Mariner Mars 1964telemetry and command systemThe Mariner Mars telecommunication system was designedto transmit video datafirom the vicinity of Mars and additionalscientific data during the.flight, as well as direct andquantitative commands to the spacecraft

Richard P. Mathison

Spacecraftcommands

Spacecraftengineeringdata Mode

control

I

! IVideodata

TT~~~~1Moe Transmissioncontrol rate control

Jet Propulsion Laboratory

High-gain\\antenna

2298 Mc/s

Redundance|co ontrol

Fig. 1. Mariner 1964 spacecrafttelecommunication system.

IEEE snectrum JULY 196576

The initial objective of the Mariner Planetary-Inter-planetary Program is the preliminary probing of theplanets Mars and Venus by unmanned spacecraft. Theprobing of Venus was successfully accomplished byMariner II. The next operation directed toward this ob-jective is the probing of Mars by the Mariner IV space-craft, which is now in transit.The primary objective of the Mariner Mars 1964

project is to conduct close-up, fly-by scientific observa-tions of the planet Mars during the 1964-1965 op-portunity and to transmit the results of these observationsback to earth. Television, cosmic dust, and a complementof fields and particles experiments are being carried byMariner IV. In addition, an earth occultation experimentis planned.A secondary objective is to provide experience and

knowledge about the performance of the basic engineeringequipment of an attitude-stabilized fly-by spacecraftduring a long-duration flight in space farther away fromthe sun than is the earth.The Mariner IV spacecraft was launched on November

28, 1964. The spacecraft is fully attitude stabilized, usingthe sun and Canopus as reference objects. It derives powerfrom photovoltaic cells, arranged on panels having body-fixed orientation, and a battery, which is used for launch,trajectory correction maneuvers, and back-up. The tele-communication system for the Mariner Mars 1964Mission is comprised of spacecraft-borne equipment andthe NASA Deep Space Net.I It is required to performthree functions: (1) tracking the position and velocity ofthe spacecraft, (2) telemetering engineering and scientificdata from the spacecraft, and (3) transmitting commandsto the spacecraft. The design of the spacecraft equipmentis based upon techniques that were used for the MarinerII spacecraft. 2,:" These techniques have been extended andmodified to improve equipment reliability, accommodatethe increased communication range required by the Mars1964 Mission, and utilize the characteristics of the Mars1964 trajectories to effect simplifications in the spacecraftequipment.

Single CW radio-frequency carriers that are trans-mitted to and from the spacecraft are used for trackingthe spacecraft and transmitting the telemetry and com-mand information. The functional arrangement of the

Fig. 2. Circulator switch circuit.

1

Inputterminal

Normaloutputterminal

231

123

Reversedoutputterminal

312

spacecraft subsystems utilized to accomplish this isshown in Fig. 1. For both the telemetry and commandfunctions, pulse code modulation, phase shift key, andphase modulation (PCM/PSK/PM) techniques in com-bination with pseudorandom sync codes provide efficient,accurate transmission of the data over interplanetarydistances.The telemnetry portion of the system is required to

transmit video data in digital form from the vicinity ofMars and both scientific and engineering data duringthe flight from earth to Mars. Since the rate at which thevideo data is gathered exceeds the capacity of the te-lemetry channel, data storage and playback are providedby a synchronous, endless-loop tape recorder capable ofstoring 20 frames of video data.The duration of the Mars 1964 Mission is approxi-

mately eight months, in contrast to the Venus 1964Mission, the duration of which was four months. Inorder to accommodate this increased equipment operatingtime, modest reliability improvements were incorporatedwithin the constraints of available power and weight.These improvements take the form of better components,extensive part screening, worst-case circuit designs, andredundant elements in the telemetry modulator and trans-mitter.By utilizing the unique characteristics of the Mars

1964 minimum-energy trajectories,4 considerable savingsin spacecraft weight and complexity were realized. Thevariation in earth, spacecraft, sun, and Canopus geometrypermitted the use of a moderately high-gain antenna thatis fixed with respect to the spacecraft and thus eliminatedthe need for antenna pointing mechanisms.The command system provides for the transmission of

both direct and quantitative commands to the spacecraftin digital form.A detailed discussion of the scientific instruments and

the data automation system that controls and interfaceswith the instruments is beyond the scope of this article.5Moreover, details of the modulation-demodulationtheory that forms a basis of the data transmission tech-niques have been adequately covered elsewhere&8 andwill not be repeated here. However, estimates of expectedcommunication performance will be included.

Radio subsystemThe radio subsystem is required to receive a modulated

RF carrier from stations of the Deep Space Net (DSN),demodulate command and ranging signals, coherentlytranslate the frequency and phase of the RF carrier by afixed ratio, modulate the carrier with telemetry andranging signals, and retransmit it back to earth. It consistsof an automatic-phase-control receiver, redundant ex-citers, redundant power amplifiers, power supplies, low-and high-gain antennas, and associated transmission andcontrol circuits. It operates at S-band frequencies, receiv-ing at 2116 Mc/s and transmitting at 2298 Mc/s.As received from earth, the up-link RF signal is phase-

modulated either singly or simultaneously by a compositecommand signal and a coded ranging signal. It is of theform

SR = A("y,r) sin [wot' + ,(i') + r,(t')]t' = t _ r(t)

c

(1)

(2)

Mathison-Mariner Mars 1964 telemetry and command system 77

20' 0. 34 7 wherer ~X I /A is the received signal level, a function of spacecraft

4 \\I \320' attitude yandspacecraftearth ranger40\' 320 w0 is the carrier frequency transmitted by the DSN station,7 / is the phase modulation by the composite command

' ,' ;--clock 1 / es / \; signal\/~;\angle I / \t \>< > r4, is the phase modulation by the coded ranging signal

r(i) is the spacecraft-earth range, a function of time tc is the velocity of propagation

This signal is demodulated by the automatic phase con-trol, double superheterodyne receiver which tracks thewoi \component of the carrier phase. The composite com-

s0o.>- / / \ 2800 mand modulation and coded ranging signals are sent to__0 45060

2

the command detector and the exciter phase modulators,-k 20X_ : 80-- -25o- t- respectively. When the receiver is phase-locked to the

70 \o 20-V / received signal, it generates for the transmitter exciter aDo, *--- hiC,//lt \\\9\>, >>lRo2 260' filtered phase reference that is coherent with the woz'

\ .- Xx)%L20\ / component of the received signal. The phase of the trans-mitted signal is then related to that of the received signalby a fixed ratio to within an error of less than I radian

20-4)< \/ /' < tX 0 < -X2400 rms. The resulting transfer function is given approxi-20'\>

With this relationship the ground stations are providedwith a signal that permits two-way Doppler tracking. 9The transmitted signal is phase-modulated by a com-

posite telemetry signal and the coded ranging signal.While the telenmetry signal modulates the carrier con-tinuously, the ranging modulation can be turned on or offby ground command.When a signal is not being transmitted to the space-

craft, transmitter frequency control is provided by anauxiliary crystal oscillator. This noncoherent mode ofoperation permits one-way Doppler tracking, angularposition tracking, and telemetry reception by the groundstations.

In order to provide increased reliability over theMariner II design, redundant exciters, power amplifiers,and power supplies have been incorporated in the trans-mitter. Each exciter consists of an auxiliary oscillator, aX 4 frequency multiplier, a phase modulator, a X 30 fre-quency multiplier, and an output isolator. Either excitercan be coupled to either power amplifier by a circulatorswitching network. Similarly, the input and output cir-cuits of the power amplifiers are coupled through circu-lator switches.The control of the switching between these elements

is provided by either ground command or on-boardfailure detection. In the case of ground command control,the receipt of the appropriate direct command causes thecontrol unit simultaneously to transfer the dc power fromthe active to the inactive element and reverse the circula-tor switch or switches. For the exciters, the modulation,phase reference, and mode control inputs are fed to bothexciters in parallel.

In the case of switching by on-board failure detection,power monitors sample both the exciter and power ampli-fier RF power outputs. When an output power dropsbelow a preset level, a gate in the control unit is enabledwhich allows cyclic pulses from the control computer andsequencer (CC& S) to toggle the relay driver circuit.Upon the receipt of one such pulse, the control unit trans-fers the dc power and RF circuits in the same manner aswhen a ground command is received. If the power outputfrom the redundant element then exceeds the threshold,the gate inhibits further transmission of pulses to thedriver circuit. The thresholds for enabling the gates tooperate are set at 3 dB below the nominal exciter andpower amplifier outputs. The cyclic pulses occur onceevery 66% hours. Thus, the maximum switching timeafter a failure is 66% hours.

Circulator switches were chosen for the control of theRF transmission paths because they appeared to offersignificant reliability advantages over conventional elec-tromechanical coaxial switches. As an RF circuit, thecirculator is simply a strip-line Y connection with nomoving parts. The Y is surrounded by a ferrite materialthat is polarized by a dc electromagnetic field. Signalflow through the device is circular, as indicated in Fig.2. By reversing the magnetic field, the signal can be madeto "circulate" in the opposite direction and hence theswitching action. In the event of a loss in electromagneticfield, the circuit will function like a transmission lineT, with the attendant power splitting and increased mis-match losses, but will not cause a complete loss of per-formance.The Mariner Mars 1964 spacecraft uses both the sun

and the star Canopus for attitude references. Sun sensors

0

10

(0

200 240 280 320 0 40 80 12X +Z axis -Z axiso Spacecraft rotation, degrees

Fig. 5. Typical low-gain antenna pattern.

provide pitch and yaw control such that the roll axis ispointed toward the sun, and the Canopus sensor pro-vides roll position control. With this type of attitude con-trol the position of earth as seen from the spacecraft (thedirection of the spacecraft coordinate system) varies asshown in Fig. 3 for a typical Mars 1964 trajectory. 4

It can be seen that the locus remains within one hemi-sphere of the spacecraft during the entire flight and withina relatively small angular region during the later portionof the flight, from 130 days before the Mars encounter to20 days past encounter. A comparison of this characteris-tic and the required minimum antenna gain vs. time-of-flight showed that the gain requirements could be metwith a combination of one low-gain and one high-gainantenna, both of which were fixed relative to the space-craft (Fig. 4). The low-gain antenna provides coverageduring the first 70 to 95 days of flight, whereas the high-gain antenna fills in the remaining period until approxi-mately 20 days past encounter.The low-gain antenna consists of a cruciform aperture

at the end of a low-loss circular waveguide, which alsofunctions as the support structure. In order to minimizepattern distortion by reflections from the spacecraftstructure, the aperture is mounted well away from thebulk of the spacecraft. As shown in Fig. 5, the antennaprovides a pattern of revolution about the roll axis witha maximum gain of 5.5 dB at 2298 Mc/s in the directionof the -Z spacecraft axis (oriented toward the sun) and aminimum gain of -6 dB with respect to circular isotropicover the entire -Z hemisphere. The pattern at 2116 Mc/sis similar.

Since the thrust vector of the mid-course motor isperpendicular to the -Z axis, the earth can be kept withinthe -Z hemisphere while the thrust vector is pointed inany arbitrary direction. Thus, the low-gain antenna pat-tern also meets the requirement for providing coverageduring mid-course maneuvers of unrestricted direction.The high-gain antenna is a 46.0- by 21.2-inch parabolic

reflector that is illuminated by a pair of turnstile elements.These elements are arranged so that a right-hand cir-cularly polarized beam is projected with a maximumgain of 23.5 dB (at 2298 Mc/s) and a half-power beam

20 160+Z axis

width of 13.50 by 7.5°. The beam is positioned so thatcoverage is provided from approximately 90 days afterlaunch until 20 days past encounter. As a result of usingthis design as opposed to the one-degree-of-freedom

Mathison-Mariner Mars 1964 telemetry and command system 79

antenna that was used on Mariner JJ,2,3 an estimated 50pounds of spacecraft weight was saved by the associatedreductions in structure, actuator, control electronics,and power requirements.

Three transmitting and receiving modes are available:

1. Transmit low gain, receive low gain.2. Transmit high gain, receive low gain.3. Transmit high gain, receive high gain.

These modes provide the required coverage during theacquisition, cruise, mid-course maneuver, and encounterphases of the flight. Selection of the proper mode is con-trolled by programmed CC&S commands or groundcommands as a backup.

In addition, two failure mode controls are provided.First, if roll position control is inadvertently lost whilethe receiver high-gain mode is being used, the loss of theCanopus sensor signal automatically switches the receiverto the low-gain antenna so that command capability canbe maintained. Second, if the spacecraft does not receivea signal from earth at least once between the occurrenceof the 66%j-hour cycle pulses, as signified by receiverphase lock, the control unit automatically switches thereceiver from one antenna to the other after the receiptof two such pulses. The receiver is subsequently cycledbetween the antennas once every 66% hours until phaselock is obtained. This later mode control provides partialredundance for some antenna failure modes.Summaries of the principal radio subsystem trans-

mission and reception parameters are given in Tables Iand II, respectively.

Telemetry subsystem: basic techniqueThe principal functions of the telemetry subsystem on

the spacecraft are to time-multiplex engineering andscientific data samples and to encode them for efficientmodulation of the spacecraft-to-earth RF carrier. Thesubsystem is specifically required to

1. Transduce engineering parameters into electricsignals.

2. Time-multiplex (commutate) engineering andscientific measurement signals.

3. Convert engineering data samples to binary words.4. Store digitally encoded video data.5. Phase-shift-key a subcarrier with the binary signal.6. Generate a cyclic, binary, pseudorandom sequence

for use in synchronizing the encoding and decodingof the telemetry data.

7. Phase-shift-key a second subcarrier with the synccode.

8. Combine the two subcarriers into a compositetelemetry signal.

The basic timing for the subsystem is derived from the2400-c/s spacecraft power frequency, which is divideddown to provide two subcarrier frequencies, one fordata and one sync. The frequency divider is arranged toprovide two data transmission (bit) rates, 331ij and 8Y3bits per second (b/s). While the 33'3-b/s rate is usedduring preflight check-out and the early flight phases upthrough a first mid-course maneuver, the 813-b/s rate isused for the remainder of the flight. In-flight selection ofthe data rate is controlled by ground command and

1. Spacecraft radio transmission parameters (2998 Mc/s)Transponder Low-Gain Channel Transponder High-Gain Channel

Parameter Value Tolerance Value Tolerance

Total transmitter powers +40.0 dBm 40.5 dB +40.0 dBm 40.5 dBCarrier modulation lossb -4.1 dB +0.9 dB -4.1 dB +0.9 dB

-1.0 dB -1.0 dBTransmission circuit lossc -1.7 dB +0.2 dB -1.3 dB +0.2 dB

-0.3 dB -0.3 dBSpacecraft antenna gaind +6.0 dB +41.8 dB +23.2 dB ±1.1 dB" Ten watts nominal output of traveling-wave-tube amplifier.b Based on modulation indexes of 0.809 rad peak for data subcarrier and 0.451 rad peak for sync subcarrier.e Includes all circuitry between the output of the TWT amplifier and the input to the antenna.d Referenced to perfectly circular isotropic pattern maximum.

IL. Spacecraft radio reception parameters (2116 Mc/s)Transponder Low-Gain Channel Transponder High-Gain Channel

Parameter Value Tolerance Value Tolerance

Antenna gain (pattern maximum)aReceiving circuit lossbEffective system noise temperaturec

Carrier APC noise bandwidth (2BLO)dCarrier threshold SNR in 2BLOTwo-way Doppler trackingeCommand reception

+6.5 dB-1.0 dB27000K

20.0 c/s

+3.8 dB+8.0 dB

±1.8 dB±t0.2 dB+17000K-610°K

1.0 dB

+21.8 dB-0.9 dB27000K

20.0 c/s

+3.8 dB+8.0 dB

4±1.1 dB±t0.2 dB+17000K-610°K

±1.0 dB

a Referenced to perfectly circular isotropic pattern maximum.b Includes all circuitry between the antenna and the input to the transponder receiver.e Includes contributions due to antenna temperature, circuit losses, and noise figure at input to preselector, 10 dB (+2 dB, -1 dB).d Tolerance included in uncertainty of system noise figure.e 3.8-dB SNR is required for +2.0-dB ground receiver degradation.

IEEE spectrum JULY 196580

CC&S command. Either rate can be selected by groundcommand, but the CC&S selects only the 8'3-b/s rate192 days before encounter. The CC&S control is to in-sure that the 813-b/s rate is used at encounter in the eventthat command capability is lost.The square-wave sync subcarrier drives a redundant

pair of pseudorandom code generators which generate acyclic 63-bit code. A set of word gates, in turn, generatesbit and word sync pulses that are used to synchronize(1) the stepping of the commutator, (2) the analog-to-digital converters, (3) the readout of data from the dataautomation system, (4) the readout of the event registersand timers, and (5) the playback of the stored video data.The word sync pulses occur once per cycle of the code,whereas the data-bit sync pulses occur once every ninecode bits, or seven times per code cycle. Thus, each dataword is seven data bits long.

In order to convey the bit and word sync timing tothe ground stations for use in synchronous demodulationof the telemetry subcarrier, the code also phase-shift-keys the sync subcarrier. The resulting composite te-lemetry signal that modulates the spacecraft-to-earthcarrier is given by

D(t) = Vd [1.79d (-kt) @ a(4fdt) + X(b) 0 a(2fdt)](4)

whereVd is the amplitude of the complex four-level waved(2fdtl9) is the binary telemetry data of amplitude + 1 and

bit rate (2/9)fda(ft) is a symmetrical square wave of amplitude i 1 andfrequencyf

X(fdt/2) is a cyclic, binary, pseudorandom sequence ofamplitude i 1, length 63 bits, and bit rate fd/2

0) represents modulo 2 additionAt the ground station, a local model of the code is

phase-locked to the received code. Word gates identicalto those in the spacecraft code generators then produceaccurate bit and word sync pulse trains. For a detaileddiscussion of the technique, the reader is referred toReferences 6, 7, and 8.Analog engineering measurements are sampled by a

solid-state commutator that provides 100 channels, 90 ofwhich are used for measurements and ten for synchroniza-tion points and subcommutation. These channels are

1l1. Registered eventsChannel Events

1 Pyrotechnics current pulseGyro turn-onSolar panel 1 open

2 CC&S eventsSolar panel 2 open

3 Pyrotechnic armPyrotechnic current pulseSolar panel 3 openRecorder end of tape signal

4 Ground command eventsSun acquiredSolar panel 4 openScan platform unlatched

divided among ten decks of ten channels each and arearranged to provide three sampling rates.The pulse-amplitude-modulated output of the com-

mutator is fed to two analog-to-digital converters, whichconvert the data samples to serial 7-bit words by a suc-cessive approximation technique. The output of the con-verter forms one of four data sources that comprise thetelemetry modes.Four modes of data transmission are provided for:

(1) engineering data, (2) engineering and science data, (3)science data, and (4) stored video data and engineeringdata. In the first mode, only engineering data from thecommutator, event register, event timer, and commandmonitor are transmitted, primarily for maneuver andcheck-out phases. In the second mode, both engineeringand science data are transmitted in an alternating se-quence of 140 engineering data bits followed by 280science data bits. This mode is intended for most of thecruise phases. In the third mode, only science data aretransmitted, as received from the data automation sys-tem. This mode is designed for use at planet encounter.In the fourth mode, stored video and engineering dataare transmitted in alternating periods of approximately9 and 1.5 hours, respectively. This mode provides forreadout of the video data taken during encounter andperiodic monitoring of the spacecraft performance afterencounter.

Event-type signals that signify the occurrence of eventssuch as motor-start, receipt-of-command, or solar-panels-open are accumulated as they occur in four sepa-rate registers. Each register accumulates different typesof events, as shown in Table III, and holds up to eightcounts before recycling. The registers are sampled inpairs at the high commutation rate in synchronism withthe commutator, so that the state or count of two registersis conveyed by one 7-bit word.An event timer measures the duration of certain events,

such as the mid-course motor firing duration, by dividingthe word sync rate by two and accumulating the numberof pulses that occur between the start and the end of theevent. This number is sampled at the medium rate also insynchronism with the commutator.During the Mars encounter, a television subsystem

which operates under data automation system controlperiodically generates video data in binary form. Thesedata and the mode 3 instrument data generated at aneffective rate of 10 700 b/s are organized in 516 168 bitframes, of which 504 400 are TV-related. Since this datarate greatly exceeds the 813-b/s radio transmission capa-bility at encounter, a data storage subsystem holds thedata for postencounter readout.Data storage is accomplished by an endless-loop tape

recorder. This machine records binary data and syncpulses on two tracks, filling one track at a time on each oftwo consecutive tape cycles. Recording is started andstopped by control signals from the data automationsystem to coincide with the encounter data frames. Inorder to prevent overrecording after the two tracks arefilled the first time, end-of-tape signals automaticallystop the recorder after the second complete tape pass.The tape is then in the correct position for subsequentplayback.

Playback, at the 8h-b/s transmission rate and syn-chronous with the telemetry bit sync pulses, is accom-plished by an automatic phase control servo which con-

Mathison-Mariner Mars 1964 telemetry and command system 81

trols the tape speed so that the recorded bit sync pulsesare kept in phase with the telemetry bit sync pulses. Bythis means the pseudorandom sync signal allows syn-chronous demodulation of the recorded data at theground stations.

In conjunction with the starting and stopping of therecorder during the record cycles, approximately 3 to 5feet of tape are used while the machine accelerates anddecelerates. No data are recorded on these segments.During the continuous playback, these blank spots pro-vide approximately 1.5 hours in which spacecraft engi-neering data are inserted for periodic monitoring of post-encounter spacecraft performance. Control of this alter-nation between the recorded and engineering data isprovided by a circuit that senses the presence or absenceof data on the tape. Table IV summarizes the characteris-tics of the tape machine.As with the radio subsystem, limited redundance has

been incorporated in the telemetry subsystem for in-creased reliability. This redundance is in the form of twopseudorandom code generator analog-to-digital con-verter pairs which operate with parallel inputs and logical"or" coupled outputs. Only one pair operates at a time,and this pair is selected by ground command.

In addition, the commutator sequencer has been de-signed so that many of the possible failure modes resultin a modification or "short counting" of the sequencerather than a complete stoppage. The number of channelsthat would be lost for a given failure depends on thelocation of the failed component so that a varying degreeof partial success can exist. For example, a short countin a low-rate deck would not affect the higher-rate chan-nels, while a short count in a high-rate deck could bypassa large number of low-rate channels.

Finally, redundant components such as resistors,diodes, and capacitors have been employed in the powertransformer-rectifier unit. Table V lists the principaltelemetry subsystem parameters.

Command subsystemCommands are transmitted from DSN ground stations

to the spacecraft by +t;vo subcarriers, which phase-modu-late the earth-to-spacecraft RF carrier. One subcarrieris phase-shift-keyed by serial binary command words,and the other by a pseudorandom sync code in a man-ner similar to that used for telemetry data transmission.The command subsystem is required to detect and de-

code the command words, of which there are two types:direct commands, which result in selected switch closures,and quantitative commands, which convey a magnitudeand polarity for spacecraft maneuvers.

Initial acquisition is achieved by slightly offsetting thefrequency of the clock at the ground stations from theaverage static frequency of the loop voltage-controlledoscillator (VCO). Under this condition, the local code isslowly shifted in phase with respect to the received codeuntil the phases match. The frequency difference is madesmall enough so that the automatic phase control loopreceives sufficient signal to acquire phase lock and theacquisition is complete.

Outputs from the command subsystem include thedirect command switch closures, the quantitative com-mand bits, bit sync pulses, alert pulses for the CC& S, andseveral telemetry signals. In the case of both direct andquantitative command output circuits, complete dc isola-

IV. Video storage characteristics

Record rate 10 000 b/sPlayback rate (synchronous) 81/3 b/sStorage capacity 5.24 X 106 bitsNumber of tracks 2Type of tape machine Endless loop

V. Telemetry parameters

Type of encoding

Channel requirements:Engineering measure-ments

Event countersWord lengthTransmission ratesWord error probability atthreshold

Required ST/(N/B)* for biterror probability = 5 X 10-3

Data channel modulationloss

Sync channel thresholdS/(N/B)

Sync channel modulationloss

Sampled data, digital PSKwith pseudonoise sync

904

7 bits331/3, 81/3 b/s

1 word in 28 (bit error proba-bility = 5 X 10-3)

7.6 i 0.7 dBb/s

-4.6 =t 0.6 dB

11.0 i 0.5 dB,,/

-10.5 dB (+0.2 dB, -0.3 dB)

*S = signal power; T = time for one bit; N = noise power; B =bandwidth.

VI. Command parameters

Number of commandsDiscreteQuantitative

Modulation type

Word lengthTransmission rateCommand threshold criteria:

Probability of correctly executing adiscrete command in one attempt

Probability of completely executinga quantitative command in oneattem pt

Probability of a bit error in a com-pletely executed quantitative com-mand

Probability of a false discrete orquantitative command being exe-cuted when another command issent

Required carrier SNR in 20-c/s band-width at command threshold

Required command channel ST/(N/B)at threshold

Command channel modulation lossRequired sync channel SNR atthreshold

Sync channel effective noise band-width

Sync channel modulation loss

291 address,

3 subaddressesDigital PSK withPN sync

26 bits1 b/s

>0.7

>0.5

tion is maintained from the interfacing spacecraft sub-systems.As an aid to acquisition and in-flight performance

monitoring, the sync channel VCO frequency and in-locksignals are telemetered. For this purpose, a special counterconverts the VCO frequency to a binary number whichis periodically sampled by the telemetry system. Table VIlists the principal command subsystem parameters.

PerformanceThe telecommunication system is required to provide

tracking, telemetry, and command performance fromlaunch to 20 days past encounter, including all of the in-termediate phases. In order to reasonably assure thiscapability, it was desired to choose the system param-eters so that the nominal received signal levels alwaysexceeded the threshold signal levels by at least the linearsum (in dB) of the adverse tolerances. 10 This criterion hasbeen met for all functions and flight phases, except for thetelemetry for a period of 10 to 26 days, depending on thelaunch date.

Figure 6 illustrates, for a typical trajectory, the nominalreceived carrier level for the spacecraft-to-earth chaninelvs. time from launch. The variations are due to both the

-130t=rE

-140 \ _ _

Low-gain ante na-150

*-160--- -~__-~--_

increasing range and the variable antenia gains, and it isapparent where the performance of the low-gain antennaleaves off and that of the high-gain antenna takes over.For the diplexed tracking feed and maser ground

station configuration, the nominal threshold carrier levelfor telemetry is -164.4 dBm at 813i b/s. A comparisonbetween this value, the nominal carrier levels, and the sys-tem tolerances (Fig. 7) shows that the design criter-ion hasbeen met over most of the flight, and the extent to whichit has not been met at the transition region. In the transi-tion region, the telemetry performance may be marginal.

This situation is a result of the antenna position com-promise that had to be made between mid-flight per-formance and postencounter performance with a rela-tively simple antenna design. Mid-flight performance wassacrificed to meet the 20-day postencounter requirement.Since the nominal carrier level is never less than the nomi-nal threshold level, it is considered to be a reasonablecompromise.The nominal received car-ier levels for the earthi-to-

spacecraft channel are shown in Fig. 8. Since the samespacecraft antennas are used for transmitting and re-ceiviig, both up and down channels exhibit similar timevariations. A comparison between the command thieshi-

Time from launch. days

Fig. 6. Received signal level vs. time, spacecraft to earth.

Fig. 7. Telemetry performance margin vs. time. Cip'exed tracking antenna with maser, 81,; b/s.

15 -r-ow-gain antenna

us \ \ I \I High-gain antennmE 10 _ ________-t

c- 5 ___ __-F]l s_ L - '--t-

Time from launch, days

Mathison- Mariiier Mars 1964 tecemetry and comilmanid system 883

Fig. 8. Received signal levelvs. time, earth to space-craft. Ground transmissionpower = 10 kW, commandmodulation on.

Fig. 9. Command perform-ance margin vs. time.Ground transmissionpower 10 kW.

30-

25

,20 __E

15E0

0rt

, -nE-vm

> -1105)

r7UOD

- -120a)'.a)

- 130tr

Low

T- T --- --Igain an.enna

-140[0

200o --220 .4Time from launch. days

120 140Time from launch, days

old carrier level of -143.3 dBm, the nominal receivedlevel, and the system tolerances (Fig. 9) shows that thedesign criterion for command has been met for all flightphases.

ConclusionThe Mariner Mars Mission for 1964 required a tele-

communication system to provide tracking, telemetry,and command capabilities over communication distancesup to 260 million km and which would operate for ap-proximately 8 months in an interplanetary space environ-ment. The design that has been described is an extensionand modification of well-proved techniques, where themodifications included required improvements in per-formance and limited redundance to provide greaterreliability.

This article is a condensed version of a paper presented at theSixth Winter Coniventioni on Military Electronics, Los Angeles,Calif.. Feb. 3 5. 1965. It presents results of one phase of researchcarried out at the Jet Propulsion Laboratory, California Institute ofTechnology, under Contract No. NAS 7-100, sponsored by theNationazil Aeronautics and Space Administrationi.The systemns and techniques described sere de%eloped by the

cilTrts not only of the author but of many other meimibers of theJPL Telecommunjicationls Division.

REFEREN(ES1. "System Capabilities and Development Schedule of the DeepSpace Instrumenitationi Facility," Tech. Memo. 33-83, Jet Pro-pulsion Laboratory, Pasadena, Calif., Apr. 24, 1964.2. JPL Statl.UMariner: Mission to Vemn,is. New York: McGraw-Hill Book Co., lnc., 1963.3. BrNden, J. N., *'Mariner (Venus '62) Flight TelecommunicaltioiSystem," Tech. Rept. 32-377, JPL, Jan. 15, 1963.4. Space Programs Summary No. 37-23, sol. 11, JPL, Sept. 30,1963.5. Space Programs Summinairy No. 37-24, vol. VI, JPL, Dec. 31,1963, p. 71.6. Martin, B. D., "The Mariner Planetary Communication SystemDesign," Tech. Rept. 32-85 (Rev. 1), JPL, Mav 15, 1961.7. Springett, J. C., "Commanid Techiliques for the Remote Con-trol of Interplanetary Spacecraft," Tech. Rept. 32-314, JPL, Aug.1, 1962.8. Springett, J. C., "Pscudo Random Coding for Bit and WorkSynchronizationl of PSK Data Transmissionl SN stems," presentedat Internat'l Telemeterinig Conf., Londonl, Engiand, 1963.9. Mathison, R. P., "Tracking Techisiques for Interplanetary Space-craft," Tech. Rept. 32-284, JPL, Aug. 1. 1962.10. Mathisoni, R. P., "Spacecraft Telecommunications SNsSternDesign," 1963 Yearbook WGLR Coa/l., Munich, Germany.

84Mathisoni -Mariner Mars 1964 telemnetry and comma:nd system

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