IAA-L-0508P



A LOW COST MERCURY ORBITER MISSION

Kenneth J. Ely, Wallace T. Fowler* & Byron D. Tapley**

Center for Space Research

The University of Texas at Austin

Abstract

Due to Mercury's small mass and position deep within the solar gravitational well, an orbiter mission poses difficult performance requirements (i.e. DV, thermal extremes). However, Yen (JPL) showed that with extended trip times, substantial Mercury missions are feasible. Low cost, quick concept-to-launch, high quality, and reliability are balanced in constraining a small spacecraft mission to Mercury. The primary science objective is high quality multispectral imagery and altimetry. The spacecraft is based on the Clementine spacecraft and sensors. UV/Visual/infrared cameras, star trackers, and a laser altimeter are proposed to determine mineralogical composition, surface structure/morphology, spectral/compositional mapping, and topography. Clementine sensors would require upgrades to enhance thermal dissipation and to increase radiation protection. The first available launch opportunity is in August 1996 and the target orbit is a 300 km polar Mercury orbit. Launch vehicle options are discussed. Options for companion or follow-on spacecraft are presented.

INTRODUCTION

This study was completed as an exercise in preparation for more detailed analyses of Mercury mission planning. The goals of this study are threefold: to suggest that Mercury, the only inner planet not orbited by a spacecraft, be considered for near term spacecraft missions, to promote the use of small satellites with focused scientific goals for planetary missions, and to present a conceptual design for a low cost Mercury orbiter mission.

Mercury's known and unknown characteristics are examined and used as justification for a Mercury orbiter. A general discussion of mission constraints and objectives defines the design methodology. A number of science objectives have been identified by the Terrestrial Bodies Science Working Group1 for future Mercury missions. For the purposes of this study these science objectives were grouped to define smaller payload sets for separate spacecraft with focused science goals, rather than a single all inclusive mission.

The selected mission scenarios are presented, followed by the spacecraft conceptual design including DV requirements, estimated mass and power budgets and chosen sensors. Until Yen2 developed a series of reverse multiple gravity assist trajectories in 1985 a Mercury orbiter was not believed feasible with current propulsion technology, due to the high orbit energy requirements. A discussion of the selected trajectory along with launch vehicle options is presented. Recommendations for further study are discussed.

MISSION JUSTIFICATION

The Planet Mercury

Mercury has long been one of the most elusive planets. Although it is the fourth brightest planet in the solar system, Mercury's small orbit and close proximity to the Sun make it very difficult to observe from Earth. Since its maximum apparent angle relative to the Sun is only 28° Mercury is always seen low in the sky-- limiting night-time observations. Turbulence found in the lower atmosphere further contributes to poor "seeing" conditions, forcing day-time observations through use of appropriate screens to filter scattered sunlight. Thus telescopic observations provide little in terms of surface markings.

Due to these poor observation conditions very little is known about the planet Mercury. We have learned in recent years that much analysis of observations made during the last century was in error. Until radar observations in 1965 showed that Mercury's sidereal period of rotation is actually only 59.6 days, it was largely believed that Mercury's rotation about its axis was synchronous with its revolution about the Sun (or 88 days). Maps of the Mercurian surface created by scientists such as G.V. Shiaparelli, T.J.J. See, G. Wegner, Chapman and Carmichel, and Dollfus (between 1881 and 1968) were proven to be crude representations at best by the Mariner 10 observations. These maps, which were based largely on refractor photographs and historical drawings, demonstrate that there is no chance of achieving much detail of Mercury's surface features from Earth bound observations3.

The three successful flybys of Mercury by the Mariner 10 spacecraft in 1974-75 increased scientists knowledge base enough only to prompt more specific questions. Portions of the planet's surface and features were now seen with a 5000 fold increase in photographic resolution, while unexpected phenomena were detected in the planet's immediate environment. Mariner 10 confirmed the 1965 discoveries that Mercury has lunar-like regolith, lunar-like surface materials, large craters and little or no atmosphere. A largely unexpected result was the discoveries, through use of a magnetometer, plasma probe and charged particle detector, that Mercury has a weak magnetic field and interacts with the solar wind in an earth-like manner. Through successive flybys Mariner 10 confirmed, that although only approximately 1% the strength of Earth's, Mercury possesses a dipole magnetic field approximately aligned with its spin axis4. It was also found that Mercury produces a bow wave in the solar wind at a distance of perhaps one radius from the surface of the planet5.

The discovery of this earth like magnetic field supports the wide belief that Mercury is a chemically differentiated planet with an iron core. Based on this evidence the diameter of this iron core is estimated to be between 70% and 80% the diameter of the planet (roughly the size of the moon). Radar measurements and radio occultations have shown Mercury to be 4878 kilometers in diameter.

As mentioned above, Mercury's surface features closely resemble those of the moon. Mercury is also similar in its albedo, surface scattering properties and in its emission of infrared radiation and radio waves4,5. Mercury has many craters and basins similar in landscape to parts of the moon. However, significant differences are apparent when observing relatively smooth plains between the craters and basins as opposed to the densely packed and overlapping craters in the highlands of the moon.

Based on Mercury's size and mass it is much denser than the moon or Mars and only slightly less dense than the Earth. Due to the Earth's high pressure interior, its bulk density is greater than the laboratory density of its constituent materials. Therefore Mercury's high bulk density implies a even greater abundance of heavy elements than Earth. Mercury is believed to be 60% iron by mass, but unlike the Earth it is not compressed, making it intrinsically the densest planet.

Mercury's orbital characteristics are well defined. Mercury's average distance from the Sun is only .39 AU. Coupled with its slow rotation period this produces the largest diurnal temperature range in the solar system (over 430°C near the subsolar point to -173°C or lower during the long nights). Mercury exhibits a 3/2 spin-orbit coupling resonance whereby it rotates exactly three times on its axis every two revolutions about the Sun. It is believed that the spin-orbit coupling is a result of tidal interaction with the Sun removing angular momentum and slowing its originally higher spin rate.

Not only is Mercury's orbit plane inclined 7° to the ecliptic (the second largest inclination in the solar system behind the 17° of Pluto), but its orbit is highly elliptic (e=0.206). Mercury at closest approach to the Sun is .308 AU while its aphelion distance reaches .467 AU. Its spin-orbit coupling and eccentric orbit lead to an extremely long "noontime", providing the harshest surface environment of any planet in the solar system. Mercury at perihelion receives 10 times the solar energy per unit surface area as the moon4.

Importance of Spacecraft Observations

Each planet in a solar system of only nine deserves significant study. Since the late 1970's, however, Mercury has been ignored when considering planetary probe concepts. This continued until August of 1986 when The Mercury Conference, sponsored by the Division for Planetary Sciences of the American Astronomical Society and International Astronomical Union Commission 16, was held in Tuscon, Arizona. Here scientists of varied disciplines gathered to exchange ideas on future groundbased and spacebased disciplines, resulting in the book, Mercury 6.

Mercury, is especially important due to its extreme characteristics. Mercury's close proximity to the Sun and its long day, approximately 176 Earth days, produce the most extreme diurnal temperature range in the solar system. In addition Mercury exists in the most intense solar radiation environment and is most affected by the solar tides. Mercury's barely existent yet complex atmosphere "represents the sole known example of a substantially magnetized small body and of a magnetosphere existing in the absence of an ionosphere."3 These unique characteristics are important to the understanding of other atmospheres (satellite and cometary) in the solar system. General processes that affect satellites and other terrestrial bodies could be better understood by studying Mercury's geology. Mercury's cratering history and earth-like magnetic field structure implies that important comparative planetological studies would be relevant to Earth.

Despite three successful flybys of Mariner 10 we still are unable to characterize the basic compositional make-up, phenomenological processes, and evolutionary history of Mercury. Although Earth-based observations and remote sensing from Earth orbit can provide some useful data, only by orbiting a spacecraft about Mercury can we answer the most important questions1.

DESIGN METHODOLOGY

As the world economies begin to recover from the recent recession, a large burden has been placed on the space industry as a whole by imposing more restrictive budgets. As a result, the design community is looking toward development of smaller space projects with focused scientific objectives employing a limited number of instruments. This design strategy has several advantages over the larger "do-it-all" spacecraft of the 1980's. It provides: significantly reduced development periods and costs, an increased number of spacecraft providing more opportunity for involvement across the scientific community, and reduced loss due to single point failures such as Mars Observer.

For this mission design the major constraint imposed was relatively low overall mission costs ($150M). Secondary constraints that support low cost as well as the design strategy described above include quick turnaround time from concept to launch and use of existing reliable hardware when available. Low cost, may seem to imply on the surface, small spacecraft. However due to the high orbit energy requirements of a Mercury mission (discussed shortly) the propulsion system dominates the spacecraft mass. Thus in order to answer all the questions posed by the science community multiple spacecraft are required. This study focuses on a single spacecraft and presents options for additional or follow-on satellites.

SCIENCE OBJECTIVES

Desired Science

As a result of the three successful Mariner 10 flybys of Mercury in 1974-75 a number of studies were focused on future Mercury missions. All the studies were aimed at placing a spacecraft in Mercury orbit and possibly landing on its surface. A report by the Terrestrial Bodies Science Working Group clearly defines the scientific objectives for such missions1. For completeness, the objectives stated in that report are summarized. Additionally, a general set of instruments necessary to accomplish these goals is defined.

Identified in the report were nine general problem areas that require investigation. These include: (1) planetary properties such as the gravity field and the figure and mass distribution; (2) magnetic investigations to determine the magnetic field source and the effect of thermal history on planetary magnetization; (3) magnetospheric processes to determine the mechanism that accelerates high-energy particles and what affect the lack of an ionosphere has; (4) atmospheric processes such as the thermal, surface and magnetospheric losses; (5) core properties such as its size, composition and state; (6) crustal chemical composition; (7) crustal evolution; (8) Mercury's thermal history; and (9) solar physics and relativity tests.

To accomplish these goals five general sets of instruments have been identified. The first includes high quality imagers and an altimeter. These instruments would not only provide images of the hemisphere missed by Mariner 10 but the data could infer geomorphological and tectonic processes as well. Multispectral imaging, orbital X-ray and gamma -ray fluorescence measurements and surface geochemistry experiments make up the next set. From this data chemical mineralogical composition and textural properties of the surface would be determined.

The third set includes high resolution synoptic spectroscopy, in situ charged particle measurements and in situ mass spectroscopy. These measurements would provide the full chemical composition of the atmosphere and clues to how it is generated.

In situ magnetic field and charged particle environment observations would give planetary scientists insight into how the magnetosphere is generated and its interaction with the time dependent atmosphere and variable solar wind. These measurements could help determine the present state of Mercury's core as well. Finally, surface exploration would be required to determine the global geophysical properties such as the gravity field, heat flow and seismicity.

Focused Science Objectives

The primary science objective for this mission concept is to obtain high quality multispectral imaging and altimetry of the entire planet. The goal is to give researchers the first complete close-up look at Mercury. It is intended to be only one of a possible series of spacecraft. Using a small set of instruments the following Mercury characteristics will be determined: mineralogical composition, surface structure/morphology, spectral/compositional mapping, and topography. A small gravity experiment may be included to determine Mercury's internal structure and gravitation. A low circular polar orbit would be best suited for these spacecraft observations.

Science Objectives for Additional Spacecraft

A possible second spacecraft would focus on Mercury's magnetosphere and atmosphere. Solar physics and relativity experiments could be included without a large mass penalty. The spacecraft incorporating the instruments necessary for these studies should insert into an elliptic polar orbit7. Science obtained would include detailed compositional and dynamical characterization of Mercury's magnetosphere and atmosphere at various altitudes and longitudes.

Follow-on missions should include either penetrators or landers to completely characterize Mercury's surface composition, seismic and thermal activity, surface mechanical properties, and surface magnetic field and charged-particle environment. A surface radio beacon would permit determination of the amplitude and period of the planet's physical libration7. Landers are given lower priority due to the large DV required to soft land on the surface. A significant amount of this science can be obtained through use of a penetrator network. Penetrators have the advantages of lower mass, design simplicity, and no propulsion requirements. However better characterization of the surface composition is recommended for a more reliable penetrator design.

MISSION SCENARIO

The Mercury orbiter mission will deliver a single spacecraft to Mercury for extended orbital study of the planet's composition, surface structure and topography. An August 1996 launch aboard either a Delta II (7925) or a Russian Proton launch vehicle will place the spacecraft on its interplanetary trajectory. The Delta II launch option places the spacecraft on a trajectory employing two swingbys of Venus and four of Mercury itself. The first three Mercury encounters enhance DV performance by reducing the heliocentric approach velocity at Mercury arrival. This trajectory provides final arrival at Mercury 5.3 years after launch. For the Proton scenario the spacecraft will make two swingbys of Venus on its way to Mercury with a trip duration of 2.1 years. The Venus encounters are separated by 225 days, the period of Venus' orbit, and are powered gravity assists. Upon arrival at Mercury, the spacecraft will initiate an insertion burn placing it in a 300 km circular polar orbit. The spacecraft will undergo initial instrument calibration and then map Mercury's surface for at least 88 days (one Mercury year).

MISSION ANALYSIS

Spacecraft Design

This spacecraft design is modeled after the design strategies used by the Clementine and ALEXIS small satellite missions8,9. Each program demanded small spacecraft mass and costs while employing aggressive scheduling constraints and a limited number of instruments. Additionally, each program accepted a certain level of risk: Clementine in using largely unproven equipment and ALEXIS in limiting redundancies.

The Mercury orbiter mission calls for an ambitious balance between low-cost, quick concept-to-launch turnaround, and high quality and reliability. To meet these criteria a complement of proven subsystem technologies is combined with an array of relatively new lightweight sensors and attitude control devices. The major subsystems (i.e. power, communications, command and data handling, propulsion, etc.) will rely largely on mature technologies. This will keep development costs at a minimum and allow for scrounging among other spacecraft programs without compromising reliability. Table 1 lists rough mass and power estimates for the major subsystems10,11. Due to the conceptual nature of this design a 15% margin was added to each of the mass and power budgets. The spacecraft has a total estimated dry mass of 159 kg. Since minimal orbit maneuvers will be required after low-Mercury orbit insertion this is considered a reasonable estimate for required deliverable mass. The propulsion system mass is based on the selected trajectory options discussed next.

The sensors chosen for this mission are the same complement presently being flown on Clementine. They consist of high and low resolution UV/Visible cameras, long and short wave infrared cameras, a laser ranger and 2 star trackers. With appropriate filters these sensors can provide the desired science. Table 2 lists the sensor attributes11. The UV/Visible cameras and star trackers are modifications of existing and tested SDIO hardware12. The infrared sensors and laser ranger are presently being tested aboard Clementine. Additionally, new lightweight attitude control devices using significantly less power are being flown on Clementine. These are considered highly reliable, and if successful on Clementine should be employed for this mission concept.

Table 1 - Mercury Orbiter Subsystem Estimates
Subsystem
Mass (kg)
Power (W)
Power
26
300
Communications
35
-20
Attitude Control
10
-40
Command & Data Handling
15
-20
Thermal
5
-20
Structure
40
N/A
Propulsion

Option 1 (Delta II)

Option 2 (Proton)


566

2635
-5
Sensors
7
-90 (peak)
Margin (15%)
21
-29
Totals

Option 1 (Delta II)

Option 2 (Proton)


726 (injected)

2794 (injected)

159 (Dry mass)

Table 2 - Sensor Attributes*
Sensor
Power (W)
Mass (kg)
High Res UV/Vis
10.4
1.25
Low Res UV/Vis
6.1
.483
Long Wavelength IR
26.7 (peak)
1.65
Short Wavelength IR
15.9
1.6
Star Tracker (1)
4.5
.37
Laser Ranger (LIDAR)
5.0 w/ 20 peak during mapping
1.0

*Adapted from Soyka 11

Trajectory Analysis

Mercury's position deep within the solar gravitational well and its small mass pose difficult performance requirements for an orbiter mission. High heliocentric approach velocities for standard transfer trajectories and Mercury's low gravitational braking potential result in severe launch and orbit insertion requirements. However Yen showed that through successive gravity assists of Venus and Mercury, mission performance could be enhanced to levels deliverable by the current stable of US launch vehicles. The use of Venus gravity assists to reduce launch requirements for Mercury missions is not a new concept; this was used by the Mariner 10 spacecraft in 1974-75. However, Yen discovered that through repeated Mercury encounters the heliocentric approach velocity at Mercury could be significantly reduced, thus reducing the large orbit insertion propulsion requirements.

Numerous launch opportunities exist between 1996 and 2010 employing this multiple gravity assist technique. Several trajectory options exist for each launch window, offering a trade-off between increased deliverable mass to Mercury orbit and extended transit times. The trajectory types are designated by the encountered planets, such as E for Earth, V for Venus and M for Mercury. Superscripts on the planet designator indicate the number of encounters with the particular planet. As an example, EV2M3 designates a trajectory from Earth to Mercury employing two Venus encounters and three Mercury encounters, the third being Mercury orbit insertion. Due to space limitations refer to Yen 2 for a detailed discussion as well as trajectory plots.

An August 1996 launch opportunity was selected primarily due to the mission constraint of quick concept-to-launch turnaround. Additionally, this opportunity provides by far the best performance trajectories of the remaining launch windows detailed by Yen through 2007. Yen detailed four trajectory options for each launch opportunity. All four options were considered for this analysis. Table 3 shows the DV requirements and flight times for the launch occurring on 7/10/96. These trajectories required an escape energy (C3) of 27.9 km2/s2. DVa indicates the Mercury orbit insertion DV required. DVpl is the total post-launch DV including navigation and other orbit operational allowances. Thus mass estimates for attitude control propellant are reflected in the propulsion system estimates.

Table 3 - Trajectory Performance*
EV2M
EV2M2
EV2M3
EV2M4
Flight Time (yr)
2.1
2.9
3.8
5.3
DVa (km/s)
4.189
3.235
2.378
1.762
DVpl (km/s)
4.528
3.957
3.278
2.855
DVtot (km/s)
8.717
7.192
5.656
4.617

* Table adapted from Yen 2

Launch Vehicle Study

Before choosing a trajectory option a launch vehicle analysis was carried out. Launch vehicle selection was constrained by maximum deliverable mass to a 300 km Mercury orbit and minimum launch costs. Also considered were the tradeoffs due to extended mission time. Table 4 lists the deliverable mass and cost estimates using the four trajectory options for various available launch vehicles. The standard medium lift US launch vehicles were considered. The Titan IV and Space Shuttle were not considered due to their prohibitively high costs. Since the world space community has recognized the cost effectiveness of cooperative efforts, the European Ariane IV and Russian Proton were also considered. These mass estimates were based on the previously cited DV calculations and Earth escape performance data detailed in JPL-D 6963 Rev. 113.

Highlighted in bold print are the viable options based on the deliverable spacecraft mass; estimated previously as 159 kg. All vehicles considered (except Atlas I) can deliver the required mass for at least one of the trajectory options. In order to best demonstrate the tradeoffs involved, two launch vehicle scenarios were considered. Use of the Delta II and the Russian Proton, employing the EV2M4 and EV2M trajectories, respectively, represents two extremes in terms of total launch mass and mission duration. These considerations are discussed subsequently. It should also be noted that the use of the Ariane vehicle may become a feasible option. A shared data arrangement between the US and ESA may result in reduced launch costs. The use of the Atlas IIAS and Titan III were given no further consideration due to their high costs.

1. Delta II Scenario

This is the most attractive option in terms of spacecraft mass and DV performance. An estimated injection mass of 726 kg is required for this scenario. For the given spacecraft dry mass, the propulsion system, although substantial, is not overwhelming. Additionally, launch costs are significantly lower than the Atlas, Titan and Ariane launch vehicles. However this extended

Table 4 - Launch Vehicle Performance and Costs
Launch

Vehicle
Launch Cost*

$M
Mesc

kg*

C3=27.9

km2/s2
M(lmo) kg

EV2M

2.1 Yrs
M(lmo) kg

EV2M2

2.9 Yrs
M(lmo) kg

EV2M3

3.8 Yrs
M(lmo) kg

EV2M4

5.3 Yrs
Delta II 7925
50-55
740
42.109
69.527
115.215
162.138
Atlas I
70-80
650
36.987
61.071
101.202
142.419
Atlas IIA
85-95
1100
62.594
103.35
171.265
241.016
Atlas IIAS
110-130
1500
85.355
140.933
233.543
328.659
Titan III/TOS
150-225
1600
91.046
150.329
249.113
350.569
Ariane 44L
100-110
1600
91.046
150.329
249.113
350.569
Ariane 44+4 stage
100-110
1700
96.736
159.724
264.682
372.480
Proton
35
3200
182.091
300.658
498.225
701.139

*Values adapted from JPL Document 693613

mission (5.3 years) is considered undesirable since the sensors are new and long duration survivability has not yet be proven. This long duration mission also incurs additional operational costs. Finally, the following question should be addressed: "Does the amount of science gained from this mission concept sufficiently justify the added cost of an extended transit time?"

2. Proton Scenario

This option is attractive due to the shorter transit time and the bargain launch cost. The $35 M launch price coupled with a mission duration of less than 3 years (2.1 years in transit + 88 days in mapping orbit) provides a strong sell. However for the same spacecraft dry mass as used in the Delta II scenario, a severe penalty is paid in prolusion system mass (Table 1). Therefore, this scenario would likely include in-transit propulsion system staging. Although the overall price of spacecraft development may increase over the first scenario, it is not likely to offset the launch and operation costs of the first scenario.

Although the Russian Proton is quite attractive in terms of cost it may pose significant political hurdles. NASA missions have typically relied on US launch vehicles. The use of SDIO developed sensors may cause additional concern about a foreign launch.

Design Considerations

A major concern with any Mercury mission is the role of the thermal subsystem in Mercury's intense radiation environment. Thermal design will likely require significant effort in the development phase of the program. Conservative mass and power estimates were accounted for in the above analysis. In addition to overall spacecraft thermal control, the SDIO sensors must be upgraded for interplanetary flight. These upgrades include improved thermal dissipation, increased reliability of electronic parts, and increased radiation hardening12.

The choice of solar cells as the primary power source requires further investigation. This option may not be feasible due to the intense radiation in the near-sun environment. Advances in solar cell radiation hardening should be considered. If not, the solar arrays would be replaced by RTG's at a substantially higher cost.

There may be an additional advantage to closely following the Clementine design. The Clementine program originally consisted of two spacecraft and thus subsystem components were procured for each. However Clementine II did not survive recent budget cuts. If the Clementine II project is not revived, this hardware could be incorporated into the Mercury orbiter, significantly reducing development costs.

COST SUMMARY

Use of a larger number of small satellites, each with focused objectives, is quickly being recognized as a cost effective alternative to the larger scientific platforms of the 1980's and early 1990's. Reductions in costs can be attributed to several factors. The Clementine and ALEXIS programs both had relatively quick turnaround times from concept to launch. Simply put, development time costs money, and therefore should be kept to a minimum. This restriction demands less complexity in design and integration and is therefore well suited for spacecraft with a limited number of instruments. As a result, use of appropriate existing hardware from previous projects is considered key to minimizing costs. In addition to simplifying the development, existing spacecaft systems are usually priced attractively due to the larger number of vendors than actual projects8. Finally, launch services incur a significant amount of the total program cost; therefore options should be considered carefully.

This study recommends following a design strategy similar to the Clementine program. The Clementine program took less than two years from concept to launch. A similar schedule for this Mercury orbiter concept is required for the selected August 1996 launch date. Clementine employs a limited number of lightweight sensors and equipment developed by SDIO. An attractive option for the Mercury orbiter would be use of available Clementine II systems. Finally, Clementine was launched aboard a refurbished Titan IIG, at a cost of approximately $30 M. Use of the Delta II and Russian Proton rocket were considered as viable launch vehicle options for the Mercury orbiter. At its present low cost and high lift capability the Proton is the most attractive option on paper.

CONCLUSIONS

It was shown that there is much scientific justification for a Mercury orbiter mission. The study of Mercury's magnetosphere and atmosphere are important to the understanding of atmospheres existing on other bodies throughout the solar system. Additionally, important comparative planetological studies of Mercury's geology, cratering history and earth-like magnetic field are relevant to Earth. Only by orbiting Mercury can we better understand the planet's processes. Small spacecraft programs with focused scientific objectives are quickly being recognized as an economical alternative to the larger scientific platforms of the past. A variety of options, including trajectories and launch vehicles, exist for a Mercury orbiter mission. Aggressive scheduling is required to meet the August 1996 launch date. To reduce costs, existing hardware may be available from previously cancelled programs, and foreign launch vehicles were considered.

REFERENCES

1. "Report of the Terrestrial Bodies Science Working Group, Volume II, Mercury," JPL Publication 77-51, Jet Propulsion Laboratory, Pasadena, California, October 14, 1977.

2. Yen, C.-W., Ballistic Mercury Orbiter Mission via Venus and Mercury Gravity Assists. In The Journal of the Astronautical Sciences, Vol. 37, No. 3, July-September 1989, pp. 417-432.

3. Chapman, C.R., Mercury: Introduction to an End-Member Planet. In Mercury, eds. F. Vilas, C.R. Chapman and M.S. Matthews, The University of Arizona Press, 1988, pp. 1-23.

4. Murray, B.C., The Planets, W.H. Freeman and Company, San Francisco, 1983, pp. 5-15.

5. Whipple, F.L., Orbiting the Sun: Planets and Satellites of the Solar System, Harvard University Press, 1988, 1981, pp. 173-182.

6. Mercury, eds. F. Vilas, C.R. Chapman and M.S. Matthews, The University of Arizona Press, 1988.

7. Stern, S.A. and F. Vilas, Future Observations of and Missions to Mercury. In Mercury, eds. F. Vilas, C.R. Chapman and M.S. Matthews, The University of Arizona Press, 1988, pp. 24-36.

8. Soyka, M.T., "A preliminary Mission Plan for the Clementine II/DSPSE II Spacecraft", The University of Texas at Austin, Spring 1993.

9. Cheaper Faces Reality. In IEEE Transactions on Nuclear Science, Vol. 40, No. 4, August 1993, pp 863-873.

10. Space Mission Analysis and Design, eds. J.R. Wertz and W.J. Larson, Kluwer Academic Publishers, Dordrecht, The Netherlands, 1991.

11. Soyka, M.T., "Sensor description of the Clementine/DSPSE Mission" Revision 1, the University of Texas at Austin, February 21, 1993.

12. Rustin, Lt. Col. P., "Clementine: Executive Overview", Texas Space Grant Consortium Summer R&D Workshop, June 1993.

13. "Launch Vehicles Summary for JPL Mission Planning," JPL-D 6936 Rev. C, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, February 1993.