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
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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
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.
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.
References
Works Cited 4.1 “Characterizing electron bombardment of Europa’s surface by location and depth” by Patterson, Paranicas, and Procker (www.elsevier.com/locate/icarus 8.1 Panning, M., Lekic, V., Manga, M., Cammarano, F., & Romanowicz, B. (2006, December 12). Long Period Seismology on Europa: 2. Predicted seismic response. Journal of Geophysical Research, 111. 8.2 Lognoné, P., Banerdt, W. P., Weber, R. C., & al., e. (2015). Science Goals of SEIS, The Insight Seismometer Package. 46th Lunar and Planetary Science Conference. 8.3 SEIS Seismometer. (n.d.). Retrieved May 31, 2015, from SRON Netherelands Institute for Space Research: http://www.sron.nl/seis-seismometer-shamroc-1967 7.4 Pike, W. T., Standley, I. M., & Calcutt, S. (2013, June). A Silicon Microseismometer for Mars. IEEE.
8.5 Firgelli Technologies Inc. (2014). Miniature Linear Motion Series - L16. Retrieved from http://www.firgelli.com/pdf/L16_datasheet.pdf 8.6 MOOG. (n.d.). Schaeffer Magnetics Division. Retrieved from www.moog.com/literature/Space_Defense/Spacecraft/Spacecraft_Mechanisms_Product_Catalog2.pdf 7.7 New Scale Technologies. (2015, February 4). M3-RS Rotary Stage. Retrieved from http://www.newscaletech.com/doc_downloads/M3-RS-datasheet.pdf
8.8 Insight. (2013, April). Insight Newsletter. Retrieved from http://solarsystem.nasa.gov/insight/docs/ISGH-SEIS-NL-IPGP-19255_InSight%20Newsletter%20April%202013_v1.pdf
8.9 Robert, O., & al., e. (2012). The Insight Very Broad Band (VBB) Seismometer Payload. 43rd Lunar and Planetary Science Conference. Retrieved from http://www.lpi.usra.edu/meetings/lpsc2012/pdf/2025.pdf
8.10 National Aeronautic and Space Administration. (2013, July 13). Radioisotope Power Systems. Retrieved from https://solarsystem.nasa.gov/rps/rhu.cfm
16.1 Esposito, Larry W., et al. Preventing the Forward Contamination of Europa. Publication. National Academy Press, 2000.
16.2 Belz, Andrea, et al. Assessment of Planetary Protection and Contamination Control Technologies for Future Planetary Science Missions. Rep. no. JPL D-72356. N.p.: NASA JPL, 2012.
Other Sources Used: • Beagle Stereo Camera System
– Griffiths, A.D., et. al. 2005. The Beagle 2 stereo camera system. Planetary and Space Science 53. (1466-1482)
– Micro Cameras & Space Exploration. http://www.microcameras.ch/site/index.php?id=42
• Pancam Camera
– Schwochert, M.A. and J. N. Maki, 2005. The Mars Exploration Rover Cameras: A Status Report. Lunar and Planetary Science http://www.lpi.usra.edu/meetings/lpsc2005/pdf/1793.pdf
• ECAM-C50
– Malin Space Science Systems. 2013. ECAM-C50. http://www.msss.com/brochures/c50.pdf
– Malin Space Science Systems. 2013. ECAM Imaging System. http://www.msss.com/brochures/ecam.pdf
• Microscopic Imager (MI)
– NASA. 2012. Europa 2012 Study Report. https://solarsystem.nasa.gov/europa/2012study.cfm
• Site Imaging System (SIS)
– NASA. 2012. Europa 2012 Study Report. https://solarsystem.nasa.gov/europa/2012study.cfm
– Reyes, Tim., 2014. Rosetta’s Philae Lander: A Swiss Army Knife of Scientific Instruments. http://www.universetoday.com/114471/rosettas-philae-lander-a-swiss-army-knife-of-scientific-instruments/
• Multiband Seismometer (MBS)
– NASA. 2012. Europa 2012 Study Report. https://solarsystem.nasa.gov/europa/2012study.cfm
• SEIS VBB Seismometer
– Insight, 2012. The SEIS Insight VBB Experiment. http://www.lpi.usra.edu/meetings/lpsc2013/eposter/2006.pdf
– Dandonneau, P-A., et-al., 2013. The SEIS Insight VBB Experiment. 44th Lunar and Planetary Science Conference. http://www.lpi.usra.edu/meetings/lpsc2013/pdf/2006.pdf
– Mimoun, D., et-al., 2012. The Insight SEIS Experiment. 43rd Lunar and Planetary Science Conference. http://www.lpi.usra.edu/meetings/lpsc2012/pdf/1493.pdf
– Robert, O., et-al., 2012. The Insight SEIS Experiment. 43rd Lunar and Planetary Science Conference. http://www.lpi.usra.edu/meetings/lpsc2012/pdf/2025.pdf
• Colibrys SF3000L
– Merchant, B. John., 2009. MEMS Application in Seismology. Sandia National Laboratories. http://www.iris.edu/hq/instrumentation_meeting/files/pdfs/MEMS_Seismology.pdf
• Telecomm
– A Comparison of the Ka-Band Deep-Space Link with the X-Band Link through Emulation by Shervin Shambayati (http://ipnpr.jpl.nasa.gov/progress_report/42-178/178A.pdf)