Chapter 11 TECHNOLOGIES for HUMAN EXPLORATION

Chapter 11 TECHNOLOGIES FOR HUMAN EXPLORATION

Scott Lowther lays out the possibilities for Mars mission heavy lift boosters. Pathfinder Chief Scientist Matt Golombek explains his mission’s results.

502

MAR 98-051

MARSSAT: ASSURED COMMUNICATION WITH MARS

Thomas Gangale

In the past, robotic missions to Mars have accepted the inevitable communications blackout that occurs when Mars is in solar conjunction. This interruption, which lasts several weeks, would seem to be unacceptable during a human Mars mission. This paper proposes a relay satellite as a means of maintaining vital communications links during conjunction, and explores candidate orbits for such a spacecraft.

The basic approach to system design is to minimize size, weight, and power of spaceborne elements of the communications system, since it is more economical to compensate with large, heavy, and power-consuming elements on Earth. Ideally, it is the Earth-to-Mars link which should drive the overall system design, with the Earth-to-relay and Mars-to-relay links impacting system design as little as possible. This ideal is approached by minimizing the length of the link between the relay spacecraft and Mars. An orbit whose period is one Martian year, but whose eccentricity and inclination both differ from that of Mars, assures communications between Earth and Mars during conjunction while minimizing the length of the link between the communications satellite and the Mars mission.

1.0 STATEMENT OF NEED At the end of April of this year, the Deep Space Network (DSN) lost contact with the Mars Global Surveyor spacecraft in orbit around Mars. An entire month passed before communications were reestablished with the vehicle. This hiatus was not caused by any hardware or software failure; rather, it was an inevitable consequence of planetary orbital mechanics. During this period in May 1998, Mars passed behind the Sun as seen from Earth, a planetary configuration known as solar conjunction. Fortunately, links were re-established at the end of May, and Mars Global Surveyor is continuing its mission. During the nearly three Mars years over which the Viking 1 Lander operated, there were three such lengthy communication blackouts due to solar conjunctions. This Mars-Sun-Earth alignment occurs at 780-day intervals on the average, varying from 766 to 803 days due principally to the eccentricity of Mars’s orbit.

Various mission profiles have been proposed for human expeditions to Mars. Among these is the conjunction class mission, which utilizes minimum-energy Hohmann trajectories both to and from Mars. However, the use of Hohmann transfers on both the outbound and inbound legs of the mission requires roughly a 500-sol layover on Mars to await the proper planetary configuration for the return flight. As can be seen in Figure 1, solar conjunction occurs near the midpoint of this 500-sol stay on Mars. Although the use of the conjunction class scenario on initial human Mars missions is an issue yet to be decided, it is likely that this type of mission profile will be flown at some point in a human Mars exploration program, not only because it is the most propellant-efficient profile, but also because it maximizes stay time on Mars while minimizing travel time to and from Mars with respect to other mission profiles using the same class of propulsion systems.

503 Figure 1

Until the Tracking and Data Relay Satellite System was completed in the 1980s, human missions historically endured short-duration interruptions in communications. These blackouts would last for a few minutes either while passing between ground stations or during reentry. There were communications losses of as much as an hour during the Apollo program when vehicles passed behind the Moon. But it is hard to imagine that a communications outage on the order of one month will be tolerated on a human Mars mission. Now, the specific duration of the interruption during a solar conjunction depends on several factors, such as the amount of link margin designed into the communications system, as well as the minimum data rate that is acceptable from a mission standpoint. Still, regardless of how much robustness is designed into the communications links, the minimum blackout period will always be on the order of weeks, not the few minutes or hours that have been experienced on past human space missions.

Communications interruption by the Sun will become even more of a problem as human Mars operations build up to permanent bases. The conjunction blackout will of course hold true for a Mars base as well as for a conjunction-class mission, but furthermore, such a base will also have to contend with oppositions, which are, from the Martian point of view, inferior conjunctions of Earth, i.e., when Earth passes in front of the Sun as seen from Mars. During oppositions, Earth will be able to receive signals from Mars, but Earth’s transmissions to Mars will be drowned out by the Sun’s radio noise.

504

Some means of assuring uninterrupted communications between Earth and Mars will have to be included in human Mars program planning. To satisfy this need, I propose a relay satellite concept which I call MARSSAT.

2.0 PRELIMINARY SYSTEM TRADE-OFFS Figure 2

Figure 2 depicts the connectivity for a Mars communications system designed to circumvent the solar conjunction blackout, including Mars surface stations, low Mars orbit vehicles, a constellation of communications relays in Mars orbit, MARSSAT, and the DSN.

The basic considerations in sizing a satellite communications system are embodied in the link budget, which includes the following parameters:

On a conceptual level, spacecraft weight (and therefore cost) is driven by the size of the antenna and the power requirements of the transmitter and receiver. These in turn are driven by the range over which the link is required to operate. Exact numbers for the size, weight, and power of specific mission elements are a subject for detailed system engineering, and thus beyond the scope of this presentation. However, the basic approach is to minimize size, weight, and power of spaceborne elements of the communications system, since it is more economical to compensate with large, heavy, and power-consuming elements on Earth.

505

Ideally, it is the Earth-to-Mars link which should drive the overall system design, with the Earth-to-MARSSAT and Mars-to-MARSSAT links impacting system design as little as possible, since these alternative links represent additional costs. Assuming that the range of the Earth-to-MARSSAT link will be on the same order as that of the Earth-to-Mars link, the MARSSAT communications equipment need only be comparable to the equipment on the near- Mars elements. The stressing case, however, will be the Mars-to-MARSSAT link, for size, weight, and power will be at a premium on all spaceborne mission elements. It is for this link that the system must be optimized, since in this case, there are no large ground stations to figure into the link budget. To achieve minimum system impact, the maximum range over which the Mars-to-MARSSAT link must operate should therefore be as short as possible. At the same time, however, a minimum angular separation between the MARSSAT spacecraft and Mars, as seen from Earth during solar conjunction, must be maintained in order to reduce the impact of solar noise on the links. In general terms, this identifies the trade space to be investigated in the system engineering process.

For Mars Global Surveyor, mission planners and telemetry engineers defined the solar communications outage as occurring when the Sun-Earth-Mars angle was within seven degrees, and they planned for a loss of signal from 30 April to 26 May during the 1998 conjunction. It was also noted that a quiescent sun could have reduced this angle to five degrees. Several factors could reduce this minimum solar separation angle for human missions. The link throughput requirements might be limited to voice communication and only that telemetry necessary to the safety of the crew, while science data taken during the conjunction period could be recorded in situ and transmitted to Earth after the conjunction. Also, the use of a laser communications system might offer advantages over a conventional radio system. In my preliminary analysis, required minimum solar separation angles between two and three degrees were assumed. These may be unrealistically small angles from the mission operations perspective, but they provided me with some stressing cases for evaluating orbit stability.

This leads to my next point. Another consideration in the location of the relay spacecraft is the stability of its position relative to Mars over the design life of the satellite, since the more stable the orbit, the less fuel must be expended to maintain the spacecraft’s position — yet another factor affecting design weight. In this paper, the eight-Mars-year (fifteen-Earth-year) cycle over which the relative positions of Earth and Mars more or less repeat is defined as the mission life of the MARSSAT vehicle. Thus the spacecraft would be required to provide communications through seven solar conjunctions.

3.0 ORBIT SIMULATION AND EVALUATION CRITERIA

3.1 MARSSAT Simulation

To investigate candidate orbits, I developed a simulation which provides two split-screen graphic displays. Display 1 (Figure 6) consists of the familiar solar system “overhead” view from above (north of) the ecliptic, and the in-the-ecliptic view along the vector of the vernal equinox. Display 2 (Figure 7) provides two views edge-on to the ecliptic plane. Both views are boresighted on Mars. The left side of the screen is a view from the perspective of the sun and sighted along the acceleration vector of Mars (the Sun-Mars line), while the right side of the screen is a view sighted along the velocity vector of Mars (perpendicular to the Sun-Mars line). The large divisions along the XY axes of the two in-the-ecliptic views represent one degree of arc, measured from 400 million kilometers, which is roughly the distance between Earth and Mars during conjunction. It is these two Mars-centered views that generate information that is useful in evaluating potential MARSSAT orbits.

506

3.2 Evaluation Criteria

Some criteria for evaluating candidate MARSSAT orbits are:

•

MARSMAX — Maximum range from Mars in kilometers.

•

SUNMIN — Minimum angular separation in degrees from the Sun-Mars line for an observer at a distance of 400 million kilometers.

•

A figure of merit, defined as SUNMIN * 107 / MARSMAX, expresses the optimization of minimum solar angular separation and maximum range from Mars.

Orbital drift (in kilometers).

•

A second figure of merit, defined as FMERIT * 106/DRIFT, expresses the optimization of minimum solar angular separation, maximum range from Mars, and minimum orbital drift.

Not investigated were DV requirements to insert satellites into candidate orbits, although it should be pointed out that this would be an important consideration in the selection of a MARSSAT orbit.

4.0 ORBIT CONCEPTS

4.1 Co-Orbital Leader/Trailer Satellites

Figure 3

The simplest orbit for a MARSSAT vehicle would be one having the same parameters as that of Mars itself but slightly out of phase, i.e., either leading or trailing Mars by a few degrees in its orbit around the Sun. Ideally, a spacecraft in this orbit would remain stationary with respect to the Sun-Mars system. Unfortunately, such a relationship can only be stable if the phase angle of the orbit is either -60° or +60° with respect to Mars, corresponding to the equilateral

Lagrange points L4 and L5. As seen in Figure 3, satellites with phase angles of -2° and +2° exhibit very poor stability, quickly departing from their assigned stations in the vicinity of Mars toward the L4 and L5 points.

507 Figure 4

Of course, the equilateral Lagrange points themselves are too far from Mars to be suitable stations for MARSSAT, since links over this 230 million kilometer range would require spaceborne communications elements whose size, weight, and power would be on the order of a DSN ground station (Figure 4).

4.2 Co-Period In-Plane Oscillating Satellites 508

Figure 5

Another simple idea, one which would avoid the instability problem of the co-orbital leader/ trailer concept, is an orbit in which a satellite would alternately lead and then trail Mars, thus tending to balance out the gravitational influence of Mars. Such a scheme can be achieved by having the satellite in an orbit whose period is the same as that of Mars, but whose eccentricity is either greater or less than that of Mars. Figure 5 depicts the behavior of a spacecraft whose orbital elements are the same as Mars, except for an eccentricity of 0.05. It can be noted that from the Martian point of view, a spacecraft in this orbit appears to orbit around Mars, although in reality it is gravitationally bound to the Sun. The MARSSAT simulation indicates excellent stability for this orbit over an eight-Mars-year period. Since the satellite circles Mars once every Martian year, the perturbation effects of Mars’s gravity tend to cancel out.

An obvious disadvantage of this type of orbit is that since it is in the same plane as Mars’s orbit, as seen from the Sun the satellite transits Mars twice each Martian year, and thus an unobstructed line-of sight with Earth during solar conjunction is not assured by a single satellite. A second satellite, one whose longitude of perihelion were 90° out of phase, would be necessary.

4.3 Co-Period Out-of-Plane Oscillating Satellite Figure 6

509 Figure 7

510

Figure 8 Figure 9

Figure 6 and Figure 7 illustrate an orbit whose period is one Martian year, but whose eccentricity and inclination both differ from that of Mars. As with the in-plane oscillating satellite, this out-of-plane orbit mimics the motion of a satellite in orbit around Mars, although the satellite is not bound by Mars’s gravity, but is actually in solar orbit. In contrast to the in- plane oscillator, however, a minimum solar angular separation from Mars is maintained as, referring to Figure 8, MARSSAT A) rises north of, B) trails, C) drops south of, and D) leads Mars in their journey together around the Sun. In such an orbit, only one satellite is necessary to assure line-of-sight with Earth during any solar conjunction. As with the in-plane oscillator orbit, excellent orbit stability is demonstrated over a fifteen-year period (Figure 9). This type of orbit therefore seems well-suited for the MARSSAT concept. Now, if we need to increase the solar separation angle over what I assumed in this study, orbit stability improves, since we increase our distance from the gravitational influence of Mars.

Figure 7 also shows that a minimum angular separation, as seen from Earth, between MARSSAT and Mars, of 2.5 degrees is achieved. The line-of-sight distance from MARSSAT and Mars is on the order of 22 million kilometers. Note that this is only one-tenth the distance that a relay satellite stationed at a Lagrange point would be from Mars. All other things being equal, signal strength is inversely proportional to the square of the distance, thus to gain the same signal strength across ten times the distance, a Lagrange point satellite would need to be 100 times more powerful, affecting the weight of the spacecraft accordingly (Figure 10). So, if we have to double the separation angle to five degrees, the MARSSAT link is still only one- fifth the range of a Lagrange point link, which represents a link budget savings of 25 to 1. 511

Figure 10

5.0 ANALYSIS

The following parameters are used to characterize MARSSAT orbits: The sixth orbital parameter — a — is understood to be equal to the semimajor axis of Mars’s orbit, since in all cases we want the period of the MARSSAT orbit to be one Martian year.

Several general observations can be made concerning MARSSAT orbits:

512

•

Increasing the delta eccentricity with respect to Mars increases the horizontal (in-plane) travel of the spacecraft as seen along the Sun-Mars line. Increasing the delta inclination with respect to Mars increases the vertical (normal to plane) travel of the spacecraft. To increase the minimum solar separation angle (SUNMIN) requires simultaneous increases in delta eccentricity and inclination.

•

For eMARSSAT < eMars and DW < 0 the spacecraft rotates around Mars in a counterclockwise direction as viewed from the Sun.

•

For eMARSSAT < eMars and DW > 0 the spacecraft rotates around Mars in a clockwise direction as viewed from the Sun. Figure 11

As long as the semimajor axis of the MARSSAT orbit is identical to that of Mars, its period is one Martian year. The other five orbital elements can be combined in a vast number of permutations to produce useful MARSSAT orbits. For a specified SUNMIN and DW, however, there is a unique combination of e, i, Dv and DL which produces an optimum F2MERIT, i.e., optimizes minimum solar angular separation, maximum range from Mars, and minimum orbital drift. This trade space can be characterized as a series of curves representing optimized F2MERIT for specified SUNMIN values plotted across the full range of DW (Figure 11).

6.0 CONCLUSION

Selection of a specific MARSSAT orbit must be left to future system engineering trade studies based on the optimum Mars-to-MARSSAT range and the optimum Mars-to-MARSSAT solar separation angle, as well as DV requirements to insert a satellite into a given orbit. Human Mars mission studies should consider the minimum data rate that would be acceptable during the conjunction period. For instance, transmission to Earth of much of the science data recorded during the conjunction might be deferred until after the conjunction, while the minimum acceptable data rate might be primarily driven by voice links and engineering data relevant to mission safety. The minimum solar separation angle, and therefore the cost of the MARSSAT system, would be reduced accordingly.

513

Most space systems have one or more levels of redundancy designed into them, and even though these additional components increase the cost of a system, they are included in the design with the hope that they will never need to be used at all! In contrast, if assured communications with Mars throughout all phases of a human mission is a firm system requirement, the necessity of MARSSAT in providing that vital link with Earth will be a certainty.

514