Conceptual Design of a 1979 Mars Rover

1971 (8th) Vol. 1 Technology Today And The Space Congress® Proceedings Tomorrow

Apr 1st, 8:00 AM

Jesse W. Moore Supervisor, Advanced Navigation Group, Jet Propulsion Laboratory, Pasadena, California

Mel Swerdling Member of Technical Staff, Jet Propulsion Laboratory, Pasadena, California

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Scholarly Commons Citation Moore, Jesse W. and Swerdling, Mel, "Conceptual Design of a 1979 Mars Rover" (1971). The Space Congress® Proceedings. 3. https://commons.erau.edu/space-congress-proceedings/proceedings-1971-8th/session-1/3

This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. CONCEPTUAL DESIGN OF A 1979 MARS ROVER

Jesse W. Moore Mel Swerdling Supervisor, Advanced Navigation Group Member of Technical Staff Jet Propulsion Laboratory Jet Propulsion Laboratory Pasadena, California Pasadena, California To Be Presented At

Eight Space Congress April 19-23, 1971 Cocoa Beach, Florida ABSTRACT

The results of a conceptual design study of a Mars Given the premise that long Martian surface roving vehicle mission in 1979 are presented. distances are to be explored, the automated sur Descriptions of the mission, science objectives, face roving vehicle appears to be a prime means vehicle configuration and subsystems are included. for achieving these exploration goals. Mission analysis parameters required to define a mission profile and sequence of events are pre To focus and bound this conceptual design study, sented. Science operations including the deploy several assumptions were defined. These limited ment of small, self-contained, long-lived meteo the depth of effort in certain areas and eliminated rology and/or seismology stations are considered. some of the tradeoffs that otherwise could have The vehicle system is described by the functional been performed. Throughout the study these requirements, vehicle configuration and weight and assumptions were adhered to, and interpretation of power allocations. Following the system descrip the results should include the effect of these study tion, seven subsystems on board the vehicle are assumptions. considered. The characteristics and capabilities of each are described. Mission operations also The key assumptions made were: were evaluated to the degree necessary to identify the areas of foremost concern. (a) Six-wheeled vehicle concept The six-wheeled vehicle concept has been INTRODUCTION shown by several previous studies(l» 2, 3) to provide the "best" mobility capability and the In future decades, the space science community "best" overall design for operation on unknown will be provided with unique and challenging oppor surfaces. tunities to explore the solar system by spacecraft of increased complexity, longer lifetime and more (b) Direct-link communications with earth autonomy. A candidate mission is an automated traverse on Mars. Roving vehicles capable of A relay communications link with an orbiting extended operation on the surface, and sufficiently vehicle was not considered, although relay mobile to traverse regions remote from acceptable communications may be practical and desir landing sites, would offer scientists the opportunity able in some instances. to explore large regions of the planetary surface. (c) One routine science operation performed per The value of landed spacecraft in performing scien day tific explorations and surveying future landing sites for manned vehicles was demonstrated by the Some science data should be collected during Surveyor lunar missions. Surveyor, a "fixed- each day of the mission. This includes pos point" surface spacecraft, collected composition sibly imaging, meteorology, etc. , but does and topographic data within local areas about the not always include operation of the life- landing sites. In many cases, however, areas of detection experiments. scientific interest are likely to be in regions remote from acceptable spacecraft landing sites, since (d) No locomotion during the Martian night acceptable landing sites are chosen from topo as scientific criteria. Further graphic as well with the earth can be many scientifically interesting areas are Since no communications more, a direct link during the Martian to exist on the Martian surface and explora effected with likely the risk to the safety of the vehicle tion of all of these areas by immobile surface night, quite high. spacecraft is clearly impractical. becomes

This paper presents the results of one phase of research carried out at the Jet Propulsion Laboratory, ;Galifornia Institute of Technology, under Contract No. NAS 7-100, sponsored by the National Aeronautics ,and Space Administration.

1-29 rover stops at a science site, a large fraction of (e) No science during locomotion the power previously used to propel the vehicle becomes available for science. This power level All scientific observations were assumed to is more than adequate to satisfy the needs of the take place while the vehicle is stationary. subsystem. This lowers the power consumption during science motion. MISSION REQUIREMENTS AND ANALYSIS (f) No TV during locomotion The mission requirements and analysis is con This implies that the TV subsystem will not be cerned mainly with parameters of the vehicle's used for near-real time driving of the vehicle. surface operation; e. g. , round-trip light time, A picture might be taken during motion, but earth elevation, and sun elevation. No effort was its transmissions to earth would be delayed expended on the launch, cruise, entry and landing until the vehicle is stopped. phases of the mission. Launch and arrival dates were determined and the rover was assumed to Other assumptions made during the study are land in a 0. 52 rad. latitude belt, centered around stated in this report as appropriate. the equator. In addition, no effort was expended on specific scientific mission design. Consider able effort is required to determine a traverse SCIENCE OBJECTIVES route, including selection of the scientific sites to be investigated. This selection will significantly From a scientific point of view, the rover under influence the design of the vehicle system (partic study may be considered as a lander with the flex ularly affected will be mobility, lifetime require ibility to investigate a large number of different and operational strategies). sites. The science objectives are, therefore, ments, essentially identical to those of a lander with the The 1979 launch opportunity was considered, with recognition that the rover has the additional capa arrival in 1980, and a vehicle operational lifetime bility not only of studying many distinct sites, but approximately one year. The minimum C3 also of escaping from the area altered by the land of trajectory for this opportunity will be launched on ing maneuver and of deploying small, independent 3, 1979» arrive at Mars on August 5, science packages for specialized investigations at November 1980 and require a GS of 8. 955 km2 /sec 2 . Allow one or more locations along the traverse route. for a reasonable spread of launch and arrival The rover could deploy independent meteorology ing launch could occur between late October and seismic stations which would make possible the dates, and 1979, and arrival at Mars could be acquisition of simultaneous data from different mid-November in August 1980. Use of higher C$ to locations, thus greatly enhancing the value of such anytime communication distance would somewhat investigations. shorten modify these dates. A listing of the major science objectives are given of the most significant mission analysis param below: One eters is the light time-delay between earth and parameter influences the amount of Search for evidence of living organisms over Mars. This (a) autonomy desired for the on-board a large surface area. operational vehicle system. With large time delays, fly-by- is not practical. Instead of travelling (b) Visually characterize many scientific sites. wire control long distances, the mission lifetime is spent for part on transmitting and receiving com (c) Search for and characterize organic com the most an arrival date of August 5, 1980 and pounds over different types of surfaces. mands. For for a lifetime of 1 yr, the roundtrip light time earth and Mars is shown in Figure 1. At (d) Determine atmospheric composition and its between round-trip light time is approximately temporal, spatial and altitude variations. arrival, the 28 min, as seen from the figure, increasing to a min and then decreasing slowly begin (e) Determine meteorological characteristics. peak of 41 ning in May 1981. (f) Determine seismological characteristics. is The Mars-earth communication visibility period the amount of auton the above objectives, the basic science another parameter influencing To achieve the vehicle. Table 1 assumed to consist of imaging, biology, omous control on-board payload was Martian day that the meteorology and seismometry shows the number of hours per molecular analysis, various latitudes on It was further assumed that the total earth would be visible from experiments. dates of interest. The alti available for rover science was approxi Mars, and the various weight angle of the earth is also given. mately 46 Kg. tude (elevation) For reliable communications with the vehicle, the be at least 0. 26 rad. above the local In an earlier assumption it was stated that no sci earth must is ence experiments are performed when the rover horizon. in motion; therefore only a small amount of power To define the functional requirements of a Mars for standby and for maintaining some previously surface roving vehicle, it was necessary to estab acquired sample in proper incubation is required lish a roving strategy. It was assumed that the by science during the moving phase. When the

1-30 Table 1. Number of Hours Per Martian Day That Earth is Visible for Various Dates and Latitudes

JULIAN DATE, LATITUDE ON VISIBILITY HRS MAX ALTITUDE OF CALENDAR DATE days MARS, radian OF EARTH EARTH, radian

2444460 -0.52 8.96 0.63 0 10.93 1. 15 AUG 9 +0.52 12.90 1.44

2444520 -0.52 10.64 0. 84 0 12.93 1.36 OCT 8 +0.52 13.97 1.25

2444580 -0. 52 13. 12 1. 17 0 13.68 1.45 DEC 7, 1980 +0.52 10.25 0.92

2444640 -0.52 15.00 1.41 0 12.93 1. 18 FEE 5 +0.52 10.86 0.65

2444700 -0.52 13.68 1.41 0 11.68 1. 15 APR 6 +0.52 9.69 0.63

2444760 -0.52 12.88 1.22 0 11.68 1.38 JUN 5 +0.52 12.48 0.86

2444820 -0.52 11.52 0.92 0 12.93 1.44 AUG 4, 1981 +0. 52 12.34 1. 18 intent of the rover was to move to an appropriate Assuming the individual traverse segments will be science site, perform prescribed scientific analy lengthened as a function of time, a range capability sis of the area, transmit the appropriate findings of the roving vehicle was estimated for a 1-yr mis to earth, and repeat the foregoing sequence for the sion. This range capability is shown plotted in duration of the mission. Figure 3. To arrive at this plot, several additional assumptions are necessary. First, it was assumed August 5, 1980 and a -0.52 rad latitude were chosen that 9 major scientific sites would be investigated, to graphically describe the profile of a single mis each requiring up to 15-20 days for a complete sion. The resulting single Martian day profile is investigation. Second, the number of kilometers shown in Figure 2. A reference time of zero is per day was increased from 0. 3 to 4. 0 in incre selected when the earth rises in the Martian sky on ments. On later days the vehicle would have to the above date. Vehicle motion is confined to the travel during a large part of the visibility period time when the earth is 0. 26 rad. above the horizon, and, in addition, the frequency of obstacles would and on this date the time is confined to approxi have to be small. With these assumptions, it is mately 8-9 hr of operations. The cross-hatched possible to achieve a range of 400-500 km during a areas indicate periods when data are transmitted 1-yr surface mission. to earth and commands from earth are received and verified. The solid areas represent 100-m traverse segments. For this day, three 100-m VEHICLE SYSTEM DESCRIPTION segments are possible. Beginning at sunset, the night science operations (meteorology, etc. ) and The basic functional elements of a Mars surface battery recharge take place. This profile might be roving vehicle are centered around the on-board followed in the early phases of the mission; how computer and sequencer (data handling subsystem) ever, as time increases the traverse segments are which performs the necessary computations, makes likely to be lengthened to a kilometer or several the decisions and executes the events in their kilometers. proper sequence. The primary inputs to the

1-31 series of experiments and transmit the results to computer are from the science, navigation, receiving station. This intermittent roving and mobility and obstacle detection functional elements. the or experiment sequence, is then repeated. The communication element, which includes the decision, demands on the control system and of a directional antenna to earth, serves To limit the pointing operators, internal hazard-sensing to telemeter data from the vehicle and receive earth-based (e.g. tilt obstacle detectors) are provided. commands from the earth. Power is required for devices all elements, and thermal and environmental con roving vehicle weight estimate summary trol are needed to reduce the effect of variations The Mars in Table 2. Twelve subsystems are iden in the Martian environment. is given tified. The total weight of the vehicle, including kg contingency, is 512 kg. Mobility and power In order to make the exploratory design investiga a 46 require the largest weight fractions. tions as meaningful as possible, the initial config subsystems requirement estimates are provided for the uration study considered a reference or baseline Power during locomotion, i. e. , when the Vehicle which: worst case vehicle is in motion over an average +0. 09 rad. This will probably occur for a (a) Is generally compatible with the Viking delivery terrain slope. small percentage of time. The mobility and landing systems. (This imposes major relatively provides for a vehicle speed of weight and C.G. limitations, as well as some power estimate on the slope. The power esti critical geometric constraints. ) 0. 35 km/hr while mates are shown in Table 3. A total of 283 W, of W are required for mobility, are (b) Provides sufficient mobility to a meaningful which 124 complement of science and communications to required for locomotion. perform a major surface exploration mission.

The exploratory configuration developed for this VEHICLE SUBSYSTEMS adapts the major transport and descent vehicle subsystems involved with vehicle systems of the Viking in order to soft land a six- The on-board considered in this study are (a) mobility, wheeled roving vehicle onto the Martian surface. motion and The entry and landing sequence is shown in Fig ure 4. The configuration uses the Viking descent 2. Weight Estimate Summary capsule, the parachute system, and the soft-landing Table terminal propulsion system in combination with the radar systems for the descent requirements. How SUBSYSTEM WEIGHT, Kg ever, it repackages the three-legged stationary Viking '75 (Surveyor-type) landing spacecraft of 59 into a six-wheeled rover using a three-segment Rover Structure control compartment body. 127a Mobility developed is shown in The design configuration 46 Figure 5. The body or chassis of the rover con Science the first two of which sists of three compartments, 32 contain the science, communications and electron Communications thermally controlled environ ics equipment in a 43 ment. The rear compartment supports two multi- Data Handling hundred watt RTG's (Radioisotope Thermoelectric nob Generator). Each compartment supports a pair of Power wheels with independent springing for landing on 9 the wheels (in the interlocked condition), and sur Navigation the extended condition). The ter face roving (in 5 minal descent propulsion modules are attached to Obstacle Detection the science compartment wheels and the hubs of 5 the aft end of the rear RTG compartment. These Antenna Pointing after landing. Descent radar anten are jettisoned 14 nas attached to the hubs of the rear wheels are Thermal Control after landing. The mast attached also jettisoned 11 to the central compartment is erected after land Cabling the high- and low-gain antennas ing. It supports 5 plus the facsimile (panoramic) camera, the local Pyrotechnics imaging and ranging sensors. Soil manipulation Total 466 Kg instruments are attached to the front compartment. 46 The roving vehicle is powered by an electric motor Contingency wheel. Center point steering is also elec in each 512 Kg trical; motion (forward or reverse) and steering is either ground-commanded or commanded on-board. survival Operation of the vehicle during roving is intermit Includes weight for landing tent in that power is first provided to travel a short distance (approximately 50-100 m) and then Includes two batteries stop. Power is then supplied to perform a new visual survey and, where appropriate, to make a Does not include sampler weight

1-32 Table 3. Locomotion Mode Power Requirements involved the evaluation of many competing factors. (Worst Case) The most important of these factors were:

(1) Mission terrain characteristics. Surface SUBSYSTEM POWER, W VOLTAGE characteristics along the routes of pertinent science objectives (roughness, slopes, hard Science 0 - ness, obstacles, crevasses).

Communications 25 AC (2) Mobility system efficiency.

Obstacle Detection 15 AC (3) Trade-off between obstacle traversing and obstacle avoidance subsystems. Navigation 10 AC (4) Mobility system reliability and redundancy. Mobility >124a DC (5) Adaptability of the concept to the delivery Controller 10 AC systems.

Data Handling 70 AC Since the accurate evaluation of these factors for the different roving concepts on Mars was beyond Thermal Control 10 AC the scope of this initial study, a gross review was made of the difference in the Martian and lunar Conversion and Net 19 ' environments as they affected these competing fac Losses tors. It was found that although there were con siderable environmental differences between Moon Totail 283 W and Mars (gravity, temperature range, atmosphere, communication distance, etc. ) none of these Provide for vehicle speed of 0. 35 km/hr while seemed of sufficient influence to appreciably operating on average 0. 09 rad. slope change the comparative merits of the various mobility concepts, as made for lunar roving studies. (b) power, (c) telecommunications, (d) navigation, (e) obstacle detection, (f) antenna pointing and In view of these results the choice of mobility con (g) data handling. Omitted from consideration cepts for the reference vehicle was narrowed to were the vehicle controller subsystem and the four- or six-wheeled vehicles as considered for the obstacle avoidance subsystem. The controller moon. Since investigations of the four-wheeled provides the commands to the drive motors for rovers indicated quite limited mobility (approxi speed and heading changes. The obstacle avoid mately half that of the six-wheeled version in ance subsystem makes use of the data from the obstacle and crevasse capability) as well as obstacle detection subsystem to provide heading reduced redundancy, the six-wheeled concept was changes to the vehicle controller. Obstacle avoid adopted for the initial reference vehicle. However, ance is a very necessary and complex subsystem it should be pointed out that a four-wheeled concept for the Mars rover and its design must be com might present advantages for missions limited to patible with the obstacle detection subsystem. comparatively smooth areas. Mobility character Avoiding obstacles while minimizing some crite istics of the referenced vehicle is given in Table 4. rion function, such as time or distance to the des tination, will require the development of path find In attempting to achieve the 0. 52-0. 61 rad. goal ing algorithms(4, 5) with significant flexibility and established for this vehicle, special effort was capability. made to reduce soil pressures as much as possible. Such design takes advantage of soil cohesion that is available on the moon and is perhaps available (a) Mobility on Mars. Soil pressure on the lunar rover, which uses 82 x 23 cm wheels, is approximately 4800 In considering the mobility approach for the Mars N/m2 . The concept presented here is expected to Roving Mission, a review was made of the various have about half to two-thirds of that soil pressure, designs investigated for the lunar roving pro which may provide for the 0. 52-0. 61 rad. slopes. gram^* 2, 3) 0 These included:

(1) Legged vehicles: single and multiple (b) Power (articulated mechanical legs). Radioisotope thermoelectric generators and her (2) Tracked vehicles: single and multiple tracks metically sealed nickel cadmium batteries are the (with single and multiple chassis). primary and secondary power sources selected for the Mars Rover application. This combination can (3) Wheeled vehicles: single and multiple wheels provide the dependability and high reliability (with single and multiple, rigid and flexible required throughout the approximate 2-yr mission. chassis/body configuration). A power profile depicting three key modes of oper As in the lunar rover investigations, the choice ation, such as mobility/locomotion, science, and a from the above concepts for the Mars requirements, typical "night time" period is shown in Table 5.

1-33 Table 4. Mobility Characteristics

CRITERIA ESTIMATES

soil 1. Maximum slope capability 0.52-0.61 rad. soft a 0. 61 rad. -wedge formed by two intersecting 2. Ground clearance (A) Straddle crater walls (B) Undercarriage clearance 0. 4 m (approx). (Within central compartment area) 3-4 m (approx) 3. Maneuverability (A) Turning radius (B) Front and rear steering

(C) Reverse drive 0.70-0.79 rad. for traversing crater walls of soft 4. Stability Approximately soil and providing for some wheel sinkage

5. Obstacle capability 0.9m (approx)

6. Crevasse capability 0. 6 to 0. 9 m. (approx)

Table 5. Mars Rover Power Profile Continuous (watts)

NIGHT OPERATION SUBSYSTEM VOLTAGE MOBILITY SCIENCE OPERATION

40 40 Computer AC 40

Centralized data 18 18 storage AC 18

Communications 0 (TWT) DC 0 55 25 25 Radio frequency AC 25 0 Navigation AC 10 0

Controller 0 (mobility) AC 10 0 12 Data handling AC 12 12 10 Thermal control AC 10 0 124a 0 Mobility (locomotion) DC 0 0 Obstacle detection AC 15 0 96b 81 Science 0 48 Battery charging DC 2 2

Conversion and 23 distribution losses 17 25 257 TOTAL: 283 273

with 2 RTGs only on an average Provides for an initial vehicle speed of 0. 36 km/hr while operating steeper local slopes. 0.09 rad. slope. Battery supplemented for greater speed or unless battery fully charged. Continuous average shown. Peaks may not be on simultaneously

1-34 The power requirements are based on what can be through the high-gain antenna, and coherently expected from technology development through received by the DSN stations. There is insufficient 1975 - 1976, as well as that which can be expected power to transmit telemetry data to earth through from the Viking '75 program. Voltages shown the low-gain antenna at 6 bits/sec (the DSN lower were assumed for conversion losses, required raw limit for coherent detection) and thus an M-ary power, and weight. Frequency Shift Keyed (MFSK) modulator is pro vided. For low-gain transmission, the data is The mobility power demand of 124 W (see Table 5) MSFK modulated onto an S-band RF carrier at is based on that available from 2 RTGs after pro about 0. 5 bits/sec and transmitted to earth through viding for supporting subsystems. The 124 W the low-gain antenna where the signal is incoher available will provide for a speed of 0. 36 km/hr on ently received and detected. It is intended that the an average 0. 09 rad. slope. A functional block MSFK mode would be used only in the event of a diagram of the Mars Rover power subsystem is failure of the pointing systems of the high-gain shown on Figure 6. antenna.

Operating the RTG on the Martian surface may For transmission of data to earth at the maximum affect performance, since this design is for the data rate, the X-band carrier would be used. To deep space environment. For this study, it was obtain sufficient pointing accuracy of the high-gain estimated that each RTG will provide approxi antenna for X-band use, a monopulse antenna mately 140 W at arrival on the surface, approxi pointing system will be required. In any event, a mately 136 W at the end of 1 yr, and 134 W at the gyro and sun sensor pointing system also would be end of mission. used for backup with the S-band mode.

A shunt regulator is provided with each RTG to An S-band receiver is provided to receive com maintain the RTG at its maximum power capability mands from earth, and to provide the reference when subsystem power demands are reduced. An for a phase coherent down link S- or X-band signal inverter is provided for changing a portion of the when required. Ranging can be performed when RTG DC power output to regulated square wave AC required. The command waveform is sent from power for engineering and science. A standby the receiver to the modulation/demodulation sub inverter controlled by a failure detector sensing system where the command subcarrier is removed the main inverter voltage and frequency outputs is and the data bits are detected. The detected data provided for redundancy. The remaining power bits are then sent to the command decoder for from the RTG is available for mobility and further operation. communications. The baseline telecommunications subsystem con Two 12 amp-hr nickel cadmium batteries are pro sisting of a 20-W X-band TWT with monopulse vided to supplement the RTGs during periods of tracking (backup S-band with gyro and sun-sensor mobility or locomotion when increased speeds are antenna pointing) was not evaluated for its capabil desired or increased slopes are encountered. The ity to provide for communication to earth during batteries are discharged in parallel through a reg launch, cruise, entry or deployment phases of the ulator for operation at the maximum power point of mission. Also, it was not evaluated for providing the RTG. A standby regulator, controlled by a a communication capability for a relay link to earth. failure detector sensing the main regulator output Evaluation of these modes of operation must be voltage, is also provided for redundancy. A single considered before final selection of a telecommuni battery may be adequate if compromises in other cations system for the Mars rover. vehicle subsystems are made (e.g. , mobility). (d) Navigation (c) Telecommunications The basic requirements of the navigation subsystem The telecommunications design associated with a are to direct the vehicle motion accurately and Mars roving vehicle mission presents some par reliably over the planned mission traverse, which ticularly difficult design tradeoffs. For this con will be designed and over-laid on the best available cept, the number of tradeoffs was limited by con photographic maps of the surface area of interest. sidering only a direct link to earth. Use of a direct General heading and location of the destination link primarily affects bit rate, antenna pointing (major scientific sites) relative to the vehicle can accuracy, and communications visibility period. be obtained from the planned traverse maps. Con With limited power on board the vehicle, the direct tinuous vehicle position and heading information link data rate is limited. Antenna pointing accura will be obtained and used in conjunction with the cies of ±21 mrad. must be achieved for reasonable traverse map information (on earth) to provide the data rates. direction to the destination.

A diagram of the direct link telecommunications Accuracy requirements of the navigation subsystem subsystem selected is shown in Figure 7. The are determined primarily from the exploration flight telemetry subsystem provides data to the accuracy required by science. If regional accuracy modulation/demodulation subsystem, which modu is required, i. e. , it is only necessary to reach an lates the data onto a subcarrier. In the normal area or region of interest, then the accuracy if it mode of operation, the subcarrier phase modulates requirements will be much less stringent than the data onto an S-band or X-band RF carrier, is required to reach a particular site within an which is amplified by a TWT, transmitted to earth area or region. It also will be desirable to

1-35 and the expected photographic coverage elevation changes which can accuracy accurately determine of Mars from the Mariner 1971 and Viking missions, with various scientific measurements be correlated landmark navigation should be considered for their value. to enhance updating the prime navigation subsystem. The method used in navigating the vehicle must be capable of a certain degree of self-contained oper (e) Obstacle Detection ation, relatively simple to mechanize and capable The design of meeting the accuracy requirements. of a Mars roving vehicle can, in under vehicle control for The effectiveness must allow for operation measured in terms of the vehicle's ability as possible in order to over part, be as many operations to detect and avoid all obstacles interfering with inefficiency of earth-based control come the vehicle motion. Obstacles which can impede communications link with a delay time on through a vehicle motion are expressed in terms of the mobil minutes. The method selected to nav the order of ity capability of the vehicle. Both short- and long- igate the vehicle is a gyrocompass/odometer (typically, 1 m and 30 m, by a surface land range obstacle detection scheme, updated periodically techniques, using tactile and electro - technique. This combination pro respectively) mark matching optical sensors, will be required. The short-range the accuracy, simplicity, and degree of self- vides sensors are used primarily to assure vehicle operation required for a Mars roving contained safety, whereas the long-range sensor(s) is (are) Other methods, such as total inertial or vehicle. used for measuring the conditions at some distance navigation, have limitations (primarily in celestial from the vehicle for obstacle warning and path accuracy capability) which make them less adjustment. attractive. which use data from comparing the Obstacle avoidance strategies, A performance analysis'"' the necessary heading com subsystem with a total the sensors, provide gyrocompass/odometer for negotiating the obstacles. shownthat the gyrocompass/ mands to the vehicle inertial subsystem has only influence vehicle safety, The primary reason is that These strategies not odometer is superior. a direct bearing on the traverse effi serves as an external reference (as but they have the odometer The likelihood of the vehicle traveling the to a calculated reference), limiting ciency. compared minimum distance to the destination is directly errors in computed distance to a small percentage governed by early obstacle detection and avoidance of total distance traveled. The frequency of updates also is reduced considerably as a result, maneuvers. providing a more time-efficient surface operation. Functionally, the range obstacle detection subsys must provide data that can be used to identify The gyrocompass performance is affected by gyro tem obstacle in the path of the vehicle at a pre misalignment and drift, and platform tilt. Exter an to compensate determined range. Once the obstacle has been nal navigation updates are required determines the error which accumulates with identified, the vehicle computer the vehicle location necessary to avoid the obstacle. The time. A diagram of the gyrocompass/odometer commands of coverage of the long-range system should navigation subsystem is shown in Figure 8. sector include an azimuth angular range sufficiently wide likely vehicle paths. Resolution is Landmark navigation or "piloting"''' °» °) is a rela to encompass by the physical characteristics of the tively simple method requiring only a TV camera determined and terrain maps of the surface area of interest. obstacles. Surface landmarks are observed in the TV field -of - view and correlated with landmarks on the terrain A possible sensor configuration for obstacle detec map. Bearings to the landmarks, relative to tion is a laser whose beam is scanned back and vehicle heading, are measured and lines-of - forth through a range of azimuths on either side of position determined. The intersection of at least the vehicle. By detecting the time of the return two lines-of-position defines the vehicle's position signals and correlating these returns with the beam and heading in map coordinates. scan position and vehicle attitude, the position and type of obstacle can be determined. The accuracy of landmark navigation, relative to of landmark frequency, the map, is a function an obstacle and the error in determining Figure 9 is a block diagram of such landmark distribution beam is shown the number of landmarks increases, detection subsystem. Here, the bearing. As rad. azimuth range decreases. Uncertainty due to being scanned through a ±0. 35 the undertainty The return signal is errors of feature location reduces with in a 1-sec time interval. independent and sampled every 0. 09 measurements, but other errors such as map detected by the receiver more change. The range of detection is coordinate offsets are not reducible and, therefore, rad. of azimuth by receiver sensitivity and Mars sur limit the accuracy. determined face reflectivity properties. The output of the input to the thresh disadvantage of this technique is that sampling gate is conditioned and The greatest with vehicle attitude. process must be performed on earth. old detection element, along the navigation by the range to are transmitted to earth, landmarks The threshold level is determined TV pictures the surface, and vehicle position and heading com the intersection of the beam and identified, a level vehicle attitude. puted. At Mars distance, this process consumes a assuming a flat surface and is a function of return signal significant amount of valuable mission time and Threshold tolerance attitude determination does not appear practical as a single means of strength, noise, and vehicle navigation. Because of the inherent accuracy.

1-36 (2) The degree to which ground commands and (f) Antenna Pointing « * on-board commands are utilized in the execu variable. High bit rate science data from the Mars roving tion of a given sequence must be vehicle requires that a high-gain antenna be pointed on-board com at the earth. Two antenna mounts can be considered (3) The sequence of issuance of con for such an antenna: (1) an azimuth-elevation sys mands must be variable both via ground tem in which north and local vertical are used as trol and by automatic on-board state references and (2) a polar or hour-angle declination determination. amount in which one axis is oriented parallel to the planet's spin axis. Because of its flexibility and (4) An on-board logical element must be included simplicity during the earth tracking periods, the which has access to appropriate rover sensor hour-angle declination mount is the selected base information and has the ability to control the line configuration. sequence and time of the vehicle's actuators.

The hour-angle declination mount, requiring two (5) The on-board logical element must be capable gyros and four gimbals, is more autonomous, more of performing fundamental arithmetic and flexible, has better degraded mode capability, and Boolean operations. requires minimum on-board computer requirements. The first gyro is used to sense north. The second (6) The data rate from rover instruments must be gyro is used to erect the hour-angle axis parallel variable and capable of being controlled by a to the planet's spin axis. The sun is then acquired telemetry processing unit. by a two-axis sun sensor which is biased to account for the sun-Mars-earth angle, pointing the antenna (7) The data from the rover instruments must be toward the earth. processed and stored for re-transmission to earth. The selected Antenna Pointing Subsystem (APS) is capable of tracking the earth within ±35 mrad. The functional block diagram of a data handling This corresponds to the -0. 75 dB points of a 1. 2 m- subsystem which meets the requirements stated diarpeter parabolic antenna at S-band. Therefore, above is shown in Figure 11. the S-band operations can be performed by the APS alone. For X-band, the required improvement in pointing accuracy can be achieved either by cali MISSION OPERATIONS brating out the bias errors or by using a monopulse loop to track the uplink. The earth-based operations involved in the control of a Mars roving vehicle mission are quite exten The APS will acquire the earth within about 15 to sive and complex(lO). This study effort did not 20 min on the first day. Subsequent daily acquisi consider the mission operations area in sufficient tions will be shorter if the rover daily orientation depth to provide detail descriptions of the opera changes are small. The APS will be capable of tional requirements. However, the operational pointing the antenna at any part of the sky which is aspects were given a general consideration and the greater than 0. 26 rad. above the local horizon. results are described below.

A block diagram of the APS is shown in Figure 10. A functional block diagram of the earth-based oper The control loops for the azimuth and elevation ations is shown in Figure 12. The mission opera drives are identical, utilizing gyros for position tions system is seen to be comprised of the error sensors to drive the control axes. When the mission-dependent computing and processing, data polar shaft is aligned parallel to the planet's spin printout and visual displays, mission control, axis, power is removed from these two loops. The science planning and evaluation and the vehicle stepper motors contain permanent magnet or command and control element. It should be noted mechanical detents which act as brakes when power that only those elements of the total mission opera is turned off. When the polar shaft erection mode tions system directly involved in the vehicle opera is completed, the hour-angle and declination axes tions are considered. The various DSN stations, are enabled and the sun is acquired, using the sun coverage, etc. , are not considered. sensors. A bias is mixed with the sun sensor out are conditioned and put to point the antenna at the earth. Data received from the vehicle formatted for driving visual displays, magnetic tape recorders, and line printers. Mission con direction (g) Data Handling trol, whose function is to provide overall of all mission operations, establishes the vehicle of the sci A number of data handling design requirements objectives based upon recommendations the command have been established which will permit a more ence planning and evaluation team and planning and efficient utilization of the rover capabilities and and control operations team. Science the mis will in effect progressively diminish the adverse evaluation is concerned with maximizing sites to effects of the long two-way transmission time sion payoff and, thus, selects the potential and con associated with the mission. These requirements be surveyed and explored. The command human operator are: trol element includes the "standby" and the command and control console. The oper to all internal and (1) All primary control functions must be capable ator must know or have access and must be of being actuated by either ground commands external vehicle status information, control measures. or on-board commands. prepared to execute emergency

1-37 The earth-based computer system, not unlike the CONCLUSIONS vehicle system, is extensive and complex. These characteristics, however, can be accommodated The successful development of a surface roving more easily because they are not subject to the vehicle for a 1979 Mars mission hinges on many physical constraints imposed by spacecraft pay load factors. The vehicle design must have the capa limitations. In a normal operation mode, the bility to operate for extended periods of time, earth-based computer system will receive and relatively independent of the earth, to withstand process a data "dump" over the high-gain commu the harshness of the Martian environment, and to nications system following the completion of a pro travel hundreds of kilometers, independent of the grammed science investigation sequence, and surface delivery spacecraft. These capabilities receive and process a limited amount of data over imply, respectively, system techniques which the omnicommunications system during motion focus on autonomy and reliability, environmental toward the destination. The system will be required sensing and control, and a safe and effective to route the various portions of the processed data motion control. to the command and control operations team, the science planning and evaluation team, and mission The 1979 opportunity is both favorable and unfavor control. Each group will analyze certain portions able in terms of implications on the mission design. of the data and, under the direction of mission con The foremost implications are that 1979 requires trol, determine the next destination and/or the a low launch energy (C-$ from 9-15 km^/sec^) and, necessity to alter the operational mode. because the earth and Mars are at extreme dis tances, the round trip light time delay is quite At all times, the earth-based system must be in a large. During the anticipated 1-yr mission, the position to take control of the vehicle. The com time delay reaches 41 min, round trip. This has puter software system must have capabilities iden serious implications on the ability of the vehicle to tical to the vehicle-based software system qualities travel large distances over the surface. It and, in addition, the capability to handle contingen requires some degree of onboard control in order cies which the on-board system cannot handle to limit the inefficiency associated with strictly alone. An earth-based "learning" process must be earth-based control. During the early phases of continuously operative in the command and control the mission, the dependence on earth-based con operations team to achieve the desired "near- trol will be higher. As mission confidence optimal" control process. Earth-based data analy increases, more reliance will be placed on the sis software will be required for evaluation and on-board control system. Consequently, the improvement of the current vehicle-based model vehicle traverse capability will be low initially and to support the learning process. increase appreciably as the mission lifetime increases. Command generation is another major requirement on the earth-based system. The command genera The nature of the roving vehicle system requires a tion software must be sufficiently flexible and effi complex interconnection of many vehicle-based and cient to provide minimum-time command capability earth-based elements, linked by the telecommuni under a diverse range of circumstances. For cations capabilities of the Deep Space Network and example, the sequence of individual vehicle com the communications system of the vehicle. Each mands necessary to perform a particular operation element must possess the qualities of functional must be assembled, verified, and transmitted to performance and reliability dictated by the mission the uplink rapidly. Software flexibility also must requirements, and each must be capable of func be provided to allow for direct access discrete con tioning in the total, integrated system environment. trol of on-board sensors or actuators. To develop the capability to perform a Mars rover Transmitted data will be used to drive the ground mission, a strong research and advanced develop display system for the three groups mentioned ment program is needed several years prior to a above. The mission control display will relate the technology cut-off. (For a 1979 rover, the cut-off overall mission status. Science operations will may be as early as 1975 or 1976). The program have software for filtering and enhancing television should be structured to give high priority to pictures and for processing (calibration and filter obstacle detection and avoidance, on-board control philosophy, ing) other data from the science instrument com computer capabilities, and mobility. plement. The command and control team will be Efforts in these areas will contribute needed tech concerned with vehicle motion and, thus, the nology for developing a vehicle with a range of ground display will provide the vehicle operator hundreds of kilometers and capable of surviving with all vehicle status information. 1 yr on the Martian surface. In each of the areas, system reliability must be a dominant factor in the development process. Several relatively detailed studies have been per formed on the subject of mission operations for a lunar roving vehicle. One is given in Reference 11. ACKNOWLEDGEMENT The referenced study, plus the cursory look at mission operations presented here, should pro The authors would like to acknowledge contributions vide considerable background for the design of a to this paper from several individuals at JPL who mission operations system for a Mars roving collectively participated in a short design study' 13 ) vehicle. of a 1979 Mars rover. The individuals were:

1-38 H. Bank, W. Dorroh, J. Gilder, T. Gottlieb, (8) Anthony, V. F. , Lewis, R. A. , Moore, J. W., G. Hornbrook, R. Imus, D. Kurtz, L. Lim, "Man-Machine Systems; An Evaluation of Roving W. McDonald, P. Roberts. Without their efforts Vehicle Navigation by Landmarks, " JPL Docu in the design study this paper would not have been ment No. 760-57, April 1970. (JPL Internal) possible. (9) Coryell, R. B. , Rubin, D. K. , "Experiments in Piloting by Landmark on the Moon, " JPL Docu REFERENCES ment 760-37, April 1, 1969. (JPL Internal) (10) Moore, J. W. , "Computer System Require (1) "Dual Mode Lunar Roving Vehicle, " prelimi ments for an Autonomous Martian Roving Vehicle," nary design study, prepared under NASA/MFSC AIAA Paper 69-980, September 8, 1969. Contract No. NAS8-24529, Grumman Aerospace Corporation, February 1970. (11) McCormick, C. W. , "Operations Profiles for Lunar Roving Missions, " JPL Document No. 760- (2) "Dual-Mode Manned/Automated Lunar Roving 46, May 1970. (JPL Internal) Vehicle Design Definition Study, " prepared under NASA/MFSC Contract No. NAS8-24528, Bendix (12) Moore, J. W. , et al, "An Exploratory Investi Aerospace Systems Division, January 1970. gation of a 1979 Mars Roving Vehicle Mission," JPL Document 760-58, December 1, 1970. (3) "Roving Vehicle Motion Control, " General (JPL Internal) AC Electronics Defense Research Labora Motors ILLUSTRATIONS tories, TR67-34, June 1967. Figure 1. Round Trip Light Time (Minutes) (4) Lim, L. Y. , A Pathfinding Algorithm for a Earth-Mars. Myopic Robot, Jet Propulsion Laboratory, Figure 2. Mission Profile, One Martian Day TR-32-1288, 1968. (August 5, 1980, -. 52 rad Latitude). Figure 3. Typical Mars Rover Range Capability (5) Kirk, D. E. , and Lim, L. Y. , "A Dual-Mode (1-yr Mission). Routing Algorithm for an Autonomous Roving Figure 4. Mars Roving Vehicle Entry and Landing Vehicle, " IEEE Trans. Aerospace & Electron Sequence. System, Vol. AES-6, No. 3, May 1970. Figure 5. Mars Roving Vehicle Configuration. Figure 6. Mars Rover Power Subsystem. (6) Lewis, R. L. , "Roving Vehicle Navigation Figure 7. Mars Rover Telecommunications Subsystem Feasibility Studies, " presented at Third Subsystem. Hawaii International Conference on System Sci Figure 8. Gyrocompass/Odometer Subsystem. ences, University of Hawaii, January 13, 1970. Figure 9- Obstacle Detection Subsystem. Figure 10. Antenna Pointing Subsystem. (7) Moore, J. W. , "Surface Navigation on Mars, " Figure 11. Mars Rover Data Handling Subsystem. presented at the ION National Space Meeting, Figure 12. Functional Block Diagram of Earth- Moffett Field, California, February 17-19, 1970. Based Mission Operations.

1-39

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