Europa Cadmus
A Scout Mission Proposal April 21, 2015
0
A Executive Summary Named after the Cadmus of Greek mythology who went in search of his sister Europa, this proposed mission will journey to Europa in search of truth regarding the habitability of one of the most probable locations of extraterrestrial life in the solar system. Science Rationale The Cadmus mission will take the next step to complement and extend the information gained by the Jupiter Icy Moons Orbiter (JIMO) and Galileo in establishing a knowledge base about the makeup, dynamics, and habitability of Europa. Cadmus will survey the dynamics of the icy crust of Europa, assess the presence of potential energy and nutrient sources for life, and assess the Europa environments suitability for life in order to ultimately determine the habitability of the moon. Galileo has collected and JIMO will collect invaluable information regarding Europa, which is agreed to be one of the most probable locations for the existence of life in our solar system. Both of these missions, however, are inherently limited in that they are both orbiters and are thus forced to observe Europa from a distance. Cadmus will take the next step by landing on the icy satellite to take direct, insitu measurements. In establishing the habitability and the most probable energy sources and environments for Europas potential inhabitants, Cadmus will lay the groundwork for future missions returning to Europa to search directly for life. Cadmus will achieve its science objectives using imagery, spectroscopy, seismology, and environmental readings. Through these tools, Cadmus will allow us to see what the surface of Europa looks like, determine the elemental
and molecular constituents of the icy crust, monitor seismic events and tidal motions from deep within the moon, and measure the environmental conditions on the surface. This information will allow scientists to determine the potential habitability of Europa. Mission Architecture The Cadmus mission architecture is low risk with a high science return. Cadmus consists of two identical, independent landers (Figure A.1) and a propulsion module launched from Kennedy Space Center (KSC) on a single launch vehicle in late 2021. The propulsion module has a propulsion system and navigational instruments that rely on the power, computers, and communications of the landers. The propulsion module transports the landers to the Jupiter system using its bipropellant chemical propulsion system to achieve orbit around Jupiter. This cruise stage is launched on a path to Venus in order to complete a Venus-Mars-Venus-Earth flyby maneuver and is scheduled to arrive in the Jupiter system in June of 2025. Upon arrival at Jupiter, the spacecraft performs multiple maneuvers until it delivers the landers to a location near Europa. At this point, the landers separate from the propulsion module and separately insert into circular orbits around Europa. They each perform a propulsive descent and landing to arrive at their respective locations of science exploration. The specific landing sites will be determined from data returned from JIMO, but are expected to be relatively flat places of recent upwelling from below the crust. Having two landers in different locations on Europa will allow for complete surface redundancy, provide a source for comparison to determine Europas uniformity, and significantly improve the chance of returning the science data. Once on the surface, the landers will perform their experiments and
1
will complete all measurements with the exception of the seismology portion of the mission within twelve days of landing. The seismology data will then be collected and returned to Earth over the remaining life of the landers. The total mission is expected to end in September 2026, roughly five years after launch. The total surface duration will at minimum be four weeks at minimum, six weeks at maximum. Technology Development Plan Because much of its hardware has already flown on past missions, Cadmus requires a relatively small amount of technology development. The most rigorous developments that must be made involve readying the landers and instruments for the harsh radiation environment of Europa. Advancements made in radiation shielding and radiation hardening for the JIMO mission will be utilized to prepare instruments and hardware to survive on Europa. Still, it is expected that the lifespan of the landers once on the surface will be approximately four to six weeks due to the radiation. Extensive radiation testing will be performed on all non-proven components to ensure that the mission goals can be met. Cadmus is intended to develop and use hybrid, inflatable high gain antennas for lander communications. Although the technology has not yet been flight qualified, such a demonstration of advanced technology will open up the technology to use on future missions. Furthermore, the technology may be transferred to the private sector for use in industry. The proposed antenna will inflate, increasing disk diameter and thus increasing data rates. Most importantly, the failure of the antenna to deploy would not compromise the mission. The baseline mission can be
completed with the use of the standard, uninflated portion of the antenna. Management With significant past experience in managing advanced planetary science missions, Jet Propulsion Laboratory (JPL) will manage the Cadmus mission. The management team will report directly to Wes Patterson of Brown University, the Principal Investigator (PI) for the mission. The project management team will have direct control over the project systems engineering group, the flight systems engineering group, the payload systems engineering group, and the mission systems engineering group. Project Management will ensure that there is a high level of mission assurance and reliability, parts reliability, analogous environmental testing, adequate testbed construction, full integration of the payload into the flight vehicle, efficient mission operations and ground system management, and sufficient planetary protection measures. The PI will also be responsible for all descopes, keeping the cost on budget, and maintaining the project schedule. The project schedule has been designed to allow sufficient planning prior to Critical Design Review (CDR) for project definition and technology maturation, as well as significant time for spacecraft and instrument construction and integration. Should the project be unable to complete the science floor, the PI is prepared to recommend project cancellation to the National Space Administration (NSA) Office of Space Science (OSS).
2
Cost The Cadmus mission cost is 50 million less than the cost cap while providing twice the reliability and science data return of a single-lander architecture. Cadmus cost estimation has been performed using advanced models and by analogy to similar past missions. A conservative estimate of development costs for advanced instruments has been assumed to ensure sufficient resources for technology maturation within the project schedule timeline. The advanced missions operations utilized by the Cadmus mission in the form of beacon tone monitoring and the use of existing ground systems has afforded some cost savings in the mission. The Cadmus mission has numerous descope options in case of reduced availability of funds. Additionally, the project can accommodate significant future unexpected growth within its 30% cost margin. The total cost to NSA is 950 million in Fiscal Year (FY) 2016 dollars, 50 million less than the cost cap. Group Roles The Cadmus team project manager was Bob Thompson, who was also the disciplinary
specialist for cost estimation. Scott Francis served as Project System Engineer, and was in charge of trajectories and mission design. Randy Olsen was Flight Systems Engineer, and was in charge of structures and mass estimation. Michael Parsons was Payload Systems Engineer, and performed all CAD modeling and visualization. Robbie Coffman served as Mission Systems Engineer, and was in charge of Power and Propulsion systems.
Figure A.1. A Cadmus lander on the surface of Europa performing surface science to determine habitability.
Figure A.2. The Cadmus mission is a critical step towared future Europa exploration.
3
B Table of Contents C. Fact Sheet5 D. Science Investigation..........6 1. Science Goals and Objectives6 a. Goals and Relation to Past and Future Exploration...........6 b. Measurements Required.7 c. Baseline and Performance Floor....8 2. Implementation..9 a. Landing Site and Sampling....9 b. Data Analysis and Archiving...10 E. Education/Public Outreach, Technology Transfer, Small Disadvantaged Businesses..11 F. Mission Implementation12 1. Mission Design..12 2. Flight System.14 a. Cruise Stage.........14 b. Lander..........17 3. Launch Vehicle..20 4. Ground Systems.23 5. Communications Approach...23 6. Mission Operations...24 7. Technology Development.25 8. Instrumentation.26 G. Management and Schedule...29 H. Cost and Cost Estimating Methodology...30 Appendices A. Planetary Protection..........32 B. List of Acronyms and Abbreviations33 C. References List..35
4
Cadmus Fact Sheet
Science Objectives Cadmus complements and extends the science of JIMO and Galileo; it is also an important step towards future Europa missions. Science Theme: Habitability. Survey the dynamics of the icy crust of Europa. Assess the presence of potential energy and nutrient sources for life on Europa. Assess the suitability for life in the Europa environment. Mission Overview The Cadmus mission platform offers the optimal opportunity to accomplish the science objectives with the minimum amount of risk. Launch Vehicle: Atlas V 551 Launch Date: Oct. 26, 2021 Lander Europa Orbit Insertion: July 2026 Launch Wet Mass: 3620.0 kg Growth Allowable: 30.4% Mission Design: Cruise stage consisting of propulsion module and two landers Venus-Mars-Venus-Earth trajectory Landers separate from propulsion module, enter Europa orbit, and land separately Communications w/ DSN via KA-band and X-bands Landers: Two identical surface landers Dry Mass: 248.2 kg each Wet Mass: 558.1 kg each Expected Science Duration: 30 days Science Payload The Cadmus mission utilizes a comprehensive suite of flight qualified and rigorously tested science instruments aimed at fulfilling all science objectives. Instruments being used include a Gas Chromatograph/Mass Spectrometer, descent imager, panoramic camera, ultrasonic corer, aqueous chemistry laboratory, near-infrared spectrometer, Raman spectrometer, 3- axis seismometer, and environmental sensors. Mission Management The Cadmus management plan and mission development schedule will ensure a project that accomplishes all mission goals and objectives while minimizing risk and staying on budget. Planetary Protection Cadmus will meet Category IVB planetary protection criteria. All efforts will be made to ensure a minimal number of organisms present on the spacecraft. Schedule and Cost Note: All costs M$FY16 w/ 30% reserves Phase Duration Start Date Cost A 8 months June 2016 1.25 B 8 months Feb. 2017 28.11 C 24 months Oct. 2017 463.95 D 15 months Nov. 2019 175.53 E 5 years Oct. 2021 80.63 Launch Vehicle 201 Education/Public Outreach and Technology Infusion/Transfer Objectives The Cadmus mission will provide funding for Educational Outreach programs for grades K-12, provide research opportunities to students at historically black colleges and universities, as well as patronize minority and/or women owned small businesses. Educational objectives will be achieved through, but not limited to, school curriculum packets, student contests, public websites, and creation of a Public Broadcasting System NOVA television special. Technology infusion and transfer will be supported.
5
D.1.a Goals and Relation to Past and Future Missions The Cadmus mission will extend the science of JIMO and Galileo as well as provide a critical link between those missions and future Europa missions. The science theme of the Cadmus mission is the study of the habitability of the Europa environment. As a scout mission, Cadmus' science investigation focuses very specifically on this theme. The Cadmus mission has been specifically designed to obtain as much information as possible about habitability, which has been identified by the JIMO Science Definition Team as the most important aspect of Europa for study by future missions1. The Cadmus mission establishes a good scientific basis for the planning of future scout and flagship class Europa science missions. As such, Cadmus is a vital step between remote sensing missions (JIMO, Galileo) and future surface and subsurface (e.g. Cryobot) life detection missions. Table D.1 shows the relation of Cadmus to past and future missions. Science Objective 1: Survey the dynamics of the icy crust of Europa. Cadmus will provide an in-situ platform for the direct measurement of Europa's tidal and seismic response, and will provide ground truthing and validation of the JIMO geological measurements. This objective relates to the study of the habitability of Europa by determining the extent to which there is an exchange of icy material between the surface and the subsurface. Because of the possibility of the existence of a liquid-water ocean more than 20 km beneath the surface of Europa, and the possibility of the presence of life in this liquid-water environment, the transport of material from the subsurface to the surface
must be understood before science measurements on the surface can be used to infer the habitability of the subsurface environment. JIMO is expected to provide geophysical measurements (gravity, magnetic field) of Europa in order to determine the interior characteristics of the moon. Additionally, radar altimetry will be used to measure the tidal shifting that occurs on a periodic basis. There is already a significant need for understanding the seismic response of Europa's icy shell in order to obtain an accurate model of the Europa system1. Science Objective 2: Assess the presence of potential energy and nutrient sources for life. Cadmus will look for signs of geothermal energy in the subsurface environment, as well as signs of compounds that can support known forms of life. Europa is an important moon because the possibility of liquid water makes it a prime candidate for potential life. In order for life to exist, however, sources of energy and nutrients must also be present. Because of Europa's distance from the sun and the extreme depth of the liquid ocean (the most likely place for life to evolve and flourish on Europa), sources of energy other than solar energy must be sought. Cadmus will provide in-situ composition analysis of the near-surface material greater than 10 cm deep to analyze samples mostly untouched by the radiation environment. JIMO and Galileo's remote sensing instrumentation was capable of performing spectroscopy of the surface. While this is useful information, it only represents the composition of the top 1 mm to 1 cm of the surface material, a highly irradiated environment.
6
Objective 3: Assess the suitability of the Europa environment to support life. Cadmus will study the physical properties of the surface material directly, obtaining observations not possible from orbital mission such as JIMO and Galileo.
Most importantly, Cadmus will prepare for future astrobiology missions that will search directly for signs of life in the Europa system by allowing scientists to define what types of life they expect to exist on Europa based on the Cadmus environmental assessment.
Table D.1 Exploration missions and future search-for-life missions including rovers, Cryobots, and sample return.
Mission Mission Type Galileo Flyby Gravimetric and radiometric data, low resolution imagery JIMO Orbiter Radar and laser mapping, gravimetric data, medium resolution mapping Point spectroscopy from orbit for determination of constituents of top 1mm to 1 cm of surface Cadmus Lander Local high resolution surface imagery (panoramic); seismographic data. Direct GCMS/Raman Spectrometer and aqueous chemistry lab analysis of samples up to 20 cm depth (samples unaltered by radiation) Future Missions Rover, Cryobot, Sample Return Multiple surface high resolution imagery, direct exploration of subsurface
1) Europa structural dynamics investigation
Science Investigations
2) Nutrient/Energy Resources in Surface Ice
Some spectroscopic imagery
Surface and ice column constituency analysis
3) Environmental conditions
None
Radiation sensors, remote sensing spectroscopy
Direct measurement of surface environmental conditions, and direct analysis of near surface ice.
Multiple site surveys, subsurface ocean exploration?
D.1.b Measurement Requirements The Cadmus mission measurement requirements will enable a more detailed understanding of Europa, its resources, and its potential for harboring life. The scientific rationale of the Cadmus mission is to take detailed measurements of the surface composition, as well as provide imagery and seismographic information that will provide for the context information necessary to determine the interchange of material between the subsurface and the surface. The Cadmus 7
mission provides the best architecture within the given constraints to assess the suitability of life to Europa. Imagery: Landing site context imagery will provide information about the geology of the site at which Cadmus has landed. This will allow the science team to determine if the surface material being analyzed has been part of a subsurface ocean. A 360-degree panoramic camera similar to those utilized on the Mars Exploration Rover (MER) mission will be used to provide the imagery. Additionally, a descent imager will allow the
science team to compare the landing site with JIMO and Galileo imagery in order to put it in a global context. Spectroscopy: Three types of spectroscopy measurements will be utilized to accomplish the mission science goals. Elemental analysis will be utilized to determine the constituents of the ice. Elements of interest will be those that are indicative of energy resources in the subsurface environment (e.g. sulfur as a sign of geothermal activity), as well as those elements that could provide nutrients to life, such as phosphates and nitrates. Additionally, molecular analyses will be performed. Specifically, potential byproducts of life will be searched for, such as complex carbon compounds. Finally, remote spectroscopy will be utilized to analyze the surface composition of the area immediately surrounding the lander. Seismology: The seismology measurements of Europa allow tides and quakes to be measured. It is theorized that these events occur regularly on a scale of days (the tidal period of Europa is 3.55 days). This information will provide ground truth to complement JIMO's radar altimetry experimental results. Environment: Measurements of the Europa environment will be taken in order to assess the suitability for life, as well as to assist the science team in understanding what forms of life may be present on Europa. Measurements of radiation, temperature, and solar flux will be taken. Additionally, the surface material will be analyzed to determine its pH and salinity. D.1.c Baseline Science and Performance Floor The Cadmus baseline mission surveys two sites and provides measurements of
Europa's crustal dynamics, surface constituency, and the suitability for life, and allows surface and subsurface scientific investigation. Baseline: The Cadmus baseline mission utilizes a two-lander architecture that allows more than one area to be surveyed. This architecture maximizes the chances for scientific return, both by providing redundancy in case of a failure, and in aiding the assessment of the homogeneity of the Europa surface. The baseline mission addresses the theme of habitability, and in turn focuses on the three objectives of crustal dynamics, surface constituency, and environmental suitability. Analysis of samples below the irradiated zone of ice is accomplished to obtain a pristine sample. The baseline mission calls for the analysis of three samples from each landing site in order to obtain a good sampling of the site and to eliminate possible outliers in the data set. The Cadmus architecture is ideal for obtaining the information necessary to evaluate the habitability of Europa. Performance Floor: The Cadmus performance floor maintains surface constituent analysis on a local scale. One landing site is investigated using one lander, with two samples from that site being analyzed in order to remove outliers from the dataset. Redundant systems on the lander ensure a low-risk and fully reliable system capable of allowing the habitability of Europa to be determined. Potential descope options to achieve the performance floor are shown in Table D.2. Without the seismology mission, the mission duration is reduced to 12 days total.
8
Table D.2. Descope options ensure the Cadmus mission stays under budget and on time.
Mass Savings Cost Savings 241 M$FY16 ~5 M$FY16 ~2 M$FY16 Decision Timeframe
Descope Option Reduce to one lander
1465 kg Eliminate seismology 4 kg Reduce # of samples per site 5 kg
Science Impact Homogeneity of Europa cannot be assessed, reduced survey of Europa Crustal ice dynamics cannot be measured except by imagery; no verification of JIMO Outliers in dataset cannot be accurately eliminated
Before CDR
Before PDR Before CDR
Table D.3. All instruments and data product requirements are directly traceable to scientific objectives.
Science Objective Survey the dynamics of the icy crust of Europa Mission Requirement Two landers Landing at one or more sites of recent upwelling In-situ sampling of crustal ice; drill for sample extraction Landing at one or more sites of recent upwelling In-situ sampling of crustal ice; drill for sample extraction Multiple samples per site Descope for Floor Baseline Instrument Requirement Seismometers Pancam Mass Spectrometer Raman Spectrometer Aqueous Chemistry Lab Point spectrometer Environmental Sensors Baseline Data Product Requirement (Per Lander)
> 2 weeks of measurements without other instruments/mechanisms operating = 1.2 Gbit
20 Images (1024x2048) =240 Mbit 3 samples to ~2 ppb resolution= 6 Mbit 3 Samples to ~2 ppb resolution= 6 Mbit 3 samples = 3 Mbit 30 images (512x512) = 120 Mbit Full readout every hour during mission ops = 120 Mbit
Identify potential energy and nutrient resources
Assess the suitability of the crustal ice to supporting life
D.2.a Landing Site and Sampling Cadmus mission design and sampling strategy offers the optimal opportunity to accomplish the science objectives with minimum risk.
The Cadmus landing sites will be chosen based on data returned from the JIMO mission. The sites chosen will be areas of recent upwelling and will thus offer a chance to study material that was below Europas icy crust not long ago. Little is known about Europas sub-crustal material and this data
9
could hold significant clues regarding Europas capability of sustaining life. The landing sites will also be chosen for their relatively smooth terrain which will significantly reduce the risk involved in landing. Taking sub-surface samples will provide information about Europan material that has not been bombarded with radiation. Since the majority of the radiation on Europa is absorbed by the top 10 cm of ice, studying material below 10 cm will allow scientists to learn about the areas that are the most habitable. Utilizing an arm on the landers will allow for interactive sample selection. When scientists receive images of the area surrounding the lander, the arm may be repositioned to allow samples to be taken in the most interesting spots. D.2.b Data Analysis and Archiving Cadmus science will be archived in the NSA Planetary Data System (PDS) within six months of the end of mission, after which it will be readily available to the science team, scientific community, and the general public without delay (Figure D.4). During the mission development phase, science team members will create algorithms and models that will be capable of analyzing the data that will be collected by the Cadmus mission. Once the landers begin their mission, their data will be collected by the DSN and stored simultaneously at the Missions Operations Center and at PDS. The science team will then begin to carefully draw results from the data using the aforementioned methods. All data will be based on standard units (SI) for more efficient communication.
Once the data has been verified and validated, it will be released to the general public. There will be no proprietary period for analysis by the science team. During Phase E, the science team will submit the results of their investigations to peerreviewed journals for publication.
Landers DSN Mission Ops Science Team PDS General Public Figure D.4: A quick and direct flow of data exists from the landers to the science team and the general public.
10
E Education and Public Outreach The Cadmus mission will provide for Educational Outreach programs for grades K-12, provide research opportunities to students at historically black colleges and universities, and support minority and women owned small businesses. K-12 Education: Materials designed to complement the science curriculum at each grade level (K-12) will be developed and distributed to schools across the country. At the grade school level, this will consist of worksheets, coloring books, paper model kits, pictures, and fact sheets. The students will be engaged to understand the science being accomplished by Cadmus, the rationale behind it, and the engineering challenges of the mission. Additionally, a website designed for children of varying ages will be posted for children to keep up to date on the status of the mission. At the high school level, several engineering and science themed contests will be held and the winners will be invited to be at JPL for the landing and duration of the Cadmus science mission. Additionally, materials and curricula will be developed for inclusion in high school level biology, chemistry, and geology classes, including information on the formation of Europa, the current state of knowledge about the structure of Europa, and what the science results could mean towards understanding the potential of life on Europa. The National Geographic Society will also co-sponsor a student essay-writing contest about the importance of exploration of the outer solar system. College Level Education: Several intern/research grant opportunities will be made available to students studying astrobiology and planetary science at historically black colleges and universities
(HBCUs). These grants will fund the students to work as interns to members of the mission science team, as well as the opportunity to perform research utilizing the scientific data returned by the Cadmus mission. Funding will be provided for travel to present peerreviewed articles at a scientific conference. Public Outreach: A website will be maintained with up to date pictures, press releases, and scientific results. Daily briefings will be released during the period of science mission operations, and coverage by all of the major networks will be encouraged. In addition, a partnership with the Public Broadcasting System / WGBH will be formed to produce a three-part "Nova" special about the planning for the mission, the design process, building the lander, and the surface science operations. NSA will partner with the National Science Foundation to produce layperson-targeted lectures at community centers across the country describing the importance of Europa exploration. Additionally, the Planetary Society will perform additional outreach. Technology Infusion/Transfer: The Cadmus mission will make the support of technology transfer a priority wherever applicable. Project management will examine the potential markets for technologies developed in the mission design process, and will work with the appropriate NSA institutions on patenting and commercialization. Minority and/or Women Owned Businesses: The Cadmus management team will make the support of small disadvantaged businesses a priority wherever possible and appropriate to meet the goals of the NSA OSS.
11
F. Mission Implementation F.1 Mission Design Cadmus will minimize risk and maximize science return per dollar by featuring flight proven technology. The Cadmus mission (Figure F.1-3) begins on October 26, 2021 with a launch from KSC aboard an Atlas V 551 launch vehicle. The launch vehicle delivers the spacecraft, composed of a propulsion module and two landers, to an interplanetary trajectory with an Earth escape energy (C3) of 16.4 km2/s2. The spacecraft then completes a Venus-MarsVenus-Earth flyby maneuver (Figure F.1-5) and arrives at the Jupiter system in July of 2025 (Figure F.1-1). As the spacecraft enters the Jupiter system, it performs a Ganymede Gravity Assist (GGA) and enters a highly elliptical orbit around Jupiter. The spacecraft then performs Perijove Raise Maneuvers (PJR), including Europa Gravity Assists (EGA), which slowly raise the periapsis of the orbit to be tangent with Europas orbit (Figure F.1-2). The Cadmus landing site selection team will analyze Europa surface data from the JIMO mission. This data will be used to augment the descent imager data in selecting suitable landing sites. While in the Jovian system, the landers separate from the propulsion module and independently perform Europa Orbit Insertion (EOI) maneuvers to arrive in 100 km circular, stable orbits around Europa. This was done for planetary protection reasons, as well as to reduce the overall launch mass. After EOI, the lander subsystems are checked to ensure proper functioning and the first lander initiates its descent to the Europan surface (Figure F.1-4). Data from the first landing is analyzed and knowledge gained is applied toward the second landing. This adaptive landing scenario helps to ensure a successful mission.
Each lander descends using chemical propulsion. So as not to land on ice that was in the path of the propulsion products, the main propulsion system is turned off at a point 3.4 meters above the surface. The lander at this point has zero vertical velocity and 5 m/s of horizontal velocity. The lander drifts horizontally as it falls, and the attitude control thrusters are used as translational thrusters to null the remaining horizontal velocity before touchdown. Each lander is equipped with an autonomous avoidance system which will use short bursts from the main thrusters to keep the lander afloat in the event that it is going to land on an obstacle. The lander has enough propellant for 60 seconds of hovering in this fashion (Figure F.1-6). It impacts the surface with a nominally vertical velocity of 3 m/s. To cushion the lander from this impact, the lander legs feature crushable aluminum honeycomb shock absorbers. The honeycomb is sized to endure impact velocities of 6 m/s. Once on the surface, the lander deploys the pancam and snaps two pictures, which are sent back to Earth along with all other diagnostics as a test of the landers health. Coring begins for sample extraction and analysis. The corer takes three samples, the locations for which are autonomous. This is done in the event that communications problems prevent ground instructions from being received by the spacecraft. Scientists will have the ability, however, to analyze the initial pancam pictures and send new locations for the second and third corings. Finally, when all three core samples have been completely analyzed, the seismometers are switched on. All data except for the seismology readings will be returned to Earth within 12 days of landing. The seismometers will continuously gather and return seismology data to Earth for the life of the lander, which is expected to be approximately four weeks and is limited by radiation.
12
Figure F.1: Cadmus mission implementation is an efficient method of studying Europa
100 90 80 70 Height (km) 60 50 40 30 20
Landing Trajectory
+ Figure F.1-1: Cadmus 3.7-year trajectory to the Jupiter system involves four planetary flybys that reduce the amount of propellant needed.
July 2028
PJR Maneuvers Jupiter Ganymede Europa Inbound Trajectory Initial Joviocentric Orbit Perijove Raise Orbits
10 0 0 50 100 150 200 250 Distance (km) 300 350 400 450
Figure F.1-4: The landers cover 401.4 km over the ground from descent initiation to touch down.
Cruise Stage V
Maneuver
Venus-Mars-Venus-Earth and Deep Space Maneuvers Jupiter Orbit Insertion Perijove Raise Maneuver Margin (10%) Total V
V (m/s)
763 320 1100 218 2401
EGAs GGA
July 2029
EOI
Figure F.1-2: Cadmus takes one year to reach Europa orbit once it enters the Jupiter system.
1st lander lands 2nd lander lands
Figure F.1-5: Having a cruise stage with a propulsion module saves each lander from requiring the propellant for a considerable amount of V.
Lander V
Maneuver V (m/s)
520 1680 80 228 2508 Europa Orbit Insertion Landing Hover (60 sec.) Margin (10%) Total V
Europa O Insertion rbit Jupiter O Insertion rbit Earthflyby M flyby ars 1st Venus flyby O ctober 2021 July 2022 July 2023 August 2023 June 2026 June 2025 July 2026 2nd Venus flyby
Figure F.1-6: A large portion of the V is required is for the landing.
M 2022 arch
Figure F.1-3: The bat chart shows that most of the missions life is spent in deep space, away from Jupiters harmful radiation. 13
F.2 Flight System F.2.a Cruise Stage The Cadmus propulsion module is a lowrisk integrated mission component utilizing a proven propulsion system, system redundancy, and generous margins to ensure successful arrival at Europa. Extensive testing and use of heritage hardware, a high level of autonomy in navigation, and the existence of a fault protection and safing mode system reduce mission cost and risk. Configuration: The cruise stage consists of two Europa surface landers and a propulsion module (Figure F.2). The landers are in a stacked configuration above the propulsion module. This configuration has a reinforced load bearing structure to withstand the Gs expected during launch.
The cruise stage has its own power control unit, converters, regulators, and wiring. The cruise stage also has a Pyro Initiation Unit (PIU) to detonate the pyros and separate the landers from the propulsion module prior to EOI, descent, and landing. Advances made during the JIMO mission in radiationhardened electronics will be used in the development of power system components. Propulsion: An all-chemical storable bipropellant propulsion system is used for the cruise stage and consists of two main thrusters using MMH and NTO pressurized with helium. The thrusters will provide up to 445N each to provide a satisfactory thrust to weight ratio throughout the cruise stage. This propulsion system was selected because of demonstrated success in the Jupiter system on the Galileo mission and on other planetary missions. The propulsion module has its own spherical MMH, NTO, and helium pressurant tanks. These tanks are centrally located to minimize center-of-mass translations as propellant is used2. All tanks are titaniumlined composite over-wrapped. The helium in the pressurant tank will start at about 3000 psia and maintain a pressure of about 225 psia. All tanks have a safety factor of two and will be supplied by Pressure Systems, Inc., which has delivered many spacecraft tanks and has never had an in-flight failure. A schematic of the propulsion module propulsion subsystem can be seen in Figure F.3. It exhibits standard bipropellant system design. Redundancy in pressure regulation is built in. This system provides the propulsion option with the most heritage and the least complexity. Attitude Determination, Control, and Navigation: The cruise stage is 3-axis stabilized through use of 12 attitude control thrusters. Redundant star trackers provide attitude determination about one axis (considered to be the roll axis) while
Figure F.2: The cruise stage is stacked to withstand launch Gs. Power: Power to all cruise stage electrical items is provided by the two lander Radioisotope Thermal Generators (RTGs)1. 14
redundant Sun sensors determine attitude about the other two axes (the pitch and yaw axes)3. Each lander has an Inertial Measurement Unit (IMU), and only one is used at any given time during the cruise phase. The IMU is capable of propagating attitude when information from the star trackers and/or the Sun sensors is unavailable. The star trackers have 10 mm thick aluminum shutters that close for added protection when the spacecraft is in the high-radiation environment of the Jupiter system. The shutter of the star tracker in use opens only periodically for rollaxis determination. Redundant navigation cameras provide autonomous navigation and reduce the need for frequent ground monitoring, thus reducing operations cost. Communication: All Cadmus communication is facilitated through the landers. A low level of risk is ensured with the high level of redundancy in the lander communication system. Each lander has two hemispherical, low-gain antennas utilizing X-band and a controllable high-gain antenna utilizing KAband. Beacon-tone monitoring is used to reduce ground operations cost, as continuous telemetry monitoring by ground personnel is unnecessary and the tones can be detected by small Earth receivers. Command and Data Handling (C&DH): All command and data handling operations for
the propulsion module are handled by only one of the two lander C&DH systems. The other C&DH system functions as a backup. The need for a large ground support team is reduced by use of a Remote Agent system (first demonstrated on Deep Space 1) which allows the spacecraft to autonomously diagnose and respond to most anomalies. Thermal Control: The cruise stage must endure the high intensity solar radiation at Venus during the flybys and also the cold of the Jupiter system once it arrives. Multi-layer insulation (MLI) covers the propellant and pressurant tanks. Heat from the lander RTGs is used to keep temperature critical components on the cruise stage warm. Radioisotope heater units (RHUs) will be used on thrusters to minimize power drain on the RTGs. A heat shield is employed between the main thruster and the rest of the cruise stage. Power and mass estimations for the cruise stage are shown in Tables F.1 through F.4. A conservative engineering approach has been exercised in the application of both contingencies and margins to mass and power calculations. A generous initial contingency of 30% was applied to the mass and required power of all payload and subsystem components. The two contingencies for any part have only been decreased if that part has demonstrated sufficient heritage or simplicity warranting the reduction.
15
Table F.1: Cruise mass breakdown shows generous contingencies to ensure an acceptable overall launch mass.
With Contingency (kg) Subsystem Totals (kg)
Table F.2: Cruise stage power drain for launch phase can easily be met by the RTGs.
Launch Phase Payload (landers) Propulsion Module Subsystems Propulsion Attitude Determination and Control Thermal Control Power Margin (30%) Total Power Required 0.0 0.7 20.0 1.9 6.8 187.6 CBE w/Contingency (W) Totals (W) 158.2 22.6
CBE (kg) Landers (2) Lander Adapters Propulsion Module Subsystems Propulsion Maneuver/attitude control thrusters (12 @ 0.66 kg) Main thrusters (2) Titanium lined composite overwrapped MMH tank (dry) Titanium lined composite overwrapped N204 tank (dry) Pressurant (He) Titanium lined composite overwrapped pressurant tank Lines, valves, fittings, pressure taps, filters, etc. Attitude Determination and Control Star sensors (2 @ 3.3 kg) Sun sensors with NavCams (2 @ 3.1 kg) Thermal Control Power (uses power from lander RTGs) Power control unit Regulator/converters Pyro Initiation Unit (PIU) Wiring Structures & Mechanisms Shielding Cruise Stage Dry Mass Propellant Loaded Mass Bioshield Adapter to Launch Vehicle Launch Wet Mass
Contingency
Totals (kg) 1116.2 58.2 361.6
113.3 7.9 7.2 7.3 7.3 11.2 42.9 15.1 6.6 6.2 0.15 0.15 0.15 0.15 0.10 0.15 0.15 0.10 0.10 9.1 8.3 8.4 8.4 12.3 49.4 17.4 14.1 7.3 6.8 18.1 25.2 4.9 6.2 3.2 7.2 0.15 0.15 0.30 0.15 5.7 7.1 4.2 8.3 172.8 18.1 1536.0 1801.1 3337.1 192.0 91.0 3620.1
Table F.3: Maximum cruise stage power drain during cruise phase can be met by the RTGs and batteries.
Cruise Phase (Maximum) Payload (landers) Propulsion Module Subsystems Propulsion Attitude Determination and Control Thermal Control Power Margin (30%) Total Power Required CBE w/Contingency (W) Totals (W) 291.0 74.2
36.3 11.7 20.0 6.3 22.3 387.5
Table F.4: Minimum cruise stage power drain during cruise phase can be met by the RTGs.
Cruise Phase (Minimum) Payload (landers) Propulsion Module Subsystems Propulsion Attitude Determination and Control Thermal Control Power Margin (30%) Total Power Required 0.0 11.7 10.0 2.0 7.1 148.4 CBE w/Contingency (W) Totals (W) 117.6 23.7
16
F.2.b Lander The lander design exhibits inherited or thoroughly tested components and a high level of redundancy to ensure completion of the Cadmus surface science mission. Configuration: The proven Viking and Phoenix designs serve as the basis for the lander configuration. A closed hexagonal bus serves as the central piece of the lander. Inside this bus are command and data handling components, batteries, and select science payload instruments, along with other items that require high radiation shielding and/or tight thermal control. The pressurant tank is also stored inside the bus for high packaging efficiency. Each of the two propellant tanks is set into the bus so that one hemisphere remains inside and the other remains exposed on the top platform of the bus. As a result, tank insulation will be biased to prevent formation of a large temperature gradient. The upper platform serves as the mounting base of the panoramic camera, the high-gain antenna, one low-gain antenna, and the lander arm. The arm, which has redundant actuators, has on its end an ultrasonic corer and several of the science instruments. Mounted to the bottom face of the bus are a descent imager, laser altimeter, terminal-descent and landing radar, and a second low-gain antenna. Three legs are attached to the bus to provide a stable configuration for landing and conducting surface science. Each of three non-adjacent side panels on the bus has connected to it a main engine and thruster pointing downward. An RTG is extended off a fourth side panel. Table F.5 provides a mass breakdown of the lander. Figure F.3 shows the lander in its deployed configuration.
Propulsion: The lander propulsion system uses the same MMH/NTO bipropellant combination as the propulsion module, allowing a single propulsion procurement for the mission. The lander system uses helium pressurant and is pressure-fed, also like the propulsion module. An oxidizer-to-fuel mixture ratio of 1.65 was selected to make the MMH and NTO tanks the same size. Combined with a common operating pressure for all propellant tanks, this reduces the costs of tank development and allows production of only one flight spare. All tanks are spherical titanium-lined composite overwrapped pressure vessels to keep propulsion system mass low. Following the successful example of Viking, the lander has three main engines. These are fired for the EOI maneuver and during descent to Europas surface. Together, the engines may be throttled to provide a lander thrust-toweight ratio of one for temporary hover capability near the surface. Attitude Determination and Control: A radiation-hardened hemispherical resonator gyro IMU provides attitude determination during periods in the cruise phase when data from the star and/or Sun sensors is unavailable. This IMU also propagates lander attitude after the lander separates from the propulsion module. To mitigate error buildup in the determined attitude, the IMU is recalibrated during high-gain communications sessions prior to landing. These sessions are frequent because attitude determination is necessary for proper articulation of the highgain antenna. The three main engine thrusters on the lander are offset from the lander center of mass and also may throttle and gimbal to ensure sufficient pitch and yaw control during all lander maneuvers. For roll control, the lander has four small thrusters mounted as far as
17
practically possible from its center of mass, also like Viking. Landing Guidance: While in its 100 km parking orbit about Europa, the lander takes pictures of the surface at a rate of several per orbit with its descent imager and transmits them back to Earth. These images will be used along with JIMO imagery data for landing site confirmation. The lander also uses this descent imager periodically in its landing sequence as a tool to determine velocity with respect to the ground. A highpower laser altimeter supplies continuous altitude information. Five kilometers above the surface of Europa a four-beam terminal-descent and landing radar provides an additional and precise method of determining velocity with respect to the surface. Robust hazard avoidance and autonomous touchdown site selection are provided by advanced flight software. This software shall be qualified through extensive computer simulation and hardware interface tests. Communications: Both KA-band high-gain and X-band low-gain lander antennas support communication directly with Earth. A hybrid inflatable high-gain parabolic antenna 1.0 meter in diameter upon inflation allows a high transmission data rate of about 6.5 kilobits per second (kbps) during encounter so that relatively little time and power are spent sending science data back to Earth. An inflatable antenna was used to save space in the stowed configuration versus a rigid antenna. The possibility of inflation failure is mitigated by the fact that the high-gain antenna (HGA) has a rigid central section 0.45 meters in diameter, enough large to provide a viable data rate throughout the mission. To satisfy pointing requirements on the surface of Europa, the dish is steerable by the lander computer. The lander also has two
hemispherical low-gain antennas (LGAs) to provide high redundancy to the communications system; these are attached on opposite panels of the lander body to ensure that the ability to communicate with ground is maintained at almost any spacecraft attitude. Use of the KA-band for high-gain communications provides a data rate many times higher than other available frequency bands for the given HGA diameter and transmission power. The lander has redundant KA-band transponders to produce a coherent signal that not only carries data but also allows for tracking and ranging by ground personnel. Low-power, low-mass X-band transceivers are used for low-gain communications. These transceivers produce noncoherent signals but the risk to navigation by ground is mitigated by the autonomous navigation capability onboard. Command & Data Handling: The lander C&DH subsystem contains a RAD750-class processor for processing commands and telemetry data, Input/Output serial buses for interfacing with other subsystems, and data storage devices to store telemetry and science data. All components are radiation-hardened and/or have sufficient radiation shielding and are inherited from JIMO. Processor speed and data storage volume are both sized to be four times more than necessary for the nominal mission to ensure satisfactory C&DH performance even if significant system degradation occurs. In the event of communication difficulties during the science phase of the mission, embedded command sequences and algorithms allow the lander to perform a nominal set of surface science independently of ground personnel. Data compression algorithms operate on the raw telemetry and science data and reduce the total data volume by a factor of more than two.
18
Thermal: The requirements that the lander be able to operate on Europa at surface temperatures as low as 70 K and additionally be able to endure solar intensity variations from Venus to the Jupiter system demand a sophisticated thermal control system. Passive techniques such as insulating gold paint on the lander body and aerogel insulation layers inside the body help reduce required power and complexity in thermal control. Thermal switches direct portions of the approximately 380 W of heat produced by the RTG either inside the lander body to keep electronics warm. A cover over the RTG prevents unacceptable heat dissipation to the Europa environment during the science phase. RHUs act in place of thermostatically controlled heaters on thrusters to reduce needed power for thermal control of the propulsion subsystem. Additionally, insulation and heater tape cover the propellant tanks to prevent the propellant from freezing. Appropriately placed heat shields on the lander body protect the structure and internal components during thruster firings. For instruments and mechanisms residing outside the body, thermal coatings and thermostatically controlled heaters provide thermal control throughout the mission. Power: One Multi-Mission RTG (MMRTG) supplies power to the lander. Two lithium-ion batteries complement the RTG and provide secondary power to the lander. They were sized with a specific power of 200 W-hr/kg, and a worst case depth of discharge of 75%. They support portions of the cruise phase requiring high power, such as maneuvers and periodic high-gain communications sessions. Each has three times more power than needed to satisfy power demands through the descent and landing phase, during which the power draw is at its highest for the entire mission. In addition, the batteries enable timely completion of the science mission.
Structures and Mechanisms: Aluminum forms the basic skeleton of the lander bus to create a lightweight support structure. Aluminum honeycomb sandwiched between aluminum face sheets provides light and stiff body panels. Three legs give the lander the ability to perform a stable soft-landing. Based on the successful Viking and Phoenix landers, each leg has a main strut and a pivoting wishbone arm, with crushable aluminum honeycomb inside the strut to absorb the shock of landing and allow the strut to shorten by several inches at touchdown. Multiple drop tests simulating worst-case touchdowns are scheduled for certifying that the legs will be able to withstand impact velocities of up to 6 m/s even at off-nominal angles. The legs are fully deployed before launch to eliminate the chance of an unsuccessful deployment in space. The HGA is restrained during launch by retractable pins but once in space is pointed with two rotational degrees of freedom by a jointed titanium member. This member is motor-driven and can extend the HGA away from the bus, making possible high gain communications with either lander at a wide range of spacecraft attitudes during cruise. During the Europa landing phase the HGA is again restrained to prevent touchdown damage. Retractable pins also prevent the robotic arm and panoramic camera mast from moving before deployment on Europa. Titanium construction endows high strength and low weight to both the motor-driven arm and the telescoping camera mast. Cutouts are made where possible for further weight reduction. At the end of the arm is an ultrasonic corer used to obtain material samples up to 20 cm
19
below the surface, beyond the irradiated top 10 cm of the terrain.
highly
Shielding: Aluminum shielding protects the science payload and vulnerable electronics from radiation throughout the mission, especially in Jupiters harsh radiation environment. Being aluminum, the body structure itself provides some shielding for internal components. Shutters protect sensitive instruments like the panoramic camera and the descent imager when they are not in use. Over the course of the mission, the expected total radiation dosage behind 100 mils of aluminum would be approximately 7 Mrad, with about 2 Mrad occurring in the encounter phase. With the aluminum skin of the spacecraft as well as the electronics chassis acting as shielding, an additional 4.5 cm of aluminum shielding will be placed around the electronics. This shielding results in a radiation dose of 30 krad on the electronics, allowing the use of 60 krad electronics after taking into account the factor of two uncertainty associated with Europas radiation environment. Contingencies and Margins: A conservative engineering approach has been exercised in the application of contingencies to mass and power calculations. In the estimation of total lander power, a large margin of 30% has been applied to the science payload and all subsystems to provide room for growth due to presently unknown factors. Sufficient fuel margin has been included in the 10% margin set on the V the lander propulsion system is designed to deliver.
F.3 - Launch Vehicle The Cadmus mission launch vehicle, the flight-proven Lockheed Martin Atlas V 551, is capable of delivering the spacecraft into a Venus-bound heliocentric trajectory with the desired Earth Escape Energy (C3). A wide range of expendable launch vehicles (ELVs) provided by NSA and their capabilities were evaluated. The intermediate launch vehicle class was selected because of its capability of launching a boosted mass of 4720 kg with a C3 of 16.4 km2/s2.4 Within this category, the Lockheed Martin Atlas V 551 launch vehicle best meets the required performance. Lockheed Martin has demonstrated continued success with the Atlas family of launch vehicles.
Figure F.9: The landers (with bioshield) and propulsion module fit easily inside the Atlas V 551 launch vehicle fairing, providing for an economical, low risk launch.
20
Table F.6: Lander power breakdown for descent and landing shows highest mission power drain, which can be met by the RTG and batteries. Table F.5: Lander mass breakdown shows generous contingencies to ensure an acceptable overall launch mass.
CBE (kg) Science Payload Descent imager Panoramic site imager Aqueous chemistry laboratory (ACL) Gas Chromatograph / Mass Spectrometer (GCMS) Near-Infrared spectrometer (NIR) Raman spectrometer (RS) Environmental sensors Seismometers Ultrasonic corer (USC) Lander Subsystems Propulsion Roll control thrusters (4 @ 0.66 kg) Main thrusters (3 @ 3.6 kg) Titanium lined composite overwrapped MMH tank (dry) Titanium lined composite overwrapped N204 tank (dry) Pressurant (He) Titanium lined composite overwrapped pressurant tank Lines, valves, fittings, pressure taps, filters, etc. Attitude Determination Inertial Measurement Unit Landing Terminal-descent and landing radar Laser altimeter Communications High-gain hybrid inflatable parabolic antenna (Ka-band) X-band low-gain hemispherical antennas (2) with transceivers Ka-band transponders (2 @ 1.5 kg) Filters, switches, and diplexers Command and Data Handling Thermal Control Power RTG Lithium ion secondary batteries Power control unit Regulator/converters Wiring Structures and Mechanisms Shielding Lander Dry Mass Propellant Loaded Mass 0.2 2.5 2.0 4.7 1.0 2.5 0.2 0.2 1.0 Contingency 30% 30% 30% 30% 30% 30% 30% 30% 30% w/Contingency (kg) 0.3 3.3 2.6 6.1 1.3 3.3 0.3 0.2 1.3 229.7 49.3 2.6 10.8 1.3 1.3 2.3 8.9 13.4 7.0 1.7 0.4 4.0 2.0 1.5 2.4 2.0 25% 25% 15% 15% 10% 15% 25% 10% 30% 30% 30% 30% 30% 30% 30% 3.3 13.5 1.5 1.5 2.6 10.3 16.8 7.7 7.7 2.7 2.2 0.5 12.9 5.2 2.6 2.0 3.1 2.6 Subtotals (kg) Totals (kg) 18.5 Descent and Landing Phase Science Payload (only Descent Imager on) Lander Subsystems Propulsion Attitude Determination Landing Communications Command and Data Handling Thermal Control Power Mechanisms Margin (30%) Total Power Required CBE w/Contingency (W)
Totals (W) 4.6 280.7
42.9 28.6 44.5 117.7 3.0 20.0 24.1 0.0 85.6 370.8
Table F.7: Minimum power drain for surface operations falls well below RTG-provided power.
Surface Operations (Minimum) Science Payload Lander Subsystems Propulsion Attitude Determination Landing Communications Command and Data Handling Thermal Control Power Mechanisms Margin (30%) Total Power Required CBE w/Contingency (W) Totals (W) 0.0 43.7
0.0 0.0 0.0 17.0 3.0 20.0 3.7 0.0 13.1 56.8
2.6 12.4 61.4
Table F.8: Lander tanks exhibit high performance and safety for risk mitigation.
Tank MMH NTO He pressurant 248.2 309.9 558.1 Inner radius (m) 0.32 0.32 0.26 Volume (m3) 0.137 0.137 0.074 Tank performance factor (m) 38100 38100 38100 Safety factor 2 2 2
40.0 2.6 2.5 3.1 5.0
15% 30% 15% 15% 15%
46.0 3.3 2.8 3.5 5.7 62.0 18.6
21
Table F.9: Maximum power drain for surface operations can be met by the RTG and batteries.
Surface Operations (Maximum)
Science Payload Panoramic site imager Aqueous chemistry laboratory (ACL) Gas Chromatograph / Mass Spectrometer (GCMS) Near-Infrared spectrometer (NIR) Raman spectrometer (RS) Environmental sensors Seismometers Ultrasonic corer (USC) Lander Subsystems Propulsion Attitude Determination Landing Communications Command and Data Handling Thermal Control Power Mechanisms Margin (30%) Total Power Required 0.0 28.6 0.0 32.0 3.0 20.0 6.5 10.0 41.7 180.6
RTG Fuel Tan GCMS Ports
He Antenna Switch
Other subsystems Diplexer Diplexer Diplexer Explosive Valve Antenna Switch Antenna Switch Regulator Filter
3000 psia initial Fill Valve
CBE w/Contingency (W)
2.6 11.7 16.9 1.3 0.7 0.3 0.2 5.2
Totals (W)
38.8
Secondary Regulator Burst Disc Relief Valve
Transponder
Transponder Transceiver Transceiver MMH 225 psia NTO 225 psia
C&DH Computer
Bipropellant Valve
112.0
Figure F.5: A highly redundant system provides reliable lander communications.
Hybrid Inflatable High Gain Antenna Low Gain Antenna Environmental Sensors Pancam Thrust Chamber Valve Main Thrusters Roll Control Thruster Pair
Figure F.4: Lander propulsion subsystem has redundancy in pressure regulation and exhibits standard bipropellant system design
Robotic Arm
He Filter Regulator Explosive Valve
3000 psia initial Fill Valve
Low Gain Antenna ACL Ports Main Thruster (1 of 3)
Raman Probe Ultrasonic Corer
Secondary Regulator Burst Disc Relief Valve
MMH 225 psia Bipropellant Valve
NTO 225 psia
Figure F.3: The configuration of the Cadmus lander shows similarity to previous successful planetary surface landers and enables the lander to efficiently perform the science mission.
Science
I/O Board RAD750-class Processor Communication
Subsystem Telemetry
I/O Board
Main Thrusters
Attitude Control Thrusters (12)
Figure F.8: Propulsion Module propulsion system features redundant attitude control thrusters
Figure F.7: The Atlas V 551 provides a reliable, low cost launch option for the Cadmus mission. Credit: International Launch Services 22
NVM Data Storage
Figure F.6: Data flow schematic shows a logical and reliable method of transmitting data to Earth.
F.4 Ground Systems The Ground Systems consist of existing assets in order to minimize cost. The ground systems for operations will utilize existing systems created through the JPL Interplanetary Network Directorate (IND), including telecommunications, interplanetary navigation, information systems, and computing. The data returned will be delivered to the mission operations center at JPL, for use by the science team, and for concurrent archival by the JPL PDS. F.5 Communications Approach The Cadmus communications approach meets signal strength and data rate requirements throughout the mission with minimal complexity and high reliability. F.5.a - Uplink and Downlink: All uplink and downlink communications are direct between Earth and the landers. The absence of relays reduces complexity and eliminates the possibility of mission failure due to reliance on separate space hardware for data return. High-gain communications are in the KA-band, with uplink at approximately 34.5 GHz and downlink at 32 GHz. Low-gain communications are in the X-band, with uplink near 7.19 GHz and downlink at 8.45 GHz. These bands and frequencies were selected to fall within those supported by DSN and also based on transponder/transceiver capabilities and achievable rates. A much higher data rate is achievable with KA-band than other supported bands, making it the best choice for high-gain communications, but the mass and power savings available with Xband transceivers make X-band more attractive for low-gain communications. Through most of the cruise phase beacon-tone monitoring is used, meaning the active
computer on board analyzes telemetry data to assess the health of spacecraft components. The computer then summarizes the overall spacecraft status with a single-frequency tone it transmits to Earth, rather than modulating the telemetry data onto a downlinked carrier signal. The frequency of the tone received immediately tells ground personnel whether spacecraft performance is nominal or degraded, and how urgently full-scale communications are needed. Beacon-tone monitoring reduces the need for constant human monitoring of telemetry, thus significantly reducing operations cost. Also, the tones are detectable by relatively small Earth dishes, so communications costs are lessened. DSN 34 m diameter antennas are employed periodically during the cruise phase (~once per month) to receive telemetry and to uplink any necessary command sequences or software updates to the spacecraft, and the 70 m DSN dishes are employed throughout the encounter phase to ensure reception of telemetry and science data. A signal-to-noise ratio of 9.8 dB or higher shall be maintained in all communications activities to provide a 3 dB link margin over the 6.8 dB necessary for a sufficiently high probability of signal decoding. Coding is used to prevent errors in data transmission and shall adhere to the Consultative Committee for Space Data Systems (CCSDS) standard. The coding scheme to be employed shall be a turbo code of appropriate rate and block size to achieve higher performance in achieving the desired bit error rate than concatenated and Reed-Solomon codes while maintaining lower coding complexity than many predecessor spacecraft. F.5.b - Critical Events Coverage: Key telemetry values are transmitted at 10 bps to Earth during all critical events, including launch, planetary flybys, trajectory correction maneuvers, Jupiter orbit insertion, perijove-
23
raise burns, separation of landers from propulsion module, Europa orbit insertion, descent, and landing. Table F.10: High-gain communications plan provides a high signal-to-noise ratio margin from launch up through encounter phase.
Parameter Encounter phase Earth-Europa distance Frequency band Uplink frequency Downlink frequency Estimated HGA parabolic efficiency HGA diameter (uninflated) Data transmission rate (uninflated) HGA gain (uninflated) Approximate bandwidth Transmitter antenna diameter (inflated) Data transmission rate (inflated) HGA gain (inflated) Approximate bandwidth HGA transmission power Ground station antenna diameter Ground station background noise temperature Ground station antenna gain Atmospheric loss factor Space loss Power received (HGA uninflated) Power received (HGA inflated) Minimum acceptable signal-tonoise (S/N) ratio S/N ratio (HGA uninflated) Margin over minimum acceptable S/N (uninflated) S/N ratio (HGA inflated) Margin over minimum acceptable S/N (inflated) Value 6.3 AU Ka 34 GHz 32 GHz 0.6 0.45 m 1.35 kbps 13600 1350 Hz 1.0 m 6.5 kbps 67400 6500 Hz 10.8 W 70 m 80 K 3.30E+08 1.5 1.29E+30 2.03E-17 1.00E-16 6.8 dB 9.85 dB 3.05 dB 9.96 dB 3.16 dB
Table F.11: HGA-provided data rate exceeds the required data rate for transmitting all first-phase science data within 12 days, even if the antenna is uninflated.
Parameter Science data volume CBE Approximate transmission time available Minimum required data rate Uninflated HGA provided data rate Inflated HGA provided data rate Percent of data returnable by uninflated HGA Value 500 Mb 4.52days 1.28 kbps 1.35 kbps 6.5 kbps 100%
Table F.12: HGA-provided data rate exceeds the required data rate for transmitting all second-phase science data within 20 days.
Parameter Science data volume CBE Approximate transmission time available Minimum required data rate Uninflated HGA provided data rate Inflated HGA provided data rate Percent of data returnable by uninflated HGA Value 1800 Mb 7.63 days 2.73 kbps 1.35 kbps 6.5 kbps 49%
F.6 - Mission Operations The Cadmus mission operations plan utilizes a low cost approach while maintaining high efficiency and competency. The high degree of autonomy of the spacecraft will ensure a small operations team and low operations cost. The use of beacon tone monitoring reduces DSN use. A relatively short science mission also minimizes mission operations cost. Typical daily operations will consist of a Flight Operations officer, on duty at JPL during all periods to monitor the beacon tone and ensure nominal spacecraft operation.5 When the spacecraft is being 24
tracked by DSN, this person will monitor realtime spacecraft data for any anomalous events. The Flight Operations officer will serve as a single point of contact between the entire JPL flight team and external teams such as DSN. The flight team will consist of a Spacecraft Team, Navigation and Control Team, Descent, and Landing Team, and a Surface Science Team. During the cruise phase, the Flight Operations officer will monitor the mission sequence of events to ensure timely completion. The spacecraft navigation, spacecraft and payload status, and instrument health will also be closely monitored. The Flight Operations officer will work closely with the Network Monitor and Control (NMC) operator. The NMC operator will be monitoring and controlling all aspects of the assigned DSN antenna. The flight team will be contacted at the first sign of any off-nominal data. Instrument and science payload checks and calibration will be performed prior to the descent, and landing phase of the mission, and again once landing has been achieved. The mission operations high level of autonomy and streamlined method of communications reduces operations costs, allowing more of the project budget to be directed towards spacecraft engineering and science. F.7 Technology Development Through rigorous testing and flight qualification, Cadmus will be ready for a 2021 launch. Most of the instruments and hardware used for the Cadmus mission are flight proven. Furthermore, the Cadmus architecture has been designed such that a failure in the nonflight proven components will not result in a
loss of the mission. For example, Cadmus intends to test a hybrid inflatable antenna that will provide a high rate of data transmission to Earth so that minimal ground and power resources will be required for communications. If the device malfunctions, however, and does not inflate, the base, uninflated portion of the antenna is sufficient to complete the baseline mission, although ground operations and resource management will be more complex.
Figure F.10: A hybrid inflatable antenna will maximize communications capabilities. (Credit: ILC Dover) The demonstration of this technology will promote its use in future missions. The technology can also be transferred to the private sector, where it may be used by industry and universities. Figure F.10 shows a conceptual image of a hybrid inflatable antenna.The most significant risk associated with the technology readiness of this mission concerns the high radiation environment on Europa. Although most of the instruments and hardware are flight proven, many of them have not flown in such high radiation levels. To mitigate this risk, heritage components were taken from the JIMO mission which will have successfully operated in the Europan 25
environment. Knowledge gained from JIMO regarding radiation hardening and radiation Table F.13: Cadmus utilizes proven and qualified components to ensure success.
Item Description GCMS TRL Basis MSL, CassiniHuygens Development Radiation hardening and testing
shielding will also be applied to environmentally qualify Cadmus components. All components will be subjected to rigorous environmental testing using JPLs radiation testing facility. Testing will also be required for the honeycomb aluminum shock absorbers in the landers legs which are meant to soften the landing impact. This method of shock absorption was proven by the Viking and Phoenix missions, but drop tests will have to be performed to determine mission specific parameters for Cadmus lander mass and impact velocity. From these tests, the amount of crushable aluminum honeycomb required for an acceptable amount of loading will be determined as well as the change in vertical height of the lander to predict the distance from the lander body to the ground after landing. Table F.13 shows the technology readiness level and basis for the major components of the Cadmus mission. F.8 Instrumentation The Cadmus mission utilizes a suite of instruments with significant flight heritage and operational testing in analogous environmental conditions. All Cadmus instruments are directly traceable to the stated scientific objectives, as shown in the Science Traceability Matrix (Table D.3). Cadmus scientific mission consists of a suite of nine instrument packages, with eight having proven mission heritage (Table F.12). Every instrument will carry a Technology Readiness Level (TRL) of at least eight by the end of Phase C, meaning that the actual system is completed and has been flight qualified through testing and demonstration. The instruments are also able to meet the strict
6
Environmental Sensors Aqueous Chemistry Lab Pancam NIR Spectrometer Raman Spectrometer Seismometers Descent Imager Ultrasonic Corer Bus (aluminum) Main Thrusters (Kaiser Marquardt R4D modified) C&DH (computer) Batteries (Li ion) Star Sensors Sun Sensors IMU Data Storage Thermal Crushable Shock Absorbers Hybrid Inflatable Antenna
6
Phoenix, MSL, JIMO, Rosetta
None Required Radiation hardening and testing Radiation hardening and testing None Required Radiation hardening and testing Radiation hardening and testing Radiation hardening and testing Radiation hardening and testing None Required Radiation hardening and testing, minor tailoring None Required Minor Tailoring None Required None Required None Required None Required None Required Drop tests Radiation hardening, rigorous testing to ensure reliability
6 6 9 6 6
Phoenix, MSL MER, Phoenix MER, Galileo, JIMO MSL Viking MER, Phoenix, CassiniHuygens, MSL FIDO ground testing Various Leasat, Intelsat 6, Milstar JIMO Various JIMO JIMO JIMO JIMO Various Phoenix, Viking
6
6 6
6
9 6 9 9 9 9 9 6
5
ILC Dover
26
operating requirements of the Cadmus mission (Table F.15). The Gas Chromatograph/Mass Spectrometer22 (GCMS) provides high spectral resolution elemental analysis of subsurface samples essential to the mission objectives of Cadmus (Table D.3). It has extensive heritage (Table F.12) in planetary surface missions and will undergo thorough testing in environmental conditions similar to that of Europa. The GCMS consists of six identical, single-use chambers - three for analyzing samples and three for reference. Each sample chamber will have a sealable chute at its top for deposit of the sample by the corer. To meet the power and mass requirements of the mission, the GCMS will take advantage of miniaturization advancements developed at JPL. The descent imager is based on previous successful instruments, and provides high resolution context imagery for choosing and verifying a landing site for the Cadmus mission. The imager will also provide the capability for active hazard avoidance for the spacecraft during landing. Cadmus panoramic camera26 provides high spatial resolution context imagery to meet the mission science objectives (Table F.14). The pancam consists of two cameras to provide stereoscopic imagery and will rest atop an extendable mast on the lander (Figure F.3). The camera has four filters for red, green, blue, and near-IR images. For context imagery, the three visible spectrum images will be overlaid to produce a full color image. The pancam will be capable of 360 rotation in azimuth and +- 90 elevation. Thorough testing in a relevant environmental conditions will prepare the imager for the mission. In order to reduce the inherently large data requirements, the pancam will utilize ICER image compression developed by NASA JPL.
The ultrasonic corer24 (USC) allows the lander to collect subsurface samples for analysis by the GCMS using much less force and power than conventional coring methods. The USC is the only instrument without flight history (Table F.12); however, it does have thorough ground testing through JPLs Field Integration Design and Operations (FIDO) technology testbed. The USC utilizes a piezoelectric actuator to make the corer similar to a jackhammer. As a result, the corer does not get dulled or need to be replaced like a conventional drill bit. The USC also does not suffer from drill walk as it involves no driving torque. The corer will be located on the lander arm, and will drill up to a depth of 20 cm and remove a sample core of 1 cm in diameter. The estimated surface hardness of Europa, which is similar to that of basalt, will allow a drilling speed of 1 cm / hour. This speed can be increased by supplying more power to the drill. The USC will be built by Cybersonics Inc. which holds the patent on this ultrasonic drilling technology. The aqueous chemistry laboratory23 (ACL) is a flight proven instrument that allows Cadmus to analyze the subsurface environment. Thorough testing in conditions similar to that on Europa will prepare the ACL for the Europa mission. The ACL consists of three single-use...