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[Table of contents i](#__RefHeading___Toc145044488)

[List of Figures iv](#__RefHeading___Toc145044489)

[List of Tables vi](#__RefHeading___Toc145044490)

[1 Executive Summary 1](#__RefHeading___Toc145044491)

[2 Enceladus Background Summary 4](#__RefHeading___Toc145044492)

[2.1 Current State of Saturn Knowledge 4](#__RefHeading___Toc145044493)

[2.2 Current State of Enceladus Knowledge
4](#__RefHeading___Toc145044494)

[2.3 Mission Rationale 5](#__RefHeading___Toc145044495)

[3 Science 7](#__RefHeading___Toc145044496)

[3.1 Science Overview 7](#__RefHeading___Toc145044497)

[3.2 Scientific Objectives 7](#__RefHeading___Toc145044498)

[3.3 Mission Objectives 7](#__RefHeading___Toc145044499)

[3.4 Science Requirements 8](#__RefHeading___Toc145044500)

[3.4.1 Photopolarimeter/Radiometer 8](#__RefHeading___Toc145044501)

[3.4.2 Laser Altimeter 8](#__RefHeading___Toc145044502)

[3.4.3 Magnetometer 8](#__RefHeading___Toc145044503)

[3.4.4 Near-infrared mapping spectrometer
8](#__RefHeading___Toc145044504)

[3.4.5 Enceladus Descent Imager (EDI) 8](#__RefHeading___Toc145044505)

[3.4.6 Seismometer 9](#__RefHeading___Toc145044506)

[3.4.7 Pancam 9](#__RefHeading___Toc145044507)

[3.4.8 Saturn Orbiting Camera (SOC) 9](#__RefHeading___Toc145044508)

[3.4.9 ChemCam 9](#__RefHeading___Toc145044509)

[3.4.10 Enceladus Surface Analyzer by Mass Spectrometry (ESAMS)
9](#__RefHeading___Toc145044510)

[3.4.11 Amino Acid Analyzer (UREY) 10](#__RefHeading___Toc145044511)

[3.5 Trade Studies 10](#__RefHeading___Toc145044512)

[3.6 Connection to Past, Present, and Future Saturn Missions
10](#__RefHeading___Toc145044513)

[4 Payload 11](#__RefHeading___Toc145044514)

[4.1 Instruments 11](#__RefHeading___Toc145044515)

[4.1.1 Lander 11](#__RefHeading___Toc145044516)

[4.1.1.1 Amino Acid Analyzer 11](#__RefHeading___Toc145044517)

[4.1.1.2 Alpha Particle X-Ray Spectrometer
12](#__RefHeading___Toc145044518)

[4.1.1.3 Camera/Descent Imager 14](#__RefHeading___Toc145044519)

[4.1.1.4 ChemCam 14](#__RefHeading___Toc145044520)

[4.1.1.5 Laser Altimeter 16](#__RefHeading___Toc145044521)

[4.1.1.6 Magnetometer 17](#__RefHeading___Toc145044522)

[4.1.1.7 Enceladus Surface Analysis by Mass Spectrometry (ESAMS)
Experiment 17](#__RefHeading___Toc145044523)

[4.1.1.8 Near-Infrared Mapping Spectrometer
18](#__RefHeading___Toc145044524)

[4.1.1.9 Photopolarimeter/Radiometer (PPR)
19](#__RefHeading___Toc145044525)

[4.1.1.10 Drilling and Sample Delivery Subsystem (SD2)
20](#__RefHeading___Toc145044526)

[4.1.1.11 Seismometer 20](#__RefHeading___Toc145044527)

[4.1.1.12 Wide Angle Camera 21](#__RefHeading___Toc145044528)

[4.1.2 Orbiter 23](#__RefHeading___Toc145044529)

[4.1.2.1 Camera 23](#__RefHeading___Toc145044530)

[4.1.2.2 Student Payload 23](#__RefHeading___Toc145044531)

[4.1.2.3 Radio Science 24](#__RefHeading___Toc145044532)

[4.2 Up-scope Options 24](#__RefHeading___Toc145044533)

[4.3 Planetary Protection 24](#__RefHeading___Toc145044534)

[4.4 Conclusions 26](#__RefHeading___Toc145044535)

[5 Mission Architecture 27](#__RefHeading___Toc145044536)

[5.1 Overview 27](#__RefHeading___Toc145044537)

[5.1.1 Launch Vehicle and Earth Escape 27](#__RefHeading___Toc145044538)

[5.1.2 Interplanetary Trajectory 28](#__RefHeading___Toc145044539)

[5.1.3 Orbiter Mission Trajectory 31](#__RefHeading___Toc145044540)

[5.1.4 Enceladus Approach 31](#__RefHeading___Toc145044541)

[5.1.5 Future Trajectory Refinement 31](#__RefHeading___Toc145044542)

[5.1.5.1 The Launch Window 32](#__RefHeading___Toc145044543)

[5.1.5.2 Saturn Orbital Insertion 32](#__RefHeading___Toc145044544)

[5.1.5.3 Additional Refinements 33](#__RefHeading___Toc145044545)

[5.1.6 Conclusions 33](#__RefHeading___Toc145044546)

[6 Systems Engineering 34](#__RefHeading___Toc145044547)

[6.1 Requirements and Constraints 34](#__RefHeading___Toc145044548)

[6.2 Baseline Design 34](#__RefHeading___Toc145044549)

[7 Spacecraft Subsystems 36](#__RefHeading___Toc145044550)

[7.1 Structures and Mechanisms 36](#__RefHeading___Toc145044551)

[7.1.1 Requirements 36](#__RefHeading___Toc145044552)

[7.1.2 Proposed Architecture 36](#__RefHeading___Toc145044553)

[7.1.2.1 Saturn Orbiter 37](#__RefHeading___Toc145044554)

[7.1.2.2 Enceladus Lander 37](#__RefHeading___Toc145044555)

[7.1.2.3 Propulsion Stage 38](#__RefHeading___Toc145044556)

[7.1.3 Protection from the Space Environment
39](#__RefHeading___Toc145044557)

[7.1.4 Materials 40](#__RefHeading___Toc145044558)

[7.1.5 Separation Method 40](#__RefHeading___Toc145044559)

[7.1.6 Conclusions 40](#__RefHeading___Toc145044560)

[7.2 Environmental Control 41](#__RefHeading___Toc145044561)

[7.2.1 Temperature Control -- External 41](#__RefHeading___Toc145044562)

[7.2.2 Temperature Control -- Internal 42](#__RefHeading___Toc145044563)

[7.2.3 Radiation 42](#__RefHeading___Toc145044564)

[7.2.4 Micrometeorite Shielding 43](#__RefHeading___Toc145044565)

[7.2.5 Conclusions 43](#__RefHeading___Toc145044566)

[7.3 Radioisotope Power Systems 43](#__RefHeading___Toc145044567)

[7.3.1 Requirements 43](#__RefHeading___Toc145044568)

[7.3.2 Subsystem Design 45](#__RefHeading___Toc145044569)

[7.3.3 Trade Studies 46](#__RefHeading___Toc145044570)

[7.3.4 Conclusions 47](#__RefHeading___Toc145044571)

[7.4 Communications 47](#__RefHeading___Toc145044572)

[7.4.1 Requirements 47](#__RefHeading___Toc145044573)

[7.4.1.1 Command Requirements 47](#__RefHeading___Toc145044574)

[7.4.1.2 Telemetry Requirements 48](#__RefHeading___Toc145044575)

[7.4.2 Subsystem Design 48](#__RefHeading___Toc145044576)

[7.4.2.1 Orbiter 48](#__RefHeading___Toc145044577)

[7.4.2.2 Lander 50](#__RefHeading___Toc145044578)

[7.4.3 Alternate Design Configurations 51](#__RefHeading___Toc145044579)

[7.4.4 Conclusions 51](#__RefHeading___Toc145044580)

[7.5 Command & Data Handling 51](#__RefHeading___Toc145044581)

[7.5.1 Requirements 51](#__RefHeading___Toc145044582)

[7.5.2 Subsystem Design 52](#__RefHeading___Toc145044583)

[7.5.2.1 Small Deep Space Transponder 52](#__RefHeading___Toc145044584)

[7.5.2.2 S-Band Transponder 53](#__RefHeading___Toc145044585)

[7.5.2.3 Power Amplifier 53](#__RefHeading___Toc145044586)

[7.5.2.4 RAD750 Processor (uses VSWorks)
53](#__RefHeading___Toc145044587)

[7.5.2.5 Solid State Recorders 53](#__RefHeading___Toc145044588)

[7.5.2.6 cPCI-1553 & RS232/422 53](#__RefHeading___Toc145044589)

[7.5.3 Trade Studies 54](#__RefHeading___Toc145044590)

[7.5.4 Alternate Design Configurations 54](#__RefHeading___Toc145044591)

[7.5.5 Conclusions 54](#__RefHeading___Toc145044592)

[7.6 Guidance, Navigation & Control 54](#__RefHeading___Toc145044593)

[7.6.1 Requirements 54](#__RefHeading___Toc145044594)

[7.6.2 Subsystem Design 55](#__RefHeading___Toc145044595)

[7.6.3 Trade Studies 56](#__RefHeading___Toc145044596)

[7.6.4 Alternate Design Configurations 57](#__RefHeading___Toc145044597)

[7.6.5 Conclusions 57](#__RefHeading___Toc145044598)

[7.7 Propulsion 58](#__RefHeading___Toc145044599)

[7.7.1 Requirements 58](#__RefHeading___Toc145044600)

[7.7.2 Subsystem Design 58](#__RefHeading___Toc145044601)

[7.7.3 Trade Studies 60](#__RefHeading___Toc145044602)

[7.7.4 Alternate Design Configurations 60](#__RefHeading___Toc145044603)

[7.7.5 Conclusions 60](#__RefHeading___Toc145044604)

[8 Cost and Schedule 62](#__RefHeading___Toc145044605)

[8.1 Cost 62](#__RefHeading___Toc145044606)

[8.2 Schedule of Activities for Project
64](#__RefHeading___Toc145044607)

[8.3 Outreach 65](#__RefHeading___Toc145044608)

[8.3.1 Student Design Payload 65](#__RefHeading___Toc145044609)

[9 Conclusions 65](#__RefHeading___Toc145044610)

[10 Appendices 67](#__RefHeading___Toc145044611)

[10.1 Science Trade Studies 67](#__RefHeading___Toc145044612)

[10.2 List of Acronyms 69](#__RefHeading___Toc145044613)

[10.3 References 69](#__RefHeading___Toc145044614)

[Figure 1: Tiger stripes in the vicinity of the South Pole^54^
5](#__RefHeading___Toc145044435)

[Figure 2: Amino acid analyzer ('UREY') instrument^3,66^
12](#__RefHeading___Toc145044436)

[Figure 3: Engineering rendering of MARDI for MSL camera head as well as
electronics box. (Courtesy of Malin Space Science Systems)
14](#__RefHeading___Toc145044437)

[Figure 4: Mass Spectrometer Gas Chromatograph
18](#__RefHeading___Toc145044438)

[Figure 5: Near-Infrared Mapping Spectrometer^13^
19](#__RefHeading___Toc145044439)

[Figure 6: Exploded view of photopolarimeter^67^
20](#__RefHeading___Toc145044440)

[Figure 7: Miniature seismometer 21](#__RefHeading___Toc145044441)

[Figure 8: Cassini's Wide Angle Camera.
23](#__RefHeading___Toc145044442)

[Figure 9: Performance comparison of 5 launch vehicles for high energy
(earth escaping) orbits. ^15^ 27](#__RefHeading___Toc145044443)

[Figure 10: Delta IVH launch profile to LEO.^8^ A final burn from the
185 km parking orbit will send the EAGLE spacecraft into the
interplanetary cruise phase. 28](#__RefHeading___Toc145044444)

[Figure 11: EAGLE's interplanetary trajectory
28](#__RefHeading___Toc145044445)

[Figure 12: Enceladus orbit is highlighted in red in this view of the
Saturn system on the date of EAGLE's arrival into the Saturn system.^30^
29](#__RefHeading___Toc145044446)

[Figure 13: Direct capture by Enceladus, as shown in the above figure,
is too costly and too long a maneuver to perform.^30^
29](#__RefHeading___Toc145044447)

[Figure 14: EAGLE's SOI burn occurs above the C-Ring, 20,000 km above
Saturn's clouds. Eagle passes between the F and G rings twice.
29](#__RefHeading___Toc145044448)

[Figure 15: Orange represent's EAGLE's commissioning orbit, and green
represents EAGLE's mission orbit. Enceladus' orbit is shown in red.
30](#__RefHeading___Toc145044449)

[Figure 16: EAGLE passes through Saturn's rings, with periapsis slightly
above the ecliptic plane. 30](#__RefHeading___Toc145044450)

[Figure 17: Rhea's orbit, in blue; Titan's orbit, in orange; Iapetus'
orbit, in purple. Note the inclined nature of Iapetus' orbit.
30](#__RefHeading___Toc145044451)

[Figure 18: Full Saturn approach trajectory.
30](#__RefHeading___Toc145044452)

[Figure 19: The alignment of the inner planets constrains EAGLE to a
Venus arrival on December 3^rd^, 2023. 32](#__RefHeading___Toc145044453)

[Figure 20: Launch configuration, including the interplanetary insertion
propulsion stage. 36](#__RefHeading___Toc145044454)

[Figure 21: Orthogonal view of the orbiter
37](#__RefHeading___Toc145044455)

[Figure 22: Top view of the orbiter 37](#__RefHeading___Toc145044456)

[Figure 23: Orthogonal view of the lander configuration.
38](#__RefHeading___Toc145044457)

[Figure 24: Top view of the lander configuration.
38](#__RefHeading___Toc145044458)

[Figure 25: Propulsion stage 39](#__RefHeading___Toc145044459)

[Figure 26: Illustration of orbiter separation.
40](#__RefHeading___Toc145044460)

[Figure 27: V-clamps 40](#__RefHeading___Toc145044461)

[Figure 28: MMRTG & SRG power distribution in transit (BOM).
45](#__RefHeading___Toc145044462)

[Figure 29: Power distribution for the Saturn orbiter 8 years after
launch. 46](#__RefHeading___Toc145044463)

[Figure 30: Power distribution for the Enceladus lander during decent to
surface. 46](#__RefHeading___Toc145044464)

[Figure 31: Plot of atmospheric attenuation v. frequency. Once can see
that, in certain cases, a signal can be almost entirely lost when
transmitted at a high enough frequency.
48](#__RefHeading___Toc145044465)

[Figure 32: Flow chart for the Command & Data Handling subsystem.
51](#__RefHeading___Toc145044466)

[Figure 33: Small deep space transponder.
52](#__RefHeading___Toc145044467)

[Figure 34: Processor design roadmap^50^
53](#__RefHeading___Toc145044468)

[Figure 35: Mission Timeline with Reviews and Milestones. Mission
activities will be broken up by the NASA standard phases.
65](#__RefHeading___Toc145044469)

[Table 1: Assumptions involved in launch vehicle performance estimates
27](#__RefHeading___Toc145044470)

[Table 2: Swingby altitudes and dates 29](#__RefHeading___Toc145044471)

[Table 3: Summary of Saturn system trajectory burns.
30](#__RefHeading___Toc145044472)

[Table 4: Summary of ECS design components
43](#__RefHeading___Toc145044473)

[Table 5: Saturn orbiter: 2 SRG exclusive (BOM max = 116 Watts each)
44](#__RefHeading___Toc145044474)

[Table 6: Enceladus lander: 1 MMRTG exclusive (BOM = 125 W)
45](#__RefHeading___Toc145044475)

[Table 7: Trade study for the RPS 47](#__RefHeading___Toc145044476)

[Table 8: Trade study considering BOM power
47](#__RefHeading___Toc145044477)

[Table 9: Link budget for the orbiter-DSN downlink using HGA.
49](#__RefHeading___Toc145044478)

[Table 10: Link budget for orbiter-DSN downlink using one LGA.
49](#__RefHeading___Toc145044479)

[Table 11: Link budget for downlink from lander to orbiter.
49](#__RefHeading___Toc145044480)

[Table 12: The data rates for all of the science instruments on the
lander. Note that several of the instruments will not be sending data
back to the Earth. 50](#__RefHeading___Toc145044481)

[Table 13: Applications of Command & Data Handling components.^505152^
52](#__RefHeading___Toc145044482)

[Table 14: Specifications of C&DH components^505152^
54](#__RefHeading___Toc145044483)

[Table 15: Summary of GNC components 56](#__RefHeading___Toc145044484)

[Table 16: Summary of the Propulsion system
59](#__RefHeading___Toc145044485)

[Table 17: Costs by Work Breakdown Structure
62](#__RefHeading___Toc145044486)

[Table 18: Total mission cost funding profile.
64](#__RefHeading___Toc145044487)

From the time that scientists and engineers received the first
high-resolution images of the Saturn system provided by the Voyager 1
and Voyager 2 flybys in 1980 and 1981, respectively, the level of
understanding of Saturn and its moons has grown significantly, primarily
due to the contributions of the Cassini-Huygens mission to Saturn. At
the time of this study, Cassini was two years into its primary science
mission, providing tantalizing glimpses into a complex and intricate
system of rings and moons. Among the most intriguing discoveries in the
Saturn system are the South Polar hot spot and erupting geysers of the
small moon Enceladus.

The composition of Enceladus is primarily water ice, with varying types
of terrain across the surface of the moon, including smooth regions with
no craters, surface fissures called tiger stripes, and other types of
surface deformations. The South Pole has active geysers spewing
primarily water vapor and water ice, but also constituents of CO~2~,
N~2~, CH~4~, and light organics. Cassini will obtain further data during
a flyby in 2008, during which the spacecraft will pass at an altitude of
approximately 1000 km above the surface. While the Cassini flybys
provide appealing clues to the complex physical and geological processes
that shaped this moon, a dedicated mission to Enceladus will answer many
of the questions posed by these brief encounters.

This concept study proposes a dedicated orbiter/lander mission to the
surface of Enceladus. This mission, the Enceladus Astrobiology and
Geophysics Lander Expedition (EAGLE), has a complex set of mission
objectives. These objectives include determining whether liquid water is
present and mapping its distribution, investigating the surface chemical
composition, including any possible organic compounds, understanding the
physical processes that shaped the surface and the interior, and finally
probing the subsurface to identify the sources of heat in the region of
the South Pole. The following sections describe a design concept capable
of accomplishing each of these goals.

One of NASA's primary objectives under the Vision for Space Exploration
is to "follow the water" as exploration extends beyond the Moon and
Mars. The potential presence of liquid water combined with the hot spots
in the South Polar Region make Enceladus a primary exploration target.
Furthermore, water is a critical component in the development of life;
if there are organic compounds on the surface of Enceladus, the hot
spots in the South Pole may be an ideal place for life to exist.

The EAGLE mission includes two primary components: the Saturn orbital
communications and imaging platform, and the Enceladus lander. The
Saturn orbital vehicle will provide communications support for the
lander vehicle. Additionally, the mission orbit for the orbital craft
permits several flybys of many of Saturn's primary moons, allowing the
camera system on the vehicle to image these moons and further the
understanding of the Saturn system.

The transfer trajectory from the Earth to Saturn involves using a
Venus-Venus-Earth flyby approach. Launching in June of 2023, the
transfer to Saturn will take approximately seven years. The spacecraft
will then enter an orbit with its periapsis at the radius of Enceladus'
orbit, and the apoapsis near Iapetus. During closest approach of the
spacecraft to Enceladus, the lander will separate from the Saturn
orbital vehicle, and orbit Enceladus for several days. During this time,
the lander vehicle will use its laser altimeter to map the surface, and
its magnetometer to investigate the subsurface structure. The surface
characteristics mapped while the vehicle is in orbit will determine the
final landing site selection. Upon completion of the orbital phase, the
lander will descend to the surface to begin its primary mission.

The Cassini flyby of Enceladus in 2008 will provide further information
about the terrain of the moon in general, as well as the terrain in the
region around the South Pole, specifically. This data will be crucial to
determining the landing site for the EAGLE lander. While the science
objectives demand that the lander investigate the most scientifically
interesting region as possible, this may not be feasible if the terrain
is dangerous or unsuitable for a precision landing. Therefore, based on
the data received from Cassini, the lander will conduct its experiments
in a region that is scientifically valuable, but does not risk the
safety of the lander or the mission.

Once the lander is on the surface of Enceladus, the primary mission will
begin. The lander will use a ChemCam instrument for laser-induced remote
sensing for chemistry and micro imaging. This instrument has the
capability of detecting hydrated minerals, as well as elemental
composition of the surface. It can also characterize the geology of the
landing site, as well as nearby locations.

Several instruments on the lander require samples from the near
subsurface; therefore, the lander payload includes the Drilling and
Sample Delivery System, which uses an integrated drill to gather samples
from a shallow depth. An incorporated carousel receives the samples and
distributes them to the Mass Spectrometer and the Amino Acid Analyzer.

Using the sample taken from the Sample Delivery System, the Amino Acid
Analyzer has the capability of identifying and measuring complex organic
molecules such as amino acids, amines, and nucleobases. It also has the
ability to identify the chirality of amino acids. The sensitivity of
this instrument exceeds that of the Mars Viking experiments by several
orders of magnitude.

The lander payload will incorporate ten small sample collection ovens,
or pyrolysers, which will heat samples obtained from the near
subsurface. The Mass Spectrometer or Gas Chromatograph will then analyze
the samples to identify a range of masses from noble gases to organics.

In order to determine the elemental chemistry of the surface at the
landing site, the payload contains an Alpha Particle X-Ray Spectrometer.
By exposing the surface material to energetic alpha particles and x-rays
from a radioactive source, the APXS determines the elemental composition
by measuring the backscattered particles. This will provide information
about the surface structure, and identify possible compounds necessary
for organics.

Finally, the miniature seismometers, placed on the legs of the lander,
will investigate the physical structure of the surface and interior of
the moon, and characterize any tectonic activity found in the
subsurface. The seismometers could also aid in the identification of the
source of the hot spots in the South Polar Region.

Upon completion of the primary mission, the extended mission will
encompass a surface operations phase of hopping to different sites, and
conducting similar experiments to those described above. The low surface
gravity of Enceladus means that the propulsive requirements to lift off
and land are very low. This hopping option during the extended mission
allows for greater mobility and increased science return for the
mission. Furthermore, this allows for a progression from caution to
ambition in the operations at the surface, which is a proven technique
implemented during the Mars Exploration Rovers missions.

The launch vehicle payload capability constrains the mass of each
spacecraft. This mission will launch on a Delta IV (4050H-19), which can
carry payload mass of 7600 kg into an Earth-escape trajectory. This mass
placed a mass constraint of 600 kg dry mass on the Saturn orbiter, and
800 kg dry mass for the lander.

Finally, the cost of the EAGLE mission falls into NASA's flagship
mission classification. The cost of this mission in FY2006 dollars is
\$838 million. The total cost for the lifetime of the mission, including
development phases A through D, launch costs, and communications support
is \$1.13 billion.

Saturn, a gas giant that is the sixth planet from the Sun, is the second
largest planet in the Solar System, and has the most complex
ring-system. Its rings consist of ice particles, rocky debris and dust
that extend from 6,630 km to 120,700 km above Saturn's equator. There is
a distinct gap in the rings, first identified by astronomer
Jean-Dominique Cassini in 1675. The gravitational influence of Saturn's
moon Mimas is believed to be responsible for the so-called Cassini
Division. However, the Voyager missions in 1980 and 1981 revealed that
within the two main groups of rings, there is an intricate structure of
thousands of gaps and ringlets. The rings likely originated from a
former moon of Saturn whose orbit decayed until it was so close to
Saturn that tidal forces ripped it apart, or that the moon disintegrated
after a collision with a comet or asteroid. Another theory is that the
rings are the remains of nebular material from the formation of Saturn.
The rings, with a thickness of almost one kilometer, are primarily made
of silica rock, iron oxide and ice particles, which range in size from a
few micrometers to several tens of meters in diameter.

Due to the large sizes of the objects in the rings, it is difficult to
classify how many natural satellites Saturn has, but NASA has identified
56 objects as moons. Seven of these are massive enough to have collapsed
into a spheroid under their own gravitation: Mimas, Enceladus, Tethys,
Dione, Rhea, Titan and Iapetus.

Saturn itself is unique, as the only planet in the Solar System with a
density less than water -- its average specific density is only 0.6873
g/cm^3^. Saturn's structure consists of a rocky core surrounded by a
layer of liquid metallic hydrogen, with a layer of molecular hydrogen
above that. Its interior is extremely hot, with temperatures reaching
12,000 Kelvin at the core. Saturn is an oblate spheroid, flattened at
the poles and bulging at its equator, as a result of its rapid rotation
and fluid state. With an equatorial diameter of 120,536 km, Saturn has a
volume that is 755 times that of Earth. It has a mass of 5.6846 x 10^26^
kg, a rotation period of 0.449375 d, and an axial tilt of 26.73^o^. Its
atmosphere is approximately 93% Hydrogen, 5% Helium and 0.2% methane,
and exhibits a banding pattern similar to that seen on Jupiter.

Although many of the physical and chemical characteristics of Saturn are
known today, the most significant questions remain about its unique
moons, and the possibility of finding liquid water and perhaps life on
an extraterrestrial body.

Evidence for unusual processes operating on Enceladus was first provided
by Voyager 2, which returned low-resolution images of a world half
covered by craters and half smooth terrain. The perception of Enceladus
has changed dramatically over the past two years due to a series of
findings by the Cassini spacecraft.

On February 17, 2005, Cassini executed its first Enceladus flyby at an
altitude of 1,176 km, taking high-resolution images that revealed
smooth, crater-free plains, a sign of geological youth. Multiple cracks
in the icy surface were also observed, indicating tectonic activity. The
cracks appear to have different composition or ice crystal size from
their surroundings.^39^ In addition to these findings, the Near-Infrared
Mapping Spectrometer (NIMS) showed that the surface is made of nearly
pure water ice,^10^ and the Cassini Magnetometer revealed a tenuous
atmosphere that distorts Saturn's magnetic field lines.^20^ During a
second flyby on March 9, 2005, the Magnetometer found that the
atmosphere is concentrated above the South polar region.^20^

In its third Enceladus flyby on July 14, 2005, with a minimum altitude
of 175 km, Cassini detected a "cloud" above the South pole. This cloud
was composed mostly of water vapor and water ice, with CO~2~, N~2~,
CH~4~, and possibly light organics identified as trace constituents.^58^
In addition to these findings, the Composite Infrared Spectrometer
(CIRS) data revealed hot spots at the South Pole aligned along the
"tiger stripe" surface cracks,^54^ and the high-resolution images (4
meters/pixel) showed house-sized ice boulders in the South polar
terrain.^39^ Furthermore, particle measurements by the Cosmic Dust
Analyzer confirmed that Enceladus is the major source of Saturn's E
ring.^53^

![](media/image4.jpeg){width="2.938888888888889in"
height="2.9145833333333333in"}

[]{#__RefHeading___Toc145044435 .anchor}Figure 1: Tiger stripes in the
vicinity of the South Pole^54^

On November 27, 2005 Cassini collected very high phase angle images of
Enceladus, revealing multiple collimated jets that form a plume of
material (mostly water), extending hundreds of kilometers above the
South Polar surface.^39^

In mankind\'s search for life and habitable environments on other
worlds, evidence of liquid water elsewhere in the Solar system is of
great importance.  In 2005, images sent back from the Cassini-Huygens
mission to Saturn revealed water vapor and ice erupting from the south
pole of Saturn\'s moon Enceladus.  The Cassini spacecraft has performed
several flybys of Enceladus, sending back images of a dynamic body with
a relatively young surface covered in ice, exhibiting at least five
different types of terrain.  Cassini will fly by Enceladus only once
more in March 2008, leaving many questions unanswered.  Thus, a
dedicated mission to Enceladus is the only effective way to fill the
gaps that Cassini will leave behind.  This mission will send a lander to
the moon\'s surface, which will: 1) determine if liquid water is present
and map its distribution, 2) investigate the existence of life and/or
organic molecules, 3) learn about the physical processes that created
the surface, and 4) probe the physical structure of the moon\'s interior
and potentially identify sources of heat for the south pole
geysers. This mission will actively engage students in NASA science by
introducing outreach projects and involving the public in the search for
life.  Landing on Enceladus plays a key role in NASA\'s mission by
advancing the search for life and contributing to the exploration of the
universe.

As described above, images sent back from the Cassini-Huygens mission to
Saturn revealed water vapor and ice erupting from the south pole of
Enceladus.  These recent discoveries indicate that this small moon of
Saturn is one of the only active bodies in the Solar system. Since
Cassini will fly by Enceladus only once more in March 2008, many
questions will remain unanswered. The very large plume at its southern
pole could imply the existence of a subsurface liquid layer, creating
possibilities for important astrobiology findings.

In order to fill the gaps that Cassini will leave behind, a dedicated
mission to Enceladus is proposed. This dedicated mission will include a
lander that descends to the moon\'s surface, and:

1.  Determines if liquid water is present and maps its distribution,

2.  Investigates the surface chemical composition, such as elemental
    abundances and organic molecules that are critical to life,

3.  Studies the physical processes that created the surface, and

4.  Probes the physical structure of the moon\'s interior, potentially
    identifying sources of heat for the south pole geysers.

The EAGLE mission will carry several instruments that address the
scientific goals listed above. Some of these goals will be achieved
during the first stage of an Enceladus-orbiting spacecraft. This
spacecraft will carry a near-infrared mapping spectrometer (NIMS), a
photopolarimeter/radiometer (PPR), an Enceladus descent imager (EDI), a
laser altimeter, and a magnetometer. On the spacecraft's second stage, a
different science payload will be carried to the surface of Enceladus.
The instruments selected for this package include miniature
seismometers, a panoramic camera (Pancam), a drilling and sample
delivery system (SD2), a laser-induced remote sensing instrument for
chemistry and micro-imaging (ChemCam), an amino acid analyzer (UREY), an
Enceladus Surface Analyzer Mass Spectrometer (ESAMS), and an
alpha-particle X-ray spectrometer (APXS).

From orbit around Enceladus, EAGLE will study the differences between
various regions of the surface. The NIMS will address EAGLE's second and
third scientific objectives by identifying short-chain organic molecules
in the surface ice, and by measuring the size and crystallinity of the
ice particles to reveal how and when they were formed. The PPR will
contribute to EAGLE's fourth scientific objective by mapping the
temperature distribution across the surface of Enceladus and identifying
hot spots. During the orbit and descent phases, the EDI and laser
altimeter will meet the third objective by mapping the surface in three
dimensions at high resolution. Scientific objectives one and four will
be met by using a magnetometer. The magnetometer will attempt to
discover evidence of liquid water by investigating Enceladus' subsurface
structure when EAGLE is in the orbit phase.

Landing on Enceladus will enable more detailed characterization of the
surface properties near the vents. Miniature seismometers will help
identify pockets of liquid water (science objective #1) and activity in
the subsurface (science objective #4). The Pancam high-resolution images
will show structures that should provide more information about the
processes shaping the surface (science objective #3), whereas the
ChemCam and APXS instruments will determine the elemental composition of
the surface by observing various targets (science objective #2). By
analyzing samples from SD2 and from the plume, UREY and ESAMS will
measure the abundances of complex and simple organic molecules,
respectively (science objective #2).

The Photopolarimeter/Radiometer will study surface texture and
temperature variations. It will map thermal emission and polarization
properties across the South Polar region, and has the capability of
taking these measurements elsewhere on Enceladus. The Photopolarimeter/
Radiometer must have an accuracy of 1 K and a 2.5-mrad field of view.
The primary instrument functional requirement is that its internal
components must be temperature-calibrated. The Photopolarimeter/
Radiometer also requires that the spacecraft have an orbit that allows
for complete surface mapping.

The objectives for the laser altimeter are to determine elevation and
changes in topography. It is also a critical part of the decent phase
for the lander, used in conjunction with the autonomous landing
software. The laser altimeter will take measurements of the time
required for the laser pulse to complete a round trip from the
instrument to the surface of Enceladus and back to the instrument. In
order to make these measurements, the instrument must be directed
towards the surface. The functional requirements for the instrument
include a range up to 750m and a resolution of 50mm. Measurements must
be taken while in orbit and during decent phase of the mission.

The magnetometer will investigate the atmosphere and subsurface of
Enceladus. To do so, it will measure variations in the interaction
between Enceladus and Saturn's magnetic field, covering a range of 0 to
±256nT. The main instrument and mission requirements are that the
magnetometer be at the end of a 9 m long boom, and will operate during
while in orbit around Enceladus.

The Near-infrared mapping spectrometer will study surface composition
and ice structure on Enceladus. It will be capable of mapping near-IR
reflectance across the South Polar region and elsewhere. This requires a
resolution of 0.5-mrad, and an orbit that will provide opportunities for
complete surface mapping. Furthermore the instrument must be passively
cooled to \~150 K.^13^

During the orbiting phase and during entry, decent and landing, the
Enceladus Descent Imager (EDI) will map Enceladus and provide imagery
for the automated landing software. It will take color images using
multiple filters with a wavelength range of 400-1100 nm. This will
require a telephoto zoom with a varied field of view from 50^o^ to
10^o^. The Enceladus Descent Imager (EDI) also requires a resolution
between 24m/pixel and 5m/pixel, and greater during descent.

The miniature seismometers will observe tectonic activity in order to
characterize the subsurface of Enceladus from 10 to hundreds of meters.
To do so, the seismometers will measure tectonic activity over a period
of time (days to weeks, depending on the activity of the region). The
Seismometers must be placed on a stable part of the lander, and ideally
measurements will be taken at a time when lander vibrations are at a
minimum -- not during periods of active drilling. However, data can be
collected occur continually.

Throughout the mission, the Pancam will assess the morphology,
topography, magnetic properties, and geologic context of the landing
site. It must be able to constrain surface photometric and physical
properties as well as generate digital terrain models. The Pancam will
also obtain information regarding the density or physical properties of
the geyser material, monoscopic and stereoscopic images and mosaics, and
multispectral images over a range of viewing geometries. The Pancam is
also required to provide adequate stereo coverage as well as view and
obtain images of geyser material. To accomplish this, the Pancam must be
raised above lander (on a mast) and have capability to swivel 360°. It
must resolve one millimeter per pixel from 10 feet away, be in line of
sight of the calibration target, and take panoramic images in all
directions from the lander. Pancam must also be able to look up towards
geyser material and take approximately 20 images a day.

The Saturn Orbiting Camera (SOC) will take wide-angle color images of
Saturn and its moons aboard the Saturn Orbiter. This requires 18 filters
covering a wavelength range from 380-1100 nm. The camera must have field
of view of 3.5^o^ in order to image moons from a distance of 13,000 km
to greater than 100,000 km while in orbit. In addition to imaging while
in the Saturnian system, SOC must also be powered on to image during
gravity assist flybys of Venus and Earth.

ChemCam's science requirements include: characterizing the geology of
the landing region, accessing any organics present in the target,
detecting water ice, and measuring surface composition near the lander.
The laser must operate at a distance of 1-13 m at a frequency range of
1-20 Hz and a wavelength of 1067 nm. The laser should be directed onto
the sample causing ablation into light emitting excited states. The
micro-imager will provide very high resolution images of targets. The
mass unit must detect elements such as H, Li, Be, B, C, N, O, Ba, Sr,
Cs.

The Enceladus Surface Analyzer by Mass Spectrometry (ESAMS) must find
the composition of Enceladus' surface and be able to detect carbon
chains if they exist. The mass spectrometer will have sensitivity at a
range of 1-300 amu. ESAMS will use SD2 to take samples from the surface
and place them in the ovens before pyrolysis before analysis. Samples
will be collected from 2-3 depths with SD2 as well as from the Falling
Sample Collector (FSC).

The Amino Acid Analyzer (UREY) will find complex organic molecules in
surface samples. It must identify and quantify amino acids and
nucleobases with parts-per-billion sensitivity. It must also measure
chirality of amino acids from solid sample(s) collected from the
surface. SD2 is needed to accomplish these goals.

See the Section 10.1: Science Trade Studies.Science Trade Studies

The first flyby of the Saturn system was performed by the Pioneer 11
spacecraft in 1979. Pioneer took low-resolution images of Saturn and
several moons, but no surface features could be discerned due to the
resolution. Our knowledge of the Saturn system was enhanced by the
Voyager 1 and 2 missions in 1980 and 1981, which took the first high
resolution images of Saturn, its rings and its moons. Voyager 1 also
performed a flyby of Titan to investigate its atmosphere. However, the
most enlightening mission thus far has been the Cassini-Huygens mission,
launched in 1997. Cassini is presently orbiting Saturn, and continues to
send back images and data that further our understanding. The
Cassini-Huygens mission has served as a great foundation for this
mission, and the images of Enceladus taken by the Cassini spacecraft are
what initially sparked interest in this moon. In 2005, the Cassini
spacecraft revealed large geysers of water vapor spewing from the South
Pole of Enceladus. The Eagle mission intends to use some of the high
resolution images taken by Cassini to help select a landing site on
Enceladus. Additional information will be gained in Cassini's next flyby
of Enceladus, scheduled for 2008.

The amino acid analyzer ("UREY") is an instrument suite under
development for the European ExoMars mission, currently slated for
launch in 2011 or 2013.^65^ Using a solid sample, the instrument suite
will be capable of identifying and measuring the abundance of complex
organic molecules including amino acids, amines, nucleobases, and
polycyclic aromatic hydrocarbons (PAHs). The latter are found in almost
all sedimentary fossils uncovered on Earth, while the other molecules
are the building blocks of protein, DNA, and RNA, the compounds
essential to life on Earth. Sensitivity to these compounds will be in
the part-per-billion to part-per-trillion range, exceeding that of the
Viking experiments on Mars by several orders of magnitude. Chirality of
amino acids will also be measured, which will distinguish between a
biogenic vs. abiogenic origin of the amino acids.

The UREY package includes four components, which work together to study
the organic chemistry of planetary surfaces.^3,66^ First, a subcritical
water extractor (SCWE) extracts the targeted compounds from the
collected sample. These compounds are then heated to the point of
sublimation in an instrument called the Mars Organic Detector (MOD),
thereby isolating the organic molecules of interest and increasing their
concentration. The vaporized organics are trapped on a cold finger in
the MOD and bombarded with ultraviolet laser light, which causes PAHs to
fluoresce with a characteristic spectrum. The cold finger is also coated
with fluorescamine, which bonds to amino acids, amines, and nucleic
acids to create fluorescing compounds. MOD identifies each of these
organics using a fluorescence spectrograph.

If MOD detects organics in a sample, then the sublimed molecules are
transferred via a "sipper" to a "lab-on-a-chip" microcapillary
electrophoresis (CE) system for further composition analysis at the
part-per-trillion level. The CE device also measures the chirality of
any amino acids in the sample. Finally, a Mars Oxidant Instrument (MOI)
is included with the SCWE/MOD/CE suite for further study of possible
oxidative surface environments, such as that discovered on Mars by the
Viking landers. Oxidation reactions would rapidly break up organic
molecules on the surface, so that if MOI finds the surface to be highly
oxidizing, then a non-detection of organic compounds by MOD will be
understandable.

The total mass of the instrument is 3.2 kg, and it occupies a volume of
11 cm x 14 cm x 15 cm. It has a peak power requirement of 28 W.

![](media/image5.wmf){width="2.9868055555555557in" height="3.175in"}

[]{#__RefHeading___Toc145044436 .anchor}Figure 2: Amino acid analyzer
('UREY') instrument^3,66^

To determine the elemental chemistry of the material at the landing
site, the payload includes an Alpha Particle X-ray Spectrometer (APXS),
similar to those on the MER and Rosetta missions, on the bottom of the
lander. By exposing the target material to energetic alpha particles and
x-rays from a radioactive ~244~Cm source, the APXS determines the
elemental composition by measuring the energy spectra of backscattered
alpha particles and emitted x-rays. There are six of these small
radioactive sources in the APXS "head".^2^

The radioactive sources emit the alpha particles, which bounce back from
the target material---after directly hitting an atomic nucleus---into a
detector in the APXS, along with some x-rays that are excited from the
target in the process. By using the laws for Rutherford Backscattering
in an angle close to 180°---conservation of energy and conservation of
linear momentum---it is possible to calculate the mass of the nucleus
hit by the alpha particle. Light elements absorb more energy, while
heavy nuclei reflect more back. The resulting energy spectrum of the
scattered alpha particles then exhibits 25% to almost 100% of the
initial particles, depending on how many were absorbed. X-ray
fluorescence is also emitted when the alpha particles bump out lower
shell electrons, forming vacancies that are filled by electrons from
upper shells. This easily detectible x-ray fluorescence is
characteristic of each element, having best resolution for heavier
elements.

The detectors then measure the energy distribution of the alpha
particles and x-rays, and the analysis determines the amounts of
different chemical elements present. Most of Enceladus' surface is water
ice; however, it is desirable to measure the surface composition to
discover what elements are in the ice, along with their proportions. The
elemental composition provides information about the formation of the
moon's surface, and identifies the amount of critical elements necessary
for life.

The x-ray mode is sensitive to major elements, such as Mg, Al, Si, K,
Ca, and Fe, and to minor elements, including Na, P, S, Cl, Ti, Cr, and
Mn. The alpha mode is sensitive to lighter elements, most importantly C
and O. The depth of analysis varies with atomic number, ranging from
about 10 to 20 micrometers for sodium, to about 50 to 100 micrometers
for iron. The detection limit is typically 0.5 to 1 weight percent,
depending on the element. The APXS is insensitive to small variations in
the geometry of the sample because all major and minor elements sum to
100 weight percent.^2^

The APXS instrument consists of a sensor head mounted on the bottom of
the lander within a cylindrical casing, with electronics mounted in the
lander's warm electronics box. The casing will be lowered from within
the lander via rack and pinion so that the sensor will be within a
maximum distance of 40 mm from the target, though closer to the point of
touching is preferred. The total strength of the six 244Cm alpha
radioactive sources is about 30 mCi. The sources are each covered with
3-µm aluminum foils that reduce the energy of emitted alpha particles
from the initial value of 5.8 MeV to about 5.2 MeV.^2^ This
significantly reduces the alpha particle scattering cross section of
carbon and removes much of the atmospheric background in the alpha
spectra. However, since the excitation spectrum is broadened, there is a
slight degradation of the alpha spectral resolution. Collimators in
front of the sources define the instrument's field of view, which is
about 38 mm in diameter at the nominal working distance of 29 mm. For
best results, the APXS must be placed with its front aperture ring in
contact with the sample surface and the axis of the sensor head must be
within 20 degrees of normal to the surface.^11^

The accumulation time for the APXS will usually be at least 10 hours per
sample analysis, although much shorter durations are acceptable when
only the x-ray mode is used. However, the total accumulation time can be
separated into several shorter accumulation periods. The APXS can store
up to 12 sets of accumulated spectra and can transmit the data to the
lander after one set is accumulated or after all sets are
accumulated.^2^

A pair of doors protects the entrance to the detector head from
potential contaminants. These doors swing inward and lock open when the
sensor head is pressed against a target or other hard surface. They can
be closed again by actuation of a release mechanism. The inner surfaces
of the doors provide a calibration reference surface for the instrument.

Proper preflight calibration is necessary to properly analyze APXS data,
so the instrument will undergo an extensive calibration program
mirroring that of the Athena APXS on MER. The Athena calibration
measurements were made in a chamber filled with a mixture of gases that
closely matched the composition of the Martian atmosphere, at the
appropriate atmospheric density. Since Enceladus has no atmosphere,
EAGLE's APXS will be calibrated within a vacuum unless Cassini's June
2008 flyby shows a change in atmospheric composition in the South Polar
region. Calibration measurements will include:

1.  spectral "library" measurements of pure elements and oxides;

2.  geochemical standards that span the full range of plausible
    Enceladus surface compositions;

3.  standard targets in both hard ice and powdered form, to investigate
    texture effects;

4.  the APXS flight calibration target;

5.  several blind certified geochemical reference standards, for
    independent assessment of the accuracy with which compositions can
    be measured. ^2^

The APXS has a mass of 0.57 kg, and is 10.5 x 8.0 x 6.5 cm.^47^ It uses
on average 0.3 W, though with associated power supplies draws 0.8 W of
raw power. Nonvolatile APXS memory stores a maximum of 64 Kbits, even
with power off, and plays back a maximum of 16 Kbits per query at 9600
baud (RS 232) when turned on. It has a data rate of 9600 kb/s. (NSSDC
ID: MESURPR-01) The lander CPU does not need to run during APXS
operation, except for on/off commands and APXS data playback.

On MER, the APXS sensor head only operated when the surrounding Martian
ambient temperature was between -100°C and +25°C. This is much warmer
than the surrounding Enceladus ambient temperature, so thermal control
measures will be necessary to provide an adequate environment.^2^

Since versions of the APXS flew on both MER and Rosetta missions, it has
extensive applicable flight heritage.

The goal of the Enceladus Descent Imager (EDI) is to map the moon during
the orbilander phase, as well as provide information to automated
software for landing. This versatile camera must be able to provide high
resolution images from an orbit 50 km above the surface, and continue
imaging as the orbilander descends, to a distance of 5 m from the
surface. It must also have low mass due to the large amount of fuel
required to take mass into Enceladus orbit. Considering these
constraints, the best instrument to fulfill the requirements resembles
Malin Space Science Systems' Mars Descent Imager (MARDI) from the Mars
Science Laboratory (MSL)^34^.

This instrument will have flight heritage from cameras aboard the Mars
Polar Lander, Phoenix, and Mars Science Laboratory; however,
technological advancements will improve EDI's performance by the time of
launch. EDI includes many features such as telephoto zoom, focus,
high-speed data acquisition, and multispectral imaging from 400-1100nm.
The zoom optics provide a field of view ranging from 10^o^ at the long
focal length position to 50^o^ at the shortest focal length. The
detector is a 1600x1200 pixel CCD which produces low noise images. EDI
will be capable of mapping Enceladus' entire surface with approximately
24 m/pixel resolution while in 50 km orbit, with the capability to
obtain high resolution images at about 5 m/pixel. During descent, EDI
will continue to image the surface. This information will be used for
navigation as described in Section 4.1.1.5.2. The entire package will
have a mass of approximately 0.6 kg, operate at an average power of 10
W, and will be 7 x 7 x 10 cm^3^ in dimension^34^.

![](media/image6.png){width="3.1458333333333335in"
height="2.0944444444444446in"}

[]{#__RefHeading___Toc145044437 .anchor}Figure 3: Engineering rendering
of MARDI for MSL camera head as well as electronics box. (Courtesy of
Malin Space Science Systems)

After the EAGLE lander reaches the surface of Enceladus, several
scientific investigations will occur. However, the mobility and scope of
scientific investigations will be limited by the icy terrain. To
overcome this constraint, the payload includes the ChemCam, which has
the ability to capture elemental composition of samples located at great
distances.

The ChemCham will meet the following science objectives:

1.  Fast analysis of the sample (6 minutes)

2.  Trace quantitative elemental composition of different samples, from
    which H, Li, Be, B, C, N, O, Ba, Sr, Cs are very sensible for
    detection.

3.  Detect hydrated minerals.

4.  Characterize geological features of the surroundings.

5.  Provide superficial and subsurface profiles of samples.

6.  Wide range of site investigations allowing several measurements at
    diverse locations, permitting a comparison of elements among
    different targets.

The ChemCam analysis is conducted mainly by two groups of instruments,
the Laser-Induced Breakdown Spectrometer (LIBS) and the Remote
Micro-Imager (RMI). The LIBS is comprised of a powerful laser and three
spectrometers. After selecting the desired sample, the laser beam will
be directed toward the target, removing the outermost layer of the
sample, and inducing atom excitement. This atomic stimulation causes
light emission, which is then measured by the UV spectra, the red
visible spectra, and the blue visible spectra. The following are the
advantages that LIBS presents to the EAGLE Mission:

The LIBS does not require sample collection for analysis.

It operates at distances several meters away from the lander.

The time to analyze the sample and recharge for next operation is short.

By removing the outer layer of the sample, it will provide a depth
profile for a better investigation.

LIBS maximizes the capability of element detection from the sample in
comparison with other instruments.

LIBS minimizes the amount of power required.

The primary advantage of RMI is the production of the highest possible
resolution images without an operational sample. RMI provides a close-up
imaging of the target at long distances, revealing sub-millimeter
structural analysis.

The ChemCam design has three separate units:

Body Unit:

Spectrograph x 3

Power Box

Data Processing Unit

Mass Unit:

Laser

Remote Micro Imager

Telescope

Focus Laser

Calibration Unit:

Target Calibration

The ChemCam executes three steps to collect and analyze data after
target selection:

Monochromatic light illuminates the target. The light is scanned and
monitored at the correct wavelength of the spectrograph, correcting the
focus based on the maximum data fit.

The RMI takes a general image with a resolution of 1024 X 1024 pixels
prior to LIBS analysis, and a 128 X 128 pixel image will be taken after
the LIBS analysis.

The LIBS laser fires at the target approximately 75 times, producing
plasma as it contacts the sample, which is recorded by the spectrograph.
This instrument may conduct further laser tests if additional analysis
is necessary.

The Mars Science Laboratory Rover, which will launch in 2009,
incorporates the ChemCam into the science payload. This instrument
quickly characterizes the chosen sample, and identifies additional
samples without having to move the whole vehicle. The specifications of
this instrument are as follows:

The laser altimeter will determine elevation, and will be especially
important during the mapping and landing phases of the mission.
Elevation will be determined by measuring the time required for the
laser pulse to complete a round trip from the instrument to the surface
and back to the instrument. The laser altimeter design will be similar
to the LD90-3800-FLP, developed by Riegl Laser Measurement Systems. The
altimeter mass will be approximately 4.4 kg, with external dimensions of
232 x 184 x 100 mm, and will require 11-28 Volts DC, with approximately
10 W built-in protecting circuitry for over-voltage and reverse
polarity^44^. The altimeter developed by Riegl Laser Measurements System
will have a range of 750 m if reflectivity \> 0.8, a resolution of 50 mm
and an accuracy of 50 mm.

The laser altimeter will be used in conjunction with Autonomous
Vision-Guided Safe and Precise Landing (AVGSPL) software to land safely
on the surface of Enceladus autonomously. The software is currently
being developed under CICT's Intelligent Systems (IS) Project. Core
components of AVGSPL were demonstrated in the landing of the Mars
Exploration Rovers (MER), using a subsystem known as DIMES - Descent
Image Motion Estimation Subsystem. The software will use a series of
images taken in the final seconds before landing, along with data from
the laser altimeter, to determine altitude. Pre-programmed actions and
reflexes will enable the lander to quickly adapt and respond to changes
in the environment and within itself^29^. Prior to landing, the vision
althgorithms construct an elevation map. Software operators are then
applied to the terrain map to locate regions that are free of steep
slopes and uneven terrain. When a safe landing site is identified, the
lander will enter what is known as "object-track mode" and a state
estimation algorithm will send commands to the autonomous flight
controller. The vision-based controller will command the lander to enter
"landing mode" once it is above the desired landing site^29^.

With the limited data available, it is difficult to select an exact
landing site at the time of this study. However, some potential areas
could be selected from the high resolution images taken by Cassini, and
the next fly-by of Enceladus in 2008 may reveal even more information.
The landing site will be selected using the high resolution images from
Cassini, combined with images taken during the first phase of the
landing mission, during which a considerable amount of time will be
spent orbiting and mapping the moon. The science objectives dictate that
the lander will land in the South Polar region of Enceladus, ideally
close to a plume. Investigation of the "Tiger Stripes" in this region is
also of primary interest to the science mission objectives\--as they may
be relatively younger ice\--and these cracks are potentially closer to
liquid water

Cassini's Dual-Technique Magnetometer (MAG) discovered Enceladus' plume
during a flyby of that moon. By measuring the interaction of Enceladus
with Saturn's magnetic field, the Cassini team was able to detect
deviations that indicated mass ejection from the surface. Following this
measurement, cameras imaged the region to reveal the plume ejecting from
the surface. Magnetometers can also investigate subsurface structure
such as underground oceans. The EAGLE lander will carry a magnetometer
while orbiting Enceladus to map the moon's interaction with Saturn's
magnetosphere more precisely, and thus infer properties of the
subsurface.

EAGLE's magnetometer will be similar to one of the magnetometers flown
on Cassini^21^. The vector/scalar helium magnetometer developed at the
Jet Propulsion Laboratory (JPL) has two modes of operation. The vector
mode measures the field in three dimensions to get a precise direction
of the magnetic field; whereas, the scalar mode measures only the
magnitude when measuring stronger magnetic fields. The instrument must
be placed at the end of a 9 m long boom in order to isolate it from the
spacecraft's mass and electrical currents. The total mass is less than 3
kg plus the mass of the boom and it uses 3 W of power^21^.

Cassini has found strong evidence for the existence of water and organic
molecules on Enceladus^10^. In order to accomplish one of EAGLE's main
objectives, the search for organic molecules, surface sampling is
necessary to confirm Cassini's findings. The Enceladus Surface Analysis
by Mass Spectrometry (ESAMS) experiment will perform in-situ
investigation in order to detect complex organics using a gas
chromatograph and time of flight (TOF) mass spectrometer. Samples will
be delivered to the experiment in two ways; from the surface by SD2
(Section 4.1.1.10), or directly from the plume using the Falling Sample
Collector (FSC) attached to the top of the lander. The main heritage for
this instrument will be from the COmetary SAmpling and Composition
(COSAC) experiment currently flying aboard ESA's ROSETTA mission to
Comet 67P/Churyumov-Gerasimenko^45^. A mass restriction of 5 kg and a
power restriction of 16 W are placed on the instrument design.

As with COSAC, samples will be collected from the surface and feed into
small ovens for pyrolysis. Ten pyrolysers sit on a carousel that will
take the delivered sample from the drill to the GC/MS input. Each small
oven will be used only once to prevent cross contamination. The ovens
are platinum cylinders measuring 6 mm in height and 5 mm in diameter.
Each oven weighs less than 1 gram but is capable of heating to 600 ^o^C
to vaporize the samples. The FCS will also be able to deposit samples
into the carousel. This will be a cup-like collection device with a
surface area large enough to collect a 5 mm^3^ sample within the first
week of landing on Enceladus. The rate of deposition from the plume onto
the surface is not yet known; however, Cassini will be able to deliver
this information as it continues to study Enceladus from Saturn orbit.

Gas from the ovens can be analyzed by gas chromatograph (GC), mass
spectrometry (MS) or in combined GC/MS mode (see Figure 4.1.4.1). For GC
mode, the gas exiting an oven will be moved using He carrier gas stored
in stainless-steel containers. Several columns will be used in the
capillary gas chromatograph, which will be developed to identify a wide
range of masses from noble gasses to organic molecules.

![](media/image7.png){width="3.0076388888888888in"
height="1.7916666666666667in"}

[]{#__RefHeading___Toc145044438 .anchor}Figure 4: Mass Spectrometer Gas
Chromatograph

An electron beam ionizes gas molecules entering the mass spectrometer.
The ions are then accelerated through a constant voltage. The velocities
of ions depend on the mass-to-charge ratio, which can be measured using
the time it takes ions to reach the detector. The mass spectrometer will
have a range between 1-300 amu, which is large enough to detect all
hydrocarbon chains expected to pass thought this instrument^45^. The
calibration process will determine the level of Earth organics in the
mass spectrometer, thereby validating any detection of non-Earth
compounds.

Surface samples will provide the best results, assuming that the freshly
deposited surface will be the least effected by radiation. Furthermore,
the FSC may complement any data by ensuring the purity of the sample.
Samples can also be taken at each hopping site to look at composition
distribution on the surface. All of these measurements will provide
information on the composition of the moon as well as the possibility of
detecting life.

The NIMS will produce a near-IR spectral map of the surface of Enceladus
from orbit. The primary goal of this study will be to examine the
composition of Enceladus' ice at various locations across the surface.
The surface is nearly pure water ice, but carbon dioxide and short-chain
organics are also present in the region of the tiger stripe features.
Some of the primary science objectives focus on further study of the
form and distribution of these organics. The wavelength range also
covers absorption bands of carbon monoxide and ammonia, the latter of
which is important both to interior models of Enceladus and to solar
system formation theory. Additional spectral features will provide
information about the grain size and structure---amorphous vs.
crystalline---of the ice; these parameters can in turn yield a better
understanding of the time and temperature at which the ice formed.
Mapping the entire surface of the moon at near-IR wavelengths will
enable comparative studies of ice properties and trace constituent
composition across the surface, which will facilitate selection of a
landing site.

The Galileo-NIMS instrument^13^ is the model for EAGLE's NIMS. Light
enters the spectrometer through a telescope of aperture 22.8 cm. By
rotating the secondary mirror, the telescope scans in a single
direction, providing 1x20-pixel strips of imaging data. This corresponds
to a field of view of 0.5 x 10 mrad per strip. The light passes through
an optical chopper, creating a dark reference, and onto a diffraction
grating, which disperses the light on its way to the focal plane
assembly. An array of filters and 17 detectors populate the focal plane.
The instrument can measure up to 408 spectral channels, spanning the
range from 0.7 μm to 5.2 μm, by moving the grating so that different
colors fall onto each detector.

The total instrument mass is 18 kg; the optics occupy a volume of 83 cm
x 37 cm x 39 cm, while the electronics occupy 20 cm x 25 cm x 13 cm. The
average power draw is 12 W, and the peak power required is 13 W. The
instrument is passively cooled to 150 K during operations, when it
achieves a data rate of 11.52 kbps.

![](media/image8.wmf){width="3.048611111111111in"
height="2.2840277777777778in"}

[]{#__RefHeading___Toc145044439 .anchor}Figure 5: Near-Infrared Mapping
Spectrometer^13^

The radiometer will image the surface of Enceladus from orbit in the
thermal infrared portion of the spectrum. The global distribution of
temperature across the surface will thus be ascertained, allowing the
identification of hot spots, which may serve as attractive potential
landing sites. In addition, a photopolarimeter is packaged with the
radiometer; this will enable additional studies of the texture and
composition of the surface ice.

The Galileo-PPR instrument^67^ is the model for EAGLE's PPR. Light
enters the spectrometer through a telescope of aperture 10 cm. Separate
detectors and operating modes exist for infrared radiometry, optical
photometry, and polarimetry. For radiometric measurements, a chopper
provides observations alternating between the scene-viewing telescope
and the space-viewing telescope; the latter observes an empty patch of
sky for temperature calibration. All electronic components are carefully
calibrated and have temperature sensors, enabling useful radiometric
measurements without cooling to cryogenic temperatures. For photometry
and polarimetry, the chopper is not used, and a polarizing beamsplitter
separates the light into orthogonal polarizations. All of these modes
yield images with a 2.5 mrad circular field of view. Five spectral bands
ranging from 15 to approximately 100 μm provide broad radiometric
coverage of the thermal infrared, while seven visible/near-infrared
bands are used for photometry, and three visible/near-IR bands for
polarimetry.

The total instrument mass is 5.2 kg; it occupies a volume of 44.8 cm x
39.1 cm x 32.7 cm. The peak power required for operating the instrument
is 11.0 W; its heaters require an additional 4.5 W. PPR can operate
within a temperature range from -20 to 5 °C; to survive while not
operating, it must be maintained between -25 and 30 °C. The data rate
for PPR is 216 bps.

![](media/image9.png){width="3.0479166666666666in"
height="2.188888888888889in"}

[]{#__RefHeading___Toc145044440 .anchor}Figure 6: Exploded view of
photopolarimeter^67^

The ESAMS experiment and the amino acid analyzer instruments on EAGLE
require samples taken directly from the surface or near subsurface. To
fulfill this requirement the payload includes the Drilling and Sample
Delivery System (SD2).^68,69^ The SD2 is an integrated driller, sample
tool, and carousel instrument that will provide samples from near the
surface to a depth of 230 mm. The carousel then distributes the samples
to the appropriate instruments.

The SD2 was developed for the ROSETTA mission, a cometary rendezvous
mission incorporating small masses, while obtaining a reasonable sample
for the other instruments to analyze. The specifications for the SD2 are
the following:

The drill bit may penetrate the surface to a maximum depth of 0.23
meters. Once the desired depth has been reached, the drill bit begins
raising material and placing it above the carousel. This collection
method takes multiple samples between the surface and the depth of
penetration, and supplies them to the instruments for analysis. After
placing the sample in the carousel, it is then transferred to a sealed
oven and begins the ESAMS and Amino Acid analysis.

The primary goals of the seismometers are to probe the physical
structure of the moon\'s interior and determine its tectonic activity.
Furthermore, information gained from the seismometers could potentially
be used to identify sources of heat for the South Pole geysers.  The
lander will hold 2 seismometers, each with a mass of 0.2 kg and
requiring only 0.1 W^70^. The Microdevices Laboratory of JPL is
currently developing these seismometers, which are intended for
extraterrestrial applications. The miniature seismometers measure 2 x 2
x 3 cm (see Figure 7), and their low mass, volume and power requirements
make them ideal for such planetary expeditions. These instruments can
provide high-quality seismic measurements, comparable to that of
conventional terrestrial seismometers.

![](media/image10.wmf){width="3.002083333333333in"
height="4.926388888888889in"}

[]{#__RefHeading___Toc145044441 .anchor}Figure 7: Miniature seismometer

The seismometers will be positioned on the legs of the lander, ideally
operated at a time when lander vibrations are at a minimum; although the
instruments allow for continuous data collection. Key elements of the
seismometer include a suspension mechanism, proof mass and capacitor
plates, and a very sensitive capacitive displacement transducer, which
will use a force-rebalance feedback system.

Seismometers are an excellent method for investigating subsurface
structure\-- especially in the near-subsurface\--where the propagation
of waves is more sensitive to density and degree of compaction. The use
of seismometers could help to identify the presence of water, as well as
different layers in the upper crust of Enceladus^70^.

The lander will have a Panoramic Camera (Pancam) attached to it, similar
to those implemented in the MER mission launched in 2003. The scientific
goals of the Pancam are to obtain color pictures of the surrounding
terrain, to assess the morphology, topography, and geologic context of
the landing site, and to obtain any information regarding the density or
physical properties of the geyser material falling back to the surface.
Simply analyzing images of the landing site can provide information on
the history of its formation. The Pancam can also help in locating
hazardous terrain to avoid during a "hop", and identify potential
relocation sites. The pictures can also be utilized in public outreach
and educational products.

Pancam is a multispectral, stereoscopic, panoramic imaging system
consisting of two digital cameras mounted on a mast 50-70 cm above the
lander deck. On the mast, Pancam can image the full 360° in azimuth and
±90° in elevation. Each Pancam camera has a 1024 x 1024 active imaging
area frame transfer CCD detector array. The cameras are designed to each
use half of their pixels, making for a total active imaging area of 1024
x 2048, while the other half is used as a storage/readout area masked
from illumination by an opaque black-painted aluminum light shield.^5^

Multispectral panoramic imaging provides information on the
mineralogical composition of surface materials. It is expected to find
mostly ice, but it has the capability of detecting any possible mineral
abundances that might have been deposited as meteorites. Current studies
of Enceladus find that the surface is mostly covered in water ice;
however, should the lander discover any rocks, the Pancam has the
ability to study their reflectance properties to constrain surface grain
size, degree of compaction, macroscopic roughness, and particle internal
structure. The Pancam can also be used for solar imaging. Both solar
filters (L8: 440 nm and R8: 880 nm) are sensitive to dust and large ice
particles, while the blue solar filter is also sensitive to small ice
particles. This will help constrain the opacity of ice particles near
the plume where they will be falling back towards the surface.

The Pancam optics have an effective focal length of 43 mm and a focal
ratio of f/20, yielding an instantaneous resolution of 0.27 mrad/pixel
and a field of view of 16°x 16°. The two Pancam ''eyes'' are separated
by 30 cm and have a 1° toe-in to provide adequate stereo parallax.^5^
Toe-in results in 100% image overlap at a distance from the camera
proportional to the toe-in angle, in this case 10 m. The 30 cm
separation maximizes the accuracy of far-field ranging, which places the
lower limit of its range at 1.5 m. If later research shows that higher
accuracy for closer images is desired, the separation distance can be
shortened.

Each camera has a mass of about 0.27 kg and operates from 218---278 K.
The thermal system on the Enceladus lander will ensure the Pancam
remains within this range. The minimum non-operating temperature for all
components is 163 K, while the maximum is 328 K. A small 3.5 W (maximum)
strip heater is attached to each electronics box and filter wheel to
provide survival and warm-up heating if required. Each camera requires
+7.1 V and --10.4 V from the lander power supply. Power usage during
full-frame imaging is 1.4 W while idle, 2.4 W during integration and
readout, and 3.8 W maximum during CCD fast flush. Each full-frame image
generates 12.98 Mbits of data. The CCD frame transfer time is 5.2 ms;
the readout rate is 200 kHz. This leads to a total readout time of
approximately 5 s per full frame image.^5^

Each camera has an eight-position filter wheel in which thirteen of the
slots were designed to take images of the surface or to look skyward,
two were designed for direct imaging of the Sun, and one was left empty
to maximize sensitivity during low-light and ambient Earth temperature
(preflight) imaging conditions. Filters range from 440 nm to 1000 nm. A
short sunshade and set of black internal baffles provide rejection of
stray and scattered light.^5^

A calibration target is included on the lander to verify and validate
the preflight calibration of the Pancam imaging system and to monitor
the stability of the calibration during the mission. It is located so
the Pancam has an unobstructed view of it when deployed and is as far
away as possible for best focus. The target is 8 x 8 cm^2^ with a 6 cm
high black vertical post for casting a shadow. Measurements in and out
of the shadowed region provide a way to quantify the direct versus
diffuse components of incident scene radiance. The surface of the
calibration target contains three gray level annuli of 20%, 40%, and 60%
reflectivity ("black," "gray," and "white" rings, with the white ring
farthest from the post), and four pieces of pigmented "color chip"
materials in the corners. The calibration target materials consist of
silicone RTV pigmented with either titanium dioxide or carbon black to
raise or lower the reflectivity, or submicron powders of hematite,
goethite, chromium oxide, or cobalt aluminate to generate the red,
yellow, green, and blue color chips, respectively. The materials were
UV-irradiated to handle 30 Martian sols for the MER mission, to minimize
any UV degradation of the reflectance levels during operations. It will
provide calibrated reflectance and color surfaces that will be used to
validate, monitor, and refine Pancam calibration.^5^

The Pancam system is a science instrument designed by members of the
Athena science team at Cornell in cooperation with engineers and
scientists at the Jet Propulsion Laboratory (JPL) and Ball Aerospace.
Pancam components were constructed by JPL, Ball, and many other vendors,
and fabrication and integration of the system occurred at JPL.

An essential element of any outer planetary mission for both scientific
reasons and public relations is the color imager. The Saturn orbiting
phase of the mission will be a bare bones satellite mainly to provide
communications. However, while in Saturn's system, the spacecraft will
be in a perfect position to collect stunning images like those received
with Cassini's Imaging Science Subsystem (ISS). The chosen orbit around
Saturn allows flybys of the inner moons out to Iapetus, including Titan,
Rhea, Hyperion, Dione, Tethys, and of course Enceladus. EAGLE will have
a wide-angle camera capable of taking color images, mainly for press
releases, but also to expand on Cassini's imaging science.

The SOC will have a 3.5^o^ field of view with a 60 μrad/pixel
resolution, similar to Cassini's ISS Wide Angle camera^41^. This should
allow a full view of Titan during several flybys and more detailed
images with closer approaches in the orbit. It will have 18 filters
covering a range of 380-1100 nm. SOC will be an f/3.5 with a 20 cm focal
length. The detector is a square array of 1024 pixels. The total mass of
the system will be around 25 kg, fitting in a cylinder approximately 80
cm in length by 40 cm in diameter (both of these figures include the
associated electronics)^41^. During the cruise stage, SOC will image
Venus and Earth during gravity assists, ideally with enough bandwidth to
record several image sequences that can be complied into video clips.

![](media/image11.png){width="2.9923611111111112in"
height="2.247916666666667in"}

[]{#__RefHeading___Toc145044442 .anchor}Figure 8: Cassini's Wide Angle
Camera.

The only other science instrument on board the Saturn orbiter besides
the Saturn Orbiting Camera (SOC) will be a payload designed and built by
students. Mass (10 kg) and power (5 W) will be reserved in the
spacecraft design to fill a volume 20 cm x 20 cm x 50 cm on the outside
of the orbiter. The science goals of the payload will be at the
discretion of the team chosen to build the instrument. Section 8.1
outlines the procedure for team selection.

The orbiter communications system is equipped with an ultra-stable
oscillator that has the ability, when coupled with the ground antennas
of NASA\'s Deep Space Network (DSN), to perform radio science
experiments. Although Enceladus has only a very thin global atmosphere,
the DSN can detect refraction, Doppler shifts, and other radio signal
modifications during occultations to determine the temperature and
composition profiles of the atmosphere. Furthermore, Doppler and two-way
ranging measurements generated by the DSN allow for investigation of the
Enceladus gravity field and ephemeredes. A host of other radio science
experiments, including tests of general relativity and experiments in
solar science, may be performed during the interplanetary cruise phase
of the mission.

Given additional mass, power, and a mode of communication, seismic
packages similar to those designed for the Mars Netlander SEIS
experiment could be deployed from orbit around Enceladus. The EAGLE
lander will carry two seismometers on its legs, but additional
information could be gained by deploying seismometers around the
equator. Three seismic packages could be deployed during the mapping
phase of the mission, distributed around the equator of Enceladus rather
than the pole. The objectives for the Netlander seismometers are to
perform both seismic and tidal measurements, in an effort to determine
if the South Pole geysers are the result of tidal stresses. Each
seismometer will have a mass of 2.3 kg, a diameter of 15 cm and will
require 1-3 W of power^49^. As in the proposed Netlander experiment,
each seismometer will send a data volume of approximately 2.5
Mbits/day^49^. An extensive seismic experiment such as this one could
better determine the interior structure of the moon, and perhaps lead to
the discovery of liquid water, severe tidal stressing, or a dynamic
core.

The purpose of planetary protection is to protect life, remnants of past
life, and the precursors of life on other solar system bodies, so that
they can be discovered by future exploration missions (forward
contamination). Planetary protection also involves backward
contamination, which protects the Earth from contamination by sample
return missions. Since the EAGLE mission does not include any sample
returns, it only needs to address protection against forward
contamination.

The EAGLE mission is a Category IV mission since it is targeting a body
with origin-of-life interest in which a lander presents a significant
chance of contamination that could jeopardize future biological
exploration. Requirements for a Category IV mission include detailed
documentation, bioassays to specify the burden, a probability of
contamination analysis, an inventory of the bulk constituent organics,
and an increased number of implementing procedures. The latter may
include trajectory biasing, the use of clean rooms (Class 100,000 or
better) during spacecraft assembly and testing, bioload reduction,
possible partial sterilization of the hardware having direct contact
with the target body, a bioshield for that hardware, and, in rare cases,
a complete sterilization of the entire spacecraft.^46^

In-depth studies for the planetary protection requirements of an orbiter
or lander for Enceladus have not yet been defined; however, for Europa,
the constraint is to reduce the probability of inadvertent contamination
of a Europan ocean to less than 1x10^-4^ per mission. Modifying the
critical factors to apply to Enceladus at a very general level, the
considerations become:

1.  Microbial burden at launch.

2.  Cruise survival for contaminating organisms.

3.  Organism survival in the radiation environment adjacent to
    Enceladus.

4.  Probability of landing on Enceladus.

5.  The mechanisms of transport to the subsurface of Enceladus.

6.  Organism survival and proliferation before, during, and after
    subsurface transfer.^38^

Preliminary calculations of the probability of contamination suggest
that microbial reduction will likely be necessary for Europa orbiters as
well as for landers. Since there is a significantly less harsh radiation
environment in the Saturnian system than the Jovian system, there is a
greater chance for contamination and thus any derived requirements for a
Europa mission will be absolutely necessary for Enceladus. This will
require the use of clean room technology, the cleanliness of all parts
before assembly, and the monitoring of spacecraft assembly facilities to
understand the bioload and its microbial diversity, including specific
problematic species. Specific methods should be developed to eradicate
problematic species. Methods of microbial reduction should reflect the
type of environments found on Enceladus, focusing on Earth extremophiles
most likely to survive on Enceladus, such as cold and radiation tolerant
organisms.

As defined by federal standard, NASA requires that planetary protection
procedures involving sterile items and sample processing must be
conducted in Class 100 clean rooms. These clean rooms have
laminar-air-flow systems to filter out contaminants, which work by
keeping the air within a space moving in one direction along parallel
flow lines at a uniform velocity through very fine filters. Planetary
protection procedures specify the types of devices that may be used for
air sampling in these environments. ^46^

Those who execute the planetary protection procedures during assembly,
testing, and launch operations must wear protective clothing to prevent
contamination. This includes using hoods, masks, surgical gloves,
booties, and bunny suits. For the EAGLE mission, active microbial
barriers will be necessary to protect against the recontamination of the
spacecraft after microbial reduction processing. Further studies must be
conducted to determine whether to operate under high pressure or ambient
pressure. High-pressure planetary protection schemes involve maintaining
a barrier at a specific, continuous static pressure level above ambient
air pressure to prevent microbiological recontamination. Air must flow
from inside a pressurized microbial barrier toward the outside. If it
operates at ambient air pressure, filters capable of retaining 99.97
percent of all particles or organisms greater than 3 x 10^-7^ m in size
must be employed to protect against microbiological recontamination. As
described by the Planetary Protection Office, passive microbial barrier
methods include the use of sterile stainless steel fallout strips and
Teflon fallout ribbon to collect samples of airborne microbial
contamination accumulating on surfaces. These strips or ribbons are
assayed periodically. ^46^

Currently, dry heat microbial reduction is the only NASA-approved method
for spacecraft sterilization. The Viking Mars landers were cleaned to
reduce their total surface biological burden to a specific level, then
packaged in a fully enclosed bioshield to be baked in an oven at 111.7°C
for 30 hours. Because some advanced equipment is heat-sensitive and
could be damaged by such high-temperature treatment, NASA researchers
are developing a low-temperature vapor-phase hydrogen-peroxide-based
sterilization process for the outer surface of the spacecraft.^46^ This
would not be sufficient for the entire payload to Enceladus, though,
since the buried and mated bioburden would not receive treatment. Since
some of the instruments were designed for the MER mission, they were
already designed to handle dry heat microbial reduction. It would be
beneficial as the EAGLE project continues to design or modify the rest
of the hardware to handle it as well. Providing hardware covers where
possible would also improve protection.

One dilemma with the EAGLE mission is that after all the components'
masses are added together, while accounting for the mass margin
including a bioshield would exceed the lifting capability of the launch
vehicle. To moderate the issue, it has been suggested to surround the
spacecraft with a Tyvek or Tedlar bag before putting it into the
fairing. The fairing could then be used to protect the "unstiff"
bioshield. The bag is necessary because it would be very difficult to
sterilize the fairing's acoustic blankets without damaging the
spacecraft. The spacecraft could then be treated in the Tyvek/Tedlar
bag, the bag could be affixed to a reasonably clean fairing, and a
mechanism could ensure that the bag splits and goes away with the
fairing after launch. The fairing would have to undergo more stringent
sterilization than the normal requirement.

The Science subsystem selected this specific payload in order to
accomplish the mission objectives. With these instruments, the Saturn
orbiter is capable of imaging Saturn's moons for public release and
possibly creating short image sequence videos. The orbilander will be
able to map Enceladus' surface and gain information on the subsurface
structure. And most importantly, the lander will be able to measure
surface composition, image the surface, and detect organic molecules and
potentially life. A large majority of these instrument specifications
are taken from instruments that have already flown or are already
designed for future missions. The Science team assumes that technologies
will improve prior to launch in 2023, which will lead to more
sensitivity and less massive instruments. This section is intended to
give an overview of the payload capability with the understanding that
many of the specifics will change as further work yields more exact
requirements, and as Cassini continues to gather information on
Enceladus during future flybys.

The mission architecture team primarily used JAQAR^®^ Space
Engineering's Swingby Calculator and Analytical Graphics^®^ Inc.'s
Satellite Tool Kit (STK) with Astrogator module to design an optimized
trajectory from Earth to the Saturn system. This trajectory minimizes
the ∆V needed to reach the desired orbit in the Saturn system by making
two gravity assists at Venus followed by one gravity assist at Earth. At
Saturn orbital insertion (SOI), the EAGLE spacecraft will follow a
trajectory similar to the Cassini/Huygens mission, passing through the
gap between the F and G rings with a minimum altitude above Saturn's
clouds of 20,000 km. This close approach reduces the fuel needed to
achieve Saturn capture. Once captured by Saturn, the spacecraft will
adjust to an elliptical mission orbit with periapsis matching the
orbital radius of Enceladus and apoapsis at the orbital radius of
Iapetus. At closest approach to Enceladus, the EAGLE spacecraft will
separate into an orbiter and a lander. The orbiter will remain in the
elliptical Enceladus-Iapetus orbit and will serve as a communication
link and imaging platform. The lander component will orbit Enceladus and
map it with onboard instruments for several days before landing and
performing analysis of its surroundings.

The EAGLE mission will launch on a Delta IV (4050H-19) from SLC-37 at
Cape Canaveral. This vehicle demonstrates the best performance out of
those considered in Figure 9.

![](media/image12.png){width="3.0194444444444444in"
height="2.0180555555555557in"}

[]{#__RefHeading___Toc145044443 .anchor}Figure 9: Performance comparison
of 5 launch vehicles for high energy (earth escaping) orbits. ^15^

The performance estimates include several assumptions. Table 1 lists the
vehicle-specific assumptions. In addition, the estimates assume a 185 km
altitude perigee and 28.6^o^ inclination for the parking orbit and
therefore a 185 km perigee for the escape orbit for all vehicles.
Estimates do not include orbital debris compliance, and all assume a
3-sigma Stage-2 commanded shutdown probability, plus additional reserves
determined by launch site procedures.

![](media/image13.wmf)[]{#__RefHeading___Toc145044470 .anchor}**Table 1:
Assumptions involved in launch vehicle performance estimates**

EAGLE can potentially remain in its parking orbit for several days,
thereby increasing the duration of the launch window. Figure 10 shows a
typical Delta IVH launch profile to LEO.

![](media/image14.png){width="2.9930555555555554in"
height="2.2895833333333333in"}[]{#__RefHeading___Toc145044444 .anchor}\
**Figure 10: Delta IVH launch profile to LEO.^8^ A final burn from the
185 km parking orbit will send the EAGLE spacecraft into the
interplanetary cruise phase.**

Once EAGLE is in the 185 km altitude parking orbit, Swingby Calculator
predicts a required C~3~ of 12.561 km^2^/s^2^ to escape Earth on the
appropriate trajectory to the Saturn system. This C~3~ corresponds to
7600 kg in the earth-escape trajectory. The current EAGLE mission
requires 6867.1 kg. This gives a mass margin of approximately 10%.

The EAGLE spacecraft will make use of multiple gravity assists to reduce
the ∆V necessary to reach Saturn. Jupiter is only in a favorable
position for gravity assist to Saturn approximately every 20 years, and
the next two favorable alignments occur in roughly 2018 and 2038.^30^
These dates are somewhat too soon and much too late, respectively, for
the EAGLE mission. Consequently, the trajectory team focused on Swingby
Calculator optimized trajectories that do not use a Jupiter assist.
Venus-Venus-Earth and Venus-Earth-Earth assists are popular for such
missions.

The relative positions of the inner planets present a nearly ideal
launch opportunity in June of 2023. The EAGLE mission will take
advantage of this alignment, which governs EAGLE's primary launch
window. The final trajectory is shown in Figure 11 and involves two
assists at Venus and an assist at Earth. Between the Venus swingbys,
there is a small deep space maneuver of 16.078 m/s. The cruise time from
launch to arrival at Saturn is 2550 days, almost exactly 7 years.
EAGLE's altitude above the gravity assist body at nearest approach is
given in Table 2.

![](media/image15.png){width="6.485416666666667in"
height="2.1368055555555556in"}

[]{#__RefHeading___Toc145044445 .anchor}Figure 11: EAGLE's
interplanetary trajectory

![](media/image16.wmf)[]{#__RefHeading___Toc145044471 .anchor}**Table 2:
Swingby altitudes and dates**

In May of 2030, EAGLE will arrive in the Saturn system (Figure 12),
traveling at approximately 10.4 km/s relative to the sun and 740 m/s
relative to Saturn. Once within Saturn's gravitational well, EAGLE will
accelerate considerably. The primary objective during the Saturn
approach phase is to counter this acceleration efficiently and in a
magnitude sufficient to allow for capture by Enceladus.

![](media/image17.jpeg){width="3.051388888888889in"
height="1.4847222222222223in"}

[]{#__RefHeading___Toc145044446 .anchor}Figure 12: Enceladus orbit is
highlighted in red in this view of the Saturn system on the date of
EAGLE's arrival into the Saturn system.^30^

Enceladus is no easy target. Indeed, the need to be captured by
Enceladus presents one of the most challenging tasks in celestial
mechanics. The moon orbits hardly one Saturn-radius beyond Saturn's
outer rings^22^ and carries a mass just more than one-thousandth that of
our moon^39^. As EAGLE sinks into Saturn's gravitational well, a 6 km/s
burn will be required to become captured by Enceladus (Figure 13). Such
a burn requires very large thrusters and produces considerable loss in
efficiency due to its long duration. Correspondingly, EAGLE will steer
clear of a direct capture by Enceladus, in favor of a second solution.

![](media/image18.jpeg){width="3.0756944444444443in"
height="1.2138888888888888in"}

[]{#__RefHeading___Toc145044447 .anchor}Figure 13: Direct capture by
Enceladus, as shown in the above figure, is too costly and too long a
maneuver to perform.^30^

![](media/image19.png){width="3.071527777777778in"
height="3.1034722222222224in"}[]{#__RefHeading___Toc145044448
.anchor}**Figure 14: EAGLE's SOI burn occurs above the C-Ring, 20,000 km
above Saturn's clouds. Eagle passes between the F and G rings twice.**

The second solution requires a closer approach to Saturn. As mentioned
earlier, efficiently reducing the craft's energy is the primary goal of
the approach phase. One can easily show that the greatest decrease in
energy can be affected when the burn is performed at maximum orbital
velocity. In light of this fact, EAGLE will perform its primary Saturn
Orbital Insertion (SOI) burn above Saturn's C-Ring, 20,000 km from
Saturn's cloud tops. This burn at periapsis reduces EAGLE's velocity by
910 m/s, down to 30.4 km/s, and allows for capture of EAGLE within its
commissioning orbit (Figure 15). In order to avoid collision with large
ring particles, the mission analysis team has designed EAGLE's
hyperbolic trajectory with a slight inclination relative to Saturn's
ring plane (Figure 16). The craft will fly between Saturn's F and G
rings. The orbiter will remain in this slightly inclined orbit during
the commissioning phase.

![](media/image20.png){width="3.013888888888889in"
height="1.6416666666666666in"}

[]{#__RefHeading___Toc145044449 .anchor}Figure 15: Orange represent's
EAGLE's commissioning orbit, and green represents EAGLE's mission orbit.
Enceladus' orbit is shown in red.

![](media/image21.png){width="3.0104166666666665in"
height="1.613888888888889in"}[]{#__RefHeading___Toc145044450
.anchor}**Figure 16: EAGLE passes through Saturn's rings, with periapsis
slightly above the ecliptic plane.**

The SOI burn places EAGLE in a highly eccentric orbit, with apoapsis
near Saturn's moon Iapetus. Iapetus' orbit was selected for apoapsis due
to the outstanding scientific questions posited by Iapetus' surface
features. At apoapsis, EAGLE burns its main thrusters once more to bring
its periapsis safely beyond Saturn's rings. This approximately 470 m/s
burn positions the EAGLE orbiter in its mission orbit, from which it
will be capable of providing communications between the EAGLE lander and
Earth. After reducing the mild inclination with respect to the ecliptic
plane, the mission orbit permits EAGLE to flyby several of Saturn's
natural satellites, including Titan, Iapetus, Hyperion, Dione, and Rhea,
all of which have attracted considerable scientific interest.

![](media/image22.png){width="3.032638888888889in"
height="1.1680555555555556in"}

[]{#__RefHeading___Toc145044451 .anchor}Figure 17: Rhea's orbit, in
blue; Titan's orbit, in orange; Iapetus' orbit, in purple. Note the
inclined nature of Iapetus' orbit.

As shown in Figure 17, Iapetus is highly inclined with respect to the
ecliptic. In order to flyby Iapetus, EAGLE's argument of periapsis must
be set such that apoapsis coincides with Iapetus' longitude of the
ascending node. The mission orbit satisfies this requirement.

![](media/image23.wmf)[]{#__RefHeading___Toc145044472 .anchor}**Table 3:
Summary of Saturn system trajectory burns.**

![](media/image24.png){width="3.279861111111111in"
height="1.8256944444444445in"}

[]{#__RefHeading___Toc145044452 .anchor}Figure 18: Full Saturn approach
trajectory.

For this mission proposal, the lander is considered to be the primary
science package, while the prime function of the orbiter is considered
to be that of communications relay. The long orbital period of the
mission orbit increases the communication link time available. The
orbiter also carries a wide-angle camera, but science data from this
camera is a secondary objective. Flybys of Titan, Iapetus, and Rhea will
be done as the opportunity presents itself. The mission design team
expects approximately three Titan flybys every two years, one Iapetus
flyby every four years, and frequent Enceladus and Rhea flybys. Such
flybys serve several purposes. The flybys may be used to adjust the
mission orbit's inclination and other parameters, as well as collect
images in the 380-1100 nm range.

The best places for any desired station-keeping would be as the orbiter
passes near Enceladus, Titan, or Iapetus. Burns at these locations will
enable closer flybys and high resolution pictures of some of Saturn's
more interesting moons. As a result of gravitational assists, maneuvers
at these locations will also provide a ∆V bonus.

Near the apoapsis of EAGLE's Saturn-centered commissioning orbit, the
spacecraft performs two maneuvers. The first, as has been previously
mentioned, is of 471 m/s in magnitude and circularizes the commissioning
orbit into the mission orbit (Figure 18). The EAGLE lander separates
from the orbiter after this burn and prepares to target Enceladus. The
second maneuver modifies the orbital period slightly to ensure arrival
at periapsis just as Enceladus arrives. Due to Enceladus' short orbital
period (1.37 days) relative to EAGLE's mission orbit period (1 month)
this targeting burn is negligible in magnitude. Two weeks after these
targeting burns, the EAGLE lander's nominal mission trajectory calls for
an Enceladus Orbital Insertion (EOI) burn of ∆V of 4.66 km/s. Nominally,
EOI will occur above the north pole of Enceladus, positioning EAGLE in a
polar orbit. In actually, this burn is sufficiently long that it will
have to be broken into several burns, the first of which intended to
merely reduce the eccentricity of the Saturn-centered orbit. Future
analysis, discussed below, will pursue these options. After performing
the EOI burn, the large motor and expended fuel tank will be jettisoned.

Once in a circular, polar orbit at a 50 km altitude, the EAGLE lander
will survey the terrain for potential landing sites over a three to four
Earth-day period (20-30 orbits). The polar orbit provides access to
Enceladus' southern regions, which are of considerable scientific
interest. Saturn's significant third-body effects, the uncertainty in
Enceladus' gravitational field and EAGLE's limited fuel supply, limit
the EAGLE lander to an orbit time on the order of days to weeks. After
Enceladus' southern regions have been mapped, the EAGLE lander will
begin its entry, descent, and landing phase.

The orbital velocity around Enceladus at 50 km altitude is only 154.42
m/s. To land, in the ideal case, the orbiter will make a 7.13 m/s burn
to transfer from a 50 km circular orbit to an elliptical orbit with
periapsis at the surface. It will then perform a 176.5 m/s burn to
reduce its velocity to zero at periapsis.

While a launch on June 12^th^ of 2023 is optimal, weather and other
uncontrollable events may prevent an optimal launch. Launches prior and
after this date require increased energy expenditure. EAGLE may launch
on any day prior to or after June 12^th^, provided that the launch
vehicle can produce the additional energy required to match the optimal
trajectory. The Delta IV (4050H-19) upper stage generates sufficient
impulse to propel EAGLE beyond Earth's sphere of influence with a C~3~
of 13 km^2^/s^2^ (The Delta IV Heavy can actually accelerate the EAGLE
payload to a C~3­~ of \~17 km^2^/s^2^. 13 km^2^/s^2^ is a conservative
value which takes into account a significant mass and fuel safety
margin). Correspondingly, EAGLE's launch opportunities exist on any date
between June 4^th^ and June 22^nd^ (Figure 19). Were a mission
controller to choose to launch before June 4^th^, EAGLE can remain in
its parking orbit with a 185 km for several days before boosting to
Venus for its first gravity assist. This option may potentially extend
the launch window's opening to late May.

![](media/image25.png){width="3.0208661417322835in"
height="2.982677165354331in"}![](media/image26.png){width="0.7771653543307087in"
height="0.8464566929133859in"}

[]{#__RefHeading___Toc145044453 .anchor}Figure 19: The alignment of the
inner planets constrains EAGLE to a Venus arrival on December 3^rd^,
2023.

Given the constraint on Venus arrival, the above graph visually depicts
the available launch window (beneath the purple curve). Departure dates
between June 4^th^ and June 21^st^ provide an arrival on December 3^rd^
within the C~3~ constraint. The key is C~3\ ­~in km^2^/sec^2^.

Currently, EAGLE's orbital insertion about Saturn closely resembles that
of the Cassini-Huygens mission. Cassini-Huygens passed through the gap
between the F and G rings twice to closely analyze Saturn's ring system.
However, EAGLE is not designed to study the rings and this insertion
maneuver risks collision with ring particles. Therefore, other potential
trajectories are being considered. As mentioned in an earlier section,
Eagle can save fuel by performing the orbital insertion burn as close to
Saturn as possible. One potential trajectory deserving future study
would have EAGLE perform the burn 5000 km above Saturn's surface, within
the inner most ring. While most efficient, this maneuver lacks heritage
and consequently adds risk. A more feasible trajectory deserving
consideration sets the orbital insertion burn at an altitude of 85000
km, between Saturn's F and G rings. While slightly less efficient than
our current trajectory, this trajectory holds several benefits over the
current. First, EAGLE must pass through Saturn's rings only once, rather
than twice. Second, the inclination of the orbit required to avoid the
rings is less, reducing the fuel required to return EAGLE to the
ecliptic plane. Future studies will analyze these potential orbital
insertion maneuvers.

Currently, all maneuvers are modeled as impulsive. In future studies,
the SOI, EOI, and EDL burns will be modeled as finite. Finite models may
increase fuel mass requirements by as much as 20%, and EAGLE has been
designed with margin to account for these non-idealities. In particular,
the mission trajectory team will likely break the 4.5 km/s ∆V EOI into
several more manageable burns. EOI-1 would reduce the eccentricity of
the Saturn-centric orbit. Enceladus will then capture EAGLE in a highly
eccentric orbit after the EOI-2 maneuver. Several finite circularization
burns will then lower the orbit to a 50km altitude, circular, near-polar
orbit. Due to Saturn's large third body gravitational effect, EAGLE's
orbit will be highly unstable and require intensive station keeping
prior to arrival in the optimally stable 40-50 km parking orbit.

The EAGLE spacecraft will launch on a Delta IV (4050H-19) with a C~3~ of
12.561 km^2^/s^2^. During the interplanetary cruise phase, the
spacecraft will perform two gravitational assists at Venus and an assist
at the Earth to minimize the ∆V necessary to reach Saturn. After a
cruise time of 7 years, EAGLE will enter into an elliptical orbit around
Saturn, passing as close as possible to the planet to maximize the
efficiency of the orbital insertion burn. The total ∆V for the
trajectory from earth escape to Saturn capture is 4.47 km/s.

Once captured, the spacecraft will execute a 0.47 km/s burn to enter a
mission orbit with periapsis at the orbit of Enceladus and apoapsis at
the orbit of Iapetus. It will then separate into two components: an
orbiter that will remain in the Enceladus-Iapetus orbit and serve as a
communications relay and imaging platform, and a lander. The lander will
execute a 4.66 km/s burn to enter into a 50 km polar orbit around
Enceladus. It will remain in this orbit for approximately 3 days to map
the surface in detail. Finally, the lander will target a site at the
tiger stripes on the South Pole. It will take a ∆V of 183.63 to reach
the surface.

The EAGLE lander will include instruments primarily to search for
evidence of liquid water and organic compounds, and investigate the
physical properties of the surface and subsurface. The systems design
goal is to satisfy the programmatic requirements, while balancing the
requirements and constraints of the other subsystems. Achieving a
successful systems design requires managing the mass, power, and cost
data budgets.

Enceladus is a moon with significant interest to the scientific
community, with the potential to generate noteworthy discoveries and
enhance the current understanding of the solar system. Therefore, the
design team decided to develop a flagship mission to investigate as many
science objectives as possible.

The EAGLE mission will launch aboard a Delta IV 4050H-19, with a launch
C­~3~ of 12.6 km^2^/s^2^. The transfer trajectory to Saturn uses
Venus-Venus-Earth assists, and the lander platform will utilize
retrorockets to descend to the surface of Enceladus.

As a flagship mission to the Saturn system, there is a high degree of
commonality with the Cassini-Huygens mission; therefore, many aspects of
the design trace their flight heritage to Cassini. While this heritage
does not significantly affect the cost of this mission, it may reduce
the required contingency in the design to the typical 5-15%.

The timeline for this mission extends far into the future; in order to
ensure the total mission costs remain within NASA's budget, the Phase-A
mission design will begin in 2013, or ten years before the scheduled
launch date. After the seven-year cruise phase, and six month primary
mission duration, the projected end-of-life date is 2030. Based on this
timeline, the total mission cost is \$1.13 billion.

The EAGLE spacecraft consists of three separate components: the Booster,
Lander and Orbiter. The three sections remain as one unit during launch
and while in transit until Saturn orbit insertion. From there, the
Orbiter continues to circle Saturn while the Lander and Booster detach
and are placed in orbit around Enceladus. After orbiting for several
days, the spent Booster stage detaches and the Lander begins its descent
to the surface of Enceladus.

The orbiter is a small spacecraft consisting of two instruments and its
own power, command, communications and GN&C subsystems. This section of
the spacecraft controls the EAGLE during transit and until
Lander/Orbiter separation. At this point, the orbiter will continue
around Saturn and serve to relay communications back to Earth. In
addition, to the Saturn orbiter will take images with the
high-resolution camera and record data from the student instrument.

The Booster section consists of the magnetometer instrument, two
thrusters and accompanying fuel tanks. This section of the spacecraft is
expendable and will be used to accomplish the burns needed for Saturn
orbital insertion and Enceladus orbital insertion. Once in orbit around
Enceladus, this section will deploy the magnetometer instrument on a
boom and take measurements for several orbits before being jettisoned.
The Booster will be dependent on the Lander for power, command and
communications.

The Lander section contains the remaining instruments as well as power,
communications, and GN&C subsystems. This section of the spacecraft
becomes independent of the orbiter after separation and is then inserted
into Enceladus orbit. The lander eventually makes its way to the surface
using three sets of thrusters. Once on the ground, the Lander will raise
its mechanical tower to obtain a better view of the surrounding area
with two sets of cameras in stereo as well as the ChemCam instrument.
Underneath the Lander, several instruments will deploy including the
APXS, which will be lowered down on an elevator arm, and the IR
Spectrometer/Amino Acid Analyzer, which will measure drill samples. The
nominal mission design will be 6 months

The EAGLE spacecraft is designed to support the science payload and
other subsystems hardware. The structure adheres to size and mass
budgets and protects the spacecraft from the harsh space environment.
The spacecraft experiences a variety of environments during transit and
on the surface of Enceladus, and it must be optimized to meet these
demands. It must also survive assembly, test, and launch operations
(ATLO), fluctuating radiation levels during the cruise phase,
micrometeoroid bombardment in Saturn's rings, and the intense cold of
deep space and on the surface of Enceladus. While satisfying these
requirements, the spacecraft fits inside the launch fairing, meets the
mass budget, and completes all scientific mission objectives during the
mission.

The need for two separate platforms and a detachable propulsion stage
drives the EAGLE design. Each platform has its own power supply,
communication, and attitude control. After determining the lander
platform could not support the High Gain Antenna (HGA) and the mass of
the fuel tanks, the propulsion stage was integrated into the spacecraft
configuration. The propulsion stage is jettisoned prior to landing on
Enceladus, saving considerable mass and ΔV. Figure 20 illustrates the
proposed launch configuration.

![](media/image27.jpeg){width="2.939583333333333in"
height="3.8368055555555554in"}

[]{#__RefHeading___Toc145044454 .anchor}Figure 20: Launch configuration,
including the interplanetary insertion propulsion stage.

The design is a robust vehicle capable of withstanding launch load
profiles while minimizing the mass. Of primary concern is the large
diameter of the HGA. The proposed structural design places the antenna
on the top (+z axis) of the spacecraft in order to minimize the stress
experienced during launch. This configuration also allows the HGA to act
as a sun shield during the Venus flybys. The propulsion system is on the
opposite side of the spacecraft (z axis) to isolate the payload from
vibration and reduce shear stress. During launch, a spacecraft can
experience up to 135 dB in vibration if a metal fairing is used. As a
result, EAGLE uses passive damping by using visco-elastic materials
(VEMs).

The spacecraft bus structure optimizes the placement of the science
payload and meets the mass budget for the lander. The most efficient
design is a three-legged configuration, leading to a hexagonal bus
structure, 1.2 m on a side. This configuration allows for a simpler
attachment of the legs and offers a large inner volume for the payload.
The orientation of the spacecraft in orbit dictates that the science
instruments utilized in orbit and the instruments for landed operations
be on the same surface. The hexagonal bus provides sufficient surface
area to support the entire payload with the proper shielding.

The primary function of the Saturn orbiter is to serve as a
communications relay for the lander. The orbiter bus and spacecraft
structure are designed to support the mass of the communications dish
and create a simple attachment for the lander. To maintain pointing
accuracy during communications windows with either Earth or the lander,
the orbiter requires an accurate attitude control system. Hydrazine
thrusters are attached to extended booms which will act as moment arms
to improve efficiency. This allows the vehicle to use less fuel per
maneuver. The two SRG power units are located on the exterior of the
orbiter to avoid overheating the payload.

![](media/image28.jpeg){width="3.2527777777777778in"
height="2.329861111111111in"}

[]{#__RefHeading___Toc145044455 .anchor}Figure 21: Orthogonal view of
the orbiter

![](media/image29.jpeg){width="3.0in" height="2.158333333333333in"}

[]{#__RefHeading___Toc145044456 .anchor}Figure 22: Top view of the
orbiter

The science payload drives the complexity of the lander design. The
majority of the science instruments need to be placed on the bottom (-z
axis) of the spacecraft for instruments to map Enceladus while orbiting,
and the landed experiments to have surface access. To accommodate all
the science instruments, the lander has a large surface area and is
short enough to be stable while on the surface. Some of the scientific
instruments are very large and extend out well below the bottom of the
lander bus. Therefore, the EAGLE design includes legs to protect the
instruments while landed. Deployable legs are not necessary on the
lander because, unlike other lander designs, the EAGLE spacecraft does
not have to fit within a heat shield during entry, descent, and landing
(EDL). Reducing the number of deployable mechanisms also increases the
chances of mission success. The static legs are also advantageous to the
science mission objectives by allowing seismometers to be placed on the
footpads of the lander. This eliminates the need for a separate
seismometer deployment mechanism. The leg spring system dampens
vibrations created by the lander instruments while gathering data from
the seismometers.

![](media/image30.jpeg){width="2.88125in" height="2.0625in"}

[]{#__RefHeading___Toc145044457 .anchor}Figure 23: Orthogonal view of
the lander configuration.

![](media/image31.jpeg){width="3.0in" height="2.1479166666666667in"}

[]{#__RefHeading___Toc145044458 .anchor}Figure 24: Top view of the
lander configuration.

One of the most important tools on the lander is a drill that supplies
samples to various instruments in the payload. The drill tip must be
placed no more than 0.5 m off the surface and contained in a 0.72 m
tube. To reduce the size of the spacecraft bus, many of the instruments
extend past the defined surfaces, lowering the center of mass. The two
instruments that need to be lowered, the drill and the APXS, are placed
slightly off center so that a small rotation with the retro-rockets will
render a new sample area in the event of hydrazine contamination.

To reduce the contamination due to the hydrazine plume, the downward
thrusters are angled away from the lander's body. The sensitive science
instruments are placed inside the Warm Electronics Box (WEB) to further
minimize contamination. This helps protect the payload from the
radiation environment, extreme cold on Enceladus, and space and geyser
debris. Due to the low gravity on Enceladus, it is not difficult to
maneuver large objects on a gimbaled boom. Therefore, ChemCam and PanCam
are connected on a robotic mast with a large range of motion. This
allows both instruments to measure the same sample and improve the
scientific results of the tools. These two instruments are shielded
during flight by the propulsion stage and, while folded, the lenses of
the instruments are facing towards the lander to protect the CCDs from
sunlight and radiation. The RTG is mounted on the top of the lander,
away from the payload to prevent overheating. However, a portion of the
generator is seeded below the surface of the lander to supply a limited
amount of heat to the inside of the lander.

The retrorockets, designed to slow down the decent of the lander, also
act as attitude control during orbit and a method of hopping during the
extended mission. The thrusters are placed in three separate clusters on
the open faces of the lander. This allows for six degrees of freedom and
reduces the amount of contamination to the surface of Enceladus.

The propulsion stage was designed entirely to support the fuel tanks.
The square structural bus provides the simplest and strongest attachment
for two MMH and two NTO tanks. The interior of the bus contains the
Pressure Control Assembly (PCA) and the helium tank, both of which are
small enough to allow for a thin rectangular support system. The
proposed design for the propulsion stage is not equipped with any
individual control or communication, and it incorporates only one
science instrument. The magnetometer, which will only be used while
orbiting Enceladus, needs to be extended on a long boom to protect it
from the interference of the spacecraft. The deployment mechanism is too
large to fit on the lander and too heavy to carry to the surface, so it
will detach with the rest of the propulsion system. The telescoping
magnetometer boom is controlled and powered through the lander and
deploys shortly after entering the Saturn system. The propulsion system
also houses the two main thrusters, located on the bottom of the lander
along the vertical Z axis of the spacecraft. The thrusters are gimbaled
to minimize the shear forces on the spacecraft.

![](media/image32.jpeg){width="3.2493055555555554in"
height="2.703472222222222in"}

[]{#__RefHeading___Toc145044459 .anchor}Figure 25: Propulsion stage

The spacecraft layout and materials were chosen to minimize the risk of
a system failure resulting from the dangerous space environment. The
primary radiation concerns will be resolved by the multi-layer
insulation covering the bus of the spacecraft. The exposed scientific
instruments will be protected during the cruise phase, either by
incorporated folding methods, or due to the shielding provided by other
parts of the spacecraft.

The Saturn orbital insertion trajectory requires the spacecraft to
travel between Saturn's the F and G Rings. This requires that the
instruments and spacecraft be protected from any dust and debris.
Honeycombed aluminum will be used to construct the bus structure due to
its exceptional ability to protect against micrometeorites. A trade
study lead to the conclusion that using honeycombed aluminum provided
more protection than the alternatives. ^23^

Spacecraft survivability during the 7 year cruise phase was another
consideration in the overall design. Extended exposure to the high
vacuum of space can result in outgassing of materials. This prevents
using materials like zinc and cadmium. Poor choices in materials can
cause short circuits, as well as material deposits on the instruments.
It was found that using aluminum 7075-O for the internal truss structure
and honeycombed aluminum for the skin is the most reliable material to
use for long duration missions.^48^

When the spacecraft is built, it will be vital to test the rate at which
it outgases. Current vacuum technology can only reach 10^-8^ torr,
whereas, the interplanetary environment is 10^-11^ torr. This means that
outgassing data must be extrapolated to obtain accurate values. Though
outgassing will not destroy the satellite, it can disrupt individual
components, leading to a mission failure. A database located on the NASA
Goddard website has a list of materials that have very low outgassing
characteristics. ^48,\ 57^

The spacecraft must also be thermally stable. Thermal conductivity and
thermal expansion coefficients are important considerations in the
design. Three examples of materials that have a net expansion
coefficient of nearly zero are Kevlar 49, UHM carbon, and Invar. Invar
is used on optical components like cameras since it has similar
properties to glass. This material can also be used to help mount
electronics within EAGLE since the coefficient of thermal expansion is
similar to silicon.^27^

The ease of manufacture of the materials must be considered. The
spacecraft will not be using composites because they are difficult to
manufacture, and therefore expensive. Composites are also too brittle to
be primary structure. The strength of the honeycombed aluminum depends
on how the force transfers between the sheets of aluminum. ^23^

A large quantity of titanium, in addition to the aluminum, will be used
for the spacecraft. The titanium will be used for the fuel tanks of the
spacecraft. There will be a future trade study to investigate the
possibility of using aluminum-lithium fuel tanks. This was found to be
effective for the External Tank on the Space Shuttle which has 1/8^th^
inch thick aluminum-lithium alloy.^7,\ 9^

The spacecraft will be attached to the fourth stage by the spring-loaded
Marmon clamp system. V-clamps with an explosive-bolt release mechanism
will connect the lander to the propulsion stage on the top, and the
lander stage on the bottom. This separation system is preferred because
the detachment forces will act strictly in the Z-axis. This will reduce
the shear stresses on the spacecraft that usually result from the use of
pyro-bolts.^8^

![](media/image33.jpeg){width="3.0256944444444445in"
height="2.6284722222222223in"}

[]{#__RefHeading___Toc145044460 .anchor}Figure 26: Illustration of
orbiter separation.

![](media/image34.jpeg){width="2.9458333333333333in"
height="2.6395833333333334in"}

[]{#__RefHeading___Toc145044461 .anchor}Figure 27: V-clamps

The primary focus of structures was designing a robust spacecraft
structure, while minimizing the mass of the vehicles. The hexagon became
the most effective shape to withstand the stresses of launch, provide
enough space inside for the equipment and instruments, and still fulfill
the other mission objectives. The rest of the satellite was designed
around the hexagonal lander. The placement of the dish ensures that it
will manage the launch loads, and the propulsion stage was designed to
be jettisoned in order to reduce the landed mass of the vehicle.

.

After the basic design was generated, the details of the structure had
to be examined. Some considerations included materials, vibrations,
acoustics, the launch faring, and the space environment. Though the
vibration and acoustic calculations could not be accomplished due to
time constraints, the spacecraft does meet the basic requirements and
should withstand all expected stresses during the mission.

The environmental control system (ECS) on EAGLE includes internal and
external temperature control for spacecraft avionics and science payload
instrumentation, protective measures for radiation exposure, and
micrometeorite shielding. The EAGLE ECS is based strongly on the TPS and
exposure data from the Cassini-Huygens mission to Saturn and Titan,
minus EDL system components from the Huygens probe. The EAGLE ECS
concept accounts for extremes experienced during transit and the
expected operational lifetimes for the Saturn orbiter and Enceladus
lander.

Extremes in equilibrium temperature, T~eq~, over the duration of the
mission are found during the transit time spent around Venus and the
operational phase at Saturn, with T~eq~ ranging from 318K at Venus to
90K at Saturn.^18^ Such variances in planetary equilibrium temperatures
and ambient temperature of 4K during other phases of the transit
trajectory require the use of MLI on all exposed surfaces of the EAGLE
spacecraft, for both the Saturn orbiter and the Enceladus lander. The
external temperature control system also accounts for the incident
sunlight on the spacecraft and the amount of time spent at Saturn, where
the incident sunlight is approximately 100 times less intense than that
at Earth.^18^

Based on the required operating temperatures of the scientific payload
instrumentation and spacecraft avionics, the targeted equilibrium
temperature for EAGLE is 220K, achieved using two different lay-ups of
Kapton-based MLI. The MLI lay-ups are based on results from the high
temperature exposure testing done in preparation for Cassini-Huygens TPS
design. Kapton is a synthetic material capable of withstanding
temperatures up to 670K.^31,32^ An all-embossed Kapton ply of 20 layers,
each spaced by Dacron netting, is the designated "high temperature
blanket" and is used on the top of the orbilander near the RTG and on
the underside of the spacecraft near the main engine and ACDS thrusters.
All layers are 0.025 mm thick.^32^ Non-uniformity in thermal emittance
of an entirely Kapton MLI prevents its use on the entire surface area of
the spacecraft.^31,32^

A standard lay-up involving a hybrid of 5 layers of embossed Kapton with
15 layers of Mylar/Dacron can withstand up to 520K, making it suitable
for the majority of the spacecraft surface area at its hottest
environmental ambient temperature point, 0.65 AU.^31,32^ The layer
spacing with the Mylar/Dacron net makes the MLI lay-up appropriate for
micrometeorite shielding as well, as discussed later in this
section.^32^

In the same manner as Cassini-Huygens, EAGLE uses the 4 m diameter high
gain antenna as a heat and sun shield during the transit time spent in
the Venus system. The antenna will be directed toward the sun during the
time EAGLE spends at distances of 0.65 AU to 1 AU. The Hexcel/epoxy and
carbon fiber/epoxy honeycomb structure is coated with high reflectivity
and low absorption paint, as specified for the Cassini-Huygens mission.
The entire external temperature control system, including MLI blanketing
and coatings, totals 11 kg.

Internal temperature control includes both active and passive systems
for maintaining the operating temperatures of the avionics and science
payload instrumentation. The MLI maintains a spacecraft equilibrium
temperature of approximately 220K. For the majority of the science
payload, this thermal environment is well within the operating
temperature range for the instruments. However, several require
operational temperatures below 200K. Passive radiative cooling systems
built into the instruments maintain proper thermal environment for these
instruments, most specifically the IR camera.

During thermal cycling caused by orbit variation and trajectory,
additional passive measures are required, namely partially embedding the
RTG power source on the lander and including Radioisotope Heater Units
(RHUs) with the payload. During the night cycle, RHUs are used to
passively heat the spacecraft interior. Based on the number of RHUs on
Cassini-Huygens, EAGLE will include 3 kg of RHUs (75 individual units)
to provide an additional 75 W (thermal) to the payload.

Active internal temperature control is achieved through the use of
temperature sensors and electric heaters embedded in the payload. Each
instrument has a temperature sensor for each operation critical
component, averaging 4 sensors per instrument. This network of sensors
is linked into the onboard computer, which monitors the temperature of
the components critical for instrument operation. Electric heaters, each
10 W thermal, and averaging one heater for every two instruments,
actively heat avionics and science instrumentation using feedback from
the temperature sensors.

The radiation environment expected during transit and the operational
lifetime of EAGLE is not expected to require a separate protection
system. The absence of Mars and Jupiter passes in the transit trajectory
reduces the possibility of the spacecraft passing through the severe
radiation belts associated with both planets.

All avionics and instrumentation on EAGLE prescribe radiation hardened,
spaceflight qualified hardware. The majority of the science payload
instrumentation has flown on the Mars Exploration Rovers (MER) and
survived radiation environments more severe than those expected at
Saturn.^18^ Should additional data analysis from Cassini's remaining
flyby of Enceladus in 2008 show a more severe radiation environment than
currently known, the current mass budget allows for the inclusion of a
polyethylene radiation shield for the lander payload.

The MLI covering the entire exterior of both the Saturn orbiter and
Enceladus lander also protects the spacecraft from micrometeorite
impacts. The MLI is capable of preventing payload damage from dust sized
particles traveling at speeds ranging from 4-10 km/s.^36^

For the complete spacecraft configuration and the Saturn orbiter
independently, the composite structure and sensitivity range of the high
gain antenna provide a shield against particles larger than dust size
(0.01g) as may be seen with the close flybys of Saturn's E, F, and G
rings.^36^ The greatest particle impact threats to the lander, while in
Enceladus orbit, are those in the E ring, likely being produced by the
geysers on the Enceladus surface. On the surface, the impact threats to
the lander are from the geyser ejecta falling back to the surface as
snow and ice particles. The MLI on the lander and RTG on the top side of
the lander protect the science payload.

The ECS on EAGLE requires the use of both active and passive systems,
including MLI, RHUs, temperature sensors, and electric heaters to
maintain equilibrium temperatures in the operating ranges for the
science payload and avionics. Table 4 summarizes the quantities, mass,
and other design characteristics of the ECS.

![](media/image35.wmf)[]{#__RefHeading___Toc145044473 .anchor}Table 4:
Summary of ECS design components

EAGLE is a deep space mission designed to do many scientific operations
over a long time scale. The power demands will be significant, thus
requiring a robust power system that can operate for longer periods of
time with a high efficiency rate, has low mass, and does not interfere
with the science instruments. Currently, the best option for deep space
power systems are nuclear based with the system relying on Plutonium 238
as fuel.

Since EAGLE will be using various instruments on the Enceladus lander
and the Saturn orbiter, Power has decided to use two separate
Radioisotope Power System (RPS) units. The Multi-Mission Radioisotope
Thermoelectric Generator (MMRTG) will be used exclusively for the
Enceladus lander and two Stirling Radioisotope Generators (SRGs) will be
used during launch, transit and on the Saturn orbiter. All three systems
will provide \>110 W; at beginning of mission (BOM) the MMRTG will
produce 125 W and the SRG will produce 116 W (BOM) each.^71^ After 14
years of operation, the expected power output of the MMRTG is 100 W
while both SRG will output 101 W.

EAGLE will require power to operate three separate stages of the
mission. Once the overall structure reaches Saturn, it will break into
two different sections: the Saturn orbiter and the Enceladus lander. The
latter will eventually land on the surface of Enceladus. Therefore, when
the structure separates, the two SRG units will stay with the Saturn
orbiter to provide power for the Command and Data Handling components as
well as the camera and Propulsion/Attitude Control systems. Overall, the
Saturn orbiter will have an average power allotment of 110W.

The Enceladus lander will have one MMRTG that provides an approximate
average of 110 W by the time the system lands on the surface of
Enceladus, eight years after launch from Earth. Since there are no
moving parts on the MMRTG, the lander will not have any additional
vibrations from the lander when the seismometer is in use.^72^ The fuel
used in both systems will be Plutonium 238, with a decay rate of 0.8%
per year.^72^ The overall power decay rate of the MMRTG is 1.6% per year
and the SRG is 1% per year.^72^ These numbers rely heavily upon the
thermal insulation systems of the units as well as the mechanical moving
parts on the SRG.

During transit, EAGLE will not pass by Mars or Jupiter, which reduces
the amount of exposure to extreme radiation environments. Since the
design of the MMRTG accounts for operations on Mars, it is safe to
assume that this unit will be able to operate under the radiation
environment of Saturn.^72^ As for the SRG, while under development, it
can be assumed that the design will allow for operations under extreme
radiation environments as well.

Below is a chart listing the power demands of each subsystem with
respect to the Saturn orbiter and the Enceladus lander stages of EAGLE:

![](media/image36.wmf)[]{#__RefHeading___Toc145044474 .anchor}**Table 5:
Saturn orbiter: 2 SRG exclusive (BOM max = 116 Watts each)**

\*Constant power demand

\*\*80 Watts initially for 100 milliseconds during burn then 50 W demand
for the duration of the burn.

\*\*\*When there is further development of the power system, a more
concrete power cushion can be suggested. In regards to the current power
SRG allocation, it would be desired to have 5 W for the student
instrument but right now it is not feasible (thus the reason for the
zero).

\*\*\*\*The Diplexer requires a continuous power of 5 W but at times,
the power need could reach 100W.

![](media/image37.wmf)

[]{#__RefHeading___Toc145044475 .anchor}Table 6: Enceladus lander: 1
MMRTG exclusive (BOM = 125 W)

At the beginning of the entire mission, Power will have a total output
of 357 Watts with a power demand of 429.6 Watts with all components
active at full power. Therefore, the final power system should include
the use of power modes, and thereby not require the addition of another
RPS.

[]{#__RefHeading___Toc145044462 .anchor}Figure 28: MMRTG & SRG power
distribution in transit (BOM).

Each subsystem will benefit from the respective power system, depending
on its location and demands during the various stages in the mission.
Having two separate power systems would benefit the mission greatly
because it reduces the fuel demand, and amount of mass. Furthermore, the
MMRTG and SRG have complimentary advantages for deep space missions. The
MMRTG generates power from thermocouples connected to eight General
Purpose Heating Sources (GPHS) that are stacked on top of each other.
What sets the MMRTG apart from the RTG units used on Cassini is the
total mass of 44.2 kg and a fuel mass of 4 kg.^71^ The Cassini RTG had a
mass of 55.5 kg and a fuel mass of 7.8 kg.^72^ Therefore, for future
missions it is imperative to reduce the amount of fuel needed and ensure
that the RPS is more efficient.

The two SRG units will go onto the over structure and will provide power
to the Propulsion/Attitude Control and the Command & Data Handling
Subsystems. The SRG has a higher efficiency than the MMRTG and requires
a smaller amount of fuel. With two 55 W Stirling engines on each SRG,
the system will receive the thermal component of the system from two
GPHS.^73^ With a reduction in the amount of fuel used (1 kg), the
overall thermal power that requires dissipation is only 500 Watts for
each SRG.^74^ For surrounding instruments that require cooler operating
temperatures, this is a friendlier neighbor. To reduce thermal and
vibration interference, the two SRG systems will be on the back of the
orbiter.

[]{#__RefHeading___Toc145044463 .anchor}Figure 29: Power distribution
for the Saturn orbiter 8 years after launch.

In order to conserve power demands from all of the instruments and
systems involved, the power system will implement the use of standby or
power modes. Only vital systems will draw power from their respective
RPS on a constant basis. The MMRTG on the Enceladus lander will only
keep the instruments on safe power mode during transit. Before EAGLE
separates, there will be a burn time where the MMRTG will act as an
additional power resource for the Propulsion/Attitude Control (P/AC)
system.

[]{#__RefHeading___Toc145044464 .anchor}Figure 30: Power distribution
for the Enceladus lander during decent to surface.

Command power needs were not available by the time of this study.
However, due to the initial high power requests from C&DH, one of the
SRG units will be used exclusively for the C&DH components, while the
P/AC will use most of the power output from the second SRG. As mentioned
before, if P/AC requires more power, the MMRTG may provide additional
support.

During the research process, Power considered many different types of
deep space powering systems. The total amount of solar energy is
drastically reduced when EAGLE reaches Saturn, eliminating the potential
use of solar panels. The Department of Energy (DOE) has poured millions
of dollars into RPS research and development programs across the
country. NASA Glenn Research Center has devoted a large amount of time
and money, in conjunction with Stirling Technology Company, to produce
deep space power systems using Stirling engine technologies.

The SRG 110, which is the primary unit suggested for use in this
mission, is in the prototype phase of development at the Glenn Research
Center. According to researchers, the finalized design will lead to a
working engineering model in the near future. Due to the low power
output of the SRG (116 Watts) compared to the RTG used on Cassini (300
Watts each), there is a brand new focus on developing the Advanced SRG.
The expected power output of this system is 160 Watts at BOM and 140
Watts after 14 years.^71^

Below are charts of comparisons between different RPS considered for
EAGLE:

![](media/image38.wmf)[]{#__RefHeading___Toc145044476 .anchor}**Table 7:
Trade study for the RPS**

\*Prototype phase

\*\*Currently in initial development

\*\*\*One Cassini RTG

Power Outage During Mission: DC Electric (Watts)

![](media/image39.wmf)

[]{#__RefHeading___Toc145044477 .anchor}Table 8: Trade study considering
BOM power

The goal is for EAGLE be the most efficient spacecraft for a deep space
mission by the time it launches in the next 20 years. Accounting for the
current shortage of Plutonium 238, it is exceedingly important to
develop missions that require a smaller amount of fuel without
sacrificing performance. RPS are the best options for future missions
and SRG development is the key to producing a highly efficient power
source with the smallest amount of total and fuel mass.

The responsibility of the Communications subsystem is to receive and
transmit commands from NASA's Deep Space Network and send some of those
commands to the EAGLE lander. The EAGLE orbiter must also receive data
from the lander and send that data telemetry to the Deep Space Network.
The design of the communications system requires an intricate balance of
power, mass, and cost while achieving a data rate and link margin
sufficient to achieve the science goals of the mission.

The EAGLE communications system will need to receive and demodulate
command waveforms transmitted from the Deep Space Network. The goal is
to do this with a bit error rate (BER) of 10^-5^; there will only be one
erroneous bit for every 100,000 transmitted bits.

The spacecraft must also be able to receive commands while it undergoes
Earth and Venus gravity assists. During this phase of the mission,
however, the thermal protection system will use the high gain antenna
(HGA) as a solar radiation shield. The spacecraft will therefore receive
commands using one of its two low gain antennas (LGAs). In the later
phase of the mission, it will receive commands predominantly through the
HGA.

The EAGLE orbiter must generate and modulate the downlink carrier
containing data from the onboard instruments and data acquired from the
Enceladus lander. During the interplanetary cruise phase of the mission,
one of the LGAs can provide the telemetry downlink until the spacecraft
is 1.5 AU from the Earth. The HGA will not be used until several months
after the Earth gravity assist.

The orbiter-DSN downlink will meet the requirements indicated below:

34m DSN antenna, 12 hour downlink per day, 300 MBits/day

70m DSN antenna, 12 hour downlink per day, 1.3 GBits/day

The EAGLE mission consists of a spacecraft which will break into an
orbiter and a lander when it reaches the Saturnian system. The lander
will orbit Enceladus, choose a landing site, and eventually land on
Enceladus. The lander contains most of the science instruments and the
Saturn orbiter is essentially an orbital relay satellite to provide
communications support for the EAGLE lander.

It is standard to put one HGA and two LGAs on a spacecraft traveling to
the outer planets and the Communications team has continued this
practice. The HGA is necessary for long range/high data rate
communications and the LGAs are necessary for communications while the
spacecraft is near the Earth and for potential emergencies later in the
mission.

There are only three frequency bands allotted to deep space
communications: S-band (2.29 -- 2.30 GHz downlink, 2.11 -- 2.12 GHz
uplink), X-band (8.40 -- 8.45 GHz downlink, 7.15 -- 7.19 GHz uplink),
and Ka-band (31.8 -- 32.3 GHz downlink, 34.2 -- 34.7 GHz uplink)^37^.
Although a higher frequency corresponds to a higher data rate, Ka-band
was not used for orbiter-DSN communications because the high frequency
produced higher levels of atmospheric attenuation than lower frequencies
(see Figure 31). S-band causes the lowest amount of atmospheric
attenuation, but it does not correspond to a data rate high enough to
meet the science goals. X-band, therefore, is the optimum frequency; it
provides a high enough data rate and low enough signal loss, so the
Communications subsystem implemented it as the uplink and downlink
frequency for the orbiter-DSN communications link.

![](media/image40.png){width="2.998611111111111in"
height="1.5409722222222222in"}

[]{#__RefHeading___Toc145044465 .anchor}Figure 31: Plot of atmospheric
attenuation v. frequency. Once can see that, in certain cases, a signal
can be almost entirely lost when transmitted at a high enough frequency.

To choose antenna size and transmitter power for the orbiter, the
Communications team established a link budget to confirm that the
communications link would include a sufficient link margin (3 dB). A
careful analysis of previous deep space missions and the current link
budget caused the team to choose 20 W as the power for all antennas
(LGAs and HGAs). To achieve a data rate of 7 kilobits/second, which
should be high enough to achieve the science goals, the diameter of the
HGA on the orbiter is 4 m and the diameter of the LGAs on the orbiter is
0.20 m. A conservative estimate of 0.60 has been used for the antenna
efficiency and a total of -2 dB has been used for the line, propagation,
and polarization loss^28^. The appropriate link budgets are given in the
following tables.

![](media/image41.wmf)

[]{#__RefHeading___Toc145044478 .anchor}Table 9: Link budget for the
orbiter-DSN downlink using HGA.

![](media/image42.wmf)

[]{#__RefHeading___Toc145044479 .anchor}Table 10: Link budget for
orbiter-DSN downlink using one LGA.

![](media/image43.wmf)

[]{#__RefHeading___Toc145044480 .anchor}Table 11: Link budget for
downlink from lander to orbiter.

All current space missions use a digital encoding scheme called phase
shift keying. This technique conveys data by modulating the phase of a
carrier wave and lowers the required digital signal to noise ratio
(E~b~/N~0~) relative to an uncoded signal. The Communications team chose
binary phase shift keying (BPSK) and Plus Reed-Solomon Viterbi Decoding
because it gives the best bit error rate (BER) performance for a given
E~b~/N~o~^55^. In this case, the system achieved a BER of 10^-5^ using a
required E~b~/N~o~ of 2.7 dB.

The EAGLE lander contains almost all of the science instruments and the
lander communications system is vital to the success of the mission. To
that end, there are two LGAs on the lander which will transmit data to
the Saturn orbiter over an S-band frequency. An HGA would have provided
a much higher data rate, but would have also required a much higher
mass, so it was not included on the lander.

The fact that an electromagnetic signal spreads over larger areas at
increased distances from a source causes a significant signal loss for
deep space communication. This free space loss is greatly reduced as the
orbiter approaches Enceladus. The bit rate can therefore increase by
very large amounts near orbital periapsis. The period of the orbit is
approximately one month, so each month will have characteristic times
when the lander can achieve high data rates. During other times, it will
only be possible to transmit at much lower data rates. Evidently the
Science team should attempt to use the high data rate instruments such
as the Alpha Particle X-ray Spectrometer and the Wide Field Cameras
(which have a variable data volume depending on how many images one is
acquiring) during the appropriate time in the orbit each month.

The downlink data rate from the lander to the orbiter is a minimum of
1.5 kilobits/second, but that will increase to megabits/second or even
gigabits/second near orbital periapsis. Downlink times between the
lander-orbiter and the orbiter-DSN will therefore vary depending on the
orbiter's position. Table 12 shows the data rates for the science
instruments on the lander.

![](media/image44.wmf)

[]{#__RefHeading___Toc145044481 .anchor}Table 12: The data rates for all
of the science instruments on the lander. Note that several of the
instruments will not be sending data back to the Earth.

An alternate design configuration would be to place a high gain antenna
on the lander. This was done on the Mars Exploration Rovers and would
have allowed, in this case, for much higher data rates with the Saturn
orbiter and the possibility of low data rate communication between the
lander and the Earth. However, HGAs are very expensive and heavy, so the
system uses two LGAs instead.

During the initial gravity assist phase of the mission, the mission
design uses the HGA as a solar radiation shield, while two LGAs provide
X-band communications with the Earth at a data rate of 1 kilobit/second.
Several months after the Earth gravity assist, the HGA will be deployed
and will commence X-band communications with the Earth at a minimum of 7
kilobits/second. Once the lander has reached Enceladus, it will
communicate with the Saturn orbiter at an S-band frequency with a
minimum data rate of 1.5 kiloibts/second. Significantly higher data
rates will be possible when the orbiter is closer to Enceladus. All
antennas operate with a power of 20 W and the entire communications
subsystem has a mass of 109.5 kg (orbiter) and 4 kg (lander).

The responsibility of the Command & Data Handling (C&DH) subsystem is to
design a system that will meet all of the C&DH requirements designated
for this mission. These requirements are to record all scientific data
on the surface of Enceladus and process them for relay to earth via the
Saturn orbiter. The primary focus is double redundancy, mainly in the
transponder, antennas and recording devices. Figure 32 illustrates the
flow of information for the C&DH subsystem to allow a clear
understanding of the transmission and how it relates to the
telecommunication subsystem.

![](media/image45.wmf)

[]{#__RefHeading___Toc145044466 .anchor}Figure 32: Flow chart for the
Command & Data Handling subsystem.

The Command & Data Handling subsystem is critical to the operation of
the spacecraft and has several key components that allow for
communication between each system on the vehicle. The current design is
designed around the science objectives, as well as the transfer of data
between other systems. The finalized scientific objectives and payload
specifications will help determine the overall design of the C&DH system
as well as the specifications of each component. shows the components of
the C&DH system.

![](media/image46.wmf)

[]{#__RefHeading___Toc145044482 .anchor}

Table 13: Applications of Command & Data Handling
components.^50,\ 51,\ 52^

The Small Deep Space Transponder (SDST) will send and receive
transmissions via the High Gain Antenna (HGA) and Low Gain Antennas
(LGAs) when communicating between the Saturn orbiter and Earth. It is
the predecessor of the Deep Space Transponder (DST) used on the
Cassini-Huygens mission, and can operate at a wide range of frequencies
to allow for redundancy in the communication subsystem. This design has
less mass and smaller dimensions, and it uses less power for
communications. Furthermore it combines the Command Detection Unit (CDU)
and Telemetry Modulation Unit (TMU). The Jet Propulsion Laboratory (JPL)
in Pasadena, California funded the development of this transponder by
Motorola Inc. The C&DH team chose the SDST for this mission in part
because it is efficient in X-band communications for deep space.
Additionally, it is radiation resistant, light in mass, and uses less
power than the DST. The testing and verification phase of development
will be completed for this deep space orbiter vehicle mission before the
preferred launch date.^51,\ 52^

![](media/image47.png){width="2.5458333333333334in"
height="2.115972222222222in"}

[]{#__RefHeading___Toc145044467 .anchor}Figure 33: Small deep space
transponder.

The S-band transponder will be used when transmitting and receiving data
between the orbiter and Enceladus lander. It will be efficient in the
frequency range of 2.2 GHz to 2.3 GHz and communicate between the
Enceladus lander and the Saturn orbiter. The basic requirements for the
S-band transponder will be to transmit at least 256 kbps at a frequency
range 2200-2300 MHz and receive at a frequency range of 2000-2120MHz
with 2dB or less noise. Its mass will be approximately 4kg or less and
have an output power of 5W. Though these specifications are preliminary,
it will use the least amount of power for the maximum data rate using
the most up-to-date technology available at the date of launch.

For a deep space mission to the Saturn system, the telecommunication
system requires amplification to transmit data over large distances. The
amplifier will have small mass and size, and will be completely space
qualified to have a high radiation tolerance. It needs to operate in the
X-Band frequencies in order to enhance the transmission power of the
High Gain Antenna (HGA) for communications with Earth. This amplifier
will be complementary to the SDST, which will connect to it using a
standard 1553 data interface. The power amplifier will be used only when
during downlink from the Saturn orbiter to Earth.

This processor will be capable of performing all the necessary actions
to control the C&DH subsystem of EAGLE. It will build on the heritage of
previous processors used on spacecraft, and it will implement VSWorks
software. For more information, see Section 7.5.3, Trade Studies.^50^

![](media/image48.png){width="2.433333333333333in" height="1.6in"}

[]{#__RefHeading___Toc145044468 .anchor}

Figure 34: Processor design roadmap^50^

The Solid State Recorder (SSR) on the lander will save data collected
from science instruments until the lander can transmit this information
to the orbiter. Two SSRs are located on the Saturn orbiter: one will
record data while the other will be used in case of contingency. The
limiting factors for data transmission depend on the location of
Enceladus with respect to the orbiter and Saturn. The orbiter will have
an additional two SSRs to save all of the data received from the lander
until the orbiter can it relay it to the Deep Space Network (DSN) on
Earth. The transmission time from the orbiter to Earth will be
determined mostly by costs of using the DSN and in part by the allowed
communications windows for transmission between the two points.

These components will allow for the high-speed transmission of data
within the Enceladus Lander, as well as within the Saturn orbiter.

The specifications of each component will vary depending on finalized
data rates, cost, size, and power. shows approximated values using
specifications given on commercial web sites.

![](media/image49.wmf)[]{#__RefHeading___Toc145044483 .anchor}

**Table 14: Specifications of C&DH\
components ^50,\ 51,\ 52^**

The Command & Data Handling team conducted a trade study to determine
the most reliable processor. Compared to the RAD6000, the RAD750 can
operate at a higher frequency, operates at a higher rate with more
Read-Only Memory (ROM), and can withstand a wider range of temperatures
and radiation exposure. Due to these advantages, the RAD750 is the best
component.^50^

The C&DH subsystem has many commonalities with previous deep space
missions, so only the level of technology will be unique. The components
will need to be efficient and reliable, but at the same time, they
should be only as advanced as required. The computing capabilities of
this subsystem should not exceed that of the amount of expected science
data recorded and analyzed. The C&DH subsystem of EAGLE will be
doubly-redundant through multiple communication antennas, an extra SSR,
and a multi-frequency transponder. A non-redundant system could be
implemented with the calculated risk of failure to any of the major C&DH
components. This design would be limited to one frequency of
communication and less data storage.^12^

The Command & Data Handling subsystem of EAGLE shall include all
components necessary to collect and effectively transmit data between
earth, orbiter, and lander, based on the requirements of the science
data and other subsystems. It will be composed of the most up-to-date
software and hardware at the date of launch, as well as parts that have
been space/radiation tested and qualified.

The Guidance, Navigation & Control (GNC) subsystem is responsible for
monitoring the spacecraft's position and attitude through its
interplanetary trajectory, and making attitude changes as necessary to
comply with both pre-planned and corrective burn sequences. For this
mission, the GNC subsystem must provide the following capabilities: the
system must determine, in real-time, the exact spacecraft position and
attitude during its interplanetary trajectory and related gravity assist
maneuvers. The system must be able to orient the spacecraft in a way
that the communications dish is always facing the sun in order to
provide radiation protection during the portion of the trajectory where
solar radiation is above acceptable levels. The system must process this
information on board to determine what corrections, if any, are
necessary to ensure successful Saturn orbit insertion; the system must
have the means to re-orient the spacecraft to allow both pre-planned and
corrective burn sequences, and to enable potential interplanetary
science-related operations. Finally, the system must provide both
station-keeping capability for the orbiter spacecraft to maintain and
correct its steady-state orbit, and surface mobility capability to the
lander spacecraft to allow for potential surface traversal activities.
The pointing accuracy is required to be on the order of degrees in order
to attain these objectives. Additional requirements include minimizing
the power required and minimizing the necessary mass.

The GNC design team studied a variety of previous NASA interplanetary
missions for examples of GNC subsystems. The most analogous missions
were Voyager and Cassini-Huygens; therefore, the GNC subsystem for EAGLE
has strong commonalities with both of these spacecraft. Guidance and
navigation of the spacecraft will be accomplished using a series of two
star trackers, each of which will take measurements of the star
background and transmit this information to a central processor. This
processor, equipped with an exhaustive database of known star positions,
will triangulate the exact position of the spacecraft based on a
comparison between the star database and primary measurements. Three
inertial reference units included in the spacecraft will relay the
orientation of the spacecraft to the central processor. Once the exact
position of the spacecraft is known, the GNC system will make
corrections to the flight path based on comparisons between the
calculated interstellar trajectory and the actual path of the
spacecraft.

In order to make such corrections, the GNC subsystem implements a
combination control system consisting of monopropellant hydrazine
thrusters and reaction wheel mechanisms. Reaction wheel systems provide
large-scale attitude corrections, and the monopropellant thruster system
accounts for precision pointing.^28^ The GNC team selected to include
both types of attitude control for a variety of reasons. First, the
reaction jets were necessary to attain the precision required for
communications between ground control and the lander and for
communication between the lander and Earth as well. Due to the length of
the mission, the reaction wheels are a secondary system that will allow
the spacecraft to conserve fuel. By using reaction wheels, the
three-axis stabilized spacecraft can be augmented with spin
stabilization when precision pointing is not required. The attitude
control system uses a total of three 10 kg reaction wheels, in a
mutually orthogonal configuration on the Saturn orbiter portion of the
spacecraft. Such a configuration provides three-axis spacecraft control
through precision angular momentum changes.

The spacecraft utilizes twenty-eight 1-N hydrazine thrusters. Sixteen of
these thrusters, organized into clusters of four, are attached to four
one-meter booms on the orbiter portion of the spacecraft, enabling full
3-axis spacecraft control. These thrusters comprise the primary control
system during the interplanetary voyage when both lander and orbiter are
connected and when the orbiter is in orbit around Saturn. This system is
partially redundant, as full attitude control can be achieved using only
8 thrusters. The remaining three clusters of four thrusters are attached
to alternating vertices of the underside of the hexagonal lander
spacecraft platform. This system provides attitude control for the
lander as it orbits Enceladus and during entry, descent, and landing.
These thrusters

will serve a primary function of providing "hopping" capability to the
structure once it reaches the surface of Enceladus. These thrusters will
also serve as backup interplanetary attitude control thrusters,
providing an additional level of redundancy prior to lander-orbiter
separation, should the orbiter thruster system fail. Both separate
attitude control systems will contain respective monopropellant tank
systems, with the orbiter system having twice the hydrazine mass, since
it will provide the main attitude control for the interplanetary voyage
and will require precision pointing while the lander relays science
data. The conservative fuel mass estimate is 200 kg of hydrazine. The
monopropellant tanks are made of titanium and have pressure fed systems
for the thrusters. Table 15 summarizes the GNC components:

![](media/image50.wmf)[]{#__RefHeading___Toc145044484 .anchor}Table 15:
Summary of GNC components

The Guidance, Navigation & Control subsystem weighed several different
qualitative options prior to choosing the current system design.
Star-tracker based position determination has become an industry
standard practice, and the process of active position determination
based on star tracker measurements has been refined and proven on many
successful interplanetary missions, with the most recent mission being
the Cassini mission to Saturn. This proven flight capability and
traceable heritage, coupled with low instrumentation mass requirements
and high levels of precision, makes star-tracker determination the best
method of position determination for this mission.

The use of star-tracking based position determination weighed heavily in
the decision to implement 3-axis thruster/reaction wheel control system
as opposed to other attitude control systems. Spin stabilization, a
commonly used method that places the craft into a controlled single- or
multi-axis spin and uses angular momentum to maintain axial
directionality, poses problems for star-tracking navigation. If the
spacecraft is constantly rotating, taking measurements of distant stars
for navigation becomes significantly more challenging. Spin
stabilization also eliminates any possibility of interplanetary
camera-based observation, and complicates any potential in-flight
scientific experimentation. For these reasons, spin-stabilized attitude
control is not the primary mode of attitude control for this mission.

Remaining viable alternatives for attitude control include reaction
jets, reaction wheels and momentum wheels. There are a few options in
terms of reaction jets, but hydrazine thrusters have an extensive
heritage, including Voyager, Viking, and Cassini, and provide the most
thrust and highest specific impulse for the lowest mass. Hydrazine
thrusters, while very precise, require consumable fuel (and therefore
fuel tanks) to operate, thus increasing overall system mass and
complicating structural layout and thermal requirements as compared to
reaction wheels. However, attitude control thrusters have the added
benefit of potentially doubling as a limited propulsion system, enabling
surface or interplanetary thrusting capability. Momentum and reaction
wheels, while less precise than thruster systems, are less massive than
thruster/fuel systems and require no consumable fuel. For these reasons,
along with recent proven performance on the Cassini mission, a
combination system using both reaction wheels and hydrazine thrusters
was selected for this mission, enabling both large scale and precise
attitude control. Therefore, the attitude control system meets all
design requirements, with little sacrifice in terms of mass.

Manipulating the number and configuration of monopropellant thrusters
could yield alternative spacecraft geometries. A system of 16 thrusters
in four clusters on the orbiter could provide greater control for the
spacecraft. This will result in increased mass but also yield greater
control authority. For the lander, 2-N thrusters and a larger fuel tank
would make the attitude control system more powerful, which would give
the lander a greater range of mobility for exploring the surface.
However, this would add more mass and would require high levels of
precision due to the low gravitational force of Enceladus. Too much
thrust may cause the lander to overshoot the target, possibly achieve
exit velocity, and leave the surface of the moon.

Both star tracker-based position determination, and monopropellant
hydrazine thruster/reaction wheel attitude control have proven flight
capability and traceable mission heritage. The combination system
enables both redundant interplanetary GNC capability prior to vehicle
separation and station keeping/surface mobility capabilities after
vehicle separation. This design successfully meets the mission
requirements while providing a desired amount of efficiency.

The propulsion subsystem is responsible for all spacecraft propulsive
needs beyond interplanetary trajectory insertion. For this mission, the
propulsion subsystem must provide the following capabilities: the
subsystem must be capable of providing time-lapse burns during the
spacecraft's interplanetary voyage in accordance with both planned and
corrective trajectory burns; the subsystem must provide the thruster
capability necessary to ensure spacecraft Saturnian orbit capture; the
subsystem must provide thruster capability to the lander spacecraft such
that the vehicle can enter temporary orbit around Enceladus. The
requirements for these burns are included in the Trajectory subsystem
section, with the largest propulsive requirement being the 4.5 km/s burn
required to enter into Enceladus orbit.

In the design of the propulsion subsystem, the Cassini spacecraft
contributed the most significant flight heritage to this mission.
However, the Cassini mission had a much lower propulsive constraint, so
EAGLE will need a unique propulsion system to meet the requirements.
This system, which exploits a chemical reaction between
Monomethylhydrazine (MMH) and Nitrogen Tetroxide (NTO) to create thrust,
has been used extensively for deep-space exploratory vehicles, most
recently on the Cassini mission to Saturn.^6^ The MMH/NTO bipropellant
system provides a specific impulse of 340 s,^4^ which is the highest
attainable by chemical propellant systems for interplanetary travel.

Two separate, pressurized fuel tanks house 1220.4 kg of fuel (MMH) and
1952.7 kg of oxidizer (NTO) components. Using the propulsive
requirements provided by the Trajectory team, along with the specific
impulse of the fuel, the Propulsion team used the rocket equation to
determine the masses of each propellant. A pressure system containing 16
kg of highly pressurized Helium gas will maintain the pressure within
these tanks. Two Pressurant Control Assemblies (PCAs) pump the Helium
out of its highly pressurized composite tank housing to provide
pressurization during regulated maneuvers.^6^ The fuel and the oxidizer
flow through the Propellant Isolation Assembly (PIA) to be combined in
the Main Engine Assembly (MEA) to provide thrust for the spacecraft.^6^
The attitude control system described in the previous section will align
the spacecraft in the correct orientation to enable successful
completion of both planned and corrective burn sequences.

The main bipropellant tanks are located in the center of the spacecraft
with the main rocket engine located on the opposite end of the
spacecraft from the communications dish. The entire propulsion system is
to be contained in this area because at Enceladus orbit insertion, the
Saturn orbiter will disconnect from the lander and travel on another
trajectory. Then the lander spacecraft will re-orient so that the
propulsion system can insert the lander into Enceladus orbit. After this
main burn, the lander will jettison the entire propulsion system,
including fuel tanks and the main rocket engine assembly; these
components will then enter into Enceladus orbit or crash into the
surface of Enceladus. The lander ejects this stage because it is a large
amount of mass and volume that serves no further purpose to the lander.
The fuel tanks will be empty and the spacecraft requires no more large
burns; the attitude control system will provide the small propulsive
requirements to land the spacecraft. The deployment of the main
propulsion system makes the wiring and plumbing a crucial component that
must be completely self-contained.

The propulsion system has a total dry mass of 360 kg, with an additional
5052 kg necessary for fuel. This system will provide the spacecraft an
*I~SP~* of 340 sec, a thrust of 445 N, and will provide all the
propulsion for the interplanetary portion of the mission as well as
Enceladus orbit insertion. This system has an average mass power load of
only 50 W, except for the main burn when there is an initial spike of
about 80 W, lasting only 100 ms, which is necessary to open the main
valve.^6^ Enceladus entry and descent will be controlled using the
hydrazine monopropellant GNC system discussed in the Section
[7.6.2]{.mark}. Table 16 provides a summary of the Propulsion system:

![](media/image51.wmf)[]{#__RefHeading___Toc145044485 .anchor}**Table
16: Summary of the Propulsion system**

The propulsion team conducted basic trade studies to determine which
propulsion system most adequately satisfied the requirements of this
mission. At the time of this study, two potential options existed that
could realistically bring a spacecraft to the Saturn-Enceladus system:
chemical propulsion and electrical propulsion.

Chemical propulsion has been flight verified on several interstellar
missions, and its successful performance has been refined and improved
several times. In the case of MMH/NTO, the propellant system provides a
significant level of thrust (445 N) at its given I~SP~ of 340 s.
However, chemical propulsion requires extremely high amounts of fuel (in
this situation 5052 additional kg of mass), consuming a considerable
proportion of the overall spacecraft mass budget. Electrical propulsion
does not require nearly as much mass and can theoretically achieve much
higher specific impulses, on the order of 1,000 s.^28^ However, electric
propulsion systems require much higher electrical budgets and, most
importantly, impart orders-of-magnitude less thrust than chemical
propulsion. Such low levels of thrust would significantly increase the
length of interstellar travel time, which in turn would significantly
complicate the design of other subsystems. The EAGLE spacecraft is
simply so massive that the most sophisticated electrical propulsion
still would not provide enough thrust to reach Saturn. Progress is being
made in electrical propulsion, but currently, there is no system
powerful enough to meet the mission requirements.

For these reasons, combined with Cassini's recent success in reaching
the Saturn system using a similar propulsion system, chemical propulsion
was selected for this mission. This design satisfies every
propulsion-related mission requirement, and has sufficient flight
heritage to enable successful system implementation.

The propulsion team considered a range of configurations for the
subsystem, many of which were determined to be inefficient and/or too
complicated. However, one significant design challenge in the current
configuration is the fact that the spacecraft jettisons the rocket
assembly and the fuel tank after depletion as if they are a separate
stage. This leads to complexity in wiring and piping and creates space
debris. An alternative would be to have the fuel tanks and main rocket
engine remain attached to the lander and utilize the main propulsion
system for Enceladus orbit insertion and the entry, descent, and landing
phase. This provides more power than the hydrazine thrusters in the
current configuration, and would not have a large fuel requirement.
After landing, it is possible that the lander could disconnect from the
main rocket assembly and then "hop" to another location to perform
scientific readings. This alternative configuration results in landing a
much larger amount of empty mass on Enceladus, which is why it was not
selected as the final design.

The propulsion system is by far the most massive of all the subsystems
on the spacecraft. For this reason, any added efficiency will have a
large effect on the available mass for other subsystems. Because of
this, the Propulsion team paid special attention to reducing mass, in
terms of both fuel and hardware. The MMH/NTO bipropellant system is
currently the most efficient propulsion system available for
interplanetary spaceflight. A more efficient and more powerful
propulsion system may be developed in the 15 years before this mission
is launched, but this cannot be planned into the design. With existing
technologies, the Propulsion team has selected what it believes to be
the best propulsion system to accomplish the requirements with the
lowest use of mass.

Several methods were employed in order to determine the cost of this
mission. The primary method was parametric cost estimation using the
Unmanned Space Vehicle Cost Model, 8^th^ edition (USCM8). This model
uses Cost Estimating Relationships (CERs) that were derived from a
historical database of actual unmanned space missions. All costs were
assumed to be nonrecurring and were calculated based on a theoretical
first unit (TFU). In addition, a nonstandard bus was assumed. With all
this considered, the USCM8 is used to find costs in FY 2000 dollars,
which are then inflated to reflect current FY 2006 dollars. Thus, the
total cost of the EAGLE mission in 2006 dollars is expected to be \$838
million, which falls under NASA's classification of a flagship mission.

![](media/image52.wmf)[]{#__RefHeading___Toc145044486 .anchor}**Table
17: Costs by Work Breakdown Structure** Note: all costs are given in
thousands of dollars.

Due to the error inherent in purely parametric estimating methods,
engineering build-up methods were supplemented where appropriate.
Specifically, the power and propulsion subsystems were estimated using a
combination of the CERs and actual costs found in the literature. In
addition, more accurate costs were determined for other subsystems by
using the component level CERs rather than subsystem level CERs.

The Program Management and System Engineering cost allocations were
based on the Program Level CER specifically for science missions such as
EAGLE. This particular CER was based on the total cost of the spacecraft
plus the cost of integration, assembly, and testing. The product
actually combines the two disciplines, but they are represented
separately based on the relative percentages of the budget they
typically occupy.

For the purposes of this project, mission design is defined as all
calculations that take place to create a flight path for the spacecraft.
Thus, the cost is predominantly representative of a handful of employees
and their respective labor costs. Some software licensing was added for
such programs as STK and Matlab.

The thermal control estimate was derived using both a mass based CER and
values found in the literature since the use of RHUs is not standard.
The cost of the RHUs was found directly by multiplying the quantity by
\$35K per RHU. It should also be noted that the Environmental Impact
Statement (EIS) required for the use of RTGs was assumed to include the
investigation of the safety of the RHU thermal system. Therefore, some
of the costs incurred by thermal were actually included in the power
estimate.

The power subsystem's cost was also estimated by using a combination of
parametric and direct methods. Due to a lack of MMRTG or SRG
representation in the heritage database, specific costs based on RTG
heritage were acquired from Mike McGrath, Director of Engineering at the
Laboratory for Atmospheric and Space Physics (LASP). Also included was
\$15 million for an EIS as required when any kind of nuclear device is
flown in space. The CER for power conditioning and distribution was used
in order to provide a more complete cost of the power subsystem.

Performance Assurance/Safety, Assembly Test & Launch Operations (ATLO),
and Calibration CERs were all used to come up with the values listed.
All of these CERs were based on the total cost of the spacecraft. As was
done for Program Management and Systems Engineering cost estimations,
ATLO and Calibration were separated out of the same CER by the relative
percentages of the budget they typically occupy.

The cost of Mission Operations was calculated by using NASA's own
Mission Operations Cost Model (MOCM). The primary driver of this CER was
the total mission cost (less mission operations) and can be adjusted to
fit communications, planetary, physics/astronomy, manned, and Earth
observation spacecraft. The planetary mission model was used for the
values given in tables 8-1 and 8-2.

The cost of the launch vehicle was found separately from any parametric
cost estimation methods. According to astronautix.com, the cost to
launch a Delta IV in the 4050H-19 configuration is \$254 million in FY04
dollars. This number was inflated to current FY06 dollars using actual
inflation values obtained from inflationdata.com. Thus, the cost for
launching a Delta IV Heavy in FY06 dollars is approximately \$270
million.

Educational outreach was assumed to be 1-2% of the total mission cost
minus launch vehicle expenses. All of these costs are distributed over
the lifetime of the mission and adjusted for inflation. Table 8-2
represents the total mission cost (TMC) in real year dollars according
to the inflation values taken from SMAD and provided by the Office of
the Secretary of Defense.

![](media/image53.wmf)

![](media/image54.wmf)

[]{#__RefHeading___Toc145044487 .anchor}Table 18: Total mission cost
funding profile.

Mission activities will be broken up by the NASA standard phases.

1\. Advanced Studies-Pre-Phase A (Currently)

> Conceptual design phase
>
> Development of performance goals and requirements
>
> Establishment of science working group
>
> Trade studies of mission concepts
>
> Feasibility and preliminary cost analysis

2\. Preliminary Analysis-Phase A (Currently)

> Proof of concept analysis

Mission operations concepts

Make vs. Buy trade studies

Payload definition

Project Schedules

3\. Definition-Phase B (2013-14)

Define baseline technical solution

Define Science Schedule

Create requirements documents

System requirement and system design reviews

4\. Design-Phase C/D (2015-23)

Detailed Design Process

Cutting Metal

Test and Analysis

PDR, CDR, Test Readiness, Launch Readiness reviews.

5\. Operations-Phase E (2023-2030+)

Launch

Mission operations

Maintenance and Troubleshooting

Failure Monitoring

End of Life Disposal

![](media/image55.png){width="3.5194444444444444in"
height="1.2458333333333333in"}

[]{#__RefHeading___Toc145044469 .anchor}Figure 35: Mission Timeline with
Reviews and Milestones. Mission activities will be broken up by the NASA
standard phases.

An occasionally overlooked but always essential part of any NASA
mission, or any scientific endeavor for that matter, is a strong
outreach component to draw in public enthusiasm. There are many ways to
communicate NASA's goals for any specific mission. The Mars Exploration
Rover website is an excellent example of an extremely successful
outreach campaign, recording 1.7 billion hits in the week after
*Spirit's* landing and setting the bar far above any previous NASA
missions. As part of the EAGLE design and as part of the NASA Academy in
general, outreach has been integrated to "inspire the next generation of
explorers as only NASA can."

As any student who has participated in a research or design project can
attest, the most valuable learning comes with hands on activities. So,
as a group of students, the NASA Academy has introduced a project that
can provide invaluable research experience for future students. On the
Saturn orbiter phase of EAGLE, there will be mass and power allocated to
an instrument designed and built by students.

The EAGLE project will send out a call for proposals to universities for
project ideas. The proposed instrument must be less than 10 kg in mass,
use less than 5 W of power, and fit within a space 20 x 20 x 50 cm^3^.
It must be an engineering or science experiment that will push the
boundaries of technology or improve humanity's understanding of the
universe. The teams should be comprised of undergraduate and graduate
students supported by one or more academic mentors. Proposals must show
an organizational structure and a timeline of development, taking into
account the fact that the development of this instrument may outlast any
one student's career at a particular institution.

All proposals received by a given date will be considered based on the
goal and organization of the project, the creativity, and the
feasibility. The team will then be matched with a NASA center to
facilitate the design and manufacturing of the instrument. The team will
be involved in all phases of the mission if possible, from blueprint to
launch, giving the students a unique perspective on NASA missions.

The EAGLE mission will be instrumental in facilitating the study of
potential sites of extraterrestrial biological activity. By
incorporating both a lander and orbiter, EAGLE will ensure both breadth
and depth by meeting the science objectives mentioned previously. This
mission is a logical next step to follow Cassini that will increase
scientific knowledge across a variety of disciplines, including geology,
chemistry, and biology. In addition, it serves as a testbed for future
planetary exploration missions such as potential missions to Europa
currently under consideration.

At an estimated \$1.12 billion, EAGLE is considered by NASA to be a
flagship mission. However, this is only one third of the cost of the
Cassini mission currently in orbit around Saturn. It will also provide
much more scientific data than can be achieved by an orbiter alone.
Additionally, multiple sources of heritage have been incorporated in an
effort to mitigate risk and keep cost.

This study is a first cut at defining the science that needs to take
place at Enceladus as well as one potential engineering design that can
be implemented to achieve that science. Future work will entail refining
this design through a series of trade studies and updating it with
knowledge gained from future Cassini discoveries, such as those
anticipated by the next Enceladus encounter in 2008.

Enceladus is currently the center of attention among the astronomical
community and Cassini has merely whet the appetite of scientists all
over the world. Enceladus is a unique moon that clearly needs more study
and may hold clues to the most fundamental question of the universe: Are
we alone? It is in this vein that the EAGLE mission is proposed.

In order to choose the best instruments for EAGLE's specific mission
objectives, a study of the instruments used on other planetary missions
was conducted. The following table lists instrument packages for nine
different missions that were used as trade studies.

+--------------------------+---------------------+---------------------+
| RANGER 4                 | SURVEYOR 3          | NEAR                |
+--------------------------+---------------------+---------------------+
| Camera                   | Camera              | MultiSpectral       |
|                          |                     | Imager (MSI)        |
+--------------------------+---------------------+---------------------+
| Gamma Ray Spectrometer   | Radar Altimeter     | Gamma Ray           |
|                          |                     | Spectrometer        |
+--------------------------+---------------------+---------------------+
| Radar Altimeter          | Temp sensors        | X-Ray Spectrometer  |
+--------------------------+---------------------+---------------------+
| Seismometer              | Strain Gages        | Infrared            |
|                          |                     | Spectrograph        |
+--------------------------+---------------------+---------------------+
|                          |                     | Laser Rangefinder   |
|                          |                     | (NLR)               |
+--------------------------+---------------------+---------------------+
|                          |                     |                     |
+--------------------------+---------------------+---------------------+
| CLEMENTINE               | PHOBOS 1/2          | SELENA              |
+--------------------------+---------------------+---------------------+
| Ultraviolet/Visible      | Camera              | Gamma Ray           |
| Camera                   |                     | Spectrometer        |
+--------------------------+---------------------+---------------------+
| Near-Infrared Camera     | Eadar Altimeter     | X-Ray Spectrometer  |
| (NIR)                    |                     |                     |
+--------------------------+---------------------+---------------------+
| [Thermal Infrared        | Temp. Sensor        | Near-Infrared       |
| spectromet               |                     | Camera (NIR)        |
| er](http://nssdc.gsfc.na |                     |                     |
| sa.gov/database/MasterCa |                     |                     |
| talog?sc=1994-004A&ex=3) |                     |                     |
+--------------------------+---------------------+---------------------+
| Penetrator               | Strain Gages        | [Laser              |
|                          |                     | Alt                 |
|                          |                     | imeter](http://nssd |
|                          |                     | c.gsfc.nasa.gov/dat |
|                          |                     | abase/MasterCatalog |
|                          |                     | ?sc=1994-004A&ex=4) |
+--------------------------+---------------------+---------------------+
| [Star Tracker            | Mass Spectrometer   | Multiband Imager    |
| Camer                    |                     |                     |
| as](http://nssdc.gsfc.na |                     |                     |
| sa.gov/database/MasterCa |                     |                     |
| talog?sc=1994-004A&ex=7) |                     |                     |
+--------------------------+---------------------+---------------------+
| [Laser                   | X-Ray Fluorecence   | Lunar Radar Sound   |
| Altimet                  | Spectrometer        |                     |
| er](http://nssdc.gsfc.na |                     |                     |
| sa.gov/database/MasterCa |                     |                     |
| talog?sc=1994-004A&ex=4) |                     |                     |
+--------------------------+---------------------+---------------------+
| [Bistatic Radar          | APXS                | Magnetometer        |
| Experime                 |                     |                     |
| nt](http://nssdc.gsfc.na |                     |                     |
| sa.gov/database/MasterCa |                     |                     |
| talog?sc=1994-004A&ex=9) |                     |                     |
+--------------------------+---------------------+---------------------+
|                          | P False color image | Upper atmosphere    |
|                          | of Enceladus        | and Plasma Imager   |
|                          |                     |                     |
|                          | enetrator           |                     |
+--------------------------+---------------------+---------------------+
|                          | Thermal Infrared    | Charged Particle    |
|                          | Spectrometer        | Spectrometer        |
+--------------------------+---------------------+---------------------+
|                          | Near Infrared       | Radio Science       |
|                          | Imaging             |                     |
|                          | Spectrometer        |                     |
+--------------------------+---------------------+---------------------+
|                          | Thermal Imaging     |                     |
|                          | Camera              |                     |
+--------------------------+---------------------+---------------------+
|                          | Magnetometer        |                     |
+--------------------------+---------------------+---------------------+
|                          | X-Ray Telescope     |                     |
+--------------------------+---------------------+---------------------+
|                          | Laser Altimeters    |                     |
+--------------------------+---------------------+---------------------+
|                          |                     |                     |
+--------------------------+---------------------+---------------------+
| VEGA 1                   | PIONEER VENUS       | BEPICOLOMBO         |
+--------------------------+---------------------+---------------------+
| LANDER                   | Gas-Phase           | Camera              |
|                          | Chromatograph       |                     |
+--------------------------+---------------------+---------------------+
| Camera                   | Malachite Mass      | UV Spectrometer     |
|                          | Spectrograph        |                     |
+--------------------------+---------------------+---------------------+
| Pressure Sensors         | Infrared Radiometer | X-ray Spectrometer  |
+--------------------------+---------------------+---------------------+
| Tem. Sensors             | Nephelomete         | Visible NEAR-IR     |
|                          |                     | Map. Spect          |
+--------------------------+---------------------+---------------------+
| UV spectrometer          | Solar Flux          | Thermal IR Spect    |
|                          | Radiometer          |                     |
+--------------------------+---------------------+---------------------+
| Gas-Phase Chromatograph  | Temperature Sensors | Laser Altimeter     |
+--------------------------+---------------------+---------------------+
| X-Ray Spectrometer       | Pressure Sensor     | Gammma Spectrometer |
+--------------------------+---------------------+---------------------+
| Malachite Mass           |                     |                     |
| Spectrograph             |                     |                     |
+--------------------------+---------------------+---------------------+
| VM-4 Hygrometer H20      |                     |                     |
+--------------------------+---------------------+---------------------+
| IPF Aerosol Analyzer     |                     |                     |
+--------------------------+---------------------+---------------------+

Based on this study, the following instruments were selected for the
science payload:

  -------------------------- ------------------- -------------------------
  Near-IR spectrometer       Magnetometer        Pancam

  Thermal IR imager          Altimeter           Mass Spectrometer

  Descent imager             Seismometer         Amino Acid Analyzer
  -------------------------- ------------------- -------------------------

Additional trade studies were performed to select instruments for
measuring the elemental composition of the surface and for collecting
collect microscopic images. The following table depicts the instruments
which were considered.

  -------------- -------------- ------------- ---------------------------------------
  Instruments    Mass           Power         Elements

  ChemCam        5.62 kg        6.7 W         Ca,Si,Al,Sr,Fe,Ba,C,Mg,Ti,Na,H,Li,K,O

  APXS           0.57 Kg        0.3 W         Si,S,Cl,Ca,Fe,Ni

  Gamma-Ray      8.6 Kg         3 W           Si,O,Fe,Mg,K,Al,Ca,S,C,H,Th,U

  Neutron        3.9 Kg         2.5 W         C,O,N,Cl,Fe,Ti,Si,Ni,Mg,Al,Ca,K,Th,U
  Spectrometer                                

                                               
  -------------- -------------- ------------- ---------------------------------------

Various combinations of the instruments were considered, but a
combination ChemCam-APXS package was selected since it will minimize the
science payload mass while providing maximum information about the
elemental composition of Enceladus. Since ChemCam includes a remote
microscopic imager, this package also means it is not necessary to
include a separate microscopic imager in the payload.

COSAC- COmetary SAmpling and Composition (from Rosetta)

ECS -- Environmental Control System

EDI-Enceladus Descent Imager

EDL -- Entry, Descent, and Landing

ESAMS-Enceladus Surface Analysis by Mass Spectrometry

FSC-Falling Sample Collector

GC-Gas Chromatograph

ISS-Imaging Science Subsystem (from Cassini)

MAG- Dual-Technique Magnetometer

MARDI-Mars Descent Imager (from MSL)

MER -- Mars Exploration Rover

MLI -- Multi Layer Insulation

MMH- Monomethylhydrazine

MS-Mass Spectrometer

MSL-Mars Science Lab

NTO- Nitrotetroxide

RHU -- Radioisotope Heating Unit

RTG -- Radioisotope Thermal Generator

SOC-Saturn Orbiting Camera

TOF-Time of Flight (mass spectrometry)

TPS -- Thermal Protection System

1.  "Alpha Particle X-Ray Spectrometer (APXS)", NASA, June 20"
    \[http://marsrovers.jpl.nasa.gov/mission/spacecraft_instru_apxs.html\]

2.  "Alpha Particle X-ray Spectrometer", a technical description by the
    Athena group at Cornell, \[retrieved 28 July 2006
    http://athena.cornell.edu/pdf/tb_apxs.pdf\]

3.  Bada, et. al. 2005, Astronomy and Geophysics 46, 26-27

4.  Barber, T.J.; and Cowley, R.T.: "Initial Cassini Propulsion System
    In-Flight Characterization," AIAA Paper 2002-4152, July 2002.

5.  Bell, J. F. et. al., "Mars Exploration Rover Athena Panoramic Camera
    (Pancam) investigation", *Journal of Geophysical Research*, Vol.
    108, No. E12, 8063, doi:10.1029/2003JE002070, 2003.

6.  Berry, Robert L; Eberhardt, Ralph N; Leeds, Michael W: "Development
    of the Cassini spacecraft propulsion system," AIAA Paper 96-2864,
    July 1996.

7.  Blum, D., Bourque C., Mehta, G., "The Alloy with a Mission" Berkley
    Report, URL:
    \[http://www.mse.berkeley.edu/classes/matsci102/F01reports/allithium.pdf\]

8.  The Boeing Company, "Delta IV Payload Planners guide," Boeing,
    Huntington Beach, Ca, 1998

9.  Braukus, M., Malone, J., "Shuttle Super Lightweight Fuel Tank
    Completes Test Series**"** *NASA* , URL:
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