SENIOR DESIGN PROPOSAL The Richter Program’s “Europa CT Scanning” RFP

Prepared by Michael Corpuz (Team Lead) Randall C. Acosta (Deputy)

Omar Alhassen Matt Bergman

Brendan J. Clarke Frank Garcia

Kasbar Gulbenli Jeremiah Kho

Sean Matthews Juan Sanchez

Department of Aerospace Engineering California State Polytechnic University, Pomona, CA, 91768

Abstract

This paper outlines how the Kronus team satisfied NASA’s RFP for a “Seismometer

Array and Delivery System Capable of Collecting Seismographic Data Sufficient to Map the

Interiors of Jupiter’s Moon II Europa” and all its requirements. After extensive design and

trade studies, a design of a solar powered, dual-spin stabilized, and liquid bi-prop orbiter

carrying eight solar-powered landers was chosen. The spacecraft will launch on a SpaceX

Falcon Heavy in 2020 and arrive to Europa in 2026 by using a Venus Earth Gravity assist

and a Jovian Satellite Tour. The landers will be placed in a Legendre-Gauss-Lobatto point

distribution and collect seismographic and camera data. The orbiter will transport the

landers to Europa as well as relay all scientific and engineering data from the landers to

Earth. Through examination of all requirements, the proposed design is compliant with all

restraints and requirements and is fully capable of completing the RFP’s mission.

I. Introduction

The official title of the Request for Proposal given to the design team by Dr. Stephen Edberg of NASA’s

Jet Propulsion Laboratory is: “Seismometer Array and Delivery System Capable of Collecting Seismographic Data

Sufficient to Map the Interiors of Jupiter’s Moon II Europa.” Europa is one of Jupiter’s Galilean moons and there is

much speculation that Europa may be able to support life in its large subsurface oceans. However, a mission to

Europa presents a multitude of challenges. Due to Europa’s distance from the sun, the spacecraft will have to deal

with low solar fluxes as well as cold temperatures. In addition, the large doses of radiation and gravitational torques

from Jupiter and the unknown topography of Europa’s surface are factors to take into account as well. This mission

is therefore classified as a NASA flagship mission, due to its scope and scale. The primary goals of this mission are:

to strategically place seismometer array on the surface of Europa that is able to record and read any seismic activity

that may occur due to the subsurface ocean of Europa, expand knowledge and understanding of interior composition

and structure of Europa, and finally demonstrate capacity for inter-planetary exploration. The primary requirements

derived from the RFP revolve around the seismometer and camera payloads as well as a specific landing layout. The

main requirements other than the payload and landing sequence, is to have a minimum 90 days of seismic and

imaging data on Europa and arrive at Europa by 2026. The full list of requirements can be found in the Appendix.

This paper outlines how the Kronus group will satisfy this RFP and all its requirements.

II. Mission Design for a Europa Orbiter

The mission design was broken into three phases: 1) the trajectory from Earth to Jupiter, 2) the tour from

Jupiter to Europa, and 3) the desired Europa orbit characteristics. Each mission phase conducted their own trade

studies in order to optimize a low delta-V (ΔV), low time-of-flight (ToF), and maximize the scientific benefit. A low

ΔV was important in order to reduce the spacecraft’s wet mass; a low ToF was necessary in order to satisfy the

Request for Proposal (RFP) stated Europa landing date of 2026.

Table 2.1: Comparison of VEGA and VEEGA Trajectories to Jupiter

Figure 2.1: Comparison of Venus, Earth, and Mars Gravity Assist Trajectories

After doing a global search of Earth, Venus, and Mars gravity-assist trajectories with JAQAR’s Swing-by

Calculator, it was determined that Venus gravity assists would provide the lowest ΔV and therefore lowest

spacecraft wet mass. However, a direct transfer from Venus to Jupiter is highly inefficient; one, or two, Earth

gravity-assists were sought in order to reduce the Jupiter arrival velocity. Both the single-Earth flyby (VEGA) and

double-Earth flyby (VEEGA) trajectories are shown in Table 2.1. The VEGA’s time of flight from Earth to Jupiter

is 4.3 years, however the second Earth gravity assist (VEEGA) requires a time of flight of 5.8 years—far too long in

order to satisfy the RFP requirements of a 2026 Europa arrival date. Therefore, the chosen trajectory for this mission

was the VEGA, with a total mission ΔV of 2952 m/s.

Figure 2.2: Earth-to-Venus Pork Chop Plot

After determining that an initial Venus flyby would be optimal for a trajectory to Jupiter, a porkchop plot

was generated in MATLAB. This plot, shown in Figure 2.2, allowed for the extraction of a one-month launch

window in March 2020, as well as the respective launch vehicle payload mass. For this mission, only two launch

vehicles were considered, the SpaceX Falcon Heavy and United Launch Alliance’s Delta IV Heavy. With a

maximum launch characteristic energy (C3) of 13 km2 s-2, the Falcon Heavy provided a launch mass of 13,500 kg

while the Delta IV Heavy provided 7,920 kg. Due to the high mission ΔV of the VEGA, the Falcon Heavy was

considered, rather than having to use 2 Delta IV Heavy launches. Although the VEEGA trajectory does not satisfy

the RFP date requirement, the reduced total ΔV does allow the same spacecraft to launch on the Delta IV Heavy,

due to the significant reduction in propellant required.

Table 2.2: Critical Dates for 2020 Venus-Earth Gravity Assist Trajectory

Critical dates for the Venus-Earth gravity assist trajectory are shown in Table 2.2. The earliest launch date,

starting at the one-month launch window, is on February 27, 2020 with a launch characteristic energy of 13 km2 s-2.

As the days in the launch window progresses, the characteristic energy reduces until an optimal launch day on

March 18, 2020. Assuming a one-month launch window, the latest launch possible would be on March 26, 2020,

with a characteristic energy of 13 km2 s-2. After launch, the first flyby encounter is at Venus on July 1, 2020, at an

altitude of 22,000 km. This maneuver is energy increasing: the arrival velocity is 6.38 km/s and the departure

velocity is 6.41 km/s. The second flyby encounter is at Earth on April 28, 2021, at an altitude of 1,300 km. Once

again, this is an energy increasing maneuver, with an increase in velocity of nearly 0.6 km/s. After a 1,140 day

transfer from Earth, the spacecraft will arrive to Jupiter on June 11, 2024, with an arrival velocity of 6.40 km/s.

The next phase was to determine the trajectory from Jupiter to Europa. Two tours were investigated: the

Banzai Pipeline, a low-radiation dose tour, and the 12-L1, a low- ΔV tour. A comparison of these two tours is shown

in Table 2.3.

Table 2.3: Comparison Between Jovian Satellite Tours

The main consideration of the tour was the total radiation dose accumulated. Due to the high radiation

environment of Jupiter, the spacecraft must be outside the region of Ganymede during the tour in order to avoid

excessive radiation. With an already challenging 90-day mission at Europa, any additional radiation dose will further

increase the radiation shielding necessary for the spacecraft. In comparison, the 12-L1 tour had a radiation dose of

124 krad, but the Banzai Pipeline only accumulated 89 krad. The additional ΔV required for the Banzai Pipeline was

worth it due to the significant decrease in radiation shielding required for all 8 landers.

Additional considerations of the tours included the time of flight (due to the RFP requirement), the number

of satellite flybys, the lowest flyby altitude, and the time of flight between each satellite encounter. Due to the

navigational challenges of a satellite tour, the risk assessment of both trajectories were considered. On average, the

Banzai Pipeline had higher altitude flybys than the 12-L1, as well as less critical flybys (encounters with a satellite

less than 500 km). Therefore, navigationally, the Banzai Pipeline was the preferred satellite tour.

One of the most challenging maneuvers of the mission is the Europa Orbit Insertion on February 1, 2026.

Due to the satellite encounter time of flight of less than 3 days, an autonomous orbit insertion burn may be necessary

for mission success. The autonomous navigation phase of the satellite tour is shown in Table 2.4. An alternative

solution this would be ground-based navigation, which would only be able to utilize 1 maneuver per satellite

encounter. In order to increase both the navigational accuracy, as well as the number of maneuvers between flybys,

autonomous navigation would be necessary.

Table 2.4: Critical Dates for Banzai Pipeline Tour

The last phase of the mission was determining the optimal science orbit around Europa, as well as the ΔV

required for a plane change. The RFP requires polar landers, therefore a polar orbit would allow easy access to most

landing sites on Europa. As an additional benefit, a polar orbit allows for global mapping coverage of Europa, which

can be used to seek safe landing site zones with a high-resolution camera. Because the orbiter is using solar arrays, a

special type of polar orbit, the full-sun orbit, allows for the spacecraft’s solar arrays to be pointing nearly directly

toward the Sun. Due to the small angular distance of the Earth and Sun, the spacecraft’s high gain antenna will also

be able to constantly communication with the Earth, except during the Jupiter eclipse.

The ΔV allocation for the entire mission in shown in Table 2.5. Trajectory correction maneuvers were

accounted for from Earth to Jupiter; these included statistical low- ΔV maneuvers between flybys, a Jupiter-arrival

trajectory correction, as well as a worst-case launch trajectory error correction. The 100 m/s ΔV for the deep space

maneuver is the worst-case maneuver, which would only be encountered if launched near the first or last days of the

launch window. Because of the optimal trajectory, if launched on March 18, 2020, there would be no deep space

maneuver necessary. A significant reduction it he Jupiter Orbit Insertion was acquired by preforming an initial 500-

km Ganymede flyby, 3 hours before the JOI burn. The JOI is the required ΔV to be captured into a highly elliptic,

15 RJ by 242.5 RJ, Jupiter orbit. This specific orbit sets the spacecraft up properly for the Banzai Pipeline trajectory.

After a perijove raise maneuver to correct for orbital perturbation, the spacecraft begins the tour with a Ganymede

flyby 95 days after the raise maneuver. Orbital trim maneuvers, which are deterministic, and statistical maneuvers (2

m/s per satellite encounter) were accounted for during the Banzai Pipeline. Lastly, the large Europa Orbit Insertion

inserts the spacecraft into a circular, polar orbit around Europa at an altitude of 100 km. 120 days of orbital

maintained ΔV was accounted for to keep the spacecraft in the proper orbit. Without this orbit maintenance, the

spacecraft’s orbit would eventually degrade into Europa’s surface in about a month.

Table 2.5: Total Europa Mission ΔV

III. Radiation Effects on a Europa Orbiter and Lander:

There are two branches of radiation: non-ionizing and ionizing. Non-ionizing radiation causes damage in

material by the production of heat (vibration) or atomic displacement, while ionizing radiation causes

malfunctioning of electronic devices, especially semiconductors. While passive electrical components (resistors,

capacitors, and inductors) are relatively immune to radiation damage, active devices, such as computer systems,

have four main categories of damage: non-ionizing thermal damage, displacement damage, total ionization dose

damage, and single event upsets. For the preliminary analysis of a spacecraft mission to Europa, both the

displacement damage and the total ionization dose damage was accounted for by considering the effects of non-

ionizing and ionizing radiation, respectively, in the Jupiter radiation environment.

When nonionizing radiation interacts with an atomic nuclei, it has a probability of displacing, or removing,

them from their lattice sites. This displacement damage will ultimately cause a reduction in the lifetime of

semiconductors, and therefore is important for solar cell power attenuation. For analytical purposes, it is a common

standard to express the damage effectiveness of a particle’s energy by using the unit equivalent 1-MeV particle

fluence.

Figure 3.1: 120 Day Particle Fluence at Europa

The accumulation of ionizing radiation for the lifetime of a mission is called the total ionization dose, or

TID. The ultimate outcome of extensive ionizing radiation on semiconductors is the decrease in functionality.

Eventually, the device will have a high probability of failure after a specified radiation rating. The lifetime of

electrical devices can be increased by considering the use of radiation-hardened components, which in general will

be able to accumulate ten to fifty times the radiation dose of their equivalent commercial parts. For a flagship

mission to Europa, it was assumed that the majority of the electronics would be rated for a TID of 100 krad (Si).

For radiation analysis of this interplanetary mission, ESA’s Space Environment Information System, or

SPENVIS, was utilized. Modeling the radiation damage of the spacecraft was a three-step process. First, the external

environment at Jupiter and Europa was modeled. Second, a program was chosen. Third, the environment of the

spacecraft was considered.

Table 3.1: Orbiter and Lander Solar Cell Cover Glass

SPENVIS contains an implemented Jupiter radiation environment package, JOREM (Jupiter Radiation

Environment and Effects Models and Mitigation). This package contains JOSE (Jovian Specification Environment),

a model for the particle environment around Jupiter. In order to access this package, the spacecraft’s reference planet

must be changed to Jupiter. Next, the coordinate generator must be implemented in order to determine the state of

the spacecraft during its mission, in reference to a Jupiter-centric coordinate system. In this case, only the 120-day

mapping and science phase at Europa was considered. Therefore, the state of the spacecraft was set at a 100 km orbit

around Europa, which is at a perijove altitude of 664,792 km and apojove altitude of 677,408 km.

Cover glass

(mils)

120 Day

Fluence (rad)

Power

Attenuation

42 m2

Mass (kg)

120 Day

Fluence (rad)

Power

Attenuation

17.1 m2

Mass (kg)

3 mils 9.44E+15 28% 8.2 9.33E+15 28% 3.3

6 mils 2.01E+15 18% 16.3 1.91E+15 18% 6.6

12 mils 4.86E+14 12% 32.6 3.91E+14 11% 13.3

20 mils 2.09E+14 9% 54.4 1.26E+14 8% 22.2

30 mils 1.40E+14 8% 81.6 6.72E+13 6% 33.2

For solar cell radiation degradation, the EQFLUX program was used. There were two cell types considered:

single and multiple junction. Because both the orbiter and lander used triple-junction cells, only the multiple

junction type was considered. EQFLUX also offers a variety of manufacturers for their solar cell types, including

Spectrolab,

AZUR, and TECSTAR. For this analysis, only Spectrolab as considered. The results are shown in Figure

3.1 and Table 3.1. The power attenuation values are derived from a best-fit curve based on the radiation degradation

parameters in the Spectrolab data sheet [1].

As expected, as cover glass thickness increases, the power attenuation decreases. However, a plateau effect

is present where the power attenuation does not decrease significantly after approximately 20 mils of cover glass,

but the mass of the cover glass still increases linearly. For this mission, it was critical to reduce the total mass of the

spacecraft while maintaining the proper radiation shielding. Therefore, it was determined that the optimal cover

glass thickness ranged between 6 and 12 mils, thus remaining low-mass at the expense of a 12 to 18% power

attenuation at the end-of-mission.

Figure 3.2: Total Radiation Dose Compared To Five Shielding Materials

In order to calculate the shielding radiation dose, the program SHIELDOSE-2Q was used. This program

used the Jovian trapped particle models to estimate the doses behind tantalum, aluminum, titanium, iron, and

copper-tungsten shielding. The targeted material was silicon, which is the material most electronics are made with.

Figure 3.2 shows the result of the SHIELDOSE-2Q program. Tantalum and copper-tungsten are both great radiation

shielding material, however they are expensive to manufacture and are dense. Although aluminum and titanium do

not shield radiation as effective as tantalum or tungsten, they were chosen as the orbiter and lander radiation vault

materials due to their low density and great structural properties.

Table 3.2 provides the complete material-mass relation between all of the five materials implemented in

SHIELDOSE. In general, aluminum and titanium required a shielding thickness about 3 to 4 times the thickness of

tantalum and copper-tungsten. However, due to the significant cost reduction, aluminum and titanium vaults were

still decided upon.

Table 3.2: Orbiter and Lander Radiation Vault Masses

Figures 3.3 and 3.4 both show the radiation accumulation effect of the landers staying up in orbit for 1

month, rather than doing an immediate autonomous landing on Europa. Two important conclusions can be derived

from these figures. First, the orbiter around Europa will receive significantly more radiation than the lander. The

total radiation dose of the orbiter after 120 days is 1560 krad, and therefore the orbiter’s radiation design factor is

3120 krad. The lander received only 567 krad after 30 days in orbit, and 90 days on the surface, and therefore has a

radiation design factor of 1134 krad. The orbiter will receive 3 times the radiation dose of the lander due to the

radiation protection the surface of Europa provides.

Figure 3.3: Total Radiation Dose Accumulated Over 120-Day Mission

The one-month mapping phase before landing puts a significant toll on the lander’s radiation

shielding. One month in a 100 km orbit around Europa is equivalent to four months on the surface of Europa. Due to

this mapping phase, the radiation shielding required by the landers nearly doubled in thickness. Figure 3.5 shows the

comparison of the total ionizing dose (TID) acquired during this Europa mission compared to previously studied

Europa missions.

In order to ensure the accuracy of the estimates from SPENVIS, a similar radiation analysis was done using

boundary conditions from the Europa Explorer and Europa Lander Mission. The Europa Explorer team estimated a

radiation dose of 1.4 Mrad (Si) behind 100 mils of aluminum shielding after a nominal 120 mission. In comparison,

SPENVIS estimated a radiation dose of 1.5 Mrad for the same mission, therefore the percent error was 7.2%. In

comparison to the Europa Lander mission, SPENVIS’ estimates for the orbiter’s total ionizing dose was 1.35%,

while the lander total ionizing dose was 1.65%. Because the radiation shielding accounted for a 100% margin, all of

these errors are well within the accounted margin for this mission.

Figure 3.4: Aluminum Shielding Thickness vs. Radiation Dose Accumulated

Figure 3.5: Comparison of Total Radiation Dose of Europa Missions

IV Mapping and Landing Phases

A. Mapping Phase

It is critical to mission success to be able to map the eight landing sites. This allows the mission operations

team to select the areas with the most solar illumination and the least amount of sloped terrain. Most of the landers

are landing near the poles, so the solar incidence angle near there is almost 90 degrees. This means that any surface

extrusion could cast a shadow over the lander, limiting its solar coverage. This high incidence angle is also an

advantage because it is easier to see hazardous topology because of the long shadows that it will cast. With this

information, the lander can avoid valleys, steep cliffs, and rough terrain that could be mission ending otherwise.

Mapping the surface also provides unique benefits. If the surface structure of Europa is better known, it

would allow scientists to more accurately update Europa’s characteristics including its shape (volume), albedo,

rotation rate, mass, and gravity values. The landing site images can be analyzed to determine seismologically and

scientifically interesting landing sites. It also allows the flight dynamics group to update their landing trajectory

before landing. These mapping photos can also be used to determine a very accurate location of the lander after

touchdown by correlating the high resolution images of the landing site that were taken by the orbiter with the

panoramic photos that will be taken by the lander.

The orbiter will be inserting into a 100 km altitude, near-polar, near-circular, full-sun orbit around Europa.

This orbit was chosen because it gives the orbiter global coverage of Europa, including the poles, and allows for the

maximum amount of sun to reach the solar arrays. For the first 30 days at Europa, the orbiter will be mapping the

surface. Figure 4.1 shows the behavior of the satellite during this time. The orbiter will be using two cameras to

map the surface of Europa, a Wide Angle Camera (WAC) and a Narrow Angle Camera (NAC). For the first 8 days,

the WAC will provide a global coverage mosaic of the surface at a resolution of 150 m/pixel. This includes a 20%

overlap per image to stitch them together. This global coverage mosaic will allow for a global characterization of

landforms and a general evaluation of the landing sites. As can be seen from Figure 4.2, the WAC swath covers the

entire surface after the 8 days. In this image, the red lines are the border of the coverage swath and the colored

circular areas are the landing areas.

Once this period of 8 days is over, the WAC will stop mapping and the NAC will begin to provide high

resolution image mosaics of the landing site. For the next 22 days, the NAC will be collecting image mosaic strips

at a resolution of 1 m/pixel through the pre-determined landing area. These strips can be seen in Figure 4.2 filling in

the landing areas. The two landers near the equator have the worst coverage, so these are the areas that are designed

to. Figure 4.3 shows a zoomed in image of the white area to see more clearly. The landing area is defined in the

RFP to be a circle of radius 137 km. Figure 4.3 shows this area and the orbiter ground tracks that are created

through that area over the period of 22 days. With this worst case scenario, the orbiter passes over that area 16 times

with a mean coverage duration of 2 minutes and a total coverage duration of 43 minutes. The orbiter collects these

strips of high resolution image matrices that are shown as white lines in Figure 4.3. Each image from the NAC

covers an area of 250 km2 and these are stitched together with a 20% image overlap. The lander areas near the poles

will have much better coverage than the one shown in Figure 4.3. With the data rate that is available to the orbiter,

the NAC is able to collect an area of 7,620 km2 per landing site within the 22 days.

All of this image data will be analyzed to select the best landing sites for each of the 8 landers. Landing

sites will be selected by weighing the technical feasibility of the landing site against the seismologically and

scientific desirability of the site.

Figure 4.1: The NAC (blue) and WAC (red) surface swath coverage by the end of the mapping phase

Figure 4.2: The WAC coverage swath after 8 days

Figure 4.3: Worst case NAC coverage of the landing sites

B. Landing Sequence

The landing sequence is also a critical part of the mission. Two architectures were analyzed for the best

and most reliable method to land on the surface.

The first method that was considered is shown in Figure 4.4. At point 1 in this figure, the orbiter inserts

into a 10 km, full-sun, polar, near-circular orbit. This orbit is harder to achieve than the 100 km orbit in architecture

2 because it is so close to the surface, so the burn needs to be very accurate. Once the orbiter is in this orbit, it will

begin to deploy the landers. The landers will then perform a deorbit burn with a solid rocket motor that will provide

a delta-V of 1.432 km/s that will cancel all of the horizontal velocity. Point 2 in Figure 4.4 shows where this burn

takes place. The behavior of the lander during this burn was analyzed in MatLab. The results of this analysis are

shown in Figure 4.5. This is a very short burn (about 12 seconds), which means that the load on the spacecraft is

very high. This high load is a concern for the sensitive seismographic equipment on board. As shown in Figure 4.5,

the flight path angle changes rapidly from zero degrees to a vertical free-fall. The altitude also decreases about 90

meters during this time. Once all of the horizontal velocity is cancelled, the ACS thrusters are then used to cancel

the remaining vertical velocity. The details of this burn are shown in Figure 4.6. This burn lasts 109 seconds and

has a relatively linear deceleration. The burn does not start until the altitude of the lander above the surface is about

5,500 meters, then the thrust stays between 120 N and 100 N. Once the lander is on the surface, the orbiter must

raise its altitude to 100 km because the 10 km altitude is very unstable. This is performed by a simple Hohmann

transfer.

Figure 4.4: Architecture 1 landing sequence

Figure 4.5: Architecture 1 deorbit burn data

Figure 4.6: Architecture 1 gravity turn burn data

The second architecture that was considered is shown in Figure 4.7. In this architecture, the orbiter is

initially in a 100 km, polar, full sun, near-circular orbit. The trajectory of each of the landers is optimized to use the

least amount of fuel, as well as to reach the surface in the least amount of time to reduce the required battery mass.

The landing sequence occurs in three sections that can be seen in Figure 4.7: deorbit burn, coast, and gravity turn

burn. When the lander is 342.9 km away from the landing site on the surface, the deorbit burn will begin. The

deorbit burn performs a delta-V of 800 m/s over a 102 second duration. The details of the lander trajectory and

behavior during this burn can be seen in Figure 4.8. This burn is performed at a constant max thrust and it takes the

lander from an altitude of 100 km to 79.3 km and a flight path angle of 0° to -15.2°. As can be seen in Figure 4.8,

there is also a relatively constant deceleration and the load factor in Earth G’s stays below 4. During this burn, the

lander will travel a total distance along the surface of 208.8 km.

Figure 4.7: Architecture 2 landing sequence

Once this burn is complete, the lander will shut off its engines and coast for 84 seconds. This lowers the altitude

from 79.3 km to 62.1 km. The lander travels 51.2 km along the surface during this period. The lander will also be

preparing for the next critical gravity turn burn.

The gravity turn burn will then place the lander on the surface. At 338 seconds, this is the longest burn and

it will provide the remaining delta-v required for the lander to reach near-zero velocity at the surface. The detailed

trajectory and behavior of the lander during this burn can be seen in Figure 4.7. During this burn, the lander is

lowered to the surface from an altitude of 62.1 km and the lander travels 82.9 km along the surface. The flight path

angle is also lowered from -22° to -90°, so the lander will land upright on its legs. This flight path angle also

Figure 4.8: Architecture 2 deorbit burn data

Figure 4.9: Architecture 2 gravity turn burn data

ensures that the lander is always pointed towards the landing site so the ACS cameras have a constant view of the

destination (More details on the ACS during this period in the ACS report section). This is also a variable thrust

burn that is designed to reduce the propellant mass required and maintain a relatively constant deceleration to keep

the load factor on the spacecraft down. The thrust curve shown in Figure 4.7 is the desired ideal thrust. This thrust

is accomplished by pulsing the engines to obtain that ideal thrust over time. The load factor during this burn stays

below 2.5 Earth G’s, which is very reasonable.

This second architecture was chosen for the landing sequence for several reasons. First, the short burn time

of the solid rocket motor is very risky because there is no time to adjust for errors in the trajectory. Second, the high

loads experienced during this burn cause the structure of the lander to be very heavy and this almost negates the

benefit for using this architecture. This high load can also damage the seismographic equipment onboard.

The trajectory of the landing sequence can also be optimized further by going through a constrained

trajectory optimization. The software package “DIDO” was used in this research to employ the Legendre

Pseudospectral Method for optimization. An example of vertical descent was used to simplify the problem as well

as prove the viability of DIDO. This program calculates the state of the lander at discrete nodes. For this problem,

30 nodes were chosen to be distributed on a Legendre-Gauss-Lobatto spacing distribution. This spacing was chosen

to have higher accuracy in the results. The equations of motion for this simple vertical descent are as follows:

Where y, v, m, k, and Tmax are altitude, velocity, mass, engine throttle (from 0 to 1), and max thrust respectively.

The cost function was selected to minimize the fuel usage. The vehicle is initialized at an altitude of 500 m and a

vertical velocity of -5 m/s. The parameters were also normalized because the optimization code runs more smoothly

and has better convergence values if the parameters are the same order of magnitude. For this example, the scaling

factors were chosen to be 500 m, 10 sec, and 1000 kg, for distance, time, and mass respectively. The minimal fuel

solution for a simple vertical descent was found and the normalized values are shown in Figure 4.10.

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

Figure 4.11: Vertical Descent Minimum Fuel Solution

Figure 4.11 shows the un-normalized values. This figure shows that the throttle does not switch on until about 1

second. This example proves the viability of using DIDO for optimization and this same method can be applied for

optimization for a trajectory in two dimensions. However, the free version of this program does not allow the use of

enough state and control variables for this optimization, so this is recommended for future research.

The eight lander locations seen in Figure 4.12 are laid out in a Legendre-Gauss-Lobatto spacing

distribution. This spacing was chosen so that there is a higher concentration of landers near the poles than the

equator and so that there is global coverage. In Figure 4.12, the colored circles are the areas that each lander can

land in. This lander distribution also avoids the high radiation areas on the trailing edge of Europa, marked with

blue ellipses [4.1].

Figure 4.12: Colored circles are lander locations and blue areas are zones of high radiation on Europa’s

trailing edge

V: Spacecraft Design

The Design process began with the need for two architectures that would be able to effectively fulfill the

mission objective to land a minimum of 7 seismometers with cameras onto Europa's surface. The two designs were

to have one safe design and one radical design. The first design incorporated a dual-spin stabilized spacecraft that

consisted of: a fully solar powered orbiter and eight solar powered soft Landers. The second design incorporated a

three-axis stabilized spacecraft, which consisted of: an orbiter and eight soft Landers that were both powered by

RTGs.

The first design, which used dual-spin stabilization, used an orbiter as a relay for the Landers to

communicate with Earth and as a transportation bus for the eight Landers. The orbiters were modeled in Solidworks

to a fine detail including accurate measurements and placement of all systems to allow for an accurate c.g. location.

The ACS system had a non-spinning top section that includes the High Gain antenna, the command and data system,

the four low gain antennae, and two cameras. The non-spinning top section allows for constant view of Europa by

the low gain antennae and cameras. The bottom section of the spacecraft spun along its major axis so that it would

stay stable while en route to Europa. The orbiter, seen in Figure 5.1, was made out of 2024 Aluminum for its superb

strength to weight ratio and finite element analysis was used to ensure the lightest structure possible while still

handling the Load factors. The three view of the orbiter design 1 was created for its stowed and un-stowed

configuration seen in Figure 5.2 and 5.3, giving its c.g. location and inertia properties. The orbiter was stowed in the

Falcon Heavy payload fairing with a static envelope of 0.3m and a dynamic envelope of 0.2m, seen in Figure 5.4.

Design 1 was chosen to be the better architecture and its full mass summary table is seen in Figure 5.5.

Figure 5.1 Orbiter 1 with all subsystems

Figure 5.2 Orbiter Stowed Configuration 3-View

Figure 5.3 Orbiter Un-Stowed Configuration 3-View

Figure 5.4 Design 1 Stowed in Falcon Heavy Payload Fairing

Figure 5.5: Design 1 Mass Summary

The second design configuration also had an orbiter and eight Landers. The Orbiter was a three axis-

stabilized spacecraft that is seen in Figure 5.6. The orbiter was powered by solar arrays on three sides with the

telecom dish balancing the fourth side. The design had all Landers mounted directly on top of the orbiter, separating

the first four Landers from the second four Landers by a truss structure. The three view of the orbiter can be seen in

Figure 5.6 and Figure 5.7, showing its c.g. locations and inertia values. The design c.g. location was not centered in

the x-direction because of the difference in weight of the telecom dish to the solar array. The unsymmetrical design

made it required to change the design of the ACS system to accommodate the c.g. location. Design 2 wasn't chosen

because of its less than ideal c.g. location and the volume of the orbiter resulting in a very small static envelope.

Figure 5.6 Design 2 Orbiter three view stowed configuration

Figure 5.7 Design 2 Orbiter three view un-stowed configuration

VI: Lander Design

The first design's Lander is a three axis stabilized spacecraft. It has a rectangular shape for simplicity and

symmetrical placement of all its internals to give it a c.g. in the center of in the x and z axis. The Lander is solar

powered with two foldable solar arrays each with diameters of 3.3 meters. The Lander is a bi-prop design with four

cylindrical tanks to fit in the volume of the structure. The Lander has four legs for stabilization that each fold up to

increase the static envelope during stowed configuration in the payload fairing and to give the Lander a damping

during landing. The payload of the Lander is a seismometer and camera, which are placed in the corner farthest

from the solar arrays to allow for maximum visibility. The configuration of the Lander in design 1 is seen in Figure

6.1. Three views of the Lander is seen in Figure 6.2 and 6.3 with locations of the c.g. and the inertia given.

Figure 6.1 Lander Configuration with all subsystems

Figure 6.2 Lander Stowed Configuration 3-View

Figure 6.3 Lander Un-Stowed Configuration 3-view

Figure 6.4 Design 2 Lander

Figure 6.5 Design 2 Lander 3-View

VII: Structural Analysis

The structural analysis started by finding the largest loads applied to the orbiter and the landers. The

largest axial and lateral loads were found to be the launch loads. The launch loads were not available for the Falcon

heavy launch vehicle so the closest thing was used, the Delta IV Heavy launch vehicle loads. The launch loads were

found to be a maximum axial load of 6g and 3g lateral load. The largest axial load on the lander was found during

de-orbit to be a load factor of 4g.

Multiple tests were run: bending stress, buckling tests, shear stress, thermal analysis, and modal analysis.

Using FEMAP to model the orbiter bottom bus, top bus, and lander, they were all set up for each test. The tests

were run with the maximum loads applied by fixing the bottom section and applying the axial loads to test buckling.

The bending stresses were found by setting one side as a constraint and applying the lateral load. Each test was run

with multiple elements: beams, rods, and tubes, till the element that provided the best balance between strength and

mass was found. The structures were each chosen to use circular tubing for their elements.

Figure 7.1: FEMAP Buckling Analysis on Orbiter upper bus

The structural analysis was done on each of the three main structures with multiple configurations. The

truss structure found in Figure 7.1 shows the upper orbiter bus in FEMAP under buckling loads. The thicknesses of

the tubes were found by iterating the tests until the maximum load applied was underneath the yield stress of the

material. The materials used were Aluminum 2024, Steel 4340, and Ti6Al4v. The aluminum 2024 was the best for

the job but after a Northrop Grumman presentation, was changed to use Aluminum 6061-T6 because of its superb

qualities for spacecraft. The stresses from the buckling test, bending stress, and max displacements are shown in

Figure 7.2.

Figure 7.2: Maximum Stresses and Displacements of each structure

The modal analysis was done on the orbiter structure by creating a constraint on the bottom of the structure.

The launch loads were applied and using NX Nastran the modal analysis test was run. The first ten modes were

created by Nastran from frequencies from 0-200 Hz. The payload planner’s guide stated that frequencies above 35

Hz were the main causes of displacements on the structure.

Figure 7.3: Modal Analysis of Orbiter Structure

VIII. Scientific Payload

A. Seismometer Payload:

The first component that was chosen was the seismometer, and the current seismometer design is based off

the Mars Insight SEIS mission that is scheduled to launch in 2016. A trade study was done in order to determine

which seismometer would be used. This trade study is shown in Table 8.1. The reason this design was chosen was

due to the fact that the seismometer will be used in an extraterrestrial surface and because it of its current operating

specifications.

Table 8.1: Trade Study for Seismometer Selection

The Insight SEIS has relatively low mass and low power consumption. It also has 3 Very Broad Band

(VBB) sensors and 3 Short Period (SP) sensors that are able to read a seismic wave in all 3 axes. Furthermore there

needs to be a minimum of 3 seismometer stations placed within an array in order to accurately read the ice shell

thickness at Europa as well as to confirm the presence of a subsurface ocean.(8.1) The current mission design has 8

seismometer stations placed in an array across the surface of Europa.

The Insight SEIS seismometer was chosen as a starting point for the team’s final mission design. The

Insight SEIS was modified with thermal protection in the form of Radioisotope Heating Units (RHU) to maintain a

temperature so that it can operate under Europa’s harsh conditions. The 3 Very Broad Band sensors used in this

seismometer are placed within a vacuum sphere further protecting them and isolating them from outside

disturbances that may occur from the lander movements. There are also 3 Short Period sensors that are placed

outside the vacuum sphere. The seismometer employs 2 electronic boxes that store the data and send it to the

lander’s computer which will then send the seismic data to the orbiter. Figure 8.1 shows the seismometer in a 3D

CAD model view with all the previously mentioned components. Figure 8.2 shows the same seismometer in a top

and side view and with dimensions (in meters). However, due to Europa’s high radiation dosages, shielding is

required for the seismometer in order for it to be operable at the surface of Europa for a 90 day mission. Thus

aluminum shielding at a rating of 2000 mil (5 cm thick) surrounds the seismometer. This thickness is essential so

that the seismometer can survive during the orbiter’s 30 day orbit trajectory around Europa and a 90 mission at the

surface of Europa. Furthermore, Table 8.2 shows the specifications of the seismometer used for this mission.

Figure 8.1: 3D CAD Model of Seismometer

Figure 8.2: Seismometer Dimensions (in meters)

Table 8.2 – Seismometer Specifications

Dimension Value Mass (kg) 2.9 Power (W) 1.5 Data Rate (kbps) 12.5 Frequency Readings (Hz) 0.001 – 50 Min. Operating Temp. (°C) -220

B. Seismometer Component Details:

- Very Broad Band Sensors: There are 3 VBB sensors used in this seismometer design, which also stems

from the Insight SEIS mission. The 3 sensors together are able to read in the three P-, S-, and L-seismic

waves. The noise sensitivity of the VBB is less than 10-9 ms-2 Hz-1/2 (100-18 g2/Hz) at a frequency of 0.001

Hz to 2 Hz.(8.2) A CAD model design representation of the VBB sensor is shown in Figure 8.3.

Figure 8.3: Very Broad Band Sensor

- Short Period Sensors: There are also 3 SP sensors that are also inherited from the Insight SEIS design.

These sensors are placed within the seismometer such that one is horizontal to the ground and two of them

are vertical to be able to read the 3 seismic wave axes. They are also able to read a better noise sensitivity

of 10-8 ms-2 Hz-1/2 (100-15 g2/Hz) at a frequency of 0.1 Hz to 50 Hz.(8.3) Furthermore these are able to resist

external loads of up to 2000g in order to survive major turbulences during mission deployment.(8.4) The

short period sensor is shown as a CAD model in Figure 8.4.

Figure 8.4: Short Period Sensor

- Electronic Boxes: There are currently 2 electronic boxes that are placed outside of the vacuum sphere of

the seismometer. These electronic boxes store the seismic data to be sent to the orbiter, and after sending

the data they erase the data stored so that they can make more available space in order to collect more

seismic data.

- Rotating Sphere: This was a previously proposed design feature in the seismometer design. The purpose

of this was that if the lander did not land in an even horizontal ground, then the seismometer would be able

to re-orient itself such that the seismometer may be able to accurately read any seismic waves from the

ground. However it was later decided that the best course of action would be to re-design the seismometer

boom so that the boom may be able to rotate the entire seismometer if the lander is not properly oriented to

the ground. A similar mechanism for the camera already exists for this purpose.

C. Camera Payload:

The camera for this mission requirement had to be a camera that was already space qualified and with low

mass and low power operational specifications. A trade study was also done to select the camera that best met these

requirements, and it is shown in Table 8.3. Thus the cameras used in the Beagle 2 mission were selected as they best

met the requirements for a Europa mission. Similar to the Beagle 2 mission, 2 cameras are used for this mission for

both redundancy factors and for easier data acquisition. And similar to the seismometer used for this mission, the

camera also has RHU units for thermal protection, and aluminum shielding for radiation protection. Figure 8.5

shows the camera used and the shielded package surrounding the camera. Table 8.4 further shows the camera’s

dimensions and specifications.

Table 8.3: Trade Study for Camera Selection

Figure 8.5: Shielded Camera Package

Table 8.4: Camera Specifications

Dimension Value Mass (kg) 0.175 Power (W) 0.9 Size (mm) 79 x 63 x 75

Temperature Range (°C) -150 to 100 FOV (°) 48 Focus 1.2 m to infinity

Data Size (Mb, per picture, compressed)

1.31

D. Scientific Payload Package Components:

The camera and seismometer have to work in accordance to each other within the deployment of a lander

on the surface of Europa. The camera package itself consists of 2 cameras placed back-to-back and surrounded by

aluminum shielding (minus the focal lens) with 2 additional RHU units. According to the requirements made by the

RFP, the camera is placed atop an extendable boom. The boom lets the cameras be able to see the local ground

above and around the lander for which the payload package is placed in. The boom uses both linear and rotary

actuators to be able to turn the camera 360° in horizontal azimuth and 90° in vertical elevation. Listen to Tame

Impala, they are an awesome band, I saw them live about two months ago. This is a random sentence no coherence

here. The boom further is supported by a separating mechanism that lowers the seismometer package (along with the

shielding) to the bottom of the ground. Thus whilst the boom pushes the camera up, a second boom mechanism

pushes the seismometer to the ground. Figure 8.6 shows the payload package design with a deployed camera boom.

Figure 8.6: Scientific Payload Package with Deployed Camera Boom

E. Payload Actuators:

The mechanisms to move the booms for this design included both linear and rotary actuators. The linear

actuators used for both the seismometer and camera booms were of the L16 Miniature Linear Motion series by

Firgelli Technologies.(8.5) This linear actuator has relatively low power and is small enough in size to be used within

the inside of the booms, as is shown in Figure 8.7. Both the 100 mm and the 140 mm stroke options are used with a

mass of 74 g and 84 g, respectively, and a similar power output of 0.96 W. Four of these actuators are used in the

camera boom, whilst another 4 are used for the seismometer boom. However these actuators are only used once after

the lander is deployed on the surface of Europa. After they are used, the boom remains extended throughout the 90

mission at the surface of Europa, and thus their output is not calculated into the final average power afterwards.

Figure 8.7: L16 Series Linear Actuator

There are 2 types of rotary actuators used for this design; one is to rotate the camera 360° across the

horizontal azimuth, whilst 2 of them are used to tilt the camera from a level of 0° to + 90°. The type M8 rotary

incremental actuator by MOOG Schaeffer Magnetics Division, as shown in Figure 8.8, is used to rotate the camera

package across the 360° of horizontal azimuth.(8.6) This actuator is moving the camera along 4° intervals so that the

camera may be able to take an overlapping mosaic of pictures at the surface of Europa. This rotary actuator has an

average power use of 5 W and it is constantly operating throughout the 90 day surface mission.

Figure 8.8: Type M8 Rotary Actuator

The second type of rotary actuator used is the M3-RS Rotary Smart Stage by New Scale Technologies.(8.7)

This is a smaller actuator that is able to rotate the camera payload up to a 90° tilt elevation angle, and it is shown in

Figure 8.9. There are 2 of these actuators used, however, during the mission phase of the camera they will only go

up to 24° in elevation which is due to the way the solar arrays are placed relative to the camera FOV. Thus the

camera will not be rotated to the local zenith angle since there would not be any significant scientific data at this

zone. The reason for the 90° tilt capability stems from the precaution that the lander may not land in a level surface,

and thus the camera must be able to tilt to see its local surface if the lander lands on an inclined slope. The average

power used by these rotary actuators is 7 W, and they have a mass of 150 g.

Figure 8.9: M3-RS Rotary Actuator

F. Damping Systems/Isolating external disturbances:

This was something that was considered and much research was done on this subject. However, the results

were not significant enough to produce concrete results. The issue stemmed from the fact that the seismometer was

placed atop of the lander and thus disturbances such as the solar arrays moving would affect the data collected by

the seismometer. To mitigate this issue, it was necessary to create a boom that allows the seismometer to be placed

directly onto the ground and thus be able to read seismic activity without any disturbances. The boom then detaches

the seismometer from the rest of the lander, and only the electronic wiring is connected to allow for data to be

transferred. Furthermore, the VBB sensors are placed within a vacuum sphere, whilst the SP sensors have their own

damping mechanisms already installed.(8.8) Figure 8.10 shows how the boom is separating the seismometer from the

rest of the payload.

Figure 8.10: Boom Extending Seismometer to the Ground

G. Radiation Dosages:

Radiation was a particularly challenging problem, especially since the mission duration at Europa was 90

days on the ground, with an additional estimated radiation dosage accumulation of 400 krad. The most sensitive

instruments of the seismometer were the VBB sensors, and they were tested to 15 krad with good operating

results.(8.9) Thus the radiation shielding applied to both the camera and the seismometer was done so with an ideal

rating for up to 15 krad of radiation dosage for 120 days (90 day ground mission, with an additional 30 day orbital

mapping mission). The entire payload of the camera and the seismometer is covered in aluminum shielding due to

aluminum being a less dense metal. Approximately 5 cm of aluminum radiation shielding (2000 mil) was used to

surround the camera and seismometer payloads. However this increased the payload mass by almost 30 kg for each

of the seismometer/camera payload packages.

H. Thermal Protection:

Another challenge in designing a scientific payload for a mission at Europa is the cold temperature

encountered. The temperature at the surface of Europa is -160 °C to -220 °C at the poles. Since RTGs were not

viable for this particular mission, it was thus proposed to use Radioisotope Heating Units (RHUs) to heat up the

scientific payload instruments at the surface of Europa such that they could operate and complete their proposed

mission. The RHUs chosen contain Plutonium-238 and generate 1 W of power, and they are small in size and mass.

They have also been used in previous space missions such as Galileo and Cassini, both of which used over 100 RHU

units in their respective designs.(8.10)

I. Final Seismometer/Camera Payload Design:

After all the components are placed together the final mass and power dimensions are calculated, and these

dimensions are displayed in Table 8.5. From this table, even though the peak power is currently at 16.26 W, the

average power used is actually 8.3 W, since once the camera boom and the seismometer boom are deployed, those

actuators will not be used again (except the rotary ones in the camera assembly). With the current shielding

protection it is estimated that the total dosage of shielding accumulated for the scientific payload package is 15 krad

for 120 days.

Table 8.5: Mass and Power Breakdown of Scientific Camera/Seismometer Payload

Component Mass (kg) Power (W) Seismometer 2.88 1.5 Camera (2) 0.35 1.8

Camera Boom (5 Parts) 0.665 0 L16 Linear Actuator (4) 0.326 0.96

Type M8 Rotary Actuator 0.30 5 M3-RS Rotary Actuator (2) 0.30 7

RHU Units 0.12 0 Rotating Stick 0.405 0

Seismometer Separator 0.072 0 Camera Shielding 7.32 0

Seismometer Shielding 17.14 0 Total Dry Payload 5.42 0

Total Shielding 24.46 0 Total 29.87 16.26

J. Scientific Payload Mission Design:

After the orbiter deploys a lander onto the surface of Europa, the seismometer is deployed by a boom

towards the ground such that it may be able to read seismic activity straight from the surface of Europa. Then the

camera boom deploys the camera upwards so that it can start taking photos. This deployment along with the relative

orientation with the lander is shown in Figure 8.11. Here the seismometer is constantly recording seismic activity

data at an average rate of 12.5 kbps. The camera, in accordance with the RFP requirements, takes one picture for

every 4° of horizontal azimuth for a 360° FOV for a total of 90 pictures per 360° mosaic. The camera will change its

orientation in terms of tilt elevation depending on where the lander is located at. For the landers at the poles the

orientation does not change much, but for the landers not at the poles of Europa, the orientation can change up to

24° in tilt elevation. This way the cameras may be able to take mosaics that range from a different terrain

environment without looking at the “dead” zenith angle zone.

Figure 8.11: Deployed Seismic and Camera Configuration atop of a Lander

The amount of seismic data collected throughout the 90 day planned mission is the same for all 8 landers

across the surface of Europa. However, the amount of data that will be transmitted back changes depending on the

amount of time available for the orbiter to communicate with the lander. Landers at the poles of Europa are able to

send back all the seismic data, but the other landers have to compress their seismic data sizes in order for them to be

able to send the seismic data back to the orbiter. I wonder if anyone is even reading this. Well I was listening to

Unknown Mortal Orchestra whilst typing this up; their new third album came out today by the way. This is just a

random sentence, no coherence here. Another constraint that limits the amount of data that can be transferred back,

and these are the low-gain antenna data rate, which is assumed to stay at a constant rate of 149 kbps for the 90 day

mission at Europa. The amount of seismic data that is sent back to the orbiter is tabulated in Table 8.6.

Similarly, the amount of data that the camera can send back is also tabulated in Table 8.6, as well as the

amount of pictures and 360° FOV mosaics that are proposed to be taken at each individual lander station. The same

constraints that plague the seismometer data hinder the camera from being able to send too many pictures back to

the orbiter. Nevertheless, the amount of data taken from the pictures is enough to satisfy the RFP requirements.

Furthermore, for the cameras located at the poles of Europa, the cameras will take a few pictures of the planet

Jupiter for aesthetic purposes. The goal of these images is to determine if the terrain at the surface of Europa

changes within a 90 day period, as well as to present pictures with artistic value to the science community and the

general public in whole. Figure 8.12 shows how the camera and the seismometer transmit their data to the orbiter, as

well as how the camera is able rotate on its boom to take multiple 360° FOV mosaics.

From Table 8.6, the total amount of data that is able to be transferred back to the orbiter is 46.5 Gigabits

(Gb) for the camera (35,460 pictures) and 502.5 Gigabits (Gb) for the seismometer. This in total is 549 Gb

transferred back to the orbiter throughout the 90 day mission at Europa’s surface.

Figure 8.12: Camera/Seismic Payload to Orbiter Communication

Table 8.6: Detailed Data Transfer for Camera and Seismometer for 90 day Mission

Lander to Orbiter Transfer Data (90 days)

Lander 1 Lander 2 Lander 3 Lander 4 Lander 5 Lander 6 Lander 7 Lander 8 Total

Time Available to Transfer Data (hr)

221.40 196.64 70.52 43.26 42.68 71.27 197.70 222.29 1065.75

Data Transfer Rate (kbps) 149 149 149 149 149 149 149 149 N/A

No. of 2 pi Mosaics 90 70 15 15 21 19 74 90 394

Total No. of Pics taken 8100 6300 1350 1350 1890 1710 6660 8100 35460

Total Camera Data Transferred (Gb)

10.6 8.3 1.8 1.8 2.5 2.2 8.7 10.62 46.5

Seismic Data Collected (per day) (Gb)

1.08 1.08 1.08 1.08 1.08 1.08 1.08 1.08 8.64

Seismic Data Collected (Gb) 97.2 97.2 97.2 97.2 97.2 97.2 97.2 97.2 777.6

Seismic Data Transferred (Gb)

97.2 97.2 35.96 21.38 20.41 35.96 97.2 97.2 502.52

Total Data Transferred (Gb) 107.82 105.46 37.73 23.15 22.89 38.21 105.93 107.82 549.00

IX. Telecommunications, Command and Data Handling

The design consists of an orbiter with eight landers, each containing it’s own telecommunications and

command and data handling system. The process for down-selecting a system capable of sending and receiving

commands via ground station began using a link budget calculation as shown in table 1, using a similar format to

that found in Brown, table 9.8, as well as Space Mission Analysis and Design by Larson and Wertz. The table

focuses on minimum and maximum range, as well as emergency uplink. Applying data rates and distance range

allows for the carrier uplink and data link performances to be determined, which then permits down selecting to a

system consisting of appropriate rates in order to carry out the mission.

Table 9-1: Link Table (Brown 9.8)

A. Telecommunications Subsystem

For space missions, a telecommunications system generally consists of X, Ka, and S- bands. Initially, an X

and Ka-band were selected for the orbiter, while an S-band was selected for each lander. The X-band would be used

for communication with the spacecraft, while the Ka-band would be used for the scientific data and images collected

to be transferred to ground stations through DSN. The S-band on the landers would be used to transfer the scientific

data and images collected by the lander, to the orbiter. The table below (table 2) lists the equipment initially

determined to be on both the orbiter and landers.

Table 9-2: Equipment List for Telecommunication

A trade study was then done on whether or not the Ka-band was necessary to carry out this mission. JPL

published a paper on the comparison between the X-band and Ka-band, where the advantages and disadvantages of

each were discussed. According to JPL, the X-band is not as power efficient as the Ka-band, which means that the

X-band requires more power to match the Ka-band. However, the Ka-band is much more sensitive to weather,

meaning that there could be a risk of power outages during the phase where data is being sent to the ground. If

power outages occur, this could last up to 30 minutes in time, with a standard deviation of one hour. A power outage

of this duration could risk the possibility of losing data. Therefore, the X-band deemed most reliable for this

mission. Table 3 shown below, lists the updated equipment for telecommunications after this trade study was

performed. The telecommunications system for the orbiter remained with the X-band, removed the Ka-band, while

the landers continued with the S-band.

Table 9-3: Updated Equipment List for Telecommunications

Once this trade study was used to down select which Band would be appropriate for the mission, rates to

allow for a data transfer needed to be determined. Provided that the distance from ground to Europa made this

challenging, using the rates of seismometer and camera used on each lander would determine the rate needed to

transfer data from lander to orbiter. Using the amount of time each lander had with the orbiter, also known as the

“window” for each lander to communicate, as well as how much data needed to be transferred, allowed for the

calculation of how much of a data rate was required. Data transferred from lander to orbiter would be done through

the use of a low gain antenna, rather than a high gain antenna. Low gain antennas have much more of a field of

view, or wider angle, which would mean data would be transferred as long as the orbiter was in view. Even though a

low gain antenna would provide a much lower frequency, it would be much more reliable than the use of a high gain

antenna. It was determined that a 149 kbps data rate for the low gain antenna would be feasible in transferring

camera and seismic data to the orbiter from the lander. Another trade study was done on low gain antennas to

determine if “stacking” antennas in the same direction would be more beneficial (able to transfer more data), than

simply using 4 LG antennas at different angles. The idea of “stacking” antennas would mean that the data rates

would increate by a factor of the amount of antennas used. Having antennas at different positions to “cover” a wider

angle of communication would allow for the lander to be able to communicate with the orbiter for a much longer

duration. The trade study proved that having a longer duration to communicate with the orbiter would allow for

more transfer of data, than to stack antennas for a much higher data rate.

Figure 9-4: Courtesy of JPL. Data Rate comparison using X and Ka-band

Figure 9-5: Courtesy of JPL Ka-Band for different MAR values

Figure 9-6: Courtesy of JPL. X-band for different MAR values

After determining the low gain antenna data rate, as well as the idea of using multiple LG antennas to cover

a wider field of view (increasing the angle of communication with the orbiter), understanding how much of a rate

the HG (high gain) antenna required was necessary. The HG antenna would be used to transfer the data from orbiter

to ground. This data would consist of the picture and seismic data received by the landers, as well as the data

collected during the orbiter’s time mapping Europa. Europa would be mapped using a Narrow Angle Camera (NAC)

as well as a Wide Angle Camera (WAC), during the orbiter’s orbit around the icy moon. In order to be able to

transfer these images during each orbit, a much higher data rate of 360 kbps is required. However, a HG antenna

requires that it be always aligned with its target in order to transfer the data, which means it must be pointed very

accurately. For this reason, it is necessary to provide back up antennas on the orbiter to continue the transfer if the

high gain antenna was unable to. LG antennas would also be attached to the HG antenna in order to ensure that there

is back up to carry out the transfer of data.

B. Command and Data Handling Subsystem

The C&DH system for both the orbiter and landers would consist of RAD750 processors, which seemed

most reliable, as they are very commonly used aboard a number of space missions. A solid state recorder (SRR)

would be present on both orbiter and landers in order to store the data collected, as not all the data would be able to

be transferred at one time. The orbiter would contain a 1 Tb drive, while each lander would only require a 1 Gb

drive to store data. Understanding the amount of data able to be transferred during data collection, and how much

would be stored, would allow for a down select of the size of the solid state required for each lander and the orbiter.

Photo sizes from NAC and WAC, along with HG data rates as well as LG rates (worst case) would allow for the

determination of a 1 Tb drive. Same concept applies to orbiters; however, on a much smaller scale, as each would

require 1 Gb of storage. Redundancy is necessary, as any error in C&DH would doom the mission. A second

processor would be needed to ensure redundancy of the computer system. For this, a second RAD750 processor

would be added to each lander.

C. Flight Modes for Space Mission

For a typical space mission, a number of modes are present onboard each spacecraft. Safety mode is very

crucial, as it is required to preserve the mission if anything was to go wrong. Understanding that anything can

happen, everything must be taken into consideration. This mode would place the spacecraft into a low power mode

with all unnecessary subsystems turned off, in order to preserve the spacecraft. A cruise mode, where low power is

used during the spacecraft’s trajectory, as well as a normal mode, where instruments are powered on during

trajectory in case any scientific data or images are to be gathered. Orbiter Mode is the final mode needed for a

successful mission as this is when the spacecraft will be in orbit, collecting the data once landers are deployed.

X. Power Subsystem

Each design requires two power subsystems, one to power the orbiter and one for the landers. Along with

this, the orbiters and the landers will require peak power estimates as well as average power estimates. The power

systems will be sized based on the average power estimates while peak power situations will be satisfied with a

battery. The power estimates for design one can be seen in table 10.1, while the power estimates for design two can

be seen in table 10.2.

Table 10.1: Design 1 Power Estimates

Orbiter Lander

Peak Power Average Power Peak Power Average Power

Subsystem W % W % W % W %

Thermal 82.00 17.07 82.00 25.68 7.2 3.97 7.2 20.48

ACS 135.78 28.27 56.74 17.77 100 55.09 0 0.00

Power 41.40 8.62 41.40 12.96 2.65 1.46 2.65 7.54

C&DH 20.00 4.16 20.00 6.26 10.8 5.95 10.8 30.73

Telecom 40.00 8.33 40.00 12.53 8.45 4.66 6.5 18.49

Propulsion 34.00 7.08 0.00 0.00 36 19.83 0 0.00

Mech 58.50 12.18 10.50 3.29 0.12 0.07 0 0.00

Payload 68.7 14.30 68.70 21.51 16.3 8.98 8.3 22.76

Total 480.38 100.00 319.34 100.00 181.52 100.00 35.45 100.00

Table 10.2: Design 2 Power Estimates Orbiter Lander

Peak Power Average Power Peak Power Average Power

Subsystem W % W % W % W % Thermal 101.10 20.87 101.10 28.44 7.2 3.36 7.2 20.57

ACS 124.00 25.59 72.00 20.25 177 82.68 0 0.00 Power 41.40 8.54 41.40 11.65 2.5 1.17 2.5 7.14 C&DH 20.00 4.13 20.00 5.63 10.8 5.05 10.8 30.86

Telecom 116.00 23.94 116.00 32.63 8.45 3.95 6.5 18.57 Propulsion 34.00 7.02 0.00 0.00 0 0.00 0 0.00

Mech 48.00 9.91 5.00 1.41 0.12 0.06 0 0.00 Payload 0.00 0.00 0 0.00 8 3.74 8 22.86 Total 484.50 100.00 355.50 100.00 214.07 100 35 100

Deep space missions have historically used nuclear power sources to power spacecraft. Radioisotope

thermoelectric generators (RTGs) have been used on most spacecraft travelling to Jupiter and beyond but Juno has

proved that solar power is feasible at this distance from the sun. Because of this, solar power was chosen for design

one while a combination of solar and RTG power was chosen for design two.

Solar power is dependent on the solar flux reaching the solar array. At Earth, the solar flux is 1370 W/m2

but this decreases based on the inverse square law. At Jupiter, solar flux drops to 51 W/m2 since Jupiter is 5.2 times

further from the sun than Earth is. This means a solar array at Earth will generate 27 times more power than at

Jupiter. This means that solar arrays on spacecraft at this distance have to be very large. Jupiter also has an intense

radiation field surrounding it and Europa lies within this field. This radiation will reduce the efficiency of the solar

arrays based on the thickness of the cover glass protecting the array. The solar arrays will also degrade gradually

with time. All of these effects increase the required size of the solar array. Low temperatures increase the solar

array’s efficiency however. At -130º C, the solar arrays will generate 20% more power.

Both orbiters design will be in the same orbit around Europa. This orbit is a full sun orbit. Not only does

this orbit maximize the amount of sunlight both orbiters will receive but also insure that the orbiter will sweep most

of the surface of the moon. The orbiters will still be eclipsed however. Jupiter is large enough and close enough to

block solar rays from reaching the solar arrays. An STK model was made of the illumination and eclipse times

experienced by the orbiter in its final orbit. This can be seen in figure 10.1

Figure 10.1: Orbiter Eclipse and Illumination Data

Jupiter will eclipse Europa and the orbiter every 69.5 hours. This eclipse occurs for 2.86 hours for each

eclipse. During the eclipse period, the orbiters’ loads will be powered by lithium ion batteries. The solar array will

be powering the orbiter’s loads and charging the batteries during illumination times. Both orbiters designs use rigid

panel solar arrays provided by Spectrolab. The chosen cells used for the solar array are ultra triple junction (UTJ)

GaAs cells. These deliver 350 W/m2 at Earth’s distance from the sun. Using the method in Elements of Spacecraft

Design by Brown, the solar array for design one must be 56 m2 and has a 24% power margin. This is distributed into

four wings of 14 m2. The solar array for design two must be 63 m2 and has a margin of 25%. This orbiter has three

solar arrays, each 21 m2 in area. Figure 10.2 shows the deployed solar arrays for both orbiter designs.

Figure 10.2: Deployed Orbiter Solar Arrays

These solar array sizes were not consistent with the Europa Clipper’s solar array size. According to the

NASA Solar Study Status Report, a 460 W spacecraft would require a 46 m2. Another method of sizing a solar array

was then attempted based on using the data sheet provided by SpectroLab and adjusting the power based on

efficiencies. This method can be seen in table 10.3.

Table 10.3: Solar Array Sizing Example

Area per wing 11.500 m2

Number of arrays 4.000

Total area 46.000

P at earth 350.000 W/m2

EOM degradation 324.557 W/m2

(1.25%/yr from SMAD without radiation)

Temperature adjust 397.582 W/ m2

(122.5% due to lower temperature)

Radiation adjust 357.824 W/m2

(12% radiation efficiency reduction)

Flux Adjustment 0.037

P at Jupiter 13.217 W/m2

Power Generated 607.973 W

Power Available 461.119 W

Excess 141.779 W

Margin 44.397 %

Using this method, a 46 m2 solar array would be able to power a 461 W spacecraft. This result matches the

result from the solar study done by NASA. The method was then applied to the design solar arrays. Design one’s

solar array has an area of 42 m2 and a power margin of 28.9%. Design two’s solar array has an area of 44 m2 and has

a power margin of 26.7%. These solar arrays weighed 168 kg and 176 kg respectively. The solar array

characteristics can be seen in table 10.4.

Table 10.4: Solar Array Characteristics Textbook Method Data Sheets Method

Area Power Generated

Power Margin Mass Area Power

Generated Power Margin Mass

Orbiter One

56 m2 396 W 24 % 224 kg 42 m2 411 W 28.9 % 168 kg

Orbiter Two

63 m2 446 W 25% 252 kg 44 m2 431 W 26.7 % 176 kg

During the eclipse, solar arrays are unable to generate power. In order to power the orbiter during these

times, lithium ion batteries are used. Lithium ion batteries have a higher specific energy than nickel-cadmium

batteries, which means a smaller battery can provide more power. NiCd batteries also suffer from the memory

effect, in which the battery’s capacity reduces after being only partially discharged. Lithium ion batteries are not

affected by this.

The battery cells chosen for design one are the SAFTVES 180. These cells were chosen because they offer

the highest specific energy out of all candidates. The battery specifications are shown in table 10.5. The battery

required by the orbiter was sized using the same illumination and eclipse data as the solar array. The capacity

required for the battery was determined to be 61 Ah at 28 V. To satisfy this, two strings of eight cells will make the

battery. Eight cells in parallel will have a voltage of 28.8 V and two parallel strings will have a capacity of 100 Ah.

This will provide 2880 Wh of energy. The orbiter requires 1710 Wh of energy so the batter will have an excess of

1170 Wh. The second design’s orbiter also uses SAFT VES 180 batteries. This design requires 68 Ah at 28 V. The

battery also requires 16 cells, eight in series and two parallel strings. This battery has an excess energy of 977 Wh.

Table 10.5: SAFT WES 180 Battery Characteristics specific energy 165 W/kg energy 180 W/h mass per cell 1.11 kg Nominal Voltage 3.6 V capacity per cell 50 Ah diameter 0.053 m height 0.25 m

The first lander design utilizes solar power. Since the landers will be on the surface of Europa, eclipses

from both Jupiter and Europa will block the solar arrays. The illumination times for each lander vary since their

locations determine when the Jupiter Eclipse occurs. For some landers, the Jupiter eclipse will occur during the

Europa eclipse and since Europa is tidally locked with Jupiter, the eclipse will always occur at this time. In order to

simplify manufacturing, all eight landers will be identical. All landers will be sized based on a worst case lighting

conditions, in which the Jupiter eclipse occurs during a Europa day. The eclipse and illumination data can be seen

below in figure 10.3.

Figure 10.3: Europa Surface Lighting and Eclipse Time

Again, the shaded regions represent eclipses. The wider bands represent eclipses due to Europa while the

smaller bands represent eclipses due to Jupiter. Together, these eclipses have a duration of 45.78 hours, leaving

39.44 hours of light to generate power. With a power requirement of 35.45 W, the solar arrays need to have an area

of 17 m2. The solar array chosen are the Ultrafelx solar array made by Orbital ATK because of its low mass and

compact stowed size. The ultrafelx solar arrays can be seen below. The solar array ranges in sizes based on

diameter. In order to meet the power requirement, two solar arrays, each with a diameter of 3.3 m are used. This

generates 43.84 W and has a power margin of 23.81%.

Figure 10.4: Stowed and deployed Ultraflex Solar Array Source: http://nmp.jpl.nasa.gov/st8/tech/solar_array3.html

During the eclipse, the lander will be powered by a battery. The lithium ion battery will use SAFT VL 9E

cells, which have a nominal voltage of 3.6 V and a capacity of 11 Ah. Although the power required by the lander is

much lower than the orbiter, the battery capacity required is almost twice as much as the orbiter. The lander will

require a 107 Ah battery since the eclipse time is long on the surface of Europa. In order to satisfy this, 11 strings of

eight cells are used to create the battery. This battery will also be used for peak power situation, the landing phase,

and during times when the power margin generated by the solar arrays is negative.

The second lander would use a RTG power source. An RTG generates power through a temperature

gradient on a thermoelectric generator. The heat source that generates the temperature gradient comes from the

radioactive decay of plutonium oxide. These generators will produce power constantly once the isotope pellets are

installed in the system. The current RTG model being used is the MMRTG, which has been used in the Mars

Exploration Rovers. The MMRTG can be seen in figure 10.5.

Figure 10.5: MMRTG Diagram

Source: https://solarsystem.nasa.gov/rps/docs/MMRTG%20Fact%20Sheet%20update%2010-2-13.pdf

The system provides 120 W at beginning of life (BOL) but suffers power degradation of 3.8% per year.

After six years, MMRTGs will only provide 90.4 W at end of mission (EOM). A modified version of the MMRTG

will have to be manufactured in order be used on the landers since the power output is twice the required power.

Since RTGs generate power based on a thermal gradient, reducing the length of the MM RTG by half will cut the

power generated by half. Reducing the number of radioisotope heating units (RHU) from eight to four will reduce

the mass of plutonium required by each system from 4.8 kg to 2.4 kg.

Although the MMRTG system can continuously produce power, a battery will still be utilized. Unlike the

solar power lander, the battery on this lander will only be used for peak power situations and landing. The capacity

of the battery is 11 Ah at 28 V. This gives an energy capacity of 308 Wh.

At 2.4 kg per lander, a total of 19.2 kg of plutonium is required to power all of the landers. This is more

than the available amount of plutonium. Along with this, an MMRTG system adds more complexity to the power

system. A solar array system can be integrated easier and is cheaper. An RTG system also has more risks of

contamination on Earth during launch and on Europa during the landing. For these reasons, the solar array lander

design was selected.

With the orbiter power source selected, the power system will need a power conditioning unit to convert

the power into a usable form and a power distribution unit to distribute the power to the necessary loads. The power

conditioning unit (PCU) is supplied from Terma. The PCU contains an array power regulator, battery

charge/discharge regulator, and a command and monitoring system. The array power regulator acts as a peak power

tracker, which uses the maximum power needed by the system. The battery charge and discharge regulator controls

the charging and discharging of the lithium ion battery. The command and monitoring system controls the other

components in the power subsystem. The power then goes into the power distribution unit. This unit transfers the

power into the appropriate loads.

The lander’s power control and distribution system is the Clyde Space Small Sat system. The Small Sat

system connects the solar array directly to the battery charge regulator. The power then goes into a power

conditioning module and then into a power distribution module. From there, the power goes to the lander’s loads.

The power system block diagram for both power systems can be seen in figure 10.6.

Figure 10.6: Power System Block Diagram

Since the lander uses solar power, the orbiter will need to supply power to the lander during the flight to

Europa. The lander would not be able to deploy the solar arrays and cannot generate its own power. The orbiter must

power the lander’s thermal system, power system, and command system. Because of this, the lander is considered

one of the orbiter’s loads, which can be seen in the power block diagram. The lander would then separate from the

orbiter when it is deployed.

On the surface of Europa, the lander’s solar array must be able to track the sun to maximize the power

generated. Using the height of the lander and the radius of the solar array, the maximum angle that the solar array

can rotate to track the sun in 37º but a maximum tracking angle of 35º was used. Along with this, a 1 km tall

obstruction placed 10 km away from the solar array was assumed to obstruct the solar array. The power generated at

versus time of day for each of the situations can be seen in figure 10.7.

Figure 10.7: Power Generated vs Time of Day.

The complete mass summary for the power system can be seen in table 10.6. Most of the mass of the

orbiter’s power system comes from the large 42 m2 solar array. For the lander, a lot of mass comes from the battery

required to power the lander.

Table 10.6: Power System Mass Summary

Orbiter

Solar Array 168 Kg

Battery 17.76 kg

Power Conditioning Unit 16.6 kg

Power Distribution Unit 13.2 kg

Total 215.56 kg

Lander

Solar Array 32.02 kg

Battery 21.12 kg

Power Control System 1.5 kg

Total 54.64 kg

XI Thermal System

The thermal environment at Europa is extremely harsh, with an approximate temperature range of -220

degrees Celsius to -160 degrees Celsius at the poles and equator respectively. Compounding the cold temperatures

with the high temperatures experienced during the Venus flyby required this mission to have a very delicately

designed thermal subsystem. As a result, a variety of different thermal control elements were explored, and can be

broken down into two general categories: active thermal control and passive thermal control. The purpose of these

elements is to ensure that the spacecraft (and all of its sub-components, including the landers) remains at an

allowable operating temperature, regardless of the extreme temperatures they are to be exposed to.

The active elements of the spacecraft include: radioisotope heater units (RHUs), resistance heaters, and

louvers. Radioisotope heater units are small devices that produce heat constantly throughout the length of the

mission. In others words, once these devices are activated, they cannot be turned off. The reason for this is because

they produce heat through means of radioactive decay. Resistance heaters, on the other hand, can act as a variable

heat source and are most commonly regulated through the use of either thermostats or solid-state controllers. For

this mission, solid-state controllers were the preferred selection for resistance heater regulation. The details of this

preference will be discussed shortly. Louvers are another active element of the thermal subsystem, and were needed

specifically during the Venus flyby. Louvers can be described as fins whose orientation can be adjusted via a

mechanical mechanism; when open, these fins increase heat expulsion through means of thermal radiation.

The passive elements of the spacecraft include: optical solar reflectors (OSRs), black & white thermal

coatings, and multi-layer insulation. Optical solar reflectors generally have low absorptivity and high emissivity as

characteristics of their thermal properties, and are usually composed of a quartz top-layer with a metallic sub-layer.

This makes OSRs a cold surface. Black & white thermal coatings serve opposite purposes each other. Black

thermal coatings are considered to be “hot” surfaces, as they retain a significant amount of the heat they absorb,

while white coatings (like OSRs) are considered to be “cold” surfaces, since there are efficient at ejecting heat while

absorbing minimal thermal energy. Multi-layer insulation, referred to as MLI for short, is composed of many thin

layers of plastic with a metallic coating. The main purpose of MLI is to ensure little or no thermal conduction

between layers, allowing portions of spacecraft to remain close to a constant temperature.

A. Key Drivers

The design of the thermal subsystem began by identifying the key drivers of the thermal requirements.

These drivers were derived from the required operating temperatures of the science instruments, as well as other key

components of the spacecraft. Table 11.1 (shown below) displays the most significant operating temperature

requirements taken into consideration. RHUs were included in the table with the minimum surface temperature and

do not have a maximum operating temperatures, as they produce heat constantly regardless of the thermal

environment.

Table 11.1 - Significant Operating Temperatures

Component Min. Temp. [K] Max. Temp. [K]

RHU 300 -

Lithium Battery* 233 323

HG Antenna 216 334

Propellant* 263 313

Star Tracker 243 323

Note: Items designated with an asterisk (*), were heavily considered due to their strict operating temperatures.

These items are the batteries and the propellant.

Although the lithium battery can operate over a range from 233K to 323K, it was determined that the

battery loses a notable amount of efficiency when operating outside the range of 273K to 283K, or approximately 0

degrees Celsius to 10 degrees Celsius, leaving a very small window for error in thermal control. Similarly, the

propellant also has very strict thermal requirements, with an operating range that only permits a ±20K temperature

swing from 283K.

B. Thermal Environment at Venus

During the Venus flyby, the spacecraft will reach its closest altitude at approximately 22,000 Km from the

surface of Venus. This is also the portion of the mission where the spacecraft is closest to the Sun, and is thus

subjected to its hottest thermal environment. Two important elements of this Venus flyby include: passing through

the shadow of Venus and using the high-gain antenna (HGA) as a sun shield. Passing through the shadow of Venus

significantly lowers the temperature of the spacecraft, and using the HGA as a sun shield further reduces the

temperature of the spacecraft; the combination of these two flyby elements together work to reduce the total solar

flux that the spacecraft is subjected to. Figure 11.1 (shown below) provides a pictorial of the Venus gravity assist.

Figure 11.1 – Pictorial of spacecraft and solar array temperatures during Venus flyby

It must be noted that the figure shown above provides temperature values that were computed assuming no

thermal protection was in place. Similarly, Figure 11.2 (shown below) is a plot displaying the temperature increase

of the spacecraft as a function of time. Note: this plot was generated using spherical approximations and assuming

no thermal protection was in place. The calculations used to compute these thermal values included shape factors,

as well the use of thermal characteristics such as emissivity, absorptivity, and albedo (reflectivity).

Figure 11.2 – Time plot of temperature during Venus flyby (Note: plot generated using spherical

approximations and assuming no thermal protection)

As can be observed from the plot, the maximum temperature of 363.6 K (approx. 90.44 degrees Celsius)

occurs approximately 3600s into the Venus flyby. At time, t = 0s, the spacecraft is approaching Venus and is at an

approximate altitude of 50,000 Km (Note: this altitude at time is not specified in the plot). It also must be noted that

this plot does not account for the cooling effects of passing through the shadow of Venus.

C. Multi-Layer Insulation

A number of different types of MLI were investigated for thermal properties, specifically regarding the

ratio of absorptivity to emissivity, α/ε. Absorptivity, α, describes the ability of the material to absorb energy from

incident electromagnetic waves. The specification of MLI that was finally selected was from Test Series A140, as

tested by Kennedy Space Center. This particular test coupon was found to have a density of approximately 37

kg/m3. In order to account for the mass of the MLI to be used on the spacecraft and landers, a MLI mass calculator

was developed (shown in Figure 11.3). This calculator was generated using Microsoft Excel 2011, and requires the

input of the spacecraft area (or lander area) that is to be covered in MLI. As can be seen in the figure below, the

initial type of MLI that was used held a higher density of 79 kg/m3. Although MLI makes up only a small portion of

the total spacecraft mass, this specification of MLI was determined to be too massive for the delta-V constraints of

the mission. As a result, a less dense specification (which still satisfied the thermal requirements) was selected for

the mission (Test Series A140).

Figure 11.3 – Image of MLI Calculator Generated in Microsoft Excel 2011

MLI degradation was an important factor that was considered for this mission. The impact factors that

were taken into account include: atomic oxygen (AO) exposure, thermal cycling, micrometeoroid impacts, and

radiation exposure. Since this mission is not an earth science endeavor, it was quickly determined that

micrometeoroid impacts and thermal cycling would not be an issue. Additionally, since there is virtually no

atmosphere at Europa, MLI degradation due to AO exposure was also deemed to not be a concern. The main

degradation factor that was taken into account was that of radiation and particle bombardment (such as the

accumulation of electrons and protons). It was determined that for a mission duration of 90 days, the total radiation

accumulation would be approximately 1500 Krad. This radiation exposure affected the thermal properties of the

MLI, as shown in Table 11.2 (shown below).

Table 11.2 – MLI Properties changes due to 1500Krad exposure, accumulated after 90 days

This is perhaps the only aspect of the mission for which radiation was found to be beneficial. As can be

seen in the table below, absorptance of the pristine (new) MLI sample was approximately 0.13, which then increased

by 28%, reaching a final value of 0.18. On the other hand, emissivity of the pristine MLI sample was approximately

0.79, which then decreased by 5%, reaching a final value of 0.75. In the table, both of these percent changes were

noted as being desirable. These material property changes would cause the MLI to become more of a “hot”

material, and since Europa has such an extremely cold thermal environment, these property changes increase the

MLI efficiency for the purposes of this mission.

XII Propulsion System

The propulsion system of spacecraft is arguably the most vital system when it comes to interplanetary

transfer. Considering the fact that this is a deep space mission with a rather short trajectory, many key drivers had to

be considered in order to ensure a timely arrival.

Since Jupiter is 5.2 AU away from the Sun, the spacecraft was designed to withstand the lowest operating

temperature of 272 K. Also, the power delivered by the solar arrays is limited at 5.2 AU from the Sun. Considering

the spacecraft operating at a low temperature and having limited power resources, the best option was to consider a

propulsion system was able to operate at low temperatures.

Figure 12.1 Propellant Equilibrium Temperatures

Moreover, some of the key drivers for the propulsion system were to maximize specific impulse (Isp), maximize

storability, and have maximum control over thrust variance. With the spacecraft operating temperature estimated to

be 272 K, or 486 °R, the propulsion system for the orbiter and lander was based on Hydrazine having a lowest

thermal equilibrium temperature of about 450 °R. The propellant combination that would give the best specific

impulse was found to be Nitrogen Tetroxide (NTO) and Hydrazine (Hyd), a liquid bipropellant to provide for the

propulsion system.

Out of the many engine candidates, the one who supplied the highest thrust and specific impulse was

Northrop Grumman’s TR-308 Liquid Apogee Engine. The TR-308 engine is able to provide a Isp of 322 seconds,

along with a maximum thrust of 471 N. Two TR-308 engines are needed in order to provide a sufficient amount of

thrust required during the 7 major interplanetary burns.

Figure 12.2 TR-308 Liquid Apogee Engine

The biggest concern for the engines was if it could handle the mission trajectory’s longest burn. This burn would

come at Jupiter Orbit Insertion and was estimated to last around 40 minutes. These engines are rated to have a

maximum firing duration of 50 minutes, yielding a 20% margin just in case the burn needs to last another 10

minutes.

After analyzing the possibilities of the Orbiter spacecraft being a Dual-Spin stabilized or a Three-Axis

controlled, it was concluded that the Dual-Spin stabilized orbiter was more beneficial. A Dual-Spin system is lighter

due to the Attitude Control System being less complex than the Three-Axis control. Since it is lighter, it only needs

12 reaction control thrusters, as opposed to 16, and it requires less Hydrazine to provide for the trajectory control

maneuvers. The Dual-Spin stabilized orbiter requires to have centrifugal tanks in order to maximize the efforts of

the Helium pressurant gas, and to minimize the residual propellant due to the spinning of the spacecraft. The

propulsion system for the orbiter is composed of the following components: pressure transducers, pyrotechnic

valves, system filters, solenoid valves, flowmeters, and latch valves.

Figure 12.3 Spacecraft Propulsion System Overview

For design purposes, two NTO and two HYD tanks were proposed. This would help balance the center of gravity for

the spacecraft while it was fully loaded with the landing units. The tanks were calculated to have a diameter of 0.8

meters, with a membrane thickness of 1.96E-03 m. The design of these titanium tanks are expected to withstand

3300 kPa of maximum tank pressure.

Since this is an interplanetary mission with 8 landing units and has a trajectory with a decently large ΔV,

most of the mass will come from the propellant. Two trajectories were analyzed in order to optimize the ΔV

required, and consequently reducing the propellant mass. The trajectory with a proposed arrival date in 2026 would

require a ΔV of 2.95 km/s, whereas an arrival in 2027 would require a ΔV of 2.41 km/s. The difference in these

proposed trajectories is about 500 km/s. The mass penalty for carrying propellant for this difference in ΔV is about

1,500 kg. The comparison of propellant mass and wet propulsion system mass can be found in the tables below. The

propellant masses listed in the table account for an extra 10% of total ΔV, losses due to 7 major startups, and an

expected 3% residual propellant mass.

Table 12.1 Spacecraft Propulsion System Mass Comparison

Using the Falcon Heavy for a 2026 arrival yielded a very comfortable margin. On the other hand, it would

not be possible to launch with the Delta IV Heavy. The only way that the mission could launch was to reduce mass

in two very important areas: payload and propulsion. A combination of reducing the total amount of landers to 7,

and reducing the ΔV by 500 km/s, gave us a rather small but positive launch margin.

The propulsion system for the landing units also followed a very similar fashion as the orbiter. Due to the

extremely cold temperatures on Europa, the propulsion system was designed to be powered by a similar liquid

bipropellant. In this case, NTO is once again being used but combined with Monomethylhydrazine (MMH). The

structural design constraints on the landing units had a significant impact on the propellant selection, which was due

to the engine chosen for the landers.

Figure 12.4 R-1E Engine

The R-1E by Aerojet Rocketdyne had the best trade between nozzle length, Isp, and thrust. Moreover, these

engines have heritage from the Shuttle program, so they have been proven to work in the past. Aside from being

lightweight, it is capable of having 330,000 pulses available to vary thrust. This is particularly important since it is

very complex to throttle engines, short pulses may be used in order to maintain the ideal thrust level. In order to

ensure that the propulsion system could deliver the pulses necessary, it’s design was composed of the following

components: pressure transducers, pyrotechnic valves, system filters, solenoid valves, flowmeters, and latch valves.

Figure 12.5 Lander Propulsion System Overview

The R-1E engine has a Isp of 280 seconds and a very light mass of 2 kg each. At 110 N each, the design of

the propulsion system requires that 4 engines be used per lander to meet the proposed landing scheme. The engines

were strategically placed along the center of gravity of each lander. The benefit of having the engines placed in the

center is to prepare for the chance of any single engine failing. If any one engine fails at the center, ACS can be used

in order to compensate for that misalignment of thrust. If the engines were placed at the corners and any one failed,

the landing unit would tumble and be uncontrollable.

The design of the propulsion system for the orbiter and lander are very similar. There are 2 tanks for MMH

and NTO in order to have a more centralized center of mass. Pairing a tank of MMH and NTO to feed two engines

worked the best for this design. For every two engines, these tanks measure about 0.2 meters with a membrane

thickness of 5.95E-04 m. The weight of each tank was calculated to be about 0.82 kg.

Table 12.2 Lander Propulsion System Mass

The landing scheme requires for the propulsion system to deliver a ΔV of 1.48 km/s. This was calculated to be about

144 kg of propellant. This propellant mass accounts for an extra 10% of total ΔV, and loses due to startups.

XIII Attitude and Articulation Control Subsystem

The path of our spacecraft during its powered flight is directly influenced by its attitude and orientation in

space. Once outside the atmosphere, changing the direction of thrust by articulating exhaust nozzles or changing the

spacecraft's attitude influences its flight path. Our spacecraft's attitude will be stabilized and controlled so that its

high-gain antenna will be accurately pointed to Earth for communications, so that onboard experiments may

accomplish precise pointing for accurate collection and subsequent interpretation of data, as well as heating and

cooling effect of sunlight and shadow may be used intelligently for thermal control.

Figure 13.1: Dual Spin Orbiter

The mission to deploy multiple landers on the surface of Europa is a tall order, let alone a successful

mission alone to just navigate Europa. Requiring over 4 years of interplanetary travel, a Jupiter orbit insertion, a

Europa orbit insertion, and deploying 8 eight landers, will add up to a very large mass, and every kilogram really

counts. To avoid an additional 60 kilograms and 88 watts, the stabilization method we chose eliminates our need to

adapt reaction wheels on our orbiter. The method chosen to stabilize for this mission is not the normal 3-axis

stabilization, nor the spin stabilization technique, but the less frequent dual spin stabilization. Spin stabilization was

an option, but when the need for constant communication with the landers while in orbit about Europa, we needed to

adapt a despun section. This called for the use of a dual spin stabilization for the simplicity of a gyroscopic

stabilization, which allowed for a minor axis of inertia to be our spin axis. The orbiter's despun section, shown in

Figure 1, contains the electronics box, high gain antenna (HGA), low gain antenna (LGA), as well as the orbiter

cameras used for mapping the surface during the first 30 days mapping phase. The Bearing and Power Transfer

Assembly (BAPTA) is the original mechanism chosen for the dual spin mechanism, however, the max weight

allowed was much less than the weight of our spacecraft. This led to the switch to the spin mechanism assembly

used on the Global Measurement Instrument. This mechanism allows for simple damping between the spun and

despun sections, and for major burns releases the damping mechanism to permit both sections to spin freely

together. For these major burns, the spacecraft will require to be spinning at a rate no less than 6 rpm, due to the

centrifugal tanks used to accommodate the spin of the spacecraft.

For attitude control, there needs to be a reference on which was is 'up'. Many different devices may be

chosen to provide attitude reference by observing celestial bodies, or using inertia as a reference. The orbiter in our

proposed designed utilizes a total of five attitude references. This consists of three celestial references as well as two

inertial references. The three celestial references consist of two star trackers and a sun sensor. The star trackers used

are the CT-602 Star Tracker manufactured by Ball Aerospace. The sun sensor is specifically used for spinning

spacecraft, and that is the Adcole Spinning Sun Sensor. The star tracker uses an automated recognition of observed

objects based on built-in star catalogs. The sun sensor also if needed could be used for yaw and pitch reference.

Most star trackers use its roll reference with Canopus, a bright star. For this too work, our star trackers are placed on

the non spinning section of the orbiter. The sun sensor, considering it is used mainly for spinning spacecraft, will be

placed on the spinning section of the orbiter. The inertial references are the same instrument, just coupled for

redundancy. Added to the orbiter will be two LN-200 Core IMUs manufactured by Northrop Grumman Corporation.

Attitude control is obtained by sensors first most, but these communicate with the actuators, which in our case are

thrusters or our reaction control system (RCS). The actuators chosen for our design are the MR-106E 22N thrusters.

The orbiter is utilizing the configuration found on the Juno Spacecraft. Juno is a spinning spacecraft, and since our

thrusters are placed on the spinning section this became our design as well. The thrusters are configured with two

Figure 13.3: Adcole Sun Sensor

Figure 13.2: CT-602 Star Tracker

tangential thrusters, and one radial. These rocket engine modules allow the orbiter to obtain its full 6DOF. These

thrusters allow for variable thrust ranging between 30.7 - 11.6 N of thrust. They draw low power which assists in our

power management on the orbiter, and also has a single fire max burn of 2,000 seconds, with a lifetime of 4,670

seconds. Through the lifetime of the orbiter, the propellant used accumulates to less than 300 kilograms (7 years).

The center of gravity of a spin stabilized space craft is desired to be along the spin axis. In our case, since

we are dual spin, we are able to stabilized along a minor axis, which assisted in the spacing of the spacecraft. The

landers of the mission were attached on the sides of the spinning section, and evenly space across the conveniently

shaped octagon body. The center of gravity was nearly perfectly aligned along the spin axis, and only shifted along

that axis throughout the mission. The center of gravity shifted during the entire mission due to attitude correction

fuel, but the large shifts occur during the Jupiter insertion burn, Europa insertion burn and the deployment of the

landers. To compensate for the shift in the center of gravity during the lander deployment, an analysis had to be

done to guarantee the shift only along the spin axis. This would not have been the case if we deployed one lander at

a time. To account for this, the landers will be deployed two at a time, one on either end of the orbiter. The

determining sensors for when the landers will be deployed will be the star tracker coupled with the sun sensor. The

deployment will be executed using pyrotechnic bolts, due to the high failure of a mechanism. After four years of

interplanetary travel, a mechanism that hasn't been used or performed for more than four years is unreliable. The

pyrotechnic bolts give the landers a delta V of .3 m/s, which is combined with the tangential velocity obtained from

the spin, and the total delta V accumulates to just over .5 m/s.

The attitude and articulation control system was a task for the orbiter, but the landers have a significantly

higher demand for precision. The landers have to overcome unpredictable terrain, and with minimal information

Figure 13.4: LN-200 Core IMU

about the predetermined landing sites. This task requires a new generation of planetary landers and especially a new

generation of attitude control systems. The landers must have the ability to recognize their intended landing point,

detect hazards, and adjust/divert on its descent. Having done research on multiple avoidance systems, none quite

were as impressive as the ALHAT system. The ALHAT system, Autonomous Landing Hazard Avoidance

Technology, was created and developed by NASA Langley, Kennedy Space Center and Johnson Space Center. They

implemented it on Project Morpheus to demonstrate its ability to land autonomously while avoiding and assessing

the safest landing site. ALHAT is a system of three LIDARs (Light detecting and ranging), that assesses a terrain,

and uses algorithms to avoid hazards such as boulders and crevices. The three LIDARs used are a 3D Flash LIDAR,

a Doppler LIDAR, and a High Altitude Laser Altimeter. LIDARs work in any lighting condition, allowing for a

consist system through the 8 landers. Some will be landing at 'night' and some will be landing during 'day'. The 3D

Flash LIDAR is used as a camera capturing images and producing maps of the terrain, the Doppler LIDAR measures

the velocity and altitude, and the Laser Altimeter measures the altitude as well, but it mainly used for located the

intended landing point as well as the final 20 meter decent. The ALHAT system is able to detect hazards larger than

30 centimeters tall, and it can detect slops greater the 5 degrees, but will be used to divert slopes only greater than 15

degrees. As well as the ALHAT sensors, the landers will all be equipped with two LN-200 Core IMUs, shown above

in Figure 4. These are the same IMUs aboard the orbiter, but these will assist in the orientation and trajectory

calculations of the landers.

The landers actuators were carefully chosen due to the time of operation and the thrust necessary. Due to

the knowledge that the landers will only be operating for a few minutes, there will be no reaction wheel assembly

onboard. Adding reaction wheels to each of the landers would add 54 kilograms, and 88 watts per lander. All power

Figure 13.5: Lander REM

provided to the landers ACS is from a battery which is only used during the descent. With no reaction wheels, there

is no redundancy in the system with actuators. However, the thrusters chosen are very reliable. The selected

thrusters are the MR-111C 4N thrusters, shown in Figure 5. These have a variable thrust of 1.3 - 5.3 N. They

consume low power, and weigh only .33 kilograms each, taking the total weight for the actuators up to just under 4

kilograms. The thruster orientation is an original design that consists of two rocket engine modules (REMs) that

consist of 5 thrusters each as seen in Figure 4, and two individual thrusters on top of the lander to help obtain the full

6DOF.

The ALHAT has 3 phases, all with individual tasks that shall happen before the next. The beginning phase is Terrain

Relative Navigation (TRN), where the spacecraft is localized relative to the intended landing point (ILP). After

being localized, the guidance, navigation, and control algorithms determine the thrust or reaction control correction

needed to eliminate position error. Since TRN only uses reconnaissance images gathered during the mapping phase

of the orbiter, a higher resolution map needs to be generated to determine a safe landing site. This begins phase 2,

Hazard Detection and Avoidance (HDA). During the HDA phase, ALHAT creates a Digital Elevation Map (DEM)

to determine new safe sites. The hazard detection algorithms are used to assess the DEM and determine the new safe

sites. Although we had an original ILP, it may or may not be in an area designated as a new safe site by the DEM.

The new safe sites are then ranked by site safety and required fuel, before a divert maneuver is determined. The new

safe site is chosen and the maneuver takes place. The third phase is Hazard Relative Navigation (HRN). This phase

is a redundant phase to HDA phase, and produces higher resolution maps that are used for comparison to ensure the

safety of the new safe site. At 20 meters above the ground, the hazard detection ends, and the final descent is done

solely by the laser altimeter. The ALHAT system has said to "... open the gate for ALHAT technologies to be used

on the next set of lander missions". This was said by Kevin Kempton, a head engineer at Langley research center.

Instrument

Altimeter Flash LIDAR Doppler LIDAR IMU

Amount 2 2 1 2

Redundant? YES YES NO YES

Action in case of FAILURE

Use Doppler LIDAR to calculate distance traveled from orbiter to calculate altitude

Hazard Avoidance will be surpassed. Lander

will land without hazard avoidance

IMUs will be sensor for

trajectory and orientation

Doppler LIDAR will be the primary sensor

for trajectory and orientation

The ALHAT sequence can be seen in Figure 6. This figure shows the 3 phases, the sensors used, and the

main points along the trajectory that are critical for ALHAT. As the lander is above 10 kilometers, it is only using

the IMUs for its orientation, and the laser altimeter for localizing itself with the ILP. This occurs during the TRN

phase discussed above. The doppler LIDAR turns on just at approximately 2 km, and begins measuring the velocity,

and the altitude above the ILP. Finally the 3D flash LIDAR turns on at 500 meters, and begins the hazard detection

phase, HDA as well as HRN. Although the landers are only used for just a few minutes, it still requires redundancy

in the sensors to assist in the landing. With the low weight and low power consumption of most of the instruments,

the design is capable of utilizing two IMUs, two 3D flash LIDARs and two high altitude laser altimeters. However,

the doppler LIDAR draws an abundance of power, leaving only one aboard the lander. With the addition of these

redundant instruments, the landers are fully redundant. For a failed lander, 3 sensors must fail. If two of any of the

sensors fail, there is a redundant factor that will allow the lander to still land, however the hazard detection phase

will be surpassed. Table 1 shows the redundancy and the procedures in the event of failures.

Once the descent and landing are complete, there needs to be a confirmation between the legs, and the

system. To avoid any incorrect signals sent to the computer issuing a touch-down, we will have multiple checks

before a complete touch-down is acquired. The first will be a signal from the accelerometer in the legs. These will

be able to detect a 3g load, which is normally used for relatively solid surfaces. Europa is predicted to have ice

formations throughout the moon, so landing on a solid surface is a high probability. An alternative signal uses a

Table 13.1: ALHAT Redundancy

damping mechanism. The damping mechanism takes kinetic energy, and transforms it to electrical energy which

then sends a voltage and a current to trigger a touch-down signal. The final alternative utilizes the damping

mechanism to detect soft landings by evaluating small movements with the on-board computer. This system is

redundant and avoids the possibility of a false reading declaring a touchdown.

XIV. Manufacturing

In order to reduce manufacturing time of the spacecraft, all components that are available from other suppliers will

be bought. Only components specific to the spacecraft, such as propellant tanks and the bus structure will be made.

Modified versions of components based on available products would be purchased from suppliers as well. The list of

bought components for a each subsystem can be seen in table 14.1.The components that will be made are listed in

table 14.2. Finally the percentage of each subsystem that would be bought or made and the total for the spacecraft

can be seen in table 14.3. The percentages are based on mass of each component.

Table 14.1: Spacecraft Buy List

Item Model Supplier Orbiter Quantity Lander Quantity

Attitude Control System

IMU LN-200 Northop Grumman 2 16

MR-106E 12 0 ACS Thruster

MR-111C Aerojet Rocketdyne

0 96

Star Tracker CT-602 Ball Aerospace 2 0

Radar Altimeter HG-8500 Honeywell 0 1

Motor - General Dynamics 1 0

SMA - Ball Aerospace 1 0

Command and Data Handling

Compact PCI Rad750 3U BAE Systems 1 8

Processor Rad 750 BAE Systems 1 8

SRR - Airbus 1 0

Memory Module S990 3U Aitech 0 8

Payload

Micro-Camera - Micro-Cameras &

Science Exploration

0 16

VBB Sensor 0 24

SP Sensor 0 24 Seismometer

E-Box

CNES

0 8

Power System

VES 180 cells 1 0 Battery

VL 9E cells Saft

0 16

PCD - Terma 1 0

PDU - Terma 1 0

SmallSat Power - Clyde Space 0 8

Rigid UTJ Spectrolab 4 0 Solar Array

Ultraflex Orbital ATK 0 16

Propulsion

TR-308 Northrop Grumman 2 0 Bi-prop Engine

R-1E Aerojet 0 32

Pyrotechnic 6 32

Solenoid 8 32 Valve

Latch

TBD

2 64

Pressure Transducer - TBD 8 48

System Filter - TBD 8 32

Flow Meter - TBD 4 16

Telecom

TWTA X-Band L-3 Communication 1 0

1 0 Switching Network X-Band General Dynamics

2 0

Ultrastable Oscillator - General Dynamics 2 0

S-Band 0 1 Transponder

X-Band General Dynamics

2 0

S-Band 0 2 Diplexer

X-Band General Dynamics

0 2

Low Gain 2 4 Antenna

High Gain General Dynamics

1 0

Thermal

RHU - DoE 8 24

White 7.52 kg Insolating Paint

Optical Black AZ Technology

7.52 kg

MLI - TBD 49.9 kg 29.76 kg

OSR TBD TBD TBD 0

Mechanisms

Explosive Bolt SB4400 Hi-Shear Tech 32 0

Antenna 1 0 Pointing Mech

Solar Array TBD

4 16

Table 14.2: Spacecraft Make List

Item Orbiter Quantity

Lander Quantity Notes

Spacecraft Bus 1 8 Structures made of aluminum and titanium

Propellant Tank 2 32 Sized for appropriate amount of propellant

Pressurant Tank 1 16 Sized for appropriate amount of pressurant

Resistance Heaters 0 8 Composites with Resistors

Propellant Lines As needed Used for main propulsion and ACS system

Cabling As needed Power and telemetry lines

Shielding As needed Protect electronics from radiation

Boom 0 1 Extendable boom for camera

Table 14.3: Make/Buy Percentage

Subsystem Make Buy

Thermal 4.4% 95.6%

Attitude Control 10.54% 89.46%

Power 12.70% 87.30%

Command and Data 6.56% 93.44%

Telecom 9.51% 90.49%

Propulsion 81.43% 18.57%

Structure 100% 0%

Payload 19.32% 80.68%

Total 23.44% 76.56%

XV End of Mission

For the end phase of the mission, the landers will remain on the surface of Europa while the orbiter

eventually falls to the surface from orbit. When the orbiter’s propellant tanks become empty and therefore become

incapable of maintaining its orbit, the orbit will naturally decay via periapsis decrease. The orbiter will then crash

into the surface of Europa between 80 to 200 days after its required 90 days of science data gathering. This was the

most efficient way to dispose of the spacecraft because of numerous factors and design constraints. The first reason

used to justify this method was that it requires no input or commands from Earth in order for disposal. This was

done so that if radiation degraded the spacecraft electronics beyond usability, the spacecraft would still be able to

dispose of itself. Secondly, End of Mission scenarios involving disposal on Jupiter or other Jovian satellites are too

costly. The ΔV required to escape the orbit around Europa was calculated to be approximately 650 m/s. This alone

adds about 1500 kg of mass to the entire spacecraft and is approximately 50% of the launch margin. It was

concluded that any other burns would render the spacecraft unable to launch and unable to meet the RFP’s

requirements of arrival time by 2026. This mission to Europa is classified as a Category IV mission, which means

that it is interested in the evolution and/or the origin of life. Therefore, planetary contamination is one of the biggest

risks and concerns for such a mission. NASA requirements for such missions state that planetary contamination

must be kept under 10-4. To address these concerns, the spacecraft payload and components will be assembled and

maintained in Class 100,000 clean rooms. The spacecraft will also be routinely sterilized of contaminants throughout

the manufacturing process and prior to launch. Viking sterilization methods, such as dry heat microbial reduction

(DHMR) and vapor phase hydrogen peroxide (VHP), will be used as well any other methods that planetary

protection office deems necessary. These techniques are supplement by the large doses of radiation the spacecraft

receives during its trajectory and orbit around Europa. Finally, it was found that in numerous technical papers done

by NASA and others that this end of mission design is the most effective way of disposal for any mission to Europa.

In addition, these papers address the various planetary protection concerns raised by this type of mission. The

National Academy Press Report Preventing the Forward Contamination of Europa (2000) states the following:

Another conclusion reached from this particular sample analysis is that to meet the requirement

that Pc ≤ 10-4 , it will be necessary to do at least one of the following:

Demonstrate that no Type C or Type D organisms are on the spacecraft; or

Demonstrate that the probability of impacting the surface is less than 10-4 for the entire time the

spacecraft is in the vicinity of Europa (regardless of whether the spacecraft is operational or not);

or

Show by probabilistic calculations that the 10-4 standard can be met through a combination of

spacecraft cleaning, selective and/or whole-spacecraft sterilization, and exposure of the spacecraft

to the radiation environment at Europa for a long enough period of time to reduce the bioload to

the required level (“near sterilization”)[16.1]

The key message from this quote is that only one of the techniques will need to be used in order to prove planetary

protection requirements are met. Another quote from a more recent NASA report from 2012 supplements these

justifications.The report entitled Assessment of Planetary Protection and Contamination Control Technologies for

Future Planetary Science Missions states the following:

Forward contamination would exploit ambient radiation of Jovian system to provide further bio-

burden reduction with assumed system-level sterilization.

Europa orbiter or lander planned contact with Europa; must demonstrate sufficiently low

probability of contamination of subsurface ocean.

JUICE or Europa multiple fly-by planned impact with Ganymede; must also demonstrate

sufficiently low probability of impact on Europa and/or contamination of subsurface Europa or

Ganymede ocean.[16.2]

These quotes prove that the chosen end of mission design concept will be compliant with NASA’s Planetary

Protection Agency’s requirements.

XVI Preliminary Cost Analysis

Two Cost Estimation methods were chosen to estimate the preliminary cost of the spacecraft: the

Unmanned Space Vehicle Cost Model, Eighth Edition (USCM8) and Project Cost Estimating Capability (PCEC)

which is an Excel Spreadsheet based off Cost Estimating Relationships (CER) of the NASA Air Force Cost Model

(NAFCOM). All cost estimations were made for financial year 2020, which is the launch year of the spacecraft, and

assuming one Orbiter and eight landers. In Table 16.1 the USCM8 excel sheet is shown.

Table 16.1: USCM8 model for spacecraft design

The UCM8 Model is a simple cost estimation table because it does not account for many of the program

management, system engineering, and other design and production costs. Since it only estimates the design of one

flight model and test model, other elements of the design had to be accounted for using separate estimations. Most

additional estimations, such as Launch Vehicle cost and Mission Operations, were taken from the ARO 481L notes.

Adding these estimations together, the final USCM8 estimate comes to approximately 3 billion dollars. Table 16.2

shows the summary table of the PCEC cost estimation.

Table 16.2: PCEC model for spacecraft design using Uncrewed Spacecraft Template

The PCEC cost model accounts for multiple phases of the design and development phases as well as many

of the program management and system engineering elements. In addition, the particular PCEC template used for

the estimation includes the cost and development of the landers as well. From PCEC, the resulting total cost

estimation for the spacecraft design is approximately 4.5 Billion dollars. Figure 16.1 shows a comparison between

the two cost model estimations as well other planetary missions.

Figure 16.1: Cost Comparison

The PCEC estimation is largest of the cost estimates, however considering the scope and scale of this

design, it is the most reasonable estimate. While Europa Clipper is less than half of the discussed design, this is

reasonable considering that Clipper only has one lander whereas this design has eight. In addition, the PCEC

estimate is the most reasonable because it is the most comprehensive, self-sufficient, and uses historical data from

previous spacecrafts when calculating its estimates.

XVII Conclusion

The design chosen to satisfy the requirements set forth by NASA JPL’s RFP is one composed of one

orbiter and eight landers. The spacecraft uses a high ΔV, low flight time VEGA trajectory as well as a Jovian Moon

Tour in order to reach Europa by 2026. In order to achieve the necessary C3 value, the SpaceX Falcon Heavy

Launch Vehicle was chosen due to the fact that it is the only launch vehicle with a suitable launch mass to

accomplish this mission. An alternate VEEGA trajectory is possible with a Delta IV Heavy Launch Vehicle,

however this comes at the cost of arriving in 2027 and having only seven landers. Once at Europa, the spacecraft

performs pre-landing mapping of Europa for one month prior to deploying the landers and gathering science data. In

order to perform the landing sequence, the landers are designed to be autonomous and use their attitude control

systems for decent. Spacecraft communications and data transfer will be performed by using high and low gain

antennas operating over the X and S bands. These bands were chosen because they are not as sensitive to weather as

bands of higher frequency and they both have high data capacities. the Both the orbiter and landers are solar

powered due to lack of feasibility of using 12 MMRTGs. The orbiter is in a full sun orbit in order to reduce possible

eclipse times while the landers use their rotatable solar arrays to charge their batteries for nighttime use. Both the

orbiter and landers use liquid biprop engines as they are much more reliable for long flight time missions. End of

mission will disposal will compose of the landers staying on the surface and the orbiter naturally decaying and

crashing into Europa. Planetary protection concerns are addressed via sterilization and enhanced Viking cleaning

techniques at Earth as well as the high doses of radiation received throughout the mission. Through examination of

all requirements, the proposed design is compliant with all restraints and requirements and is fully capable of

completing the RFP’s mission.

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