1.0 Executive Summary
1.1 Brief overall description of Lander/Probe
In Greek mythology, Europa is a mortal concubine of Zeus. Zeus carried Europa off
to a remote island and had several demigod children with her. Our probe takes the name
of Zeus in the spirit of life which is embodied in this story. Life forming on a distant
deserted island is exactly what may have occurred on Jupiter’s moon Europa. Along the
same lines as in mythology, it is hoped that the union of the moon Europa with our Zeus
spacecraft will present to the world the fruits of nature in the form of life on another
world.
According to the 1997/98 AIAA Undergraduate Team Space Design Competition
RFP, “The objective of this project is to produce a complete system design for a
spacecraft that can land on the surface of Jupiter’s moon Europa and examine the ice and
water of which it is made.” To accomplish this objective, the lander/orbiter and the
ice/water probe need to possess certain instruments and experiments. The spacecraft
needs to re-map certain areas of Europa that have been selected from data from the
previous Europa Orbiter mission. This mapping will be done from a 20km orbit and will
include photography and surface topography. Once the spacecraft lands, surface tests
must be conducted to determine the physical and chemical properties of Europa.
Photography and seismic testing are also essential portions of the lander/orbiter’s
mission. Once the spacecraft lands, the ice/water probe can be released. This probe has
the duty of providing chemical analysis, taking photographs, and determining the
pressure and temperature distribution beneath the surface of Europa, as well as
conducting life detection experiments.
The Zeus spacecraft is furnished with the RIEGL Laser Altimeter LD90-31K to aid in
mapping the surface of Europa. The lander is equipped with 1024x1024x24-bit CCD
cameras, which are also used for surface mapping from orbit, as well as providing
photographs of the surface once the spacecraft has landed. The lander carries a robotic
arm, which is modeled after the Mars Surveyor 2001 lander. An optical spectrometer is
attached to this arm to perform chemical analysis of the Europa’s surface. To observe the
ice shifting and “Moonquakes,” a JPL Microseismometer is imbedded in one of the feet
of the landing gear to perform these measurements. Before any of the scientific
operations can occur, the spacecraft must be able to maneuver to an orbit around Europa.
Zeus’s three main engines provide a maximum thrust of 488 N and use
monomethelhydrazine and nitrogen tetraoxide propellants. The ten monopropellant
hydrazine thrusters on this spacecraft provide a maximum thrust of 6.15 N each. The
power is generated by a Radioactive Power Source (RPS), which converts thermal energy
to electrical power. The communications system consists of a parabolic high gain
antenna constructed out of an aluminum honeycomb. A low gain antenna is used for low
data rate communications and in case of high gain antenna malfunction. The spacecraft
guidance and navigation is performed by autonomous star trackers, Fine Sun Sensors
from the Swedish Space Corporation, an Analog Devices/ ADXL05 accelerometer, and a
Litton LN-200 Fiber Optic Inertial Measurement Unit. To organize all of these
instruments, a redundant computer is used to conduct the operations of the spacecraft.
The ice/water probe, like the lander/orbiter, is also required to perform scientific
operations. The chemical analysis is executed by capillary electrophoresis system. This
system separates molecules based on their movement through a fluid under the influence
of an applied electric field (Weinberger, 1993). A set of cameras is used to take pictures
near the ice/water boundary. The cameras are equipped with halogen lamps to provide
the light within the underdwellings of Europa. These cameras view the inside of the
moon through ports provided for the transducers used to measure the pressure and
temperature distribution of Europa. The ice/water probe contains instruments vital to the
success of this portion of the mission. Like the lander/orbiter, the ice/water probe houses
an RTG for power. In addition to the RTG, the probe contains radioactive heating units
(RHU) which aid in melting the ice for its journey to the ice/water boundary. Once the
probe has melted its way down to about 30 m from the ice/water boundary, the Probe
Arresting Sub-System (PASS) is activated. This system consists of titanium blades that,
when released, hold the upper portion of the probe in the ice while the lower portion
continues to fall into the hopefully present water.
The purpose for the PASS is to ensure contact with the ice for the communications
between the ice/water probe and the lander. The communications system between the
probe and the lander uses sonar through the ice. This system consists of hydrophones and
an acoustic modem. Sound, instead of an electrical impulse, is sent through the ice. This
2
means that a cable between the probe and the lander is not necessary. The sonar system
can also be used to determine the depth of the ice and the distance from the probe to the
lander.
1.2 Mission Events Sequence
Europan Orbit Entry
Ice Thickness
Measurements
Pictures
Surface
Altimeter
Measurements
Landing
Surface
Chemical
Analysis
Seismic Readings
Pictures
Pressure, Temperature
and Composition Readings
PASS Activation
Sub-Surface Ocean Entry
Pressure and
Temperature
Readings
Photographs
Water Composition
Experiments
Figure 1.1: Mission Events Sequence
3
Probe Release
1.3 Mission Time-Line
The following is a time-line explaining the order and date of key aspects of the mission.
YEAR
Mission Event
1999 2000 2001 2002 2003 2004 2005 2006 2007 2008 2009 2010 2011 2012 2013 2014 2015 2016
Initial Approval/ Funding of Mission
Completion of R&D, tests and evaluation
Construction of Spacecraft
Preparation of Launch/ Launch
Space journey up to arrival near Jupiter
Surface scans begin/ Landing on Europa
Reach Ice/Water Interface (assuming between 100m up to 5km)
4
2.0 Introduction
The exploration of the world and universe has always been a preoccupation for
humans. Since the dawn of our species, as we traveled across continents to settle into
different territories, to the modern space age, we have spent tremendous resources on
expanding our horizons. In the last few years of the 20th century, we are on the cusp of
discovering life outside of our planet. New telescopes are able to see planets forming
around distant stars, and planets and moons in our own solar system show promise of
having what it takes to form life as we know it—liquid water. With the discovery of
oceans covered by a thick ice crust on Europa, the possibility of finding life is more
imminent than at any other time.
As part of this study of our solar system, and our quest for life outside of our planet,
NASA has proposed a possible mission to Europa. This mission is going to do a detailed
analysis of the surface and subsurface of Europa. No probe prior to this one will have
probed these features in as much detail. This will be the first time a probe will be landed
on a Jovian moon. The choice of Europa is based on the interest surrounding the liquid
water, which is theorized to be present.
As a moon of Jupiter, Europa is brought under the gravitational forces of the second
largest body in the solar system. Since Jupiter comprises as much mass as all other
bodies in orbit around the sun, any object which comes in close proximity to this planet
experiences huge tidal forces. These huge tidal forces continually oscillate through
Europa’s orbit, due to the oblateness of Jupiter. Further tidal forces are encountered as
nearby moons are passed. This motion continually generates heat in the center of Europa.
It is theorized that this heating is what causes the liquid ocean under the crust to exist.
The scientific community has only recently accepted the concept of a floating ice
shell. The first real evidence of this situation has come from high resolution, <1km,
imaging from the Galileo probe. The fracture level and local plate shifting shown in
these images quickly lead scientists to conclude that the ice was indeed floating. The
nearly complete absence of craters also points to floating ice, since large objects drawn
into the Jovian gravity will bombard any object orbiting Jupiter.
The surface of Europa is a thick and relatively smooth coating of ice. There are no
mountains or canyons. There is practically no appreciable atmosphere either, since
5
Europa doesn’t generate enough gravity to hold one. Current gravity analysis and plate
motions lead scientists to conclude that Europa has a 1-10km ice crust followed by
approximately 100km of ocean. The brine content of the ice and the water is not known
at this time. At the core of the moon is a rocky center—probably iron.
To determine the accuracy of these predictions our probe will employ several
experiments meant to study these features. From orbit a detailed photographic and
topographical study will be carried out over a few pre-selected regions. This will
augment a previous study done by an orbiting NASA probe before our arrival. Besides
using this information for scientific purposes, NASA will use this data for final selection
of the landing site. Upon making a soft landing, a probe designed to penetrate through
the ice will be released. This probe will melt through the ice and study the subterranean
oceans. Pictures will also be taken of this undersea environment and transmitted to earth.
Chemical analysis will also be carried out to better determine the composition of
Europa. The lander will perform one set of chemical analysis, and the probe will perform
another. The lander experiment is modeled after the spectrometry experiment, which is
planned for the Mars Lander 2001. Along with studying the elemental composition, the
probe will also conduct experiments to determine if life has ever existed on Europa.
With the current array of science payloads, the Zeus probe will be able to provide
new insights into Europa, and by extension our own planet.
6
2.1 General Performance Parameters of the Configuration
Table 2.1: Weights and power requirements of major components
UNIT
PROBE
Power Generation
Capillary Electrophoresis
Probe Shell
PASS
Photographic Equipment
Acoustic Modem
Computer
Pressure Transducers
Thermocouple amp
PROBE TOTAL
WEIGHT (kg)
POWER (Watts)
20.72
5.5
14
8.571
0.6
1.1
1
0.1
0.25
51.84
44-62.5 (generated)
10.75
20
0.6 to 10.1
1.66
5
0.1
0.1
7
LANDER
Laser Altimeter
Arm/Spectrometer
Fixed Camera(3)
Zoom Cameras(4)
Seismometer
Main Propellant Tanks
Main Engine Propellant
RCS Tank
RCS Propellant
RCS Thrusters
Helium Tank
Helium
Engines
Valves
Radioisotope Power
Source
Communications System
Star Tracker (2)
Sun Sensors(4)
Accelerometer
IMU
Computers
Radiator Fins
Heat Pipes
Jacket
Acoustic Modem
Radiation Shielding
1.5
2.5
0.25
2
0.1
48
956
3.63
30.25
3.6
31.3
5.89
13.5
5
7.96
1
0.25
1.5
2
0.1
5
2
Generates 163W
35
2.8
1
5x10-3
0.7
2
7.172
1.42
3.34
1.1
10
53(High Gain), 20 (Low Gain)
7
1.4
0.35
10
20
-
LANDER TOTAL
DRY MASS
TOTAL(inc. propellant)
1210.6
306.5
1262.5
3.0 Design Evolution
The design of the spacecraft has gone through several stages over the course of the
project. The arrangement of all of the components and the components themselves were
selected to minimize size and complexity. To keep launch costs to a minimum, we
decided to limit ourselves to the payload capability of the Atlas IIA launcher. As the
design progressed, the launch vehicle became the Delta III, which has even less stringent
mass and size constraints. Several major iterations were done to perfect the design, and
each is presented below.
From the beginning the structure has been a truss made of tubular graphite epoxy
members connected by titanium joints. Similarly all designs have had a set of three
8
engines located centrally. The ice water probe has also remained at the center of the
configuration to make it easier to mount and to protect from radiation. The landing pad
design has also been held constant. To increase traction on the ice surface, a series of
small spikes line the landing pads. This will reduce any skidding which may occur if the
spacecraft lands on a slight incline.
Figure 3.1: Initial Design Concept
Figure 3.1 shows the initial design of the spacecraft. Obscured from view is the
probe. This crude drawing also lacks the instrumentation section, which would have
been mounted on a palette directly on top of the tanks. At this point two primary factors
drove us towards this design. First of all, we assumed that we wanted a full 360 degrees
of motion for the main antenna. This explains why it is so high above the main lander
body. At this point the antenna was also 3m in diameter to accommodate a data
throughput rate which proved too extravagant. The tanks are mounted directly to the
structure at two hard points on each tank. Each leg of the landing gear is a single member
which connects into the center of the main structure.
There were several problems with this design. First of all, the landing gear base was
too small. The structure lacked inherent stability. The tank mounting was also a problem
since local stresses at the hard points would have been too large at launch. The antenna
in such a position added too much mass to the structure and further contributed to
stability problems. After some analysis it became apparent that it would be possible to
land with an orientation that would allow for effective communication with only 180
9
degrees of rotation. The mounting of the instrumentation on a palette above the tanks
proved unwise also. First of all, the instrumentation was not large enough to need their
own separate section of the spacecraft; it is adequate for them to be stored throughout the
spacecraft. Second of all, in this configuration very heavy radiation shielding would be
necessary to protect the instrumentation. At this stage we had a radio echo sounder
(RES) for ice thickness measuring and a far range spectrometer. Both of these were
abandoned when NASA released news that a separate mission would carry out these
tasks. At this stage there was no robotic arm or seismometer. The sonar transducer was
also mounted above the ice/water probe.
Figure 3.2: Second Iteration of Design Evolution
After more careful analysis a second major design was arrived at. This design
incorporated the lessons learned from the first iteration, and some basic design changes
that occurred between the two iterations. The second iteration can be seen in Figure 3.2.
By the time the second iteration was started the RES system and the far range
spectrometer were removed from the design. The sonar transducer was mounted in one
of the landing pads. A seismometer was also added to the design, on one of the landing
pads. Both of these instruments need a firm contact with the surface. This is provided by
the lander’s weight.
10
The tank arrangement was altered slightly, and the formerly symmetric shape was
stretched to accompany the side-mounted antenna. The antenna, which is now only two
meters in diameter, was moved from its top position to the side to minimize the height of
the probe and to reduce the amount of additional structural support needed for the dish.
The computer and electronics are mounted internally, surrounded by tanks. This provides
a shield from the radiation and reduces the amount of radiation shielding that is
necessary. The tanks are now mounted on a set of equatorial support straps. These straps
are attached to the structure via hard points. The landing gear has been spread out to
extend further away from the center of gravity.
While this design is more stable, there were still some problems that had to be
overcome. First of all, the landing gear legs were only supported in one plane. This
caused their thickness to be too large. Extra bracing was decided to be the solution for
this problem. The tanks still had mounting problems. While the mounting stresses were
distributed throughout the tank, the hard points were on the same side of the tank. This
means that excessive moments would be generated under accelerations. To reduce these
moments, the hard points were moved to be on opposite sides of the tank. This design
also did not explicitly place the RCS fuel tanks or the helium storage tank. At this stage
the thermal analysis was not completed, so thermal control systems were not mounted
either.
All the design lessons from the second iteration were applied to the final iteration, see
Figures 3.3 through 3.6. Furthermore any equipment which was not included explicitly
in the second design, such as thermal management equipment or storage tanks, are now
present. All equipment on the lander can be seen in these two views. This system has an
initial total weight of 1262.5 kg, including propellant. The length from dish reflector to
the far tank is 4.1m and the height of the system is 2.2m. These parameters are well
within proposed Delta III launch limits.
11
Figure 3.3: Isometric view of final spacecraft configuration
12
Figure 3.4: Side view of final configuration
13
Figure 3.5: Top view of final configuration
14
Figure 3.6: Internal View of the Final lander Configuration
15
4.0 Lander/Orbiter
4.1 Overall Description
The lander/orbiter is the primary bus off which every major system operates. The
lander/orbiter contains the fuel and engines required to make orbital changes necessary
for course corrections, Jovian orbital insertion and Europan orbital insertion. The main
engines are a set of three bi-propellant rocket engines located at the bottom of the probe
designed to run on monomethalhydrazine (MMH) and nitrogen tetraoxide (NTO). The
spacecraft itself is three-axis stabilized and uses hydrazine thrusters for attitude control.
Orientation is determined with a series of star trackers and sun sensors mounted around
the outside of the probe. A set of inertial measurement units and accelerometers is also
used to help the guidance and control system.
The structural arrangement can be seen in Figure 3.3. From the tip of the antenna
reflector to the back of the tanks is 4.1m and the height from the helium tank to the base
of the legs is 2.2m. At launch the spacecraft will have a total mass of 1262.5kg. The
primary structure is made of tubular graphite epoxy rods fastened with titanium joints.
The tanks are situated to provide radiation shielding for most of the sensitive
instrumentation. The tanks themselves are connected to the main structure via equatorial
support rings. There are six main bi-propellant tanks, one monopropellant tank and one
helium storage tank. The engines are mounted on a separate set of struts that connect into
the main structure at the bottom.
The ice/water probe is mounted in the center of the probe to provide maximum
radiation shielding. The probe is held in position with a metal strap similar to the ones
which are used to support the tanks. This strap is fastened together via incendiary bolts,
which will be activated after landing to release the tension in the strap and allow the
probe to fall through to the surface. Storing the probe in the center of the bus required us
to remove the excess heat energy generated by the probe's radioactive heating system.
This energy is radiated into a metal jacket which surrounds the probe. The energy is then
drawn out to a series of radiators on the side of the lander, and radiated into space. Power
is generated by an AMTEC radioactive power source (RPS). This will provide electrical
power for all on-board components.
16
Equipment, such as the main computer and communications power electronics, is
mounted on the structure using brackets attached to the titanium joints. Sensitive
electronics were kept in the center of the probe as much as possible, however any
instrument that needed a clear field of view was moved to the exterior. Cameras are
mounted throughout the spacecraft to provide imaging during orbital and surface
operations.
From a 20km orbit, the spacecraft will take photographs and take detailed
topographic measurements of selected areas of the surface. A laser altimeter mounted on
the spacecraft will provide the topographic measurements. Both of these instruments
have a resolution of up to 1m. The laser altimeter is pointed by reorienting the
spacecraft, but the cameras can be repositioned by panning.
After landing the spacecraft will release the ice/water probe and begin studying the
surface features. A seismometer mounted in one of the landing pads will provide threedimensional seismic data. A sonar transducer and hydrophone array mounted in other
pads will be used to study fracture level of the ice beneath the lander, as well as provide
two-way communications with the ice/water probe. Photography will continue on the
surface via cameras mounted throughout the probe. A robotic arm with a sampling claw
will be used to perform spectrographic experiments on the surface. Small pressure
transducers and thermocouples are also used to determine the temperature of the ice and
the space surrounding the spacecraft.
4.2 Science Operations
4.2.1 Surface Topography Mapping with Laser Altimeter
A laser altimeter onboard the lander will be used to map the topography of Europa.
When the lander camera records an image, the apparent horizontal locations of features
within that image are distorted by elevation variations. These distortions can be easily
removed if the topography is known. The detailed analysis of three-dimensional shapes
of geologic features can also yield a greater understanding of the processes that formed
and later modified them. Moreover, this instrument will aid in determination of altitude
at landing and landing site selection.
17
The basic principles behind the laser altimeter’s operation are simple. The instrument
sends out a laser pulse and then times how long it takes for the reflection to return. The
altitude is the time of flight divided by twice the speed of light.
The laser altimeter chosen for our mission is the RIEGL Laser Altimeter LD90-31K.
The following table provides the specifications of this instrument.
Table 4.1: Laser Altimeter Characteristics (RIEGL)
Characteristics
Maximum Measuring Range
 Surface reflectivity  80%
 Surface reflectivity  10%
Minimum Distance
Accuracy
Dimensions (mm)
Power (W)
Weight (kg)
 > 1500 m
 > 500 m
2m
20 cm
200 x 130 x 76 (L x W x H)
6
Approx. 1.5
4.2.2 Surface Test
To deal with surface chemical analysis, a robotic arm has been attached to the side of
the lander, see Figure 4.1. This arm has been modeled after the Mars Surveyor 2001
lander. The arm is made of graphite composites and has a dual articulating shoulder and
elbow joint. The end of the arm has a sample collection shovel and a spectrometer sensor
head. Chemical analysis is done via a wide-wavelength spectrometer inside of the
instrument section of the spacecraft. The spectrometer is connected to the sensor by a
fiber optic cable. A possible spectrometer would be the Ocean Optics S2000
spectrometer, which is a 2048 element CCD sensitive to wavelengths between 200 and
1100 nm(Ocean Optics). It is possible to change the CCD to make it selective to another
range of wavelengths.
18
Elbow Joint
2-Degree of
Claw
Freedom Shoulder
Optics for
Spectrometer
Figure 4.1: Sketch of the lander arm
4.2.3 Photography
Photography will be essential while in orbit and on the ground. The requirements for
each of these are separate. A series of small light weight cameras are mounted on the
side of the helium storage tank. One set is used for the orbit photography and the other
set is used for pictures after landing. The difference between the different sets of camera
is the optics used for focusing and zooming.
Each of the cameras is a 1024x1024x24-bit CCD camera. The camera has shutter
times from 1 sec to 500 s. The cameras are radiation hardened to protect their sensitive
components, such as CCDs, from the Jovian radiation environment. An illustration of the
two sets of cameras can be seen in Figure 4.2. This location provides for a complete field
of view around the spacecraft.
19
Helium Tank
Camera Optics
Figure 4.2: The cameras are the small cylinders on the tank sides
Each set of lenses is adjustable by the camera to improve focus or to create a more
magnified view. The lenses are moved via a set of small motors which move the optics
along the tube. Field of view is selected by rotating a flat mirror which sits at a 45 degree
angle with the main line of the camera, see Figure 4.3.
Panning Mirror
Lenses
CCD
Guide
Tracks
Radiation
Shell
Figure 4.3: Interior and exterior view of camera, respectively
20
Before a photograph is stored, the camera will try to focus the image. After focusing,
an optimum shutter time is computed. Once the shutter time is determined, the image is
captured. All of these operations are done onboard the camera, but they are initiated by
the main computer. The image is then sent to the data compression application specific
integrated circuit (ASIC). This ASIC performs a JPEG compression on the photograph.
This compression can have any user-defined ratio, however the default compression ratio
is 8:1. After compression, the ASIC can further reduce the size of the image by reducing
the color depth. The color depth can be reduced from 24 bits to 16,12, 8 or 4 bits per
pixel. Once the final image has been computed by the ASIC, the information is sent to
the main computer for handling. The image can either be stored in memory or in storage.
Details about the compression ASIC can be found in Appendix 11.2.
Fixed cameras are also located at various positions throughout the probe to provide
pictures of different key systems. These cameras will be used to determine the status of
different components during specific parts of the mission. They could also be used to
take more pictures of the surface after landing. These cameras use fixed focal length
lenses and have 1024x1024x24 bit CCDs.
4.2.4 Seismic Sounding
The importance of seismology in understanding the interior structure and tectonics of
a planet or moon cannot be overstated. It is the only tool available that can furnish
detailed global and regional information on the compositional structure and physical state
of the interior. Whereas gravity, dynamics, and magnetic measurements can supply key
information on some aspects of interior structure, they cannot provide a substitute for the
precise radial (and, to a lesser extent, lateral) structure information that can be derived
from seismic data [Banerdt,1998].
The physical quantity that is measured with seismic instruments is normally the
displacement of the ground with respect to time (time series) resulting from propagation
of seismic waves through the interior. The choice of planetary body and specific
measurement objectives dictate to some extent the required response characteristics of a
seismometer, e.g., its frequency range, sensitivity, and dynamic range. While instrument
21
sensitivity as high as possible is desired for a body of unknown seismic activity in order
to be prepared for a very low level of activity, the local level of ground noise generally
limits the usable instrumental sensitivity. We are fortunate with Europa, however,
because expected background noise is extremely low on a body with no atmosphere. For
bodies with atmospheres, some noise due to wind is expected.
The number and distribution of seismic stations recording simultaneously on a
planetary surface greatly influence the quality and quantity of information that can be
derived from the acquired data. Generally, the larger the number of stations, the more
detailed information on the planetary interior can be obtained. Due to the nature of our
mission, a large number of stations is impossible. Nevertheless, even if we have only one
station, we can still expect certain useful information. Data from one seismic station will
tell us at least: (1) the level and other properties of the background noise and (2) the
characteristics of the signal that may prove to be from natural seismic events. In the
initial stages of exploration of a planetary body, such information is highly valuable in
designing instruments that are to be deployed later to maximize the information return in
future missions. In addition, if the internally generated seismic signals can be
unambiguously differentiated from noise, the level of seismic activity of the planet can be
determined, and if there is a seismic event large enough to excite detectable normal
modes, even a single station with sufficiently low frequency response may also provide
deep structural information.
The mission requirements were assessed and a seismometer chosen accordingly. The
seismometer of choice is the JPL Microseismometer (JPLMS). The JPLMS is a state-ofthe-art technology implementing the practice of micromachining. Due to the fabrication
process of the seismometer, an extremely small instrument can be produced. Below is a
picture of the instrument.
22
Figure 4.4: JPL Micromachined Seismometer
In addition, the specifications for this instrument can be seen below in Table 4.2.
Table 4.2: Characteristics of the JPL seismometer
Seismometer Parameters
Suspension
Transducer
Configuration
Mass
Power Consumption
Size
Acceleration Sensitivity
Frequency Range
Micromachined Silicon, 10Hz Resonance, 6x10-9
m/sec2/Hz Noise Floor
UHF Capacitive Displacement, 5x10-13 m/Hz Sensitivity
Tetrahedral (3 Components of acceleration plus 1
redundant)
<0.1kg
100mW
5cm on edge
Better than 10-8 m/sec2
0.01-100 Hz
4.3 Lander Subsystems
4.3.1 Propulsion
4.3.1.1 Main Propellant sizing
The main propellant system is designed to satisfy requirements for reaching an orbit
around Jupiter, and then Europa. It also takes into account the propellant required for
landing from a 20 km orbit. Because of the complexity of calculating the orbit insertion
23
for Jupiter, it is assumed that the spacecraft will reach Jupiter using the same VenusEarth gravity assists as Galileo. Assuming a series of gravity assists in the Jovian system,
Helium
Helium
Helium
MMH
MMH
Combustion
Chamber
Figure 4.5: Simplified diagram of the main propellant system
a total v of 2.5 km/s will be required to reach a 20 km orbit around Europa. The v
required for landing from this orbit will be 1.59 km/s. If an additional contingency v of
0.12 km/s is assumed, then the total v will be 4.22 km/s.
To calculate the required propellant for this v an initial dry mass of 365kg, including
an estimated 65kg for tanks, was used. The engine Isp was assumed to be 320 s. Using
the rocket equation:
  v  
  1
m p  m f exp 

I
g
  sp  
(4.1)
where v is in m/s, g is the gravitation constant for earth, and mf is the dry mass. With
these numbers and a contingency factor of 10%, the total mass of propellant is 1138 kg.
After performing the below calculations for tank weight, the calculation was rerun for the
actual dry mass, of 306.5kg. This yields a final propellant mass of 956kg.
To achieve an Isp of 320 s, a bipropellant rocket engine will be required. The chosen
propellants are monomethelhydrazine (MMH) and nitrogen tetraoxide (NTO). An
24
oxidizer-to-fuel ratio of 1.65 is assumed. Using the original calculation for propellant
mass, this means that there will be 429.4 kg of MMH and 708.6 kg of NTO. To aid in
tank construction, all the tanks will have the same volume. There will be a total of six
storage tanks, three for NTO and three for MMH. Each engine will be connected to its
own tank set as shown in Figure 4.5. This reduces the complexity of this system. If a 3%
spillage is assumed, then a tank volume of 0.1682 m3 is required. The internal diameter
of this tank will be approximately 0.6848 m.
The tank thickness is calculated using Equation 4.2:
t
 PDi m p g 



 tens,allow  4
Di 
FS
(4.2)
where FS is the factor of safety,  is the allowable tensile stress, P is the pressure, Di is
the internal diameter, mp is the propellant mass, and g is the local acceleration. An axial
launch of 6g’s is used. The allowable stress for graphite epoxy composite is 900 MPa.
The tank contents are at 17 bar. A tank wall thickness of 1 mm will be required. To seal
the tank, a thin inner liner of aluminum is applied. Thermal insulation MLI is provided
on the outside of the tank which is approximately 2.5 cm thick. Additional structure
inside the tank, such as insulation and heaters will further increase the tank mass by 2.25
kg. The mass of each tank will therefore be 8 kg.
These tanks are charged by an external helium tank which contains helium at 320 bar.
The amount of helium needed was found by calculating the mass of helium at 30 bar
needed to completely fill all six propellant tanks. With an 3% contingency factor, it will
take 5.89 kg of helium. The density of helium at 320 bar, using the Redlich-KuongSoave equation of state is 44.44kg/m3. The helium tank size will therefore be 0.13 m2.
This tank has a required inner diameter of 0.631 m. Again, using Equation 4.2 and
assuming a helium mass of 5.89 kg, the required tank thickness will be 11.5 mm. The
mass of the helium tank, with liner and structure will be 31.3 kg.
All these tanks have variable-opening valves, which control the flow of propellant
and helium throughout the system. The on-board computer determines the required firing
rates for each engine, and then regulates the propellant flow rate by accordingly varying
the helium exit pressure and the propellant storage line percent opening. By adjusting
25
these two values, the computer can accurately control the oxidizer-fuel ratio and thrust
level for each engine. All the valves have an assumed mass of 5 kg.
The mass of all six propellant tanks is 48 kg. The total mass of the system without
the propellant will be 90.2kg.
4.3.1.2 Main Engine Sizing
The main engines are used for orbit changes and insertion, and for landing. These
engines must provide enough thrust to complete orbit insertion in a reasonably short time.
This will significantly simplify calculating orbital trajectories. The engines also must
create enough thrust to soft land on Europa.
To soft land on Europa the engines should be able to match the gravitational
acceleration of Europa (1.3 m/s2). The mass of the vehicle at landing, assuming 30% of
the RCS propellant and 12% of the main propellant are left, is approximately 400kg. In
order to match the gravitational pull of Europa, 520 N of thrust needs to be developed.
The ability to land with an engine out would mean that each engine should develop at
least 260N of thrust.
Because the orbit insertion maneuvers will require a more substantial amount of
energy dissipation, and the spacecraft will weigh more during these maneuvers, engines
with a nominal thrust of 450 N were chosen. This will ensure the spacecraft is capable of
performing these burns in an acceptable amount of time. Such engines would typically
have a base diameter of 35cm, a height of 70cm and weigh 4.5kg.
4.3.2 Structures
No vehicle can perform its mission without a properly designed structure. The
structural design of any serious concept is a lengthy process. The Zeus probe/lander
mission to Europa is no exception. Table 4.3 lists many of the design concerns affecting
the structural design for the Zeus probe/lander mission. Based on this table the Zeus
lander has to be able to perform with integrity under many unique conditions.
26
Table 4.3: Concerns in structural design
Considerations/Design Concerns
Launch and maneuver loading
Extreme Radiation levels
Attachments/Mating to Delta III fairing
Orientation/Placement of key components
Thermal control
Europan surface conditions
Minimizing cost, size, complexity
Resulting Issues
Structural assembly optimization
Sensitive component placement to
minimize shielding required to survive
Jovian radiation.
Overall configuration constraints.
Specific component mission
conflicts/requirements.
Placement of Ice/Water probe and thermal
control devices.
Avoid tip-over and surface non-conformity
Component/Equipment choice; an overall
design concept motivator.
First, the structure must be designed to survive the 6g launch load. The structure
must also be designed to withstand accelerations and loading during full thrust maneuvers
with the main engines. This would occur during orbit insertion around Jupiter, during
landing maneuvers, and during minor course corrections throughout the flight. Finally
the structure will have to be able to cope with surviving a landing on Europa. Such a
landing could occur with up to a 2 ge decceleration (worse case).
Instrument/Equipment placement must take into account its function
requirements, and allowable size of the lander for fitting into the Delta III fairing.
Configuration also comes into play when attempting to minimize the amount of radiation
shielding used. The computers and sensitive electronic instruments are of primary
concern. In addition, these items must be placed to optimize mission effectiveness, while
complementing the structure’s roll in the mission, as defined above. All of these
concerns must be dealt with and constrained by the need to minimize the cost, complexity
and size of the entire vehicle.
These considerations lead to specific requirements for material properties, and a
detailed analysis of support members to ensure mission reliability, minimal costs, and
minimal weight of entire structure. The truss member material was chosen for the best
overall ultimate strength to density ratio. Materials investigated were Titanium,
27
Aluminum 7075-T6, Steel (Quenched and Tempered), and Graphite/Epoxy 8-ply quasiisotropic composite. Figure 4.6 & 4.7 compares the ultimate tensile and compressive
strengths to density ratios of these materials in order to assist in the proper material
selection.
Ultimate Stress (MPa)
Chart of investigated material properties
1000
800
Aluminum
Gr/Ep
Steel
Titanium
600
400
200
0
Ultimate Tensile Stress
Ultimate Compression Stress
Figure 4.6: Chart of investigated material properties
Stress to Density
Ratio(m^2)/sec^2
Chart of Stress Ratios for Investigated Materials
0.25
0.2
Aluminum
0.15
Gr/Ep
0.1
Steel
Titanium
0.05
0
Ultimate Tensile Stress
Ultimate Compression Stress
Figure 4.7: Chart of Stress Ratios for Investigated Materials
The Graphite/Epoxy composite and the Aluminum 7075-T6 materials showed
great promise according to this initial analysis. To discern the better of the two materials,
further analysis was required. Using Excel 97, it was possible to determine the material,
28
which would require the least mass. This was done using a hollow, uniform, 1 meter rod
with a variable external diameter, a 4cm internal diameter, an allowable deflection limit
of .25cm under a 1 kN load. Figure 4.8 illustrates this analysis, and shows that the
Graphite Epoxy composite is the best overall choice for the truss members. However,
since composites typically perform poorly under shearing conditions, commonly
experienced at the joint between two members, titanium will be used at the joints.
Titanium is used, since it has the combination of low elasticity, 115 GPa, and a high
shear strength of 830 MPa.
Mass (kg)
Mass of Tested Material Bar for the Test Load
0.022
0.0215
0.021
0.0205
0.02
0.0195
0.019
0.0185
Aluminum
Gr/Ep
Titanium
Tested Material
Figure 4.8: Mass of Tested Material Bar for the Test Load
4.3.3 Lander Power
The lander system generates power with an alkali metal thermal-to-electric converter
(AMTEC) radioactive power source (RPS). To determine the amount of electrical power
necessary for this mission, the lander was considered under different loading conditions.
Six loading conditions were considered, see Table 4.4. For a detailed breakdown of the
power for each component, see Section 2.1.
29
Table 4.4: Power Configurations for the Spacecraft
Mode
1
Description
Transfer Orbit
2
Europa Orbit with Data
Acquisition
3
Europa Orbit with Data
Acquisition and High Speed Data
Transmission
4
Landing
5
Post Landing Operations
Components
Computer, Low
Data Rate
Communications,
RCS, GN&C
Computer RCS,
GN&C,1-Camera,
Laser Altimeter,
Low Data Rate
Communications
Computer RCS,
GN&C,1-Camera,
Laser Altimeter,
High Data Rate
Communications
Computer, RCS,
Engines/Valves,
GN&C, Camera,
Laser Altimeter
High Data Rate
Communication, All
Cameras,
Seismometer, Sonar,
Computer,
Arm/Spectrometer
Power
50.75W
62.25W
112W
43.25W
81.51W
From the above table it is clear that we need at least 112W of electrical power. This
power is more than adequately developed by our RPS. Our RPS unit will have three
general purpose heat sources (GPHS) with six AMTEC cells on each side, and is
modeled after the planned RPS for the Pluto Express mission. The total weight of the
system including the GPHS modules is 7.96 kg. The power generator supplies 163
electrical Watts and operates at a total system efficiency of 21.8%. More information can
be found in Table 4.5 (Sievers and Ivanok, 1995). Zeus has a maximum power load of
130.25 Watts up until the end of the mission. For more information on this design see
Appendix 11.1 or (Sievers and Ivanok, 1995).
30
Table 4.5: Predicted AMTEC-RPS Performance
Generator Mass
Thermal Power
Output Voltage
Output Current
Output Power
System Efficiency
Specific Power
7.96 Kg
750 Watts
28 V
3.9 amps
163.5 Watts
21.8%
19.3 Watts/Kg
4.3.4 Data Management
Communications between the Zeus and Earth take place via a parabolic high gain
antenna mounted on the side of the spacecraft (see Figure 4.9). This antenna is designed
to operate both while the spacecraft is in orbit and after landing. To steer the dish, a set
of two step motors provides for a full range of required motion. In orbit the dish can be
steered with the actuators or can be steered via the Attitude Control System (ACS)
reorienting the entire craft.
2 meter, High Gain
Antenna Dish
Transmitter Horn
Low Gain Antenna
Figure 4.9: Antenna position on probe
To calculate the required electrical power for transmission, the standard link equation
was used. The link equation is (Space Mission Analysis):
Eb
 P  Gt  Ls  Gr  L  288.6  10 log Ts  10 log R
N0
31
(4.3)
where Eb/N0 is the bit energy to noise spectral density ratio, P is output power, Gt is the
transmitter gain, Ls is the space loss, Gr is the receiver gain, L is the pointing error loss,
Ts is the temperature loss parameter and R is the transmission rate in bps. Each term on
the right is in decibels.
Gr of the 34m DNS antennae is 73.7 (Yuen 1982). A reasonable pointing error loss is 0.8
dB (Yuen 1982). The temperature parameter was taken directly from the Galileo
communications calculations, and it is 27.5 K noise (Yuen 1982).
The equation for the other parameters are:
Gt  159.9  20 log D  20 log f Hz  10 log 
(4.4)
Ls  147.55  20 log S m  2 log f Hz
(4.5)
where D is the dish diameter in meters, f is the frequency in Hertz,  is the dish
efficiency, and S is the mean distance between the earth and Europa in meters
(7.78x1011m). For this system a dish diameter of 2m was chosen, an X-band transmitting
frequency of 8.4GHz is also used. A transmission rate of 100kbps will be adequate for
our misssion with a signal-to-noise ratio of nine. A dish efficiency of 65% was used
since the dish is designed as a Cassegrain Reflector. A general schematic of a Cassegrain
reflector antenna can be seen in Figure 4.10.
Parabolic Main
Hyperbolic sub-
Reflector
reflector
Symmetry Line
Feed
Figure 4.10: Cassegrain Reflector
Equation 4.3 was solved for P, using the listed parameters. The required transmitter
power is 16.77 dB. Assuming transmitter power efficiency of 40%, the communications
system will require 53 W of electrical energy. Similarly, a low gain antenna, mounted on
32
the dish reflector, is used for low speed communications. This transmits at an S-band
frequency of 2 GHz. Data transfer rates of only a few hundred bps can be expected.
Aluminum Shell
Aluminum
Aluminum Shell
Honeycomb
Figure 4.11: Aluminum Honeycomb Composite
The dish itself is made of an aluminum honeycomb, see Figure 4.11. The interior of
the dish is a fine mesh of aluminum honeycomb weighing 12.5 kg. On either side, there
is a thin sheet of aluminum weighing less than a kilogram each. A support ring structure
mounts this system to the main structure and motors, see Figure 4.12. Including the ring
support and the reflector apparatus, the dish will have a weight of 18 kg. The power
electronics are located in the same instrumentation box as the computer and other
instruments. The power electronics are a redundant set of amplifiers, filters, waveguides
and processors. The weight of the power electronics will be approximately 17 kg. This
gives a gross weight for the communications system of 35 kg.
33
Support Ring
High Gain Antenna
Positioning Motors
Graphite-Epoxy tubes
Figure 4.12: Support for the antenna
4.3.5 Guidance, Navigation & Control
The purpose of GN&C of the spacecraft is to stabilize and control its attitude
(orientation) in space. The position and velocity information is obtained by using the
two-way Doppler derived from radio tracking and the Deep Space Network (DSN)
antennas. Attitude control is a necessary part of the spacecraft’s mission for many
reasons. GN&C detects deviations from the desired attitude and restores correct attitude
of the spacecraft. Another function of GN&C is to allow for accurate pointing of the
high gain antenna to earth, as well as, exact pointing of the instruments during imaging
and scientific observations.
Due to the degree of precision needed for our scientific studies and landing, a fairly
robust attitude control system is required. Since the different phases of the mission
demand very distinct pointing requirements, a spin-stabilized spacecraft is less desirable
than a 3-axis stabilized spacecraft. In addition, spin-stabilization is difficult to implement
in a spacecraft that must be landed. For this reason, a three-axis stabilization was chosen.
A tradeoff study was done between different GN&C devices. Included in this study
were control moment gyros (CMG), momentum wheels, ion thrusters and attitude
34
thrusters. Moment wheels and CMG’s were ruled out because of the duration of the
mission. Since they are mechanic devices they are prone to failure over such a long
operation period. Ion thrusters were also ruled out. While these thrusters show promise
in future space applications, their continuous use over such a long-term mission will not
have been proven before the projected launch date of the probe. After all of these
considerations, attitude thrusters were chosen as our RCS.
A description of how guidance, navigation and control is brought together is
described below, as well as shown schematically in Figure 4.13. Knowledge on position
and acceleration of the spacecraft is provided by the sensors and processed by the
computer algorithms to determine the required commands to actuate torque of the
spacecraft. Our spacecraft will have a highly autonomous GN&C system to eliminate
communication problems caused by long time lag and will reduce demands on the DSN.
Star Tracker
IMU
Attitude
Attitude
Determination
Control
Main
Engines
Sun Sensor
RCS
Thrusters
Flight Computer
Acceleromete
r
D
U
O
P
W
L
N
I
L
N
I
K
High Level
Commands
N
Maneuver
K
Verification
Figure 4.13: GN&C System Block Diagram
35
One of the sensors chosen for Zeus was a space-qualified, miniature, autonomous star
tracker (AST). It is basically a “star field in, attitude out” device capable of determining
its attitude without requiring any apriori attitude knowledge. In addition to this attitude
acquisition capability, an AST can perform attitude updates autonomously and is able to
provide its attitude “continuously” while tracking a star field. The sensor, named
Cassiopeia, is a device comprised of a baffle, refractive radiation hard optics, CCD,
camera electronics in radiation hard application-specific integrated circuits, power
supply, flight computer and hardware, and housing. The following table lists the
specifics of the Cassiopeia AST.
Table 4.6: Autonomous star tracker characteristics
Characteristics
Field of View (FOV), deg
Sensitivity, mv
Number of stars tracked
10 x 10
6.8
Up to 20
Accuracy of attitude output (arcsec, 1)
Pitch and Yaw
Roll
Update Rate, Hz
Power, W
Weight, kg
Operating temperature, °C
4
14
10
7
1.4
-30 to +50
A second sensor chosen was the Fine Sun Sensor (Figure 4.14 from the Swedish
Space Corporation. It is intended for 3-axis stabilized applications and has the following
specifications:
Table 4.7: Fine Sun Sensor
Characteristics
Field of View, deg
Accuracy, deg
Power Consumption, W
Weight, kg
Rectangular 35 x 35
Better than 1° in entire FOV
<0.35 W per FSS unit
<0.25 kg per FSS unit
36
Below is a picture of the actual sun sensor used on Zeus:
Figure 4.14: Swedish Space Corporation Fine Sun Sensor
An Analog Devices/ ADXL05 accelerometer is used to augment the IMU. The
accelerometer is oriented along the nominal line of action of the main engines. The
following table (Table 4.8) lists the specifications for the accelerometer.
Table 4.8: Accelerometer
Characteristics
Weight, kg
Power, W
Size, cm
0.005
0.35
Diameter = 0.94, Length = 2.4
A Litton LN-200 Fiber Optic Inertial Measurement Unit (IMU) was also selected for
the mission. This is a 3-axis stabilized micro-electric mechanical (MEM) system that
utilizes 3 MEM gyros. See Figure 4.15 for a picture of this. Table 4.9 contains the
specifications for this instrument.
37
Table 4.9: Inertial Measurement Unit
Characteristic
Weight, kg
Size, cm
Power, W
Operating Temp, °C
Operating Range/ Angular acceleration
0.7
Diameter = 8.9 cm, Height = 8.5 cm
10
-54 to +85
1000 °/sec
Figure 4.15: Inertial Measurement Unit from Litton
The reaction control system (RCS) is the set of actuators, which are used to control
the spacecraft orientation. Several different types of actuators exist for controlling a
spacecraft. One type of control system is the reaction wheel. The principle behind the
reaction wheel is that the total angular momentum of the spacecraft must be conserved.
By changing the spin of the wheel, the angular momentum of the wheel changes. Since
this is an internal component, the angular momentum of the spacecraft will also change.
The same thing can be done by changing the orientation of the wheels, as in a Contol
Moment Gyro (CMG). It is possible to have a three axis stabilized system with these
reaction wheels. Since external torques have to be introduced to de-spin the wheels
periodically, a set of thrusters will still be required for our spacecraft. To reduce mass,
complexity, improve reliability, and simplify our attitude control system, we will only
use thrusters.
In designing this system, it is important to determine the required applied torques and
the amount of propellant that will be used. A spacecraft mass of 1295 kg was assumed,
38
and it was further assumed that this mass was evenly distributed throughout a cylinder.
This provides principle moments of inertia of: Ixx=Iyy=1367.7 kg*m2, and Izz=1420.3
kg*m2. A maximum angular acceleration of 0.006 rad/s2 was arbitrarily assumed. With
these numbers, the maximum required torque is 9.225 N·m.
The thrusters are assumed to be 1.5 m from the center of gravity, which means that a
6.15 N thruster is needed to attain the required angular acceleration. Such a thrust can
easily be produced from a monopropellant hydrazine thruster. The hydrazine thruster of
this thrust class has a maximum Isp of 230 s, a mass of 0.3 kg, and is approximately 3.2
cm in diameter and 20.3 cm in length (Marquardt) (see Figure 4.16). A monopropellant
thruster was chosen over a bipropellant thruster to reduce the complexity of the ACS
system.
Figure 4.16: A typical ~5N thruster from Marquardt
To determine the required amount of propellant for this system, it is necessary to
estimate the burn times and thrusts for the duration of the mission. This will give us the
total impulse of the system. If the total impulse is known then the propellant mass will
simply be:
mp 
Total Impulse
I sp g eath
A list of assumed burns is listed in Table 4.10:
39
(4.6)
Table 4.10: List of RCS Thruster Burns
# Burns
2000
500,000
500
10,000
Duration(sec)
1
0.01
3
.1
Thrust(N)
9
1.6
9
4.5
With these assumed burns, there is a total impulse of 44000 N·s. Because the Isp of a
thruster is approximately 110 seconds for short pulses, the above equation was applied
separately to the small and large impulse burns. The results were then added. Applying
Equation 4.6 and adding an extra 15% for contingency, there will be a propellant mass of
29.3 kg. To take into account the fact that about 3% will be trapped in the tank,
additional propellant will be included. This will give us a propellant mass of 30.25 kg.
The entire quantity of hydrazine will be included in one central tank. This tank will
be divided into two sections via a diaphragm which separates the propellant from the high
pressure helium. The high pressure helium will force the hydrazine out of the tank
whenever the valve is opened as shown in Figure 4.17. The minimum pressure which the
tank will attain is 24 bar. This pressure is high enough to ensure that the engines are
getting propellant at an adequate pressure. To achieve this pressure at the end of mission,
an initial pressure of 96 bar is assumed with an initial blowdown ratio of four.
Heliu
m
Hydrazin
e
Valve
Thruster
Figure 4.17: Schematic of RCS propellant suppy
40
The volume of propellant is 0.03 m3. The volume of helium will be approximately
one-third the volume of fuel, or 0.01 m3. This makes the total volume of the tank 0.04
m3. If an additional 1% is assumed for the diaphragm, then the diameter of the tank will
be 0.4257 m.
The wall thickness is calculated using Equation 4.2. The tank is assumed to be made
of graphite epoxy composite, with an allowable tensile stress of 900 MPa. The
gravitation acceleration used for these calculations is the launch acceleration of 6g’s.
With these numbers, a tank thickness of 1 mm will be required.
4.3.6 Lander Computer
The lander computer is a doubly redundant, radiation hardened computer that will
control each aspect of the mission. This computer is responsible for managing all
communications between Earth and the lander. The lander computer is also responsible
for maintaining communications with the ice/water probe for the ground station. The
computer has 64MB of memory, which is continually error checked. This checking
constantly corrects for single bit errors that could arise due to radiation exposure. The
memory is used to store currently running programs and recently acquired data. The
computer also has 2GB of solid state storage. A diagram of the computer control system
can be found in Figure 4.18.
41
High Gain Antenna
Steering
Power Management
Attitude Sensors
ACS Actuator Control
Serial Bus
Computer #1
Computer #2
Solid State Storage
Internal Data Bus
Main Computer
High Speed Data Bus
Science Instruments
Compression ASIC
Telemetry
Figure 4.18: Flowchart of Computer Control System
The computer is actually two identical computers with 32 MB of memory each.
These computers are capable of functioning independently or in parallel. The entire
mission can be handled by one of the computers, however a division of tasks between the
two computers makes for a seamless operation. To deal with read/write issues between
the computer memory and the solid state recorder, each of the components manages
communications between each other on-board memory management units (MMU). Each
MMU controls the read/write privileges for the data storage on each of the systems. In
order for the memory to be accessed, the respective computer requests the information
from the MMU, which in turn does error checking and access privilege verification. The
operating system also checks to make sure that there are no conflicting memory or
storage uses. These systems serve to reduce the possibility of having a system crash due
to data corruption and the MMU’s further alleviate some processing burden from the
CPU. The entire computer system is capable of restarting in the event of a catastrophic
system crash.
42
The main computer is connected to the other components either through the serial bus
or the high speed data bus. There are two serial buses which are used to control all
instrumentation on the probe. The serial bus activates and sends signals to all the
actuators and science payloads on board. A 1 Gigabit per second data bus is used to pass
around all science data from the experiments. A direct link between the compression
ASIC and the experiments reduces some of the bus traffic. This bus is also used to send
and receive signals from both the low and high gain antennae.
4.3.7 Orbital Mechanics
The trajectory of the mission is an important part of the mission’s design. An
interplanetary transfer to Jupiter using the same Venus-Earth assists as Galileo was
assumed due to the complex nature of this analysis procedure. The V necessary for the
Jupiter orbit insertion, gravity assists in the Jovian system and insertion into a 20 km orbit
around Europa is approximately 2.5 km/sec.
The V calculations for the 20 km orbit around Europa were performed assuming the
spacecraft starts at a 20 km circular and lands on Europa’s surface at via a Hohmann
transfer (see Figure 4.19). A non-equatorial orbit was chosen based on the scientific data
return since equatorial orbits have poor planetary coverage characteristics [Desai,1993].
A detailed calculation is shown in Appendix 11.4. From these calculations a total V of
1.44 km/s will be required
V
1
E
V
2
Figure 4.19: Transfer from 20 km circular orbit to Europa’s surface
1569 km
43
A margin of error of 10% was accounted for as was implemented in V studies for
research done by NASA’s Vehicle Analysis Branch on a Mars excursion vehicle as part
of a manned Mars mission analysis [Desai, 1993] resulting in a total V is 1.58 km/s.
Combining V’s for Jupiter orbit insertion, gravity assists, 20 km orbit insertion and
landing on Europa’s surface the total v is ~ 4.1 km/s.
4.3.8 Orbiter Thermal Management
A thermal design must be required to shunt and radiate the heat generated by the ice
water probe, located in the center of the orbiter configuration, to space. The design
chosen allows the probe to radiate its heat onto a long aluminum tube surrounding the
heat-generating end of the probe, namely the first 48.6 cm. The space between the probe
and the aluminum wall is 2mm. This allows approximately all the heat to be radiated
onto the wall. The wall is painted with a black epoxy, to allow maximum heat transfer.
A jacket filled with Dowtherm fluid surrounds the aluminum wall. Dowtherm was
chosen mainly due to its high boiling temperature and medium density. Two heat pipes
are inserted into the jacket and transfer heat to radiators located outside of the orbiter.
The heat pipes selected use water as the working fluid. The radiator is a roll out fin
design as presented in the AIAA 86-1323 paper. A section view can be seen in Figure
4.20.
44
Figure 4.20: Cross section of cooling system
The radiator contains an internal vapor space and liquid condensate return channels
which are activated by the capillary action and the roll up action of the spring loaded thin
walled vapor chamber segments. The heat transport is accomplished by the evaporation
and condensation processes while the deployment and roll-up operations rely on the
working fluid pressure and the spring stiffness. This design was chosen due to its small
size and compactness, especially during storage in the booster. The characteristics of the
radiator fins chosen (there are two) are presented in Table 4.11.
Table 4.11: Radiator Properties
Length
Width
Mass
Emissivity
Fin Efficiency
Radiating Temperature near Jupiter
Radiating Temperature near Venus
45
1 meter
65.2 cm
3.586 kg
0.82
0.8
380 K
472 K
The radiator fin has an initial operating temperature of 373 K. The dimensions were
chosen from the equation:

Q  A   Tw 4  Ts 4

(4.7)
where Q is the heat radiated, η is the fin efficiency, ε the fin emissivity, and σ the StefanBoltzmann constant.
The constraint on the design was that the temperature radiates the heat above the
initial operating temperature. The heat to be radiated by each fin was designed to be
around 1000 W. A sink temperature of TS of 50 K was used to include view factors from
other parts of the craft. It was assumed that a flight trajectory similar to Galileo would be
used to get the Zeus spacecraft over to Europa. This would involve a gravity assist fly-by
past Venus and should be taken into consideration.
The value of direct solar flux next to Venus is:
GS ,V  GS R 2 
1358 W m   2598 W m
2
0.7232
2
(4.8)
where GS is the solar flux at Earth and R is the distance of Venus from the Sun in AU.
The solar irradiation on the fins at a near Venus orbit is:
Q Sun  G S,V A fin  cos   1467 W
(4.9)
where QSun is the solar irradiation,  is the solar absorptivity, Afin is the area of one side
of one fin, and  is the angle of incidence of the incoming solar radiation. For the sake of
argument, the solar absorptivity was assumed to be 1 (same as black epoxy paints). The
design condition of maximum solar irradiation onto the fins occurs when the angle of
incidence is 60. Therefore the fin radiating temperature may be estimated by noting that
by conservation of energy, the heat radiated out must equal the heat input to the fin.
Solving the equation

QSun  Qheat pipe  A   Twmax  Ts 4
4

(4.10)
gives the maximum expected temperature that the fins are to experience. Qheat pipe is the
heat input by the heat pipes, Twmax is the maximum expected temperature. Irradiation
from Venus and solar reflectivity from Venus’s surface was not considered in the design.
46
This was because it was believed that the Venus fly-by would be so quick, that its effect
would be small. The result is a fin radiating temperature of 472 K.
The heat pipe design chosen was a double wall artery using copper as the wicking
material and water as the operating fluid. This design was chosen because of its
operating temperature ranges, namely from 380 K to 480 K (107 C to 207 C). The
estimated design characteristics are listed below:
Table 4.12: Heat Pipe Characteristics[Faghri, 1995]
Length
Diameter
Mass
Length of the Evaporator Section
Length of the Condensing Section
0.732 m
1.5 cm
0.088 kg
0.12 m
0.12 m
The evaporator section of the heat pipes will be inserted into the heating jacket.
The condensing section will be attached to the evaporator section of the roll out fin. A
jacket of Dowtherm will surround the condensing section of the heat pipe and the
evaporator section of the radiator.
The cooling jacket surrounding the condensing end of the heat pipe will be 1 cm
thick. It will also surround the evaporator section by 1 cm. This allows the heat to be
transferred from the cold end of the pipe to the hot end of the radiator. To allow
operation at the given temperature ranges liquid Dowtherm will be used (Dowtherm has a
boiling point of 531 K, and a melting point of 285 K). The density of Dowtherm is 953
Kg/m3. A 1 mm thick aluminum shell will surround the jacket. The properties of the
heat pipe condensing end jackets are given below:
Table 4.13:Characteristics of Heat Pipe Condensing End
Dimensions surrounding heat pipe condensing
end
Dimensions surrounding radiator evaporator
end
Volume of system
Mass of system (with aluminum housing shell)
47
Length = 0.12 m; Di = 1.5 cm, Do =
3.5 cm
Length = 0.1m; Height = 0.652m;
Widthin = 3cm; Widthout = 5cm
0.00140 m3
1.42 Kg
The heating jacket surrounding the ring absorbing the heat from the probe also uses
Dowtherm as its operating fluid. It will be 2 cm thick. It will surround the evaporator
end of the heat pipe by 1 cm. An aluminum shell wall thickness of 1 mm will surround
the Dowtherm. The properties and characteristics of the jacket are given below:
Table 4.14: Characteristics of Jacket Surrounding the Probe
Dimensions surrounding the ring
Dimensions surrounding the heat pipe
evaporator end
Volume of system
Mass of system (includes aluminum housing
shell)
Height = 0.486 m; Di = 7.5 cm; Do =
11.5 cm
Length = 0.105 m; Di = 1.5 cm; Do
= 3.5 cm
0.00298 m3
3.34 Kg
The aluminum shell will have a low emissivity, which will allow for minimal heat
loss due to radiation. However, there will still be some heat loss across the pipes and
heat jackets. Furthermore the efficiency of the transfer of heat by convection by the
Dowtherm in the jackets was not determined in this project. Therefore a temperature
gradient of 15 K is assumed to exist between the mean radiating temperature of the
radiating fins, and the inner wall temperature of the ring jacket surrounding the probe. A
more detailed analysis is required, where the convection of Dowtherm is considered in a
finite element/difference method.
5.0 Probe
5.1 Overall Probe Description
The most challenging phase of this mission takes place beneath the surface of Europa.
The RFP specifically requires a subsurface probe which can reach the point where the ice
sheets that cover Europa become water--the so called “ice/water boundary”. Current
observations have been limited to flybys of Galileo, which can only study surface
features. With current techniques, it is possible to make some predictions about the
features beneath the surface, however, a direct study needs to be undertaken to explore
Europa’s subsurface. Using an ice/water probe, direct measurements of these features
48
will be possible. These measurements will allow researchers to understand the evolution
of this moon and hence other moons in the solar system and the Earth itself.
Figure 5.1: interior view of ice/water probe
The ice/water probe (shown in Figure 5.1), henceforth called “the probe”, has the
main task of carrying out all of necessary experiments required to determine the
subsurface conditions of Europa. The probe has an overall length of 128cm and an
internal diameter of 14cm. The total mass of the probe is approximately 52kg. A
geological profile will be conducted to determine level of ice fracture, and thus the
amount of movement of the plates. The probe will also perform detailed pressure and
49
temperature measurements. This data can be used to give scientists a better
understanding of the inner regions of this moon.
From fracture, temperature, and pressure profiles, scientists can have a better
understanding of what is actually causing the formation of the current landscape. Present
theories range from internal heating due to tidal forces, or the actual tearing forces of
Jupiter pulling the plates apart. The level of mobility of the ice is a function of its
temperature and pressure, and this study will provide an insight into the importance of
each of these effects.
Besides these basic measurements, a detailed chemical analysis will be performed.
The chemical analysis will be performed periodically during the entire decent phase, and
at the ice/water boundary. The probe will look for all types of chemicals from simple
monatomic ions to amino acids. This will provide scientists with a time history of the
impacts from various planetary objects. The study of amino acids will also be used to
determine if life does exist on Europa, or had existed some time in the past.
The probe is divided into sections, based on their functions. These divisions have
been introduced to improve fabrication and testing of each component, as well as
allowing for the isolation of essential systems from one another. The four main sections
are: power/heat generation, instrumentation, Probe Arresting Sub System (PASS), and
communications. Detailed descriptions of each of these systems are presented below.
5.2 Science Operations
The instrumentation segment of the probe contains all of the critical electronics and
instruments which are required to complete the mission (see Figure 5.2). These systems
are located in the center of the probe, and are designed to
50
Transducer/Thermcouples/Optical viewing
window
High Voltage
Camera
Converter
Computer &
Capillary Lamp
Storage Board
Capillary Electrophoresis
Buffer Tanks
Capillary Plate
Figure 5.2: Illustration of the instrumentation section of the probe
be autonomous. A series of standard pressure transducers and thermocouples line the
wall of the instrumentation section, to provide pressure and temperature data for logging.
A camera is also installed in the probe, which will be used when the ice/water boundary
is reached. The chemical analysis will be performed with a capillary electrophoresis
system. Measurements from each of these systems will be recorded at regular intervals
during the descent, but they will be used at high sampling rates once the probe is in
water.
5.2.1 Chemical Analysis
The most important part of the probe’s mission is to carry out the chemical
analysis of the Europan water and ice. The materials that will be studied vary from
simple ions (H+, SO4-,etc) to metals and organic molecules. The study of the inorganic
molecules could be done with a gas chromatography or similar analysis, which has been
used on previous missions. The search for organic molecules, and specifically molecules
that indicate life, requires a different kind of analysis.
Using capillary electrophoresis, such analysis of all these chemicals can be
achieved with only one apparatus. Capillary electrophoresis is the process of separating
51
charged molecules based on their movement through a fluid under the influence of an
applied electric field (Weinberger, 1993). This is done by exposing the sample to a
uniform electric field. Various species in the solution will travel up the capillary at
different rates. This variance in rates allows the scientists to identify the chemical
breakdown of the solution. Such analysis is routinely used on Earth for studying all types
of chemicals.
The system used in the probe consists of a small capillary attached to two reservoirs
(see Figure 5.3). The total capillary volume amounts to about 10 pL, and each reservoir
has a capacity of 1 mL. All of these components can be etched onto a single plate of
glass. This makes the capillary system both small and light.
Voltage Source
Electrode
Electrode
Capillary
Sample
Injection Port
Detector
Reservoirs
Figure 5.3: A schematic of the capillary electrophoresis sytem
Both ends of the capillary are connected to small reservoirs containing buffer
solution and electrodes which create the electric field across the capillary length. This
power supply is capable of providing up to 0.3 mA at 30 kV to the capillary. The high
voltage is required to get a clear separation. The number of plates, a measure of the
resolution of a capillary system, for this system is 20 plates/V. This resolution relates the
output to the width of a given chemical’s signature(a larger number is better).To
maximize the resolution, a high voltage, on the order of kilovolts, has to be used. The
52
polarity of the field can be changed “on the fly,” at the power supply. For good results,
Joule heating must be kept to a minimum. Therefore, at the high voltages require,
~10kV, the current must be kept under 1A.
The sample is drawn in via a set of small pumps. The sample goes through two
levels of filtration before being injected into the capillary. The first filter removes large
particles, while the second filter removes small particles. These particles will interfere
with the analysis since their diameter may be larger than the diameter of the capillary.
Buffer solutions are added to their respective reservoirs in a similar fashion. The buffer
solution aids in the separation of the particles by changing the overall charge of the
sample. A set of three different buffer solutions will be stored in the system and will be
used for various analyses. After the analysis, the reservoirs and capillary will be flushed
with a cleaning solution. The fluid from the capillaries is removed via another set of
syringe pumps and the cleaning solution is added through the same system as the sample.
Detection levels vary for different setups. It is possible to study the contents of
the stream via optical and electrical methods. To easily adapt each experiment to the
given chemical of study, an optical technique is employed. For the chemicals being
studied, visual and ultraviolet (UV) absorption spectrums are prime indicators. For this
reason, a set of visible light and UV lamps, each with an output power of 5 W, and a
CCD are attached to the capillary. The lighting system which will be used will vary for
the given experiment. The arrangement inside the probe is shown in Figure 5.4. The
voltage output from the CCD as a function of time for an experiment are stored in the
main computer and compressed before transmission.
53
Capillary High Voltage
Electronics
UV & Visual Lamps
CCD Detector
Capillary Plate
Buffer Storage Tanks
Sample Intake Pump
Buffer Syringe Pump
Figure 5.4: Arrangement of CE system inside probe
Each lamp will require 5 W of electric power, and the power supply for the capillary
will typically require about 5 W. This puts the total power requirement at 10 W.
The minimum molarity for detection varies from experiment to experiment. It is
difficult to determine the minimum levels of detection without having an actual physical
system. For this reason, a reasonable guess was made, based on previous electrophoresis
experiments. Based on previous experiments, tests for organic molecules, such as sugars
and amino acids, have minimum levels of detection on the order of 10 attomoles (aM).
For small ion analysis, the minimum levels are not as impressive, with a minimum
detection level of 1 nM. These are assumed to be the order of magnitude for the
minimum levels of detection for our system.
Besides testing at the ice/water interface, chemical tests will be performed at various
locations during the probe’s journey through the ice. By studying the chemical
composition of the ice at varying levels, it will be possible to see the history of impacts
54
from solar debris. The levels of various metals and other compounds will tell scientists
something about the recent past of Europa.
5.2.1.1 Life Detection
Life detection will be determined by a chemical analysis of the oceans and ice plates
of Europa. To test for life it is best to have an understanding of what is being looked for.
Unfortunately, there is no way of knowing what life will look like, even at the molecular
level. The realm of science requires that we make initial guesses, and then hope that
these guesses provide us the proper tools for study.
One way of testing for life would be to take in a sample, feed it with a nutrient rich
solution, and then check for changes in turbidity, or cloudiness. If the solution turbidity
increases over time, then it is likely that an organism is reproducing and functioning
inside of this sample. Without any knowledge of the life forms that may exist on Europa,
this experiment has too many variables. Even on Earth, almost 90% of all organisms are
unknown. In some cases, what is a rich environment for one species is literally toxic for
others. For example, some forms of bacteria are poisoned by oxygen rich environments.
The added complexity of guessing the environment on Europa makes designing a valid
experiment along these lines infeasible.
An experiment based on DNA replication, using polymerase chain reactions (PCR)
would be a better approach. PCR is used repeatedly in criminal analysis. By introducing
a sample with DNA into a specific solution of amino acids, it is possible to make billions
of copies of DNA from a single DNA fragment inside the solution. By testing for DNA it
would be possible to determine if DNA based life existed. Again, only a single DNA
strand would be required for this experiment to work. The problem arises in how to
culture the DNA. When the general structure of the DNA is known, it is easy to
determine what the necessary solution should be. Since nothing is known about this
system, there is no way of deciding what the exact solution should be. This means that
the experiment will have to contain a large number of possible amino acid buffers to
thoroughly test for DNA based life. Due to space constraints this in unacceptable.
One final way of testing for life is to look for the presence of amino acids. Using the
Earth as a test case, it appears that all life requires amino acids. Amino acids are the
55
building blocks of all life. If we assume that life on Europa is similar in structure to life
on this planet, then it too will need an abundance of amino acids. If a test for amino acids
comes up positive, then there is a good chance that life existed at one time on Europa.
The problem with this test is that amino acids can form spontaneously without there
being any life present. This has been shown to occur in the laboratory as well as in
nature.
A way of solidifying the test results is to look at the distribution of the amino acids.
Amino acids are grouped into two different categories, L and D. These two different
classifications depend on their structure. It is a commonly known fact in organic
chemistry that when amino acids form spontaneously, they create nearly equal
percentages of L and D amino acids. When life forms enter into the equation, they take a
certain preference between these two structures. For example, all life on Earth uses L
amino acids. These life forms therefore begin to synthesize the amino acid forms that
they require. This throws the balance of amino acids far into one of the two structures.
Since life on Earth requires L amino acids, the overwhelming majority of amino acids in
a sample taken on Earth will be L amino acids.
A study of the balance between the L and D amino acids is required to guarantee that
life is being observed. Each different type of amino acid will separate out distinctly. If
amino acids are present it will be possible to see the relative concentrations of each. If an
imbalance towards either amino acid structure is found, then it is sufficient to assume that
the environment, at one time, had life.
5.2.2 Probe Photography
Once the ice/water boundary is reached, photographs of the region will be taken.
An optical port in the ring which houses the pressure transducers and thermocouples will
provide a clear optical viewing port for the camera. The window will be made of
plexiglas designed to tolerate the pressures at the ice/water boundary. Directly behind the
window will be a lens system which will allow the computer to focus and zoom in on
56
various objects. Images formed in the focal plane of the lens will be detected by a
1024x1024 pixel CCD and sent to the main computer.
To illuminate the surrounding area, halogen lamps are located near the optical port.
These lights are connected to small capacitors and can be operated in a continuous 10W
power mode or in a short burst (flash)mode providing intense light for a few millisecond
exposures. The switching from standard mode to continuous mode will occur if a
command is received from the ground station.
The camera itself is a 1024x1024x12 bit intensified charged couple device (CCD)
camera. An intensified camera amplifies the light received by the CCD by having an
intensifier plate in front of it. The light strikes the intensifier plate, which has a certain
threshold level. If the light level is high enough, then the intensifier plate sends out a
burst of photons onto the CCD. With the intensifier plate, gains of up to 1000 times can
be attained. This is used to reduce the necessary illumination requirements. By using the
intensifier, the size of the light source can be reduced drastically. To ensure that the
correct gain is used for a given picture, the computer does a statistical analysis and
adjusts for any discrepancies.
Photography will be delayed until the probe has successfully reached the ice/water
boundary. Once through the boundary, the probe will take pictures at regular intervals.
The initial series of pictures will be taken once every few minutes. The picture
scheduling time can be varied via programs sent to the probe from the ground control
stations. The probe’s computer, assuming an 8:1 compression ratio will provide memory
storage for fifty-pictures. The probe uses the same compression ASIC as is used on the
lander or data compression. Appendix 11.2 describes this compression scheme.
5.2.3 Pressure and Temperature Transducers
Pressure and temperature measurements provide useful information about the subsurface Europan features. By measuring temperature at regular intervals, scientists will
be able to determine the amount of internal heating which occurs on Europa, as well as
study other types of energy dissipation. The pressure measurements will create a profile
which will allow scientists to better understand subsurface ice-plate motions.
57
A special ring will be manufactured, to allow for easy mounting of these instruments.
The thermocouples are going to be etched directly onto the surface of the ring. This
provides the probe with a high number of thermocouples without compromising
structural integrity.
These thermocouples will completely cover a thin band on the
upper portion of the ring. These signals will be integrated and amplified via an internal
amplifier. To protect the thermocouples from damage due to rubbing or scraping, a thin
layer of titanium will be applied over top of the thermocouples. This will prevent the
thermocouples from actually coming into contact with any objects or fluids, while still
providing a reasonably accurate temperature.
Pressure will be measured via a series of miniature pressure transducers. These
transducers will have a maximum pressure rating of 10 MPa, which is larger than the
anticipated maximum pressure the probe will experience.
5.3 Probe Subsystems
5.3.1 Probe Power
The probe contains a stack of nine GPHS blocks which generate the heat to melt
through the ice. The heat generated by the blocks is transferred to the wall mainly by
conduction from the block on the bottom. The heat is then transferred to the ring of the
heat jacket by radiation. On the top of the ninth block, a copper disc is mounted holding
seven 10 Watt AMTEC cells. The ninth block and the bottom of the copper disk are
thermally isolated from the rest of the probe by MLI. Furthermore the MLI effectively
allows no thermal radiation transfer between the top block and copper plate and the rest
of the GPHS blocks and probe wall. Finally MLI is inserted between the top block and
the rest of the blocks to allow for no conduction between them. Therefor the heat
generated by the ninth block must traverse the copper plate and the AMTEC cylinders on
top of them. On top of the AMTEC cylinders is another copper plate. This conducts the
heat to its edges, where a ring of interstitial material slows down the heat transfer by a
specified amount determining the hot and cold ends of the AMTEC cells. The interstitial
material here is pyroceram. The probe computer (the next section above the AMTEC
58
cells) will be thermally isolated from the wall by MLI. The heat generated by the ninth
block will be radiated from the probe wall to the heat jacket surrounding it.
As mentioned in the Orbiter Thermal Control, the hot end design margin for thermal
analysis occurs at a near Venus flight. The analysis in that section resulted in a hot wall
temperature of the jacket to be 472 K (199˚ C). The probe wall is thermally separated
into 3 sections. The first wall is 48.6 cm long. It extends up to the copper plate, the top
of the ninth GPHS brick. Here surrounding the copper plate is a ring of high temperature
aerogel. The ring of aerogel extends out to the surface of the probe and thermally
separates the walls of the two sections. The middle section of the wall extends from the
top of the ninth brick (48.6 cm) to the top of the computer housing section (92.4 cm).
Here another ring of aerogel separates the middle section from the top section of wall.
The temperatures of the bottom two sections of titanium shell must be determined for a
near Venus flight. By conservation of energy:
Q gen,1  Qgen, 2  Qrad ,1 at the bottom  Qrad ,1 around the shell  Q2,rad
(5.1)
where Qgen,1 is heat generated by the bottom eight GPHS bricks, namely 2000W, Qgen,2 is
the heat generated by the top single GPHS brick, Qrad, 1 at the bottom is the heat radiated out
by the very bottom of the probe, Qrad,1 around the shell is the heat radiated by the 1st section’s
circumferential shell, and Q2,rad is the heat radiated out be the second shell. The third top
most shell is effectively considered isolated from the other two.
As will be determined below, the AMTEC cells will not be working on the Venus flyby as the temperature of the cold base of the AMTEC cells will be above the operating
limit. Thus Qgen,2 is 250 W. Substituting the relations into Equation 5.1 gives:
2250 W m    r T
2
2
1
1
4

 TS 
4

2rh1 T1  T j
4
4
1 1  1  j  1
  2rh T
2
2
4
 Tj
4

1 2 1  j 1
(5.2)
where σ is Stefan-Boltzmann’s constant; ε1, ε2, and εj are the emissivities of the bottom
shell, middle shell, and heat jacket respectively; T1, T2, and Tj, are the temperatures of the
first, second, and jacket surfaces; TS is the space sink temperature for the bottom,
aussumed to be 125 K; r is the radius of the probe (7.5 cm), and h1 and h2 are the heights
of the bottom and middle shells respectively. By applying black epoxy paint along all the
surfaces, the emissivity of the surfaces can be greatly increased. The emissivities of all
three surfaces will be 0.9. At the near Venus situation, the inner wall temperature of the
59
jacket was found to be (from the Orbiter Thermal Control section) 487 K. Equation 5.2
may be solved by solving separately


Qgen,1  1r T1  TS 
2
4
Qgen, 2 
4

2rh1 T1  T j
4
4

(5.3)
1 1  1  j  1

2rh2 T2  T j
4
4

(5.4)
1  2 1  j 1
The solution to these equations gives a bottom wall temperature of T1 = 689 K and
wall temperature of the middle section T2 = 535.7 K. Equation 5.3 shows that some of
the heat generated by the bottom shell gets radiated out into space. By back substituting
T1 = 689 K, and noticing that the sum of the terms Qrad,1 around the shell and Q2,rad is the rate
of heat transferred to the heat jacket, it is found that Qrad,bot = 203 W, the heat transferred
to the jacket is Qtran = 2047 W. The same wall temperatures can be found for the case of
solar input. Here (as will be shown below) the AMTEC cells will be operating, causing a
heat sink of 60 W. This results in Qgen,2 being only 190 W. Furthermore, the estimate of
inner temperature of the jacket was 395 K. A summary of the results is presented below
in Tables 5.1 & 5.2:
Table 5.1: Conditions with solar heating near Venus
Temperature of the inner wall of the jacket
Temperature of the bottom section
Temperature of the middle section
Heat radiated out the bottom
Heat transferred to the jacket
Heat to be carried by heat pipes
487 K
689 K
536 K
203 W
2047 W
1024 W
Table 5.2: Conditions with no solar heating
Temperature of the inner wall of the jacket
Temperature of the bottom section
Temperature of the middle section
Heat radiated out the bottom
Heat transferred to the jacket
Heat to be carried by heat pipes
60
395 K
665 K
477 K
176 W
2014 W
1007 W
It is assumed that all of the heat from the top most GPHS brick will go up across the 7
converters. In reality, some will be transferred across the MLI, and some across the high
temperature aerogel isolating the ninth brick from the other bricks and the aerogel
isolating the copper disk from the probe wall. Aerogel has a melting temperature of
1200˚ C, however its upper operating temperature limit is well below that due to
shrinkage occurring around 500 to 600˚ C. It will be shown that the operating
temperature ranges around the hot end of the cylinder disk are of the order of 950˚ C. It
is hoped that new developments in aerogel technology will allow operation up to its
melting point. If a super high operating temperature aerogel will not be available, then
some more traditional high temperature solid insulation will be needed, such as
diatomaceous silicate. However, these are not as good insulators as a super high
temperature aerogel is projected to be, and would require about four times greater
thickness to achieve the same thermal qualities. Assuming the thermal conductivity of
aerogel to be 0.017 W/(Km), and a temperature gradient between the ninth and eigth
block on the order of 700 K, a thickness of 2 cm aerogel is required to limit the exchange
to 5 W. This justifies the assumption that all heat will be transferred to the copper plate.
The super high aerogel surrounding the bottom plate will go up to the base of the cells,
5.9 cm from the center of the copper disk. The copper disk will be 3mm thick, to allow
for thermal conductance.
It was assumed that the copper plates would offer no thermal resistance to the heat
transfer and that the plates would have a uniform temperature. Therefore the only
resistance encountered in this simplified analysis is the resistance of the cells and the
thermal insulation at the top of the probe.
61
MLI
AMTEC cell
Copper conducting disk
MLI
Pyroceram
AMTEC Cells
Copper conducting disk
High temperature aerogel
High temperature
aerogel
Probe titanium wall
GPHS Blocks
Figure 5.5: Internal probe thermal layout
The resistance of a single cell is given as approximately 15 K/W. The seven cells are
in parallel and a thermal model of the system is given below.
62
R pyroceram
Q generated
Qout
Thot
Tcold
Q converted
Twall
Rcell
Figure 5.6: Thermal model of probe power generation
The equation relating the temperature of the hot end, Thot, and the cold end, Tcold,
is
Q gen  Thot  Tcold  R7 cells
(5.5)
where Qgen is the heat generated by the one GPHS brick, R7 cells is the resistance of seven
cells in parallel. The resistance of the seven cells in terms of a single cell Rcell is obtained
from the parallel resistance equation:
1
R7 cells

1
7

Ri Rcell
(5.6)
Equations 5.5 and 5.6 give the temperature change across the hot and cold ends of the
cells. A cell resistance of 15 K/W gives a total 7 cell parallel resistance of 2.1428 K/W.
A thermal input of 250 W gives a change in temperature of 535.7 K and a thermal input
of 222W (the output of one block after 15 years of decay) gives a change in hot and cold
temperatures of 475.7 K Equations 5.7, 5.8, and 5.9 relate the cold end temperature to
the wall temperature.
Qout  Tcold  Twall  Rwall  Q gen  Qconv
(5.7)
Rwall  L k ins  Ains 
(5.8)
Ains  2rt ins
(5.9)
63
where:
Twall andTcold are the wall and cold end temperatures
Qout is the heat transferred to the wall
Rwall is the resistance of the wall, which only includes the wall insulation
Qconv is the amount of heat energy converted to electrical by the converters
L is the width of the ring of insulation
kins is the thermal conductivity of the insulation
Ains is the circumferential area of the insulation
r is the radius of the insulation (equal to the radius of the probe, 7 cm)
tins is the thickness of the insulation
The operating temperatures of the cells require a cold end range of 150˚ - 400˚ C and
hot end temperature of 450˚ C and is limited by material limits. The cells chosen are 3.81
cm in diameter and 10.8 cm in length. There optimum operating temperatures are at a
cold end of 350˚ C and a hot end of 900˚ C. at this condition they operate at 25%
efficiency. For this design, it cannot be planned to have the cells operate at their
optimum efficiency, because they would overheat on the flight to Europa (remember the
wall temperature is estimated to 477 K, with a maximum if given direct solar irradiation
at a near Venus situation of 528 K). So it was desired to have the hot end of the cell be
less than 1000˚ C during its hottest moment. The resistance of the wall insulation was
desired to be such that it would allow the cold end to be above 150˚ C, but not let the hot
end to be above 1000˚ C. This results in requiring a thermal resistance of the insulation
ring at the cold end to be 1.0526 K/W by equations 5.5 and 5.7. Given that the thickness
of the insulation ring will be the same as the thickness of the copper plates, 0.5 cm, and
the insulation ring width can not be more than 1 cm, a thermal conductance on the order
of 3 W/(K M) is required. Pyroceram has a conductance of that order and operates at the
required temperature ranges. The properties of Pyroceram are given in Table 5.3
Table 5.3 : Properties of Pyroceram
Density
Thermal Conductivity
Width
64
2600 Kg/m3
4 W/(m K)
9mm
Finally the internal temperatures of the probe may be calculated. A relation between
the efficiencies of the cells and the operating temperatures is not available. Assuming
that the cells only operate at an efficiency of 80% at a hot end temperature of 657°C, the
electrical power output of the cells can be determined. The projected power output of the
cells is 44 W. The tables below summarize these findings.
Table 5.4: Internal temperatures with no direct solar flux
Heat generated
Wall temperature
Cold End Temperature
Hot End Temperature
Approximate Electrical Watts produced
250 W
476.5 K
403.5° C
939° C
62.5 W
Table 5.5:Worst case scenario; internal temperatures near Venus
Heat generated
Wall temperature
Cold End Temperature
Hot End Temperature
Approximate Electrical Watts produced
250 W
527.5 K
454.5° C
990° C
0W
Table 5.6: Internal temperatures at EOM
Heat generated
Wall temperature
Cold End Temperature
Hot End Temperature
Approximate Electrical Watts produced (assuming
an 80% efficiency of the cells)
222 W
265 K
181.5˚ C
657˚ C
44 W
A finite element thermal analysis will be required for more accurate predictions.
65
5.3.1.1 Probe Power Requirements
The power requirements for different modes are listed in Table 5.7.
Table 5.7: Electrical power requirements for various probe conditions
Mode
1
2
Description
Surface Communications
Science Data Acquisition
3
PASS Sequencing
4
Scientific Data Acquisition with
Continuous Lighting
Components
Modem, Computer
CE, Camera,
Pressure &
Temperature, Lamp
from capacitor
PASS Actuators,
Computer
CE, Camera,
Pressure &
Temperature,
Directly powered
lamp
Power
6.66W
16.55W
25W
26W
5.3.2 Probe Arresting Subsystem (PASS)
The Probe Arresting Sub-System (PASS) is a dual component system which allows
the probe to remain near the ice/water boundary. Once the probe has penetrated the
boundary, it is required to stay within the vicinity of the exit hole. Without a system to
anchor the probe near the ice/water boundary, communications would quickly be lost. If
an anchor system is not used, then an active control and propulsion system would be
necessary. Due to size and weight constraints such an option is not viable.
66
Acoustic Modem
Transducer
PASS I
PASS II
Figure 5.7: Side view of the upper part of the probe
The PASS consists of two components, PASS I and PASS II (see Figure 5.7). The
PASS I system is designed to anchor the probe in the ice several meters before the
ice/water interface. The PASS II system separates the probe at a junction point and
tethers the lower part of the probe to the anchored section.
The probe’s on-board computer, within about 30 m of the ice/water interface,
activates the event sequence for the anchoring and separation. The distance between the
interface and the probe is determined acoustically. The lander transmits a characteristic
acoustic pulse from its transducer. When the probe receives this pulse it begins listening
for the pulse to return after being “bounced” off the interface. Knowing the speed of
sound in ice and the delay time, the probe can accurately calculate the distance traveled
by the signal. If the distance is within 30 m, then the event sequence will be activated. A
diagram of this process can be found in Figure 5.8.
67
Receiving Signals
Filter and Parse
Send to Computer
If signal =
depth signal
Begin/End timing
If end of
timing
Calculate Distance
If Distance
> 30 m
Reset Timer
Activate PASS
Figure 5.8: PASS Activation Sequence
The first step in the event sequence is to anchor the probe. The PASS I system,
shown in Figure 5.9 contains two titanium blades which are mounted on rotating
actuators. These actuators use 20 W of power and are capable of producing 11.3 N-m of
torque. These blades have highly sharpened edges to allow them to be embedded as
deeply as possible. Once the rest of the probe drops away, the PASS I system will
become frozen inside of the ice as the liquid water refreezes around it. This refreezing
provide most of the support for the lower section, especially if poor penetration occurs.
68
Blades
Transducer Mount
Actuators
Actuator & tether
mounting plate
Figure 5.9: Side view of the PASS I section (isometric view inset)
Once the PASS I system is anchored in the ice, the PASS II system releases the
bottom portion of the probe. The two pieces of the probe are connected via a series of
explosive bolts, which disintegrate when a charge is put through them. The PASS II
system can be seen in Figure 5.10. Once the PASS II system is activated, these bolts will
have a current applied, and they will release the probe. The lower portion will continue
to melt through the ice and slowly drop away from the PASS I system. To tether the
systems together, a 50 m Spectra-1000 tether will be used. This cable not only provides
support, but also provides power and a communications link to the PASS I section—
which houses the acoustic modem transducer. The electricity is passed via a series of
copper cables. The signal from the hydrophone is sent down another series of copper
cables to the computer system. Copper was chosen over fiber-optic cables because they
69
are more durable and the data transmission rate is not high enough to be limited by the
use of copper cabling. The total weight for the entire PASS system is 27 kg.
Incendiary Bolts
Incendiary Bolts
Spectra 1000 Cable Spool
Hollow Core
Figure 5.10: Side view of the PASS II internal structure
5.3.3 Data Management
5.3.3.1 Computers
A central computer system is used to control each of the main parts of the probe. The
computer will be a 32-bit computer, with the capability of storing up to 64 MB of data.
This data will be stored in solid state memory recorders. There will be 16 MB of
memory available for temporary data storage and program storage. Besides the basic
board, a series of ASIC’s will be used to reduce the computation time from the main
processor. This will create a much smaller architecture. The main ASIC’s are for data
compression and communications.
The data compression ASIC is exactly like the ASIC on the lander. This ASIC will
implement a progressive JPEG compression algorithm to compress the image data. The
70
compression ratio can be varied from the ground; however, a default compression ratio of
8:1 will be used otherwise. Unlike the lander, the data from the probe is not suited to
post-processing. Therefore, the data collected by the pressure transducers and capillary
electrophoresis system will be compressed with a lossless compression algorithm. This
will limit data compression to a ratio of 2:1.
The communications ASIC is required to allow for high speed data transmission via
the acoustic modem. A series of filters and key shifting algorithms are used to achieve
high transfer rates with low signal to noise ratios. This is explained in greater detail in
the communications section.
Energy management and experiment control are both part of programs that reside in
the computer. These programs run continuously and can have their schedules changed by
the ground station to better suit the needs of the mission. In the event of a necessary
restart, the programs are all stored in the memory recorders, and can be recovered once
the system is rebooted.
5.3.3.2 Acoustic Modem
A communication system between the lander and probe that provides a satisfactory
transmission rate and a long range had to be investigated. Because of the unknown
thickness of the ice, it may be necessary to pass a signal over a distance of up to 5 km.
The thickness of the ice varies considerably between different regions, however a worse
case scenario must be considered.
One possible solution is to use a communications cable. Such a cable could have
reinforcements to take any stresses, and then have a core of fiber optic cables or copper
cables. While fiber optic cables are run successfully over hundreds of kilometers for
telephone and digital communications, such systems are not subject to restrictions in
cable sizes and they don’t operate in the extreme conditions found on Europa.
Furthermore, many signal boosters are used to increase the fidelity of these
communications line. For the probe however, space constraints only leave the option of
having a relatively small bundle of thin fiber optic cables. The very low temperatures of
the Europan environment create many problems for operating fiber optic cables. Another
problem which arises is line stretching. If the fiber optic cable is stretched there is a large
71
amount of attenuation. If the strand is actually fractured, then the signal is terminated at
the fracture. Because of these shortcomings, fiber optics was not pursued.
To deal with the variable distances that have to be planned for, and to deal with the
problems of having a robust design, an acoustic approach was applied to the probe
communication system. This system allows for communication at up to 5000 bps over a
distance up to 5 km. By using various encoding routines and well known filtering
techniques, it is possible to obtain a satisfactory communications speed with low signal to
noise ratios. The acoustic approach has been applied extensively in ocean
communications on Earth, and has provided communications speeds up to 9600 bps over
distances on the order of 100 km.
Both the probe and the lander have a sonar transducer and a series of
hydrophones. This will allow for two-way communications between the two crafts. The
systems’ transducers are capable of vibrating at 10 kHz, and the hydrophones are
designed to be most efficient at that frequency. This gives us a maximum
communication rate of 5000 bps. This rate will only occur near the surface, where the
attenuation is the smallest. At a depth of 5 km, the transmission rate will more
realistically be 300 baud. The transducer on both the lander and the probe will be 11 cm
half-spheres. On the probe, a section near the top will be reserved for a hydrophone.
This hydrophone is used to receive data from the lander. The weight of this system is
about 2 kg and will require 1.66 W of electrical power.
Although the actual transducer and hydrophone would probably be custom
manufactured to withstand the rigors of the Europan environment and to conform to the
probe design, an idea of the size and weights of these systems can be estimated on the
basis of currently available equipment. A transducer capable of operating at the projected
frequencies and depths would have a size of 11 cm across and a height of 7.5 cm and
weigh 1.1 kg. A picture can be seen in Figure 5.11. The actual transducer will be
hemisphere shaped to better direct the acoustic energy towards the surface. The
hydrophone for this system is very small, with a mass of 4 g and a volume of 3.3x2.2x1.5
cm3. A photograph can be seen in Figure 5.12
72
.
Figure 5.12: A photo of a typical sonar transducer
Figure 5.12: A photo of a typical hydrophone
To determine the power requirement, a calculation was made, which takes into
account background noise and signal losses. This can be used to determine acoustic
power and, therefore, the electrical power required.
It was assumed that the minimum signal that could be received by the hydrophones at
either end, would have to be 20 dB above the background noise. On earth, the
background noise for the ambient sea, assuming no wind or shipping, is 25 dB at 10 kHz
(Urick, 1983). This is caused by circulation, seismic disturbances and wave interactions.
With the ice, there is extra noise from the cracking and expanding due to internal pressure
changes and temperature changes. This can raise the level up to 40 dB on earth. Since a
lot of this noise is due to wind, an increase of only 10dB is used. This would be typical
for ice that is relatively uniform and with a rising temperature. To take into account any
additional noise, an additional 5 dB is used in this calculation. This provides us with a
background noise level of 40 dB for our calculations.
The signal loss over a distance, in meters, is given by the equation:
TL= 20 log(D)
73
(5.10)
For a distance of 5000 m, the signal loss will be 74 dB. In ice, the signal loss at 10
kHz due to absorption is 0.875 dB/km. The loss for our system at 5 km will, therefore, be
another 4.4 dB.
The equation relating the original signal to the signal received can be found as:
SL-TL=NL-DI-DT
(5.12)
Were SL is the source level, NL is the ambient noise level, DI is the receiving directivity
index and DT is the detection threshold. For our system, a DI of zero was assumed. The
minimum sound level that must be emitted at 5 km, to achieve viable communications
with the surface, is 150 dB. A contingency factor of 1.2 is applied, and, hence, a
minimum level of 170 dB must be generated at the transducer. The equation relating
decibel levels to acoustic power is:
SL=170.8+10log(P)
(5.12)
For our case, the power, P, is 0.8317 W. If we assume an efficiency of 50% in power
conversion, we will require 1.66 W of electrical power for this system.
5.3.4 Probe Structure
Because the ice/water probe may have to function several kilometers below the surface, it
is important to determine the necessary wall thickness of the probe shell so that the probe
doesn’t collapse under the extreme pressures. This will require a thin-wall buckling
calculation. Factors that affect the buckling properties are: material properties, cylinder
dimensions, and internal supports. Although the probe has an overall length of
approximately 1.5 m, it is actually a series of much smaller segments. With the current
design, the largest segment is only 30 cm long. It is advantageous to make the thickness
of the skin uniform to ease fabrication costs. Therefore, the calculations were performed
with a cylinder length of 30 cm.
A first approximation can be made using the relation for the semi-infinite cylinder
model.

K p 2 E  t  2
Pe a
1.04 L 1   2

,
K



p
t
12 1   2  L 
at


For this formula to be valid,
74

(5.13)
2
a
L
  
t
a
(5.14)
where L is the length, a is the radius, Pe is the pressure, E is the modulus of elasticity,  is
Poisson’s ratio and t is the thickness. If Equation 5.14 does not hold, then a more robust
model is necessary. A more robust model would take into account the end conditions and
uses the formula:
Pe a

Et


L
 n2
n2
a 2

2
 ah 2
La 4

121   2  n 2 a 2  n 2 2
L
(5.14)
where n is the mode of buckling.
The material used in our probe is Ti- 5Al –2.5Sn, a Titanium alloy. This was chosen
for its strength, light weight and corrosion resistance. It has a density of 4.49 g/cm3,
modulus of elasticity of 107Gpa, and Poisson’s ratio of 0.3. The pressure at 5km of
depth may be approximately 6 MPa. A safety factor of 2 was used.
With these parameters, the semi-infinite calculation was first made, resulting in a
thickness of 3.5 mm. With this thickness, Equation 5.14 does not hold. The calculation
was then made with Equation 5.15. Figure 5.13 shows the thickness as a function of
buckling mode. The largest value of thickness should be chosen. The probe wall
thickness is therefore going to be 3.28 mm, and the shell will weigh 14kg.
Skin Thickness vs. Buckling Modes
3.50
3.28
3.00
2.88
2.52
Thickness (mm)
2.50
2.24
2.03
2.01
2.00
1.86
1.72
1.50
1.61
1.00
0.50
0.06
0.00
0
2
4
6
8
10
12
Buckling Mode(n)
Figure 5.13: Plot of necessary skin thickness versus buckling mode
75
6.0 Mission Life Cycle Cost
6.1 Introduction
For this design, the minimum life cycle cost is divided into four phases. These phases
are design, development, fabrication, integration, and test (DDFI&T), launch vehicle,
flight operations, and data retrieval. Tables <cost 1>, <cost 2>, <cost 3>, and <cost 4>
show the cost breakdown, in major components, for the DDFI&T, launch vehicle, flight
operations, and data retrieval sections, respectively. There are a variety of ways to
determine these costs. One way is to acquire the exact costs from the manufacturers for
the “off-the-shelf” items and estimate costs for the remaining elements of the spacecraft.
Choosing this method is time consuming and the design and development costs would be
mere guesses. In some instances, the costs for the individual components are not
available. Another method is outlined in Space Mission Analysis and Design (<ref.
SMAD>). Though this process is thorough, it uses a traditional method of cost analysis
used in the past when applied to the large spacecraft/satellites designed in the 70’s and
80’s, which results in a design and development cost on the order of $1 billion. Also, this
method is geared more towards satellites rather than interplanetary probes. A simple and
efficient way to generate preliminary spacecraft costs is to use available cost information
on recent interplanetary missions . For this project, the majority of cost predictions are
based on the estimates for the Mars Pathfinder Project outlined in "Mars Pathfinder
Mission Operations Concepts" (Sturms, et al.).
The Mars Pathfinder is similar to Zeus in certain ways. The operations teams for
both missions will be smaller than for interplanetary spacecraft such as Galileo or
Cassini. This ensures lower flight operations and data retrieval costs since individual
employees earn anywhere from $40,000 to $140,000 per year for the duration of the
mission. This saves the mission thousands of dollars per year. The Pathfinder's rover
and Zeus' Ice/water Probe are also similar in that they are both autonomous and require
significant design and development.
However, there are important differences between the two missions as well. The
power systems differ; Zeus contains a radioisotope power source while Pathfinder uses
solar power. The lengths of time for flight and data retrieval are much greater for Zeus
than for Pathfinder. The instruments for Mars Pathfinder were heavily developed by their
design team; for Zeus, a majority of the instruments are "off-the-shelf" or require
relatively modest design and development efforts.
There are other differences in addition to those described above that require
consideration. Likewise, there are other similarities besides those outlined earlier.
76
Nevertheless, the similarities are such that the Mars Pathfinder costs can be used for
preliminary estimates of the Zeus Project costs.
Table <cost 1>: Estimated DDFI&T component costs
Major Components for
DDFI&T
Project Management
Mission Director and Operations Teams
Flight Systems
Ice/water Probe
Lander Instruments and Science
experiments
Ground Data and Mission Operations
Reserves and others
Total
Cost
(FY98$M)
11.3
18
60
40
16
11.5
25
181.8
Table <cost 2>: Estimated launch vehicle cost
Launch Vehicle
Cost
(FY98$M)
70
Delta 3
Table <cost 3>: Estimated flight operations component costs
Major Components for
Flight Operations
Management
Experiment Team
Engineering Team
Ground Data System support
Reserves
TOTAL (per year)
Cost per year
(FY98$M)
1.5
2
3
1
1
8.5
Table <cost 4>: Estimated data retrieval component costs
Major Components for
Data Retrieval
Management
Cost per year
(FY98$M)
1.5
77
Experiment Team
Engineering Team
Ground Data System support
Reserves
TOTAL (per year)
3
0.3
1
1
6.5
6.2 Spacecraft Design, Development, Fabrication, Integration, and Test
For a three-year DDFI&T sequence, the cost is divided into seven sections. These are
Project Management, Mission Director and operations teams, flight systems, instruments
and science experiments, Ice/water Probe, Ground Data and Mission Operations Systems,
and reserves and other expenses. The Zeus Project is designed to qualify under NASA's
Discovery missions. Hence, the cost for this project, not including launch costs, from
conception to 30 days after launch, needs to be capped at $181.8M in FY98$. By
modeling the DDFI&T cost with that of the Pathfinder's preliminary costs, the goal to
make Zeus a Discovery mission is within reach.
The Project Management cost is standard. For a preliminary cost analysis, the cost
from Pathfinder can be used and adjusted for inflation. This results in a Project
Management cost of $11.3M.
The Mission Director and operations teams are strictly based on the organization of
the Mars Pathfinder, with adjustments made due to differences in the two missions. The
Missions Director is a liaison between the operations teams and the Project Manager.
There are five operations teams: the Experiment team, the Engineering team, the
Multimissions Operations System Office (MOSO) support team, the Ground Data System
(GDS) maintenance crew, and the Deep Space Network (DSN) operations team. The
Experiment team includes members devoted to the development of the experiments and
the operations of the Ice/water Probe. The Engineering team is responsible for mission
and flight planning. The MOSO support team handles data system operations, data
administration, and image processing (Sturms, et al.). The GDS maintenance crew
maintains the Ground Data System. Finally, the DSN operations team acts as the link
between the project and the Deep Space Network. MOSO support and DSN operations
are outside systems provided by NASA, and the services of these systems are provided at
no cost to the project. The Experiment team, Engineering team, and GDS maintenance
crew will be the same for the Zeus and Pathfinder Projects for initial estimates.
Therefore, the cost for the Mission Director and operations teams is approximately $18M.
78
The flight systems segment of DDFI&T consists of software and hardware
development for flight, construction of the main engines and thrusters, and other flight
related areas. The Mars Pathfinder Project devoted a significant amount of time to the
development of the flight system; the Zeus Project uses existing software and hardware,
reducing design and development cost. Hence, the flight systems cost for this project can
be estimated to be $60M.
The Ice/water Probe is approximately $40M. The main cost driver for the Probe are
the radioisotope power source and heat distribution and control system, which may cost
$20M. The components that require design and development are the capillary
electrophoresis system and the Probe Arresting Sub-System. The cost listed above is
merely a very preliminary cost representing a 60% increase when compared to the rover
of the Pathfinder Project. The rover DDFI&T cost was approximately $25M in FY96$.
Since the Ice/water Probe for the Zeus Project is not as complex as the Pathfinder's rover,
the cost of the Probe itself may be less than the rover. However, the expense of the
power sources and testing the probe greatly increases the overall cost of the Ice/water
Probe. Therefore, the $40M estimate is merely a cap on the DDFI&T cost of the Probe.
The instruments and science experiments for the lander of the Zeus spacecraft are
mostly "off-the-shelf" products. Therefore, the instruments and experiments may not
require extensive design and development. An estimate for this cost is $16M in FY98$.
The Ground Data and Missions Operations Systems (GD&MO)includes software and
hardware costs for the Ground Segment (GS) of this mission, as well as testing and
training for the GS personnel. In most missions, the software cost is usually the driving
cost in this section. Therefore, the GD&MO cost for the Pathfinder Project can be used
for this project. Hence, the GD&MO cost estimate is $11.5M in FY98$.
The cost for reserves and other expenses that are either unforseen or improperly
estimated is set at $25M.
6.3 Launch Vehicle Cost
The Zeus spacecraft will most likely leave Earth's atmosphere by way of the Delta III
launch vehicle. Since information on the cost of a Delta III launch is not yet available,
only an estimate of the launch vehicle cost can be used. The Delta II ranges from $45M
to $50M. Delta III can carry a larger payload and can enter a higher orbit than the Delta
II. This leads one to believe that the Delta III would cost more. Hence, an estimate of
$70M in FY98$ can be made.
79
6.4 Flight Operations Costs
The flight operations cost is divided into five segments: management, Experiment
team, Engineering team, GDS support, and reserves. The management cost for this
project is $1.5M in FY98$ per year, which is comparable to that of the Mars Pathfinder
Project. Comparisons with the Pathfinder mission result in $2M per year and $3M per
year estimates for the Experiment and Engineering teams, respectively. For preliminary
cost estimates only, the GDS support and the allocated reserves are estimated to be $1M
per year each. This results in flight operations cost of $8.5M in FY98$ per year.
6.5 Data Acquisition Cost
The data acquisition cost is divided in the same manner as the flight operations. The
majority of the data acquisition cost is the same as the flight operations cost since flight
operations and data acquisition do not coincide with each other except when the
spacecraft is in orbit around Europa. The difference between these two costs resides in
the cost for the Experiment team and the Engineering team. The Experiment team is
more active during data acquisition than when the spacecraft is in flight. Therefore, this
cost increases for data acquisition to $3M per year. The Engineering team consists
mainly of flight personnel. Only a fraction of the personnel is devoted to mission
operations, i.e., data acquisition costs do apply. Hence, the Engineering team cost is
$0.3M per year. The overall data acquisition cost is $6.5M in FY98$ per year.
80
7.0 Technological Readiness
The underlying idea behind the Zeus probe is to build this spacecraft for a minimum
cost and with the greatest simplicity. To achieve this goal, most of the systems were
chosen on the basis of either their use on previous missions or missions to be launched
during the next few years. By planning on using equipment designed for previous or
currently design spacecraft, a large portion of the design problems and costs may be
eliminated.
The primary communication systems and electronics used throughout the ice/water
probe and the spacecraft bus are off the shelf components. The scientific payloads on the
lander are derived from hardware which will be used on spacecraft designed to be
launched at the end of the century and very early 21st century. The same can be said for
the ice/water probe. Despite our drive towards this goal, there are still certain systems
which are cutting edge by today’s standards.
The radioisotope heating units and power converter cells will most likely be used in
the Pluto/Kuiper Express mission under study for launch in 2004. The technological
hurdles of developing these systems has already been tackled and solved. Furthermore,
the ability of the ice/water probe to melt through the deep ice will have to be tested and
validated under arctic conditions. This will probably constitute the majority of the
research money necessary for the ice/water probe.
The acoustic modem communication system for probe-lander communication is an
innovation that is previously untried. The computer hardware, transducers and
hydrophones are all readily available from current ships which use these systems
routinely. The challenge will be to make these systems ready for operation under
Europa’s conditions. Redesign and testing of these components will take approximately
two years. Fabrication should not be long since the manufacturing processes are readily
available.
The capillary electrophoresis system will be the first of its kind used on a spacecraft.
No space mission prior to this one has used this system for chemical analysis.
Fabrication of the glass plate has already been accomplished in the middle of the 1990’s.
Despite this construction, changes will most likely have to be made to make the system
space ready. The automation of this experiment is also unprecedented. Since all these
81
systems exist in laboratories the human element is always present. This will be the first
time that a computer will have to make decisions usually made by a human during an
experiment. While these automation systems are not impossible to build, a certain
amount of time should be estimated. It is assumed that it will take approximately one
year to flight ready the capillary system and it will take the entire three research years to
develop all the elements of the automation system.
The data compression ASIC is another new addition on a spacecraft. While previous
spacecraft compressed data in the past, this is the first time that a specially designed
processor will be used to employ a lossy compression on scientific data. While no
problems are associated with designing such a processor, the delay time between design
and fabrication will not make this chip ready until ~2002. Design of the chip should take
no more than six months since it will be a combination of digital signal processor and
math co-processor.
While all of the technology on board the Zeus is cutting edge and proven on earth,
some of the more radical hardware will be refined for this space mission.
8.0 Outstanding Issues
Because of lack of resources and time, several aspects of the mission and design were
not completed to our satisfaction. The first major unresolved issues were the thermal and
radiation management for the spacecraft. Radiation management was handled by using
radiation hardened instruments, providing shielding and by placing hardware in the
shadow of the tanks. No real analysis was done to determine exposures for the given
instruments. Thermal management was also not carried out fully due to time and
resource constraints. A first approximation of system weight and temperatures was
made using crude estimating equations. For desired static and dynamic structural
analysis and thermal analysis, a finite element code should be used.
Another major issue that was not resolved fully was the trajectory design. Although
the RFP specifies that we should assume a Europan orbit, for the sake of completeness
the spacecraft was designed in a configuration to be launched from earth. The analysis
that would be necessary to calculate a detailed course was beyond the scope of this
project, and a crude estimate based on previous missions was used. Similarly the landing
82
v calculation was made using a crude first guess. A more accurate analysis of each of
these phases is needed.
In recent weeks it has become increasingly apparent that the surface of Europa is not
flat and smooth locally. This lack of smoothness may pose stability problems for landing
of the spacecraft. Designing a series of legs capable of dealing with variations in the
surface topography would add to the complexity of the entire system. Instead of
designing a system to deal with these surface features, we have decided to attempt to
remove the surface features locally. Since this roughness is on the order of less than
0.5m, the spacecraft will simply melt the local region flat before landing. This will be
accomplished by hovering above the landing site for a short period of time until these
features have been melted down. By performing this task, we ensure stability of the
spacecraft after landing without complicating the design.
One final issue is the launch vehicle. As the probe was refined the mass approached
the weight limit of the Delta II launch family. Launching on the Delta II would be more
advantageous than launching on the Delta III because of even lower launch costs and
because this system has been used numerous times in the past. If the 5m fairing is used
on the Delta III, then no structural modifications would have to be made. If only the
current fairings are used, then the dish location would have to be altered. To make a
definite conclusion on the feasibility of using the Delta II, the docking mechanism would
have to be designed to determine the weight of the system. If the spacecraft mass and
docking mass is small enough, then the Delta II will be the launch vehicle.
83
9.0 Conclusion
The Zeus spacecraft is designed to provide a better understanding of the features of
Europa. Launch of the Zeus is projected for 2005 aboard a Delta III booster. It will
reach Jupiter via a Earth-Venus gravity assist. Once in orbit around Jupiter, a series of
gravity assist maneuvers using Jovian moons and small burns will place it in a 20km orbit
above Europa. It is assumed that these maneuvers will take up to six years. Orbital
experiments are expected to be run by 2011
From orbit the spacecraft will do high resolution photography and topographic
imaging with a laser altimeter. This will be used by NASA to refine their initial landing
site selection. After landing the spacecraft will study the seismic, physical and chemical
properties of the surface. Surface pressure and temperature, as well as chemical
composition will be measured by a collection of instruments.
Subsurface features are measured by an ice/water probe which is capable of
descending at a rate of 100meters per month. This melting is done via a 2kW RHU
located at the bottom of the probe. A portion of this heat source is used for power
generation. The probe will measure chemistry, pressure and temperature as a function of
depth during its descent. Upon reaching the ice/water boundary photography will be
added to the list of experiments being run. The probe remains in position by tethering the
lower portion to the upper portion which is imbedded in the ice 30m above the boundary.
Two-way communications with the lander are handled via an acoustic modem.
The end of mission life is projected to be anywhere between 2012 and 2016,
depending on the depth of the ice at the landing site. The beginning of scientific return
will occur during the orbital phase of the mission, which should be in 2011.
The Zeus probe uses someoff the shelf hardware and tried methods to create a
relatively inexpensive instrument for scientists to probe Europa’s features. While certain
issues remain to be resolved, the basic concept shows the feasibility and the opportunity
of success in carrying out a mission of this type.
84
10.0 References

Bate, et.al, 1971, Fundamentals of Astrodynamics, Dover Publishing.

Beauchamp, Patricia, 1994, Pluto Integrated Camera-Spectrometer (PICS): A Low
Mass, Low Power Instrument for Planetary Exploration, AIAA Conference on Small
Satelites Paper.

Beauchamp, Patricia, 1996, The Sciencecraft Process, NASA Technical Report
Abstract.

Bentley, 1993, ‘Ice Thickness, Bed Topography and basal-reflection strengths from
radar sounding, Upstream B, West Anarctica,’ Annals of Glaciology, Vol.20, pp.148159.

Bentley, D. P., Tisdale, D., ‘Development Testing of TSS-1 Deployer Tether Control
System Mechanisms,’ ’ Tethers in Space Toward Flight, AIAA, Washington, D. C.,
pp. 352-360.

Bogorodsky, V.V., Bentley, C.R., Gudmandsen, P.E., 1985, Radioglaciology

Brown, Charles; "Spacecraft Propulsion"; American Institute of Aeronautics and
Astronautics, Inc; 1996

Bryson, Arthur; "Control of Spacecraft and Aircraft"; Princeton University
Press, 1994

Callister, William D.; "Materials Science and Engineering", 3rd Edition; John
Wiley & Sons, Inc,; 1994

Chela-Flores, Julian, 1996, Habitability of Europa: Possible degree of evolution
of Europan biota, Europa Ocean Conference, Capsitrano Conference, No.5, p.21

Clark, Robert S., 1996, Innovations in a small package: NASA’s Planetary Integrated
Camera Spectrometer.

Corliss, William R. & Harver, Dourglas G.; Radioisotope Power Generation;
Prentice-Hall, Inc; Engelwood CLiffs, NJ; Copyright 1964

David, Leonard, 1996, Incredible Shrinking Spacecraft, Aerospace America.

DEOS, ‘Satellite Laser Ranging’
http://www.neonet.nl/providers/dutl…tudelft.nl/ers/tracking1/index
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
Desai, P.N., et.al, 1993, Aspects of Parking Orbit Selection in a Manned Mars
Mission, NASA Langley Technical Report Abstract.
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Dyson, Freeman, J., 1995, 21st Century Spacecraft, Scientific American, September
1995, Vol.273, No.3.
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Europa Fact Sheet (EFS),
website=http://www.jpl.nasa.gov/galileo/europa/,"Europa Fact Sheet"

Follas, Ronald B. ‘Mars Orbiter Laser Altimeter’
http://ltpwww.gsfc.nasa.gov/eib/mola2.html

Gagliardi, Robert; "Satellite Communications",2nd Edition; Van Nostrand
Reinhold, 1991

Garwin, ‘Shuttle Laser Altimeter II’
http://denali.gsfc.nasa.gov/research/laser/sla02/sla02

Gogeneni, S., Chuah, T., Allen, C., Jezek, K., Moore, R.K., 1997, ‘An Improved
Coherent Radar Depth Sounder.’

Harty, Richard B.; "Power System Comparison for the Pluto Express Mission";
Proceeding of the 30th Intersociety Energy Conversion Engineering Conference,
Vol 1., pp 619-624, Orlando, Florida; July 31-August 3, 1995

Howard, George E., Gray, John, Levi, Alejandro, 1989, ‘Damage Inspection and
Verification of Tethers (DIVOT),’ Tethers in Space Toward Flight, AIAA,
Washington, D. C., pp. 337-351.

Ivanenok III, Joseph F. and Sievers, Robert, K.; "AMTEC Radioisotope Power
System for the Pluto Express Mission"; Proceedings of the 30th Intersociety
Energy Conversion Engineering Conference, Vol. 1, pp 661-666, Orlando, Florida,
July 31-August 3,1995

Ivanenok, Joseph F & Siewers, Robert K.; "Radioisotope Powered AMTEC systems";
29th intersociety Energy Conversion Conference, Vol 1, pp. 649-656; Monterey,
California; August 7-11, 1994

Kargel, J.S., 1996, Europa's crust and ocean: Origin, composition, and possible
implications for IO's early condition, Europa Ocean Conference, Capistrano
Conference, No.5, p. 42

Kuhn, Reinhard and Hoffstetter-Kuhn Sabrina; "capillary Electrophoresis:
86
Principles and Practice"; Springer Verlag, 1993

Larson, Wiley J & Wertz, James R.; Space Mission Analysis and Design; Microcosm
Inc. In Torrance California and Kluwer Academic Publishers in dordrecht,
Netherlands, Copyright 1992

Lengyel, Bela; "Lasers",2nd Edition; Wiley-Interscience, 1971

McKane, Larry and Kandel, Judy; "Microbiology, Essentials and Applications";
McGraw-Hill, 1985

Murrill, Mary Beth, 1995, Revolutionary new miniature sensor system developed,
Stargate, 1996, from website: http://stargate.jpl.nasa.gov/pics.

Nasa, 1979, website= http://www.jpl.nasa.gov/galileo/europa/hst.html, "Hubble
Finds Oxygen Atmosphere on Europa"

Neal, C.S., 1979, ‘The Dynamics of the Ross Ice Shelf Revealed by Radio Echo
Sounding,’ Journal of Glaciology, Vol.24, No.90, pp.295-312.

Raju, G., Xin, W., Moore, R.K., 1990, ‘Design and Development, Field Observations
and Preliminary Results of the Coherent Anarctic Radar Depth Sounder (CARDS) of
the University of Kansas, USA,’ Journal of Glaciology, Vol.36, No.123, pp.247-259.

Robin, G. DE Q., 1975, ‘Radio Echo Sounding: Glaciological Interpretations and
Applications,’ Journal of Glaciology, Vol.15, No.73, pp.49-63.

Rodgers, D.H., Alkalai, L., Beauchamp, P.M., 1996, The Kuiper Express: A
Sciencecraft, NASA Technical Report Abstract.

Scala, E., ‘Tethers in Space, and Micrometeroids,’ ’ Tethers in Space Toward Flight,
AIAA, Washington, D. C., pp.372-378.

Scott, Craig; "Modern Methods of Reflector Antenna Analysis and Design"; Artech
House, 1990

Shock, Alfred & Or, Chuenc T & Kumar, Vasanth; " Radioisotope Power System
Based on Derivative of Existing Stirling Engine"; Proceedings of the 30th
Intersociety Energy Conversion Conference, Orlando, Florida; July 31-August
3,1995

Shock, A.; "Design Evolution and Verification of the General-Purpose Heat
Source"; #809203 in the Proceedings of the 15th Intersociety Energy conversion
Engineering Conference; Washington, DC; AUgust 6-11, 1989
87

Shock, Alfred & or, Chuenc Tak & Kumar, Vasanth; "Design Modifications for
Increasing teh BOM and EOM Power Output and Reducing the Size and Mass of the
RTG for the Pluto Mission"; 29th Intersociety Energy Conversion Conference,
Vol 1,pp 541-547; Monterey, California; August 7-11,1994

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1997

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
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Conferenc, Capsitrano Conference, No.5,pp.67

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Express"; Proceeding of the 30th Intersociety Energy Conversion Engineering
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
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
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
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
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
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
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
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Appendix 11.1 Overview of Present Day Radioisotope Power Sources
The power load required by the lander is 135 electrical watts. The lander power
source must be able to supply this up to the end of mission, estimated to be 15 years, and
88
under the given environmental conditions. The best suitable form of technology
currently available for this mission is a radioisotope power generating source. The
radioisotope power generator converts the heat energy released from the nuclear decay of
some radioisotope into electrical energy. The radioisotope most commonly used for such
an application is Plutonium 238. This is due to its ideal characteristics of half life,
specific power, and radiation emission. Plutonium 238 fuel comes in the form of
plutonium oxide, PuO2. The characteristics of PuO2 are:
Table 11.1: Properties of Plutonium Oxide: Pu238O2
Half Life of Plutonium 238
Density of PuO2
Specific power of PuO2
Alpha energy emitted per decay
Gamma energy emitted per decay
87.4 years
11.46 g/cm3
490 W/kg
5.47 MeV
0.01 MeV
(Properties are taken from (Corliss & Harvey, Appendix 1. Also available are the properties of other compounds of
Pu238 as well as other isotopes of possible interest.)
The conventional and space proven fuel capsule for housing Plutonium 238 isotope
fuel is the 250-Watt General Purpose Heat Source (GPHS). The GPHS was developed
back in 1980 with the aim to develop a standard universal module design for a 250 Watt
isotopic heat source to be used in a wide range of space applications. The GPHS design
would successfully eliminate the time and cost invested in designing a new and unique
radioisotopic heat source each time a space project would come requiring radioisotopic
power. Figure 11.1 from Shock, 1980 shows the general components of the GPHS, and
Figure 11.2 shows its dimensions (note that the height, 2.65 cm, is half the height of the
GPHS). Two GPHS will be used for the lander’s power generator.
89
Figure 11.1: General Purpose Heat Source Module
Figure 11.2: General Purpose Heat Source Module Sectioned at Midplane
90
There are a few options available to convert thermal energy to electrical. These are
the thermophotovoltaic converter, the thermo couple converter, a dynamic Stirling engine
converter, and the AMTEC converter. Due to the long mission flight, the power
converter must be reliable up until the end. Although there have been great
breakthroughs in Stirling energy conversion technology, particularly by Stirling
Technology Company (Montgomery, Ross, Penswick, 1996), which has drastically
improved reliability and decreased torquing effects making dynamic energy converting
technology a reality for space use, an operation duration of fifteen years has yet to be
demonstrated. Furthermore the engine is bulkier then some of the other systems such as
the AMTEC converter. In the past, most space missions have used the traditional p-n
thermocouple to convert the heat energy into electrical. By placing two different semiconducting materials across a temperature difference, an electrical potential difference
can be generated between the two semi-conductors. Thermocouples have proven very
successful and reliable in the past. However, efficiencies of the conventional
thermocouple are on the order of 5-6%. This would require a larger heat source than the
newer more efficient converters and so this design was discarded. Thermophotovoltaic,
TPV, concept converts IR emitted radiation directly into electricity. They operate with
around 13% thermal to electrical energy conversion efficiency, and require more testing
in determining long term effects of the Plutonium heat source on the cell. The AMTEC
design operates by converting heat energy into sodium fluid flow work in a regenerative
cycle in a heat pipe, which induces a pressure and concentration potential across a
sodium ion selective membrane, creating an electrochemical difference which can be
converted into electrical current power. Designs based on this principle can run at
efficiencies as high as 25% (under certain conditions). Furthermore, the current and
newest designs as produced by AMPS, are small, lightweight, and compact. Underwood,
1995 contains a more in depth trade study of the different technologies for the Pluto
Express mission. An AMTEC design was chosen for our power converter.
Appendix 11.2 Compression ASIC
One of the problems of deep space probes is maintaining a large data stream. For this
reason, a lot is packed into a small package. At the same time, the package can not grow
91
to be too unwieldy or the possibility of mission failure increases. For our probe, a limited
data transfer rate of 120 kbps is available. This was a trade off between desired rates and
power/size requirements for the mission. To pack as much into this 120 kbps, it is
important to take measures which reduce the data throughput or increase the bandwidth
transferable on this channel. This would involve compression.
For compression our probe uses a custom Application Specific Integrated Circuit
(ASIC) which has the JPEG compression algorithm coded on it. This algorithm allows
for both variable size images as well as variable color depth images. It also has a unit
which can convert color images to grayscale images before encoding. All of these
function have logical defaults, but are configurable by the ground via the main computer.
A sample of the advantages of using these compression routines can be seen in these
images recently acquired from the Galileo spacecraft, see Figure 11.3. These images are
of a crater on the surface under various stages of compression from no compression all
the way to 7.7:1 compression.
Figure 11.3: Image under various compression states
All of these images were generated by taking the original raw image, in an uncompressed
Tagged-Image Format File (TIFF), and using a preprogrammed progressive JPEG
compression in a popular graphics program at various degrees of compression.
92
From the comparison of these high detail regions, compression ratios as high as 7.7:1
and probably higher are capable without any real loss in surface detail. If close detail is
not necessary then even higher compression ratios are possible. While there is some
aliasing in large homogeneous areas, such as the lower left corner of this image, these are
regions with little differing data. This means that this aliasing does not reduce the
scientific data presented. The progressive JPEG algorithm is designed to minimize the
losses in areas of highly varying pixel information.
While this is acceptable for imaging data, to compress scientific data a lossless
compression is necessary. In a lossless compression data is not “thrown away” as in the
JPEG compression. This means that what you put in is what is gotten out. While this is a
great feature of lossless compression, compression ratios of highly variable data is
limited to between 2:1 and 2.5:1 compression ratios. These depend on the randomness of
the numbers, where higher randomness creates fewer patterns and thus less compression.
Unfortunately if a compression of a digital signal is to be undertaken, then it must be
done in a lossless way.
To get around this problem, our probe has another ASIC dedicated to pre-processing
the scientific data before transmission. While errors on single values create problems, if
compared to an entire field, the error becomes negligible. This is readily apparent in the
above images. To take advantage of this fact, the data collection ASIC will collect the
scientific data and create “images” of the ground based on this data. These images will
be created by writing to memory as the probe scans across and moves along the ground.
This creates the same effect as would be found with a scanning electron microscope
image.
Each pixel will hold the value of the measurement as a 24-bit integer. This is similar
to the coding used for image files. This gives a gradation of 16.7 million possible
increments for an instrument. So for example, if the laser altimeter measures altitudes
between 0 and 20 km, the resolution of the recorded image will be 1.2m. This is more
than adequate for all instruments onboard. After processing the data into these images,
the ASIC passes the data and the selected parameters on to the JPEG ASIC, which
compresses the data to levels of at least 5:1.
93
By dealing with our scientific data in this way, the data transmission rate becomes
more than acceptable for our measurements. Another benefit is a minimized computer
storage necessity. This “imaging” is often done on the ground, but is easy to implement
in an automated way on the spacecraft. The benefits, which considering the compression
capabilities, far exceeds the mass and power of the processor required for this operation.
Appendix 11.3 Calculation of Power Required for Ice Melting
The primary mission of the Zeus probe/lander involves melting through the icy crust of
Europa. Determining the heat requirement involved a lengthy analysis several
assumptions. Also included in this analysis are the physical and thermal properties of
Europa. These properties were gleaned from many abstracts from the Europa Ocean
Conference, and are as follows:
  917
kg
m3
Cp  1925
J
kg  K
  147
.  10
6
m2
sec
J
sm K
kJ
  284
kg
  2.6
Using this information and the following equation:
Ri  Ro 


U 

2 

Heat  4T      ( Ri  Ro )  e



(11.1)
  Cp    Ri  Ro 2  U  2    Ri  Ro 2  U    
Where U is the average descent velocity, Ri is the radius of the probe and Ro the radius
of the melting region, see Figure 11.4. Delta T is the rise in temperature required.
With this equation it became possible to use a program to iteratively solve for
variations in probe radius as well as descent velocity. From this program it was possible
to determine the decent velocity for our probe radius. The assumed heat source for these
calculations was a radioisotope heat source of 2 kW. With this heat source, a probe
94
radius of 7cm and a melting zone thickness of 5mm, the descent rate is 108 m/Month.
That equates to a little over 1 km per year.
Probe Wall
Ri
Melting Region
Ro
Ice wall
Figure 11.4: Radii used in the program
Appendix 11.4 Hohmann transfer calculations
The following equations correspond to the V calculations [Bate,1971] required for
the Hohmann transfer. The known constants for these calculations are:
e = 3.203E12 m3/s2
Me = 4.8E22 kg
Re = 1569 km
G = 6.67259E-11
Nm2/kg2
Orbital period , =
Rapoapsis = 1589 km
Rperiapsis = 1569 km
e = GMe
3.06822E5 sec
The velocity of the 20km circular orbit is found using:
ve 
e
r
 1419.73m / s
(11.2)
The velocity at the apogee of the orbit is found by:
h2
ra 

1  e cos 
 h  2.249  10 9
95
(11.3)
e
ra  rp
ra  rp
 0.006331
h  ra v a  v a  1415.26km / s 2
v1  vc  va  4.45m / s
(11.4)
(11.5)
(11.6)
The velocity at the perigee of the orbit is found by:
vp 
h
 1.43km / s
rp
(11.7)
Europa surface velocity is 0.032 km/s.
 v 2  v p  ve  1.43km / s
2
2
(11.8)
therefore the total v for the entire transfer is:
vtot  v1  v2  1.44km / s
Appendix 11.5 Europan Environment Statistics
Discovery:
Diameter (km):
Mass (kg):
Mass (Earth = 1)
Surface Gravity (Earth = 1):
Mean Distance from Jupiter (km):
Mean Distance From Jupiter (Rj):
Mean Distance from Sun (AU):
Orbital period (days):
Rotational period (days):
Density (gm/cm3)
Orbit Eccentricity:
Orbit Inclination (degrees):
Orbit Speed (km/s):
Escape velocity (km/s):
Visual Albedo:
Surface Composition:
Jan 7, 1610 by Galileo Galilei
3,138 km
4.8 (1022) kg
0.0083021
0.135
670,900 km
9.5
5.203
3.551181 days
3.551181 days
3.01 gm/cm3
0.009
0.470°
13.74 km/s
2.02 km/s
0.64
Water Ice
96
(11.9)