1. SENIOR DESIGN PROPOSAL
The Richter Program’s “Europa CT Scanning” RFP
Prepared by
Michael Corpuz (Team Lead)
Randall C. Acosta (Deputy)
Omar Alhassen
Matt Bergman
Brendan J. Clarke
Frank Garcia
Kasbar Gulbenli
Jeremiah Kho
Sean Matthews
Juan Sanchez
Department of Aerospace Engineering California State Polytechnic University, Pomona, CA, 91768

2. Abstract
This paper outlines how the Kronus team satisfied NASA’s RFP for a “Seismometer
Array and Delivery System Capable of Collecting Seismographic Data Sufficient to Map the
Interiors of Jupiter’s Moon II Europa” and all its requirements. After extensive design and
trade studies, a design of a solar powered, dual-spin stabilized, and liquid bi-prop orbiter
carrying eight solar-powered landers was chosen. The spacecraft will launch on a SpaceX
Falcon Heavy in 2020 and arrive to Europa in 2026 by using a Venus Earth Gravity assist
and a Jovian Satellite Tour. The landers will be placed in a Legendre-Gauss-Lobatto point
distribution and collect seismographic and camera data. The orbiter will transport the
landers to Europa as well as relay all scientific and engineering data from the landers to
Earth. Through examination of all requirements, the proposed design is compliant with all
restraints and requirements and is fully capable of completing the RFP’s mission.
I. Introduction
The official title of the Request for Proposal given to the design team by Dr. Stephen Edberg of NASA’s
Jet Propulsion Laboratory is: “Seismometer Array and Delivery System Capable of Collecting Seismographic Data
Sufficient to Map the Interiors of Jupiter’s Moon II Europa.” Europa is one of Jupiter’s Galilean moons and there is
much speculation that Europa may be able to support life in its large subsurface oceans. However, a mission to
Europa presents a multitude of challenges. Due to Europa’s distance from the sun, the spacecraft will have to deal
with low solar fluxes as well as cold temperatures. In addition, the large doses of radiation and gravitational torques
from Jupiter and the unknown topography of Europa’s surface are factors to take into account as well. This mission
is therefore classified as a NASA flagship mission, due to its scope and scale. The primary goals of this mission are:
to strategically place seismometer array on the surface of Europa that is able to record and read any seismic activity
that may occur due to the subsurface ocean of Europa, expand knowledge and understanding of interior composition
and structure of Europa, and finally demonstrate capacity for inter-planetary exploration. The primary requirements
derived from the RFP revolve around the seismometer and camera payloads as well as a specific landing layout. The

3. main requirements other than the payload and landing sequence, is to have a minimum 90 days of seismic and
imaging data on Europa and arrive at Europa by 2026. The full list of requirements can be found in the Appendix.
This paper outlines how the Kronus group will satisfy this RFP and all its requirements.
II. Mission Design for a Europa Orbiter
The mission design was broken into three phases: 1) the trajectory from Earth to Jupiter, 2) the tour from
Jupiter to Europa, and 3) the desired Europa orbit characteristics. Each mission phase conducted their own trade
studies in order to optimize a low delta-V (ΔV), low time-of-flight (ToF), and maximize the scientific benefit. A low
ΔV was important in order to reduce the spacecraft’s wet mass; a low ToF was necessary in order to satisfy the
Request for Proposal (RFP) stated Europa landing date of 2026.
Table 2.1: Comparison of VEGA and VEEGA Trajectories to Jupiter
Figure 2.1: Comparison of Venus, Earth, and Mars Gravity Assist Trajectories

4. After doing a global search of Earth, Venus, and Mars gravity-assist trajectories with JAQAR’s Swing-by
Calculator, it was determined that Venus gravity assists would provide the lowest ΔV and therefore lowest
spacecraft wet mass. However, a direct transfer from Venus to Jupiter is highly inefficient; one, or two, Earth
gravity-assists were sought in order to reduce the Jupiter arrival velocity. Both the single-Earth flyby (VEGA) and
double-Earth flyby (VEEGA) trajectories are shown in Table 2.1. The VEGA’s time of flight from Earth to Jupiter
is 4.3 years, however the second Earth gravity assist (VEEGA) requires a time of flight of 5.8 years—far too long in
order to satisfy the RFP requirements of a 2026 Europa arrival date. Therefore, the chosen trajectory for this mission
was the VEGA, with a total mission ΔV of 2952 m/s.
Figure 2.2: Earth-to-Venus Pork Chop Plot
After determining that an initial Venus flyby would be optimal for a trajectory to Jupiter, a porkchop plot
was generated in MATLAB. This plot, shown in Figure 2.2, allowed for the extraction of a one-month launch
window in March 2020, as well as the respective launch vehicle payload mass. For this mission, only two launch

5. vehicles were considered, the SpaceX Falcon Heavy and United Launch Alliance’s Delta IV Heavy. With a
maximum launch characteristic energy (C3) of 13 km2
s-2
, the Falcon Heavy provided a launch mass of 13,500 kg
while the Delta IV Heavy provided 7,920 kg. Due to the high mission ΔV of the VEGA, the Falcon Heavy was
considered, rather than having to use 2 Delta IV Heavy launches. Although the VEEGA trajectory does not satisfy
the RFP date requirement, the reduced total ΔV does allow the same spacecraft to launch on the Delta IV Heavy,
due to the significant reduction in propellant required.
Table 2.2: Critical Dates for 2020 Venus-Earth Gravity Assist Trajectory
Critical dates for the Venus-Earth gravity assist trajectory are shown in Table 2.2. The earliest launch date,
starting at the one-month launch window, is on February 27, 2020 with a launch characteristic energy of 13 km2
s-2
.
As the days in the launch window progresses, the characteristic energy reduces until an optimal launch day on
March 18, 2020. Assuming a one-month launch window, the latest launch possible would be on March 26, 2020,
with a characteristic energy of 13 km2
s-2
. After launch, the first flyby encounter is at Venus on July 1, 2020, at an
altitude of 22,000 km. This maneuver is energy increasing: the arrival velocity is 6.38 km/s and the departure
velocity is 6.41 km/s. The second flyby encounter is at Earth on April 28, 2021, at an altitude of 1,300 km. Once

6. again, this is an energy increasing maneuver, with an increase in velocity of nearly 0.6 km/s. After a 1,140 day
transfer from Earth, the spacecraft will arrive to Jupiter on June 11, 2024, with an arrival velocity of 6.40 km/s.
The next phase was to determine the trajectory from Jupiter to Europa. Two tours were investigated: the
Banzai Pipeline, a low-radiation dose tour, and the 12-L1, a low- ΔV tour. A comparison of these two tours is shown
in Table 2.3.
Table 2.3: Comparison Between Jovian Satellite Tours
The main consideration of the tour was the total radiation dose accumulated. Due to the high radiation
environment of Jupiter, the spacecraft must be outside the region of Ganymede during the tour in order to avoid
excessive radiation. With an already challenging 90-day mission at Europa, any additional radiation dose will further
increase the radiation shielding necessary for the spacecraft. In comparison, the 12-L1 tour had a radiation dose of
124 krad, but the Banzai Pipeline only accumulated 89 krad. The additional ΔV required for the Banzai Pipeline was
worth it due to the significant decrease in radiation shielding required for all 8 landers.
Additional considerations of the tours included the time of flight (due to the RFP requirement), the number
of satellite flybys, the lowest flyby altitude, and the time of flight between each satellite encounter. Due to the
navigational challenges of a satellite tour, the risk assessment of both trajectories were considered. On average, the
Banzai Pipeline had higher altitude flybys than the 12-L1, as well as less critical flybys (encounters with a satellite
less than 500 km). Therefore, navigationally, the Banzai Pipeline was the preferred satellite tour.

7. One of the most challenging maneuvers of the mission is the Europa Orbit Insertion on February 1, 2026.
Due to the satellite encounter time of flight of less than 3 days, an autonomous orbit insertion burn may be necessary
for mission success. The autonomous navigation phase of the satellite tour is shown in Table 2.4. An alternative
solution this would be ground-based navigation, which would only be able to utilize 1 maneuver per satellite
encounter. In order to increase both the navigational accuracy, as well as the number of maneuvers between flybys,
autonomous navigation would be necessary.
Table 2.4: Critical Dates for Banzai Pipeline Tour
The last phase of the mission was determining the optimal science orbit around Europa, as well as the ΔV
required for a plane change. The RFP requires polar landers, therefore a polar orbit would allow easy access to most
landing sites on Europa. As an additional benefit, a polar orbit allows for global mapping coverage of Europa, which
can be used to seek safe landing site zones with a high-resolution camera. Because the orbiter is using solar arrays, a
special type of polar orbit, the full-sun orbit, allows for the spacecraft’s solar arrays to be pointing nearly directly

8. toward the Sun. Due to the small angular distance of the Earth and Sun, the spacecraft’s high gain antenna will also
be able to constantly communication with the Earth, except during the Jupiter eclipse.
The ΔV allocation for the entire mission in shown in Table 2.5. Trajectory correction maneuvers were
accounted for from Earth to Jupiter; these included statistical low- ΔV maneuvers between flybys, a Jupiter-arrival
trajectory correction, as well as a worst-case launch trajectory error correction. The 100 m/s ΔV for the deep space
maneuver is the worst-case maneuver, which would only be encountered if launched near the first or last days of the
launch window. Because of the optimal trajectory, if launched on March 18, 2020, there would be no deep space
maneuver necessary. A significant reduction it he Jupiter Orbit Insertion was acquired by preforming an initial 500-
km Ganymede flyby, 3 hours before the JOI burn. The JOI is the required ΔV to be captured into a highly elliptic,
15 RJ by 242.5 RJ, Jupiter orbit. This specific orbit sets the spacecraft up properly for the Banzai Pipeline trajectory.
After a perijove raise maneuver to correct for orbital perturbation, the spacecraft begins the tour with a Ganymede
flyby 95 days after the raise maneuver. Orbital trim maneuvers, which are deterministic, and statistical maneuvers (2
m/s per satellite encounter) were accounted for during the Banzai Pipeline. Lastly, the large Europa Orbit Insertion
inserts the spacecraft into a circular, polar orbit around Europa at an altitude of 100 km. 120 days of orbital
maintained ΔV was accounted for to keep the spacecraft in the proper orbit. Without this orbit maintenance, the
spacecraft’s orbit would eventually degrade into Europa’s surface in about a month.
Table 2.5: Total Europa Mission ΔV

9. III. Radiation Effects on a Europa Orbiter and Lander:
There are two branches of radiation: non-ionizing and ionizing. Non-ionizing radiation causes damage in
material by the production of heat (vibration) or atomic displacement, while ionizing radiation causes
malfunctioning of electronic devices, especially semiconductors. While passive electrical components (resistors,
capacitors, and inductors) are relatively immune to radiation damage, active devices, such as computer systems,
have four main categories of damage: non-ionizing thermal damage, displacement damage, total ionization dose
damage, and single event upsets. For the preliminary analysis of a spacecraft mission to Europa, both the
displacement damage and the total ionization dose damage was accounted for by considering the effects of non-
ionizing and ionizing radiation, respectively, in the Jupiter radiation environment.
When nonionizing radiation interacts with an atomic nuclei, it has a probability of displacing, or removing,
them from their lattice sites. This displacement damage will ultimately cause a reduction in the lifetime of
semiconductors, and therefore is important for solar cell power attenuation. For analytical purposes, it is a common
standard to express the damage effectiveness of a particle’s energy by using the unit equivalent 1-MeV particle
fluence.
Figure 3.1: 120 Day Particle Fluence at Europa

10. The accumulation of ionizing radiation for the lifetime of a mission is called the total ionization dose, or
TID. The ultimate outcome of extensive ionizing radiation on semiconductors is the decrease in functionality.
Eventually, the device will have a high probability of failure after a specified radiation rating. The lifetime of
electrical devices can be increased by considering the use of radiation-hardened components, which in general will
be able to accumulate ten to fifty times the radiation dose of their equivalent commercial parts. For a flagship
mission to Europa, it was assumed that the majority of the electronics would be rated for a TID of 100 krad (Si).
For radiation analysis of this interplanetary mission, ESA’s Space Environment Information System, or
SPENVIS, was utilized. Modeling the radiation damage of the spacecraft was a three-step process. First, the external
environment at Jupiter and Europa was modeled. Second, a program was chosen. Third, the environment of the
spacecraft was considered.
Table 3.1: Orbiter and Lander Solar Cell Cover Glass
SPENVIS contains an implemented Jupiter radiation environment package, JOREM (Jupiter Radiation
Environment and Effects Models and Mitigation). This package contains JOSE (Jovian Specification Environment),
a model for the particle environment around Jupiter. In order to access this package, the spacecraft’s reference planet
must be changed to Jupiter. Next, the coordinate generator must be implemented in order to determine the state of
the spacecraft during its mission, in reference to a Jupiter-centric coordinate system. In this case, only the 120-day
mapping and science phase at Europa was considered. Therefore, the state of the spacecraft was set at a 100 km orbit
around Europa, which is at a perijove altitude of 664,792 km and apojove altitude of 677,408 km.
Cover glass
(mils)
120 Day
Fluence (rad)
Power
Attenuation
42 m2
Mass (kg)
120 Day
Fluence (rad)
Power
Attenuation
17.1 m2
Mass (kg)
3 mils 9.44E+15 28% 8.2 9.33E+15 28% 3.3
6 mils 2.01E+15 18% 16.3 1.91E+15 18% 6.6
12 mils 4.86E+14 12% 32.6 3.91E+14 11% 13.3
20 mils 2.09E+14 9% 54.4 1.26E+14 8% 22.2
30 mils 1.40E+14 8% 81.6 6.72E+13 6% 33.2

11. For solar cell radiation degradation, the EQFLUX program was used. There were two cell types considered:
single and multiple junction. Because both the orbiter and lander used triple-junction cells, only the multiple
junction type was considered. EQFLUX also offers a variety of manufacturers for their solar cell types, including
Spectrolab,
AZUR, and TECSTAR. For this analysis, only Spectrolab as considered. The results are shown in Figure
3.1 and Table 3.1. The power attenuation values are derived from a best-fit curve based on the radiation degradation
parameters in the Spectrolab data sheet [1].
As expected, as cover glass thickness increases, the power attenuation decreases. However, a plateau effect
is present where the power attenuation does not decrease significantly after approximately 20 mils of cover glass,
but the mass of the cover glass still increases linearly. For this mission, it was critical to reduce the total mass of the
spacecraft while maintaining the proper radiation shielding. Therefore, it was determined that the optimal cover
glass thickness ranged between 6 and 12 mils, thus remaining low-mass at the expense of a 12 to 18% power
attenuation at the end-of-mission.
Figure 3.2: Total Radiation Dose Compared To Five Shielding Materials

12. In order to calculate the shielding radiation dose, the program SHIELDOSE-2Q was used. This program
used the Jovian trapped particle models to estimate the doses behind tantalum, aluminum, titanium, iron, and
copper-tungsten shielding. The targeted material was silicon, which is the material most electronics are made with.
Figure 3.2 shows the result of the SHIELDOSE-2Q program. Tantalum and copper-tungsten are both great radiation
shielding material, however they are expensive to manufacture and are dense. Although aluminum and titanium do
not shield radiation as effective as tantalum or tungsten, they were chosen as the orbiter and lander radiation vault
materials due to their low density and great structural properties.
Table 3.2 provides the complete material-mass relation between all of the five materials implemented in
SHIELDOSE. In general, aluminum and titanium required a shielding thickness about 3 to 4 times the thickness of
tantalum and copper-tungsten. However, due to the significant cost reduction, aluminum and titanium vaults were
still decided upon.
Table 3.2: Orbiter and Lander Radiation Vault Masses
Figures 3.3 and 3.4 both show the radiation accumulation effect of the landers staying up in orbit for 1
month, rather than doing an immediate autonomous landing on Europa. Two important conclusions can be derived
from these figures. First, the orbiter around Europa will receive significantly more radiation than the lander. The
total radiation dose of the orbiter after 120 days is 1560 krad, and therefore the orbiter’s radiation design factor is
3120 krad. The lander received only 567 krad after 30 days in orbit, and 90 days on the surface, and therefore has a
radiation design factor of 1134 krad. The orbiter will receive 3 times the radiation dose of the lander due to the
radiation protection the surface of Europa provides.

13. Figure 3.3: Total Radiation Dose Accumulated Over 120-Day Mission
The one-month mapping phase before landing puts a significant toll on the lander’s radiation
shielding. One month in a 100 km orbit around Europa is equivalent to four months on the surface of Europa. Due to
this mapping phase, the radiation shielding required by the landers nearly doubled in thickness. Figure 3.5 shows the
comparison of the total ionizing dose (TID) acquired during this Europa mission compared to previously studied
Europa missions.
In order to ensure the accuracy of the estimates from SPENVIS, a similar radiation analysis was done using
boundary conditions from the Europa Explorer and Europa Lander Mission. The Europa Explorer team estimated a
radiation dose of 1.4 Mrad (Si) behind 100 mils of aluminum shielding after a nominal 120 mission. In comparison,
SPENVIS estimated a radiation dose of 1.5 Mrad for the same mission, therefore the percent error was 7.2%. In
comparison to the Europa Lander mission, SPENVIS’ estimates for the orbiter’s total ionizing dose was 1.35%,
while the lander total ionizing dose was 1.65%. Because the radiation shielding accounted for a 100% margin, all of
these errors are well within the accounted margin for this mission.

14. Figure 3.4: Aluminum Shielding Thickness vs. Radiation Dose Accumulated
Figure 3.5: Comparison of Total Radiation Dose of Europa Missions

15. IV Mapping and Landing Phases
A. Mapping Phase
It is critical to mission success to be able to map the eight landing sites. This allows the mission operations
team to select the areas with the most solar illumination and the least amount of sloped terrain. Most of the landers
are landing near the poles, so the solar incidence angle near there is almost 90 degrees. This means that any surface
extrusion could cast a shadow over the lander, limiting its solar coverage. This high incidence angle is also an
advantage because it is easier to see hazardous topology because of the long shadows that it will cast. With this
information, the lander can avoid valleys, steep cliffs, and rough terrain that could be mission ending otherwise.
Mapping the surface also provides unique benefits. If the surface structure of Europa is better known, it
would allow scientists to more accurately update Europa’s characteristics including its shape (volume), albedo,
rotation rate, mass, and gravity values. The landing site images can be analyzed to determine seismologically and
scientifically interesting landing sites. It also allows the flight dynamics group to update their landing trajectory
before landing. These mapping photos can also be used to determine a very accurate location of the lander after
touchdown by correlating the high resolution images of the landing site that were taken by the orbiter with the
panoramic photos that will be taken by the lander.
The orbiter will be inserting into a 100 km altitude, near-polar, near-circular, full-sun orbit around Europa.
This orbit was chosen because it gives the orbiter global coverage of Europa, including the poles, and allows for the
maximum amount of sun to reach the solar arrays. For the first 30 days at Europa, the orbiter will be mapping the
surface. Figure 4.1 shows the behavior of the satellite during this time. The orbiter will be using two cameras to
map the surface of Europa, a Wide Angle Camera (WAC) and a Narrow Angle Camera (NAC). For the first 8 days,
the WAC will provide a global coverage mosaic of the surface at a resolution of 150 m/pixel. This includes a 20%
overlap per image to stitch them together. This global coverage mosaic will allow for a global characterization of
landforms and a general evaluation of the landing sites. As can be seen from Figure 4.2, the WAC swath covers the
entire surface after the 8 days. In this image, the red lines are the border of the coverage swath and the colored
circular areas are the landing areas.
Once this period of 8 days is over, the WAC will stop mapping and the NAC will begin to provide high
resolution image mosaics of the landing site. For the next 22 days, the NAC will be collecting image mosaic strips
at a resolution of 1 m/pixel through the pre-determined landing area. These strips can be seen in Figure 4.2 filling in

16. the landing areas. The two landers near the equator have the worst coverage, so these are the areas that are designed
to. Figure 4.3 shows a zoomed in image of the white area to see more clearly. The landing area is defined in the
RFP to be a circle of radius 137 km. Figure 4.3 shows this area and the orbiter ground tracks that are created
through that area over the period of 22 days. With this worst case scenario, the orbiter passes over that area 16 times
with a mean coverage duration of 2 minutes and a total coverage duration of 43 minutes. The orbiter collects these
strips of high resolution image matrices that are shown as white lines in Figure 4.3. Each image from the NAC
covers an area of 250 km2
and these are stitched together with a 20% image overlap. The lander areas near the poles
will have much better coverage than the one shown in Figure 4.3. With the data rate that is available to the orbiter,
the NAC is able to collect an area of 7,620 km2
per landing site within the 22 days.
All of this image data will be analyzed to select the best landing sites for each of the 8 landers. Landing
sites will be selected by weighing the technical feasibility of the landing site against the seismologically and
scientific desirability of the site.
Figure 4.1: The NAC (blue) and WAC (red) surface swath coverage by the end of the mapping phase

17. Figure 4.2: The WAC coverage swath after 8 days
Figure 4.3: Worst case NAC coverage of the landing sites

18. B. Landing Sequence
The landing sequence is also a critical part of the mission. Two architectures were analyzed for the best
and most reliable method to land on the surface.
The first method that was considered is shown in Figure 4.4. At point 1 in this figure, the orbiter inserts
into a 10 km, full-sun, polar, near-circular orbit. This orbit is harder to achieve than the 100 km orbit in architecture
2 because it is so close to the surface, so the burn needs to be very accurate. Once the orbiter is in this orbit, it will
begin to deploy the landers. The landers will then perform a deorbit burn with a solid rocket motor that will provide
a delta-V of 1.432 km/s that will cancel all of the horizontal velocity. Point 2 in Figure 4.4 shows where this burn
takes place. The behavior of the lander during this burn was analyzed in MatLab. The results of this analysis are
shown in Figure 4.5. This is a very short burn (about 12 seconds), which means that the load on the spacecraft is
very high. This high load is a concern for the sensitive seismographic equipment on board. As shown in Figure 4.5,
the flight path angle changes rapidly from zero degrees to a vertical free-fall. The altitude also decreases about 90
meters during this time. Once all of the horizontal velocity is cancelled, the ACS thrusters are then used to cancel
the remaining vertical velocity. The details of this burn are shown in Figure 4.6. This burn lasts 109 seconds and
has a relatively linear deceleration. The burn does not start until the altitude of the lander above the surface is about
5,500 meters, then the thrust stays between 120 N and 100 N. Once the lander is on the surface, the orbiter must
raise its altitude to 100 km because the 10 km altitude is very unstable. This is performed by a simple Hohmann
transfer.
Figure 4.4: Architecture 1 landing sequence

19. Figure 4.5: Architecture 1 deorbit burn data
Figure 4.6: Architecture 1 gravity turn burn data
The second architecture that was considered is shown in Figure 4.7. In this architecture, the orbiter is
initially in a 100 km, polar, full sun, near-circular orbit. The trajectory of each of the landers is optimized to use the

20. least amount of fuel, as well as to reach the surface in the least amount of time to reduce the required battery mass.
The landing sequence occurs in three sections that can be seen in Figure 4.7: deorbit burn, coast, and gravity turn
burn. When the lander is 342.9 km away from the landing site on the surface, the deorbit burn will begin. The
deorbit burn performs a delta-V of 800 m/s over a 102 second duration. The details of the lander trajectory and
behavior during this burn can be seen in Figure 4.8. This burn is performed at a constant max thrust and it takes the
lander from an altitude of 100 km to 79.3 km and a flight path angle of 0° to -15.2°. As can be seen in Figure 4.8,
there is also a relatively constant deceleration and the load factor in Earth G’s stays below 4. During this burn, the
lander will travel a total distance along the surface of 208.8 km.
Figure 4.7: Architecture 2 landing sequence
Once this burn is complete, the lander will shut off its engines and coast for 84 seconds. This lowers the altitude
from 79.3 km to 62.1 km. The lander travels 51.2 km along the surface during this period. The lander will also be
preparing for the next critical gravity turn burn.
The gravity turn burn will then place the lander on the surface. At 338 seconds, this is the longest burn and
it will provide the remaining delta-v required for the lander to reach near-zero velocity at the surface. The detailed
trajectory and behavior of the lander during this burn can be seen in Figure 4.7. During this burn, the lander is
lowered to the surface from an altitude of 62.1 km and the lander travels 82.9 km along the surface. The flight path
angle is also lowered from -22° to -90°, so the lander will land upright on its legs. This flight path angle also

21. Figure 4.8: Architecture 2 deorbit burn data
Figure 4.9: Architecture 2 gravity turn burn data
ensures that the lander is always pointed towards the landing site so the ACS cameras have a constant view of the
destination (More details on the ACS during this period in the ACS report section). This is also a variable thrust
burn that is designed to reduce the propellant mass required and maintain a relatively constant deceleration to keep

22. the load factor on the spacecraft down. The thrust curve shown in Figure 4.7 is the desired ideal thrust. This thrust
is accomplished by pulsing the engines to obtain that ideal thrust over time. The load factor during this burn stays
below 2.5 Earth G’s, which is very reasonable.
This second architecture was chosen for the landing sequence for several reasons. First, the short burn time
of the solid rocket motor is very risky because there is no time to adjust for errors in the trajectory. Second, the high
loads experienced during this burn cause the structure of the lander to be very heavy and this almost negates the
benefit for using this architecture. This high load can also damage the seismographic equipment onboard.
The trajectory of the landing sequence can also be optimized further by going through a constrained
trajectory optimization. The software package “DIDO” was used in this research to employ the Legendre
Pseudospectral Method for optimization. An example of vertical descent was used to simplify the problem as well
as prove the viability of DIDO. This program calculates the state of the lander at discrete nodes. For this problem,
30 nodes were chosen to be distributed on a Legendre-Gauss-Lobatto spacing distribution. This spacing was chosen
to have higher accuracy in the results. The equations of motion for this simple vertical descent are as follows:
Where y, v, m, k, and Tmax are altitude, velocity, mass, engine throttle (from 0 to 1), and max thrust respectively.
The cost function was selected to minimize the fuel usage. The vehicle is initialized at an altitude of 500 m and a
vertical velocity of -5 m/s. The parameters were also normalized because the optimization code runs more smoothly
and has better convergence values if the parameters are the same order of magnitude. For this example, the scaling
factors were chosen to be 500 m, 10 sec, and 1000 kg, for distance, time, and mass respectively. The minimal fuel
solution for a simple vertical descent was found and the normalized values are shown in Figure 4.10.

23. Figure 4.10: Vertical Descent Minimum Fuel Solution (Normalized Units)
Figure 4.11: Vertical Descent Minimum Fuel Solution

24. Figure 4.11 shows the un-normalized values. This figure shows that the throttle does not switch on until about 1
second. This example proves the viability of using DIDO for optimization and this same method can be applied for
optimization for a trajectory in two dimensions. However, the free version of this program does not allow the use of
enough state and control variables for this optimization, so this is recommended for future research.
The eight lander locations seen in Figure 4.12 are laid out in a Legendre-Gauss-Lobatto spacing
distribution. This spacing was chosen so that there is a higher concentration of landers near the poles than the
equator and so that there is global coverage. In Figure 4.12, the colored circles are the areas that each lander can
land in. This lander distribution also avoids the high radiation areas on the trailing edge of Europa, marked with
blue ellipses [4.1]
.
Figure 4.12: Colored circles are lander locations and blue areas are zones of high radiation on Europa’s
trailing edge
V: Spacecraft Design
The Design process began with the need for two architectures that would be able to effectively fulfill the
mission objective to land a minimum of 7 seismometers with cameras onto Europa's surface. The two designs were
to have one safe design and one radical design. The first design incorporated a dual-spin stabilized spacecraft that
consisted of: a fully solar powered orbiter and eight solar powered soft Landers. The second design incorporated a

25. three-axis stabilized spacecraft, which consisted of: an orbiter and eight soft Landers that were both powered by
RTGs.
The first design, which used dual-spin stabilization, used an orbiter as a relay for the Landers to
communicate with Earth and as a transportation bus for the eight Landers. The orbiters were modeled in Solidworks
to a fine detail including accurate measurements and placement of all systems to allow for an accurate c.g. location.
The ACS system had a non-spinning top section that includes the High Gain antenna, the command and data system,
the four low gain antennae, and two cameras. The non-spinning top section allows for constant view of Europa by
the low gain antennae and cameras. The bottom section of the spacecraft spun along its major axis so that it would
stay stable while en route to Europa. The orbiter, seen in Figure 5.1, was made out of 2024 Aluminum for its superb
strength to weight ratio and finite element analysis was used to ensure the lightest structure possible while still
handling the Load factors. The three view of the orbiter design 1 was created for its stowed and un-stowed
configuration seen in Figure 5.2 and 5.3, giving its c.g. location and inertia properties. The orbiter was stowed in the
Falcon Heavy payload fairing with a static envelope of 0.3m and a dynamic envelope of 0.2m, seen in Figure 5.4.
Design 1 was chosen to be the better architecture and its full mass summary table is seen in Figure 5.5.
Figure 5.1 Orbiter 1 with all subsystems

26. Figure 5.2 Orbiter Stowed Configuration 3-View
Figure 5.3 Orbiter Un-Stowed Configuration 3-View

27. Figure 5.4 Design 1 Stowed in Falcon Heavy Payload Fairing
Figure 5.5: Design 1 Mass Summary

28. The second design configuration also had an orbiter and eight Landers. The Orbiter was a three axis-
stabilized spacecraft that is seen in Figure 5.6. The orbiter was powered by solar arrays on three sides with the
telecom dish balancing the fourth side. The design had all Landers mounted directly on top of the orbiter, separating
the first four Landers from the second four Landers by a truss structure. The three view of the orbiter can be seen in
Figure 5.6 and Figure 5.7, showing its c.g. locations and inertia values. The design c.g. location was not centered in
the x-direction because of the difference in weight of the telecom dish to the solar array. The unsymmetrical design
made it required to change the design of the ACS system to accommodate the c.g. location. Design 2 wasn't chosen
because of its less than ideal c.g. location and the volume of the orbiter resulting in a very small static envelope.
Figure 5.6 Design 2 Orbiter three view stowed configuration

29. Figure 5.7 Design 2 Orbiter three view un-stowed configuration
VI: Lander Design
The first design's Lander is a three axis stabilized spacecraft. It has a rectangular shape for simplicity and
symmetrical placement of all its internals to give it a c.g. in the center of in the x and z axis. The Lander is solar
powered with two foldable solar arrays each with diameters of 3.3 meters. The Lander is a bi-prop design with four
cylindrical tanks to fit in the volume of the structure. The Lander has four legs for stabilization that each fold up to
increase the static envelope during stowed configuration in the payload fairing and to give the Lander a damping
during landing. The payload of the Lander is a seismometer and camera, which are placed in the corner farthest
from the solar arrays to allow for maximum visibility. The configuration of the Lander in design 1 is seen in Figure
6.1. Three views of the Lander is seen in Figure 6.2 and 6.3 with locations of the c.g. and the inertia given.

30. Figure 6.1 Lander Configuration with all subsystems
Figure 6.2 Lander Stowed Configuration 3-View

31. Figure 6.3 Lander Un-Stowed Configuration 3-view
Figure 6.4 Design 2 Lander

32. Figure 6.5 Design 2 Lander 3-View
VII: Structural Analysis
The structural analysis started by finding the largest loads applied to the orbiter and the landers. The
largest axial and lateral loads were found to be the launch loads. The launch loads were not available for the Falcon
heavy launch vehicle so the closest thing was used, the Delta IV Heavy launch vehicle loads. The launch loads were
found to be a maximum axial load of 6g and 3g lateral load. The largest axial load on the lander was found during
de-orbit to be a load factor of 4g.
Multiple tests were run: bending stress, buckling tests, shear stress, thermal analysis, and modal analysis.
Using FEMAP to model the orbiter bottom bus, top bus, and lander, they were all set up for each test. The tests
were run with the maximum loads applied by fixing the bottom section and applying the axial loads to test buckling.
The bending stresses were found by setting one side as a constraint and applying the lateral load. Each test was run

33. with multiple elements: beams, rods, and tubes, till the element that provided the best balance between strength and
mass was found. The structures were each chosen to use circular tubing for their elements.
Figure 7.1: FEMAP Buckling Analysis on Orbiter upper bus
The structural analysis was done on each of the three main structures with multiple configurations. The
truss structure found in Figure 7.1 shows the upper orbiter bus in FEMAP under buckling loads. The thicknesses of
the tubes were found by iterating the tests until the maximum load applied was underneath the yield stress of the
material. The materials used were Aluminum 2024, Steel 4340, and Ti6Al4v. The aluminum 2024 was the best for
the job but after a Northrop Grumman presentation, was changed to use Aluminum 6061-T6 because of its superb
qualities for spacecraft. The stresses from the buckling test, bending stress, and max displacements are shown in
Figure 7.2.

34. Figure 7.2: Maximum Stresses and Displacements of each structure
The modal analysis was done on the orbiter structure by creating a constraint on the bottom of the structure.
The launch loads were applied and using NX Nastran the modal analysis test was run. The first ten modes were
created by Nastran from frequencies from 0-200 Hz. The payload planner’s guide stated that frequencies above 35
Hz were the main causes of displacements on the structure.
Figure 7.3: Modal Analysis of Orbiter Structure

35. VIII. Scientific Payload
A. Seismometer Payload:
The first component that was chosen was the seismometer, and the current seismometer design is based off
the Mars Insight SEIS mission that is scheduled to launch in 2016. A trade study was done in order to determine
which seismometer would be used. This trade study is shown in Table 8.1. The reason this design was chosen was
due to the fact that the seismometer will be used in an extraterrestrial surface and because it of its current operating
specifications.
Table 8.1: Trade Study for Seismometer Selection
The Insight SEIS has relatively low mass and low power consumption. It also has 3 Very Broad Band
(VBB) sensors and 3 Short Period (SP) sensors that are able to read a seismic wave in all 3 axes. Furthermore there
needs to be a minimum of 3 seismometer stations placed within an array in order to accurately read the ice shell
thickness at Europa as well as to confirm the presence of a subsurface ocean.(8.1)
The current mission design has 8
seismometer stations placed in an array across the surface of Europa.
The Insight SEIS seismometer was chosen as a starting point for the team’s final mission design. The
Insight SEIS was modified with thermal protection in the form of Radioisotope Heating Units (RHU) to maintain a
temperature so that it can operate under Europa’s harsh conditions. The 3 Very Broad Band sensors used in this
seismometer are placed within a vacuum sphere further protecting them and isolating them from outside
disturbances that may occur from the lander movements. There are also 3 Short Period sensors that are placed

36. outside the vacuum sphere. The seismometer employs 2 electronic boxes that store the data and send it to the
lander’s computer which will then send the seismic data to the orbiter. Figure 8.1 shows the seismometer in a 3D
CAD model view with all the previously mentioned components. Figure 8.2 shows the same seismometer in a top
and side view and with dimensions (in meters). However, due to Europa’s high radiation dosages, shielding is
required for the seismometer in order for it to be operable at the surface of Europa for a 90 day mission. Thus
aluminum shielding at a rating of 2000 mil (5 cm thick) surrounds the seismometer. This thickness is essential so
that the seismometer can survive during the orbiter’s 30 day orbit trajectory around Europa and a 90 mission at the
surface of Europa. Furthermore, Table 8.2 shows the specifications of the seismometer used for this mission.
Figure 8.1: 3D CAD Model of Seismometer
Figure 8.2: Seismometer Dimensions (in meters)

37. Table 8.2 – Seismometer Specifications
Dimension Value
Mass (kg) 2.9
Power (W) 1.5
Data Rate (kbps) 12.5
Frequency Readings (Hz) 0.001 – 50
Min. Operating Temp. (°C) -220
B. Seismometer Component Details:
- Very Broad Band Sensors: There are 3 VBB sensors used in this seismometer design, which also stems
from the Insight SEIS mission. The 3 sensors together are able to read in the three P-, S-, and L-seismic
waves. The noise sensitivity of the VBB is less than 10-9
ms-2
Hz-1/2
(100-18
g2
/Hz) at a frequency of 0.001
Hz to 2 Hz.(8.2)
A CAD model design representation of the VBB sensor is shown in Figure 8.3.
Figure 8.3: Very Broad Band Sensor
- Short Period Sensors: There are also 3 SP sensors that are also inherited from the Insight SEIS design.
These sensors are placed within the seismometer such that one is horizontal to the ground and two of them
are vertical to be able to read the 3 seismic wave axes. They are also able to read a better noise sensitivity
of 10-8
ms-2
Hz-1/2
(100-15
g2
/Hz) at a frequency of 0.1 Hz to 50 Hz.(8.3)
Furthermore these are able to resist
external loads of up to 2000g in order to survive major turbulences during mission deployment.(8.4)
The
short period sensor is shown as a CAD model in Figure 8.4.
Figure 8.4: Short Period Sensor
- Electronic Boxes: There are currently 2 electronic boxes that are placed outside of the vacuum sphere of
the seismometer. These electronic boxes store the seismic data to be sent to the orbiter, and after sending

38. the data they erase the data stored so that they can make more available space in order to collect more
seismic data.
- Rotating Sphere: This was a previously proposed design feature in the seismometer design. The purpose
of this was that if the lander did not land in an even horizontal ground, then the seismometer would be able
to re-orient itself such that the seismometer may be able to accurately read any seismic waves from the
ground. However it was later decided that the best course of action would be to re-design the seismometer
boom so that the boom may be able to rotate the entire seismometer if the lander is not properly oriented to
the ground. A similar mechanism for the camera already exists for this purpose.
C. Camera Payload:
The camera for this mission requirement had to be a camera that was already space qualified and with low
mass and low power operational specifications. A trade study was also done to select the camera that best met these
requirements, and it is shown in Table 8.3. Thus the cameras used in the Beagle 2 mission were selected as they best
met the requirements for a Europa mission. Similar to the Beagle 2 mission, 2 cameras are used for this mission for
both redundancy factors and for easier data acquisition. And similar to the seismometer used for this mission, the
camera also has RHU units for thermal protection, and aluminum shielding for radiation protection. Figure 8.5
shows the camera used and the shielded package surrounding the camera. Table 8.4 further shows the camera’s
dimensions and specifications.

39. Table 8.3: Trade Study for Camera Selection

40. Figure 8.5: Shielded Camera Package
Table 8.4: Camera Specifications
Dimension Value
Mass (kg) 0.175
Power (W) 0.9
Size (mm) 79 x 63 x 75
Temperature Range (°C) -150 to 100
FOV (°) 48
Focus 1.2 m to infinity
Data Size (Mb, per picture,
compressed)
1.31
D. Scientific Payload Package Components:
The camera and seismometer have to work in accordance to each other within the deployment of a lander
on the surface of Europa. The camera package itself consists of 2 cameras placed back-to-back and surrounded by
aluminum shielding (minus the focal lens) with 2 additional RHU units. According to the requirements made by the
RFP, the camera is placed atop an extendable boom. The boom lets the cameras be able to see the local ground
above and around the lander for which the payload package is placed in. The boom uses both linear and rotary
actuators to be able to turn the camera 360° in horizontal azimuth and 90° in vertical elevation. Listen to Tame
Impala, they are an awesome band, I saw them live about two months ago. This is a random sentence no coherence
here. The boom further is supported by a separating mechanism that lowers the seismometer package (along with the
shielding) to the bottom of the ground. Thus whilst the boom pushes the camera up, a second boom mechanism
pushes the seismometer to the ground. Figure 8.6 shows the payload package design with a deployed camera boom.

41. Figure 8.6: Scientific Payload Package with Deployed Camera Boom
E. Payload Actuators:
The mechanisms to move the booms for this design included both linear and rotary actuators. The linear
actuators used for both the seismometer and camera booms were of the L16 Miniature Linear Motion series by
Firgelli Technologies.(8.5)
This linear actuator has relatively low power and is small enough in size to be used within
the inside of the booms, as is shown in Figure 8.7. Both the 100 mm and the 140 mm stroke options are used with a
mass of 74 g and 84 g, respectively, and a similar power output of 0.96 W. Four of these actuators are used in the
camera boom, whilst another 4 are used for the seismometer boom. However these actuators are only used once after
the lander is deployed on the surface of Europa. After they are used, the boom remains extended throughout the 90
mission at the surface of Europa, and thus their output is not calculated into the final average power afterwards.
Figure 8.7: L16 Series Linear Actuator

42. There are 2 types of rotary actuators used for this design; one is to rotate the camera 360° across the
horizontal azimuth, whilst 2 of them are used to tilt the camera from a level of 0° to + 90°. The type M8 rotary
incremental actuator by MOOG Schaeffer Magnetics Division, as shown in Figure 8.8, is used to rotate the camera
package across the 360° of horizontal azimuth.(8.6)
This actuator is moving the camera along 4° intervals so that the
camera may be able to take an overlapping mosaic of pictures at the surface of Europa. This rotary actuator has an
average power use of 5 W and it is constantly operating throughout the 90 day surface mission.
Figure 8.8: Type M8 Rotary Actuator
The second type of rotary actuator used is the M3-RS Rotary Smart Stage by New Scale Technologies.(8.7)
This is a smaller actuator that is able to rotate the camera payload up to a 90° tilt elevation angle, and it is shown in
Figure 8.9. There are 2 of these actuators used, however, during the mission phase of the camera they will only go
up to 24° in elevation which is due to the way the solar arrays are placed relative to the camera FOV. Thus the
camera will not be rotated to the local zenith angle since there would not be any significant scientific data at this
zone. The reason for the 90° tilt capability stems from the precaution that the lander may not land in a level surface,
and thus the camera must be able to tilt to see its local surface if the lander lands on an inclined slope. The average
power used by these rotary actuators is 7 W, and they have a mass of 150 g.
Figure 8.9: M3-RS Rotary Actuator

43. F. Damping Systems/Isolating external disturbances:
This was something that was considered and much research was done on this subject. However, the results
were not significant enough to produce concrete results. The issue stemmed from the fact that the seismometer was
placed atop of the lander and thus disturbances such as the solar arrays moving would affect the data collected by
the seismometer. To mitigate this issue, it was necessary to create a boom that allows the seismometer to be placed
directly onto the ground and thus be able to read seismic activity without any disturbances. The boom then detaches
the seismometer from the rest of the lander, and only the electronic wiring is connected to allow for data to be
transferred. Furthermore, the VBB sensors are placed within a vacuum sphere, whilst the SP sensors have their own
damping mechanisms already installed.(8.8)
Figure 8.10 shows how the boom is separating the seismometer from the
rest of the payload.
Figure 8.10: Boom Extending Seismometer to the Ground
G. Radiation Dosages:
Radiation was a particularly challenging problem, especially since the mission duration at Europa was 90
days on the ground, with an additional estimated radiation dosage accumulation of 400 krad. The most sensitive
instruments of the seismometer were the VBB sensors, and they were tested to 15 krad with good operating
results.(8.9)
Thus the radiation shielding applied to both the camera and the seismometer was done so with an ideal
rating for up to 15 krad of radiation dosage for 120 days (90 day ground mission, with an additional 30 day orbital

44. mapping mission). The entire payload of the camera and the seismometer is covered in aluminum shielding due to
aluminum being a less dense metal. Approximately 5 cm of aluminum radiation shielding (2000 mil) was used to
surround the camera and seismometer payloads. However this increased the payload mass by almost 30 kg for each
of the seismometer/camera payload packages.
H. Thermal Protection:
Another challenge in designing a scientific payload for a mission at Europa is the cold temperature
encountered. The temperature at the surface of Europa is -160 °C to -220 °C at the poles. Since RTGs were not
viable for this particular mission, it was thus proposed to use Radioisotope Heating Units (RHUs) to heat up the
scientific payload instruments at the surface of Europa such that they could operate and complete their proposed
mission. The RHUs chosen contain Plutonium-238 and generate 1 W of power, and they are small in size and mass.
They have also been used in previous space missions such as Galileo and Cassini, both of which used over 100 RHU
units in their respective designs.(8.10)
I. Final Seismometer/Camera Payload Design:
After all the components are placed together the final mass and power dimensions are calculated, and these
dimensions are displayed in Table 8.5. From this table, even though the peak power is currently at 16.26 W, the
average power used is actually 8.3 W, since once the camera boom and the seismometer boom are deployed, those
actuators will not be used again (except the rotary ones in the camera assembly). With the current shielding
protection it is estimated that the total dosage of shielding accumulated for the scientific payload package is 15 krad
for 120 days.
Table 8.5: Mass and Power Breakdown of Scientific Camera/Seismometer Payload
Component Mass (kg) Power (W)
Seismometer 2.88 1.5
Camera (2) 0.35 1.8
Camera Boom (5 Parts) 0.665 0
L16 Linear Actuator (4) 0.326 0.96
Type M8 Rotary Actuator 0.30 5
M3-RS Rotary Actuator (2) 0.30 7
RHU Units 0.12 0
Rotating Stick 0.405 0
Seismometer Separator 0.072 0
Camera Shielding 7.32 0
Seismometer Shielding 17.14 0
Total Dry Payload 5.42 0
Total Shielding 24.46 0
Total 29.87 16.26

45. J. Scientific Payload Mission Design:
After the orbiter deploys a lander onto the surface of Europa, the seismometer is deployed by a boom
towards the ground such that it may be able to read seismic activity straight from the surface of Europa. Then the
camera boom deploys the camera upwards so that it can start taking photos. This deployment along with the relative
orientation with the lander is shown in Figure 8.11. Here the seismometer is constantly recording seismic activity
data at an average rate of 12.5 kbps. The camera, in accordance with the RFP requirements, takes one picture for
every 4° of horizontal azimuth for a 360° FOV for a total of 90 pictures per 360° mosaic. The camera will change its
orientation in terms of tilt elevation depending on where the lander is located at. For the landers at the poles the
orientation does not change much, but for the landers not at the poles of Europa, the orientation can change up to
24° in tilt elevation. This way the cameras may be able to take mosaics that range from a different terrain
environment without looking at the “dead” zenith angle zone.
Figure 8.11: Deployed Seismic and Camera Configuration atop of a Lander
The amount of seismic data collected throughout the 90 day planned mission is the same for all 8 landers
across the surface of Europa. However, the amount of data that will be transmitted back changes depending on the
amount of time available for the orbiter to communicate with the lander. Landers at the poles of Europa are able to
send back all the seismic data, but the other landers have to compress their seismic data sizes in order for them to be
able to send the seismic data back to the orbiter. I wonder if anyone is even reading this. Well I was listening to
Unknown Mortal Orchestra whilst typing this up; their new third album came out today by the way. This is just a
random sentence, no coherence here. Another constraint that limits the amount of data that can be transferred back,

46. and these are the low-gain antenna data rate, which is assumed to stay at a constant rate of 149 kbps for the 90 day
mission at Europa. The amount of seismic data that is sent back to the orbiter is tabulated in Table 8.6.
Similarly, the amount of data that the camera can send back is also tabulated in Table 8.6, as well as the
amount of pictures and 360° FOV mosaics that are proposed to be taken at each individual lander station. The same
constraints that plague the seismometer data hinder the camera from being able to send too many pictures back to
the orbiter. Nevertheless, the amount of data taken from the pictures is enough to satisfy the RFP requirements.
Furthermore, for the cameras located at the poles of Europa, the cameras will take a few pictures of the planet
Jupiter for aesthetic purposes. The goal of these images is to determine if the terrain at the surface of Europa
changes within a 90 day period, as well as to present pictures with artistic value to the science community and the
general public in whole. Figure 8.12 shows how the camera and the seismometer transmit their data to the orbiter, as
well as how the camera is able rotate on its boom to take multiple 360° FOV mosaics.
From Table 8.6, the total amount of data that is able to be transferred back to the orbiter is 46.5 Gigabits
(Gb) for the camera (35,460 pictures) and 502.5 Gigabits (Gb) for the seismometer. This in total is 549 Gb
transferred back to the orbiter throughout the 90 day mission at Europa’s surface.
Figure 8.12: Camera/Seismic Payload to Orbiter Communication

47. Table 8.6: Detailed Data Transfer for Camera and Seismometer for 90 day Mission
Lander to Orbiter Transfer
Data (90 days)
Lander 1 Lander 2 Lander 3 Lander 4 Lander 5 Lander 6 Lander 7 Lander 8 Total
Time Available to Transfer
Data (hr)
221.40 196.64 70.52 43.26 42.68 71.27 197.70 222.29 1065.75
Data Transfer Rate (kbps) 149 149 149 149 149 149 149 149 N/A
No. of 2 pi Mosaics 90 70 15 15 21 19 74 90 394
Total No. of Pics taken 8100 6300 1350 1350 1890 1710 6660 8100 35460
Total Camera Data
Transferred (Gb)
10.6 8.3 1.8 1.8 2.5 2.2 8.7 10.62 46.5
Seismic Data Collected (per
day) (Gb)
1.08 1.08 1.08 1.08 1.08 1.08 1.08 1.08 8.64
Seismic Data Collected (Gb) 97.2 97.2 97.2 97.2 97.2 97.2 97.2 97.2 777.6
Seismic Data Transferred
(Gb)
97.2 97.2 35.96 21.38 20.41 35.96 97.2 97.2 502.52
Total Data Transferred (Gb) 107.82 105.46 37.73 23.15 22.89 38.21 105.93 107.82 549.00
IX. Telecommunications, Command and Data Handling
The design consists of an orbiter with eight landers, each containing it’s own telecommunications and
command and data handling system. The process for down-selecting a system capable of sending and receiving
commands via ground station began using a link budget calculation as shown in table 1, using a similar format to
that found in Brown, table 9.8, as well as Space Mission Analysis and Design by Larson and Wertz. The table
focuses on minimum and maximum range, as well as emergency uplink. Applying data rates and distance range
allows for the carrier uplink and data link performances to be determined, which then permits down selecting to a
system consisting of appropriate rates in order to carry out the mission.

48. Table 9-1: Link Table (Brown 9.8)
A. Telecommunications Subsystem
For space missions, a telecommunications system generally consists of X, Ka, and S- bands. Initially, an X
and Ka-band were selected for the orbiter, while an S-band was selected for each lander. The X-band would be used
for communication with the spacecraft, while the Ka-band would be used for the scientific data and images collected
to be transferred to ground stations through DSN. The S-band on the landers would be used to transfer the scientific
data and images collected by the lander, to the orbiter. The table below (table 2) lists the equipment initially
determined to be on both the orbiter and landers.

49. Table 9-2: Equipment List for Telecommunication
A trade study was then done on whether or not the Ka-band was necessary to carry out this mission. JPL
published a paper on the comparison between the X-band and Ka-band, where the advantages and disadvantages of
each were discussed. According to JPL, the X-band is not as power efficient as the Ka-band, which means that the
X-band requires more power to match the Ka-band. However, the Ka-band is much more sensitive to weather,
meaning that there could be a risk of power outages during the phase where data is being sent to the ground. If
power outages occur, this could last up to 30 minutes in time, with a standard deviation of one hour. A power outage
of this duration could risk the possibility of losing data. Therefore, the X-band deemed most reliable for this
mission. Table 3 shown below, lists the updated equipment for telecommunications after this trade study was
performed. The telecommunications system for the orbiter remained with the X-band, removed the Ka-band, while
the landers continued with the S-band.

50. Table 9-3: Updated Equipment List for Telecommunications
Once this trade study was used to down select which Band would be appropriate for the mission, rates to
allow for a data transfer needed to be determined. Provided that the distance from ground to Europa made this
challenging, using the rates of seismometer and camera used on each lander would determine the rate needed to
transfer data from lander to orbiter. Using the amount of time each lander had with the orbiter, also known as the
“window” for each lander to communicate, as well as how much data needed to be transferred, allowed for the
calculation of how much of a data rate was required. Data transferred from lander to orbiter would be done through
the use of a low gain antenna, rather than a high gain antenna. Low gain antennas have much more of a field of
view, or wider angle, which would mean data would be transferred as long as the orbiter was in view. Even though a
low gain antenna would provide a much lower frequency, it would be much more reliable than the use of a high gain
antenna. It was determined that a 149 kbps data rate for the low gain antenna would be feasible in transferring
camera and seismic data to the orbiter from the lander. Another trade study was done on low gain antennas to
determine if “stacking” antennas in the same direction would be more beneficial (able to transfer more data), than
simply using 4 LG antennas at different angles. The idea of “stacking” antennas would mean that the data rates

51. would increate by a factor of the amount of antennas used. Having antennas at different positions to “cover” a wider
angle of communication would allow for the lander to be able to communicate with the orbiter for a much longer
duration. The trade study proved that having a longer duration to communicate with the orbiter would allow for
more transfer of data, than to stack antennas for a much higher data rate.
Figure 9-4: Courtesy of JPL. Data Rate comparison using X and Ka-band

52. Figure 9-5: Courtesy of JPL Ka-Band for different MAR values
Figure 9-6: Courtesy of JPL. X-band for different MAR values
After determining the low gain antenna data rate, as well as the idea of using multiple LG antennas to cover
a wider field of view (increasing the angle of communication with the orbiter), understanding how much of a rate

53. the HG (high gain) antenna required was necessary. The HG antenna would be used to transfer the data from orbiter
to ground. This data would consist of the picture and seismic data received by the landers, as well as the data
collected during the orbiter’s time mapping Europa. Europa would be mapped using a Narrow Angle Camera (NAC)
as well as a Wide Angle Camera (WAC), during the orbiter’s orbit around the icy moon. In order to be able to
transfer these images during each orbit, a much higher data rate of 360 kbps is required. However, a HG antenna
requires that it be always aligned with its target in order to transfer the data, which means it must be pointed very
accurately. For this reason, it is necessary to provide back up antennas on the orbiter to continue the transfer if the
high gain antenna was unable to. LG antennas would also be attached to the HG antenna in order to ensure that there
is back up to carry out the transfer of data.
B. Command and Data Handling Subsystem
The C&DH system for both the orbiter and landers would consist of RAD750 processors, which seemed
most reliable, as they are very commonly used aboard a number of space missions. A solid state recorder (SRR)
would be present on both orbiter and landers in order to store the data collected, as not all the data would be able to
be transferred at one time. The orbiter would contain a 1 Tb drive, while each lander would only require a 1 Gb
drive to store data. Understanding the amount of data able to be transferred during data collection, and how much
would be stored, would allow for a down select of the size of the solid state required for each lander and the orbiter.
Photo sizes from NAC and WAC, along with HG data rates as well as LG rates (worst case) would allow for the
determination of a 1 Tb drive. Same concept applies to orbiters; however, on a much smaller scale, as each would
require 1 Gb of storage. Redundancy is necessary, as any error in C&DH would doom the mission. A second
processor would be needed to ensure redundancy of the computer system. For this, a second RAD750 processor
would be added to each lander.
C. Flight Modes for Space Mission
For a typical space mission, a number of modes are present onboard each spacecraft. Safety mode is very
crucial, as it is required to preserve the mission if anything was to go wrong. Understanding that anything can
happen, everything must be taken into consideration. This mode would place the spacecraft into a low power mode
with all unnecessary subsystems turned off, in order to preserve the spacecraft. A cruise mode, where low power is
used during the spacecraft’s trajectory, as well as a normal mode, where instruments are powered on during

54. trajectory in case any scientific data or images are to be gathered. Orbiter Mode is the final mode needed for a
successful mission as this is when the spacecraft will be in orbit, collecting the data once landers are deployed.
X. Power Subsystem
Each design requires two power subsystems, one to power the orbiter and one for the landers. Along with
this, the orbiters and the landers will require peak power estimates as well as average power estimates. The power
systems will be sized based on the average power estimates while peak power situations will be satisfied with a
battery. The power estimates for design one can be seen in table 10.1, while the power estimates for design two can
be seen in table 10.2.
Table 10.1: Design 1 Power Estimates
Orbiter Lander
Peak Power Average Power Peak Power Average Power
Subsystem W % W % W % W %
Thermal 82.00 17.07 82.00 25.68 7.2 3.97 7.2 20.48
ACS 135.78 28.27 56.74 17.77 100 55.09 0 0.00
Power 41.40 8.62 41.40 12.96 2.65 1.46 2.65 7.54
C&DH 20.00 4.16 20.00 6.26 10.8 5.95 10.8 30.73
Telecom 40.00 8.33 40.00 12.53 8.45 4.66 6.5 18.49
Propulsion 34.00 7.08 0.00 0.00 36 19.83 0 0.00
Mech 58.50 12.18 10.50 3.29 0.12 0.07 0 0.00
Payload 68.7 14.30 68.70 21.51 16.3 8.98 8.3 22.76
Total 480.38 100.00 319.34 100.00 181.52 100.00 35.45 100.00
Table 10.2: Design 2 Power Estimates
Orbiter Lander
Peak Power Average Power Peak Power Average Power
Subsystem W % W % W % W %
Thermal 101.10 20.87 101.10 28.44 7.2 3.36 7.2 20.57
ACS 124.00 25.59 72.00 20.25 177 82.68 0 0.00
Power 41.40 8.54 41.40 11.65 2.5 1.17 2.5 7.14
C&DH 20.00 4.13 20.00 5.63 10.8 5.05 10.8 30.86
Telecom 116.00 23.94 116.00 32.63 8.45 3.95 6.5 18.57
Propulsion 34.00 7.02 0.00 0.00 0 0.00 0 0.00
Mech 48.00 9.91 5.00 1.41 0.12 0.06 0 0.00
Payload 0.00 0.00 0 0.00 8 3.74 8 22.86
Total 484.50 100.00 355.50 100.00 214.07 100 35 100

55. Deep space missions have historically used nuclear power sources to power spacecraft. Radioisotope
thermoelectric generators (RTGs) have been used on most spacecraft travelling to Jupiter and beyond but Juno has
proved that solar power is feasible at this distance from the sun. Because of this, solar power was chosen for design
one while a combination of solar and RTG power was chosen for design two.
Solar power is dependent on the solar flux reaching the solar array. At Earth, the solar flux is 1370 W/m2
but this decreases based on the inverse square law. At Jupiter, solar flux drops to 51 W/m2
since Jupiter is 5.2 times
further from the sun than Earth is. This means a solar array at Earth will generate 27 times more power than at
Jupiter. This means that solar arrays on spacecraft at this distance have to be very large. Jupiter also has an intense
radiation field surrounding it and Europa lies within this field. This radiation will reduce the efficiency of the solar
arrays based on the thickness of the cover glass protecting the array. The solar arrays will also degrade gradually
with time. All of these effects increase the required size of the solar array. Low temperatures increase the solar
array’s efficiency however. At -130º C, the solar arrays will generate 20% more power.
Both orbiters design will be in the same orbit around Europa. This orbit is a full sun orbit. Not only does
this orbit maximize the amount of sunlight both orbiters will receive but also insure that the orbiter will sweep most
of the surface of the moon. The orbiters will still be eclipsed however. Jupiter is large enough and close enough to
block solar rays from reaching the solar arrays. An STK model was made of the illumination and eclipse times
experienced by the orbiter in its final orbit. This can be seen in figure 10.1
Figure 10.1: Orbiter Eclipse and Illumination Data
Jupiter will eclipse Europa and the orbiter every 69.5 hours. This eclipse occurs for 2.86 hours for each
eclipse. During the eclipse period, the orbiters’ loads will be powered by lithium ion batteries. The solar array will
be powering the orbiter’s loads and charging the batteries during illumination times. Both orbiters designs use rigid

56. panel solar arrays provided by Spectrolab. The chosen cells used for the solar array are ultra triple junction (UTJ)
GaAs cells. These deliver 350 W/m2
at Earth’s distance from the sun. Using the method in Elements of Spacecraft
Design by Brown, the solar array for design one must be 56 m2
and has a 24% power margin. This is distributed into
four wings of 14 m2
. The solar array for design two must be 63 m2
and has a margin of 25%. This orbiter has three
solar arrays, each 21 m2
in area. Figure 10.2 shows the deployed solar arrays for both orbiter designs.
Figure 10.2: Deployed Orbiter Solar Arrays
These solar array sizes were not consistent with the Europa Clipper’s solar array size. According to the
NASA Solar Study Status Report, a 460 W spacecraft would require a 46 m2
. Another method of sizing a solar array
was then attempted based on using the data sheet provided by SpectroLab and adjusting the power based on
efficiencies. This method can be seen in table 10.3.
Table 10.3: Solar Array Sizing Example
Area per wing 11.500 m
2
Number of arrays 4.000
Total area 46.000
P at earth 350.000 W/m
2
EOM degradation 324.557 W/m
2
(1.25%/yr from SMAD without radiation)
Temperature adjust 397.582 W/ m
2
(122.5% due to lower temperature)
Radiation adjust 357.824 W/m
2
(12% radiation efficiency reduction)
Flux Adjustment 0.037
P at Jupiter 13.217 W/m
2
Power Generated 607.973 W
Power Available 461.119 W
Excess 141.779 W
Margin 44.397 %

57. Using this method, a 46 m2
solar array would be able to power a 461 W spacecraft. This result matches the
result from the solar study done by NASA. The method was then applied to the design solar arrays. Design one’s
solar array has an area of 42 m2
and a power margin of 28.9%. Design two’s solar array has an area of 44 m2
and has
a power margin of 26.7%. These solar arrays weighed 168 kg and 176 kg respectively. The solar array
characteristics can be seen in table 10.4.
Table 10.4: Solar Array Characteristics
Textbook Method Data Sheets Method
Area
Power
Generated
Power
Margin
Mass Area
Power
Generated
Power
Margin
Mass
Orbiter
One
56 m2
396 W 24 % 224 kg 42 m2
411 W 28.9 % 168 kg
Orbiter
Two
63 m2
446 W 25% 252 kg 44 m2
431 W 26.7 % 176 kg
During the eclipse, solar arrays are unable to generate power. In order to power the orbiter during these
times, lithium ion batteries are used. Lithium ion batteries have a higher specific energy than nickel-cadmium
batteries, which means a smaller battery can provide more power. NiCd batteries also suffer from the memory
effect, in which the battery’s capacity reduces after being only partially discharged. Lithium ion batteries are not
affected by this.
The battery cells chosen for design one are the SAFTVES 180. These cells were chosen because they offer
the highest specific energy out of all candidates. The battery specifications are shown in table 10.5. The battery
required by the orbiter was sized using the same illumination and eclipse data as the solar array. The capacity
required for the battery was determined to be 61 Ah at 28 V. To satisfy this, two strings of eight cells will make the
battery. Eight cells in parallel will have a voltage of 28.8 V and two parallel strings will have a capacity of 100 Ah.
This will provide 2880 Wh of energy. The orbiter requires 1710 Wh of energy so the batter will have an excess of
1170 Wh. The second design’s orbiter also uses SAFT VES 180 batteries. This design requires 68 Ah at 28 V. The
battery also requires 16 cells, eight in series and two parallel strings. This battery has an excess energy of 977 Wh.

58. Table 10.5: SAFT WES 180 Battery Characteristics
specific energy 165 W/kg
energy 180 W/h
mass per cell 1.11 kg
Nominal Voltage 3.6 V
capacity per cell 50 Ah
diameter 0.053 m
height 0.25 m
The first lander design utilizes solar power. Since the landers will be on the surface of Europa, eclipses
from both Jupiter and Europa will block the solar arrays. The illumination times for each lander vary since their
locations determine when the Jupiter Eclipse occurs. For some landers, the Jupiter eclipse will occur during the
Europa eclipse and since Europa is tidally locked with Jupiter, the eclipse will always occur at this time. In order to
simplify manufacturing, all eight landers will be identical. All landers will be sized based on a worst case lighting
conditions, in which the Jupiter eclipse occurs during a Europa day. The eclipse and illumination data can be seen
below in figure 10.3.
Figure 10.3: Europa Surface Lighting and Eclipse Time
Again, the shaded regions represent eclipses. The wider bands represent eclipses due to Europa while the
smaller bands represent eclipses due to Jupiter. Together, these eclipses have a duration of 45.78 hours, leaving
39.44 hours of light to generate power. With a power requirement of 35.45 W, the solar arrays need to have an area
of 17 m2
. The solar array chosen are the Ultrafelx solar array made by Orbital ATK because of its low mass and
compact stowed size. The ultrafelx solar arrays can be seen below. The solar array ranges in sizes based on

59. diameter. In order to meet the power requirement, two solar arrays, each with a diameter of 3.3 m are used. This
generates 43.84 W and has a power margin of 23.81%.
Figure 10.4: Stowed and deployed Ultraflex Solar Array
Source: http://nmp.jpl.nasa.gov/st8/tech/solar_array3.html
During the eclipse, the lander will be powered by a battery. The lithium ion battery will use SAFT VL 9E
cells, which have a nominal voltage of 3.6 V and a capacity of 11 Ah. Although the power required by the lander is
much lower than the orbiter, the battery capacity required is almost twice as much as the orbiter. The lander will
require a 107 Ah battery since the eclipse time is long on the surface of Europa. In order to satisfy this, 11 strings of
eight cells are used to create the battery. This battery will also be used for peak power situation, the landing phase,
and during times when the power margin generated by the solar arrays is negative.
The second lander would use a RTG power source. An RTG generates power through a temperature
gradient on a thermoelectric generator. The heat source that generates the temperature gradient comes from the
radioactive decay of plutonium oxide. These generators will produce power constantly once the isotope pellets are
installed in the system. The current RTG model being used is the MMRTG, which has been used in the Mars
Exploration Rovers. The MMRTG can be seen in figure 10.5.

60. Figure 10.5: MMRTG Diagram
Source: https://solarsystem.nasa.gov/rps/docs/MMRTG%20Fact%20Sheet%20update%2010-2-13.pdf
The system provides 120 W at beginning of life (BOL) but suffers power degradation of 3.8% per year.
After six years, MMRTGs will only provide 90.4 W at end of mission (EOM). A modified version of the MMRTG
will have to be manufactured in order be used on the landers since the power output is twice the required power.
Since RTGs generate power based on a thermal gradient, reducing the length of the MM RTG by half will cut the
power generated by half. Reducing the number of radioisotope heating units (RHU) from eight to four will reduce
the mass of plutonium required by each system from 4.8 kg to 2.4 kg.
Although the MMRTG system can continuously produce power, a battery will still be utilized. Unlike the
solar power lander, the battery on this lander will only be used for peak power situations and landing. The capacity
of the battery is 11 Ah at 28 V. This gives an energy capacity of 308 Wh.
At 2.4 kg per lander, a total of 19.2 kg of plutonium is required to power all of the landers. This is more
than the available amount of plutonium. Along with this, an MMRTG system adds more complexity to the power
system. A solar array system can be integrated easier and is cheaper. An RTG system also has more risks of
contamination on Earth during launch and on Europa during the landing. For these reasons, the solar array lander
design was selected.
With the orbiter power source selected, the power system will need a power conditioning unit to convert
the power into a usable form and a power distribution unit to distribute the power to the necessary loads. The power
conditioning unit (PCU) is supplied from Terma. The PCU contains an array power regulator, battery
charge/discharge regulator, and a command and monitoring system. The array power regulator acts as a peak power
tracker, which uses the maximum power needed by the system. The battery charge and discharge regulator controls

61. the charging and discharging of the lithium ion battery. The command and monitoring system controls the other
components in the power subsystem. The power then goes into the power distribution unit. This unit transfers the
power into the appropriate loads.
The lander’s power control and distribution system is the Clyde Space Small Sat system. The Small Sat
system connects the solar array directly to the battery charge regulator. The power then goes into a power
conditioning module and then into a power distribution module. From there, the power goes to the lander’s loads.
The power system block diagram for both power systems can be seen in figure 10.6.
Figure 10.6: Power System Block Diagram
Since the lander uses solar power, the orbiter will need to supply power to the lander during the flight to
Europa. The lander would not be able to deploy the solar arrays and cannot generate its own power. The orbiter must
power the lander’s thermal system, power system, and command system. Because of this, the lander is considered

62. one of the orbiter’s loads, which can be seen in the power block diagram. The lander would then separate from the
orbiter when it is deployed.
On the surface of Europa, the lander’s solar array must be able to track the sun to maximize the power
generated. Using the height of the lander and the radius of the solar array, the maximum angle that the solar array
can rotate to track the sun in 37º but a maximum tracking angle of 35º was used. Along with this, a 1 km tall
obstruction placed 10 km away from the solar array was assumed to obstruct the solar array. The power generated at
versus time of day for each of the situations can be seen in figure 10.7.
Figure 10.7: Power Generated vs Time of Day.
The complete mass summary for the power system can be seen in table 10.6. Most of the mass of the
orbiter’s power system comes from the large 42 m2
solar array. For the lander, a lot of mass comes from the battery
required to power the lander.

63. Table 10.6: Power System Mass Summary
Orbiter
Solar Array 168 Kg
Battery 17.76 kg
Power Conditioning Unit 16.6 kg
Power Distribution Unit 13.2 kg
Total 215.56 kg
Lander
Solar Array 32.02 kg
Battery 21.12 kg
Power Control System 1.5 kg
Total 54.64 kg
XI Thermal System
The thermal environment at Europa is extremely harsh, with an approximate temperature range of -220
degrees Celsius to -160 degrees Celsius at the poles and equator respectively. Compounding the cold temperatures
with the high temperatures experienced during the Venus flyby required this mission to have a very delicately
designed thermal subsystem. As a result, a variety of different thermal control elements were explored, and can be
broken down into two general categories: active thermal control and passive thermal control. The purpose of these
elements is to ensure that the spacecraft (and all of its sub-components, including the landers) remains at an
allowable operating temperature, regardless of the extreme temperatures they are to be exposed to.
The active elements of the spacecraft include: radioisotope heater units (RHUs), resistance heaters, and
louvers. Radioisotope heater units are small devices that produce heat constantly throughout the length of the
mission. In others words, once these devices are activated, they cannot be turned off. The reason for this is because
they produce heat through means of radioactive decay. Resistance heaters, on the other hand, can act as a variable
heat source and are most commonly regulated through the use of either thermostats or solid-state controllers. For
this mission, solid-state controllers were the preferred selection for resistance heater regulation. The details of this
preference will be discussed shortly. Louvers are another active element of the thermal subsystem, and were needed

64. specifically during the Venus flyby. Louvers can be described as fins whose orientation can be adjusted via a
mechanical mechanism; when open, these fins increase heat expulsion through means of thermal radiation.
The passive elements of the spacecraft include: optical solar reflectors (OSRs), black & white thermal
coatings, and multi-layer insulation. Optical solar reflectors generally have low absorptivity and high emissivity as
characteristics of their thermal properties, and are usually composed of a quartz top-layer with a metallic sub-layer.
This makes OSRs a cold surface. Black & white thermal coatings serve opposite purposes each other. Black
thermal coatings are considered to be “hot” surfaces, as they retain a significant amount of the heat they absorb,
while white coatings (like OSRs) are considered to be “cold” surfaces, since there are efficient at ejecting heat while
absorbing minimal thermal energy. Multi-layer insulation, referred to as MLI for short, is composed of many thin
layers of plastic with a metallic coating. The main purpose of MLI is to ensure little or no thermal conduction
between layers, allowing portions of spacecraft to remain close to a constant temperature.
A. Key Drivers
The design of the thermal subsystem began by identifying the key drivers of the thermal requirements.
These drivers were derived from the required operating temperatures of the science instruments, as well as other key
components of the spacecraft. Table 11.1 (shown below) displays the most significant operating temperature
requirements taken into consideration. RHUs were included in the table with the minimum surface temperature and
do not have a maximum operating temperatures, as they produce heat constantly regardless of the thermal
environment.
Table 11.1 - Significant Operating Temperatures
Component Min. Temp. [K] Max. Temp. [K]
RHU 300 -
Lithium Battery* 233 323
HG Antenna 216 334
Propellant* 263 313
Star Tracker 243 323

65. Note: Items designated with an asterisk (*), were heavily considered due to their strict operating temperatures.
These items are the batteries and the propellant.
Although the lithium battery can operate over a range from 233K to 323K, it was determined that the
battery loses a notable amount of efficiency when operating outside the range of 273K to 283K, or approximately 0
degrees Celsius to 10 degrees Celsius, leaving a very small window for error in thermal control. Similarly, the
propellant also has very strict thermal requirements, with an operating range that only permits a ±20K temperature
swing from 283K.
B. Thermal Environment at Venus
During the Venus flyby, the spacecraft will reach its closest altitude at approximately 22,000 Km from the
surface of Venus. This is also the portion of the mission where the spacecraft is closest to the Sun, and is thus
subjected to its hottest thermal environment. Two important elements of this Venus flyby include: passing through
the shadow of Venus and using the high-gain antenna (HGA) as a sun shield. Passing through the shadow of Venus
significantly lowers the temperature of the spacecraft, and using the HGA as a sun shield further reduces the
temperature of the spacecraft; the combination of these two flyby elements together work to reduce the total solar
flux that the spacecraft is subjected to. Figure 11.1 (shown below) provides a pictorial of the Venus gravity assist.
Figure 11.1 – Pictorial of spacecraft and solar array temperatures during Venus flyby

66. It must be noted that the figure shown above provides temperature values that were computed assuming no
thermal protection was in place. Similarly, Figure 11.2 (shown below) is a plot displaying the temperature increase
of the spacecraft as a function of time. Note: this plot was generated using spherical approximations and assuming
no thermal protection was in place. The calculations used to compute these thermal values included shape factors,
as well the use of thermal characteristics such as emissivity, absorptivity, and albedo (reflectivity).
Figure 11.2 – Time plot of temperature during Venus flyby (Note: plot generated using spherical
approximations and assuming no thermal protection)
As can be observed from the plot, the maximum temperature of 363.6 K (approx. 90.44 degrees Celsius)
occurs approximately 3600s into the Venus flyby. At time, t = 0s, the spacecraft is approaching Venus and is at an
approximate altitude of 50,000 Km (Note: this altitude at time is not specified in the plot). It also must be noted that
this plot does not account for the cooling effects of passing through the shadow of Venus.

67. C. Multi-Layer Insulation
A number of different types of MLI were investigated for thermal properties, specifically regarding the
ratio of absorptivity to emissivity, α/ε. Absorptivity, α, describes the ability of the material to absorb energy from
incident electromagnetic waves. The specification of MLI that was finally selected was from Test Series A140, as
tested by Kennedy Space Center. This particular test coupon was found to have a density of approximately 37
kg/m3
. In order to account for the mass of the MLI to be used on the spacecraft and landers, a MLI mass calculator
was developed (shown in Figure 11.3). This calculator was generated using Microsoft Excel 2011, and requires the
input of the spacecraft area (or lander area) that is to be covered in MLI. As can be seen in the figure below, the
initial type of MLI that was used held a higher density of 79 kg/m3
. Although MLI makes up only a small portion of
the total spacecraft mass, this specification of MLI was determined to be too massive for the delta-V constraints of
the mission. As a result, a less dense specification (which still satisfied the thermal requirements) was selected for
the mission (Test Series A140).
Figure 11.3 – Image of MLI Calculator Generated in Microsoft Excel 2011
MLI degradation was an important factor that was considered for this mission. The impact factors that
were taken into account include: atomic oxygen (AO) exposure, thermal cycling, micrometeoroid impacts, and
radiation exposure. Since this mission is not an earth science endeavor, it was quickly determined that
micrometeoroid impacts and thermal cycling would not be an issue. Additionally, since there is virtually no

68. atmosphere at Europa, MLI degradation due to AO exposure was also deemed to not be a concern. The main
degradation factor that was taken into account was that of radiation and particle bombardment (such as the
accumulation of electrons and protons). It was determined that for a mission duration of 90 days, the total radiation
accumulation would be approximately 1500 Krad. This radiation exposure affected the thermal properties of the
MLI, as shown in Table 11.2 (shown below).
Table 11.2 – MLI Properties changes due to 1500Krad exposure, accumulated after 90 days
This is perhaps the only aspect of the mission for which radiation was found to be beneficial. As can be
seen in the table below, absorptance of the pristine (new) MLI sample was approximately 0.13, which then increased
by 28%, reaching a final value of 0.18. On the other hand, emissivity of the pristine MLI sample was approximately
0.79, which then decreased by 5%, reaching a final value of 0.75. In the table, both of these percent changes were
noted as being desirable. These material property changes would cause the MLI to become more of a “hot”
material, and since Europa has such an extremely cold thermal environment, these property changes increase the
MLI efficiency for the purposes of this mission.
XII Propulsion System
The propulsion system of spacecraft is arguably the most vital system when it comes to interplanetary
transfer. Considering the fact that this is a deep space mission with a rather short trajectory, many key drivers had to
be considered in order to ensure a timely arrival.
Since Jupiter is 5.2 AU away from the Sun, the spacecraft was designed to withstand the lowest operating
temperature of 272 K. Also, the power delivered by the solar arrays is limited at 5.2 AU from the Sun. Considering

69. the spacecraft operating at a low temperature and having limited power resources, the best option was to consider a
propulsion system was able to operate at low temperatures.
Figure 12.1 Propellant Equilibrium Temperatures
Moreover, some of the key drivers for the propulsion system were to maximize specific impulse (Isp), maximize
storability, and have maximum control over thrust variance. With the spacecraft operating temperature estimated to
be 272 K, or 486 °R, the propulsion system for the orbiter and lander was based on Hydrazine having a lowest
thermal equilibrium temperature of about 450 °R. The propellant combination that would give the best specific
impulse was found to be Nitrogen Tetroxide (NTO) and Hydrazine (Hyd), a liquid bipropellant to provide for the
propulsion system.
Out of the many engine candidates, the one who supplied the highest thrust and specific impulse was
Northrop Grumman’s TR-308 Liquid Apogee Engine. The TR-308 engine is able to provide a Isp of 322 seconds,
along with a maximum thrust of 471 N. Two TR-308 engines are needed in order to provide a sufficient amount of
thrust required during the 7 major interplanetary burns.

70. Figure 12.2 TR-308 Liquid Apogee Engine
The biggest concern for the engines was if it could handle the mission trajectory’s longest burn. This burn would
come at Jupiter Orbit Insertion and was estimated to last around 40 minutes. These engines are rated to have a
maximum firing duration of 50 minutes, yielding a 20% margin just in case the burn needs to last another 10
minutes.
After analyzing the possibilities of the Orbiter spacecraft being a Dual-Spin stabilized or a Three-Axis
controlled, it was concluded that the Dual-Spin stabilized orbiter was more beneficial. A Dual-Spin system is lighter
due to the Attitude Control System being less complex than the Three-Axis control. Since it is lighter, it only needs
12 reaction control thrusters, as opposed to 16, and it requires less Hydrazine to provide for the trajectory control
maneuvers. The Dual-Spin stabilized orbiter requires to have centrifugal tanks in order to maximize the efforts of
the Helium pressurant gas, and to minimize the residual propellant due to the spinning of the spacecraft. The
propulsion system for the orbiter is composed of the following components: pressure transducers, pyrotechnic
valves, system filters, solenoid valves, flowmeters, and latch valves.

71. Figure 12.3 Spacecraft Propulsion System Overview
For design purposes, two NTO and two HYD tanks were proposed. This would help balance the center of gravity for
the spacecraft while it was fully loaded with the landing units. The tanks were calculated to have a diameter of 0.8
meters, with a membrane thickness of 1.96E-03 m. The design of these titanium tanks are expected to withstand
3300 kPa of maximum tank pressure.
Since this is an interplanetary mission with 8 landing units and has a trajectory with a decently large ΔV,
most of the mass will come from the propellant. Two trajectories were analyzed in order to optimize the ΔV
required, and consequently reducing the propellant mass. The trajectory with a proposed arrival date in 2026 would
require a ΔV of 2.95 km/s, whereas an arrival in 2027 would require a ΔV of 2.41 km/s. The difference in these
proposed trajectories is about 500 km/s. The mass penalty for carrying propellant for this difference in ΔV is about
1,500 kg. The comparison of propellant mass and wet propulsion system mass can be found in the tables below. The
propellant masses listed in the table account for an extra 10% of total ΔV, losses due to 7 major startups, and an
expected 3% residual propellant mass.

72. Table 12.1 Spacecraft Propulsion System Mass Comparison
Using the Falcon Heavy for a 2026 arrival yielded a very comfortable margin. On the other hand, it would
not be possible to launch with the Delta IV Heavy. The only way that the mission could launch was to reduce mass
in two very important areas: payload and propulsion. A combination of reducing the total amount of landers to 7,
and reducing the ΔV by 500 km/s, gave us a rather small but positive launch margin.
The propulsion system for the landing units also followed a very similar fashion as the orbiter. Due to the
extremely cold temperatures on Europa, the propulsion system was designed to be powered by a similar liquid
bipropellant. In this case, NTO is once again being used but combined with Monomethylhydrazine (MMH). The
structural design constraints on the landing units had a significant impact on the propellant selection, which was due
to the engine chosen for the landers.
Figure 12.4 R-1E Engine

73. The R-1E by Aerojet Rocketdyne had the best trade between nozzle length, Isp, and thrust. Moreover, these
engines have heritage from the Shuttle program, so they have been proven to work in the past. Aside from being
lightweight, it is capable of having 330,000 pulses available to vary thrust. This is particularly important since it is
very complex to throttle engines, short pulses may be used in order to maintain the ideal thrust level. In order to
ensure that the propulsion system could deliver the pulses necessary, it’s design was composed of the following
components: pressure transducers, pyrotechnic valves, system filters, solenoid valves, flowmeters, and latch valves.
Figure 12.5 Lander Propulsion System Overview
The R-1E engine has a Isp of 280 seconds and a very light mass of 2 kg each. At 110 N each, the design of
the propulsion system requires that 4 engines be used per lander to meet the proposed landing scheme. The engines
were strategically placed along the center of gravity of each lander. The benefit of having the engines placed in the
center is to prepare for the chance of any single engine failing. If any one engine fails at the center, ACS can be used
in order to compensate for that misalignment of thrust. If the engines were placed at the corners and any one failed,
the landing unit would tumble and be uncontrollable.

74. The design of the propulsion system for the orbiter and lander are very similar. There are 2 tanks for MMH
and NTO in order to have a more centralized center of mass. Pairing a tank of MMH and NTO to feed two engines
worked the best for this design. For every two engines, these tanks measure about 0.2 meters with a membrane
thickness of 5.95E-04 m. The weight of each tank was calculated to be about 0.82 kg.
Table 12.2 Lander Propulsion System Mass
The landing scheme requires for the propulsion system to deliver a ΔV of 1.48 km/s. This was calculated to be about
144 kg of propellant. This propellant mass accounts for an extra 10% of total ΔV, and loses due to startups.
XIII Attitude and Articulation Control Subsystem
The path of our spacecraft during its powered flight is directly influenced by its attitude and orientation in
space. Once outside the atmosphere, changing the direction of thrust by articulating exhaust nozzles or changing the
spacecraft's attitude influences its flight path. Our spacecraft's attitude will be stabilized and controlled so that its
high-gain antenna will be accurately pointed to Earth for communications, so that onboard experiments may
accomplish precise pointing for accurate collection and subsequent interpretation of data, as well as heating and
cooling effect of sunlight and shadow may be used intelligently for thermal control.

75. Figure 13.1: Dual Spin Orbiter
The mission to deploy multiple landers on the surface of Europa is a tall order, let alone a successful
mission alone to just navigate Europa. Requiring over 4 years of interplanetary travel, a Jupiter orbit insertion, a
Europa orbit insertion, and deploying 8 eight landers, will add up to a very large mass, and every kilogram really
counts. To avoid an additional 60 kilograms and 88 watts, the stabilization method we chose eliminates our need to
adapt reaction wheels on our orbiter. The method chosen to stabilize for this mission is not the normal 3-axis
stabilization, nor the spin stabilization technique, but the less frequent dual spin stabilization. Spin stabilization was
an option, but when the need for constant communication with the landers while in orbit about Europa, we needed to
adapt a despun section. This called for the use of a dual spin stabilization for the simplicity of a gyroscopic
stabilization, which allowed for a minor axis of inertia to be our spin axis. The orbiter's despun section, shown in
Figure 1, contains the electronics box, high gain antenna (HGA), low gain antenna (LGA), as well as the orbiter
cameras used for mapping the surface during the first 30 days mapping phase. The Bearing and Power Transfer
Assembly (BAPTA) is the original mechanism chosen for the dual spin mechanism, however, the max weight
allowed was much less than the weight of our spacecraft. This led to the switch to the spin mechanism assembly
used on the Global Measurement Instrument. This mechanism allows for simple damping between the spun and

76. despun sections, and for major burns releases the damping mechanism to permit both sections to spin freely
together. For these major burns, the spacecraft will require to be spinning at a rate no less than 6 rpm, due to the
centrifugal tanks used to accommodate the spin of the spacecraft.
For attitude control, there needs to be a reference on which was is 'up'. Many different devices may be
chosen to provide attitude reference by observing celestial bodies, or using inertia as a reference. The orbiter in our
proposed designed utilizes a total of five attitude references. This consists of three celestial references as well as two
inertial references. The three celestial references consist of two star trackers and a sun sensor. The star trackers used
are the CT-602 Star Tracker manufactured by Ball Aerospace. The sun sensor is specifically used for spinning
spacecraft, and that is the Adcole Spinning Sun Sensor. The star tracker uses an automated recognition of observed
objects based on built-in star catalogs. The sun sensor also if needed could be used for yaw and pitch reference.
Most star trackers use its roll reference with Canopus, a bright star. For this too work, our star trackers are placed on
the non spinning section of the orbiter. The sun sensor, considering it is used mainly for spinning spacecraft, will be
placed on the spinning section of the orbiter. The inertial references are the same instrument, just coupled for
redundancy. Added to the orbiter will be two LN-200 Core IMUs manufactured by Northrop Grumman Corporation.
Attitude control is obtained by sensors first most, but these communicate with the actuators, which in our case are
thrusters or our reaction control system (RCS). The actuators chosen for our design are the MR-106E 22N thrusters.
The orbiter is utilizing the configuration found on the Juno Spacecraft. Juno is a spinning spacecraft, and since our
thrusters are placed on the spinning section this became our design as well. The thrusters are configured with two
Figure 13.3: Adcole
Sun Sensor
Figure 13.2: CT-602
Star Tracker