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

2

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

3

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

Sub-Surface Ocean Entry

Pressure andTemperature

Photographs Water CompositionExperiments

Europan Orbit Entry

Ice ThicknessMeasurements

SurfaceAltimeter

Pictures

Pressure, Temperatureand Composition Readings

PASS Activation

Landing

SurfaceChemicalAnalysis

Seismic Readings Pictures Probe Release

Figure 1.1: Mission Events Sequence

4

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 2016Initial Approval/ Funding of MissionCompletion of R&D, tests and evaluationConstruction of SpacecraftPreparation of Launch/ LaunchSpace journey up to arrival near JupiterSurface scans begin/ Landing on EuropaReach Ice/Water Interface (assuming between 100m up to 5km)

5

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

6

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.

7

2.1 General Performance Parameters of the Configuration

Table 2.1: Weights and power requirements of major components

UNIT WEIGHT (kg) POWER (Watts)

PROBEPower Generation 20.72 44-62.5 (generated)Capillary Electrophoresis 5.5 10.75Probe Shell 14PASS 8.571 20Photographic Equipment 0.6 0.6 to 10.1Acoustic Modem 1.1 1.66Computer 1 5Pressure Transducers 0.1 0.1Thermocouple amp 0.25 0.1PROBE TOTAL 51.84

8

LANDERLaser Altimeter 1.5 1Arm/Spectrometer 2.5 0.25Fixed Camera(3) 0.25 1.5Zoom Cameras(4) 2 2Seismometer 0.1 0.1Main Propellant Tanks 48 -Main Engine Propellant 956 -RCS Tank 3.63 -RCS Propellant 30.25 -RCS Thrusters 3.6 -Helium Tank 31.3 -Helium 5.89 -Engines 13.5 5Valves 5 2Radioisotope PowerSource

7.96 Generates 163W

Communications System 35 53(High Gain), 20 (Low Gain)Star Tracker (2) 2.8 7Sun Sensors(4) 1 1.4Accelerometer 5x10-3 0.35IMU 0.7 10Computers 2 20Radiator Fins 7.172 -Heat Pipes 1.42 -Jacket 3.34 -Acoustic Modem 1.1 -Radiation Shielding 10 -

LANDER TOTAL 1210.6DRY MASS 306.5TOTAL(inc. propellant) 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

9

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

10

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.

11

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.

12Figure 3.3: Isometric view of final spacecraft configuration

13

Figure 3.4: Side view of final configuration

14

Figure 3.5: Top view of final configuration

15

Figure 3.6: Internal View of the Final lander Configuration

16

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.

17

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 three-

dimensional 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.

18

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%

• > 1500 m• > 500 m

Minimum Distance 2 mAccuracy ±20 cmDimensions (mm) 200 x 130 x 76 (L x W x H)Power (W) 6Weight (kg) 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.

19

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 µs. 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.

Claw

Optics for

Spectrometer

2-Degree of

Freedom Shoulder

Elbow Joint

20

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.

Figure 4.3: Interior and exterior view of camera, respectively

Helium Tank

Camera Optics

Lenses

GuideTracksCCD

RadiationShell

Panning Mirror

21

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

22

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-of-

the-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.

23

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 Micromachined Silicon, 10Hz Resonance, 6x10-9

m/sec2/Hz Noise FloorTransducer UHF Capacitive Displacement, 5x10-13 m/Hz SensitivityConfiguration Tetrahedral (3 Components of acceleration plus 1

redundant)Mass <0.1kgPower Consumption 100mWSize 5cm on edgeAcceleration Sensitivity Better than 10-8 m/sec2

Frequency Range 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

24

for Jupiter, it is assumed that the spacecraft will reach Jupiter using the same Venus-

Earth gravity assists as Galileo. Assuming a series of gravity assists in the Jovian system,

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:

−

∆= 1expgI

vmm

spfp (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

Helium

MMH

CombustionChamber

Helium

Helium

MMH

25

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:

+=

πσ i

pi

allowtens D

gmPDFSt

4,

(4.2)

where FS is the factor of safety, σ 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-Kuong-

Soave 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

26

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.

27

Table 4.3: Concerns in structural design

Considerations/Design Concerns Resulting IssuesLaunch and maneuver loading Structural assembly optimizationExtreme Radiation levels Sensitive component placement to

minimize shielding required to surviveJovian radiation.

Attachments/Mating to Delta III fairing Overall configuration constraints.Orientation/Placement of key components Specific component mission

conflicts/requirements.Thermal control Placement of Ice/Water probe and thermal

control devices.Europan surface conditions Avoid tip-over and surface non-conformityMinimizing cost, size, complexity 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,

28

Aluminum 7075-T6, Steel (Quenched and Tempered), and Graphite/Epoxy 8-ply quasi-

isotropic 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.

Figure 4.6: Chart of investigated material properties

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,

Chart of investigated material properties

0

200

400

600

800

1000

Ultimate Tensile Stress Ultimate Compression Stress

Ult

imat

e S

tres

s (M

Pa)

Aluminum

Gr/Ep

Steel

Titanium

Chart of Stress Ratios for Investigated Materials

0

0.05

0.1

0.15

0.2

0.25

Ultimate Tensile Stress Ultimate Compression Stress

Str

ess

to D

ensi

ty

Rat

io(m

^2)/

sec^

2

Aluminum

Gr/Ep

Steel

Titanium

29

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 of Tested Material Bar for the Test Load

0.01850.019

0.01950.02

0.02050.021

0.02150.022

Aluminum Gr/Ep Titanium

Tested Material

Mas

s (k

g)

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.

30

Table 4.4: Power Configurations for the Spacecraft

Mode Description Components Power1 Transfer Orbit Computer, Low

Data RateCommunications,

RCS, GN&C

50.75W

2 Europa Orbit with DataAcquisition

Computer RCS,GN&C,1-Camera,Laser Altimeter,Low Data Rate

Communications

62.25W

3 Europa Orbit with DataAcquisition and High Speed Data

Transmission

Computer RCS,GN&C,1-Camera,Laser Altimeter,High Data Rate

Communications

112W

4 Landing Computer, RCS,Engines/Valves,GN&C, Camera,Laser Altimeter

43.25W

5 Post Landing Operations High Data RateCommunication, All

Cameras,Seismometer, Sonar,

Computer,Arm/Spectrometer

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).

31

Table 4.5: Predicted AMTEC-RPS Performance

Generator Mass 7.96 KgThermal Power 750 WattsOutput Voltage 28 VOutput Current 3.9 ampsOutput Power 163.5 WattsSystem Efficiency 21.8%Specific Power 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.

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):

RTLGLGPN

Esrst

b log10log106.2880

−−+++++= θ (4.3)

Low Gain Antenna

2 meter, High Gain

Antenna Dish

Transmitter Horn

32

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, Lθ 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:

ηlog10log20log209.159 +++−= Hzt fDG (4.4)

Hzms fSL log2log2055.147 −−= (4.5)

where D is the dish diameter in meters, f is the frequency in Hertz, η 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.

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

Parabolic Main

Reflector Hyperbolic sub-

reflector

FeedSymmetry Line

33

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.

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.

Aluminum Shell

Aluminum ShellAluminum

Honeycomb

34

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

High Gain Antenna Support Ring

Positioning Motors

Graphite-Epoxy tubes

35

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.

Figure 4.13: GN&C System Block Diagram

Star Tracker

IMU

Sun Sensor

Acceleromete

Attitude

Determination

Flight Computer

Attitude

Control

High Level Maneuver

Commands Verification

U

P

L

I

N

D

O

W

N

L

I

Main Engines

RCSThrusters

36

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

CharacteristicsField of View (FOV), deg 10 x 10Sensitivity, mv 6.8Number of stars tracked Up to 20

Accuracy of attitude output (arcsec, 1σ) Pitch and Yaw Roll

414

Update Rate, Hz 10Power, W 7Weight, kg 1.4Operating temperature, °C -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

CharacteristicsField of View, deg Rectangular 35 x 35Accuracy, deg Better than 1° in entire FOVPower Consumption, W <0.35 W per FSS unitWeight, kg <0.25 kg per FSS unit

37

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

CharacteristicsWeight, kg 0.005Power, W 0.35Size, cm 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.

38

Table 4.9: Inertial Measurement Unit

CharacteristicWeight, kg 0.7Size, cm Diameter = 8.9 cm, Height = 8.5 cmPower, W 10Operating Temp, °C -54 to +85Operating Range/ Angular acceleration ±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,

39

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:

eathspp gI

mImpulse Total= (4.6)

A list of assumed burns is listed in Table 4.10:

40

Table 4.10: List of RCS Thruster Burns

# Burns Duration(sec) Thrust(N)2000 1 9500,000 0.01 1.6500 3 910,000 .1 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.

Figure 4.17: Schematic of RCS propellant suppy

Heliu

Hydrazin

Valve

Thruster

41

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.

42

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.

Computer #1 Solid State Storage

Internal Data Bus

Computer #2

Main Computer

Compression ASIC

Power Management Attitude Sensors ACS Actuator ControlHigh Gain AntennaSteering

Science Instruments Telemetry

Serial Bus

High Speed Data Bus

43

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

Figure 4.19: Transfer from 20 km circular orbit to Europa’s surface

E

∆V

∆V

44

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.

45

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 1 meterWidth 65.2 cmMass 3.586 kgEmissivity 0.82Fin Efficiency 0.8Radiating Temperature near Jupiter 380 KRadiating Temperature near Venus 472 K

46

The radiator fin has an initial operating temperature of 373 K. The dimensions were

chosen from the equation:

( )44 TsTwAQ −⋅= ηεσ (4.7)

ZKHUH�4�LV�WKH�KHDW�UDGLDWHG�� �LV�WKH�ILQ�HIILFLHQF\�� �WKH�ILQ�HPLVVLYLW\��DQG� �WKH�6WHIDQ�

Boltzmann 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:

( ) 2

2

22

, mW 2598723.0

mW 1358 === RGG SVS (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:

W1467cosAGQ finV,SSun =θ⋅α= (4.9)

where QSun is the solar irradiation, α is the solar absorptivity, Afin is the area of one side

of one fin, and θ 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

( )44max TsTwAQQ pipeheatSun −⋅=+ ηεσ (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.

47

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 0.732 mDiameter 1.5 cmMass 0.088 kgLength of the Evaporator Section 0.12 mLength of the Condensing Section 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 condensingend

Length = 0.12 m; Di = 1.5 cm, Do =3.5 cm

Dimensions surrounding radiator evaporatorend

Length = 0.1m; Height = 0.652m;Widthin = 3cm; Widthout = 5cm

Volume of system 0.00140 m3

Mass of system (with aluminum housing shell) 1.42 Kg

48

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 Height = 0.486 m; Di = 7.5 cm; Do =11.5 cm

Dimensions surrounding the heat pipeevaporator end

Length = 0.105 m; Di = 1.5 cm; Do= 3.5 cm

Volume of system 0.00298 m3

Mass of system (includes aluminum housingshell)

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

49

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

50

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

51

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

Transducer/Therm-couples/Optical viewingwindow

Camera

Capillary Lamp

Capillary Plate

Computer &

Storage Board

High Voltage

Converter

Capillary ElectrophoresisBuffer Tanks

52

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.

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

Voltage Source

Detector

Capillary

Reservoirs

Sample

Injection Port

Electrode Electrode

53

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 1µA.

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.

54

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

Buffer Syringe Pump

CCD Detector

UV & Visual Lamps

Capillary High Voltage

Electronics

Capillary Plate

Buffer Storage Tanks Sample Intake Pump

55

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

56

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

57

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 sub-

surface 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.

58

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

59

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�&����7KH�SUREH�ZDOO�LV�WKHUPDOO\�VHSDUDWHG

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:

radradradgengen QQQQQ ,2shell thearound 1,bottom at the 1,2,1, ++=+ (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 fly-

by 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:

( ) ( ) ( ) ( )111

2

111

2mW 2250

2

4422

1

441144

12

12

−+−

+−+

−+−=

j

j

j

jS

TTrhTTrhTTr

εεσπ

εεσπ

πσε (5.2)

ZKHUH� �LV�6WHIDQ�%ROW]PDQQ¶V�FRQVWDQW�� 1�� 2, and j 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

60

jacket was found to be (from the Orbiter Thermal Control section) 487 K. Equation 5.2

may be solved by solving separately

( ) ( )111

2

1

441144

12

11, −+−

+−=j

jSgen

TTrhTTrQ

εεσπ

πσε (5.3)

( )111

2

2

4422

2, −+−

=j

jgen

TTrhQ

εεσπ

(5.4)

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 487 KTemperature of the bottom section 689 KTemperature of the middle section 536 KHeat radiated out the bottom 203 WHeat transferred to the jacket 2047 WHeat to be carried by heat pipes 1024 W

Table 5.2: Conditions with no solar heating

Temperature of the inner wall of the jacket 395 KTemperature of the bottom section 665 KTemperature of the middle section 477 KHeat radiated out the bottom 176 WHeat transferred to the jacket 2014 WHeat to be carried by heat pipes 1007 W

61

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�&��KRZHYHU�LWV�XSSHU�RSHUDWLQJ�WHPSHUDWXUH�OLPLW�LV�ZHOO�EHORZ�WKDW�GXH�WR

shrinkage occurring around 500 to 600�&���,W�ZLOO�EH�VKRZQ�WKDW�WKH�RSHUDWLQJ

temperature ranges around the hot end of the cylinder disk are of the order of 950�&���,W

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.

62

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.

Probe titanium wall

High temperature aerogel

High temperature

aerogel

GPHS Blocks

Copper conducting disk

Copper conducting disk

AMTEC CellsPyroceram

AMTEC cell

MLI

MLI

63

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

( ) cells 7RTTQ coldhotgen −=(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:

celli RRR

711

cells 7

== ∑ (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.

( ) convgenwallwallcoldout QQRTTQ −=−= (5.7)

( )insinswall AkLR ⋅= (5.8)

insins rtA π2= (5.9)

R pyroceramQ generated

Q converted

Qout

Rcell

Thot

Tcold

Twall

R pyroceramQ generated

Q converted

Qout

Rcell

Thot

Tcold

Twall

64

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�������&�DQG

hot end temperature of 450�&�DQG�LV�OLPLWHG�E\�PDWHULDO�OLPLWV���7KH�FHOOV�FKRVHQ�DUH�����

cm in diameter and 10.8 cm in length. There optimum operating temperatures are at a

cold end of 350�&�DQG�D�KRW�HQG�RI�����&���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�&�GXULQJ�LWV�KRWWHVW�PRPHQW���7KH�UHVLVWDQFH�RI�WKH�ZDOO�LQVXODWLRQ�ZDV

desired to be such that it would allow the cold end to be above 150�&��EXW�QRW�OHW�WKH�KRW

end to be above 1000�&���7KLV�UHVXOWV�LQ�UHTXLULQJ�D�WKHUPDO�UHVLVWDQFH�RI�WKH�LQVXODWLRQ

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 2600 Kg/m3Thermal Conductivity 4 W/(m K)Width 9mm

65

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 250 WWall temperature 476.5 KCold End Temperature 403.5° CHot End Temperature 939° CApproximate Electrical Watts produced 62.5 W

Table 5.5:Worst case scenario; internal temperatures near Venus

Heat generated 250 WWall temperature 527.5 K

Cold End Temperature 454.5° CHot End Temperature 990° C

Approximate Electrical Watts produced 0 W

Table 5.6: Internal temperatures at EOM

Heat generated 222 WWall temperature 265 KCold End Temperature 181.5Û�&Hot End Temperature 657Û�&Approximate Electrical Watts produced (assumingan 80% efficiency of the cells)

44 W

A finite element thermal analysis will be required for more accurate predictions.

66

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 Description Components Power1 Surface Communications Modem, Computer 6.66W2 Science Data Acquisition CE, Camera,

Pressure &Temperature, Lamp

from capacitor

16.55W

3 PASS Sequencing PASS Actuators,Computer

25W

4 Scientific Data Acquisition withContinuous Lighting

CE, Camera,Pressure &

Temperature,Directly powered

lamp

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.

67

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.

PASS II

PASS I

Acoustic Modem

Transducer

68

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.

Receiving Signals

Filter and Parse

Send to Computer

Reset Timer

Activate PASS

Calculate Distance

Begin/End timing

If signal =depth signal

If end oftiming

If Distance

> 30 m

69

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

Blades Transducer Mount

Actuators

Actuator & tether

mounting plate

70

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.

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

Spectra 1000 Cable Spool

Incendiary BoltsIncendiary Bolts

Hollow Core

71

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

72

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

73

.

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) (5.10)

74

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.

( )( )at

LK

L

tEK

t

aPp

pe22

2

2104.1

,112

νν

π −=

−= (5.13)

For this formula to be valid,

75

t

a

a

L ≥

2

(5.14)

where L is the length, a is the radius, Pe is the pressure, E is the modulus of elasticity, ν 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:

( )[ ] ( )

( )( )

( )[ ]2222

4

2

2

2

222

112 nnn

n

Et

aP

La

La

ah

La

e

++

−+

=π

ππ

ν (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

0.06

2.01

3.28

2.88

2.52

2.242.03

1.861.72

1.61

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

0 2 4 6 8 10 12

Buckling Mode(n)

Th

ickn

ess

(mm

)

Figure 5.13: Plot of necessary skin thickness versus buckling mode

76

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.

77

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 forDDFI&T

Cost(FY98$M)

Project Management 11.3Mission Director and Operations Teams 18Flight Systems 60Ice/water Probe 40Lander Instruments and Scienceexperiments

16

Ground Data and Mission Operations 11.5Reserves and others 25Total 181.8

Table <cost 2>: Estimated launch vehicle cost

Launch Vehicle Cost(FY98$M)

Delta 3 70

Table <cost 3>: Estimated flight operations component costs

Major Components forFlight Operations

Cost per year(FY98$M)

Management 1.5Experiment Team 2Engineering Team 3Ground Data System support 1Reserves 1TOTAL (per year) 8.5

Table <cost 4>: Estimated data retrieval component costs

Major Components forData Retrieval

Cost per year(FY98$M)

Management 1.5

78

Experiment Team 3Engineering Team 0.3Ground Data System support 1Reserves 1TOTAL (per year) 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.

79

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.

80

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.

81

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

82

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

83

∆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.

84

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.

85

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.148-

159.

• 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

86

• Desai, P.N., et.al, 1993, Aspects of Parking Orbit Selection in a Manned Mars

Mission, NASA Langley Technical Report Abstract.

• Dyson, Freeman, J., 1995, 21st Century Spacecraft, Scientific American, September

1995, Vol.273, No.3.

• 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

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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

89

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 87.4 yearsDensity of PuO2 11.46 g/cm3

Specific power of PuO2 490 W/kgAlpha energy emitted per decay 5.47 MeVGamma energy emitted per decay 0.01 MeV

(Properties are taken from (Corliss & Harvey, Appendix 1. Also available are the properties of other compounds ofPu238 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.

90

Figure 11.1: General Purpose Heat Source Module

Figure 11.2: General Purpose Heat Source Module Sectioned at Midplane

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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 semi-

conducting 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

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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.

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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 1.2µm. 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.

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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:

ρ

κ

=

=⋅

= ×

=⋅ ⋅

=

−

917

1925

147 10

2 6

284

3

62

kg

m

CpJ

kg K

m

J

s m KkJ

kg

.sec

.Κ

h

Using this information and the following equation:

( ) ( ) h×××+××+×+×××

×

×+×Κ××∆=

×+×

ρππρ

π κ

URoRiURoRiCp

eRoRiTHeatRoRi

U

22

2

2

)(4(11.1)

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

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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.

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:

µe = 3.203E12 m3/s2 Me = 4.8E22 kg Re = 1569 km G = 6.67259E-11

Nm2/kg2

Orbital period ,ω =

3.06822E5 sec

Rapoapsis = 1589 km Rperiapsis = 1569 km µe = GMe

The velocity of the 20km circular orbit is found using:

smr

v ee /73.1419==

µ (11.2)

The velocity at the apogee of the orbit is found by:

9

2

10249.2cos1

×=∴+

= he

h

ra υµ

(11.3)

Ri

Ro

Probe Wall

Ice wall

Melting Region

96

006331.0=+−

=pa

pa

rr

rre (11.4)

2/26.1415 skmvvrh aaa =∴= (11.5)

smvvv ac /45.41 =−=∆ (11.6)

The velocity at the perigee of the orbit is found by:

skmr

hv

pp /43.1== (11.7)

Europa surface velocity is 0.032 km/s.

skmvvv ep /43.1222 =+=∆∴ (11.8)

therefore the total ∆v for the entire transfer is:

skmvvvtot /44.121 =∆+∆=∆ (11.9)

Appendix 11.5 Europan Environment Statistics

Discovery: Jan 7, 1610 by Galileo GalileiDiameter (km): 3,138 kmMass (kg): 4.8 (1022) kgMass (Earth = 1) 0.0083021Surface Gravity (Earth = 1): 0.135Mean Distance from Jupiter (km): 670,900 kmMean Distance From Jupiter (Rj): 9.5Mean Distance from Sun (AU): 5.203Orbital period (days): 3.551181 daysRotational period (days): 3.551181 daysDensity (gm/cm3) 3.01 gm/cm3Orbit Eccentricity: 0.009Orbit Inclination (degrees): 0.470°Orbit Speed (km/s): 13.74 km/sEscape velocity (km/s): 2.02 km/sVisual Albedo: 0.64Surface Composition: Water Ice