ARO 483 -- Aeolus Tech AIAA Proposal FINAL

Europa CT Scanning Program: Multiple-Flyby Mission
Design
Thirupathi Srinivasan1
, Timothy Hofmann2
California State Polytechnic University, Pomona, CA, 91768
Hayk Azatyan3
, Wesley Eller4
, Jonathan Guarneros5
, Luis Leon6
, Ling Ma7
, Christopher Prum8
, Matthew
Ritterbush9
, Charles Welch10
California State Polytechnic University, Pomona, CA, 91768
The growing interest in exploring Jupiter’s moon, Europa, over the last decade by the
scientific community has prompted various studies of unmanned, robotic exploration of the
moon. The in-situ scientific data provided by such robotic probes would supplement that
provided by the future Europa Clipper mission. To carry out this task, the Europa CT
Scanning RFP by the Jet Propulsion Lab requires the design and development of a seven-
lander mission that provides seismographic and imaging data across logarithmic locations on
Europa for 90 days. A multiple-flyby mission design involving dual-carrier satellites and seven
landers addresses such RFP requirements. This design involves staggered launches similar to
the Voyager and Pioneer missions, with the first satellite containing three landers and
scientific payload, and the second satellite transporting four landers. The two carrier satellites
will execute multiple flybys of Europa. These seven landers will utilize MEMs seismometers
and imaging systems from past missions for the primary in-situ scientific data. This low-risk
mission design allows for redundancy in telecommunications and lander deployment, and
significant mass margins at the expense of $4.9 billion total cost.
1
Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA ,91768
2
Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768
3
4
Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768.
5
6
Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768.
7
8
9
10

Nomenclature
a = Albedo
e = Orbital Eccentricity
D = Diameter
Fs = Radiation view factor
Gs = Direct solar flux
H = Altitude
Ka = Albedo correction
Pmax = Maximum nominal power
Pmin = Minimum nominal power
q = Energy rate input
qIR = IR emission rate
R = Radius
T = Temperature
Tmax = Maximum temperature
Tmin = Minimum temperature
Tspace = Space temperature
Tsur = Surface temperature
α = Absorptivity
ε = Emissivity
σ = Stefan-Boltzmann constant
η = Efficiency
I. Introduction
he dual-launch, multiple-flyby mission design constitutes two carrier satellites and seven “soft” landers. The
scientific objective of the mission is to provide in-situ seismographic and imaging data from the surface of Europa at
seven latitudinal and longitudinal locations as dictated by a logarithmic trend. Secondary scientific objectives include
optical reconnaissance of the Europan surface and measurements of the Jovian magnetic field. The primary scientific
data is expected to be relayed to Earth regularly during the 90-day operational mission phase for the landers. Due to
the short development period of this design and early launch date in late-2019, much of the instruments and spacecraft
T

components are those from past missions and commercial-off-the-shelf (COTS) components. This was done to
expedite the production, V&V, I&T, and ALTO phases.
The selection of this design was based upon the following prioritized, primary design drivers:
1. Europa surface mission operation start date before Dec. 31st
, 2026
2. Non-Europa disposal
3. Survivability of at least seven landers and carrier satellites for the mission duration
4. Periodic data transmission from the lander to carrier satellites, and to mission control on Earth
5. Safe and reliable deployment of the landers, and its’ scientific instruments
6. Detection of P-,S-, and L waves and mosaic with at least 2π steradian coverage for every 4o
solar elevation
7. Logarithmic placement of the landers on Europa as per RFP requirements
Satellite #1 will be launched in mid-October 2019, and will carry three polar landers, an optical payload package,
and a magnetometer. The primary payload for this satellite are the three polar landers. These polar landers are named
as such for they orbit Europa in a 90o
inclination (or polar) orbit prior during the initial and detailed reconnaissance
phases. The secondary scientific payload for Satellite 1 include the HiRise and MARCI cameras, which are used for
preliminary landing site reconnaissance, and the Galileo-based magnetometer (MAG).
Satellite #2 will be launched in late-December 2019, and will carry four non-polar landers as its primary scientific
payload. These non-polar landers orbit Europa in a 60o
inclination for the detailed reconnaissance phase before
landing. It also carries the magnetometer as its secondary scientific payload. Both satellites are launched from Falcon
Heavy launch vehicles. They follow the VEGA trajectory, with Jupiter arrival in November 2024, and a 1.5 year
Jovian tour for the pump-down phase. The pump-down phase involves multiple flybys of the four main Jovian
satellites including Ganymede, Callisto, Io, and Europa prior to lander deployment. The satellites are placed in a final
slightly staggered, elliptical orbits around Jupiter (e = 0.19), with an Europa flyby every 3.6 days. Both satellites utilize
flex-rolled up solar arrays (FRUSAs) modeled off the Mega-ROSA technology by Deployable Space Systems to allow
for packing within the payload fairing.
The polar and non-polar landers contain a Silicon Audio Geolight MEMs seismometer and a multi-spectral Beagle
camera on a helical boom. A single axis of the MEMs seismometer are placed within the “foot” of each of the four
lander legs to sense P-, S- and local waves. The fourth seismometer is included for redundancy. The landers will be
deployed during the closest approach of Europa by the respective carrier satellites, and will execute the Europa Orbit

Insertion burn. Unique technologies for the landers include the quantum-well power system, which alleviates the need
for RTGs that can potentially contaminate the Europan surface, and the use of toroidal tanks for uninterrupted
shielding of the electronics vault on all sides. Likewise, the satellites use the cylindrical propellant and pressurant
tanks for shielding the electronics vault that contain the C&DS components.
II. System Description
A. Concept of Operations
The key segments of the mission include launch, interplanetary travel, the Jovian tour, primary mission phase, and
disposal. During each of these phases there are many key requirements that must be met, and operations that must be
performed for a successful mission. The overall mission concept informs these requirements and operations.
The general mission concept will be a two-satellite, multiple-flyby concept launching from Kennedy Space Center
in late 2019 on a Venus-Earth Gravity Assist (VEGA) trajectory. Upon arrival at Jupiter, each satellite will perform
its own Jupiter Orbit Insertion before setting upon its Jovian tour, lasting 1.89 years. At the end of their tour the
satellites will be in Europa-synchronous orbits with periods matching that of Europa, and orbit eccentricities of 0.186.
This will ensure a pass of Europa every 3.55 days (the period of Europa) for each satellite allowing near constant
communication with the landers.
The landers will be deployed from their respective satellite when the satellites perform their final Europa gravity
assist before entering their multiple-flyby orbits. Satellite 1 will carry three landers, which will orbit Europa with polar
inclinations starting on October 16th
, 2026, while Satellite 2 will carry four landers which will orbit with inclinations
of 60°. The non-polar landers will begin their orbits on October 17th
, 2026. Following Europa Orbit Insertion, the
mapping phase begins, and within one month, all landers will have made their descent to the surface of Europa, and
will be operational by November 17th
, 2026, 43 days before the operational deadline.
Following the 90 day mission, the satellites will continue to orbit Jupiter in their flyby orbit. It has been determined
that there is no risk of impact with Europa over the course of the next five years of flybys in the proposed orbit.
Eventually, the satellites orbits will decay enough for them to impact Io or to sink beneath Jupiter’s surface, however
this would be many years after the end of this mission. Other disposal plans are available, but only with the addition
of extra ΔV. The mission can handle extra fuel mass due to the high mass margins, however this change would be

unnecessary as will be discussed in the disposal section. Below is a depiction of the mission concept from launch to
disposal (Fig. A.1) as well as a list of the mission phases and their definitions (Table A.1)
Fig. A.1 Richter program concept of operations diagram depicting all mission phases from launch to
disposal.
Table A.1 Richter Program Phase Descriptions and Timeline
Phase Sub-Phase Description Dates
LaunchPeriod
Satellite 1 Launch
Countdown to launch, launch, and insertion
into 400 km parking orbit.
16 Oct 2019
Earth Parking Orbit
(Satellite 1)
/
Pre-launch Prep
(Satellite 2)
In Orbit: Contact made with DSN. All major
flight subsystems deployed, science
instruments calibrated.
On Ground: Launch pad prep for Satellite 2,
Satellite 2 final systems check.
16 Oct 2019 - 26 Dec
2019
Satellite 2 Launch
Satellite 2: Countdown to launch, launch,
deployment of major flight subsystems,
science instruments calibrated, contact made
with DSN
Both Satellites: Injection into heliocentric leg
of VEGA trajectory.
26 Dec 2019 – 29 Dec
2019

A.1 Launch Period
Both satellites will be launching from Kennedy Space Center. Satellite 1 will launch on a Falcon Heavy on 14 Oct
2019, and will enter into a 400 km LEO parking orbit, where it will remain until Satellite 2 launches on 26 Dec 2019.
The trajectory was designed for a satellite launching on 26 Dec 2019, so to accommodate two satellites, the first will
wait in Earth orbit until it can match the departure date of the second satellite. A staggering of the satellites will be
necessary to prevent collision en route. Even with just a few minutes of distance the odds of collision are extremely
Phase Sub-Phase Description Dates
Inter-planetary
Cruise
Regular system health tests, Venus and Earth
gravity assists, Deep Space Maneuver, clean-
up maneuvers. During Venus approach HGA
will point toward sun to provide shading for
sensitive equipment.
Dec 2019 – Mar 2024
Jupiter Approach
Final clean-up on approach to Jupiter, JOI,
preparation for data reception.
Mar 2024 – Nov 2024
Pump-down
Jovian Tour
Gravity assists from Ganymede, Europa and
Io to lower orbital energy, HiRISE imaging
during close approaches (mainly Europa).
Sets up Europa flyby orbit for both satellites.
Nov 2024 – Oct 2026
Lander Deployment
On last Europa gravity assist, landers deploy
from satellites.
16-17 Oct 2026
LanderOperations
Primary Mapping
After EOI, a single polar lander maps Europa
in bands at 200 km altitude with MARCI.
Data sent to satellites. Satellites send
promising sites to individual landers. All
landers engage periapsis lowering burn.
17 Oct 2026 – 1 Nov
2026
(14 days)
Down-selected
Landing Sites
Mapping
At periapsis (2 km) each lander uses MARDI
to gain higher resolution landing site images.
Information processed on lander.
1 Nov 2026 – 6 Nov
2026
(5 days)
Descent
De-orbit burn, LIDAR and MARDI provide
real-time data to lander, hazard avoidance,
deployment of lander legs, touchdown.
Descents will be staggered.
6-7 Nov 2026
(91 seconds per lander)
Science Mission
Camera deployed, seismometers recording,
regular system health checks,
communications with satellites.
7 Nov 2026 – 6 Feb
2027
(90 days)
Satellite Operations
Regular communications with all landers,
data transmission to DSN, orbital station-
keeping to counteract Jupiter and Europa
effects, regular system health checks
17 Oct 2026 – 16 Feb
2027
Disposal
Extended mission (dependent on
lander/satellite condition), leave satellites in
flyby until orbit degrades into Jupiter’s
atmosphere
Extended Mission (Feb
2026 – May 2026)
Disposal
(Feb 2026 – Feb
2031)*
* Disposal found to not impact Europa for five years. This was maximum possible propagation time for STK
running on student computer.

low. However, it was decided that Satellite 1 will depart its parking orbit one full orbit before Satellite 2 is set to pass
through the orbit. This will provide a buffer zone between the two spacecraft, while keeping their ΔV’s consistent.
A.2 Interplanetary
Each satellite passes Venus and Earth on their trajectory to Jupiter. Throughout the journey regular health reports
will be generated semiannually as a means of troubleshooting all subsystems before they have the chance to fail.
Immediate damage reports will be transmitted to the DSN upon collisions with space debris, or upon a system fault.
During interplanetary travel, and most importantly on the approach to, and shortly after encountering Venus, the
satellites will be subjected to drastically different temperature environment. The temperature at Venus gravity assist
is potentially harmful to many components of the system. The solar heat flux is about 2631 W/m2
at Venus, compared
with 1570 W/m2
at Earth, and ~50 W/m2
at Europa. The drastic variation in heat flux leads to a drastic variation in
temperature, meaning that different measures must be taken in order to cool the satellites at Venus, and to warm it at
Europa. As far as cooling the satellites at Venus, louvers will be installed close to the electronics vault to provide
ventilation, and the electronics vault will also be more thermally isolate from heat flux effects than the rest of the
spacecraft. Another measure being implemented is turning the satellites HGA toward the sun on approach to Venus
to eliminate much of the heat flux on the majority of the satellite, and landers.
Several clean-up maneuvers are scheduled to take place preceding and following main interplanetary events, the
largest of which is the Earth escape burn performed by the launch vehicle. Fuel allowances have been made to
accommodate such burns, however the amount necessary per burn, and the date of the burns are not set due to these
burns only being necessary should the gravity assists or initial burn not cause the desired route to be taken. An
overview of the interplanetary trajectory is shown in Fig. A.2. Note, the only difference between Satellite 1 and
Satellite 2 trajectories is the launch date. The rest of the interplanetary mission will see the satellites close together,
due to Satellite 1 staying in a 400 km LEO orbit until the launch of Satellite 2.

Lastly, each satellite will perform its JOI burn on 26 Nov 2024, concluding its interplanetary travel with a final
burn of 950 m/s, which will occur over a period of roughly 2.5 hours at an altitude of 12.8 Jupiter radii from the surface
of Jupiter. The JOI burn places each satellite into a highly elliptical, 4° inclined orbit with respect to Jupiter. The
eccentricity, and period of the orbit will be lowered significantly from the gravity assists in the Jovian tour phase of
the pump-down.
A.3 Pump-Down
For Satellite 1 pump-down consists of a total of 22 gravity assists: Five of both Ganymede, and Io, and twelve of
Europa. Satellite 2 performs 21 gravity assists: Six of Io, seven of Ganymede, and eight of Europa. Both satellites
encounter Ganymede five times, then Europa once before departing paths. These first six gravity assists reduce the
apojove from being more than 11 million km from Jupiter, to less than 2 million km, reducing the orbital period from
roughly 300 days to just 13 days. Upon each pass of Europa, the Satellite 1 will be oriented so that the MARCI, MLA,
and HiRISE are focused on the surface of Europa. The benefit of doing this is to obtain early detailed imaging of
some of the potential landing sites, in some cases more than a year before lander deployment. As Satellite 1 undergoes
Fig. A.2 Satellite mission trajectory map generated using STK with the Astrogator module and Planetary
Data Supplement.

quite varied passes of Europa in both altitude, and inclination, it is ideally suited for this task. Figure A.3 shows the
passes that Satellite 1 makes of Europa.
Figure A.4 illustrates the steps taken on each pass of Europa during pump-down, as well as the lander deployment
scheme for both satellites. It’s important to note that the scheme for each flyby of Europa can be implemented for
flybys of Ganymede and Io as well to provide secondary data not critical to mission success, but possibly of some
scientific value.
Fig. A.3 Two views of Europa showing Satellite 1 passes covering diverse positions around Europa. Most
passes occur on Jupiter facing side of Europa.
North pole
South pole
Fig. A.4 Satellites mission phases at Jupiter showing pump-down, lander deployment, and flyby orbit

Upon each satellites final gravity assist of Europa before entering their multiple-flyby orbit, they will deploy their
lander payload. Satellite 1 is carrying the polar landers, which need to orbit at 90° inclination. Should they be deployed
at closest approach, a massive plane change maneuver would be needed to change their inclination. Instead the plan
is deploy the polar landers 50,000 km from Europa. This will allow for a small burn to change the inclination by the
amount needed (~25°). In doing this the landers can also be spaced far enough away to provide some collision buffer.
The deployments of the polar landers will occur on 16 Oct 2026. At the moment of deployment the landers will sync
their clocks with each other, so that seismic data may be collected accurately upon landing. Satellite 1 will also send
a transmission to Satellite 2 at the moment of deployment letting it know to tell the non-polar landers the sync time.
In contrast, Satellite 2 is transporting the non-polar landers. These landers require no change of inclination with
respect to Europa, and therefore may be deployed closer to the approach of Europa. In order to provide some spacing
between lander orbits, the deployment zone will be between 5000 km altitude at the start of deployment to 300 km at
the end. The window for deployment is roughly 45 minutes, providing 10 minutes between the launches of each lander,
or should the landers deploy in pairs, 30 minutes between launches. The deployment of the non-polar landers will
occur 17 Oct 2026.
A.4 Multiple Flyby Concept
As the landers’ operations begin, the satellites have entered their last true phase. While in the multiple flyby orbit
the satellite spends most of its time pointed toward Europa. Each satellites orbit has been designed to provide coverage
of all landers during each orbit in the event of a critical failure in the other satellite. Figure A.4 shows that each lander
has a block of time in which it may communicate with either satellite. This time-block given to each lander is roughly
3 hours, which is what is needed to transmit the expected science data from each lander. Also included in the orbit of
the satellites is time for communications to Earth. The mission will be requesting 24 hours per week from the DSN to
transmit important scientific data during the science mission. Since each orbit is roughly 3.5 days, 12 total hours of
communication have been planned into each the orbits of the satellites. Of course, should one fail, a single satellite
would need to communicate for the full 12 hours. This is no problem, as each orbit has a long duration in which no
data reception or transmission is occurring, so if needed, some of this idle time can be converted into communication
time.
Something to note is that the flyby orbit has a natural migration. Upon arrival Satellite 1 will be closest to Europa
on one side of the orbit, while Satellite 2 will be closest at the opposite end of the orbit. As the satellites encounter the

edge of Europa’s sphere of influence the duration of their orbital periods are reduced slightly. This causes them to
migrate farther away at the point in the orbit where they were closest to Europa. Over the span of 1.3 months the orbit
has migrated enough that the satellite is now closest to Europa at the far end of the orbit. At this point again, the
satellite encounters the edge of Europa sphere of influence, however instead of shortening the period, this encounter
lengthens it. A longer period causes a migration in the opposite direction. This process occurs for both satellites, and
repeats itself several times over the lander mission phase. This means that the depiction of the communications in Fig.
A.4 is a representation of only one orbit, and that each orbit following this one would see a slight shift in the placement
of the lander communication segments.
The multiple flyby concept creates a very complex mission schedule, especially with seven landers in need of
communication and in need of deployment. The first choice for the satellites was to have them orbit Europa in the
same orbits now occupied by the landers, therefore Satellite 1 would be a polar orbiter, and Satellite 2 would be
inclined 60°. As a result of this orbiter concept, it became necessary to dispose of the satellites on Europa via a crash
landing. This brought up concerns at SDR due to planetary protection, which was a known risk of disposing of the
satellites on Europa. Due to the concern expressed, several alternate orbits were proposed for the satellites.
The first alternative was to maintain the 200 km orbits for the satellites, and perform a burn at mission end to
escape Europa and dispose either in a higher orbit, or on Jupiter. The key disadvantage to this approach was the high
ΔV involved. The escape burn alone would add about 650 m/s.
The second alternative was to place the satellites in highly elliptical orbits around Europa, with the periapsis at
200 km, and the apoapsis at 2000 km or higher. The advantage of this is a much lower ΔV for EOI, and for the escape
burn. This approach made mapping landing sites uneven, as well as added fuel mass to the landers which would have
to perform a larger de-orbit burn.
Table A.2 Satellite/Orbiter mission concept trade study
Satellite Mission
Concept/Disposal
Planetary
Protection?
ΔV Penalty (m/s) Complexity Mass Margin
Circular Orbiters/Crash
Landing on Europa
No 0 Low +250 kg
Circular Orbiters/Jupiter
Disposal
Yes
Orbiters = ~ +650
Landers = 0
Low -1500 kg
Elliptical Orbiters/Jupiter
Disposal
Yes
Orbiters = ~ +300
Landers = ~ +100
Medium -400 kg
Multiple Flyby/Degrading
Orbit Disposal
Yes
Satellites = ~ -400
Landers = +1600
Medium-High
Satellite 1: +2500 kg
Satellite 2: +2000 kg

The third alternative is the currently chosen mission concept of leaving the satellites in Jupiter orbit, while the
landers perform EOI, and mapping. This concept drastically decreases satellite mass, at the cost of greatly increasing
lander mass. It also means a more complex mission concept as seen above, however this concept provides the best
mass margin while achieving planetary protection measures, and it was easiest to implement. Table A.2 shows the
benefits and weaknesses of the four mission concepts under consideration after SDR.
A.5 Lander and Satellite Operations
A.5.1 Primary Mapping
After deployment from the satellite all landers will enter into a 200 km orbit around Europa. Of the three polar
landers, one will proceed with mapping starting on 17 October 2026 and will last fourteen days: seven days for
mapping, and seven days for transmission from the polar lander to the satellites, and then from the satellites back to
all landers, after data processing. The polar lander is chosen for mapping over the non-polar lander because over the
course of several days in orbit, the polar lander will see all of Europa, whereas the non-polar landers will never see
either of the poles, which are both landing sites. Normally, a satellite would be selected to map a region for a space
mission, however, due to the planetary protection concerns mentioned in Section A.4 of this report, the satellites will
never be close enough to Europa for a long enough period of time to do any long-term mapping.
The Mars Color Imager (MARCI) camera will be used which provides images with a resolution of 5.3 km/pixel.
Even at this resolution, mapping the entirety of the moon would take much longer than time constraints allow.
Fig. A.5 Initial Mapping Phases Operations for Polar and Non-Polar Landers

Therefore mapping will occur in bands, which will cover the latitudes upon which the possible landing sites are
located. Figure A.5 depicts the orbits of the two lander types, and their operations during the initial mapping phase.
The non-polar landers are largely idle in this phase, besides sending periodic health transmissions.
When the polar lander completes its sweep, the landing site data is transmitted to both satellites, which analyze
the data and find promising landing sites in each of the bands. Once landing sites have been determined, and have
been checked for logarithmic placement along the longitude of Europa (see Fig. A.6), one landing site is transmitted
to each lander. Note that the landing sites in Fig. A.6 are not the final landing sites, they are the desired landing sites.
Should one of the sites depicted prove too treacherous, new landing sites will be chosen. Once these sites have been
chosen the three polar landers will be sent the navigational data for L1, L6, and L7. These landing sites all above 60°
latitude, meaning they are unreachable by the non-polar landers. It makes little difference which of the three landers
lands in a particular site. The other four landers will be sent the navigational data for L2 through L5. These are all
lower than 60° latitude meaning that the non-polar landers can land at any of these sites.
With the landing site
information received, the
landers proceed with a 43
m/s burn at apoapsis to
lower their periapsis to 2
km directly above their
intended landing site. This
will happen in a staggered
manner, where one lander
will proceed with this
maneuver at a time to ensure constant communication in case of an issue. All landers will be in a 200 km x 2 km orbit
with periapsis above their landing site on 1 November 2026.
A.5.2 Down-selected Landing Sites Mapping
The initial mapping selects 540 km diameter regions of Europa for each lander to find a landing location in. The
RFP sates that each lander must be emplaced within a 5° (136 km) diameter circle with the center at the perfect
Fig. A.6 Potentially Landing Sites in Logarithmic Spacing

logarithmic placement point. Thus, the initial mapping phase would not allow for a high probability of being in range
for logarithmic placement.
The second mapping phase will provide more detailed topography information for each landing site. The previous
lander mission segment brought the landers orbits to 2 km periapsis directly above that landers intended landing
location. On approach of periapsis each orbit, the Mars Descent Imager (MARDI) camera and Mercury Laser
Altimeter (MLA) on each lander will begin taking detailed imagery in the 540 km x 540 km region. The MARDI
camera and MLA will begin taking data at an altitude of 20 km above the surface of Europa. At this altitude the
MARDI images will have a resolution of about 10 m/pixel. As the lander passes periapsis the images will improve in
resolution to 1.5 m/pixel. Images, and altitude readings will be taken until the lander has achieved a 20 km outgoing
altitude, at which point the payload will enter rest mode until the next approach of periapsis. The region where data is
being taken will pass extremely quickly; the entire 200 km x 2 km orbit of each lander has a period of just 20 minutes.
Therefore the time spent imaging each orbit will be less than 1 min. Over the five days in orbit the landers will pass
their respective landing sites more than 300 times however, so a suitable landing spot will be found in the necessary
timeframe.
The goal is to limit the potential landing zone to a 54 km x 54 km circle around the logarithmically spaced
landing point. (Fig. A.7) This will put the landing restriction well within the requirement given in the RFP. Due to the
Fig. A.7 Detail Mapping Diagram

large amount of data this will produce for each lander, and the short phase duration, the data will not be sent to the
satellites for processing. Instead, each lander will process its own data and determine its ideal landing site. The 2 km
periapsis of this phases orbit subjects the landers to much higher gravitational forces, which will require fuel to
counteract. This phase is only 5 days, therefore the amount of extra fuel needed is rather small. Despite this, a ΔV
budget of 35 m/s has been included for this orbital maintenance for this phase alone.
A.5.3 Descent
At the beginning of this stage in the landers operations, the landing sites while have been determined. On 6
November 2026, the landers will begin the descent phase, one at a time. Staring with the polar landers, each lander
will engage in the largest burn of the phase, the de-orbit burn. This burn cancels out most of the orbital velocity of the
lander, and occurs just before periapsis. The reason it does not cancel out all orbital velocity is to provide continued
forward motion in the event that an unforeseen obstacle lies at the intended landing site.
During the descent the Light Detection and Ranging (LIDAR), and MARDI will provide continuous data to the
lander to aid in obstacle avoidance, and ideal landing site location. There is no possibility of remote navigation for the
descent phase as the whole process from orbit to touchdown occurs in a span of just 91 seconds. As the MARDI
imager approaches the surface the resolution improves continuously, therefore any objects not detected from orbit will
be noticeable on the descent, and can be avoided using ACS. A scheme of the descent phase from orbit to touchdown
Fig. A.8 Lander descent depicting stages of hazard avoidance, leg deployment, and landing
Descent
(11/06/2026 - 11/07/2026)

is shown in Fig. A.8. The process depicted in Fig. A.8 will be covered in the ACS section of this report. Over the 91
second descent the lander must complete all steps in this sequence, or risk mission failure. The landing orientation and
placement are of high importance for the success of the mission. Should the lander touch down on a highly sloped
surface, it has the possibility of tipping. Should the lander only have two legs touch down the seismometer data would
be incomplete, as only parts of all three axes would recording due to the placement of the seismometers.
As stated previously, the descents will be happening one by one. That is, one lander performs its entire descent
phase before the next lander is cleared to begin its own. This phase of the mission is the most difficult and most crucial
to the success of the mission. Should one lander fail, the mission has failed according to the RFP. If a lander does fail
though, it might be beneficial to rearrange the locations of the landers slightly to achieve better coverage with the
landers which have yet to land. For this reason, the overall descent phase will begin with the polar landers landing at
sites L1, L7, and L6, in that order. L1 is crucial due to its placement at the North pole. The next closest landing site,
L2, is 130° longitude, and 60° latitude from L1, meaning any seismic activity close to the North pole will not be record
with great precision. L7 is important for mostly the same reason. Once the polar landers have landed, and transmitted
a health report to Earth, mission control will signal the start of the non-polar descent. The time period between
consecutive landings will be roughly 90 minutes assuming no problems occur. Most of this will be idle time waiting
for the status report to reach Earth, and then waiting for the authorization to proceed from Earth. Both signal require
about 37 minutes to travel to their destination.
The first landing site to be filled will be L2, followed by L3, L4, and lastly L5. The spacing between L1 and L2,
and between L2 and L3 are quite large, so having landers at L2 and L3 is critical. Should one of these landings fail,
another lander will need to take its place, or the landing scheme for the remaining non-polar landers will need to be
shifted to make up for the failure. A failure in landing is not an option for mission success however, so to ensure that
a failure during landing does not occur extensive testing of the software paired with the MARDI, and LIDAR will
need to be done in all possible landing scenarios.
A.5.4 Science Mission
Beginning on 7 November 2026, the landers will begin recording seismic activity, as well as taking pictures. Each
lander will have the opportunity every 3.55 days to communicate either satellite. Fig. A.9 illustrates the multiple flyby
concept again, in which the lander’s communication windows are labelled for each satellite. Each lander has been

assigned a segment of the orbit for communication that enables optimum signal transmission. Between each lander
communication segment the satellite has some time allocated to transmit data to DSN.
Due to the migration of the
orbit, explained in Section A.4 of
this report, Fig. A.9 serves as a
template for the communication
windows rather than a set in stone
plan for communications
throughout the mission. The
segments will have to migrate
around the orbit just as the orbit
migrates around Jupiter.
A.6 Disposal
The current disposal concept calls for leaving the landers on Europa. Disposing of them elsewhere is impossibly
expensive in terms of the addition fuel mass required. To ensure no contamination of Europa, the landers, and satellites
will be pre-baked, and will be maintained in clean rooms prior to launch. The disposal of the satellites calls for leaving
them in their flyby orbits. This allows for an easy extension of the mission, but also far less expensive than the
alternatives (discussed later), and is proven to not impact Europa for at least five years (hardware propagation
limitation) after mission end. Due to the migratory nature of the flyby orbit, the satellites never approach Europa closer
than 8000 km. Even on these approaches the satellites are generally well above, or well below the moon as well. The
flyby orbits were input into STK and run for five years with no close encounters. Over the course of a much longer
timeframe, the orbit is expected to decay to a point at which it would either impact Io (fairly unlikely) or drop beneath
Jupiter’s atmosphere.
The high radiation environment makes communications with the satellites unlikely after long periods of time
following mission end, therefore if a low ΔV disposal plan was desired, which did not impact Europa, communication
would likely be lost before the disposal was confirmed. Only a large ΔV disposal is possible in a limited duration,
meaning a complete redesign of the propulsion system, and the possibility exceeding the launch capacity of the Falcon
Heavy launch vehicle.
Fig. A.9 Uplink/Downlink Communication Windows

B. Trajectory Design
In choosing and designing a trajectory for the Richter Program it was necessary to minimize mission duration,
mission ΔV, launch C3, and total ionizing dosage (TID), while making sure to allow ample time for conceptual design
and manufacturing.
B.1 Trajectory Selection
Many types of trajectories were considered as a means of travelling to Jupiter. To meet the RFP’s operational
requirements, seismographic and optical science data must be transmitted before the start of 2027, meaning that the
majority of the trajectories under consideration were discarded due to long mission durations. Table B.1 shows a
selection of the most optimal Venus-Earth-Earth Gravity Assist (VEEGA), Venus-Earth Gravity Assist (VEGA), and
Earth Gravity Assist (EGA) trajectories.
After consideration of the mission task of emplacing seven landers on the surface of Europa, Option 5 was chosen
as the best candidate trajectory due to its low ΔV to JOI, and relatively low C3. These factors will yield a high payload
capacity. Option 5 also launches late enough to
provide to a 4.5 year conceptual design and
production window.
Another important benefit of the selected VEGA
trajectory is its early Jupiter arrival date of
December 2024. According to the Europa Study
2012 Report2
, the longer a spacecraft can stay in the
Galilean moon system, the lower its ΔV will be, at
the cost of higher TID. (Table B.2)
Table B.1 Consideration of several trajectory options on the basis of mission duration, ΔV to JOI, and
launch C3 [1]
Option # Type
Earth Departure
Date
Jupiter Arrival
Date
Time to JOI
(years)
ΔV to JOI
(km/s)
Launch C3
(km2
/s2
)
1 EGA 07/19/2020 07/19/2024 4.00 1.82 27.1
2 EGA 07/23/2020 01/27/2025 4.51 1.48 27.1
3 EGA 08/26/2021 08/26/2025 4.00 1.61 27.0
4 VEGA 11/24/2019 01/09/2025 5.13 1.73 15.6
5 VEGA 12/26/2019 12/01/2024 4.93 1.23 18.9
6 VEGA 03/08/2020 11/19/2025 5.70 1.69 26.1
7 VEEGA 03/14/2020 06/30/2026 6.29 0.88 11.5
8 VEEGA 03/22/2020 02/24/2026 5.93 0.86 9.8
Table B.2 Reductions in ΔV due to increased tour
length, with consideration for TID [2]
Tour Duration ΔV, JOI-to-EOI TID (Mrad)
0 >5.5 ~0
0.25 4 ~0
0.5 3 ~0
1 2.5 0.1-0.5
1.5 1.5 0.8-1.2
2.5 1.3 1.7

B.2 Launch Vehicle Selection and Launch Window
Due to the RFP requiring seven landers as well as a carrier satellite, launch vehicle selection is important. The
preliminary wet mass of the program was found to be extremely high, at around 13,000 kg; this, while implementing
mass saving technologies such as Flex-Rolled-Up Solar Arrays (FRUSA), and deployable HGAs. The enormous mass
made it impossible to use any of the standard launch vehicles currently in use for interplanetary travel. (Fig. B.1) The
selected VEGA trajectory has a C3 of 18.9 km2
/s2
, so based on the payload capability graph, the maximum possible
payload mass with current launch vehicles is only 7,500 kg, which is considerably below what is needed. Due to the
fact that the mass could not be reduced much more, it was decided that exploring less proven launch vehicles was
necessary.
Thus, the current launch vehicle option is the Space X Falcon Heavy. It boasts an impressive C3 of approximately
12,700 kg for a C3 of 18.9 km2
/s2
. (Fig. B.2). Even this was too little for the initial mass estimates for the satellite and
landers though, and even if small mass reductions were possible, the fact that the Falcon Heavy has yet to be launched
casts some doubt on the accuracy of the payload capacity curve in Fig. B.2.2. In order to maintain a higher mass
margin over the estimated payload capacity a dual-launch design was pursued while using a Falcon Heavy for both
launches. This allowed for redundancy in the design as well as ensuring positive mass margins. After completing mass
analyses on the two satellites, the wet masses were calculated to be 10,073 kg for Satellite 1, and 10,612 kg for Satellite
2. These masses include the mass of the lander payload for each satellite, and are well below the predicted launch
capability for the Falcon Heavy launch vehicle.
Fig. B.1 Payload capacity of currently used launch
vehicles3
C3 = 18.9 km2
/s2
Max Payload Capacity
= 7.5 Tonnes
Fig. B.2 Falcon Heavy estimated payload
capacity4
Max Payload
Capacity =
12.7 Tonnes
C3 = 18.9 km2
/s2

Each Falcon Heavy will launch with one satellite, each satellite carrying either three or four landers for a total of
seven. The first launch will occur October 14, 2019. This launch date may be moved several months earlier, or up to
four weeks later, however, the first launch date has been selected so as to provide sufficient time to prepare the launch
pad for the second launch on December 24, 2019. This launch has a window of one week beginning on December 22,
2019.
The first Satellite, along with the three polar landers, will optimally be launching on October 14th
, and will be
entering a 400 km parking orbit until the second satellite, with the four 60° inclined landers, launches on December
24th
. Once both satellites have achieved the 400 km orbit, they will embark on the same VEGA trajectory.
B.3 Interplanetary Trajectory
The trajectory being employed for both satellites is to be a VEGA trajectory. (Table B.3) Assuming the satellites
launch on the correct dates, no major burns will be necessary until September 2, 2022, approximately 35 days after
Earth Gravity Assist. This maneuver will ensure proper alignment for achieving Jupiter Orbit Insertion on November
26th
2024. Small maneuvers will be
needed to correct for any
perturbations caused by Venus
flyby, or cleanup from Earth escape,
however these burns are accounted
for in the ΔV estimates for each
satellite. Should the date of Earth departure be rescheduled, within the launch window, total mission ΔV could increase
by as much as 150 m/s.
Satellites 1 and 2 will have staggered Earth escape burns to ensure safe distance is maintained throughout
trajectory. Satellite 2 is planned to wait one full orbit after Satellite 1 to perform its Earth escape burn. This will have
a slight effect on total mission ΔV, but the effects will be negligible due to the extra orbiting time being less than two
hours.
Venus, and Earth flyby altitudes are rather low, but it is necessary for keeping mission ΔV low, and achieving the
2026 arrival date at Europa. Raising the altitude of the Earth flyby to 500 km increasing the required mission ΔV by
more than 600 m/s, therefore it was determined that the lower flyby altitude would be preferable.
Table B.3 VEGA trajectory interplanetary event summary
Event
Satellite 1
Date
Satellite 2
Date
V∞ or ΔV
(km/s)
Flyby Altitude
(km)
Launch 14 Oct 2019 24 Dec 2019 4.35 -
Venus
Flyby
5 Dec 2020 5 Dec 2020 8.32 357
Earth Flyby 16 July 2022 16 July 2022 13.58 200
DSM 2 Sep 2022 2 Sep 2022 .23 -
JOI 26 Nov 2024 26 Nov 2024 .95 12.8 Rj

The Jupiter Orbit Insertion for either Satellite is performed when the Satellite reaches its perijove of 12.8 Jupiter
radii. This distance was chosen so that for the first several orbits, each Satellite would be outside of the high intensity
radiation environment. Performing the burn at a lower altitude would have provided some ΔV saving as well as time
saving, however the radiation environment inside of Io’s orbit is extremely harsh. Another advantage of performing
JOI at this altitude is that an initial
Ganymede flyby may be performed about
fifteen hours after JOI. This Ganymede
flyby will reduce the apojove of the
Satellites orbit by the same amount that an
increase in ΔV of 450 m/s would, making
it crucial for reducing fuel mass. Table B.4 lists the major burns for the satellites, and gives a total ΔV for each satellite.
B.4 Jovian Tour (Satellites)
The general concept for the trajectory at Jupiter entails using the Galilean moons to slow down over the course of
about 2 years, to achieve an orbit similar in semi-major axis to Europa, but slightly offset. The logistics of this orbit
will be discussed later.
As previously stated, a first gravity assist maneuver will occur using the gravity field of Ganymede to reduce the
ΔV of the JOI burn by 450 m/s, as well as the Satellites initial period of orbit about Jupiter by more than 5 months.
Both Satellites encounter Ganymede for their first five gravity assists which serve to lower the period of revolution
from 143 days for the first orbit to just 15 days. After this point each Satellite encounters Europa, however after this
they divert. Satellite 1 performs 22 gravity assists of Europa, Ganymede and Io before entering its flyby orbit on
October 16, 2026, for a total pump-down phase duration of 1.89 years. (Table B.5) Satellite 2 performs 21 gravity
assists, only the first few being duplicates, and enters its flyby orbit on October 17th
of 2026, meaning its total pump-
down phase duration is equivalent to that of Satellite 1. (Table B.6)
Table B.4 Satellite 1 and 2 maneuver summary.
Maneuver Satellite 1 Satellite 2
DSM 238 m/s 238 m/s
JOI 950 m/s 950 m/s
Pump-down phase 142 m/s 65 m/s
Disposal 43 m/s 43 m/s
Orbit Maintenance 20 m/s 17 m/s
Reserve 67 m/s 37 m/s
Total 1457 m/s 1347 m/s

Table B.5 Detailed flyby and maneuver summary for Satellite 1
Phase
Flyby/Man-
euver
In/
Out
Date
Altitude(km)/
ΔV (m/s)
Period
(days)
TOF
(days)
Total TOF
(days)
Jupiter
Approach
JOI I 26 Nov 2024 07:20:14 ΔV = 950.3 300 - 0.00
Pump-
down
Ganymede1 O 26 Nov 2024 22:54:56 Alt. = 110 143 .6 .6
Ganymede2 O 19 Apr 2025 00:57:43 Alt. = 450 50 143 143.6
Ganymede3 O 8 Jun 2025 02:49:48 Alt. = 1000 28 50 193.6
Target G4 8 Jun 2025 03:37:49 ΔV = 0.38 - .03
Ganymede4 O 6 Jul 2025 18:24:31 Alt. = 2300 22 28 221.63
Perijove
Raise
16 Jul 2025 11:44:48 ΔV = 82.0 - 9.71
Ganymede5 O 28 Jul 2025 02:45:45 Alt. = 600 14 11.63 242.97
Europa1 I 11 Aug 2025 06:44:56 Alt. = 500 13 14.17 257.14
Target E2 11 Aug 2025 07:06:30 ΔV = 1.28 - .02
Europa2 I 5 Sep 2025 03:08:13 Alt. = 1200 10.5 24.83 281.99
Europa3 I 15 Sep 2025 18:45:56 Alt. = 440 8.8 10.6 292.59
Target E4 15 Sep 2025 19:07:22 ΔV = 0.47 - .02
Europa4 I 3 Oct 2025 12:49:18 Alt. = 250 7.4 17.7 310.31
Europa5 I 2 Dec 2025 21:59:13 Alt. = 450 6.5 60.4 370.71
Target E6 22 Dec 2025 23:32:03 ΔV = 4.8 - 20.0
Europa6 O 5 Jan 2026 12:37:56 Alt. = 350 5.7 14.5 405.21
Europa7 I 21 Jan 2026 11:55:06 Alt. = 270 4.7 15.9 421.11
Io1 O 22 Jan 2026 02:50:54 Alt. = 910 4.27 .63 421.74
Target I2 26 Jan 2026 06:57:38 ΔV = 0.05 - 4.16
Io2 O 12 Feb 2026 08:50:18 Alt. = 1500 3.75 17.08 442.98
Plane Change 13 Feb 2026 15:34:43 ΔV = 15.63 - 1.29
Io3 O 17 Mar 2026 23:30:18 Alt. = 450 3.08 32.33 476.6
Target I4 20 Mar 2026 20:30:56 ΔV = 11.9 - 2.87
Io4 O 8 Apr 2026 04:54:05 Alt. = 450 2.5 18.33 497.8
Europa8 O 19 Apr 2026 02:35:53 Alt. = 330 2.7 10.92 508.72
Target I5 19 Apr 2026 03:35:21 ΔV = 2.39 - 0.04
Io5 I 25 Apr 2026 12:31:47 Alt. = 2550 2.8 6.38 515.14
Europa9 I 29 Apr 2026 06:37:10 Alt. = 975 3.1 3.75 518.89
Target E10 10 May 2026 13:26:40 ΔV = 3.43 - 11.3
Europa10 I 23 Jul 2026 12:22:02 Alt. = 170 3.33 70.94 601.13
Target E11 23 Jul 2026 13:20:35 ΔV = 25.1 - 0.04
Europa11 I 14 Sep 2026 18:33:21 Alt. = 310 4.08 53.2 654.38
Target E12 18 Sep 2026 03:31:01 ΔV = 3.15 - 3.37
Europa12 I 16 Oct 2026 19:13:18 Alt. = 400 3.54 28.67 686.42
Fig. B.3 Satellite 1 tour diagram showing pump-down flybys of Ganymede, Europa and Io. Left: View
from Jupiter’s north pole. Right: View from Jupiter’s equatorial plane, with north pole towards top of
image.

Table B.6 Detailed flyby and maneuver summary for Satellite 2
Phase
Flyby/Man-
euver
In/
Out
Date
Altitude(km)/
ΔV (m/s)
Period
(days)
TOF
(days)
Total TOF
(days)
Jupiter
Approach
JOI I 26 Nov 2024 07:40:14 ΔV = 950.3 300 - 0.00
Pump-
down
Ganymede1 O 26 Nov 2024 23:11:56 Alt. = 110 143 .6 .6
Target G2 26 Nov 2024 23:42:56 ΔV = 1e-6 - 0.02
Ganymede2 O 19 Apr 2025 00:57:43 Alt. = 450 50 143 143.6
Target G3 19 Apr 2025 01:45:45 ΔV = 0.016 - 0.03
Ganymede3 O 8 Jun 2025 02:49:48 Alt. = 1000 28 50 193.6
Target G4 8 Jun 2025 03:37:49 ΔV = 5.8e-4 - 0.03
Ganymede4 O 6 Jul 2025 17:36:36 Alt. = 2300 22 28 221.63
Target G5 6 Jul 2025 18:24:15 ΔV = 9e-5 - 0.03
Ganymede5 O 28 Jul 2025 04:44:17 Alt. = 800 14 21.42 243.08
Europa1 I 11 Aug 2025 04:59:29 Alt. = 750 13 14.17 257.25
Target G6 11 Aug 2025 05:18:52 ΔV = 3e-3 - .01
Ganymede6 O 9 Sep 2025 04:48:44 Alt. = 6000 12.8 28.95 286.21
Target E2 9 Sep 2025 05:34:29 ΔV = 4.6e-4 - 0.03
Europa2 I 3 Oct 2025 09:32:43 Alt. = 500 10.6 24.17 310.41
Target E3 3 Oct 2025 09:51:31 ΔV = 5e-5 - 0.01
Europa3 I 14 Oct 2025 01:11:25 Alt. = 500 8.88 10.59 321.01
Target E4 14 Oct 2025 01:30:21 ΔV = 0.01 - 0.01
Europa4 I 31 Oct 2025 19:28:47 Alt. = 150 7.54 17.75 338.76
Target E5 31 Oct 2025 19:47:48 ΔV = 3.4e-3 - 0.01
Europa5 I 31 Dec 2025 04:23:16 Alt. = 280 6.92 60.63 399.4
Target G7 4 Jan 2026 03:51:13 ΔV = 63.9 - 4
Ganymede7 I 23 Feb 2026 13:00:48 Alt. = 770 5.75 44.42 447.82
Io1 I 19 Mar 2026 08:40:07 Alt. = 600 5.13 23.79 471.61
Target I2 19 Mar 2026 08:57:31 ΔV = 0.9 - 0.01
Io2 I 9 May 2026 15:59:06 Alt. = 4000 4.79 51.29 522.91
Target I3 9 May 2026 16:11:18 ΔV = 0.064 - 0.01
Io3 I 26 Jun 2026 10:20:25 Alt. = 390 3.88 47.75 570.67
Io4 I 15 Jul 2026 21:25:17 Alt. = 440 3.25 19.46 590.13
Target I5 15 Jul 2026 21:41:43 ΔV = 1.2e-4 - 0.01
Io5 O 1 Aug 2026 14:10:50 Alt. = 640 2.94 16.7 606.84
Target I6 1 Aug 2026 14:28:23 ΔV = 3.1e-4 0.01
Io6 O 10 Aug 2026 09:53:04 Alt. = 310 2.45 8.79 615.63
Europa6 O 18 Aug 2026 08:24:34 Alt. = 230 3.05 7.96 623.59
Target E7 18 Aug 2026 09:04:37 ΔV = 1.1e-5 - 0.03
Europa7 O 8 Sep 2026 16:59:34 Alt. = 390 3.3 24.89 648.51
Europa8 O 17 Oct 2026 21:46:44 Alt. = 380 3.54 39.2 687.72
Fig. B.4. Satellite 2 tour diagram showing pump-down flybys of Ganymede, Europa and Io. Left: View
from Jupiter’s north pole. Right: View from Jupiter’s equatorial plane, with north pole towards top of
image.

B.5 Lander Trajectory
Each Satellite carries a specific type of lander. Satellite 1 carries three polar landers, while Satellite 2 carries four
non-polar landers. The only real difference between the two types of landers is the amount of fuel being carried, as
the polar landers will need extra fuel to place themselves into polar orbits around Europa. Due to the differences in
their orbits around Europa, as well as the fact that they are on different satellites, they need different trajectories.
The polar landers will be separating from Satellite 1 on 15 October 2026, the day before Satellite 1 is scheduled
to perform its final pump-down flyby of Europa. This will allow for a low burn of about 50 m/s to achieve a 90°
inclination when approaching Europa, compared to almost 200 m/s extra which would be added onto the Europa Orbit
Insertion (EOI) burn to achieve a combined plane change. Once the landers have reached their periapsis about Europa
of 200 km, they will perform an EOI of 1600 m/s. The value for EOI is rather high, especially compared with other
missions attempting a Europa lander. The reason it is so high is due to the fact that the burn incorporates matching
Europa’s angular velocity with respect to Jupiter, as well as slowing down to the appropriate circular velocity of 1.349
km/s. Other major burns for the landers include a periapsis lowering burn to enter into a 200 km x 2 km orbit around
Europa, as well the horizontal velocity cancelling burn to enter the descent phase, as well as several descent burns to
ensure a smooth, soft landing.
The non-polar landers will enter 200 km circular orbits similar to the first landers, however instead of their
inclination being 90°, it will be 60°. Satellite 2, which carries the non-polar landers, has been set up to perform its
final pump-down flyby of Europa on October 17 2026, at a 60° inclination, therefore the non-polar landers may be
dropped off much closer to Europa than the polar landers were. Satellite 2 will drop the landers off between 5000 km
and 300 km away from Europa, with about an eight minute delay between consecutive launches to ensure safe
distances between landers. After the landers have been deployed, each follows a similar path to the polar landers. A
ΔV summary for both types of landers is shown in Table B.7.
Table B.7 Polar and non-polar maneuver summary. Some values vary slightly between
individual landers, so they are shown as approximate values.
Maneuver
Polar Landers Non-polar Landers
Date ΔV (m/s) Date ΔV (m/s)
Pre-arrival Plane Change 15 Oct 2026 ~50 - -
EOI 16 Oct 2026 ~1600 17 Oct 2026 ~1600
Lower Periapsis 3 Nov 2026 42 4 Nov 2026 42
De-orbit 16 Nov 2026 1432 17 Nov 2026 1432
Powered Descent 16 Nov 2026 72 17 Nov 2026 72
Total 3296* 3246*
* ΔV values include 100 m/s reserve for descent, clean-ups, and orbital maintenance

B.6 Satellite Disposal
As the landers are actually landing on Europa, it would be incredibly expensive to dispose of them off-moon.
Therefore, the landers will remain on Europa. Communication will end when radiation dosages become too high for
the sensitive instruments sometime after the 90 day operational period.
To comply with planetary protection, the satellites are engaged in a multiple-flyby trajectory. Each encounter
Europa about once every orbit. The encounters are never closer than 10,000 km however. For this reason the disposal
plan for the satellites is to leave them in their flyby orbits.
Using STK, each satellites flyby orbit was propagated for five years. At no point in the five year propagation did
either satellite approach Europa closer
than the previously mentioned 10,000
km. The reason for this is the resonance
of the flyby orbits with Europa.
Because both satellites are in orbits
with periods of 3.54 days, whereas
Europa’s is 3.55 days, the satellites
encounter the edge of Europa’s gravity,
rather than the center. This provides the
satellite with a gentle nudge, so that in
subsequent orbits the satellite would be
moving away from the moon. Over time the orbit is expected to slowly decay to the point that it will crash into either
Io or Jupiter, both of which have lower scientific value than Europa. This will take years, by which time the radiation
environment will have rendered the satellites communication systems useless.
Another method of disposal is to use gravity assists to aid in the disposal of the satellite at Jupiter. An attempt at
creating this type of disposal was made, but it proved extremely expensive in terms of ΔV. It also needed more than
two years to even begin disposal, by which time the radiation could have already fried all necessary components for
communication.
Lastly, over the course of the five year propagation mentioned before, each satellite will have a pass of Europa
roughly 500 times. According to the Office of Planetary Protection, “Requirements for flybys, orbiters, and landers to
Fig. B.5 Satellite flyby orbits after five year propagation. Neither
of the orbits change drastically from one pass to the next, and both
maintain similar periods throughout.

icy satellites, including bioburden reduction, shall be applied in order to reduce the probability of inadvertent
contamination of an ocean or other liquid water body to less than 1 x 10-4 per mission”[4]. Technological restrictions
restricted STK from being able to propagate further than five years, however based on the pattern outlined above
regarding Europa’s effects on the satellites, it is dubious that either satellite would crash on Europa even in 1000
passes. Obviously contamination is still a concern, however a more powerful computer is needed to run the flyby orbit
simulation.
B.7 Alternate Trajectory
Despite the lengthy trade study that was conducted to find the optimal trajectory for this mission, there are still
issues with the chosen trajectory. The launch dates of October 2026, and December 2026 are only 4.5 years from the
time of submitting this proposal. With seven landers and two satellites needing to be manufactured, and new
technologies to be implemented this launch date will be a struggle to meet. The other issue is the reliance of the
mission on a launch vehicle which has to be launched, and which will not be launch until 2018 at the earliest [6].
The most beneficial alternate trajectory would be to proceed with Option 8 from Table B.1. This is a VEEGA
trajectory which launches four months after the original trajectory giving more time for production. This trajectory
also has a much lower C3 of 9.8 compared with 18.9, meaning increased payload capacity with all launch vehicles,
and it has a lower ΔV to JOI, meaning less fuel mass. The lower C3 increases the payload capacity of the Delta IV
Heavy to about 9,300 kg, and due to a ΔV decrease of 350 m/s, the total wet launch mass of the satellites with landers
are 8,595 kg, and 9,094 kg for Satellites 1 and 2 respectively. Compare that to their wet masses with the current VEGA
trajectory (Satellite 1 = 10,073 kg, Satellite 2 = 10,612 kg). The reason this trajectory cannot currently be implemented
is the Jupiter arrival date of February 24, 2026. The current trajectory arrives two years early, and needs to in order to
lower its orbital energy using a minimal amount of fuel. This alternate trajectory would require much more fuel to
slow down than the VEGA did, which would like push the mass margins for the Delta IV Heavy into the negatives.
Assuming the ΔV was kept the same from the VEGA trajectory to this, the landers would not start transmitting data
until early 2028.
This trajectory is recommended in order to alleviate the risk associated with launching on the Falcon Heavy,
however it would require an extension of over one year of the mission duration outlined in the RFP. This is a reasonable
request as the Europa Clipper mission is not planned to arrive until the early 2030’s.

C. Payload and Instrumentation
The satellite payloads include an optical instrument package, a laser altimeter, and a magnetometer. Satellite 1,
which carries three polar landers has this entire payload suite. Satellite 2 only has a magnetometer because it transports
four landers, which create volumetric constraints on payload placement. Payloads for both lander types (polar and
non-polar) include optical payload package and seismometer.
C.1 Satellite Instrument Overview
Satellite 1 includes an optical instrument package comprised of a scaled-down HiRise camera and the Mars Color
Imager (MARCI) camera, and also the Mercury Laser Altimeter (MLA) and magnetometer. The payload on Satellite
1 is used to achieve the following scientific and engineering objectives: (1) observe surface features of Galilean moons,
especially Europa, (2) generate topographical map and surface profile of scientifically interesting areas of Europan
surface, (3) observe magnetic field interaction between Jupiter and its four major satellites, and (4) photograph the
landers’ landing sites (where possible) to provide locational context for seismic activity data. Satellite 2 will only be
responsible for transmitting information on magnetic field interaction in the event that Satellite 1 fails. It must be
noted that the primary objective of Satellite 2 is not to satisfy scientific needs through passive observation, but by
ensuring the safe transportation and deployment of its four non-polar landers. A margin of 30% is allocated for direct-
to-Earth (DTE) transmission data rates during the 90-day operations phase of the landers. This allows lower priority
scientific data obtained by the satellites (such as data on Jovian magnetosphere, and images of the landers) to be
transmitted alongside higher priority lander data.
C.1.1 Satellite 1 Optical Instrument Package
The HiRise and MARCI camera on Satellite 1 have been equipped on the Mars Reconnaissance Orbiter. This
optical package was selected primarily for preliminary terrain mapping of Europa’s surface prior to lander deployment
and lander mapping phase, and is used in conjunction with the Mercury Laser Altimeter (MLA). Deliverables for this
package throughout the course of the satellite lifespan include the generation of topographic or elevation maps of
Europa’s surface and possibly the surfaces of other Jovian satellites during the Jovian tour/pump-down phase. Unlike
the MLA, which is used primarily for preliminary mapping (during Jovian tour), the HiRise and MARCI cameras will
be used for the entirety of the satellite operations phase. Images taken by the HiRise camera during the 90-day lander
mission operations phase will not all be transmitted directly to Earth. Instead, these images will be stored in the solid-

state recorder, and will be transmitted sparingly due to the large volume of data. Images from the MARCI camera will
be transmitted more frequently, from the time of capture during preliminary mapping to satellite disposal.
The HiRise camera on the Mars Reconnaissance Orbiter (MRO) is a reflector telescope which allows for a
resolution of 0.3 meters/pixel at 300 km altitude, and 10 km swath at 200 km altitude. It is estimated from the Europa
Study 2012 Report that 0.5 m/pixel resolution at a 200 km altitude would suffice for mapping, especially given that
the HiRise is used for preliminary mapping and at 0.3 m/pixel resolution would take approximately twice the volume
of data. Due to its high resolution imagery, the HiRise will also be used to view the non-polar landers during the
closest flyby approaches of Europa after lander deployment, but before the satellite increases its periapsis to its 90-
day operational orbit. The MARCI camera was also selected as a means to map regions at a lower resolution, so that
interesting regions could be down-selected for mapping using the HiRise during subsequent flybys. The satellites will
contain the wide-angled (WA) MARCI camera, while the landers, as will be discussed in the lander optical payload
section, will contain the medium-angled (MA) MARCI camera. Due to radiation sensitivity, the existing
configurations of the cameras in MRO are not planned to be operational beyond the 90-day mission at Europa.
Radiation mitigation plans include moving the primary computing/processing and flash storage devices on these
cameras into the radiation vault where possible. Fig. C.1 below shows the optical payload package, and Table C.1 lists
the information.
(a) HiRISE Camera [7] (b) MARCI camera (left-MA, right-WA) [8]
Fig. C.1 HiRISE and MARCI cameras

Table C.1 Optical Payload Package Specifications
MRO HiRISE MARCI
 Resolution: 0.3 m per pixel at 300 km
 Narrow Angle, Push-broom Imager [9]
o 40,000 pixel width
o FOV = 1.1o
o Focal length = 12 m
 SNR > 100
 Data precision: 14 bit ADC
 Data Storage: 28 Gbits
 Spectral range: 400 – 600 nm, 550 – 850 nm,
800 – 1000 nm [9]
 Used for High-Res mapping of landing sites
 FOV = 1.14o
x 0.18o
[9]
 IFOV = 1 x 1 μrad
 Two types/modes: Wide-Angle (WA) & Medium-Angled
(MA) [10]
o WA
 5 visible & 2 UV spectral bands
 Resolution of 1 to 10 km per pixel at 400 km
 FOV = 140o
o MA
 8 spectral bands between 425 and 1000 nm
 40 m/pixel at 400 km altitude
 FOV = 6o
o Both cameras  1000 x 1000 pixel images
 Low mass: 0.527 kg (WA), 0.510 kg (MA) [10]
 Low volume: ~6 x 6 x 12 cm
 Low resolution = less data
o Reduces uplink data rate during mapping
 Electronic shutter that changes from transparent to opaque
when voltage is applied
C.1.2 Satellite Laser Altimeter & 3-Axis Fluxgate Magnetometer
The laser altimeter in Satellite 1 is used least frequently of all its payloads. It is only meant for obtaining an
elevation map of Europa so that engineers can evaluate and select landing sites during the pump-down/preliminary
mapping phase before lander deployment. It is switched on when encountering Europa less than 800 km in range. It
is not planned to be used during the 90-day lander mission unless required by the scientific community.
Satellite 1 and 2 also contain a magnetometer, modeled on the Galileo magnetometer (MAG). The Galileo MAG
was chosen over the magnetometer used in the JUNO mission due to lower mass. Mass was the primary criteria for
the magnetometer as it was to be placed at the end of the flex-rolled up solar array (FRUSA). Increasing the mass
would increase solar array flexure during ACS maneuvers. A separate boom was considered, but not used for the
magnetometer as it serves as another obstacle during lander deployment. The magnetometer was incorporated to
enhance the current understanding of Jupiter’s magnetosphere, to understand magnetic perturbations, and to expand
on Galileo’s discoveries. Due to mass and volume constraints, Satellite 2 will only contain the magnetometer as part
of its scientific payload (aside from its four non-polar landers).

(a) MLA [11] (b) 3-axis Fluxgate Magnetometer [12]
Fig. C.2 Altimeter and Magnetometer
Table C.2 Laser Altimeter and Magnetometer Specifications
MLA 3-axis Fluxgate Magnetometer
 For surface profile and topography measurements
o To identify terrain slope meeting landing criterion
(terrain slope < lander tipping angle)
 Error: 1.0 m when line-of-sight < 1,200 km [13]
 Probability of detection > 95% at 200 km nadir-
pointing; > 10% at 800 km slant range [13]
 May need to be modified for reflectivity/light
diffraction on Europa’s icy surface
 Dynamic Range: 1024 nT [12]
 Sensitivity: 0.03 nT
 Sampling rate: 16 Hz [12]
 Long time drift: < 0.3 nT/o
C
 Noise: ~40 pT [12]
 Similar to DTU Space, National Space Institute’s
3-Axis Fluxgate Magnetometer
C.2 Lander Instrument Overview
The lander payload is used to achieve the following scientific objectives: (1) observe seismic activity, and thereby
identify internal structure and composition of Europa, (2) observe local surface activity on Europa, and (3) photograph
local Europan terrain and surface features at variable locations. The lander payload includes an optical instrument
package and a MEMs seismometer. The optical instrument package is composed of the Beagle 2 Stereo camera, two
MARCI cameras, and the MARDI descent imager, of which the latter two are used during the initial and detailed
mapping phases. The Beagle 2 stereo camera and MEMs seismometer are used during the 90-day mission operations
phase as required by the RFP. The payloads remain the same for both polar and non-polar landers.
C.2.1 Lander Optical Instrument Package
The optical payload for the polar and non-polar lander is used during mapping, descent, and scientific operations.
Because of its usage in wide range of critical mission phases (especially detailed mapping and descent), it was essential
that the optical instruments have redundancies in quantity, and proper placement.
The medium-angled MARCI cameras are used primarily for the initial mapping phase as specified in the concept
of operations. It is used for mapping seven bands around Europa around logarithmically spaced latitudes specified by

the RFP. Ten percent of the down-selected 540 km landing sites are then further mapped by the Mars Descent Imager
(MARDI camera during the detailed mapping phase. This corresponds to 54 km diameter region mapped with a
resolution of 1.5 m per pixel. The two MARCI cameras serve as redundancy during this detailed mapping phase if the
MARDI camera fails. The MARDI and MARCI cameras are also used for Hazard Detection (HD) during the deorbit,
descent, and landing (DDL) phase. It must be noted that the MARDI camera, despite being a descent imager used
during the landing phase of the Mars Science Laboratory (MSL) Curiosity rover, is viable as a mapping camera for its
variable resolution and large data storage. It has not been used before for terrain mapping alone. Thus, the MARDI
needs to be adapted for this mission as a mapping camera as well.
The Beagle 2 camera serves as the primary imaging payload used during the 90-day scientific mission phase of
the landers. It is a wide-angled, colored camera as required by the RFP. It was selected for its sensitivity to both the
visible and infrared spectrum, wide field of view of 48o
, variable focusing from 0.6 m to infinity, and moderate imaging
resolution of 1024 by 1024 pixels. The large field of view and moderate resolution allows for lower data rates in
comparison to MER Panoramic Camera (PanCam), without significantly sacrificing image quality. This camera is set
atop a helical boom found in the Mars Pathfinder rover, which uses a one-time deployment mechanism. The camera
and helical mast are stowed in a radiation shielded canister during cruise and up to lander touch-down on Europa’s
surface. Drive motors exist on the camera platform for both panning and tilting. This allows for creating a mosaic at
every 4o
of solar elevation at Europa with at least 2π steradian coverage. The total images captured by the Beagle 2
camera during the duration of the 90-day mission is 1440 pictures to satisfy this RFP requirement. Figure C.3 and
Table C.3 provide images and key specifications of the lander optical payload.
(a) MARDI [C8] (b) Mars Pathfinder Helical Boom [C9]
Fig. C.3 MARDI and Helical Boom

Table C.3 MARDI and Beagle 2 Camera Specifications
MARDI Beagle 2 Camera
 Compact, Wide angled, refractive camera [16]
o For detailed mapping
 Resolution: 1.25 mrad/pixel, 1000 x 1000 px [16]
o 1.5 m/px at 2 km, 1.5 mm/px at 2 m altitude
 Panochromatic electronically shuttered CCD
 Image capture rate: 50 images/second
 Resolution: 1024 x 1024 pixels
 Spectral range: 440 – 1000 nm [17]
 FOV = 48o
[C11]
 24 filters
 A/D conversion: 10 bits/pixel [17]
 Pixel size: 14 μm x 14 μm
C.2.2 Lander Seismometer Instrument
The primary instrument for the lander, and arguably the entire mission, is the seismometer. Two possible
seismometers were considered: a commercial-off-the-shelf (COTS) MEMs seismometer, and the Mars Insight mission
SEIS instrument. Due to the importance of this instrument, and the lack of redundancy in landers, it is necessary that
the selection of this payload be discussed. Table C.4 presents the highlights of the conducted trade study.
Table C.4 MEMs and Mars Insight SEIS Seismometer Comparison
Silicon Audio GeoLight 7 MEMs
Seismomter
Mars Insight SEIS Instrument
Advantages
 Small packing factor (single axis chip is
2mm x 2mm)  possible to place in
lander “feet”/legs
 100 mHz to 100 Hz flat response [18]
 Low noise floor of 1 ng’s/√Hz  noise at
low freq [C7, C13]
 Low power  25 mW/channel [18]
 No attenuation between 0.1 and 100 Hz
 Low power consumption  ~1 W
 10-3
to 10 Hz flat response [20]
 Low noise floor  -9 m-s-2
/√Hz
 Contains 3 Very Broad Band (VBB) probes,
and 3 Short Period (SP) seismic probes, and
temp. sensors [14]
 In production, and to be used space qualified
through Mars Insight mission
 Flight-ready flight software by CNES [20]
Disadvantages
 Currently not in production by Silicon
Audio
 MEMs chips may be susceptible to
radiation environment prior to landing
 Not space flight qualified
 Unknown radiation tolerance
 Large volume (~ 1 ft3
)
 Only tested for low radiation exposure (15
krad) [21]
o Adding radiation shielding increases mass
 Large mass  3 kg [20]
The Silicon Audio GeoLight 7 MEMs seismometer was selected and incorporated into the payload package
due to its small packing factor, ability to gauge short-period and broad band frequencies, low noise floor, and low
power. Although this seismometer had the disadvantage of not being in production, this can be mitigated by
duplicating or purchasing the technology from Silicon Audio. Additionally, the small size of this seismometer as
shown in Fig. C.4 will allow it to be placed inside of the base (or “foot”) of the lander’s legs. With four legs on the
lander, and a single, three-axis MEMs seismometer inside each of the leg’s base, the lander will have three redundant
seismometers to use. Thus, at minimum, only one leg needs to have good “footing” or inertial coupling with the

Europan surface to be able to read data. This seismometer also expedites the manufacturing, testing, and
implementation phases for all seven landers as it does not contain mechanical assemblies, and does not require a
complex deployment mechanism (aside from lander leg extension). This seismometer chip will be rad-hardened and
also protected from radiation by the thick aluminum metal on the lander legs.
C.3 Payload Summary
Table C.5 lists the mass, power, and operating temperature statements for the selected orbiter and lander payloads.
It must be noted that the operating temperature requirements for selected payloads, such as the MEMs seismometer
and the HiRise will be expanded beyond the range allowed by their technologies to meet environmental constraints.
Spacecraft Payload Mass (kg)
Power Consumption
(W)
Operational
Temperature
Requirement (o
C)
Satellite 1
HiRISE
35 (reduction
from 65)
38 -10 to 20 (11)
MARCI (WA) 0.527 3 (12)
-40 to 70
MLA 7.4 23 -15 to 25 (13)
Satellites 1 & 2 MAG 4.7 4 -30 to 60
Polar and Non-
Polar landers
MARCI (MA) 0.51 312
-40 to 70
MARDI 0.6 10 -40 to 70
Beagle 2 Cam &
Helical Boom
5.5 5.6 -150 to 100
MEMs seism. 0.25 ~1 -200 to 10
11
Requires advancement in technology to increase operating temperature requirement from current 0 to 20o
C range.
12
Only during imaging. ~2 W during standby
13
Advancement in tech. assumed to decrease lower-end of optimal operating temperature to -15o
C from current 15o
C.
Fig. C.4 Silicon Audio GeoLight 7 MEMs
Seismometer [C12]

D. Structural Design
D.1 Satellite Mechanical Design
The goals of the satellite design process was to develop a spacecraft
that could act as a carrier craft for the seven landers to be placed on the
surface of Europa, while also acting as the primary communication and
data interface for the landers. The large payload and lander deployment
sequence drove the structural and power requirements. The large payload
of seven landers required a large structure capable of maintaining its
integrity under launch loads, which amount to approximately 7 gees
actual, or 9 gees with a safety margin. The mass restrictions placed on
launch payloads by launch vehicles with a C3 greater than 30 pushed the
design towards a modular design that could be spread across two
spacecraft and therefore decrease the payload carried on a single
spacecraft. The two craft system carries three polar landers and optical
equipment on one craft and four non-polar landers on the other.
The structure of the satellite is conformal to the carried propulsion tanks,
which are the primary volume constraint. The frames mounted on the outside
of the structure are designed to be mounts for the landers, as can be seen in
Fig. D.1 and Fig. D.2. In the assembled configuration, the top panel of the
lander is bolted to the primary structure, and internal brackets move launch
loads due to the lander through the panel into the structure. These loads are
then passed onto the launch fairing itself. The structure is constructed
through the use of several key technologies, including spin-forming,
hollowing and large scale CNC milling. The central cylinder is made by spin-
Fig. D.1 Lander 1 Loaded Cruise
Configuration
Fig. D.2 Conformed Structure

forming, and the structure is then hollowed to remove mass, forming an isogrid structure. The
brackets are milled to fit the contour and bolted to the primary structure (bolts not pictured). The
lip bracket designed to hold the lander will, along with a lengthwise bracket (see Fig. D.1) and
blast bolts (not pictured), support launch loads. Deployment is conducted by blast bolts which both
detach the lander and separate it from the primary satellite structure. This distance allows the lander
to trigger its propulsion system without effecting the attitude of the satellite.
The design of the satellite was also driven by the difficulty of ACS on missions of this duration,
It was imperative that the CG of the spacecraft shift as little as possible over the course of the
mission. The structure is therefore internally symmetrical, and the propellant tanks are arranged
around the center of the structure. As the propellant tanks empty therefore, the CG is driven by the
payload mass, and shifts slightly away from the unloaded side of the structure of Satellite 1, and
stays extremely central for Satellite 2. This is pictured in Figure D.4 for Satellite 1 and Figure D.5
for Satellite 2. This design optimized the ACS control requirements, and therefore increased the
likelihood of mission success. Serious attention was also paid to the possibility that the plume from
the ACS thruster clusters may impinge upon the deployed solar arrays. To avoid this the thruster
clusters were designed without upward facing thrusters, so that any ACS burn will require the firing
of two clusters, but there will be very little interaction between the arrays and the plumes except in
the most rapid of maneuvers.
Fig. D.3 Deployed
configuration
Fig. D.4 Wet and Dry CG Locations of Satellite 1 Fig. D.5 Wet and Dry CG Locations of Satellite 2

D.2 Environment
The environment encountered during the cruise and particularly the Jovian
tour portions of the satellite trajectory will be harsh. Extreme thermal gradients
and powerful radiation fields are the two greatest dangers. The satellite was
designed to provide the maximum amount of protection to its payload during
this period. The most sensitive part of the spacecraft are the internal electronics
of the landers, and the telecommunications and power equipment inside the
spacecraft. Neither of these will survive without adequate protection, so the
spacecraft was designed to supply as much integrated protection as possible.
The propellant tanks were placed around the electronics vault so as to provide
protection from the radiation environment, which not only provided nearly all
the required protection, but allowed the vault to be made much lighter than
would otherwise be possible. This was most useful in the lander design,
discussed in detail later in section D. The propellant tanks also act as thermal insulators during the Venus flyby, where
surface temperatures of the satellites are in excess of 320°K. All electronics are extremely vulnerable at these
temperatures, however, the propellant is in its most useful state at above 250°K and below 380°K. This means the
tanks are an ideal insulator for the electronics during hot periods. During cold periods, such as when the spacecraft is
eclipsed by Jupiter while doing its series of Europa flybys, the tanks will again serve as insulation for the vault, by
reradiating the heat produced by the RHUs which are placed directly on them. This minimizes the number of
Radioactive Heating Units (RHUs) required and minimizes cost and mass.
D.3 Analysis
The analysis on the satellite was run on CATIA’s Generative Structural Assembly Analysis module, with a
conformal node mapping system which was quality checked for aspect ratio, skewness, and Jacobian. The solver used
was the Elfini solver, which tracked solution convergence, along with solutions for displacement, stress, nodal energy
and frequency. These solutions were calculated for several sets of conditions. Longitudinal loading was applied to the
top of the spacecraft, with a 5 gee load (safety factor of 1.5) and a 9 gee load (safety factor of 4). Under 5 gee loads,
the spacecraft had no points of failure stress. However, the load paths were apparent, and the payload attach fitting
points were placed to coincide with the termination of these paths. This minimized the absorbed strain energy in the
Table D.1 Vibration Analysis

structure. The spacecraft was then analyzed with three lateral loads: 3 gee, 5 gee, and 9 gee, or safety factors of 1.5, 4
and 7.5. Under the moderate loading of 3 gees, there was again no points of stress that indicate failure. However, the
load paths were again analyzed to ensure that supports were placed at the termination points of the load paths. For
each of these conditions, displacement, strain energy and principal stresses were analyzed.
The fixed base normal mode frequencies were analyzed, and are presented in table D.1. A sample of the results
of the displacement solution for a loading
scenario of vertical takeoff with no lateral
loading is also presented in Fig. D.6. The
results of this analysis were that the
overall structure would provide
satisfactory safety margins for the
payload and launch system.
D.4 Evolution of the Lander Design
When initially developing the
shape and structure of the lander, a few
ideas were considered. One idea was to have a soft-
lander with movable legs to conform to the surface
terrain of Europa, and a central body in which to house
all of the necessary components. The very first model
consisted of a tripod configuration, with the main body
elevated off the ground, shown in Fig D.7.
Another idea that was considered was a cube
lander, with rigid legs attached to each of the eight
corners of the cube. This too would be a soft lander, but
would utilize reaction wheels for attitude control during
landing, as well as being a possible means of mobility on the surface of Europa. By loading the reaction wheels and
Fig. D.6 Top Loading Displacement Solution
Fig. D.7 Initial Legged Lander Design

then quickly unloading them, the lander could tip onto its side,
allowing it to move around if necessary. The first model of the
cube lander is shown in Fig. D.8.
When reassessing each of these designs, it was determined that
the center of gravity of the legged lander was much too high, and
posed a considerable risk of the lander tipping over. Also, a larger
base area within the body was needed in order to store and protect
many of the electrical components and to lower the center of
gravity. Apart from these design flaws, it was decided that the legged lander was still a suitable candidate for the
final design.
The cube lander, however, was decided against, mainly because of its reliance on reaction wheels to function.
Failure mode analysis conducted on the cube landers ability to traverse the uneven terrain determined that instead of
trying to correct for any errors after the lander has touched down, it would be less risky if a suitable landing site was
determined prior to touchdown. For this reason, the cube lander was decided to not be a suitable candidate for the final
design.
When redesigning the legged lander, the first design drivers were to lower the center of gravity, protect sensitive
components from radiation, and to allow for a maximum packing factor for all of the internal components. Three
designs that came from these drivers were a plus-shaped lander, a square lander, and an octagonal lander. For each of
the three designs, the propellant and pressurant tanks were to be used as radiation protection for the internal electrical
components. The tanks were spheroids in shape and were placed around the sides of the electronics vault, shown on
the plus and square landers in Fig. D.9. The initial seismometer that was to be used in the mission was the SEIS
Prop.Tank
Fig. D.8 Initial Cube Lander Design
Fig. D.9 Plus, Square, and Octagonal Lander Designs

seismometer. Using the SEIS severely limited the packing ability, because of its large, round shape, but was used
because no other instrument was determined to perform the functions necessary for the mission.
The plus lander was designed so that the components could be compartmentalized in each of the arms of the
plus. This way, radiation sensitive components could be protected as needed, science payload could have access to
the surface of Europa, and non-radiation sensitive materials would not require the extra mass to protect, each
independent of one another. The square lander was created as a way to reduce the width of the plus lander, and to
centralize all of the components. Although the overall dimensions of the square lander were smaller than the plus
lander, the packing efficiency was lower. The octagonal lander was created to increase the packing factor of the
lander, and was overall the best choice because of its smaller size, lower structural mass, and more central and
evenly distributed component mass.
Next, two major design changes were implemented. First, the spheroid propellant tanks were replaced with torus-
shaped tanks. This change greatly increased the effectiveness of the tanks in protecting the sensitive electrical
components from radiation. The sensitive electrical components were placed into a vault in the center of the toroidal
propellant tanks, which also greatly increased the packing factor. The second design change was the use of the MEMS
seismometer instead of the SEIS. Because of the great reduction in size, the seismometers could be taken out of the
body of the lander and placed into the legs. Placing the seismometers in the legs of the lander allowed for better contact
with the surface of Europa, and therefore better seismographic readings. It also freed space within the body of the
lander allowing the size and mass to be reduced. After these changes were implemented, the configuration was
finalized with the major features of the lander being a legged soft-lander with an octagonal shape, with toroidal
propellant tanks, a centrally located electronics vault, and MEMS seismometers located within the legs. A more
detailed description of the final design is given in section D.5.
D.5 Structural Design of Polar Lander
The polar lander was designed to land on or near the poles of Europa to collect seismographic data and take pictures
of its surroundings illustrated in Fig. D.10. The main design and dimensions depended on the size of the propulsion
and pressurant tanks. Given the volume of the toroidal tanks to be 0.19430 m3
and pressurant to be 0.02839 m3
, the
tanks were designed to meet these volumes while maintaining a reasonable size to fit inside the lander body. In order
to be able to fit the tanks, the lander body was designed to have a width of 1.260 m and a height of 0.757 m.

Fig. D.10 Polar lander final product
The most important payload of the lander are the MEMs seismometer and the camera in Figure D.11. The MEMs
seismometers are located on the foot of the leg. Three of the seismometers measure one axis for the required seismic
waves and the fourth one is for redundacy. The seismometers will be installed at angles so that any three seismometers
will act in conjunction to provide the 3 axes of measurement required. The camera is extended with a helical boom
between the pairs of pressurant tanks and is mounted above the radiation vault.
Fig. D.11 Polar lander important payload
Due to extreme exposure to radiation, the polar lander was designed to protect the electronics and other delicate
instruments in layers. The first layer in the body which includes 1.0 mm thickness of Aluminum and 0.5 mm of
Polyethylene. The top panel of the body includes the same materials but instead has 2.2 mm of Aluminum and 3.5
mm of Polyethylene. The next layer of protection are the toroidal tanks to protect the sides, which are made of Titanium
and have a thickness of 0.65 mm. The pressurant tanks are designed to have a capsule shape to better fit inside the
lander body and are also made of Titanium with a thickness of 3.81 mm. The pressurant tanks are mounted on top of
the toroidal tanks to protect the electronics from the top as illustrated in Fig. D.12.

Fig. D.12. Propulsion and Pressurant Tank Layout
Finally the last layer of protection is the radiation vault which contains the electronics inside and is surrounded by
the propulsion and pressurant tanks. The design of the radiation vault is a cylinder which is 410 mm tall and has a
radius of 320 mm. The sides of the radiation vault are made of 0.1 mm of Copper and 0.5 mm of Titanium. The top
and bottom lids of the vault are made of 0.5 mm Copper, followed by 1.5mm of Titanium and 2.0 mm of Aluminum.
D.5.1 Polar Lander Dimensions
The polar lander is bigger than the non-polar lander due to requiring more fuel. The maximum height and width
of the lander during its stowed configuration are 1.256 m and 1.740 m shown in Fig. D.13. During its mission
configuration the lander has a maximum height and width of 1.563 m and 2.376 m shown in Fig. D.14. One important
design feature for our lander is that all the instruments have clear fields of view, so each instrument is positioned and
mounted specifically to not obstruct each other. The total mass of the landers
during launch is 710 kg and total dry mass is 241 kg. The important thing is that the C.G. locations always remain in
the center for stability and better attitude control.
Fig. D.13 Polar Lander Stowed Configuration

Fig. D.14 Polar Lander Deployed Configuration
D.5.2 Non-Polar Lander Dimensions
The non-polar landers are smaller than the polar landers due to requiring less propellant. The maximum
height and width of the lander during its stowed configuration are 1.237 m and 1.707 m, respectively, shown in Fig.
D.15. During its deployed configuration the lander has a maximum height and width of 1.554 m and 2.343 m
respectively, shown in Fig. D.16. The total mass of the lander at launch is 681 kg and total dry mass is 235 kg. Again,
all of the instruments have clear fields of view, so the location of each instrument has been positioned and mounted
Fig. D.16 Non-polar Lander Deployed Configuration
Fig. D.15 Non-polar Lander Stowed Configuration

specifically to not obstruct any other instrument. Another important characteristic of both landers is that the C.G.
locations always remain near the center of the body, which allows for better stability and attitude control. The moments
of inertia for each lander in the stowed and deployed configurations are shown in Tables D.2.
Table D.2 Lander Moments of Inertia (kg-m2
)
Ixx Iyy Izz
Polar Lander: Stowed Configuration (Wet) 119.9 120.2 180.2
Polar Lander: Deployed Configuration (Dry) 47.0 47.4 61.9
Non-Polar Lander: Stowed Configuration
(Wet)
107.1 107.7 158.7
Non-Polar Lander: Deployed Configuration
(Dry)
39.5 39.9 51.4
E. Propulsion Subsystem Design
E.1. Propulsion Subsystem Design
This extensive mission has over a dozen main burns which result in a significant amount of propellant
required for all spacecraft on the
mission. Table E.1 highlights the total
amount of propellant used for the
mission.
For both satellites the largest
single change in propellant mass was
during the course of the Jupiter
insertion burn, where more than two-
thirds of the fuel will be burned.
The propellant burned has a huge effect on the amount of propellant needed for future burns. After the JOI
burn the spacecraft loses a lot of mass and it takes less propellant to accelerate/decelerate the spacecraft, as well as to
maneuver the spacecraft using ACS.
E.2 Propulsion Trade Study
A propulsion system trade study for the lander, shown in Table E.3, was conducted to determine which propulsion
system was most viable for our mission. For each lander burn there was a propulsion system selected to do that burn.
The trade study was conducted for multiple propulsion system combinations, where each main lander burn would use
a different propellant, in order to figure out the most efficient way to land on Europa. It compared solid rocket motors
to, monopropellant, and bipropellant propulsion systems. For each propulsion system combination, the final
Table E.1 Total Propulsion Propellant Masses
Spacecraft
Hyd. Mass
(kg)
NTO
Mass
(kg)
He Mass
(kg)
Total
Mass
(kg)
Satellite 1 1439 2043 10.1 3492.1
Satellite 2 1433 2034 10.1 3477.1
Polar
Landers
192 274 1.36 467.36
Non-polar
Landers
183 260 1.29 444.29

propulsion system mass was calculated and compared to the other propulsion system combinations. The solid rocket
motor combination with either the monopropellant or bipropellant system proved to be more massive than all the other
combination of systems. The monopropellant system for all the lander burns was slightly heavier than the bipropellant
system. Therefore, it was determined that the bipropellant system for all the major burns for the lander would be
selected. An important note is that the mission segments listed in this table are from a preliminary mission architecture.
Though a new architecture has been chosen, with slightly different main burns, the results from Table E.2 were
conclusive enough to continue on with a Biprop system for the current mission.
Table E.2 General Lander Propulsion System Trade Study
Mission Segment Drop From Satellite
Cancel Sat
ΔV
Slow down to
the ground
Lander Wet
mass
Design #1
Mono + OODM+SRM Solid prop
Extra Mono fuel
for burn
194.88 kg 79.58 kg 5.32 kg 279.80 kg
Design #2
Mono + OODM Biprop Solid
214.32 kg 98.31 kg 5.28 kg 317.91 kg
Design #3 Mono Only 230.52 kg
Design #4 Biprop Only 220.25 kg
E.3 Propulsion Subsystem Part Lists and Schematics
Tables E.3-4 show the parts lists for the satellite and lander spacecraft. Individual satellites and landers
essentially have the same parts lists. The only variance is in the amount of propellant carried onboard and for the
satellites, the amount of landers that are carried to Europa.
Table E.3 Satellite 1 and 2 Propulsion Part List
Part Use MFG QTY
Mass
(kg)
Isp
(sec)
Thrust
(N)
Power
Req.
(W)
MR-111C
Thruster
ACS AEROJET 12 0.33 215-229 1.1-5.3 16.5
R-42DM
Main
Main
Engine
AEROJET 1 7.3 327 890 46
He Tanks
Fuel
Tank
Aeolus 2 128 N/A N/A 0
Hyd Tank
Fuel
Tank
Aeolus 1 38 NA N/A 0
NTO Prop
Tank
Fuel
Tank
Aeolus 1 38 N/A N/A 0
Total 211.63 79

Table E.4 Lander Propulsion Part List
Part Use MFG QTY
Mass
(kg)
Isp
(sec)
Thrust
(N)
Power
Req.
(W)
MR- 111C
Thruster
ACS AEROJET 12 0.33 215-229 1.3-5.3 13.64
R-4D
Main
Main Engine AEROJET 1 3.4 300 490 46
He Tanks Fuel Tank Aeolus 4 6.5 NA NA 0
Hyd Prop
Tank
Fuel Tank Aeolus 1 7.65 NA NA 0
NTO Prop
Tank
Fuel Tank Aeolus 1 7.65 NA NA 0
Totals 48.66 73.28
E.2.1 Dual Mode System
The satellites main engine is an AEROJET R-42 DM Bipropellant Engines. The specifications are shown in Table
E.5. AEROJET MR-111C 4N thrusters are used for ACS (Table E.6). For the landers the main burn engine is the R-
4D 490N thruster (Table E.7) which slows the lander to about 0.2 m/s as it touches down on the surface of Europa.
Pictures of each of the chosen engines are shown in Figures E.1-3The damage sustained by the landers at this velocity
is negligible. The landers are also using the same 4N thrusters as the satellites for ACS. This will assure that the main
burn engine is pointing in the direction of the greatest velocity reduction for the lander spacecraft. The propulsion
system for the landers and the two satellites are all comprised of dual mode systems shown in Figures E.4-5.
Table E.5. Satellite Main Engine Specifications [24]
Satellite Main Engine
Engine R-42 DM
Propellant Hydrazine/NTO MON-3
Thust/Steady State 890 N
Inlet Pressure Range 25.5-13.8 bar
Chamber Pressure 9.6 bar
Expansion Ratio 200 to 1
Flow Rate 277 g/sec
Valve Aerojet Solenoid
Valve Power 45 W
Mass 7.3 kg
Fig. E.1 Satellite Main Burn Engine R-
42 DM

Table E.6. Lander Main Engine Specifications
Lander Main Engine
Engine R-4D
Epansion Ratio 44 to 1
Flow Rate 158 g/sec
Valve Power 8.25 W
Mass 3.4 kg
Figure E.2. Lander Main Burn Engine
R-4D
Table E.7. Lander Main Engine Specifications [24]
Lander Main Engine
Engine R-4D
Flow Rate 158 g/sec
Valve Power 8.25 W
Mass 3.4 kg Fig. E.3 Lander Main Burn Engine R-
4D
Table E.6. ACS Engine Specifications [25]
ACS Engine
Engine MR-111C
Propellant Hydrazine MON-3
Thust/Steady State 5.3-1.3 N
Flow Rate 158 g/sec
Valve Power 8.25 W
Mass 3.4 kg
Fig. E.2 ACS Engine

Fig. E.4. Satellite Propulsion Schematic
Fig. E.5. Lander Prop Schematic
E.3 Propellant Tanks
The satellites propellant tanks were chosen to be in the shape of capsules. This shape is very convenient and
is very easy to manufacture. The tanks are made from a titanium alloy, Ti6Al 14V. In order to manage propellant
sloshing, PMDS were used. Inside the tank a bladder is used which takes advantage of surface tension to mitigate
propellant sloshing.

E.3.1 Toroidal Tanks
Toroidal propellant tanks were used on the
lander spacecraft in order to increase the radiation
shielding of all the electronics inside the electronic
vault. They are manufactured using a resin mold
transfer method. The tanks are made from the same
titanium alloy as the tanks used for the satellites
(Ti6Al14V). Carbon fiber filament is wound on the
outer surface of the titanium vessel. One of the
main manufacturers is San Diego composites based in San Diego, Ca. A toroidal tank is pictured in Fig. E.6.
The main issue of the toroidal tanks is the structural integrity of the inner periphery of the tank. The weakest part
of the toroidal tanks as shown in Fig. E.7 is
the inner periphery [26,27]. The hoop stress
is the highest at this point. In order to
mitigate this problem the tanks have to have
variable thickness as shown in Fig. E.8. The
inner periphery is made relatively thicker
than the outer diameter of the tank to reduce
the risk of a failure along the inner periphery
of the toroidal tank.
Fig. E.8 Toroidal Tank Wall Thickness Variance
Fig. E.6 Toroidal Tank
Fig. E.7 Hoop Stress Analysis on Toroidal Tanks

F. Thermal Subsystem Design
F.1. Thermal Design Mission Overview
One trade study conducted for the
thermal system compared different
components. (Table F.1) To determine
which components to use for the thermal
system the mass, power, and mission
necessity to the design were weighed. The
components that were proved best qualified were the coating, MLI, and RHUs. The coating is necessary in order to
ensure the correct amount of solar flux being reflected and
absorbed. The coating will help to dissipate heat at Venus
and absorb heat around Europa. The MLI is necessary to
ensure heat stays within the spacecraft to ensure the
components do not exceed their thermal limit. The RHUs
are necessary to ensure the correct thermal gradient for the
quantum wells to work efficiently [28]. Another trade study
was conducted to compare missions similar to this mission
and compare the components used. (Table F.2)
F.2. Thermal Design Mission Overview
The primary purpose of the thermal system analysis within the mission is to keep all components and sub-
components of both the satellites and the
landers within their functional temperature
range. It is essential to keep all subsystems
operational for the entire mission by
conducting detailed thermodynamic
analysis of the internal systems. Some of
the major risks involve the Venus fly-by,
deep space maneuver, and the Europa mission phase. Major analysis needed to be conducted for Earth, Venus, and
Table F.3 Satellites Thermally Constrained Components
Subsystems Components Temperature
Range (o
C )
Power Batteries -10 to 40
Power Charge Controller -10 to 40
Telecommunication Transponder -40 to 60
CD & S Solid-State Receiver -25 to 60
ACS IMU -54 to 71
ACS Reaction Wheels -30 to 70
ACS Star Sensor -20 to 50
Payload HiRISE -10 to 20
Payload MLA -15 to 25
Table F.1 Thermal Subsystem Trade Study
Table F.2 Comparison of Thermal Components on
Past Missions

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