations. 9.1.1. Decontaminntion requirc~n~o~ts. Since the beginning of the Solar System exploration. it has been required, according to the article IX of the OUTER SPACE TREATY (London/Washington, January 27, 1967) to preserve planets and Earth from biological contamination. Consequently. COSPAR (Committee of Space Research) has established own protection its planetary recommendations (Debus et ul., 1997). They encompass preventing the planetary environment from biological contamination by terrestrial micro-organisms, and to protect the Earth from contamination by sample return missions. Protection of the planetary environment facilitates also exobiological investigations of the planets by minimising the possibility that an exobiological finding would have originated at the Earth. According to the COSPAR recommendations, specifications have been written for the Mars-96 mission in order to realise biological cleanliness (Debus et ul.. 1997). All exposed surfaces of small stations and Penetrators entering into Mars atmosphere were decontaminated under a level of 300 spores per square meter. The following sterilisation procedures were carried out : (a) A11 equipment, subsystems, experiments have been biocleaned with sporicides or sterilised. The required sterilisation methods were Hydrogen Peroxide Plasma Gas (Johnson and Johnson Medical Sterrdd 100 procedure), gamma radiation, ultraviolet exposure and dry heat. After decontamination. all equipment was sent in NPO Lavochkin in Moscow in order to be integrated on the probes. A specitic sterile packaging system has been implemented for experiments in order to conduct tests without recontamination. (b) Integration of small stations and Penetrators have been done in sterile class 100 laminar flow clean rooms. The biological cleanliness was ensured using specific access procedures, sterile clothes for operators, biocleanings and ultraviolet exposures. The control was made by way of microbiological assessments and eventually, the planetary protection requirement was satisfied. For example. the level of spores on each Small Station was 250 and 270. respectively.

736


(c) Finally, after integration, the Small Stations and Penetrators have been covered with a bioshield in order to prevent any recontamination. (d) The orbiter instrumentation was not sterilised. The general philosophy about the landers’ integration was to clean all equipment during their construction (biocleaning), or after integration (sterilisation). Equipment, instruments and subsystems were sterilised, or biocleaned to a level of less than 300 spores per square meter. The required sterilisation methods were dry heat, gamma or beta rays and hydrogen peroxide gas plasma. Biocleaning was performed during construction using clean installations, alcohol, UV and sporicides cleaning with bioburden level verification. A special packaging system including two nested plastic bags was used in order to conduct acceptance tests of the instruments without impacting the biological cleanliness.

10. Discussion and aspects of future Mars landers During the past decade Mars has been the focal point of planetary exploration. Mars missions will also have a central role for years to come. The chain of Martian exploration projects was resumed (after Viking) in 1988, when the Phobos spacecraft was launched, and continued by the launch of Mars Observer in 1992. The new members in the series are Mars Global Surveyor (MGS) and Mars Pathfinder. Unfortunately, the Mars-96 mission did not have the possibility to support the Martian exploration in the form of scientific return. However, the Mars-96 mission contributed a lot in the form of technological development, scientific planning and international co-operation for space research. The state of the art concepts of the Small Station and the Penetrator proved that it is possible to deploy on the Martian surface a lander with an ambitious and versatile set of scientific instruments. In addition to that, the Small Station also proved that an ambitious scientific lander mission to Mars can be implemented with a considerably lower amount of mass than has earlier been the case. New surface missions to Mars are being planned. NASA has the Mars exploration program starting in 1996 and extending to the year 2005, including two launches for Mars every 26 months. Russia has plans for its own Mars-1998 mission, ESA and NASDA (Japan) have their Mars exploration programs. The new Mars landers are very likely to carry small rovers onboard. NASA’s Pathfinder already has one. A combination of a lander and a small rover working in the neighbourhood of the lander would be very practical. The concept of the rover using the lander as a service centre for energy, computation power, transmission capability etc., is a highly attractive scenario. When looking even further into the future, it is possible to envisage a concept of multiple small rovers and crawlers carrying out scientific measurements, and using some stationary vehicle as a source of energy and data bandwidth. They would also collect samples in sealed containers and place the containers

in some specific locations,

where the samples would be retrieved by a later sample return mission. The chain of Martian exploration will receive additional links in years to come. Inspiring technical inventions will be made, and Mars will reveal new secrets to the scientific community. The Small Station developed for the Mars96 could be used as such as a part of a forthcoming Mars mission. Using a proven and existing design would considerably decrease the cost of a lander mission. The concept of the Small Station will definitely be a seed of many future Mars landers. Additional note:

This paper was written before Pathfinder arrived at Mars and therefore none of the Pathfinder results have been taken into account in the discussions herein. Acknowledgements.

The authors wish to thank all members of the international Mars-96 Small Station partners for participating in the project, for constructive team work and for fruitful cooperation over the years. Despite the failure of the Mars-96 launch, a new lander concept. applicable to future missions, was created.

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