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 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768 4 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768. 5 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768 6 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768. 7 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768 8 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768 9Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768 10 Cal Poly Pomona Student, Aerospace Engineering, 3801 W. Temple Ave., Pomona, CA, 91768

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

La

un

ch P

erio

d

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

Inte

r-p

lan

eta

ry

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

Pu

mp

-do

wn

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

La

nd

er O

per

ati

on

s

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 14 th, 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/oC

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 (oC)

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 20oC range. 12 Only during imaging. ~2 W during standby 13 Advancement in tech. assumed to decrease lower-end of optimal operating temperature to -15oC from current 15oC.

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

Propellant Hydrazine/NTO MON-3

Thust/Steady State 490 N

Inlet Pressure Range 29.3-4.1 bar

Chamber Pressure 7.45 bar

Epansion Ratio 44 to 1

Flow Rate 158 g/sec

Valve Aerojet Solenoid

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

Propellant Hydrazine/NTO MON-3

Thust/Steady State 490 N

Inlet Pressure Range 29.3-4.1 bar

Chamber Pressure 7.45 bar

Epansion Ratio 44 to 1

Flow Rate 158 g/sec

Valve Aerojet Solenoid

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

Inlet Pressure Range 12.1-3.4 bar

Chamber Pressure 7.45 bar

Epansion Ratio 44 to 1

Flow Rate 158 g/sec

Valve Aerojet Solenoid

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 (oC )

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

Europa. By having temperature parameters under control at these three major locations, both of the satellites will be

ensured to survive the trip to Europa, and the landers will be able to finish their mission on Europa safely.

In order to accomplish the entire mission, all sub-systems payloads need to be under thermally controlled at

equilibrium temperature as they approach Venus, and upon arrival at Europa. Tables F.3 and F.4 illustrate the most

thermally constrained

components of the satellites and

landers, respectively.

Based on the comparison

shown in the tables above, for

both the satellites and the

landers, the most thermally

constrained components are the

batteries and the charge

controller with survival

temperature ranges of -10 to 40 ℃. The reason these are considered the most thermally constrained despite the HiRISE

and MLA actually having tighter temperature restrictions is because these instruments are not vital to mission success.

Measures will be taken to maintain thermal constraints for these payloads, however it is more important to protect the

batteries and charge controllers

As for the satellites, upon approach of Venus the high solar flux is a huge threat since most of the electronics are

designed to be kept within the temperature range of -10 to 50 ℃. In order to overcome this problem, each of the

satellites is equipped with two radiators, with areas of 1.25 m2 each. In addition to that, three louvers are also attached

to dissipate heat. Most importantly, the satellites are 45% covered with Aluminized Teflon Coating (ε = 0.81) on the

outer surface area, which serves as a primary source of heat reduction at Venus. With all of the considerations taken

into account, a maximum temperature of 47.9 ℃ will occur at the closest point to Venus during its flyby. As a result,

all of the components and electronics meets the thermal constraints, except for the batteries and the charge controllers

in the power sub-system. These two components will be placed in specially designed thermal protection vaults with

low thermal conductivities (i.e. aerogel) to isolate from incoming heat flux sources both from within and outside the

spacecraft. Additionally, these components will be placed near the louvers and RHUs to allow for rapid thermal

Table F.4 Landers Thermally Constrained Components

Subsystems Components Temperature

Range (oC )

Power Batteries -10 to 40

Power Charge Controller -10 to 40

Power Quantum-Well

Generators

-5 to 40

Telecommunication Transponder -40 to 60

CD & S Solid-State

Receiver

-25 to 60

ACS Sun Sensor -15 to 60

Payload MARDI -40 to 70

Payload MEMs

Seismometer

-200 to 100

alleviation at Venus or Europa if the temperature falls below the required temperatures.

When the satellites arrive at Europa, there will be a Tmax of 13.8 ℃ for Satellite 1 and 14.8 ℃ for Satellite 2 due to

a difference in the number of RHUs. The RHUs serve as the primary source of providing heating power under the

freezing conditions at Europa. Each RHU provides 1 Watt of heating power. By having a total of 104 RHUs for

Satellite 1, and 109 RHUs for Satellite 2, both of the satellites are able to maintain an overall temperature of -5 ℃

under extreme conditions.

As for the landers, they will encounter maximum temperatures of 49.1 ℃ during the Venus fly-by with telecom,

propulsion, and payload subsystems turned off to limit heat dissipation. Since most of the electronics are similar to

those found in the satellite, the most thermally constrained components are the batteries and the charge controller.

Again, these components will be placed in separate vaults to shield them from solar flux. For additional heat protection,

the High Gain Antenna is pointed at the sun to prevent direct contact heat flux to the landers from Venus. The landers

will also be 20% covered with Aluminized Teflon Coating to reduce heat.

It is essential for the landers to survive for a minimum of 90 days under the harsh environment on Europa, which

includes high radiation dosages and extremely low temperatures. The landers are also 35% covered with Multi-Layer

Insulator (MLI) on the outer surface which is specially designed to reduce the rate of incoming heat-radiation from

Europa. A total of 18 RHUs are used to keep

all payloads and electronics above -15 ℃. To

sum up, all of the protection methods are

enforced to minimize risks and ensure

mission success.

Table F.5 demonstrates a summary of the

maximum and minimum temperatures for

the satellites throughout the entire mission.

The maximum temperature is

understandably at Venus due to its

proximity to the sun, whereas the minimum

temperature occurs at Europa, when in the shadow of both Jupiter and Europa.

Table F.6 provides the maximum and minimum temperatures for the landers. The worst case cold temperature at

Table F.5 Satellites Worst Case Temperatures Table

Location Tmax (oC ) Tmin (oC )

Earth 36.6 -16.4

Venus 47.9 -

Europa (Satellite 1) 13.8 -26.9

Europa (Satellite 2) 14.8 -25.9

Table F.6 Landers Worst Case Temperatures Table

Location Tmax (oC ) Tmin (oC )

Venus 49.1 -

Europa 20.7 -16.2

Venus for both the satellites and landers are not applicable since they do not exceed the worst case minimum

temperature at Europa. For Satellite 1, there is a slight temperature difference of 1 ℃ at Europa due to the different

amount of RHUs.

F.2 Thermal Payload Configuration Schematics

It is essential to keep the main electronic section warm inside of the satellites during the Europa phase. There are

a total of 104 RHUs for Satellite 1, 56 of them will be surrounding the electronics vault in order to keep the most

important components of

the satellite within

operating temperatures.

There will also be 6

RHUs around the optical

instruments, as well as 5

RHUs around the main

engine just to keep the

engine at its equilibrium

temperature. Ten RHUs

are used around the Reaction Wheel Assembly (RWA). Lastly, the star tracker will be surrounded by 3 RHUs. To

protect from thermal radiation, 35% of outer surface area will be covered with MLI, as well as a 45% of Aluminized

Teflon coating for heat reduction at Venus. The location of all of the crucially placed RHUs are shown in Fig. F.1.

For the lander 14 RHUs out of a total of 18 are placed around the electronics and radiation vault (not shown) to

keep the most important

payloads at temperature

equilibrium due to Europa’s

harsh cold temperature. The

remaining four RHUs will be

distributed evenly to the four

Thruster Cluster Assemblies.

Louvers will only be turned on

Fig. F.1 Satellites Thermal Payload Configuration

Fig. F.2 Landers Thermal Payload Configuration

during Venus fly-by to cool the landers. The landers will have more MLI coverage than the satellites; they will be

covered with 65% outer surface area since they will have to last 90 days minimum at Europa’s high radiation surface.

Lastly, landers are made of Aluminum 2024 just like the satellites, as well as a 20% outer surface coverage of

Aluminized Teflon Coating to prevent heat overload during Venus fly-by. A model of the lander, with RHU

placements is shown in Fig. F.2.

F.3. Transient Case Analysis

There are six different main modes during the entire mission. Each of the modes requires a different power input

that will result in a different temperature output. All of the transient case temperature analyses for the satellites are

illustrated in Fig. F.3.

Fig. F.3 Satellite 1 and 2 Temp. vs. Power Flight Modes Analysis

A peak temperature occurs at 68 ℃ in mode II, with a power usage of 510.1 W. Mode II involves the

satellites performing their deep space maneuver, which requires the most power in the propulsion and

telecommunication sub-systems. This will only occur for a short transitional period of time compared the overall

mission length. The lowest temperature occurs at Mode V, with a power input of only 347.1 W. This mode only

involves the magnetometer and part of the telecom subsystems from the satellite, shortly after the landers have

-60.0

-40.0

-20.0

0.0

20.0

40.0

60.0

80.0

399.1 510.1 465.1 581.1 347.1 558.9

Tem

pe

ratu

re (℃

)

Tmax_Earth

Tmin_Earth

Tmax_Venus

Tmax_Europa

Tmin_Europa

Power (W)

Mode I Mode II Mode III Mode IV Mode V Mode VI

deployed. As a result, a low temperature occurs at -28 ℃. RHUs will be providing heating power to keep all

electronics and payloads warm under cold conditions. Both satellites will experience similar situations, except that

Satellite 2 will have a slightly higher temperature than Satellite 1 by 3℃ due to the difference in the amount of

RHUs.

Landers will experience the highest temperature at mode II, when they utilize the most power during their

deployment from the satellite to Europa and during landing. A peak of nearly 65℃ will occur as a result of 250 W of

power for a short period of time. Louvers will keep the landers cooled down during this transitional period. The lowest

temperature occurs at -40 ℃ at Mode I, from the launch phase to cruise, when the landers sit inside of the satellites.

To keep all payloads functional, RHUs will be regulating the temperatures by providing heating power. The various

power modes for the landers can be seen in Fig. F.4.

Table F.5 is a summary of the thermal payloads listing their mass and power. The thermal payload for Satellite 1

has a mass of 32.52 kg and a total power requirement of 164.2 W. Satellite 2 has a higher mass of 32.72 kg due to the

added RHUs and power of 169.2 W. As for the landers, the total mass is 2.62 kg and 20.6 W of power.

Fig. F.4 Landers Temp. vs. Power Flight Modes Analysis

-130.0

-80.0

-30.0

20.0

70.0

120.0

60.0 250.0 130.0 95.0

Tem

pe

ratu

re (℃

)

Power (W)

Tmax Tmin

Mode I Mode II Mode III Mode IV

Table F.5 Mass and Power Chart for Thermal Payloads List

Spacecraft Thermal Components Mass (Kg) Power (W) QTY

Satellite 1 & 2

RHU (Pu-238) 0.04 - 104 for Sat. 1

109 for Sat. 2

Radiator (C-Ag Teflon) 6.4 25 2

Louver 1.5 3.4 3

MLI 3.4 - 35%

Coating 7.66 - 45%

Total (Satellite 1) 32.52 60.2

Total (Satellite 2) 32.72 60.2

Lander

RHU (Pu-238) 0.04 - 18

Louver 0.4 1.3 2

MLI 0.89 - 65%

Coating 0.21 - 20%

Total (Lander) 2.62 2.6

G. Power Subsystem Design

G.1 Power Subsystem Design Summary

The Europa mission provides a unique challenge for power in the deep space environment. The main challenges

to overcome were the eclipse time of the satellite behind Europa, the high radiation environment, and the low solar

flux at this distance from the Sun. Solar panels proved feasible for the satellites as discussed in the next section, but

did not prove feasible for the landers. The landers will be powered by a Quantum Well thermoelectric system. This

system generates power from temperature gradients, which will be generated by our radioactive heating units and the

ambient cold of space. Both these of systems can provide adequate power for our 90-day mission and can survive the

harsh environment of Europa and the Jovian system.

G.2 Power Requirements

The payloads and main systems of our spacecraft have different power requirements, which provided the starting

point for designing the power system. A list of the subsystems and power requirements are shown below in Table 1.

Subsystem Orbiter Subsystem

Power (W)

2nd Orbiter Subsystem

Power (W)

Single Lander Subsystem

Power (W)

Thermal control 138.5 138.5 7.25

ACS 130.1 130.1 5.91

Power 10 10 5

CDS 75 35 35

Communications 196.3 196.3 96.3

Propulsion 115.2 115.2 45

Mechanisms 4.94 4.94 2.15

Max Total 535 392 196.6

Nominal Total 75 75 36.3

Table G.1 Power System Allocation Overview

After the power requirements of each subsystem were determined, the orbit and trajectories were analyzed to

determine what kind of solar and eclipse environments the spacecraft would be in. This determined the duration of

power for the each system and in what environments the most power would be needed. The basic power requirements

of the satellites and landers based off orbit and eclipse times were calculated to determine the amount of time for

which solar panels could not power the satellite and the lander.

The main points to note are the maximum eclipse times. The maximum eclipse times for the satellites and landers

are 2.42 hours and 42.4 hours respectively. These eclipse times along with the nominal operating power from Table 1

determine what kind of power system we will need. It is rather apparent from these values that solar panels are not an

option for the landers.

G.3 Power System Selection

The selected power system took

careful deliberation due to the large

variety of options and unique

power environment. The main

selections based off of past space

missions of similar scope were

Radioisotope power sources or

Photovoltaic Arrays. Fig. G.1

illustrates the feasibility of several

different power systems for the use of any sized mission.

The red line represents approximate length of the lander

segment of this mission, which is about 120 days, including

mapping. The main challenge confronted once the

requirements were determined was the volume constraints

of the spacecraft interior, as well as the volume constraints

for stowing in the Falcon Heavy payload fairing. The low

solar flux due the far distance from the Sun at Europa was

another major concern. The solar flux near Europa averages

Fig. G.1 Power Source Feasibility given Mission Duration and

Required Power [34]

Fig. G.1 Example RTG layout [36]

about 50 W/m2. This led to a very large mass and volume of solar panels to accommodate the high power requirements.

Trade study research determined that solar arrays were feasible for the satellites but would not be feasible for the

landers. For the landers the first consideration was using a Radioisotope Thermoelectric Generator (RTG) as seen in

Fig 2.

This device uses the radioactive decay of materials to generate heat and then converts that heat into electricity.

However, a difficultly arose when researching the availability of radioactive fuels for use in this mission. The RTG

could adequately power the landers but there were concerns about the validity of its use for this mission due to limited

information being available on the production and sizing of RTGs. Planetary protection concerns also arose when

discussing putting seven sources of radioactive waste onto the surface of Europa. This led to the investigation of

alternative power systems; specifically,

into a technology called Quantum Wells as

seen in Fig. G.3.

Quantum wells act as the

“thermoelectric generator” portion of

RTG. Because of this new quantum well

technology, the mass and amount of

radioactive fuel needed to power the

system was greatly reduced. Quantum

wells are able to generate electricity from

heat much more efficiently than previous

technologies. This allowed for the thermal

system of the spacecraft to inform the

power system design. The thermal system

would provide heat to the spacecraft using

RHUs and the quantum wells would

convert the temperature gradients

generated between the spacecraft interior

and exterior into power. Using these

Fig. G.2. Quantum Well Material Layout [35]

quantum wells 10 kg of mass could be saved per lander, meaning a total decrease of 70 kg for the power system over

the next best power option.

G.4 Battery Pack Sizing

The battery pack sizing for both the lander and satellites was dictated mainly by the eclipse times. However, for

the landers, since the power system is not dependent on sunlight, the battery pack will not be specifically sized for the

eclipse. For both satellites, with

the given eclipse times, power

demands, redundancy, and 150

W-hr/kg energy density, the

battery pack for the satellite was

sized at 26 kg. For the landers the

battery pack was sized for the

situation in which the quantum

wells failed and a last data transmission was required to transmit the remaining data to the satellites. Given the power

demand and data transmission time, the battery pack for the lander was sized at 0.816 kg. The specifications for the

battery pack, along with the calculated masses of the battery packs can be found in Table G.2. Both these packs were

sized with a double redundancy meaning they are sized to last twice as long as they need to be. The battery packs will

be placed in the radiation vault so they can be thermally controlled and shielded from incoming radiation.

G.5 Power Modes

The main power modes for both

satellites and landers were

determined from the orbit and the

different subsystems that would be

powered. A plot of the different

power modes is shown below in

Figures G.4 and G.5 for both

satellites and for the landers,

respectively.

Table G.2 Battery Sizing for Lander and Satellites

Battery Calcs

(Lander)

Battery Calcs

(Satellites)

Power Required (W) 196 435

Time required (hr) 0.167 2.4

Voltage Required (V) 28 28

Power loss (V) 0.97 0.97

DoD (%) 0.55 0.55

Energy Density (W-hr/kg) 150 150

Capacity (Amp-hr) 2.18 69.88

Energy Required (W-hr) 61.23 1956

Mass of Pack (kg) 0.408 13.04

With Contingency (kg) 0.816 26.09

Fig. G.4 Satellite 1 and 2 Power Modes during Mission

0

100

200

300

400

500

600

700

Launch toInnerCruise

Cruise DSM CruiseAttitude

Corrections

EuropaInitialRecon

EuropaNominal

EuropaAttitude

Corrections

Po

we

r (W

atts

)

Power Modes

Satelllite 1

Satellite 2

Fig. G.5 Lander Power Modes during Mission

G.6 Power System Simulation Model

The last step taken to model the power systems for the spacecraft was to model the power system using a simulation

model. The program Simpowersystems from Matlab was used to model the entire system. The model for the satellite

was set up to simulate the solar panels and the delivery of power to all of the subsystems. The model for the lander

was set up using a DC power source for the quantum wells. These models were not able to run fully due to a system

error with the most recent version of the software. However, the system was still able to show that given the input

power of both the quantum wells and solar panels that all subsystems would be able to receive adequate power

throughout the duration of the mission lifecycle. The power system block diagram for satellite 2 can be seen below in

Fig. G.6. The diagram is identical for satellite 1, but with three landers instead of four. The power system block

diagram for the landers can be seen below in Fig. G.7.

0

50

100

150

200

250

300

Launch to Cruise Deployment toEuropa/Landing

Europa DataMeasurementw/Telecomm

Europa DataMeasurement

w/out Telecomm

Po

we

r (W

atts

)

Power Modes

Lander Power Modes

Fig. G.6 Satellite Power System Block Diagram

Fig. G.7 Lander Power System Block Diagram

H. Telecommunications and CDS Subsystem Designs

H.1 Telecommunications and CDS Overview

The telecommunications and CDS subsystem are crucial to the Europa CT Scanning Mission. Its primary purpose

will be to allow the satellite and landers to perform their missions and record and transfer all of the required data back

to Earth. During the Lander Reconnaissance and Descent Phase, a symbol rate of 111 ksps will be transmitted. Once

the Mission Operation Phase begins, the two satellites will be in an orbital trajectory that will allow each lander to

communicate with each satellite at least one time per satellite orbit. The S-band will be used to allow the lander to

transmit data from its LGA to the satellite’s HGA. The symbol rate for this downlink will be 140 ksps. A total of 23.5

GB will be produced from the seismometer and camera for each lander. Therefore, the satellite will receive a total of

164.5 GB and will transfer this data using the X-band to NASA’s DSN. The symbol rate for this transfer will be 242

ksps.

The software will be a centralized topology architecture. Its functions will consist of navigation, spacecraft control,

payload functions, safekeeping mode, telecommunications, and spacecraft monitoring functions. Both the lander and

satellite will have similar software architectures. The key differences will be the lander’s payload data processing

functions and the autonomous descent functions.

The hardware selected for both the satellite and lander are also nearly identical. The key differences will be an

additional HGA for the satellites. Hardware was selected based on its heritage and radiation protection. The total

satellite hardware mass will be 107.18 kg and the power consumption will be 268.2 W. The total lander hardware

mass will be 22.75 kg and the power consumption will be 59.22 W.

H.1.1 Design Drivers

The design drivers for the telecommunications and CDS subsystems were derived from three key points in the

RFA. The first key point was the length of the mission. Since the mission consisted of only 90 days, it is expected that

the telecommunications be designed for a high data rate transfer. The second was the number of locations the payload

was required to be at. From this requirement, seven separate landers were proposed to deploy the payload at each

landing site and relay the data back to a carrier satellite. This meant that the relay satellite’s telecommunications had

to be able to communicate with each lander at least one time during the 90 day mission. The final key point was the

location of the mission. Europa has an extensive amount of radiation in the area that could easily damage the

telecommunications and CDS. Therefore, this subsystem had to be designed with redundancy and robustness in mind.

H.1.2 Telecommunications and CDS Operations during Lander Reconnaissance and Descent Phase

During the four stages of the Lander Reconnaissance and Descent Phase, the CDS will store imaging and

engineering data. Table H.1 displays the data sizes and what will be captured during each stage.

Table H.1 Data Acquisition and Sizing for Lander Reconnaissance and Descent Phase

Stage Transmission Data Size

1: Initial Mapping Polar Lander captures low resolution images

with the MARCI at 5300 m/pixel and 40 m/pixel 66.7 GB

2: Uplink/Downlink of Processed

Images

Polar Lander transmits low resolution images

of the selected landing sites 16.3 GB

3: Detailed Mapping Landers captures high resolution images of

landing sites with the MARDI at 1.5 m/pixel 621.7 GB

4: Descent and TRN Landers captures descent images and

engineering data during descent 0.58 GB

Data acquired from stage 2 and stage 4 will be critical to the mission. The 16.3 GB from stage 2 will be transmitted

from the polar lander’s LGA to the satellite HGA. The S-band will be used for satellite to lander transmissions. Once

the data is received, the satellite will relay the data back to the other landers to assign their selected landing sites. This

data transmission is expected to take two orbital periods which equates to roughly one week. With 10% added for

overhead and Reed-Solomon Coding, the symbol rate expected for this stage will be 111 ksps. The data acquired from

stage 4 will be discussed in the mission phase as it will be transmitted during that phase.

H.1.3 Mission Operation Phase

The telecommunications and CDS subsystem will see the most usage during the Mission Operation Phase. During

this phase, each lander will record its own seismic and imaging data. As this data is collected, it will be transmitted

from the lander’s LGA to either Satellite 1 or Satellite 2’s HGA. The S-band frequency will be used for downlink and

uplink communications. The mission operation’s satellite trajectories allow each lander to communicate with Satellite

1 and Satellite 2 at least one time per Europa orbit. The maximum distance each lander will have between Satellite 1

or Satellite 2 will be approximately 400,000 km. It was estimated that a maximum of 3 hours will be allotted for each

lander to communicate with a satellite per orbit. This equates to approximately 76 hours of communication time per

lander over the course of the whole mission.

Once the lander data has been received, Satellite 1 and Satellite 2 will transmit the data back to NASA’s Deep

Space Network (DSN). Their HGA will transmit the data and the DSN 34m antenna will be expected to receive the

data. The X-band frequency will be used to downlink the data. The satellites and DSN maximum distance is expected

to be 9.27 x 108 km, however this will not occur during the lander mission phase.

Table H.2 shows the data sizes that will be acquired from each payload. During this time, stage 4 data of the Lander

Reconnaissance and Descent Phase will be transmitted as well. Each lander produces a total of 23.5 GB of data. With

the 3 hours per orbit time allotted,

a symbol rate of 140 ksps is

required. This symbol rate

includes an extra 30% for

engineering and additional data.

The Reed –Solomon coding scheme is used and overhead is taken into account. The total mission data size will be

164.5 GB. This mission will request a communication time of 24 hours per week with the DSN. This amount of time

is reasonable because the radiation environment will not allow the equipment to last very long, and the mission time

length is short. This gives a symbol rate of 242 ksps. With the equipment designed for mission downlink, it was

calculated that the DSN can uplink to the satellites at a rate of 120 ksps. It is important to note that all calculations

were done assuming that only one satellite is operational. Therefore, the second satellite can be considered redundant,

however it will remain fully operational in the event of a failure with Satellite 1.

H.1.4 Software Design

The software architecture was designed based on the functions needed to perform the mission. These functions

include navigation, telecommunications, payload functions, thermal control, and data storage. The software was

designed with a centralized topology architecture. The central function will input data and output the required system

functions. Around this centralized point are the functions required for the mission. Fig. H.1 and Fig. H.2 show the

software block diagrams for the satellites and the landers. Their architectures are nearly identical except for the

functions involving their respective payloads. The lander also has two extra functions: the lander must be able to

process the landing site images it collects for the Lander Reconnaissance and Descent Phase and it also must filter the

seismic data in order to send only the important data for the mission.

The safekeeping functions for the satellite and lander will also be important to design. This function must relate to

all the other subsystems and must be able to perform during all conditions. As the satellite monitor’s the subsystem

health, when a mission critical failure occurs, the satellite will enter safekeeping mode. When safekeeping mode occurs

for the satellite, the satellite will automatically reduce power to a minimum and orient the HGA to point towards the

sun. Next it will rotate the spacecraft with a slight wobble to allow the HGA to have a large enough beam width until

Table H.2 Lander and Satellite Data Sizes and Symbol Rates

Lander to Satellite Satellite to DSN

Seismometer Data (GB) 7.7 54

Beagle Camera Data (GB) 15 106

Descent and TRN Data (GB) 0.58 4.06

Communication Time 76.1 hr/mission 24 hr/week

Symbol Rate (ksps) 140 242

the DSN can communicate with the satellite and further examine the mission critical failure that had occur. Once the

landers are deployed and are in the mission operation mode, the safekeeping mode will be slightly different. When a

mission critical failure occurs on the lander, the lander will similarly reduce its power to a minimum and constantly

transmit a distress signal. Once the satellite encounters this signal, it will be able to transmit this data back to the DSN

for further assessment.

Fig. H.1 Satellite 1 and 2 Software Design Block Diagram

Fig. H.2 Lander Software Design Block diagram

H.2 Telecommunication and CDS Hardware

The hardware selected for both the lander

and satellites are almost identical. The

difference comes in the quantities of each piece

of hardware and the extra HGA on the satellite.

The hardware mass and power for the satellites,

and landers can be found in Table H.3 and h.4

respectively. The hardware was selected based

on its heritage and its rad-hardness. This was

important for the mission to ensure the equipment can survive in the intense radiation environment to and around

Europa.

The amplifier selected has the ability to transmit frequencies ranging from 1.5 GHz to 30 GHz [39] and the

ability to output up to 450 W of power. These characteristics met the requirements for the telecommunications design.

Table H.3 Satellite 1 and 2 Telecommunication and CDS

Hardware Mass and Power Table

Hardware Manufacturer Quantity Mass

(kg)

Power

(W)

Low Gain

Antenna

Aeolus

Technology 2 0.3 0

High Gain

Antenna

Aeolus

Technology 1 71.14 0

TWTA

Amplifier Bosch 4 3.4 213.7

Small Deep

Space

Transponder

General

Dynamics 2 3.2 19.5

RAD750 BAE 2 1.22 25

NEMO SSR Airbus 2 6.5 10

Total 107.18 268.2

Fig. H.3. Satellite 1 Hardware Block Diagram

Fig. H.4. Satellite 2 Hardware Block Diagram

This amplifier has an expected lifetime of over 15

years, making it perfect for this mission. The

traveling wave tube design was selected because

it was the most simple and reliable design. For the

satellites, a maximum power of 213.7 W will be

input to the amplifier. This will give a transmitting

power of 48.33 dB. For the landers, a maximum

of 4.72 W will be input to the amplifier. This will

give a transmitting power of 4.5 dB.

Table H.4. Lander Telecommunications CDS Hardware

Mass and Power Table

Hardware Manufacturer Quantity Mass

(kg)

Power

(W)

Low Gain

Antenna

Aeolus

Technology

2 0.3 0

TWTA

Amplifier

Bosch 2 3.4 4.72

Small Deep

Space

Transponder

General

Dynamics

2 3.2 25

RAD750 BAE 2 1.22 25

NEMO SSR Airbus 1 6.5 10

Total 22.74 59.22

Fig. H.5. Lander Hardware Block Diagram

The SDST is a transceiver developed by General Dynamics and JPL. It allows multiple telecommunication

equipment to be combined in a compact and light package [40]. The SDST has already proven its effectiveness in

many other NASA missions such as the Dawn, MRO, MER, and Messenger.

The RAD series computers are single board computers that are designed for space flight and radiation protection.

The RAD750 is the latest computer developed by BAE. It can withstand u to 100 krad and requires only 25 W of

power [41]. The RAD750 has already seen use in missions such as the MRO, Curiosity, and Juno.

The NEMO SSR is a space qualified solid state recorder (SSR) developed by Airbus Defence and Space. It is non-

volatile and can store up to 0.5 TB while consuming only 10 W of power [42]. It is relatively small and lightweight at

6.5 kg. The NEMO SSR can withstand up to 40 krad. It was space qualified in 2014 with over 20 months of successful

operation.

All of the telecommunication equipment, along with the payload, power, and propulsion system are shown linked

together in the hardware diagrams for the satellites and landers. (Figures H.3-5)

H.3 Telecommunications Uplink/Downlink Analysis

The uplink and downlink analysis were performed by following the telecommunication analysis in Elements of

Spacecraft Design. The first important variable for our analysis was the type of antenna. This affected the antenna

gain. The lander’s LGA provided an antenna gain of 7 dB. The 4m HGA provided 48.38 dB. Once the hardware was

selected, the frequency and symbol rate were decided based on the network provided. The S-band and X-band were

chosen and the symbol rate was given by analyzing the data being transmitted. Finally, the transmitter power was the

last variable to change for the analysis. The transmitting power had to be increased in order to achieve a positive

margin for the uplink and downlink analysis. Once the margins were positive, the transmitting power was converted

from dB to watts in order to calculate how much power was required from the amplifier.

H.3.1 Lander to Satellite Downlink Analysis

The lander to satellite downlink analysis in Table H.6 had significantly large margins with a carrier performance

margin of 19.87 dB and a command performance margin of 6.55 dB. The maximum distance between the lander and

satellite was 400,000 km. This distance was short enough for a strong downlink. In order for the downlink to work

properly, a transmitter power of 4.5 dB was selected. This ensured that the landers would only require a 2.8 W.

Table H.5. Lander to Satellite Downlink Analysis at Maximum Distance

Inputs Carrier Performance

Set Frequency 2.29 GHz Carrier/Total Power Ratio -8.52849 dB

Set bit error rate 0.00001 Receiver Carrier Power -160.736 dB

Set range 400000 km Noise Bandwidth 14.77 dB-Hz

Set Symbol rate 139.9 ksps Noise Ratio Received 39.87141 dB

Transmitter Power 4.5 dB Noise Ratio Required 20 dB

Cable loss -0.06 dB Margin 19.87141 dB

Antenna Diameter 7

EIRP 11.44 dB Command Performance

Free space path loss

-211.678

dB Command/Total Power

Ratio -0.65668

dB

Atmospheric attenuation -0.15 dB Power Received -152.865 dB

Polarization loss -0.2 dB Symbol Rate Effect -51.4582 dB-Hz

Receiving Gain 48.38 dB Eb/No Achieved 11.05504 dB

Pointing loss 0 dB Eb/No Required 4.5 dB

Theta 0 deg Margin 6.555038 dB

Receiver Cable Loss 0 dB

Total Received Power -152.208 dB

Receiver Noise Temperature 21 K

System Noise Density -215.378 dB/Hz

H.3.2 Satellite to DSN Downlink Analysis

The Satellite to DSN downlink analysis in Table H.7 required the most amount of power from the amplifier. A

transmitter power of 21.1 dB was needed to achieve a carrier performance margin of 16.67 dB and a command

performance margin of 0.99 dB. The power required was 130 W. The antenna was also sized to achieve the maximum

diameter but still remain within the design constraints. A diameter of 4 m was selected and achieved an antenna gain

of 48.34 dB.

Table H.6. Satellite to DSN Downlink Analysis at Maximum Distance

Inputs Carrier Performance

Set Frequency 8.4117 GHz Carrier/Total Power Ratio -8.52849 dB

Set bit error rate 0.00001 Receiver Carrier Power -163.934 dB

Set range 9.27E+08 km Noise Bandwidth 14.77 dB-Hz

Set Symbol rate 241.5 ksps Noise Ratio Received 36.67384 dB

Transmitter Power 21.1 dB Noise Ratio Required 20 dB

Cable loss -0.06 dB Margin 16.67384 dB

Antenna Diameter 4

EIRP 69.37888 dB Command Performance

Free space path loss -290.279 dB Command/Total Power Ratio -0.65668 dB

Atmospheric attenuation -0.15 dB Power Received -156.062 dB

Polarization loss -0.2 dB Symbol Rate Effect -53.8292 dB-Hz

Receiving Gain 66.2 dB Eb/No Achieved 5.486474 dB

Pointing loss -0.35508 dB Eb/No Required 4.5 dB

Theta 0.581333 deg Margin 0.986474 dB

Receiver Cable Loss 0 dB

Total Received Power -155.405 dB

Receiver Noise Temperature 21 K

System Noise Density -215.378 dB/Hz

H.3.3 DSN to Satellite Uplink Analysis

The DSN to satellite uplink analysis in Table H.8 was created to use the satellite’s telecommunication properties

required for downlink and the current DSN 34 m antenna properties. With these properties inserted into the analysis,

the symbol rate was increased until an appropriate uplink could be achieved. A symbol rate of 50 ksps was selected

to provide a carrier performance margin of 23.99 dB and a command performance margin for 5.14 dB.

Table H.7. DSN to Satellite Uplink Analysis at Maximum Distance

Inputs Carrier Performance

Set Frequency 7.15 GHz Carrier/Total Power Ratio -8.52849 dB

Set bit error rate 0.00001 Receiver Carrier Power -166.615 dB

Set range 9.27E+08 km Noise Bandwidth 14.77 dB-Hz

Set Symbol rate 50 ksps Noise Ratio Received 33.99249 dB

Transmitter Power 43 dB Noise Ratio Required 10 dB

Cable loss -0.06 dB Margin 23.99249 dB

Antenna gain 65.5157 dB

Antenna Diameter 34 Command Performance

EIRP 108.4557 dB Command/Total Power Ratio -0.65668 dB

Free space path loss -288.868 dB Power Received -158.744 dB

Atmospheric attenuation -0.15 dB Symbol Rate Effect -46.9897 dB-Hz

Polarization loss -0.2 dB Eb/No Achieved 9.6446 dB

Receiving Gain 48.33 dB Eb/No Required 4.5 dB

Pointing loss -25.6548 dB Margin 5.1446 dB

Theta 0.068392 deg

Receiver Cable Loss 0 dB

Total Received Power -158.087 dB

Receiver Noise Temperature 21 K

System Noise Density -215.378 dB/Hz

I. ACS/GN&C Subsystem Design

All satellite and landers utilize a 3-axis attitude control system (ACS) to meet the pointing accuracy of the

telecommunications and optical payload, and to improvise lander precision. Common ACS components between the

satellites and landers include sun sensors, and thruster cluster assemblies (TCAs).

I.1 Satellite ACS Design Overview

Satellites 1 and 2 both have a 3-axis control system in order to meet the high gain antenna (HGA) and HiRise

camera pointing requirements of 0.01o and 0.001o respectively. The reaction wheel assembly in particular addresses

the HiRise pointing requirement. Thus, the RWA is primarily used during the operation of the HiRise camera during

the preliminary mapping phase, and the post-lander-deployment mapping phase where images are taken of the landers

and their surrounding environment. The TCA is composed of four Honeywell Constellation reaction wheels, one of

which is redundant as shown in Fig. I.1. Three of these reaction wheels are spun simultaneously to rotate about one

axis. In order to minimize power consumption to 22 watts per reaction wheel, these wheels are spun at 3000 RPM

instead of the maximum operational rate of 6000 RPM [43].

Fig. I.1. RWA from Satellite Cut-out and Uncoupled Longitudinal Thrusters on TCAs

Four thruster cluster assemblies (TCAs) exist on the four sides of the satellite, with two opposing pairs containing

two thrusters, and the other two pairs containing four thrusters. In order to minimize plume effects on the HGA,

landers, and solar arrays, the longitudinal thrusters were designed to point downwards (along –X direction) as shown

in Fig. I.1. Thus, these vertical thrusters are uncoupled, which generates an undesirable thrust along the longitudinal

axis. The increase in ΔV due to this thrust can be minimized using low burn times. This TCA configuration is not

risky as it was proven, and used successfully on Cassini [2].

I.1.1 Satellite Navigation

The navigation and flight path control system for the satellite is based upon that of previous missions. Orbit

determination, including velocity and position estimation, is accomplished using the two SED26 star trackers (one for

redundancy), 10 wide-angle Adcole digital sun sensors, and measurements of Doppler shift via uplink and downlink

from the DSN. Doppler shift is used for velocity measurement, whereas a ranging pulse is used for angular and distance

measurement [44]. Angular measurement is completed using the Very Long Baseline Interferometry (VLBI), which

requires the use of two DSNs, and a differenced Doppler. Optical Navigation (opnav) will be used near Jupiter and

the other Jovian satellites for the Jupiter insertion burn and flyby maneuvers, by means of the MARCI [44]. Flight

path control is maintained using trajectory correction maneuvers (TCMs) during flybys and orbit trim maneuvers

(OTMs) when in the 90-day mission orbit around Jupiter [44].

I.1.1 Satellite Maneuver Analysis

The twelve Aerojet MR-111C thrusters on the TCAs serve as both the thrust vectoring system during large burns,

as well as providing the fine-tuned, precision burns for scientific operations – such as HiRise mapping. This behavior

is attributed to the thrusters’ low minimum impulse bit of 0.08 N-s. Due to the large moments of inertias of the satellites

about each axis, the total maneuver time for a180o turn using a bang-bang control system is between 12.5 and 15

minutes. These large maneuver times were a result of low burn times of 40 to 60 seconds, which were selected to

minimize propulsion consumption. Table I.1 provides the rotation maneuver analysis for both satellites. Dry mass

moments of inertia is used as most of the propellant is used during the DSM, JOI, and pump-down burns.

Table I.1 Satellites Axial Rotation Maneuver Analysis

Sat. # Axis Dry Mass Moment

of Inertia (kg-m2)

Burn

Time (s)

Coast

Time (s)

Total Maneuver

Time (s)

Required

Propellant, mp (kg)

1

X 89,169 40 688.2 768.2 0.1424

Y 87,291 40 673.2 753.2 0.1424

Z 154,608 60 783.1 903.1 0.2137

2

X 89,008 40 689.5 769.5 0.1424

Y 87,176 40 674.1 754.1 0.1424

Z 154,578 60 783.3 903.3 0.2137

I.1.2 Satellite RWA Desaturation

One of the largest generated torque occurs during the Earth parking orbit at 400 km for Satellite 1. This atmospheric

torque of 2.86 x 10-3 N-m is important due to the lengthy two-month period that this satellite must be in Earth orbit.

This torque causes the reaction wheels in Satellite 1 on low-Earth orbit to saturate every 4.86 hours. Desaturation is

completed using the TCAs. The amount of propellant consumed during this two month period is 60 kg by Satellite 1.

Desaturation is similar to Satellite 2 at about 5 hours, but not propellant-expensive because of the short duration in

Earth-orbit.

I.1.3 Lander Deployment Method

In order to prevent a shift in the center of gravity (C.G.), all

landers on the satellites and orbiters are deployed simultaneously,

and synchronously. This will be done using pyrotechnic bolts. For

satellite 1, it is unavoidable that the third lander opposite the

HiRise camera be deployed without imparting momentum into the

satellite. However, this momentum change can, and must be,

corrected using the RWA and TCAs. One mitigation strategy is to

have the lateral thrusters on the TCAs adjacent to the HiRise, and

longitudinal thrusters beneath the HiRise fire when the lander

opposite the HiRise is deployed. This is shown in Fig. I.2. All

remaining (and opposing) landers in satellite 1 and 2, will be

deployed in pairs to prevent significant lateral shift of the C.G.,

which needs to be maintained for ACS thrusters.

I.2 Lander ACS/GN&C Overview

The ACS and GN&C system on the lander is arguably one of the most crucial subsystems as it is necessary to

ensure a safe landing on the uneven and uncertain Europan surface. The seven landers incorporate a 3-axis system,

whose major actuating component are the four TCAs. Unlike the satellites, longitudinal thrusters are coupled to allow

for quicker maneuvers during descent, and because thruster plumes do not interfere with any major components. Much

of the payload, including the MARCI and MARDI cameras will double up as an ACS component during the 92-second

lander deorbit, descent, and landing (DDL) phase. A GoldenEye 3D Flash LIDAR is used for altimetry and IMU for

inertial parameters estimation (including acceleration, velocity, and position, in increasing degree of error).

I.2.1 Lander Navigation

Lander navigation is crucial element from the time of deployment and until the completion of the DDL phase. The

landers are equipped with 3 sun-sensors to allow for attitude positioning during mapping. Obit determination will be

done through opnav using the MARCI cameras. Doppler shift and transmitted ranging pulses between the satellites

Fig. I.2. Third Lander Deployment

and Countering Reactionary Impulse

and landers will be used for position and velocity estimation. A navigation methodology known as Terrain Relative

Navigation (TRN), will be used during the DDL phase. It is adopted from NASA’s Autonomous Landing and Hazard

Avoidance Technology (ALHAT) method, which is a TRL 7 method designed for lunar landing, and may need to be

refined for Europa landing.

The TRN method is comprised of three phases: global position estimation, local position estimation, and velocity

estimation [45]. The 54 x 54 km mapped region obtained during the detailed mapping phase will serve as the required

a-priori reference map for global and local position estimation. This mapped region will be correlated, or matched,

with on-the-fly images obtained using the MARDI and MARCI cameras and LIDAR altitude data for global

estimation. The LIDAR will be used local position estimation, and velocity estimation [46]. The onboard estimation

algorithms include the Image to Map Correlation (IMC) algorithm for position estimation, and Descent Image Motion

Estimation Subsystem (DIMES) algorithm for velocity estimation [45]. Figure I.3 shows the general landing approach

to generate an optimized trajectory prior to, and during the DDL phase. Monte-carlo simulations are required during

the CDR phases for verifying the landing site accuracy of the TRN method.

I.2.2 Landing Site Down-selection

The selection of the landing site is completed prior to the DDL phase, and requires

the use of the 54 x 54 km mapped region. The criterion for a possible landing site is that

the effective terrain slope, or inclination, must be less than the lander tipping angle of

49.2o. This angle includes a 50% factor of safety to account for possible slippage during

landing. It must be noted that this tipping angle includes the presence of obstacles, such

as rocks and boulders as shown in Fig. I.4.

I.2.3 Lander Deorbit, Descent, and Landing Phase

The DDL phase involves four major steps as shown in Fig. I.5: deorbit by canceling a large portion of the orbital

velocity, a gravity turn and free-fall phase, hazard detection and avoidance (HDA) phase [2], and landing phase.

Fig. I.4. Lander Tipping Angle

Fig. I.3. General Approach for Generating Optimized Trajectory

Process 54 km x 54 km image map from recon.

phase

Select 1 km x 1 km region that meet

landerconstraints

Obtain MARDI cam. & LIDAR data

Match/correlate image to mapped environment to

obtain :- Horizontal Vel

- Relative Position

From LIDAR:- Vertical Velocity

- Altitude

Generate optimized

trajectory to desired region

Deorbit occurs once an optimum trajectory has been selected by the process described in Fig. I.3. All In the event

that deorbit occurs precisely at the periapsis of the landers’ final 200 km by 2 km orbit, the orbital velocity is 1.4 km/s.

The deorbit burn would cancel out most of this velocity until only a 50 m/s horizontal velocity remains. This low

velocity allows for a gravity turn maneuver. This is followed by a deorbit cleanup maneuver that further reduces

horizontal velocity to 10 m/s. The MARDI (and in the event of failure, the MARCI), and LIDAR are then turned on

for the remainder of the DDL phase. The lander is left to free-fall from an altitude of about 1.785 km to 1.285 km,

after which the powered descent stage begins.

The powered descent stage incorporates both the hazard avoidance and detection phase, and the vertical descent

phase. The hazard avoidance phase involves primarily the LIDAR, and is used to avoid geographical obstructions

impeding the flight path and exceeding the landing site criterion. An extra 50 m/s ΔV is included for both polar and

non-polar landers to account for any burns during hazard avoidance. The horizontal velocity is cancelled by the time

the lander reaches the landing site, and the vertical velocity is decreased to 0 m/s before thruster cut-off. This cut-off

occurs about 0.2 m above the surface. The lander will touch the ground at a velocity of 0.2 m/s. This short distance is

Fig. I.5. Deorbit, Descent, and Landing (DDL) Phase

Alt. (km)

2.0

1.0

1.5

0.5

Deorbit BurnHvel = 50 m/s

Hvel = 1.4 km/sDeorbit Cleanup Manuever (w/ ACS)Hvel = 10 m/s

- Alt = 1.785 km- LIDAR On (20 Hz)- MARDI On (x2) (1 Hz)- Freefall to 1.3 km- Hvel = 10 m/s-Vvel = 24.8 m/s

- Alt = 1.285 km- Hazard Avoidance Manuevers & divert- Vvel = 43.9 m/s- Ignition

-Legs Deployed-Reorient to terrain-Vvel = 20 m/s- Hvel = 5 m/s

- Horz Vel Cancellation -Vertical decel.-Vvel = 20 m/s- Hvel = 0 m/s

- Vertical decel.-Vvel = 0.2 m/s

Deorbit Powered Descent

Global + Local Position Estimation & Velocity Est.

required to avoid the possibility of slipping on the terrain and/or breaking the lander legs due to a moderate landing

speed. The entire DDL phase lasts 91.1 seconds.

I.2.4 Lander Maneuver Analysis

The primary burns for the lander include the Europa Orbit Insertion burn and the powered descent burns. Table I.2

provides the 180o rotation maneuver analysis during the 91-second DDL phase. Due to the short period of the DDL

phase, the total maneuver times were limited to about 15 seconds. This allows for seven 360o maneuvers in each axis

during DDL. The analysis uses dry mass moment of inertias as most of the propellant will be consumed during EOI.

Table I.2 Lander Axial Rotation Maneuver Analysis

Lander

Type Axis

Dry Mass

Moment of

Inertia (kg-m2)

Burn

Time (s)

Coast Time

(s)

Total

Maneuver

Time (s)

Required

Propellant, mp

(kg)

Polar

X 47.04 3 6.02 12.02 0.0053

Y 47.4 3 6.04 12.04 0.0053

Z 61.93 3 10.6 16.56 0.0053

Non-Polar

X 39.5 2 6.91 10.9 0.0036

Y 39.9 2 7 11.0 0.0036

Z 51.4 2 9.6 13.6 0.0036

I.2.5 Satellite and Lander ACS System Equipment Summary

The satellites and landers have different ACS equipment. Tables I.3 and I.4 provide a summary of these equipment.

It must be noted that the satellites have 10 digital sun sensors, each with a field of view of 128o by 128o [47]. One sun

sensor was placed on each face of the satellite except on the face where the HGA is located. This allows for sun angle

detection at any attitude, which is essential for emergency mode, where the satellite must be able to reorient itself

relative to a reference object. This placement allows for 4 redundant sun sensors on each satellite.

It must be also be understood that the lander dos not have physical redundancy in its ACS instruments. Instead, it

has functional redundancy, with MARCI and MARDI cameras serving as redundant instruments during the TRN

phase execution. Extra equipment could not be accommodated by each lander, as it significantly increased the mass

and size of the landers, which cause them to impinge the payload fairing.

Table I.3 Satellites Equipment Summary

Item Quantity Mass (kg) Power

(W) Total Line

Mass (kg) Total Line

Power (W) Supplier

Thrusters MR-111C

(including valves)

12 (4 for

redundancy) 0.33 8.25 5.28

16.5 (nominally,

1 pair used at

once) Aerojet

Sun Sensor

(Digital Sun

Sensor) 10* (4 back-up) 0.3 ~1 3 10

Adcole

Aerospace

Star Tracker

(SED26) 2 (1 back-up) 3.47 13.5 6.94 13.5 Sodern

IMU (Litton LN-

200s) 2 (1 back-up) 0.75 12 1.5 12

Northrop

Grumman

Reaction Wheels

(RWA)

(Honeywell

Constellation)

4 (1 back-up) 8.5 22

(nominal) 34

66 (3 operate at

a time due to

angled

placement)

Honeywell

Corporation

Total 30 - - 50.72 118 (nominal)

Table I.4 Landers Equipment Summary

Item Quantity Mass (kg) Power

(W) Total Line

Mass (kg) Total Line Power

(W) Supplier

Thrusters – MR-111C

(including valves) 12 0.33 8.25 3.96

16.5 (nominally, 1

pair used at once) Aerojet

Sun Sensor (Fine Sun

Sensor) 3 ~1 ~1 3 3

Adcole

Aerospace

IMU (Litton LN-200s) 1 0.748 12 0.748 12 PCB

Electronics

TRN Camera (MARDI) 1 0.6 10 0.6 10 Malin Space

Science

Systems

LIDAR (GoldenEye 3D

Flash Lidar) 1 6.5 50 6.5 50

Advanced

Scientific

Concepts

(ASC)

Total 15 - - 13.3 91.5

J. Space Environment Assessment

J.1 Radiation Overview

One of the major restrictions for the mission was the harsh radiation environment that the Jovian system produces.

The requirements from the RFP which had a major impact on the radiation design were the lander mission start date,

mission phase length, and disposal phase. The mission completion date created a major issue because it did not allow

for efficient trajectory that will allow reduced radiation accumulation. The ideal trajectory would be an orbiter around

Europa, but the orbiter would receive a large dosage amount and the concept created disposal problems. The mission

length of 90 days was an issue due to the amount of shielding that was needed in order to allow for a dose factor of

two to the lowest rad-hardened components on the spacecraft. This extra shielding would also cause the landers to

increase in mass. The last problem for the radiation design was the amount of time added to mission for disposal.

Disposal is needed in order to satisfy planetary protection, and protecting the satellites components through this stage

is crucial to ensure the satellite does not impact Europa. The solution used in order to mitigate the radiation problem

was by using the propellant tanks as additional shielding for the sensitive components. This reduced the amount of

outer shielding used which reduced the structural mass. Another reduction method was using a nesting technique. The

nesting technique dictates that the most sensitive components of the spacecraft are in the center of the spacecraft and

the less sensitive components are placed farther from the center. The less sensitive components are also used as

additional shielding for the less sensitive components.

J.1.1 Radiation Environment

The radiation environment at

Jupiter as well as Europa creates a

problem for the mission because of the

placement of the landers on the

surface, as well as for determining the

feasibility of an orbiter or flyby

satellite. The environment around

Jupiter can be seen in Fig. J.1, where

Europa is in the less harmful section of

Jupiter’s radiation environment. Europa’s surface also creates a problem in the placement of the lander. The trailing

Fig. J.1 Jovian Radiation Environment

hemisphere of Europa is heavily bombarded with radiation, therefore efforts were made to limited the number of

landers placed in this high radiation region.

J.1.2 Radiation Material Shielding Selection

The material selected for the outer shell was chosen to ensure the most radiation protection from electrons and

reduce the amount of secondary radiation produced by some materials. The method used to reduce these effects is to

have a series a materials, one being a low-Z material, meaning that it is a low atomic number but has higher chance

of producing secondary

radiation rays, and a high-Z

material to prevent any

secondary rays. A trade

study was also conducted in

order to show the efficiency

of the material thickness and

the amount of reduced

radiation. The most efficient

material was aluminum as a

low-Z material and titanium,

which is used as the material for the propellant tanks as well as a layer in the inner and outer shell of the spacecraft.

A comparison of several different materials, including aluminum and titanium is shown in Fig. J.2. A high density

polymer is also used which is lightweight and produces a significant reduction in radiation.14

J.1.2 Radiation Environment Model

A model using Spenvis, Oltaris, GIRE ,and NOVICE online tools were used in conjunction with results from the

2012 Europa Study report in order to calculate the thicknesses needed in the outer shells and the radiation vault.15 The

Oltaris model used a spherical shell model which simulates the outer shell of the satellites and landers. A thick slab

14 Podzolko, M.v., I.v. Getselev, Yu.i. Gubar, I.s. Veselovsky, and A.a. Sukhanov. "Charged Particles on the Earth–

Jupiter–Europa Spacecraft Trajectory." Advances in Space Research 48.4 (2011): 651-60. Web.

15 Kang, Shawn, MIchael Cherng, Tom Jordan, and Insoo Jun. Total Ionizing Dose Environment for a Jovian Mission

Using Geant4 (n.d.): n. pag. Web

Fig. J.2 Material Trade Study

800

8000

80000

800000

0 2 4 6 8 10

Rad

iati

on D

osa

ge

(kR

ad)

Depth (g/cm2)

Aluminum

Tantalum

Polyethylene

Al-Li-2195

CuW

AlBe

Titanium

was also used in order to model the

spacecraft inner vault. Other models

were used which were derived from

different papers displaying the

radiation reduction with depth in

aluminum for a 90 day mission

period.16 The Europa study report was

also used in order to correlate data

from different models and see if the

models correlate.17 The radiation

model comparison is shown in Fig. J.3, and the Europa Study Report’s radiation predictions are shown in Table J.1.

J.1.3 Shielding Estimates

The method used in

order to calculate the

shielding estimates were

using the Oltaris tool

which gave the closest

estimation of radiation

compared to the Europa

study report. Using the

material trade study the

shielding estimates were

found to reduce the radiation to below the design point of 200krad which gives a radiation factor of two for the overall

mission.

16 Paranicas, C., B. H. Mauk, K. Khurana, I. Jun, H. Garrett, N. Krupp, and E. Roussos. "Europa's Near-surface

Radiation Environment." Geophysical Research Letters Geophys. Res. Lett. 34.15 (2007): n. pag. Web. 17 Administration, National Aeronautics And Space. "Europa Study 2012 Report." (n.d.): n.

pag. Solarsystem.nasa.gov. Web. URL:

https://solarsystem.nasa.gov/europa/docs/ES%202012%20Report%20B%20Orbiter%20-%20Final%20-

%2020120501.pdf

Fig. J.3 Radiation Model Comparison for 90-day Mission Period

1

10

100

1000

10000

100000

1000000

0 1 2 3 4 5 6 7

Ra

dia

tio

n D

osa

ge

(kra

d)

Aluminum Depth (g/cm2)

ES StudyGIRENOVICE&MCNPXOLTARIS

Table J.1 Europa Study 2012 Report Ionizing Dose

Aluminum

Thickness

(mil)

Cruise

(krad)

Tour

(krad)

Photo

Recon

(krad)

Telecom

Relay

(krad)

Total

(krad)

100 5.1 125 358 383 872

200 2.9 52 157 168 380

400 1.8 20.2 57.8 61.7 142

600 1.5 12.1 30.5 32.5 76.6

800 1.3 9.0 19.3 20.6 50.2

1000 1.2 7.5 13.8 14.8 37.4

1200 1.2 6.8 10.8 11.6 30.3

1400 1.1 6.3 9.0 9.6 26.0

1600 1.1 6.0 7.8 8.4 23.3

J.1.4 Radiation Tolerance

Another reduction method used in order to ensure the safety of the sensitive components is to increase the

rad-hardness of those components.

No component should be below 100

krads to enable mission plausibility

and it is expected that development in

components will increase to above

300 krads. Electrical screenings are

used in order to reduce the variability of the radiation response on electronic parts. The most sensitive components

from the different subsystems were found and used as a reference on what system need to be placed where on the

spacecraft.18 The shielding masses needed for the highest sensitivity parts can be located in Table J.3.

18 Administration, National Aeronautics And Space. "Europa Study 2012 Report." (n.d.): n.

pag. Solarsystem.nasa.gov. Web.

0

20

40

60

80

100

120

140

160

180

Do

sage

(kr

ad)

Cruise

Tour

Mission

Disposal

0

20

40

60

80

100

120

140

160

180

200

Cruise

Tour

Mission

Disposal

Fig. J.4 Satellite Radiation Accumulation Fig. J.5 Lander Radiation Accumulation

Table. J.2 Satellite and Lander Shielding Estimations

Satellite Outer

Shell

Satellite

Avionics Box

Lander Outer

Shell

Lander

Avionics Box

0.5 mm

Aluminum

0.5 mm

Titanium

1.0 mm

Aluminum

2.0 mm

Aluminum

1.0 mm

Polyethylene

0.1 mm

Copper

0.5 mm

Polyethylene

1.0 mm

Titanium

0.5 mm

Copper

Table J.3 Subsystem Lowest Radiation Component

Subsystem Lowest Tolerance

Parts

Part Tolerance

(krad)

~Shield

mass (kg)

Payload Instrument/Detectors 300 5.5

ACS Star Tracker/Sun

Sensor

300 Enclosure

C&DH Avionics 300 8

Propulsion Pressure Transducers 100 20

Telecom Telecom Components 300 Enclosure

J.1.5 Radiation Charging

The space environment radiation can create an unstably charged structure which has been a problem for previous

missions enduring the Jovian environment. The solution to this problem would be to limit the differential charging of

external materials by using very low-charging materials. The material used for the structure was a carbon-loaded

Kapton thermal blanket and germanium-coated carbon-loaded Kapton material for the high gain antenna. The

materials used also have treatments that allow the charging of the surface to bleed into the spacecraft ground.

J.2 Spacecraft Torques

The spacecraft is subjected to different types of torques that create issues for the attitude control system. In order

to accommodate for these problems the torque values must be found in order to ensure a sufficient amount of fuel for

spacecraft corrections throughout the entire mission. The assumptions for the spacecraft torque calculations are that

the solar panel reflection factor is 0.3, and the gravity gradient is caused by a 5° rotating offset. The altitudes at which

the worst case torques were taken were at a 400 km Earth parking orbit, 200 km Europa flyby, and 367 km Venus

flyby. The torques calculations were completed using the method from the SMAD book19 and are tabulated in Tables

J.4 and J.5 for the satellite, and the lander respectively. Highlighted values indicate the most detrimental torque

contribution at the particular location.

J.3 Outgassing Effects & Thruster Plume

J.3.1 Outgassing Effectsand Corrections

One space environment phenomena which does not destroy or malfunction the spacecraft but can jeopardize the

mission are outgassing effects. Outgassing is the escape of embedded gas and loose particles from a solid as a result

<https://solarsystem.nasa.gov/europa/docs/ES%202012%20Report%20B%20Orbiter%20-%20Final%20-

%2020120501.pdf>. 19 Wertz, James Richard., and Wiley J. Larson. Space Mission Analysis and Design. Torrance, CA: Microcosm, 1999.

Print.

Table J.4 Satellite Worst Case Torques

Torque Type Earth

(N-m)

Europa

(N-m)

Venus

(N-m)

Solar 4.42 × 10−5 1.63 × 10−6 8.44 × 10−5

Atmospheric 2.86 × 10−3 ~0 6.86 × 10−2

Gravity-

Gradient 3.67× 10−5 5.048× 10−3 3.67 × 10−5

Magnetic 5.67× 10−5 1.4× 10−7 5.32 × 10−5

Table J.5 Lander Worst Case Torques

Torque Type Worst Case

Value (N-m)

Solar 6.58 × 10−8

Atmospheric ~0

Gravity-

Gradient 7.90× 10−8

Magnetic 6.30× 10−4

of reduced surface pressure. The effects of outgassing are the accumulation of condensed particulate that obscure

surfaces such as optical instruments. Local clouds that are formed by outgassing can affect sensitive instrument

readings and also degrade the performance of thermal control surfaces. There are some corrections that can be made

in order to reduce the outgassing effects. One such correction is to use multi-layered insulation which can help trap

and store significant reservoirs of water. The effects of outgassing produces a large product of water which the MLI

can help trap and reduce. The selection of the material used for the spacecraft has a major impact on outgassing effects.

The material chosen should not have higher than a total mass loss of 1% or a collected volatile condensable mass of

1%. Decreasing sun exposure would decrease sun pressure by decreasing incidence angle and shadowing over surfaces

in sensors field of view. This would allow for decreased outgassing effects to sensors and optical instruments.

J.3.1 Thruster Plumes Effects and Corrections

The thruster plumes also create problems when conducting a long mission such as this and even moreso when

considering landing on a surface. Thruster plumes can directly impact the surface or scatter a part of the plume from

one surface to another. These thruster plume impacts can generate turning moments that must be corrected for with

ACS or by creating localized heating which can create outgassing effects or disrupt sensitive equipment readings.

Thruster plumes can also be absorbed by solar arrays and thermal control surfaces. The effects of these absorptions

are decreases in power production and increases in spacecraft temperature. The propellant used for this particular

mission is hydrazine which creates a highly condensable ammonia byproduct which is hazardous to sensitive

equipment. The methods used in order to help alleviate the thruster plume effects include ensuring the thruster nozzle

and primary exhaust are as far from optical and spectrometry instruments as possible. All thrusters must also be

shielded from direct view of payload and sensitive instruments. The thruster plume effects for landers create a problem

for instruments used to land in the desired location such as LIDAR and laser altimeters. The correction for this problem

is the location of these sensitive instruments to ensure correct descent as well as non-sensitive lasers to ensure descent

at close range to surface is unhindered. The study below (Fig. J.7-9) shows the effects of a hydrazine thruster.As seen

in the analysis, the farther away from the thruster a component is located, the less thermal and pressure change there

is. Therefore the sensitive equipment can be place at the edges of the lander which creates the largest distance to

reduce plume effects.20

20 He, Xiaoying, Bijiao He, and Guobiao Cai. "Simulation of Two-phase Plume Field of Liquid Thruster." Science China Technological Sciences Sci. China Technol. Sci. 55.6 (2012): 1739-748. Web.

K. Mass and Power Statement

K.1 Mass Statement

Mass was a major limiting factor in this mission, and because of this, the masses of the landers and the satellites

had to be as small as possible in order to reach Europa and accomplish all of the mission objectives. For the polar

landers, the on orbit dry mass was 241.03 kg, and the total wet mass was 707.19 kg. The non-polar lander’s on orbit

dry mass was 234.38 kg, and the total wet mass was 677.5. The non-polar lander had a lower mass than the Polar

Lander due to a lower propellant requirement during the deployment phase. For both landers, the propulsion system

Fig.J.8 Thruster Pressure Gradient

Fig. J.9 Thruster Particle Distribution a.) Temperature Distribution b.) Pressure Distribution

Fig.J.7 Thruster Temperature Gradient

a. b.

accounted for nearly half of the systems dry mass. 48.6% of the Polar Lander’s dry mass consisted of the propulsion

system, and 47.7% of the

Non-Polar Lander’s dry mass

consisted of the propulsion

system. A detailed mass

statement is shown in Table

K.1.

The on orbit dry mass of

Satellite 1, including the three

attached landers, was 6275.3

kg, and the total launch mass, including the payload attachment fitting, was 10176.1 kg. The on orbit dry mass of

Satellite 2, including the four attached landers, was 6846.9 kg, and the total launch mass was 10732.8 kg. The structure

and the propulsion subsystems were the most massive subsystems of each of the satellites. The Space X Falcon Heavy

was the chosen launch vehicle, and has an estimated payload capability of 12700 kg. This left a positive launch mass

margin of 2523.9 kg and 1967.2 kg for Satellites 1 and 2, respectively. A detailed mass statement for the satellites is

shown in Table K.2.

Table K.2 Mass Statement for the Satellites

Subsystem (w/ contingency) Satellite 1 Mass (kg) Satellite 2 Mass (kg)

Structure (15%) 2019.4 2019.4

Thermal (20%) 69.5 69.3

ACS (10%) 54.3 54.3

Power (10%) 260.2 260.1

Cabling (15%) 314.2 312.7

Prop. Sys. (w/o tanks) 1148.1 1145.9

Telecomm (10%) 100.9 100.9

CDS (10%) 16.9 16.9

Payload (5%) 2291.8 (w/ 3 wet landers) 2867.4 (w/ 4 wet landers)

OODM 6275.3 (w/ 3 wet landers) 6846.9 (w/ 4 wet landers)

Propellant 3481.8 3466.9

Total Wet Mass 9757.1

(w/ 3 polar landers)

10313.8

(w/ 4 non-polar landers)

PAF 419 419

Launch Mass 10176.1 (w/ 3 landers) 10732.8 (w/ 4 landers)

Fal-H Launch Capability ~12,700 ~12,700

Launch Margin +2523.9 +1967.2

Table K.1 Mass Statement for the Landers

Subsystem

(w/ contingency)

Polar Lander

Mass (kg)

60˚ Inclined Lander

Mass (kg)

Structure (20%) 37.2 36.1

Thermal (15%) 6.31 6.31

ACS (10%) 16.95 16.95

Power (10%) 11.8 11.8

Cabling (10%) 19.7 19.7

Prop. Sys. (w/o

propellant) 117.25 111.7

Telecomm (5%) 14.5 14.5

CDS (5%) 9.38 9.38

Payload (10%) 7.94 7.94

OODM 241.03 234.38

Propellant 466.16 443.12

Total Wet Mass 707.19 677.5

K.2 Power Statement

Because each of the landers contained the same payload, and electronics, the power requirements were exactly the

same. The maximum power required by a lander was 273.42 W, with a nominal operating power requirement of 131

W. The maximum power requirement for Satellite 1 was 679.84 W, with a nominal operating power of 420 W. Satellite

2 had a maximum power requirement of 612.84 W, with a nominal operating power of 420 W. Satellite 1 was

responsible for mapping the surface of Europa to determine landing sites, therefore, it carried the MARCI and HiRISE

cameras, and a Mercury Laser Altimeter, where Satellite 2 did not. For this reason, the maximum power required for

Satellite 1 was higher than Satellite 2. Both satellites had the same nominal power requirements. A detailed power

statement of the landers and the satellites is shown in Table K.3.

Table K.3 Power Statement for the Landers and Satellites

Subsystem Satellite 1 Subsystem

Power (W)

Satellite 2 Subsystem

Power (W)

Single Lander Subsystem

Power (W)

Thermal control 154 154 63

ACS 122 122 91.5

CDS 45 45 35

Communications 237.9 237.9 26.3

Propulsion 45 45 46

Mechanisms 4.94 4.94 2

Payload 71 4 9.6

Max Total 679.84 612.84 273.4

Nominal Total 420 420 131

L. Program Overview

L.1. Life Cycle

The program life cycle began in late 2014 with conceptual and preliminary design, shown as Phase A/B in

Fig.L.1, extending until mid-2015. Next, the critical design phase, C, begins and continues until late 2017. Commercial

off-the-shelf components are utilized in the design in order to expedite the production of the spacecraft and meet the

mission time constraints imposed by the RFP. The 23 month manufacturing, integration, and testing phase begins and

continues until the launch date in October 2019, shown in Phase D. The mission phase then begins, including deep

space travel, Jupiter orbit insertion, deployment of the landers, data collection, and disposal, and lasts for 90 months.

After all data collection is achieved, the disposal phase begins in June 2027, and lasts until 2030, at which the entire

lifecycle will be completed.

Fig. L.1 Program Life Cycle

L.2. Risk Analysis

One of the biggest risks involved with the lander design was the use of the toroidal propulsion tanks, because of

their technical readiness level of 7. In order to ensure that they are capable of surviving and performing as expected

throughout the mission, extensive

stress analysis at maximum

loading conditions must be

performed. Figure L.2 shows a risk

waterfall for the use of the toroidal

propulsion tanks.

Another risk involved with the

design is the use of two satellites

to receive and transmit sensor data

from the landers. If a lander is unable to communicate with one of the orbiters, collected data may not be able to be

transmitted back to Earth. In order to mitigate this risk, each satellite has been designed communicate with all seven

Fig. L.2 Toroidal Propulsion Tank Risk Waterfall

landers and to be able complete

the mission independent of

each other after the deployment

phase. Also, redundant

communication systems have

been implemented in each

satellite and all landers. Figure

L.3 shows a risk waterfall for

the lander and orbiter

communication.

The satellites and landers are subjected to the largest loads during launch. The possibility of the system being

damaged because of excessive

vibration due to large fixed-based

modal frequencies is a formidable

risk. In order to mitigate this risk,

finite element analysis to analyze

vibrational modes has been

completed, and extensive structural

testing will be performed prior to

launch. Figure L.4 shows a risk

waterfall for structural damage due to excessive vibration.

M. Cost Analysis

The cost analysis for a mission becomes vital to ensure mission cost feasibility and determine whether the cost is

acceptable in the current economy. The cost models used were the USAF 65’, USCM8, and PCEC. The cost models

all have cost estimation from different time period which would allow for an inflation factor used in order to be

equivalent to the timeframe of the mission. The inflation factor used was 3.1% which was the average from 1926 to

2015. The year estimated for the mission would be in the year 2020 when the mission would launch. The launch

vehicle cost is taken from the Space X Corporation, which estimates its Falcon Heavy at $100M per flight in the year

Fig. L.3 Communications Risk Waterfall

Fig.L.4 Launch Vibration Risk Waterfall

2015. Software cost does not come into account for these cost models and must be estimated in order to more

accurately estimate cost. The average line of code is ~ $17.50 with an average of 106 lines of code for each the satellite

and lander. The overall software cost for the lander and satellite assuming that satellite and landers have similar

software comes out to $35M.

M.1 USCM8 Cost Model

Overall cost for mission with added software cost of $35M, ground ops and tracking cost of $280M, and launch

vehicle cost of $200M comes out to $3.93B using the USCM8 cost model.

M.2 USAF 65’ Cost Model

Overall cost for mission with added software cost of $35M and launch vehicle cost of $200M comes out to $4.90B

using the USAF 65’ cost model.

Table M.1 USCM8 Cost Breakdown

Subsystem Sat. Cost ($M) Seven Lander Cost ($M)

Spacecraft Bus, structure,

thermal, ACS, Power, RCS,

TT&C

2105 710

Surveillance/Imaging

Camera* 9.7 1.6

Microwave Payloads

(Altimeter or Seismometer)* 12.1 1.8

Integration, Test, & Assembly 272 27

Ground Equip. & Flight

Support 188 29

Program Level costs 108 11

Total ~2700 ~774

Table M.2 USAF 65’ Cost Breakdown

Subsystem Orbiter 1 Cost

($M)

Orbiter Cost 2

($M)

Seven Lander Cost

($M)

DT&E 990 990 1159

Facilities 182 182 189

AGE Production 171 171 107

Hardware

Production 56 55 40

Operations 168 160 130

Total 1522 1562 1625

M.3 PCEC Cost Model

Overall cost for mission with added software cost of $35M and launch vehicle cost of $200M total cost comes

out to $4.49B using the PCEC cost model.

M.4 Cost Analysis Overview

This mission to Europa can be compared to missions that have either had landers or have been in the same region

of space. This comparison shows a fairly large disparity in costs between this mission and previous missions. (Fig.

M.1) The reason for this is because of the narrow requirements given for the mission that required (as per this design)

the need for two satellites which is similar to the cost of two separate missions. The other reason for the large difference

is because of the need to land seven landers, which is considerably more than the one lander most comparable missions

boast.

Fig. M.1 Cost Model Comparison

0

500

1000

1500

2000

2500

3000

3500

4000

4500

5000

FY2

0 (

$M

)

USAF

Juno

Cassini

Europa Study

MSL

USCM8

PCEC

Table M.3 PCEC Cost Breakdown

Subsystem Satellites Cost

($M) Seven Lander Cost ($M)

DDT&E 1490 1645

D&D 549 384

STH 523 609

Flight Unit 503 611

Production 503 611

Total 1994 2257

N. Requirements Compliance

Table N.1 shows the requirements compliance for the provided RFP, with the compliance status descriptions outlining

the main highlights for compliance.

Table N.1 Compliance Matrix

Requirement: Compliance Status & Descriptions

- At least 7 landers on

Europa by Dec. 31st, 2026

- Lander/orbiter system

operate for at least 90 days

Satisfied - VEGA traj. w/ arrival in 2026.

- Seven landers & two carrier satellites

- High radiation design point = 400 krad for lander, 300 krad for

satellites lasts 90+ days

- Solar panels for orbiter (~85% efficiency EOL); Quantum Wells for

lander power last 90+ days

- Seismometer senses P-, S-,

and L waves

- Camera capable of panning,

tilting, auto-exposure, and focus

from 10 m to infinity, 2π

steradian coverage, one image for

4o of solar elev.

Satisfied - MEMs seismometer on lander legs (found one with large broadband)

- BeagleCam 2 camera meets or exceeds camera req’s

- Helical boom with tiltable & pannable camera mount 360o panning

& tilting

- Assumed to accumulate 2000 images over mission (overapproximation

to allow room for error). Only 1440 images necessary for 2π coverage

- Lander locations on

Europa: Logarithmically placed

apart

Satisfied - Three polar landers able to land in highly inclined landing sites

- Four 60˚ inclined landers able to cover all other landing sites

- Regular data transfer to Earth - Provide data relay satellite

operational plans & design

details

Satisfied: - Seismographic data and pictures sent to orbiter from lander via LGA.

Data rate = ~85.7 – 122.5 kbps w/ positive downlink & carrier margins for

landers (depending on location) - Downlink to DSN from orbiter via HGA Symbol rate = ~241.5 ksps w/

positive downlink & carrier margins

- Delivery system safely places

landers on Europa - Determine necessary

Software/hardware - Design imaging system

to map lander locations

(if applicable)

Satisfied: - MARCI & MARDI camera and Mercury Laser Altimeter (MLA) on

landers “scout” possible lander locations for 2 weeks before descent - MARDI also used during descent to avoid dangerous terrain

- Prepare strawman mission

description Satisfied - ConOps provided with detailed analyses

- Determine economic

breakpoints for data gen.,

storage & transmission

Satisfied - Cost analysis complete.

III. Conclusion

The Europa CT Scanning Program RFP is met by using seven landers atop two carrier satellites. The mission is

inspired from the Cassini-Huygens and Voyager missions. Given the high scientific potential of this mission, the

number of landers, and the unlikeliness of developing a similar mission in the near future (aside from Europa Clipper)

low-risk, planetary protection, and redundancy were selected to be of importance. By utilizing two satellites, the

mission design allows for redundancy in lander deployment. In the event that one satellite fails, the other satellite is

capable of deploying its landers. It also allows for greater redundancy within the landers in terms of subsystem

components by lessening the volume and mass constraint. Lastly, the staggered orbits of the two satellites around

Jupiter serve as redundancy in telecommunications. At any point in time, one lander can view one of the two satellites

to relay information. Lastly, dual satellite mission and the larger mass margin associated with it allow the satellites to

be disposed outside of Europa. Disadvantages of this mission include the large cost of about $5.0 billion, and the short

time period to design, develop, manufacture, and assemble seven landers’ and two satellites’ components. This latter

risk is partly mitigated by using commercial off-the-shelf (COTS) components. However, the usage of the flex-rolled-

up ROSA solar array and the quantum wells as power system does pose a limitation due to the low TRL levels.

Possible ways to mitigate this include flight testing these power systems on smaller satellites, such as Cubesats, which

will be funded by Aeolus Technologies for rapid test-and-development. Despite these issues, the advantages of this

design ensure the regular transmission of in-situ scientific data back to Earth, which is required for complementing

Europa Clipper’s scientific data and for satisfying the scientific community.

Acknowledgments

The authors of this proposal would like to thank the following people for their support and assistance with this

project: Steve Edberg, Donald Edberg, Bjorn Cole, Try Lam, Mohammed O. Khan, Jessica Samuels, Sumita Nandi,

Bradley Clement, Christopher Delp, William Mcalpine, Jerry Horsewood, Damon Landau, and Aditya Chakraborty.

References

Trajectory

[1] Landau, Damon. “Trajectory Options to Jupiter 2015 - 2025” Jet Propulsion Laboratory. (2007) Web. n.d. 02

Feb. 2015

[2] “Europa Study 2012 Report”, NASA JPL D-71990, 01 May. 2012.

[3] “Delta IV Payload Planners Guide”, United Launch Alliance (2007).

[4] “Chapter 5 Detailed Planetary Protection Requirements”, NASA NPR 8020.12D. (2011)

[5] MAnE, Mission Analysis Environment, Software Package, Ver. 3.5, Space Flight Solutions, Hendersonville,

NC, 2011.

[6] "Arabsat Contracts Go to Lockheed Martin, Arianespace and SpaceX." Spaceflight Now. Web. 2 June 2015.

[7] “PIA07087: Telescopic Camera for Mars Reconnaissance Orbiter, Front End”, Photojournal, Jet Propulsion

Laboratory. 21 Dec. 2004. Web. 10 Jan. 2015. URL: http://photojournal.jpl.nasa.gov/catalog/PIA07087

[8] “Mars Color Imager (MARCI)”, Mars Surveyor 98. NASA, Web. 02 Feb. 2015. URL:

http://mars.nasa.gov/msp98/msss/marci.html.

[9] Delamere, Alan et al. “MRO High Resolution Imaging Science Experiment (HiRise): Instrument

Development.” NASA. 16 Apr. 2003. Web. 10 Jan. 2015.

[10] Malin, M.C., et al. “Mars Color Imager (MARCI) on the Mars Climate Orbiter”, Journal of Geophysical

Research Vol. 106, No. E8 (2001): 17,651-17,672. Web. 02 Feb. 2015. URL:

http://www.msss.com/mro/marci/references/1999JE001145.pdf

[11] “Mercury Laser Altimeter (MLA)”, MESSENGER. JHU/APL. 2015. Web. 10 Jan. 2015. URL:

http://messenger.jhuapl.edu/instruments/MLA.html

[12] Naemi, Anna Willer. “3-axis Fluxgate Magnetometer Model FGE”, DTU Space: National Space Institute.

DTU Space. 13 Dec. 2013. Web. 19 Feb. 2015. URL:

http://www.space.dtu.dk/English/Research/Instruments_Systems_Methods/3-

axis_Fluxgate_Magnetometer_Model_FGM-FGE.aspx

[13] Cavanaugh, John F et al. “The Mercury Laser Altimeter Instrument for the MESSENGER MISSION”,

NASA Technical Resports Server (NTRS). n.d. Web. 07 Jun. 2015. URL:

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20060020062.pdf

[14] “Mars Science Laboratory (MSL) Mars Descent Imager (MARDI)”, Malin Space Science Systems. Malin

Space Science Systems. n.d. Web. 10 Jan. 2015.

[15] “Mars Pathfinder Instrument Descriptions”, Mars Pathfinder. n.d. Web. 10 Jan. 2015. URL:

http://mars.jpl.nasa.gov/MPF/mpf/sci_desc.html#IMP

[16] “Planetary Landers and Entry Probes”, Cambridge University Press, 2007, Web.

[17] Griffiths, A.D, A.J. Coates, J.-L. Josset, G. Paar, and M.R. Sims, “The Scientific Objectives of the Beagle 2

Stereo Camera System”, University College London, 2003, Web. 10 Jan. 2015.

[18] “GeoLight 7 Seismic Sensors”, Silicon Audio, Scribd. n.d. Web. 10 Jan. 2015.

[19] Merchant, B. John. “MEMS Applications in Seismology.” Sandian National Laboratories. 11 Nov. 2009.

Web. 03 Mar. 2015.

[20] “SEIS Instrument”, Centre National D’Etudes Spatiales, 04 Mar. 2015. Web. 10 Jan. 2015.

[21] Robert, O. J et al. “The Insight Very Broad Band (VBB) Seismometer Payload.” 43rd Lunar and Planetary

Science Conference (2012): n.d. Web. 10 Jan. 2015.

[22] “Instrumentation: MARCI ”, Malin Space Science Systems. Malin Space Science Systems. n.d. Web. 10 Jan.

2015.

[23] “Instrumentation: MARDI”, Malin Space Science Systems. Malin Space Science Systems. n.d. Web. 10 Jan.

2015. URL: http://www.msss.com/mars/surveyor/ms98/lander/mardi/section4/section4.html

[24] "Bipropellant Rocket Engines." Bipropellant Rocket Engines. Aerojet Rocketdyne. Web. 12 Jan. 2015.

[25] "Monopropellant Rocket Engines." Monopropellant Rocket Engines. Aerojet Rocketdyne. Web. 12 Jan.

2015.

[26] Fowler, Calum. "Review of Composite Toroidal Pressure Vessel Research Research, Design and Development

for On‐Board CNG and Hydrogen Storage." (2014). Print.

[27] "NASA Technical Reports Server (NTRS)." - A Feasibility Study of Developing Toroidal Tanks for a Spinning

Spacecraft. Part 2: Evaluation of Fluid Behavior in Spinning Toroidal Tanks. Web. 2 June 2015.

[28] "Radioisotope Heater Unit (RHU)." Radioisotope Heater Unit (RHU). N.p., n.d. Web. 24 May 2015.

[29] "Spacecraft Thermal Control: 9. Louvers." Spacecraft Thermal Control: 9. Louvers. N.p., n.d. Web. 24 May

2015.

[30] La-9078-Ms, and Uc-33A. The Light Weight Radioisotope Heater Unit (LWRHU): A Technical

Description of the Reference Design (n.d.): n. pag. Web.

[31] "Spacecraft Thermal Control: 6. Radiators." Spacecraft Thermal Control: 6. Radiators. N.p., n.d. Web. 24

May 2015.

[32] "Aerospace Fabrication & Materials, LLC - Products." Aerospace Fabrication & Materials, LLC - Products.

N.p., n.d. Web. 24 May 2015.

[33] X-762-73-113. SPACEFLIGHT PERFORMANCE OF SILVER COATED FEP TEFLON AS A

THERMAL (n.d.): n. pag. Web.

[34] Bass, John C., Norbert Elsner, Saeid Ghamaty, Velimir Jovanovic, Daniel Krommenhoek, “High Efficiency

Quantum Well Thermoelectrics for Waste Heat Power Generation”, Hi-Z Technology. Web. URL:

http://www1.eere.energy.gov/vehiclesandfuels/pdfs/deer_2005/session6/2005_deer_krommenhoek.pdf

[35] Jurgensmeyer, Austin Lee. “High Efficiency Thermoelectric Devices Fabricated Using Quantum Well

Confinement Techniques”. Colorado State University, Mastere’s Thesis. URL: http://projects-

web.engr.colostate.edu/ionstand/graduates/papers/Jurgensmeyer%20Thesis.pdf

[36] “Basic Elements for Static RTGs”, Rockwell International. URL:

http://fti.neep.wisc.edu/neep602/SPRING00/lecture5.pdf

[37] Shor, Roberta, Robert H. Lafferty, and P.S. Baker, “Strontium-90 Heat Sources”, Oak Ridge National

Laboratory, ORNL-IIC-36. URL: http://web.ornl.gov/info/reports/1971/3445605716035.pdf

[38] “Advanced Solar Array Systems”, NASA, 31 Oct. 2012. Web. 18 Feb. 2015 URL:

http://www.nasa.gov/offices/oct/home/feature_sas.html#.VW91ClxViko (Solar Array)

[39] Bosch, “Traveling Wave Tube Amplifiers (TWTAs) for Space Applications” URL:

http://www.electronicnote.com/site/PDF/TWTA%20Brochure.pdf

[40] General Dynamics, “Small Deep-Space Transponder (SDST) Reliable X-Band and Ka-Band Deep Space

Transmission” URL: http://www.gd-ais.com/Documents/Space%20Electronics/SDST%20-%20DS5-813-

12.pdf

[41] BAE Systems, “RAD750 6U CompactPCI single-board computer” URL:

http://www.baesystems.com/download/BAES_052280/Space-Products--RAD750-6U-cPCI

[42] Airbus Defence and Space, “Fully Qualified Flash Memory Optimizes Satellite Data Storage” URL:

http://airbusdefenceandspace.com/newsroom/news-and-features/fully-qualified-flash-memory-optimizes-

satellite-data-storage/

[43] “Constellation Series Reaction Wheels”, HoneyWell, Honeywell. n.d. Web. 03 Mar. 2015. URL:

http://www51.honeywell.com/aero/common/documents/Constellation_Series_Reaction_Wheels.pdf

[44] “Chapter 13. Spacecraft Navigation”, Basics of Space Flight: Section II, Jet Propulsion Laboratory, n.d. Web.

08 Jun. 2015. URL:http://www2.jpl.nasa.gov/basics/bsf13-1.php

[45] Johnson, Andrew E. and Tonislav I. Ivanov, “Overview of Terrain Relative Navigation Approaches for

Precise Lunar Landing,”, Jet Propulsion Laboratory, Pasadena, CA, n.d.

[46] Johnson, Andrew E. and Tonislav I. Ivanov, “Analysis and Testing of a LIDAR-Based Approach to Terrain

Relative Navigation for Precise Lunar Landing,” Jet Propulsion Laboratory, Pasadena, CA, n.d.

[47] “Fine Sun Sensors FSS”, ADCOLE CORPORATION. Adcole Corporation. (2015). Web. 03 Mar. 2015. URL:

http://www.adcole.com/aerospace/high-accuracy-fine-sun-sensors/fine-sun-sensor-fss/

[48] “LN-200S: Inertial Measurement Unit”, Northrop Grumman. (2013) Web. 03 Mar. 2015. URL:

http://www.northropgrumman.com/Capabilities/LN200sInertial/Documents/ln200s.pdf

[49] “SED26 Star Tracker”, Sodern. n.d. Web. 03 Mar. 2015. URL:

http://www.sodern.com/sites/docs_wsw/RUB_154/FICHE_SED26_A4_page12_Avril2013.pdf