Once every seven Martian days (sols), the spacecraft will approximately retrace its ground track. After each seven sol "repeat" cycle (88 orbits), the ground track pattern will be offset by 59 km eastward from the tracks on the previous cycle. Such a repeat cycle scheme will eventually allow 99.9% global coverage to be built up from repeated instrument swaths across the planet.
The figure on the opposite page shows a timeline of the mapping phase, including the Martian seasons and the expected data rate profile for return of the recorded science data. Under normal conditions, no major maneuvers will occur during this phase of the mission.
Specifically, the baseline mapping orbit chosen for the MGS mission will utilize a descending node orientation of 2:00 p.m., index altitude of 378 kilometers, near circular eccentricity of 0.007, inclination of 92.87 degrees, and argument of periapsis of -90o. This combination results in a frozen orbit with a period of approximately 117.65 minutes. In other words, the frozen orbit condition causes periapsis to always remain at -90o, a location over the Martian south pole.
GIF = 237 MBytes
Use of a 378 km index altitude for mapping operations satisfies the MOC low altitude requirement. Index altitude is a measure of the mean semi-major axis, and reflects the difference between the semi-major axis length and the equatorial radius of Mars (3397.2 km). This parameter is related to, but not the same as the average height of the spacecraft above the surface of Mars. The 378 km number was chosen in large part for its desirable properties with respect to ground track repeat cycles and spacing.
Setting the semi-major axis to a specific value constrains the valid choices for eccentricity and inclination in terms of the spacecraft maintaining a Sun-synchronous orientation. Although many possible solutions exist, MGS will fly in a 92.87° inclination orbit with an eccentricity of 0.007. These parameters satisfy the MOLA constraint for a near polar, near circular orbit with minimal altitude variations.
For the MGS frozen orbit, the range of areocentric altitudes over a single orbit will always remain between 354 km (south pole) to 410 km (north pole), satisfying the MOLA constraint for less than 105 km altitude variation. Without the frozen orbit, the variations in eccentricity and periapsis location will cause the altitude range to exceed the horizon sensor's (MHSA) operational limits of 335 km minimum to 455 km maximum. Such an excursion would also violate the MOLA's 105 km altitude variation constraint.
A frozen orbit with a +90° argument of periapsis (over the Martian north pole) is also mathematically possible. However, this option was not chosen because it would force the eccentricity to taken on a higher value than 0.007. Consequently, the orbit would appear less circular in shape and the altitude variations over the course of the orbit would correspondingly increase.
Figure 6-2: Mapping Orbit Ground Track Pattern for 7 Sol Repeat Cycle
Technically, the mapping orbit will be targeted for a Q value of 6917/550. This parameter denotes the number of revolutions completed by the spacecraft in one Martian day. Therefore, if the mapping orbit is perfectly maintained, the ground track pattern will repeat exactly after a "super-cycle" of 6917 orbits in exactly 550 sols. However, maintaining this pattern is not possible due to navigational uncertainties in controlling and predicting the orbit. Atmospheric drag represents the most difficult orbit perturbation to measure.
In practice, the Q value of the mapping orbit will be roughly 88/7. Since 88/7 represents only an approximate fractional reduction of 6917/550 (12.57143 vs. 12.57636), the ground track pattern will not repeat exactly on every cycle. Instead, after every 88 orbit repeat cycle, the ground track pattern for that cycle will appear offset slightly to the east of the pattern laid down by the previous cycle. This offset amounts to 58.6 km when measured at the equator. For example, the equatorial location of ground track number 89 will lie 58.6 km to the east of track number 1. This 58.6 km differential is called the "orbit- walk."
Keep in mind that this 58.6 km orbit-walk at the equator does not represent the spacing between ground tracks on two successive orbits. During each mapping orbit lasting 117.65 minutes, Mars spins under the orbit from west to east at a rate of approximately .24o/minute. This rate of planetary rotation causes the spacecraft to fly over locations about 28.62° to the west of the locations flown over on the previous orbit. At the equator, 28.62° translates to a ground track spacing of 1,697 km between two successive orbits. Ground tracks laid down by later orbits within the 88 orbit repeat cycle will gradually fill the 1697 km gap between two consecutive orbits.
The next table lists the orbital elements of the mapping orbit, both mean values and osculating at periapsis. All values are referenced to the Mars centered, Mars mean equator and IAU vector of epoch coordinate system, reference epoch 1 January 1998 at 0:00:00 ephemeris time (ET).
Orbital Element Mean Value Osculating at Periapsis Semi-Major Axis (a) 3775.144 km 3776.159 km Eccentricity (e) 0.00719 .00405 Inclination (i) 92.866° 92.869° RA of Asc. Node (Omega) Varies (see below) Varies (see below) Argument of Periapsis (omega) -90.0° -90.0°
Because MGS will circle the red planet in a Sun-synchronous orbit, the right ascension of the ascending node (Omega) will advance with time at a rate that matches the mean motion of Mars around the Sun. The expression below provides a way to calculate Omega at any arbitrary time during the mission.
Omega = -98.382323° + 0.5240429(t - to) - 2.644269x10-6(t - to)
In the previous expression for right ascension of the ascending node in the mapping orbit, t is in Earth days, to is 10 September 1997 at 0:00 ET, and Omega is in degrees referenced to the Mars mean equator, IAU vector coordinate system. Keep in mind that the 2:00 p.m. orientation is with respect to the position of the fictitious mean Sun, not the location of true Sun. With respect to the true Sun, the descending node time of the orbit will vary from about 1:08 p.m. to 2:41 p.m. depending on the Martin season and Mars' exact location in its orbit around the Sun.
Figure 6-3: Earth and Sun Relative Positions to Mapping Orbit
If an orbit appears "face-on" to the target (view from the target perpendicular to the plane of the orbit), the beta angle will measure 90° and no obscuration of the target occurs as seen from the spacecraft will occur. In other words, the spacecraft will spend no time "behind" Mars relative to the target. Beta values for no obscuration can fall lower than 90° with a high enough orbital altitude. If an orbit appears "edge-on" to the target (view from the target parallel to the plane of the orbit), the beta angle will measure 0° and the target will be obscured for a maximum about of time. As a general rule of thumb, longer obscurations will result when the target is closer in position to the mapping orbit plane than when farther away.
Because the mapping orbit lies fixed at a 2:00 p.m. orientation relative to the fictitious mean Sun at Mars, a 90° beta angle means that the target lies at either 8:00 a.m. (90° west of the orbit) or 8:00 p.m. (90° east of the orbit). On the other hand, a 0° beta angle indicates that the target's position in the sky also lies at 2:00 p.m. relative to the mean Sun at Mars. In general, this type of analogy requires the use of caution because 15° of beta angle does not always translate into one hour of mean solar time differential. The reason is that while the fictitious mean Sun always lies in the Martian equatorial plane, the target may not lie there.
During the mapping phase, the position of the Earth will change drastically with respect to the mapping orbit plane. Earth occultation durations will last for roughly 40 minutes out of every 117.65 minute orbit at the start of mapping in April 1998. At this time, the Earth's position will be at about 12:30 p.m. relative to the fictitious mean Sun at Mars, roughly 15° to the west of the mapping orbit plane.
Figure 6-4: Earth Occultation Durations During Mapping
As time progresses forward from the start of mapping, the Earth will gradually move eastward toward the mapping orbit plane, and occultation durations on every orbit will increase as the beta angle drops to 0° on 29 October 1998. On that date, the orbit plane will appear "edge-on" to the Earth, and occultation durations will reach a maximum of 41.3 minutes out of every revolution.
About three and one half months later on 19 February 1999, the Earth's position as seen from Mars will pass through the orbit plane yet another time, but this time on its way back west from the east. Again, occultation durations will reach a maximum due to the "edge-on" orbital geometry. Then, the occultations will slowly decrease in duration as the Earth continues to move westward, away from the mapping orbit plane.
Fourth months later on 19 June 1999, the Earth will reach a position nearly 70° to the west of the mapping orbit, roughly 9:05 a.m. relative to the fictitious mean Sun. This date is significant because despite not yet reaching a "face-on" orientation to Earth, the orbit lies at sufficiently high altitude that the spacecraft will no longer passes behind Mars relative to Earth. In other words, Earth occultations will no longer occur during any part of the mapping orbit, and a clear line of sight will always exist between the Earth and spacecraft.
The Earth's position as seen from Mars will continue to move westward until 22 September 1999 when it reaches an extreme of 8:50 a.m. relative to the fictitious mean Sun. Then, the Earth will begin to move back toward the east. Occultations will begin to occur again on 26 November 1999 and will increase in duration for the remainder of the mapping phase of the mission.
Figure 6-5: Solar Eclipse Durations During Mapping
Despite using a Sun-synchronous mapping orbit, the position between the MGS orbit plane and the Sun will vary as a function of time. Consequently, both the solar beta angle and the amount of time that Mars eclipses the Sun during an orbit will also vary. The reason is that the 2:00 p.m. Sun-synchronous orientation is defined with respect to the fictitious mean Sun.
In reality, Mars' obliquity and the eccentricity of Mars' orbit around the Sun cause the position of the true Sun to oscillate about the mean Sun. This position will vary depending on the Martian season and the location of Mars in its orbit around the Sun. During mapping phase, the true Sun's position will vary from a maximum of roughly 51 minutes ahead (east of) to a minimum of about 41 minutes behind (west of) the fictitious mean Sun. Because the mapping orbit plane lies two hours ahead of the fictitious mean Sun, the true Sun will always be to the west during mapping despite the positional variations.
The longest amount of time during one orbit that Mars eclipses the Sun occurs when the true Sun reaches its most eastern position (closest to the mapping orbit) in late April, early May 1998. At that time, eclipses will last for 40.7 minutes out of every 117.65 minute revolution. On the other hand, the shortest eclipse durations occur when the true Sun arrives at its most western position (farthest away from the mapping orbit) in late July, early August 1999. Then, eclipses will only last for about 35.5 minutes every orbit.
Date Event Comments 06 February 1998 Winter Solstice Start of southern summer 30 April 1998 Sun Beta Angle Maximum Beta angle at -15.909° 04 May 1998 Solar Conjunction Sun-Earth-Mars angle < 2° Command Moratorium Begins ?? May 1998 Conjunction Sun-Earth-Mars angle at 0.0° 21 May 1998 Solar Conjunction Sun-Earth-Mars angle > 2° Command Moratorium Ends 22 June 1998 Earth to Mars Distance Distance at 2.518 AU Maximum 15 July 1998 Vernal Equinox Start of northern spring 29 October 1998 Edge-on-Orbit (Diametric Earth beta angle at 0.0° Occultation) 17 December 1998 Mars Aphelion Solar distance at 1.666 AU 04 January 1999 Earth Beta Angle Maximum Beta angle at 4.891° 29 January 1999 Summer Solstice Start of northern summer 19 February 1999 Edge-on-Orbit (Diametric Earth beta angle at 0.0° Occultation) 02 May 1999 Earth to Mars Distance Distance at 0.578 AU Minimum ?? May 1999 Opposition Sun-Earth-Mars angle at 180° 01 August 1999 Autumn Equinox Start southern spring 01 August 1999 Sun Beta Angle Minimum Beta angle at -39.182° 22 September 1999 Earth Beta Angle Minimum Beta angle at -77.656° 25 November 1999 Mars Perihelion Solar distance at 1.382 AU 25 December 1999 Winter Solstice Start of southern summer
A set of three reaction wheels will point the spacecraft under normal circumstances. However, due to atmospheric drag, gravity gradient, and solar pressure torques, the momentum in the wheels will build up and must be periodically unloaded by autonomous thruster firings of the attitude control jets. Most of the unloading burns will occur near periapsis (around the south pole). For typical orbit trim maneuvers (OTM) with a velocity increment in the +X or -X direction, the spacecraft will turn under reaction wheel control to point the desired attitude control thruster in the proper burn direction. A typical OTM will last for several seconds.
As described in Section 2, the S&E-1 and S&E-2 data streams used for science data collection will be Reed-Solomon encoded, resulting in 250 encoded symbol bits for every 218 raw data bits collected by the Payload Data Subsystem (PDS) from the science instruments. Output streams will leave the PDS already Reed-Solomon encoded. Further detail regarding the bit allocations within the data stream can be found in Appendix ??
During the standard, daily 10 hour DSN tracking pass for playback of S&E1 recorded data, the spacecraft will circle Mars every 118 minutes and complete about five orbits. For most of the mapping phase, the spacecraft will enter zones of Earth occultation as it passes behind Mars relative to the direct line of sight to Earth. The time that the spacecraft spends on the "back side of Mars" in the Earth occultation zone will typically eliminate 40 minutes out of every orbit from use for playback of the recorded science and engineering data.
Some additional time when the spacecraft exits occultation will be used for the atmospheric occultation experiment. This process works by transmitting to the Earth through the atmosphere for a range of atmospheric altitudes from zero to 200 kilometers. Because telemetry modulation will be turned off for this experiment to provide maximum signal strength, the duration of this experiment will subtract from the available playback time. Normally, the exit experiment will typically require about five minutes of transmission per orbit. Occultation exit will typically occur near the Martian north pole.
After the spacecraft exits the zone of Earth occultation and completes the radio science experiment, the DSN will require a short amount of time to gain a "signal lock" for modulated telemetry. This time will amount to a maximum of five minutes on the first observed orbit of a tracking pass, but drop to one minute on each subsequent orbit. Then, the spacecraft will be free to downlink its recorded science and engineering data to Earth as it orbits from north to south around the "front" side of Mars.
Another radio occultation experiment will occur on every playback orbit as the spacecraft near the Martian south pole, but before passing behind Mars relative to the Earth and entering the occultation zone. Like the occultation exit experiment previously described, the entry experiment will consume about five minutes of available time for transmission of recorded data.
All of this information leads to the fact that the total available playback time on each orbit of a DSN tracking pass can be determined by taking the orbit period, and subtracting the time the spacecraft spends in Earth occultation, the time allocated for the radio science experiment, DSN lockup time, solid state recorder (SSR) management activities, and navigational uncertainty. With these subtractions, only 67.5 minutes per playback orbit out of the total 117.65 will be available for data return. Therefore, over a period of five playback orbits that elapse during a standard 10 hour DSN tracking pass, a theoretical maximum of 337.5 minutes will be available for downlink (five 67.5 minute segments).
However, several design constraints will reduce the true amount of available downlink time per tracking pass from the 337.5 minute maximum. First, for the purpose of design simplicity, the playback script will only downlink recorded data in discrete 67.5 minute segments. Second, DSN tracking stations will require five minutes to gain a "signal lock" on modulated telemetry the first time the station attempts to listen to telemetry on any given tracking pass. Therefore, in order to avoid potential loss of data during this initial five minute lock-up period, the DSN station must observe the spacecraft for at least a partial 67.5 minute orbit segment before the first whole segment of the actual data playback. Satisfying this constraint will result in only four guaranteed, discrete 67.5 minute segments (270 minutes total) available for playback per 10 hour tracking pass. The net result is that the playback rate must exceed the data record rate by a 5.33 to 1 ratio (24 hours of data in 270 minutes).
Figure 6-6: Typical Profile for Playback Orbit
In order not to "waste" the first partial segment of a tracking pass, the spacecraft will downlink real-time S&E1 data during that time (see Section 2 for a description of that data rate). This data will be the same as that being recorded for the next day's playback. The first five minutes of the real-time data will serve as a "throw-away filler" for the initial DSN signal "lock-up" period. Any additional real-time data returned during the first partial downlink segment of the pass will provide engineering telemetry for the ground monitoring teams to assess the current spacecraft "health" and status. This data is important because the engineering data downlinked during the playback session will represent the spacecraft status up to one day prior.
The possibility also exists that a tracking pass will contain a partial 67.5 minute playback segment after the four whole segments used to downlink the recorded data. This additional time will provide an opportunity to gather more real-time S&E-1 telemetry, or to uplink command updates. Such a scheme will provide additional data for health assessment and command verification.
From 19 June 1999 to 26 November 1999, the MGS mapping orbit will appear "face on" relative to Earth, and no periods of Earth occultation will occur. However, such a geometry will not substantially increase the amount of transmit time available per orbit because of solar eclipse periods. During times when the spacecraft passes through Mars' shadow, the batteries will provide power to run the electronics. Under the current design, running the transmitter throughout the solar eclipse period will drain the batteries to an unacceptably low depth of discharge level.
MGS will utilize playback rates 21.3, 42.7, and 85.3 ksps that correspond to record rates of 4.0, 8.0, and 16.0 ksps, respectively. These rate pairs were selected to match the 5.33 to 1.0 ratio, and to cover the range of expected telecommunications capability as the Earth-to-Mars distance varies over the mapping phase. At maximum Earth to Mars range, the link will only support the 21.3 ksps playback rate. At other times, the operations team will select the highest downlink rate supportable by the link margin. The next table lists the rate mode pairs available for use. Detailed information regarding the bit allocations for each instrument within the S&E1 data stream appear in Appendix ??.
Data Mode Record Rate Playback Rate Single SSR Storage Low Rate (LRC) 4,000 sps 21,333 sps 104 hours Medium Rate (MRC) 8,000 sps 42,666 sps 52 hours High Rate (HRC) 16,000 sps 85,333 sps 26 hours
Substantial record capacity margins exist in the SSRs as each one can store up to 104 hours of data at the 4.0 ksps record rate, and 26 hours at the 16.0 ksps rate. If both SSRs are functional, then it will be theoretically possible to store between 42 and 208 hours of data, depending on the specific record rate. However, downlinking more than 24 hours of recorded data will require a tracking pass longer than 10 hours in duration.
The next table shows an estimate of the dates of usage for the three different data rates over the course of the mapping phase. Availability of a DSN site to support a given rate assumes a 34m-HEF antenna, and greater than a 10 hour view period at the station where the spacecraft elevation, as seen from the antenna, lies at least 10° above the horizon.
Mission Dates S&E1 Record Rate S&E1 Playback Rate Stations Supported (10 Code hr. pass) 5-Apr-98 to LRC1 21,333 sps Goldstone yes Canberra 23-Sep-98 no Madrid yes 23-Sep-98 to MRC1 42,666 sps Goldstone yes Canberra 9-Dec-98 yes after 25-Nov-98 Madrid at LRC1 until 28-Oct-98 Madrid at MRC1 after 28-Oct-98 9-Dec-98 to MRC2 42,666 sps Goldstone no Canberra 6-Jan-99 yes Madrid no 6-Jan-99 to HRC1 85,333 sps Goldstone no Canberra 10-Nov-99 yes Madrid no 10-Nov-99 to MRC3 42,666 sps Goldstone no Canberra 26-Jan-00 yes Madrid no 26-Jan-00 to LRC1 21,333 sps Goldstone yes Canberra 21-Feb-00 yes Madrid yes after 16-Feb-00
For about half of the mapping phase, at least two DSN sites can support MGS tracking at the 10 hour per track requirement . This capability will facilitate the resolution of tracking conflicts with other missions and will provide flexibility in rescheduling should one station go "off-line." However, throughout 1999, only Canberra will possess the capability to support a 10 hour tracking pass because Mars will appear only at southern declinations as seen from Earth during that year. In this case, the problem lies not with the link margin, but with the physical amount of time that Mars will spend above the horizon as seen from the two northern tracking stations. However, due to Mars' close proximity to the Earth during the early to middle part of that year, receiving at the high data rate (85,333 ksps) will be possible from both Goldstone and Madrid with a shorter tracking pass duration.
Only one playback rate will be used over each tracking pass, and that the rate will not increase as the elevation of the spacecraft as seen from the DSN station increases over the first half of the track. The playback can be completed more quickly (or more data could be returned), by adjusting and maximizing the playback rate as a function of time during the tracking pass, but such a scheme will increase the complexity of planning the data return.
Only the Mars Orbiter Camera (MOC) and Thermal Emissions Spectrometer (TES) will utilize this 40 ksps real-time mode. During periods of S&E2 operations, the SSRs will continue to record data in the standard S&E1 mode for daily playback (described in the previous section) to avoid losing Magnetometer (MAG) and Mars Orbiter Laser Altimeter (MOLA) data. Another reason for simultaneous S&E1 recording is that real-time data collection cannot occur in the Earth occultation zone, primarily on the night side of the planet. However, the bit allocations for the different instruments within the S&E1 data stream will be slightly different when simultaneously transmitting S&E2 than during normal record operations. Detailed information regarding the bit allocations for MOC and MOLA within the S&E2 data stream and the modified bit allocations within the S&E1 data stream appear in Appendix ??.
The next table shows an estimate of the dates of usage for the two different S&E2 data rates over the course of the mapping phase. Availability of a DSN site to support a given rate assumes a 34m-HEF antenna, greater than an eight hour view period at the station where the spacecraft elevation, as seen from the antenna, lies at least 10° above the horizon.
Mission Dates S&E2 Playback Rate Stations Supported (>8 Hr. Pass) 5-Apr-98 to 40,000 sps Goldstone yes Canberra 23-Dec-98 listen-only before 21-Oct-98 Canberra yes after 21-Oct-98 Madrid listen-only before 19-Aug-98 Madrid yes after 19-Aug-98 16-Dec-98 to 80 ,000 sps Goldstone yes Canberra yes 28-Jul-99 Madrid yes after 23-Dec-98 28-Jul-98 to 80,000 sps Goldstone until 18-Aug-99 1-Dec-99 Canberra yes Madrid no 1-Dec-99 to 40,000 sps Goldstone yes Canberra yes 21-Feb-00 Madrid yes after 22-Dec-99
Although a 10 hour view period will be requested from the DSN for the every third day real-time S&E2 track, an eight hour period will suffice when either the link margin will not support a 10 hour DSN pass, or the physical viewing geometry (horizon to horizon) will not support a 10 hour pass. The reason is that the return of real-time data will not require four discrete 67.5 minute transmit segments as in the case of S&E1 playback. Using three discrete segments, plus two partial segments will work as long as the total real-time transmit time totals 270 minutes.
Near the beginning of the mission's mapping phase, the range from Earth to Mars will reach a maximum (2.518 AU on 22 June 1998). For several months before and several months after, the link margin will not support a data rate of 40,000 sps in the normal diplex (two-way) mode from Canberra and Madrid. For S&E2 operations from those two stations during times of maximum Mars range, the DSN will operate in a "listen-only" mode to increase the effective gain of the signal.
Each campaign will be dedicated to one or more disciplines in the fields of atmospheric studies, diametric occultations, global color imaging, geodesy, and gravity field measurements. In general, the science campaign schedule calls for as many of the "sub-campaigns" corresponding to the various individual disciplines to occur at the same time in order to minimize the amount of continuous tracking requested from the Deep Space Network (DSN). The next table shows the current science campaign schedule for MGS.
Mission Dates Campaign Code Disciplines Mars Season and Comments 2-Mar-98 to A Atmosphere #1 Northern Winter / 9-Mar-98 Southern Summer Observe carbon dioxide clouds in the northern polar regions and the initiation of dust and winter storms. 29-Jun-98 to B Atmosphere #2 Northern Winter / 6-Jul-98 Color Imaging #1 Southern Summer Observe residual southern polar ice cap, close of dust storm period. 26-Oct-98 to C Atmosphere #3 Northern Spring / 2-Nov-98 Color Imaging #2 Southern Fall Observe Diametric mid-spring water sources Occultation #1 and carbon dioxide Gravity #1 clouds in the southern polar regions. Earth is in the plane of the mapping orbit. 5-Jan-99 to D1 D2 D3 D4 Geodesy Look 1, Start of Southern Winter 12-Jan-99 Pass 1 Geodesy Atmosphere in both 20-Jan-99 to Look 2, Pass 1 hemispheres is clearest 27-Jan-99 Geodesy Look 1, at this time. Earth is 3-Feb-99 to Pass 2 Geodesy in the plane of the 10-Feb-99 Look 2, Pass 2 mapping orbit. Each 18-Feb-99 to Diametric geodesy "week" must be 25-Feb-99 Occultation #2 separated by multiplies of 88 orbits. 3-May-99 to E Atmosphere #4 Northern Summer / 10-May-99 Color Imaging #3 Southern Winter Observe Gravity #2 atmosphere profiles near the equator and water transport across the equatorial zone. 27-Sep-99 to F Color Imaging #4 Start of Northern Fall 4-Oct-99 Gravity #3 Earth out of mapping orbit plane angle is at a maximum. 13-Dec-99 to G Atmosphere #5 Start of Southern Summer 20-Dec-99 Observe carbon dioxide clouds in the northern polar regions and the initiation of dust and winter storms.
Similar to the strategy for collecting real-time data during the "extra" pass every third day as described previously, the solid state recorders (SSRs) will continue to record data in the standard S&E1 mode for daily playback while the spacecraft simultaneously transmits the real-time, S&E2 data stream back to Earth. During each day of a campaign, the current plan involves transmitting S&E2 real-time for two-thirds of the day, and then downlinking the previous 24 hours of recorded S&E1 data in the remaining one-third of the day.
Primarily, the major advantage of using AEM to assist in autonomously returning recorded data is manifested in significant savings in operational planning time due to minimizing the number of scripts needed over the entire mapping phase for playing back data. The following sub-sections describe the implementation of AEM in returning the science data.
In addition to the SSR playback, time in the EVPs must also be allowed for the radio science experiments at Earth occultation exit and entry, when the spacecraft appears to pass through the atmosphere from 0 to 200 km and from 200 to 0 km, respectively, as viewed from Earth. For the first 15 months of mapping, the radio science durations each orbit typically require 6 to 9 minutes of transmission time. The radio science experiments require an additional 200 seconds (100 seconds above 200 km at egress and at ingress) each orbit for calibration. For the worst case, when the Earth occultation periods are longest, there is sufficient time in the EVP to sequence all of the above activities.
Although AEM initiates scripts on eclipse entry and exit detection, the data return must be timed relative to Earth occultation transitions. Figure 6.9 shows the timing relationship between solar eclipse and Earth occultation. The figure shows that the eclipse and occultation periods generally overlap for most of the mission, except for late in mapping when the orbit is never occulted from Earth. The figure also shows that occultation exit and entry can occur before or after the corresponding eclipse exit and entry, respectively. AEM allows the ability to utilize both eclipse entry and exit detection to trigger separate scripts or to use either one or the other to initiate a single script.
From an operations perspective, it is desirable to develop a single script, triggered off either the eclipse entry or eclipse exit detection, that contains all of the commands to perform the radio science measurements and the recorder playback. However, the relative timing between eclipse and occultation transitions ensures that not all of the commands would be executed before the script was triggered again. To avoid this complication, a two script design approach can be utilized. The first script performs the radio science experiment at egress and the playback of the recorded data, and is timed relative to Earth occultation egress. The second script contains the radio science experiment at ingress and is timed relative to Earth occultation ingress.
Figure 6-7: Timing Relationship Between Earth Occultation and Solar Eclipse
As noted above, since occultation egress can occur before eclipse egress and similarly for the ingress case, the previous eclipse transition must be used to trigger the appropriate scripts. Thus, the radio science at egress and recorder playback script is triggered off of eclipse entry detection and contains a delay time equal to the time between eclipse entry and occultation exit minus the time required to warm up the telecom subsystem. The radio science at ingress script is correspondingly triggered off of eclipse exit detection, and contains a delay time equal to the time between eclipse exit and occultation minus 100 seconds for radio science calibration above 200 km.
Late in the mapping phase, when the spacecraft orbit is never occulted from the Earth, the playback could be performed continuously and not broken into four 67.5 minute portions. However, the spacecraft has power constraints which prohibit continuous transmission in eclipse, so that the strategy outlined in this subsection is still valid. Because there are no radio science experiment opportunities, a single script approach can be utilized however, which performs 67.5 minutes of playback triggered off of the eclipse exit detection.
Normal commanding is performed on the spacecraft using stored command sequences. These sequences are generated by the ground using sequence generation software tools, uploaded to the spacecraft, and initiated at the desired time by an "absolute" time-tag. The commands contained in the sequence are executed relative to each other using offset times (i.e., "relative" time-tags). By using AEM, the mapping sequences need only contain the commands to enable the appropriate pre-loaded AEM script on-board the spacecraft, prior to the desired orbit, and not the entire set of commands required to return the data for each scheduled DSN pass. Short-range DSN schedules are generated 8 weeks in advance of the sequence load. Using this DSN allocation file, the appropriate enable commands prior to the start of each scheduled playback pass and high-rate real-time pass (and disable commands upon completion of the pass) are placed in the sequence.
By utilizing AEM, the duration of the mapping stored sequences are no longer driven by the accuracy of the navigation timing uncertainties. Instead the sequence durations are driven exclusively by the availability of an accurate conflict free DSN allocation file and the sequence development time required by the ground to generate and validate the sequences, which should also be significantly reduced due to the decreased number of commands that need to be generated.
Four-hour windows shall be provided in the mapping sequences for the OTMs. This allows the maneuver time to be selected from two orbits in the event that large gravity anomalies are present on one of the orbits. Spacecraft and payload activities which conflict with the maneuver must be kept out of the window.
The current plan is to design a solar conjunction sequence that will provide real-time spacecraft telemetry when the solar array is illuminated and the opportunity to acquire navigation Doppler data during a single DSN tracking pass each day. A special AEM script will be developed to turn the transmitter on and return real-time telemetry for this period. This strategy allows for daily health assessment of the spacecraft and the payload, radiometric monitoring of the orbit, and limited science observations on the day side of the planet whenever telemetry can be received on the ground.
Realtime performance at the lowest science collection rate of 4 ksps (S&E-1 LRC) may not even be achievable throughout the period. Because of this, and because of the desire to have at least some spacecraft health status each day, a composite data return strategy of 250 bps and 4 ksps has been designed. For the first ten minutes of each orbit, engineering-only telemetry will be returned at 250 bps. For the rest of the orbit, 4 ksps S&E-1 will be returned in realtime. Radio Science atmospheric occultation measurements will also be performed during conjunction. During the Radio Science segments of each orbit, the transponder will be 1-way with telemetry and ranging modulation off, while for the rest of the orbit, the transponder will be configured with telemetry and ranging on, and with 2-way tracking enabled. As the conjunction geometry becomes more extreme, the uplink carrier may not be transmitted from the ground and one-way Doppler data, referenced to the USO, will be measured at the DSN. This will reduce noise on the downlink, as the signal makes only one pass through the solar corona, rather than two. Later, as the geometry improves, two-way data collection will resume. The data return profile described above will be repeated for each orbit during solar conjunction, even though tracking requirements are only for one 34-m HEF pass per day. If anomalies or any other circumstances indicate that additional coverage is warranted, downlink from the spacecraft will be available without emergency commanding.
When commanding again becomes possible after the conjunction, normal mapping data return will begin.