Communication system and operation for lunar probes under lunar surface

Communication System andOperation for Lunar ProbesUnder Lunar Surface

T. MIZUNO

H. SAITO, Senior Member, IEEE

M. ICHIKAWAInstitute of Space and Astronautical ScienceJapan

In the Japanese LUNAR-A mission, penetrators will be

deployed to the moon for global seismic measurements. The

unique communication system between the subsurface probes

under the lunar surface and the lunar orbiter is described.

Radio wave propagation through a crater which is formed at the

penetration is investigated by means of scaled measurements in

a simulating environment. Acquisition and tracking sequence is

optimized within limited power capacity of the probe to maximize

contact time between the probe and the spacecraft.

Manuscript received March 16, 1998; revised April 8, 1999;released for publication August 30, 1999.

Refereeing of this contribution was handled by T. Busch/V. Skormin.

IEEE Log No. T-AES/36/1/02592.

Authors’ address: Institute of Space and Astronautical Science,3-1-1 Yoshinodai, Sagamihara, Kanagawa 229, Japan.

0018-9251/00/$10.00 c° 2000 IEEE

I. INTRODUCTION

The United States and Russia had frequent surveysof the moon between the 1950s and the 1970s. TheApollo project performed seismic measurementswith a small network. However, global seismicmeasurements of the moon require different types ofprobe technology than soft landing, the cost of whichwould be enormous for multipoint explorations.Penetrating probes to the surface of the moon

and planets are suitable technology for multipointexplorations. The Institute of Space and AstronauticalScience (ISAS), Japan, will perform Lunar Penetratormission, Lunar-A in 1999 [1]. The scientific purposesof the mission are global seismic measurement andheat flow measurement. The probe of Lunar-A is14 cm in diameter, 90 cm in length, and 13 kg inweight. The probe will be released from a lunarorbiter and deorbit by means of a small solid-motor.Then the probe will hit the lunar surface and penetrateinto the lunar regolith. Similarly, in the Deep Space 2mission by the United States [2], a micro probe willbe deployed and penetrate to the Martian regolith in1999. The scientific purpose of the micro probe is todetect water in the martian regolith. The micro probeis 4 cm in diameter, about 20 cm in length, and 1.5 kgin weight.This paper describes the communication system

between the probe under the lunar surface andthe mother spacecraft in Lunar-A mission. Thetechnical difficulty consists of the following twopoints: communication with the probe under thelunar surface, and acquisition and tracking of theprobe from the mother spacecraft in a short visibletime within the limited power capacity of the probe.Section II summarizes the mission outline. Thecommunication system and development of antennaof the probe are described in Section III and IV. Thenlink performance and acquisition method are outlinedin Section V.

II. MISSION OUTLINE

A spacecraft for the Japanese lunar missioncalled Lunar-A will be launched in August of 1999from Kagoshima Space Center of ISAS using theM-V launch vehicle. The M-V launch vehicle cansend a spacecraft of 540 kg into a lunar transferorbit. The mother spacecraft, which is shown inFig. 1, is designed to be spin-stabilized. It is 2.2 m inmaximum diameter and 1.7 m in height. The attitudeand spin rate of the spacecraft is controlled by N2H4mono-propellant reaction control system (RCS), whileorbital maneuvering near the moon is accomplishedby a bipropellant (N2O4 and N2H4) 500N engine.The outline of the Lunar-A mission is described

schematically in Fig. 2. The spacecraft will be firstinserted to an elliptic Earth orbit of apogee radius

IEEE TRANSACTIONS ON AEROSPACE AND ELECTRONIC SYSTEMS VOL. 36, NO. 1 JANUARY 2000 151

Fig. 1. Configuration of Lunar-A spacecraft.

4£ 105 km in August of 1999. Fig. 3 shows thetrajectory of the spacecraft from the Earth departureto the lunar orbit insertion (LOI). After severalresolutions, the spacecraft will swing by the moon(SWB) in October of 1999 for sake of lunar-gravityassist [3—5] to be inserted to a larger Earth orbitof apogee radius 1£ 106 km. The spacecraft willapproach the moon with a small relative velocity.On Feb. 11, 2000 the spacecraft will be insertedinto an elliptic lunar orbit (LOI) by the bipropellantengine. The apolune altitude is about 650 km and theperilune altitude is about 200 km. The gravity-assistmaneuvering of the spacecraft significantly reducesthe amount of propellant required to insert thespacecraft into the lunar orbit. However, it will takea half year from the launch to the LOI. Then thespacecraft will lower the perilune to about 45 kmfrom the lunar surface. Two probes will be deployedone by one from the mother spacecraft. It will requireabout one month to deploy two probes.One probe will be deployed on the near-side of

the moon on Feb. 27, 2000. The site for the firstprobe at the near-side is at the Oceanus Procellarum(336:5± longitude, ¡2:7± latitude). This site is near theApollo 12 site since the Lunar-A data is comparedwith the Apollo seismic network data. The spacecraftwill wait for a half month between the first and thesecond deployment while the moon rotates by a halfrevolution. The far-side probe will be deployed at aposition near the antipodal point of the Apollo 12 site

Fig. 2. Lunar-A mission sequence.

Fig. 3. Trajectory of Lunar-A in Sun-Earth fixed coordinate.Spacecraft will depart Earth in August of 1999 and then performfirst lunar SWB in October of 1999. LOI will be in February

of 2000.

(Mendeleev Crater, 141:7± longitude, 2:4± latitude)on Mar. 13, 2000. After deploying the probes, thespacecraft will be injected into a nearly circular orbitof about 220 km altitude. The inclination and the orbitperiod are 20± and 130 min, respectively.The penetrating probe is a missile-shaped

instrument carrier, which is about 14 cm in diameter,90 cm in length and 13 kg in weight. Fig. 4 describesthe sequence of the deployment [1]. The probe willbe separated from the spacecraft when true anomalyof the spacecraft is about 90± behind the position ofdeployment on the moon. When the probe approachesthe point above the deployment site, the probe willignite a small solid-motor to cancel the oribitalvelocity. The body axis of the probe is parallel tothe lunar surface. The probe will fall onto the lunarsurface. During the free fall descent, the probe willbe reoriented to become vertical to the lunar surfaceusing a cold gas side-jet. This control is based upon

152 IEEE TRANSACTIONS ON AEROSPACE AND ELECTRONIC SYSTEMS VOL. 36, NO. 1 JANUARY 2000

Fig. 4. Sequence of probe deployment.

Fig. 5. (a) Module of probe with deorbit motor and attitudecontroller. (b) Penetrating probe.

Rhumb-line attitude maneuver. A sun sensor is usedas an attitude sensor to detect the attitude maneuver.After the attitude maneuver a residual notation will becanceled by a cold gas side-jet and an accelerometer.The deorbit motor and the attitude controller attachedto the probe will be jettisoned just before the probehits the lunar surface. The detailed sequence of thedeployment is presented in [1]. Fig. 5 shows the probeand the probe module with the deorbit solid-motorand the cold gas side-jet.The final impact velocity of the probe will be

about 300 m/s. It is known that the lunar surface iscovered with lunar regolith of about 10 m thickness[6]. According to numerous impact tests [7] using

model probes and simulating targets of the lunarregolith, a probe is predicted to penetrate the lunarregolith to depth of about 1 m. It will encountera shock of about 10,000G at impact on the lunarsurface.The surface temperature of the moon varies

between 140±C at daytime and ¡140±C at night. Thepenetration depth is important to stabilize temperatureof instruments in the probe. According to the resultsof the Apollo heat flow experiment [8], an insulatingregolith blanket of only 30 cm thickness is sufficientto mitigate temperature of the regolith within ¡20±C§3±C. Thus instruments of the probe at depth of 1 m donot require any temperature control.The data gathered by the scientific instruments

will be compressed and stored in a recorder in theprobe. The probe will transmit the data to the motherspacecraft, when the spacecraft flies above the probe.A UHF (f = 400=450 MHz) band will be used forthe communication between the probe and the motherspacecraft.Communication between the spacecraft and the

Earth station will use S-band (f = 2:0=2:2 GHz). TheUsuda Deep Space Center, Japan (64 m antenna) willbe used as the ground station to perform telemetry,commanding, ranging, and Doppler measurementfor a deep space mission. The telemetry data rateis typically 8 kbit/s and the command data rate is4 kbit/s.

III. COMMUNICATION SYSTEM OF PROBE

A. System Selection

The surface temperature of the moon variesbetween 140±C and ¡140±C. The probes in theLunar-A have to be placed under the lunar surfacefor a mild thermal environment. The impact velocityof 300 m/s causes a shock of about 10,000 G at thepenetration. Propagation loss of radio wave is very

MIZUNO ET AL.: COMMUNICATION SYSTEM AND OPERATION FOR LUNAR PROBES UNDER LUNAR SURFACE 153

TABLE ICommunication Between Probes and Mother Spacecraft

Forward Link Return Link

Frequency [MHz] 450 400.877Transmitting Power [dBm] 36.4 30.6Transmitting Antenna Gain [dBi] ¡1:0+ ¡13¤Orbit Altitude [km] 200—240Receiving Antenna Gain [dBi] ¡13¤ ¡1:0+Modulation FSK-PM PCM-PSK-PMSubcarrier Frequency [kHz] 1.4/1.7 (tone) 16.384Modulation Index [rad] 0.67 0.67Coding – Convolutional Code (k = 7, r = 1=2)Bit Rate [bps] 50 128—1024

Note: *Antenna coverage of probe is in Table III.+Antenna coverage of spacecraft is 50± half-angle cones toward spin and antispin axis.

low in the lunar regolith. We select the system ofa subsurface antenna, which is equipped in the tailpart of the probe to transmit and receive radio wavethrough the lunar regolith.Velocity vector of the probe at penetration is

vertical to the average surface. Attitude of the probeat rest in the regolith after impact is affected seriouslyby the attack angle in motion, which is defined as anangle between the body axis and the velocity vectorprior to impact. Control accuracy of the attack angle isexpected to be as good as 8±. This 8± deviation of theattack angle causes the deviation of the rest-attitudeas large as 60± from the vertical direction. Therefore,the antenna of the probe has to be an omni-directionalantenna with about 60± half-angle cone in the regolithto prepare for any possible rest-attitude.The mother spacecraft is a simple spinner with dry

mass of 225 kg. The spin axis is nominally toward thesun direction to generate enough power by its solarpanels. The probe in the moon may appear in anyline-of-sight direction with respect to the spacecraftbody due to the spin motion and the orbital motion.It is not practical to install a high or a medium gainantenna with despun or tracking functions. Theantennas in the mother spacecraft to communicatewith the probes are low gain antennas. Low gainantennas (UANT-A, B) provide gain of ¡1 dB forcoverage of 50± half-angle cones toward the spin axisand the antispin axis.The communication between the mother spacecraft

and the probe is the communication link between twolow gain antennas. Such communication link shoulduse as low a frequency band as possible to maximizethe link capability. However lower frequency bandrequires a larger antenna which may not be installedinside the probe. The selected band of frequency forcommunication between the probe and the motherspacecraft is UHF band. The frequency for forwardlink (command) is 450 MHz and the frequency forreturn link (telemetry) is 400 MHz. The forward linkand the return link utilize righthand and lefthandcircular-polarized waves, respectively.

The communication system between the probeand the mother spacecraft is summarized in Table I.The return link signal (telemetry) from the probe isencoded with convolutional code (r = 1=2, k = 7) andis decoded by a Viterbi decoder (field programmablegate array) in the mother spacecraft. The return link(telemetry) data rate varies from 128 to 1024 bit/s.The forward link (command) data rate is 50 bit/s.The lunar global seismic measurement, which

is the main scientific purpose of Lunar-A, requiresdetermination of the positions of probes in themoon with accuracy of about 2 km. Measurementof two-way Doppler frequency is utilized todetermine the positions of probes. The probe carriesa transponder for this purpose. The mother spacecrafthas onboard a sweep oscillator for link acquisition anda stable oscillator for two-way Doppler measurement.

B. Antenna of Probe

The antenna of the probe transmits and receivesthe radio wave through the regolith. Dielectricproperty of the lunar regolith is one of the mostimportant factors for the antenna functions. TheApollo 11—17 missions brought many samplesof regolith back to the Earth. The intensivemeasurements of lunar soils and rocks were performedby many researchers [6, 9, 10]. Reference [6] showsa good summary of the measurements at frequency of450 MHz, which is the frequency band of Lunar-Amission. The relative dielectric permittivity of lunarsoils and rocks is strongly dependent upon densityand independent of chemistry; above 1 MHz, it isalso independent of frequency and temperature range.The loss tangent (dielectric dissipation factor, tan±)is strongly dependent upon density and content of(TiO2 +FeO). Average density of the regolith fromthe antenna to the surface is estimated to be 1:6»1:85 g/cm3, taking into account the density profile [6]and formation of a crater [7]. Content of (TiO2 +FeO)at the deployment site is estimated to be less than 30%[6]. The series of measurement at 450 MHz [6, 9, 10]


Fig. 6. Antenna structure of probe.

give the relative dielectric permittivity of 2:9§ 0:2 andthe loss tangent of 0:0» 0:02.Numerical simulations predict that the probe forms

an impact crater which is part of a spherical surface.Typically, the diameter is 1.65 m and the depth is0.47 m. The size of the crater is comparable withthe wavelength of UHF band (wavelength 0.75 mfor 400 MHz) and influences the radiation patternpropagating through the crater. The effects of thecrater are discussed in Section IV.The structure of the antenna is shown in Fig. 6.

The antenna is a cross-dipole antenna filled withlow loss GFRP material to provide enough stiffnessfor 10,000 G shock. Equivalent axial compressionforce of 50 T and shearing force of 27.8 T areconsidered at the structural design. The antennaelements are surrounded with GFRP (permittivity3.4) and the antenna as a whole is expected to bein the lunar regolith (permittivity 2:9§ 0:2). Halfwavelength at 400 MHz frequency in the GFRPmaterial is about 22 cm, which is larger than diameter14 cm of the probe. The antenna elements are benttwice to increase the element length for purposeof impedance matching. The antenna of the probesatisfies impedance matching in the wide frequencyrange between 400 MHz and 450 MHz, when it is inthe lunar regolith of permittivity 2:9§ 0:2.

C. Electronics Equipments of Probe

The onboard equipments of the probe are requiredto be light in weight, small in size, and consumeless power. Recent microelectronics technologies inthe field of cellular phone and multimedia would beuseful for space microelectronics. These technologiesinclude field programmable gate array, surface mounttechnology, and low power devices of 3 V operation.The environmental conditions of the probes

are unique compared with common satellites. The

Fig. 7. Simulating environment for antenna of probe in lunarregolith.

temperature in the regolith is constant, while satellitesexperience eclipse and sunshine. Radiation conditionunder the regolith is mild because of protection bythe regolith layer. There are experimental evidencesthat plastic packaging is more robust for extremelyhigh impact environment than ceramic packaging orCAN-type packaging.We apply these advanced, non-space technologies

to the probe after evaluation tests includingtemperature test, thermal cycle test, burn-in test,structure analysis, radiation test, and the penetrationtest [11]. All electronics equipments in the probe arefilled with epoxy resin including glass microballoon tobe robust against the extremely high impact.

IV. ANTENNA PERFORMANCE IN LUNAR REGOLITH

A. Simulating Environment for Antenna in LunarRegolith

A test facility which simulates wave radiationfrom an antenna in the lunar regolith is requiredto develop the probes. As described in SectionIIIB, dielectric permittivity of the lunar regolith in450 MHz is 2:9§0:2. Silica sand (SiO2) of dielectricpermittivity 2.4 and ceramic powder (Kyocera F1120)of dielectric permittivity 2.9 were selected for thesimulating regolith. Although F1120 powder has theequal permittivity to the lunar regolith, only a smallamount of F1120 ceramic powder can be produced.Ceramic powder F1120 is used only at adjustment forimpedance matching of the antenna.A simulating environment for radio wave emitted

from the lunar regolith was constructed at the beachin the Noshiro Testing Center, Japan. A pit (10 m£10 m, 2 m depth) was dug at the beach and was filledwith dry silica sand of 70 T as shown in Fig. 7. Apressurized plastic air dome of 10 m in radius coversthe test site to keep the silica sand dry.Undesired wave reflections from boundaries of the

facility must not interfere radiation from the antennaof the probe. The air dome is made by plastic sheetof 1 mm in thickness. No standing wave is observed


Fig. 8. Configuration of probe under crater. Dotted lines areschematic drawing of ray paths from antenna, includingtransmitting ray through surface, reflecting ray at surface,

multiple-reflecting ray in crater.

in the air dome. The other boundaries are side andbottom surfaces of the silica sand region. The sidesurfaces are layers of sand bags filled with dry beachsand. The side surfaces are not parallel to each otherand are diffusive to eliminate effects of cavity. Threematching layers of 10 cm in thickness are installedat the bottom surface of the silica sand in order toeliminate wave reflection from the bottom surface.Each layer which is separated by a plastic sheetcontains a mixture of dry silica and dry beach sand.The mixture ratio of each layer gradually changeslayer by layer from pure silica to the dry beach sand.The reflection from the matching layers is measuredto be 30 dB lower than one of metal reflection. Nostanding wave is observed in the region of the silicasand. From these experimental results, this test sitecan be considered a simulating environment forantennas in the lunar regolith.Far field patterns of antennas are required to

measure for the Lunar-A mission. Let us assume atransmitting antenna of diameter D and a receivingantenna of diameter d. The centers of two antennasare separated by distance r. Condition for region offar field radiation [12] is given as

r > 2(D+ d)2=¸ (1)

where ¸ is the wavelength in vacuum. When thecondition (1) of far field measurement is satisfied, theerror of measurement is less than 0.2 dB.Fig. 8 shows the configuration of the probe under

a crater. Radiation waves which propagate directlyfrom the antenna to flat surface outside the cratercannot be emitted above the surface, since totalreflection takes place at the surface. The critical angleof total reflection in the silica sand is 40± for incidentangle to the surface. Numerical simulations [13]indicate that multiple reflections take place insidethe crater. The aperture of the crater is considered

an effective antenna. If full scale configuration ofthe antenna is assumed, the aperture of an effectiveantenna is a diameter of the crater (D = 1:65 m) and¸= 0:67 m for 450 MHz. Diameter of a receivingantenna would be about 0.6 m. Then the conditionof far field radiation (1) is r > 15:1 m, which is notsatisfied by the air dome of 10 m in radius. Scaledmeasurements are required for far field radiationsat the air dome. We select one-third scale, in whichD = 0:55 m, ¸= 0:23 m and frequency of 1.3 GHz.The receiving antenna for measurement is a helicalantenna of diameter d = 0:22 m. Then the conditionof far field radiation is r > 5:2 m at one-third scaleconfiguration. As described in the next section, theactual distance r between the crater and the receivingantenna is 5.6 m. Far field measurement at this testsite is valid when one-third scale model is applied.

B. Radiation Pattern Scattered by Crater

At penetration to the lunar surface the probeforms a crater of 1.65 m in diameter and 0.47 min depth, and it rests at depth of about 1 m fromflat regolith surface to the tail. Rest-attitude of theprobe deviates inside 60± half-angle cone from thevertical line depending on the attack angle prior to thepenetration. This rest-attitude influences performanceof the antenna.Antenna models of one-third scale at frequency

1.3 GHz band are utilized to measure the farfield pattern with a crater. In the one-third scaleexperiments, the crater is simulated by portion of asphere made from styrene foam. Radio wave fromantenna of the probe is scattered by the crater and isradiated to free space. The aperture of the crater withthe subsurface probe can be considered an effectivetransmitting antenna. A receiving helical antennaof 1.3 GHz is installed at a wood frame, which isrotatable in radius of 5.6 m for 0±—180± in elevationto measure the far field pattern. We measure theabsolute value of the receiving power and determinedthe absolute gain pattern from the crater with thesubsurface antenna.The following results (Figs. 9—11) are expressed

in terms of the full-scale antenna in the lunar regolithtaking into account a slight difference of the boresightgain between the one-third model and the full-scalemodel. The numerical simulations [13] show that theantenna gain in the lunar regolith (permittivity 2.9) is2 dB smaller than one in silica sand (permittivity 2.4).We substract 2 dB from the experimental gain in silicasand to take into account the difference in permittivity.The rest-attitude and the position of the probe are

described in Fig. 8 in terms of the angle ³ between thebody axis and the vertical line, the depth v from theflat surface to the antenna, and the horizontal distanceh from the center of the crater to the antenna. If thefalling probe has an attack angle prior to the impact,


Fig. 9. Antenna gain of probe for vertical rest-attitude (³ = 0±,v = 90 cm, b = 0 cm) with crater (solid line) and with flat surface

(broken line).

Fig. 10. Antenna gain of probe for titled rest-attitude with crater(³ = 30±, v = 90 cm, h= 15 cm).

Fig. 11. Antenna gain of probe for tilted rest-attitude with crater(³ = 60±, v = 60 cm, h= 30 cm).

then in the regolith it turns aside in the horizontaldirection by “rudder effect” and rests at the tiltedattitude. The far field pattern measurements areperformed as functions of elevation angle µ in thevertical plane including the body axis, as well as inthe other vertical plane that is perpendicular to theformer.Scattering effects by a crater are examined for an

ideal case where the probe hits the surface with noattack angle and rests vertically in the regolith (³ = 0±,v = 90 cm, h= 0 cm). The v and h are presented interms of full scale. Fig. 9 shows the far field patternwith a crater and without a crater (a flat surface).While there is no degradation due to the crater atthe zenith gain (µ = 90±), the gain off the zenith(µ < 70±, µ > 110±) decreases due to the crater effects.The craters are considered concave lenses taking

dielectric permittivity of the regolith into account. Theoff-axis ray is diverged after it passes the concavelens, while on-axis ray intensity remains the same.Based on Snell’s law where a plane wave propagates

from the regolith to the vacuum through the boundary,total reflection takes place for the case where anincident angle at the surface is larger than the criticalangle of 40±. The concave surface of the craterincreases incident angles for the waves, and decreasesthe total wave energy radiating out of the surface.Effects of the tilted rest-attitude of the probe are

shown in Fig. 10 (³ = 30±, v = 90 cm, h= 15 cm),and Fig. 11 (³ = 60±, v = 60 cm, h= 30 cm). Thesolid line indicates the pattern in the vertical planeincluding the body axis. The peak of the pattern shiftsfrom the zenith toward the tilt direction of the probe,and the antenna gain is degraded. On-axis part of thepattern from the antenna reflects at the crater surfacebecause of total reflection, and only off-axis partscan radiate through the crater surface to free space.The numerical analyses [13] indicate that multiplereflections inside the crater also influence the radiationpattern especially for the cases of tilted rest-attitude.

V. LINK PERFORMANCE AND ACQUISITIONPROCESS

A. Link Budget

After all the probes are deployed to the moon, themother spacecraft executes orbit maneuvers into thedata relay orbit, the altitude of which is optimized interms of data relay communication for the probes. Anorbit of a lower altitude than 100 km may be unstabledue to the gravity anomaly. The link margin becomesinsufficient for an orbit of a higher altitude than350 km. In general, an orbit of lower altitude provideshigher bit rate but shorter visible (contact) time. Theinitial 20—30 s of visible time is devoted to executecommunication handshakes. On the other hand,an orbit of higher altitude provides longer visibletime but lower bit rate. The nearly circular orbit of220 km altitude is found to maximize communicationcapability.As the spacecraft passes above the probe, the

slant range and the antenna gain of the probe varyas functions of elevation angle µ from the probe tothe spacecraft. The antenna pattern of the probe isdescribed in Figs. 9—11. The link budget is calculatedfor various elevation angles µ. Table II is an exampleof link budget, where tilt of the probe ³ = 60±,elevation angle µ = 45±, orbit altitude of 240 km,and return link bit rate of 256 bit/s. Acquisition ofcommunication link starts from the forward link. Theforward link is designed to have larger margin thanthe return link. In the case of Table II, the forwardlink has a margin of 8.0 dB which is determined bythe carrier channel. The return link has a margin of4.3 dB which is determined by the data channel.Table III shows the antenna coverage of the probe

which gives appropriate communication link margin(5 dB return link margin for bit rate 256 bit/s and


TABLE IIExample of Link Computation

Forward ReturnLink Link

Transmitting RF PerformanceTransmitter Power dBm 36.4 30.6Transmitting Feeder Loss dB ¡2:1 ¡1:4Transmitting Antenna Gain dBi ¡1:0 ¡13:3Transmitting EIRP dBm 33.3 15.9

RF Loss in PropagationPolarization Loss dB ¡1:3 ¡1:3Free Space Loss dB ¡135:7 ¡134:7Dielectric Loss of Regolith dB ¡0:8 ¡0:7

Receiving RF PerformanceReceiving Antenna Gain dBi ¡13:3 ¡1:0Receiving Feeder Loss dB ¡1:7 ¡3:1Received Power dBm ¡119:5 ¡124:9Noise Spectral Density No dBm/Hz ¡172:1 ¡172:0Received Power/No dB-Hz 52.6 47.1Total Rcvd Power/No dB-Hz 52.6 45.9

Receiving Carrier ChannelModulation Loss dB ¡1:2 ¡1:2Receiver Hardware Loss dB 0.0 0.0Noise Band Width dB-Hz 27.0 27.8Required Eb/No dB 16.4 12.0Margin dB 8.0 4.9

Receiving Data ChannelModulation Loss dB ¡7:8 ¡7:8Receiver Hardware Loss dB ¡2:5 ¡5:2Noise Band Width dB-Hz 17.0 24.1Required Eb/No dB 13.4 4.5Margin dB 11.9 4.3

Note: Orbit altitude 240 km, elevation µ = 45±, rest-attitude ofprobe ³ = 60±, bit rate of return link 256 bit/s.

TABLE IIIAntenna Coverage Probe

Rest-Attitude (³) Antenna Coverage*

0± (vertical) Elevation 90± § 29± Cone30± Elevation 80± § 24± Cone60± Elevation 65± § 18± Cone

Note: *5 dB return link margin for 256 bit/s. Orbit altitude is240 km.

orbit altitude 240 km). As the probe tilts (³ = 30±

or 60±), the antenna coverage of the probe tilts andbecomes narrower. This is because the antenna gain ofthe probe decreases and the slant range increases.

B. Link Visibility

When the mother spacecraft executes acquisitionof the probes and data relay, the spacecraft is in anearly circular orbit of about 220 km altitude. Theperiod and the inclination of the orbit are 130 minand 20±, respectively. The moon revolves on its ownaxis with a period of 27.3 days. A visible pass isdefined as a time duration in which the spacecraft

Fig. 12. Link visibility in elevation-azimuth diagram from probe.Probe located at latitude ¡2:7±, longitude 336:5± (Oceanus

Procellarum). Antenna coverage of probe for rest-attitude ³ = 0±,30± and 60± are shown by circles. Series of trajectories in visiblepass group of May 29—31, 2000 are shown by lines. Solid linescorrespond to region where spacecraft watches position of probewithin antenna coverage of spacecraft. Broken lines correspond toregion where probe is out of sight from antenna coverage ofspacecraft. Solid marks indicate position of spacecraft at

every 60 s.

flies above the probe and can communicate with theprobe. The next visible pass may occur in 130 min.There are a number of sequential visible passes. Thenvisible passes are terminated since the moon rotateson its axis. We define these sequential passes as avisible pass group. Duration of a visible pass andthe number of passes in a visible pass group dependupon the rest-attitude of the probe and the attitudeof the mother spacecraft. Typically, the duration ofa visible pass is about 1—3 min, and a visible passgroup consists of about 10—30 visible passes. The nextvisible pass group occurs in 14 days (a half period ofthe lunar rotation).Fig. 12 is the elevation-azimuth diagram which

indicates direction of line-of-sight from the probe tothe mother spacecraft. The origin of Fig. 12 indicatesthe zenith (elevation 90±). The abscissa is east—westdirection and the ordinate is north—south direction.The probe is located at latitude ¡2:7±, longitude336:5± (Oceanus Procellarum). The series of thetrajectories of the spacecraft in one visible pass groupduring May 29—31, 2000 are shown in Fig. 12. Thetrajectory of the spacecraft is a circular orbit with220 km altitude and 20± inclination. The intervalbetween each pass is 130 min (a period of the orbit),and 30 sequential passes are depicted in Fig. 12.The solid marks on the trajectory indicate positionsof the mother spacecraft at every 60 s. The motherspacecraft passes along the trajectory from south—westto north—east with velocity of about 30 s per 10 deg.Solid lines of the trajectory correspond to the regionwhere the spacecraft watches the position of the probewithin the antenna coverage of the spacecraft. Brokenlines are the region where the probe is out of sightfrom the antenna coverage of the spacecraft. Thespacecraft is in sun-pointing attitude mode to obtainthe maximum power.


Also the antenna coverages of the probe are shownby circles in Fig. 12. The antenna coverages of theprobe are indicated for the following three cases asexamples. In the first case of ³ = 0± the rest-attitude ofthe probe is vertical (³ = 0±) and the antenna coverageis a cone of elevation 90± §29± (see Table III). In thesecond case of ³ = 30±, tail of the probe is tilted with³ = 30± in west direction and the antenna coverageis a cone of elevation 80± §24±. In the third caseof ³ = 60±, tail of the probe is tilted with ³ = 60± inwest direction and the antenna coverage is a cone ofelevation 65± § 18±.To satisfy visibility for communication, antenna

coverage of the spacecraft needs to cover the positionof the probe as indicated by the solid lines of thetrajectory. Also position of the mother spacecraftneeds to be within antenna coverage of the probe asshown by the circles. For the vertical rest-attitude ³ =0±, one visible pass group consists of about 14 passeswhich is longer than 60 s. The maximum visible timeis 170 s. For the maximum tilt of the probe ³ = 60±,one visible pass group consists of about 8 passeswhich is longer than 60 s. The maximum visible timeis 100 s.The antenna coverage of the mother spacecraft

is 50± half-angle cones toward the spin and antispinaxis. The visibility terminates when the positionof the probe is outside the antenna coverage ofthe spacecraft. In the case of Fig. 12, the motherspacecraft is in the nominal sun-pointing attitude. Toavoid termination of the visibility due to the spacecraftantenna coverage, we plan in the mission operation tooptimize the spacecraft attitude for each visible passgroup on condition that the sun aspect angle is lessthan 45±.

C. Acquisition of Probe

At the initial acquisition phase, the motherspacecraft has to search the probe in the moonthrough the communication link. The position of theprobe deployed to the lunar surface can be controlledwith site dispersion of 20 km. This uncertainty ofthe probe position corresponds to 6± half-angle cone,when the spacecraft searches the probe from 220 kmaltitude. At the initial acquisition phase, rest-attitudeof the probe is unknown and uncertainty of antennacoverage of the probe is 43± half-angle cone fromthe zenith (Table III). Effect of the rest-attitudedominates uncertainty of the link visibility in theinitial acquisition.The mother spacecraft has to search the probe by

transmitting the forward link continuously when thespacecraft flies above the probe. Before the probe isseparated from the mother spacecraft, the predictedstarting time of the first visible pass group is recordedin the onboard instrument of the probe. The probeswitches on its receiver prior to this predicted time.

Fig. 13. Process of communication between spacecraft and probe.

The receiver of the probe switches on in the waythat switch-on duration covers uncertainty of startingtime of the visible pass. This uncertainty is estimatedto be several tens of minutes, which is determinedmainly by the orbit maneuvering error and the timeinterval between the probe deployment and the firstcontact (one month). Once the communication linkis established, the next visible pass in 130 min canbe predicted very accurately. Receiving power ofthe return link is measured at the spacecraft duringthe link contact. The antenna pattern of the probe isestimated pass by pass along the trajectory.The second visible pass group takes place in 14

days. There is no orbit maneuvering between visiblepass groups. The determination error of orbit period isabout §0:05%. The starting time of the second visiblepass group can be predicted at the first visible passgroup with error of §10 min. This predicted time isrecorded in the onboard instrument of the probe at thefirst visible pass group. The acquisition for the secondvisible pass group is executed in the similar way tothe first visible pass group. The antenna pattern of theprobe which is measured at the first visible pass groupis utilized to ensure the link acquisition at the secondvisible pass group.

D. Process of Communication

Let us describe the process of communication ateach visible pass. The probes depend upon primarylithium batteries (6.8 V, 500 WH) as their powersource for 12 months operation. For the sake ofpower saving, the transmitter of the probe mustemit radio wave for limited period when the motherspacecraft can receive the return link. Handshakeprocess between the spacecraft and the probe isapplied to ensure the communication link as shownin Fig. 13. The spacecraft sends signals to the probeand then the probe replies to the spacecraft. First,handshake at the level of carrier wave is established


between the transmitters and the receivers (step 1—6).Then handshake at the level of data channel isestablished between the hardware (modulators anddemodulator/decoder) and between the processors(step 7—12). After the communication link isestablished, the data protocol of HDLC (high leveldata link control procedure) is applied to ensure thelink quality (step 13).The process of communication is as follows.

1) The mother spacecraft transmits continuouslyforward link without modulation. The frequency offorward link is swept in order that the receiver of theprobe locks on the frequency. Typical values of thesweep range and the sweep period Tsweep are 42 kHzand 10 s, respectively.2) The receiver in the probe is switched on by the

internal timer prior to the predicted time when themother spacecraft is expected to be inside the regionof visibility. The receiver continues to be on for 30 sand be off for next 30 s.3) The spacecraft enters into the region of

visibility. Within a period Tsweep(» 10 s) of thefrequency sweep, the receiver of the probe locks onthe forward link.4) The transmitter of the probe is switched on

to transmit the return link without modulation. Ittakes 2 s to stabilize the frequency. The frequencyof the return link is swept in accordance with theforward link, since the carrier wave of the return linkis coherent to the forward link.5) The receiver of the mother spacecraft locks on

the return link within a time of (3/4) Tsweep(» 8 s).6) The radio wave source of the forward link in

the spacecraft switches from the sweeping oscillatorto the stable oscillator with stability of 10¡7 in orderto measure two-way Doppler frequency to determinethe position of the probe. Handshake process isestablished at the level of carrier wave.7) Modulation of the forward link starts at the

spacecraft. Preamble signal of 50 bits and the signalof set response mode (SRM; 48 bits) are transmittedto the probe. The signal of SRM is a command thatsets parameters of the return link such as the bit rate.8) Command decoder (CMD) hardware at the

probe receives the preamble and then locks on. ThenSRM signal is recognized by the processor of theprobe.9) Modulation of the return link starts at the

probe. Preamble signal of 1024 bits is sent to thespacecraft, and then the signal of unnumberedacknowledgment (UA; 48 bits) is transmitted. Thesignal of UA is an acknowledgment of SRM.10) Demodulator (DEM) at the spacecraft receives

the preamble and then locks on. Then UA signalfrom the probe is recognized by the processor at thespacecraft. Handshake process is established at thelevel of data channel.

11) The spacecraft transmits commands to theprobe.12) The probe transmits telemetry to the

spacecraft.13) In steady state of the communication, the

data protocol of HDLC is applied to ensure the linkquality. Unit of data to transmit and receive betweenthe spacecraft and the probe is defined as a frame. InHDLC protocol a frame contains information on thelink control; the sequential number of the frame thatis transmitted at present and the sequential number ofthe frame that is expected to be received at next step.Based on these two sequential numbers, the spacecraft(or the probe) can check whether frames that thespacecraft (or the probe) had already transmittedhave been received properly by the probe (or thespacecraft). When some frames are found not to bereceived, the same frames are transmitted again.14) When handshakes at carrier wave or

data channel are discontinued, the process ofcommunication returns to a certain point to recoverthe handshakes. If these trials fail for a certain timeperiod, the transmitter and the receiver of the probeswitch off to terminate the process of communication.The starting time of the next visible pass is set to theinternal timer of the probe.

It takes 20—30 s to execute this handshake process(step 3—10) depending on the bit rate of the returnlink. Time indicated in Fig. 13 is for the case of256 bit/s return link. Most of the time is spent tosweep the forward link frequency in order thatreceivers of the probe and the spacecraft lock on. Therange of sweep must cover the Doppler frequency, andfrequency variation of the receiver due to temperaturevariation as well as the penetration shock. Thefrequency of the receiver in the probe is roughlyestimated by the spacecraft after several visiblepasses. Then the sweep range of the forward linkchanges to a narrow value to make the acquisitiontime shorter.

VI. CONCLUSIONS

The communication system with subsurface probesof the moon in Japanese Lunar-A mission is describedin this paper. Radio wave propagation through a craterwhich is formed at the penetration is investigatedby means of scaled measurements. Acquisition andtracking of the probe from the mother spacecraftare optimized within limited power capability of theprobe.The Institute of Space and Astronautical Science

and the National Space Development Agency, Japan,are planning Selene II mission by the Japanese H-2launch vehicle. In Selene II mission twelve subsurfaceprobes will be deployed to the moon around 2006.We will apply to Selene II mission the communication


system that is described in this paper. There wouldbe several points that can be improved in the systemof communication. Spacecraft in Selene II will beable to have a high or a medium gain antenna witha two-axes gimbals to track point of the probe in themoon. The transmitter of Lunar-A spacecraft has tosweep frequency of the forward link in order thatthe receivers of the probe and the spacecraft lock onthe frequency. In Selene II mission, receivers of theprobes and the spacecraft would be digital receiverswhich execute fast Fourier transform to lock on thefrequency. This would eliminate the sweeping time ofthe forward link. These improvements of the systemwould contribute effectively to enhance capability ofthe communication with subsurface probes in SeleneII mission.

ACKNOWLEDGMENT

We express special thanks to Mr. W. Kudo, staffof the Noshiro Testing Center for their contributionin our experiment, and the Lunar-A team engineers atNippon Electric Corporation.

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Takahide Mizuno received the B.S., M.S., and Ph.D. degrees in electricengineering from Yokohama National University, Yokohama, Japan, in 1988,1990, and 1993, respectively.Since 1993, he has been working as a research associate at the Institute of

Space and Astronautical Science, Sagamihara, Japan. He is working on thedevelopment of subsurface antenna for LUNAR-A. His current research includesfree electron laser experiments.

Hirobumi Saito (SM’80–M’81) received the B.S., M.S., and Ph.D. degrees inelectrical engineering from Tokyo University, Tokyo, Japan, in 1976, 1978, and1981, respectively.Since 1981, he has been at the Institute of Space and Astronautical Science,

Sagamihara, Japan, where he is a professor. He is working for the planetaryexploration projects including the lunar mission LUNAR-A and the asteroidsample return mission MUSES-C. His current research includes experiment offree electron laser, and miniaturization of advanced spacecraft. He was a visitingresearch scientist of Massachusetts Institute of Technology for 1984—1986, and ofJet Propulsion Laboratory, California Institute of Technology for 1991—1992.

Mitsuru Ichikawa received the B.S. of electrical engineering at Nihon University,Japan, in 1960.He worked at the Institute of Industrial Engineering, Tokyo University,

from 1960 to 1964. Since 1964 he has been working at the Institute of Spaceand Astronautical Science, Japan. His research includes on-board antennas ofsatellites, tracking radars for rocket launching, and large antennas of groundstations.


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Communication system and operation for lunar probes under lunar surface