CubeSat Constellation Mission Design

8/20/2019 CubeSat Constellation Mission Design

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AE 4357 Team L • Nov 2015 • Georgia Institute of Technology

Contents

1 Executive Summary 2

2 Mission Objectives & Impact 4

3 Science Traceability Matrix and Requirements Flowdown 8

4 Performance Requirements 9

5 Program Requirements 10

6 Mission Implementation 11

7 Payload System 16

8 Flight System 20

9 System Engineering 29

10 Risk Identification & Mitigation 37

11 Management, Schedule & Cost 40

12 References 47

13 Nomenclature 49

List of Figures

1 GT SWARM Mission Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52 Objectives shared by NEO stakeholders. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63 GT SWARM Science Traceability Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84 GT SWARM Science Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95 GT SWARM Instrument Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 96 GT SWARM Mission Operations Requirements . . . . . . . . . . . . . . . . . . . . . . . . 107 GT SWARM Launch Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108 GT SWARM Cost Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109 GT SWARM Concept of Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1110 Didymoon Mass Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

11 STK Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1512 Mission Life Cycle Functional Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1713 Nanocamera . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1814 Thermal Imager Picture and Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . 1915 IR Spectrometer material composition band . . . . . . . . . . . . . . . . . . . . . . . . . . . 2016 NASA Goddard Mini Ion-Neutron Mass Spectrometer . . . . . . . . . . . . . . . . . . . . 2017 Configuration for CubeSat Architecture trade study . . . . . . . . . . . . . . . . . . . . . . 2118 Pugh Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2119 Imager and Diver Functional Diagram with both payloads . . . . . . . . . . . . . . . . . . 2220 SWARM Imager Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22


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21 SWARM Imager Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2322 EPS Functional Diagrams . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2423 SWARM Thermal Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2524 SWARM Telecommunications Subsystem Diagram . . . . . . . . . . . . . . . . . . . . . . 2725 SWARM C&DH Subsystem Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2826 Mass and Volume Budgets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2927 Power Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

28 SWARM Power Variables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3229 SWARM Solar Cell Output As A Function of Temperature [18] . . . . . . . . . . . . . . . 3230 Telecommunications Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3431 ∆V Budgets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3632 Likelihood Consequence Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3833 Quantitative Risk Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3934 Organizational Chart for Group L . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4035 Cost Estimate Based on Small Satellite Cost Model . . . . . . . . . . . . . . . . . . . . . . 4336 Quick Cost Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4437 Historical Cost Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4538 SWARM Component Level Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4639 Grass Roots Engineer Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

List of Tables

1 Pass/Fail Criteria for SWARM Mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82 Constraints prescribed by ESA on the CubeSat Opportunity Payload Intersatellite Net-

work Sensors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83 AIDA Mission Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134 Near-Earth Asteroids similar to SWARM’s target asteroid Didymoon . . . . . . . . . . . . 145 Different approximated circular orbital parameters. . . . . . . . . . . . . . . . . . . . . . . 15

6 Average access times for the Imager CubeSat. . . . . . . . . . . . . . . . . . . . . . . . . . 167 Thermal Imager Picture and Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . 198 Architecture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219 SWARM Thermal Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2410 Specifications for the CubeSatShop Nanosatellite Micropropulsion System . . . . . . . . 2611 Miniature Integrated Star Tracker (MIST) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2612 Transceiver . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2713 Technology Readiness Levels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2914 Thermal Budget Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3515 Thermal budget for Cold Case (500 m, β = 0) . . . . . . . . . . . . . . . . . . . . . . . . . . 36

16 Thermal budget for Hot Case (500 m, β = 40) . . . . . . . . . . . . . . . . . . . . . . . . . . 3617 Data volume for a unit of output data from each of SWARM’s science payloads. . . . . . 3718 SWARM Overall Development Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

1. Executive Summary

Currently, there are several existing theories for the origins of Earth’s water, how chemicalelements are disseminated through space, and the origin of the planetary bodies of the SolarSystem. To investigate these problems and to narrow down on specific theories of water origin

or particle disbursement, SWARM will study the formation of astronomical bodies by analyzing the


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dissemination of chemical elements after an asteroid impact. In this proposed mission, the impactcrater will be produced when the NASA Double Asteroid Redirection Test (DART) spacecraft impactsthe Didymoon asteroid at a very large velocity. SWARM (Solar System Water and Rock Measurementswill complement the European Space Agency’s (ESA) Asteroid Impact Mission (AIM) by performinghigh-resolution observations of an impact event on the Didymos asteroid system’s secondary body tounderstand how astronomical bodies form. Additionally, the elemental and isotopic composition ofthe secondary body will be analyzed to identify the origin of the binary system and its constitutive

material. The SWARM mission formation is composed of two 3U CubeSats (Diver and Imager). TheDiver CubeSat will analyze the elemental composition of the surface and subsurface material of theDidymoon asteroid by collecting samples from the impact ejecta plume. The Imager CubeSat wilperform high-resolution thermal and optical observations of the impact event. The proposed AIMDART, and SWARM mission will be the first space mission perform high-velocity impact dynamicsmodeling using in-situ operations. Moreover, this mission will support and extend existing models ofplanetary impact and high kinetic energy collisions with the experimental data from AIM.

In an Op-Ed Space News article from April 22, 2015, the Chief Technology Officer of TerminalVelocity Aerospace and former NASA Chief Technologist Dr. Robert Braun said, “a compelling questcould be defined focused on the diversity and distribution of liquid resources across the solar system.”[1] Moreover, Andres Galvez, the European Space Agency (ESA) General Studies Program Managerstated the following during a press conference about the future of ESA’s projects: “It is increasinglyclear, directly or indirectly, these [solar system formation] processes might well have played a crucialrole in the development of life on our planet. On one side, through massive collisions and massextinctions, on the other by a ’mass input’ of the building blocks of life, possibly including Earth’swater.” As affirmed through both assessments, there exists a critical need for a mission to continue toisolate the origin of Earth’s large amounts of water as well as its rock-like and potentially non-chondriticstructure (Wahlund 2015). Although several missions have gathered data to provide evidence for thesepoints of origin, there is still no consensus primarily because there are not enough in situ measurementson Near Earth Objects (NEO) to fully support any given hypothesis.

In addition, the theories regarding the origins of the Solar System are still unconfirmed. Junko

Kominami, a professor of Astronomy at the Tokyo Institute of Technology, notes that there are severalhypotheses for the origins of the terrestrial planets, but that no hypothesis is advantageous over anotherHowever, Kominami believes that the process of colliding and combining minor asteroids is criticalto understanding of the formation of planets [2]. Near Earth asteroids are the result of billions ofyears of asteroid collisions. The encounters of these asteroids with other large celestial bodies willcause imprints into their structures and cause other pieces to break. Currently, there have not beenany space experiments or studies focused on directly observing an asteroid collision event. As a resultthis mission will provide an opportunity to directly observe a collision using ground based and in situobservations. Observations on the internal structural properties and orbital dynamics will be made

before and after the impact to determine the impact’s effect on the astronomical body.SWARM’s solution is to use a optical nanocamera, thermal imager, and mass spectrometer to

characterize the secondary body of the asteroid system. These payload items will be used on two3U CubeSats that will be transported to the asteroid system from another mothership-like satellitecalled AIM. CubeSats are the best option for studying binary asteroid systems for several reasons. FirstCubeSats are low cost. Because of the small amount of mass aboard each CubeSat, launch costs andoverall cost of the satellites will be minimized. Secondly, the mission lifetime is around 3 monthsAs a result, building a large satellite would be unnecessary in terms of cost and size since largersatellites are typically used for missions that will last a year or more. Additionally, CubeSats introduceredundancy to the proposed mission. Because the AIM and DART mission is so complex, the CubeSatswill serve as both redundant observers of the event and contributing science additions to the missionFinally, CubeSats are a key future component of multiple different space missions. Because of this


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it is important to begin testing CubeSat devices and technology in deep space. CubeSats require thetechnical testing to ensure that in the future, CubeSats will have the ability to survive traveling furtherdistances from Earth’s orbit without suffering from large amounts of radiation damage. As a resultthis mission will feature CubeSats that will demonstrate deep-space optical communication technologyand create an inter-satellite communication network with CubeSats and a lander. As the first missionof its kind, this mission represents where the future of space technology is heading.

SWARM will cost approximately $5 million including cost of the CubeSats, integration, assembly

test, program management, launch support, operations support, and ground equipment. This is basedon 4 different cost comparisons: historical cost model, small satellite cost model, Quick cost modeland a grass roots cost model. The grass roots model was the most influential model because the othermodels were poor approximations of this mission. There are no other missions similar to this missionthat a historical cost model would represent an accurate representation of the overall cost. Additionallythe small satellite cost model is more representative of larger satellites near 100 kg or more. Also, thequick cost model is a general model that isn’t specific to a particular subset of missions. As a result,the grass roots model presents the most accurate estimate although the other models are taken intoaccount to provide other cost estimates. In total, $5 million for a CubeSat mission with a high scientificmission value is relatively inexpensive compared with other missions of similar scientific significance

This mission is essential now because of the recent water discovery on Mars. In September 2015NASA released that their team, in conjunction with Dr. Lujendra Ojha of Georgia Tech, discoveredwater on Mars. Specifically, after extracting spectral information from pixels of the CRISM satellite’sinstrument data, Ojha and his team determined that the salts: magnesium perchlorate, magnesiumchlorate, and sodium perchlorate, had water molecules interspersed in their crystal structures. [5]In other words, there is strong evidence that the salt particles lined along the steep slopes of Marsmountainous structures were deposited by flowing water. As a result of this discovery, there is alarge push to discover the origins of water on Earth and see if other celestial bodies carry water. Thismission can not only help isolate where the origin of Earth’s water, but also whether asteroids wereonce capable of providing the same capabilities as Mars. Additionally, one of the top science missionsfrom the National Academy of Science is to discover the origins of the Solar system. Therefore, one of

the top grants in the United States is geared towards isolating data that can provide more accuratemodels for the Solar System.

The return of this mission consists of two major items. First, SWARM’s data about the structuraland kinetic dynamics of the binary asteroid system will either confirm or deny several different modelsthat are proposed for the origin of the Solar System. Second, the elemental data about Didymoon wilisolate whether asteroids of this class can transport water-like particles depending on the collisions ofthe asteroids, its spin, its rotational evolution, and its evolutionary coupling. Both of these returns willprepare future CubeSat missions by specifying which models should be studied further to confirmtheir accuracy. For example, if nearby binary asteroid systems don’t have properties that are capable oftransporting water, then there is no further reason to explore these systems for these particles.[6]

2. Mission Objectives & Impact

2.1 Science Objectives

SO1: The impact response of a small body as a function of impact conditions and physical parameters is crucial to understanding the role of collisions in the history of the Solar System.

Dr. Gene Shoemaker, one of the founders of the field of planetary science, proved that asteroid impactis an important process in planetary formation [8]. This mission objects to further study the roleof asteroidal impact on the formation of astronomical bodies by using collisional data from DART’s


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impact event to infer the Solar System’s collisional history. Furthermore, collisional data from DART’simpact is essential in analyzing and modeling the effects of man-made collisions on asteroids and theirtrajectories. This is important in the field of planetary defense, where future missions would entailman-made collisions in order to deflect asteroids.

SO2: The molecular and isotopic compositions of a binary asteroid system can provide informationon the formation of binary systems, the evolution of the Solar System, and the origins of water on

Earth.

The molecular composition of a solid body provides information about the evolutionary steps the bodymay have gone through before arriving at its present state. In addition to the body’s evolutionary stepsthe isotopic composition of an astronomical body can also provide insights into its place of originThe most well-known example with regards to this usage of isotopic composition is the Deuterium-toHydrogen (D/H) ratio. It has been determined that bodies with similar D/H ratios originated from thesame part of the protosolar nebula [20]. The D/H ratio can also potentially enable scientists to settlethe debate between the chondritic (asteroidal) or cometary origin of Earth’s water [ 7].

2.2 Mission Objectives

Figure 1: GT SWARM Mission Objectives

2.2.1 MO1In order to fulfill the first science objective of using collisional data to understand the role of collisions inthe history of the Solar System, the mission purports to collect collisional data through high-resolutionthermal and visual imaging of the impact crater’s formation and evolution. This information willenable scientists to measure the changes in crater size and depth over time, as well as the velocitysize, temperature distribution, and other characteristics of the ejecta plume. This collisional data is keyto understanding the extent to which DART’s impact perturbs the asteroid’s orbital parameters andinternal structure. about the effects of the collision will in turn allow extrapolation to other bodies inthe Solar System whose craters can be analyzed to obtain an estimate of their collisional history. Withthe data returned from SWARM and AIM as a starting point, scientists will be able to create planetary

evolution models that take into account impact events across astronomical time scales.

2.2.2 MO2

Various methods are available through which molecular and isotopic abundances can be measuredAccording to Dr. Rob Staehle at the Jet Propulsion Laboratory, “measurements made in situ by a massspectrometer would give the highest fidelity results.” SWARM’s low orbital velocity due to the targetasteroid’s low mass makes this mission an ideal one for which mass spectroscopy can be employedThe benefits of a low orbital velocity on mass spectrometry can be attributed to the fact that a lowvelocity provides the spacecraft with adequate residence time to collect samples required for massspectrometry.


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Figure 2: Objectives shared by NEO stakeholders.

2.3 Knowledge Gain Resulting from SWARM

As mentioned in the previous sections, SWARM is a part of the AIM mission, which is in turn part ofa larger effort called the Asteroid Impact & Deflection Assessment (AIDA) mission. The three mainpillars of the AIM mission are as follows:

• Technical demonstration of deep-space optical communications and inter-satellite links• Asteroid deflection assessment by means of kinetic impact• Planetary science study on the history of the Solar System and the formation of celestial bodies

particularly binary system formation through the YORP spin-up mechanismThe three areas collectively offer the potential for knowledge gains that serve all NEO explorationstakeholders whose objectives include planetary defense, science, human exploration, and resourceutilization [9]. SWARM will contribute to each of the three aspects of the AIM mission, thereby enablingmany of the knowledge gains shown in Figure 2.

2.3.1 Formation-flying and Inter-satellite Link Technology

Through the technical demonstration on inter-satellite communications links, progress can be madetowards larger satellite constellations that function autonomously through communicating with eachmember, much like the eponymous swarm. This is especially important in the current trend towardsaccomplishing more with multiple smaller satellites, potentially enabling radical new methods inconstellation-based adaptive synthetic aperture arrays (SAR) for specific high-resolution remote sensingoperations. A very practical example would be the technologies required for precise formation-flyingand real-time communications between the multitude of satellites required to image double-lobedactive galactic nuclei (DRAGNs) and measure energy transport from super-massive black holes to theintergalactic medium. Towards ESA’s agenda, this mission would “increase Europe’s competitivenessthrough the qualification of technologies and operations relevant to other missions, in particular in thearea of autonomous guidance, navigation and control, and spacecraft TT&C (Telemetry, Tracking, andCommand)” [9].

2.3.2 Impact Models and Scaling Laws

With DART’s impact velocity currently estimated to be around 6.1 km/s which is consistent withaverage asteroidal impact velocities, SWARM and AIM will allow verification of hyper-velocity impact


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modeling and scaling laws at an unprecedented scale. Through the observation of DART’s impactevent on Didymoon and the measurement of the subsequent crater and plume evolution, the momentumtransfer efficiency β can be determined, which is defined as:

β = momentum change

momentum in put (1)

Determination of the momentum transfer efficiency will require the initial state of the asteroid, theimpact velocity, the impact location and angle, and detailed properties of the entire asteroid. SWARMwill measure the impact conditions from a distance orders of magnitude closer to the impact site thanthe AIM spacecraft. The momentum transfer efficiency provides a benchmark for the assessment of thekinetic impact technique for deflecting potentially hazardous asteroids (PHA), and will subsequentlyenable the development of higher-fidelity impact models and scaling laws. These models will not onlyfacilitate the design of similar concepts in the future, but will also enable planetary scientists to bettersimulate the early Solar System where collisions between asteroids were commonplace occurrencesthat led to the formation of other planetary bodies.

2.3.3 Asteroid Chemical/Minerological Properties

With planetary science being one of AIM’s secondary goals, SWARM objects to fulfill this goal bycharacterizing the chemical properties of the Didymos system. This mission objective provides anavenue to perform for the first time molecular analysis on samples from a binary asteroid systemto ascertain the elemental, chemical, and mineralogical composition of the system. Compositionainformation is key to constraining the formation of the Didymos system, and can be extrapolated toother bodies to provide insights into the collisional history, evolution, and lifetime of other asteroidpopulations. Furthermore, the ability to determine the composition of binary asteroid systems will helpestablish the suitability of these systems as candidates for future explorations and asteroid deflectiontests [10].

2.4 Mission Pass/Fail Criteria

At a minimum, Diver and Imager need to be deployed successfully to ensure the mission is carried outThe Diver CubeSat needs to intersect the impact ejecta and college sample material to provide an in situmeasurement of the composition without remote sensing. In situ measurements are more accurate andprovide the ability to take higher resolution photos of the ejecta to understand more about the kineticimpact. However, remote sensing could be used to accomplish this criteria if necessary. One goal ofSWARM is to determine the ejecta plume size and evolution. This is significant to understanding binaryasteroid system origins. Therefore, if this data is corrupted, then our mission has failed. Another goalof SWARM is to measure the impact the crater size, depth, and rate of growth. The crater providesmore information about how significant collisional evolution is within celestial bodies. AIM cannottake this data with its instruments so SWARM needs to this data to accomplish its mission. Anothergoal is to characterize the temperature change at the point of impact. Heat is transferred due to achange in the kinetic energy of DART and Didymoon. Therefore, a heat analysis of the point canshow how the impact transferred kinetic energy between the 2 bodies. The D/H ratio will shine lighton whether water particles were once on the asteroid or if the body has the capability to producewater-like particles. Determining the molecular abundance of molecules and atoms will show at a basiclevel what is expected from a binary asteroid system. Finally, the most critical goal is to transfer all ofthe scientific data back to a ground station through AIM. The intersatellite link between SWARM andAIM will need an appropriate data rate to prevent data from being lost in the transfer.


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Criteria Minimum Baseline Full

1 Successful Deployment of Imager and Diver CubeSat x x x2 Intersect impact ejecta plume and collect sample material x x3 Measure impact ejecta plume and collect sample material x x x4 Measure impact crater size, depth, and rate of growth x x x5 Characterize temperature change at the impact site during impact x6 Determine the D/H ratio from plume sample x x7 Determine molecular abundances of C, H, O, N, and Si x8 Transfer all scientific products back to Earth through AIM using intersatellite links x x

Table 1: Pass/Fail Criteria for SWARM Mission

3. Science Traceability Matrix and Requirements Flowdown

3.1 Science Traceability Matrix

Figure 3: GT SWARM Science Traceability Matrix

3.2 Programmatic Constraints

Table 2: Constraints prescribed by ESA on the CubeSat Opportunity Payload Intersatellite Network Sensors.

Total Volume 2 x 3U CubeSats

Total Mass Up to 9kgSize 3U for each CubeSat

Design Lifetime Interplanetary cruise ( 2 years) Operations (3 months)Inter-satellite Link S-band full duplex transceiver

Data rate Up to 1 MbpsTotal data volume Up to 1 Gbit allocated for the whole mission

Separation Conditions 0.5 - 2 m/s


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4. Performance Requirements

4.1 Science Requirements

Figure 4: GT SWARM Science Requirements

4.2 Instrument Requirements

Figure 5: GT SWARM Instrument Requirements


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4.3 Mission Operations Requirements

Figure 6: GT SWARM Mission Operations Requirements

4.4 Launch Requirements

Figure 7: GT SWARM Launch Requirements

5. Program Requirements

5.1 Cost Requirements

Figure 8: GT SWARM Cost Requirements


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Figure 9: GT SWARM Concept of Operations

5.2 Cost Management and Scope Reduction

a) The mission will actively seek external sources of funding to account for additional costs.

b) Provided that the Science Objectives and Mission Objectives are preserved and that due considerationhas been given to the use of budgeted subsystem contingency and planned scheduled contingencythe mission shall pursue risk management and descope as a means to control cost.

c) Scope reductions from baseline to minimum science requirements or potential scope reductionsaffecting these requirements shall be authorized by the Principal Investigator, Systems Engineerand the Program Office.

6. Mission Implementation

6.1 Concept of Operations

The CONOPS Diagram can be found in Figure 9 below.

6.2 Mission Description — AIDA

SWARM consists of two CubeSats that will be CubeSat Opportunity Payload Intersatellite NetworkSensors (COPINS) aboard the European Space Agency’s (ESA) Asteroid Impact Mission (AIM), amission targeting binary asteroid 65803, Didymos, which in turn comprises half of the Asteroid Impact& Deflection Assessment (AIDA) mission. The other half of the AIDA mission is a NASA asteroid


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impacter called the Double Asteroid Redirection Test (DART). Table 3 summarizes the planned phasesfor each of the four spacecraft involved in the AIDA mission. Upon completion of the scientificmeasurements, the SWARM CubeSats will utilize inter-satellite links to relay all collected scientific data

back to Earth through AIM which is itself is an objective of the AIM mission.

6.3 Mission Description — SWARM

6.3.1 Launch

As shown in Figure 9 and Table 3, SWARM’s journey to the binary asteroid system 65803 Didymos begins aboard AIM’s Soyuz-Fregat launch vehicle at the Guiana Space Center. DART will be launchedafter AIM to enable the latter to characterize the asteroid and provide DART with detailed topographicaland guidance information. At the point of AIM’s arrival at the binary asteroid 65803 Didymos, itwill be approximately 0.11AU away from Earth [3] and about 35 kilometers from the surface of theDidymos system’s moonlet — Didymoon. Upon arrival at the Didymos system, AIM will performcharacterizations and measurements at 35 km and 10 km before maneuvering to within 1 km of theDidymoon surface to first deploy the SWARM CubeSats and then the MASCOT-2 lander equippedwith a low-frequency radar.

6.3.2 Deployment

Following deployment, SWARM will perform intersatellite link experiments to fulfill one of the primaryAIM objectives, after which the Imager CubeSat will be placed into a lower orbit over the Didymoonsurface at an altitude of 50 meters in order to take high-resolution pictures of the surface regolith andfeatures. The Imager CubeSat will also measure the elemental composition of the surface using usingan onboard infrared spectrometer. Together, these operations will assist AIM in constraining the initialconditions of the satellite prior to DART’s impact, which will be crucial in the generation of mass andshape models of the Didymos binary system.

6.3.3 Impact Event

DART’s trajectory has been designed such that the impact coincides with a close encounter between theDidymos system and Earth in October 2022. This is when SWARM begins performing its main missionobjectives. The Imager CubeSat, positioned such that it maintains line of sight with the impact sitewill take advantage of its proximity to Didymoon to image the collision at resolutions higher than thatwhich AIM can achieve. Measurements including crater size, depth, and velocity will enable existingcrater scaling laws to be verified in addition to enabling the determination of the momentum transferefficiency β [15]. Throughout the impact, the Imager CubeSat will observe the temperature changesat the target impact site using an onboard thermal imager. This measurement will enable planetaryscientists to determine the binary system’s thermal inertia and test the YORP spin-up hypothesis

regarding the Didymos system’s formation. In addition, the composition of the crater will be furtheranalyzed at this time to characterize subsurface elements uncovered by the impact event.

6.3.4 Mass Spectrometry

While the Imager CubeSat records various aspects of the impact event, the Diver CubeSat will proceedto fly through the impact ejecta plume to collect samples and perform mass spectrometry, revealingdetails about the asteroid’s molecular and isotopic composition. Numerical simulations performed

by Schwartz have provided initial characterization of the ejecta cloud dynamics resulting from theimpact event and contributed to SWARM’s working environment definition, thereby facilitating mission


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Table 3: AIDA Mission Overview

Date Phase Spacecraft Notes/Operations

Oct2020

Launch AIM,

SWARM Soyuz-Fregat, Guiana Space Center, Kourou

July2021 Launch DART

May2022

Arrival AIM,

SWARM- Distance from Earth: 1.1 million km (0.11 AU)- Altitude above Didymoon: ∼ 35 km

May2022

Earlycharacterization

AIM

- Optical communication experiment- Generate high-resolution 3D model of Didymos system- Determine physical properties using visual imaging system- Altitude above Didymoon: ∼35 km

June2022

Detailedcharacterization

AIM- Thermal infrared imaging- Surface and shallow subsurface sounding- Altitude above Didymoon: ∼ 10 km

15 Aug2022

CubeSatrelease

AIM,SWARM

- Imager and Diver CubeSat released

- Deployment velocity: 0.5 - 2 m/s- Altitude above Didymoon: ∼ 1 km- SWARM initial diagnostics check- Inter-satellite link experiment

22 Aug2022

Lander release AIM- MASCOT-2 lander deployed

(of MASCOT lander (Hayabusa-2) heritage)- Altitude above Didymoon: ∼1 km

AIM

- Seismic experiments- Sound deep interior structure of Didymoon- Constrain Didymoon internal structure- Altitude above Didymoon: ∼1 km

Sep

2022

Detailed

characterization SWARM(Imager)

- Characterize surface regolith and features

using high-resolution visual imaging system- Measure elemental composition of surfacethrough Near Infrared Spectroscopy (NIRS)

DART - DART mass: >= 300 kg

- Impact velocity: ∼6.1 km/s

AIM- AIM altitude above Didymoon: ∼100 km- ∼ 90 ◦,to prevent damage from impact ejecta- Capture visual and thermal images of impact event

Oct2022

Impact eventSWARM(Imager)

- Capture visual and thermal images of impact event- Crater size, velocity, temperature- Ejecta plume distribution, velocity, temperature- Determination of crater composition using NIRS

AIM- AIM altitude above Didymoon: ∼ 10 km- Measure plume ejecta evolution and temperature- Measure changes in Didymoon orbital period

SWARM(Diver)

- Locate and maneuver to low-velocity region of ejecta plume- Collect samples from ejecta plume- Perform mass spectrometry- Transmit data to AIMOct

2022 Post-impact

SWARM(Imager)

- Transmit data to AIM


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Table 4: Near-Earth Asteroids similar to SWARM’s target asteroid Didymoon

Asteroids Diameter (m) Mass (kg) Standard Gravitational Parameter (km3s−2)

2009 WM1 280 2.9 × 1010 1.93453299942 Apophis 330 4 × 1010 2.66832101955 Bennu 250 6 × 1010 4.00248

2004 FH 30 2.8× 107 0.001867824

2002 AA29 60 2.3× 108

0.015342842010 TK7 200 1.38 × 1010 0.9205704

69230 Hermes 375 6.7 × 1010 4.4694362012 DA14 45 1.3× 108 0.008672042011 XC2 86 8.8× 108 0.05870304

Didymoon 170 unknown unknown

Figure 10: Didymoon Mass Analysis

planning in terms of the positioning of the CubeSats during the impact observation phase and thesample collection phase [11]. Constraining the molecular abundances of such species as C, H, O, NCH, H+, and Si, will enable SWARM to achieve its second science objective, as defined in Section 2.1and provide scientists with the data required to answer questions regarding the origins of the Didymos

binary system and other binary systems in the Solar System. Measurement of the D/H ratio of the

plume ejecta will enable inferences to be made and hypotheses to be answered regarding the origins ofvarious celestial bodies and also the origins of Earth’s water [7].

6.4 Trajectory and Maneuver Design

SWARM will be deployed from the AIM spacecraft on August 6, 2022 at a 1 km altitude from Didymoonwith an initial velocity condition relative to AIM of 0.5-2 m/s provided by a AIM’s CubeSat deployerBecause Didymoon’s mass is currently unknown, asteroid systems of similar size and type in its vicinitywere used to estimate Didymoon’s mass. Other Near-Earth Asteroids (NEA) similar to Didymoon arelisted in Table 4 below.


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(a) STK Top View (b) STK Isometric View

Figure 11: STK Simulation

Table 5: Different approximated circular orbital parameters.

Orbital Altitude (m) Orbital Period (hrs) Orbital Velocity (m/s)

50 0.534 0.1634113.63 9.0776 0.1059500 m 16.89 0.0517

Because the mass and diameter of an astronomical body are related by approximately a cubicrelationship, a cubic regression was performed on the data shown in Figure 10. As a result of thiscubic interpolation, the mass of Didymoon was approximated to be 6.89× 109 kg. Using this valuefor the mass, the standard (two-body) gravitational parameter was calculated to be 0.459 m3/s2. Thisestimated parameter is used throughout the orbit design and analysis of the proposed mission designThe uncertainties that stem from this estimation are incorporated with additional contingency in thepropulsion and ∆V budgets. The precise impact location of the DART spacecraft is an importantparameter upon which SWARM’s trajectory is dependent. Although the DART impact location iscurrently unknown, the CubeSats’ final impact observation orbits are defined relative to the impact

crater and Didymoon’s rotation axis. It should be noted that the angular rotation rate of Didymoonabout it’s axis is estimated from gathered light curve data of the Didymos system. These parameterswill be acquired or estimated during the early and detailed characterization phase of the AIM-SWARMmission as the DART impact date approaches. As the final DART impact parameters are determinedpropulsive burns will be performed to maneuver the SWARM CubeSats into their ideal science orbitsGiven various DART impact locations, the Systems Tool Kit (STK) software package was used to analyzeSWARM’s trajectories. Figure 11 displays multiple trajectory view-points displaying important aspectscorrelated with orbital paramters and Didymoon after DART’s impact. Furtherore, the importantaspects include SWARM’s trajectory in respect to Didymoon, the impact of the crater with its plumesimulated distribution, and SWARM’s link to Aim.

As part of the trajectory analysis, the orbits’ altitude and orientation were varied in order to

optimize the coverage of the impact crater as well as the line of sight to the AIM mothership fortelecommunications. Analysis was performed to study the effect of different altitudes, inclinationangles, and longitudinal ascending node on access and coverage of the hypothetical impact locationand the AIM mothership. Table 5 displays the different approximated circular orbital altitudes, periodsand velocities.

With the Imager CubeSat capturing optical and thermal images of the crater and ejecta plume, itsorbit needs to be relatively close to the Didymoon surface at 50 meters in altitude. The Diver CubeSat,on the other hand, will perform mass spectrometry for which collection of impact ejecta samples wil

be required. Plume analysis is still inquiring further studies to improve understanding of asteroiddebris propagation and how much ejecta mass would escape Didymoon’s weak gravitational potential


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Table 6: Average access times for the Imager CubeSat.

3 Month Encounter Count 208Average Encounter Duration (hr) 3.129

Total Encounter Duration (hr) 650.739

Also taking safety into consideration, we chose CubeSat 2 to have a stationary orbit with respect to theimpact crater at a calculated 396 meters in altitude. Furthermore, multiple propulsive orbit correction

burns will be completed to ensure the maintenance of the final CubeSat orbits. Table 6 shows the accessand coverage of the Imager CubeSat with respect to the impact crater for the mission duration of threemonths. With the calculated coverage parameters, the Imager CubeSat will have more than enoughtime to collect and transmit the required science data.

6.5 Mission Operations

In terms of timing during the spacecraft mission lifetime, the following functional diagram in Figure12 outlines each component’s activities during different cycles of the mission. In terms of coveragefor the worst case scenario, SWARM will be in contact with AIM during different 25 minute periods

through the 3 month life cycle of the mission. This will allow SWARM to uplink its data to AIM to bedown-linked back to Earth. AIM will be carrying the CubeSats up until AIM arrives at the asteroidsystem, when the the Diver and Imager CubeSats are deployed. For guidance and navigation, the startrackers are powered on right before arriving at the asteroid system to begin the start up proceduresand gathering pointing knowledge.

Before launch, the transceiver is turned on to perform S band downlink tests between SWARM andAIM to confirm the spacecrafts can communicate to each other. As a result, the power is turned onto90% system of charge to power the transceiver. The transceiver uses the most power out of several othercomponents. In terms of power, the operational requirements are shown at different times dependingon the instrument running. For example, powering on the spectrometer shows an increase in power to15.7 W. In terms of thermal control, the damper and survival heaters are engaged early to keep the

temperature in the spacecraft at a temperature responsible for the components. Finally, the payloadis engaged for measuring and the ground systems are engaged with AIM in order to provide directcommands to AIM, which will then communicate with SWARM.

7. Payload System

As described in Section 6, SWARM will utilize four different observation and science payloads toclose its science requirements and ensure that SWARM contributes more than just redundancy tothe AIM mission. The Imager CubeSat will carry the visual and thermal imaging system. This suiteof remote sensing payloads will enable the Imager CubeSat to not only characterize the surface of

Didymoon before the impact, but also observe the impact event as it occurs and the resulting featureswith unrivaled fidelity due to its close proximity.The Diver CubeSat will contain two science payloads — a mass spectrometer and an IR spectrometer

The notion of flying through an ejecta plume caused by a high velocity impact to collect samplesfor molecular analysis is difficult to justify for a multi-million dollar flagship mission such as AIMHowever, with CubeSats generally being more expendable than their motherships, SWARM willsignificantly increase the scientific value of the AIDA mission, with regards especially to the NEOstakeholders shown in Figure 2.


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Figure 12: Mission Life Cycle Functional Diagram

7.1 Imager CubeSat

7.1.1 Visual Imaging System

AIM’s requirements for its visual imaging payload are outlined as follows [10]:• 1 meter spatial resolution from a few kilometers away from Didymoon’s surface• 1024 x 1024 pixel array to cover an asteroid of diameter 800 m from 1 km• 0.3 m vertical resolution for generation of Digital Terrain Model (DTM)While adequate for its general characterization purposes of the Didymos binary asteroid system

AIM’s visual imaging system will lose significant spatial resolution when observing the impact eventfrom its 100-km vantage point (see Table 3). Another issue with the AIM imaging system is thepossibility for obstruction, as was the case with the Deep Impact mission. In this mission, images ofthe crater resulting from Tempel 1’s collision were unsatisfactory due to an obstruction caused by theunexpected abundance of fine ejecta [10]. As a result, SWARM will utilize the NanoCamera C1U from

Cubesatshop. This component possesses a 2048x1536 pixel array to ensure a high resolution of thephotos. At 650 km, the camera has a resolution of 80 meters per pixel. Thus, for the science altitudeof 50 m, the camera can achieve a spatial resolution of approximately 6 mm per pixel. This spatialresolution allows for the accurate characterization of the surface regolith and ejecta plume and satisfiesthe instrument requirement in Table 5.

A high resolution photo takes more data than a lower resolution, so in this case, data volume isthe limiting factor. The storage will allow SWARM to process an image before transmitting to theAIM spacecraft to avoid corrupting the data before it is sent. Additionally, the nanocamera has theoption of PNG compression in order to reduce the size of the data packets while still maintainingsufficient detail in the image itself. This is discussed further in Finally, during image processing, the


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(a) Nanocamera Picture (b) Nanocamera Specifications

Figure 13: Nanocamera

nanocamera only uses 660 mW of power for a duration of 90 s which is relatively low compared tothe other subsystems power requirements. Imager will be taking photos of several different featuresof Didymoon including the crater, plume, and regolith. These photos will be taken at different timeperiods that will be before, during, and after DART’s impact. These photos fall in line with the overallscience objectives by providing information about how a kinetic impact on the asteroid affects itsdynamics to understand how early Solar System collisions may have caused the universe to look the

way it does. Figure 13 describes some of the specifications of the nanocamera.

7.1.2 Thermal Imager

The thermal imager provided by FLIR Tau is unmatched in the market. The FLIR Tau 2 thermal imagercamera has radiometry capabilities and an increased sensitivity of 20 mK with frame rates on the orderof 600 Hz. Additionally, the software comes with integrators that allow the camera to have directcompatibility between different camera formats. The greatest development in the software is the SmartScene Optimization, where Tau will automatically adjust the contrast, pitch, and resolution dependingon the temperature, pressure, current levels, etc. Some of the specifications of the thermal imager areanalyzed below, as well as shown in figure 14.

• Sensitivity: 20 mK at f/1.0. f/1.0 means that the lux value of the camera is 1.0 lux. Therefore, theminimum illumination rating necessary to take a video at 60 frames per second is 1 lux. Becausethe frame rate is slightly higher than average, this imager requires more light to produce anacceptable image. 20 mK means that the temperature needed to take a picture of at a distance of20 m is 270 K. The temperature range of the imager is −25◦ - 100◦C. Therefore, the temperaturefalls within the appropriate range necessary for the imager to work at Didymoon.

• Pixel Pitch: 17 µm. A pixel pitch of 6-8 micrometers is the typical standard for a thermal imagerTherefore, a pixel pitch of 17 micrometers means the pixels are very sensitive to light and thatthe image will respond to all of its surroundings. Moreover, the image quality will be greater


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because the pixel resolution will increase since the pixels are smaller but more accurate in theirinterpretation of the light source that is taken in.

• Time to Image: 4 seconds. This is a relatively quick time. Diver will be moving through the plumeat 0.015 m/s As a result, with a crater in Didymoon about 5 m wide, Diver could take as many as80 images. This is sufficient to provide detail about several different aspects of the crater and theresulting plume during an orbit. Additionally, this is assuming a conservative crater that is small

The expected plume diameter is around 20 m.

Overview Tau 640 Thermal Imager

Thermal Imager Uncooled VOx MicrobolometerFPA / Digital Video Display Formats 640 x 512

Analog Video Display Formats 640 x 480 (NTSC); 640 x 512 (PAL)1Pixel Pitch 17 µm

Spectral Band 7.5 - 13.5 µmFull Frame Rates 30 Hz (NTSC)

25 Hz (PAL)Exportable Frame Rates 7.5 Hz NTSC; 8.3 Hz PAL

Sensitivity (NEdT)


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Figure 15: IR Spectrometer material composition band

Figure 16: NASA Goddard Mini Ion-Neutron Mass Spectrometer

7.2.2 Compact Ion and Neutral Mass Spectrometer

The Heliophysics Division of the NASA Goddard Space Flight Center has developed a compact Ion andNeutral Mass Spectrometer (INMS) for in situ measurements of molecular and isotopic compositionof the ejecta plume. Displayed in Figure 16 , the INMS can detect and characterize neutrals and ionsof particular energy levels based on Time of Flight (TOF) binning. From the INMS, the data packageincludes 400 mass bins for neutrals and ions and a sampling rate as fast as 10 millisecond per frame.

This component is desirable because it is a mass spectrometer built specifically for small satellites

[16]. This instrument weighs only 560 grams, requires a nominal power of 1.6 W, and takes up about1.5 U volume. The INMS component was recently launched on the ExoCube 3U CubeSat mission inearly 2015 with another mission launch planned for 2016.

For both of the stated missions incorporating the INMS, the end user was able to specify additionalrequirements for the instrument. As originally designed, the INMS allows for the in situ measurementof ions and neutrals H, He, N, O, N2, and O2. This set of ions and neutrals were chosen by the ExoCubemission towards the global and storm-time characterization of the thermosphere and exosphereHowever, for future missions such as with SWARM, additional technology development and researchcan be performed on the current INMS to measure the ions and neutrals of interest at the Didymoonasteroid. Per the SWARM mission, this allows for the accurate determination of the deuterium-tohydrogen ratio and other species of interest that will characterize the composition and spectral class of

the Didymoon asteroid.

8. Flight System

8.1 Spacecraft Architecture

Determination of the scientific instruments to meet the mission objected was evaluated throughtrade analysis emphasizing observation, composition, kinematics, and plume sample collectionsDevelopment of six configurations, three for each Cubesat, were generated in respect to instrumentation


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Figure 17: Configuration for CubeSat Architecture trade study

Table 8: Architecture

Architecture 1 Architecture 2 Architecture 3 Architecture 4

Configuration 1 and 6 Configuration 2 and 4 Configuration 3 and 5 Configuration 1 and 5

options considered. Figure 17 display the configurations development.The following configurations are used to develop four mission architectures, each containing one

diver and one imager. Configuration 1 and 6 are the baseline configurations for which represents our baseline architecture design. The architectures were developed to meet to best meet all science, datavolume and mass expectations. Different instruments within the given modes

The overall functional diagram showing the interactions between each subsystem of the architectureis listed in Figure 19 below. The power bus provides power to each of the subsystems by drawing powerfrom the Electrical Power subsystem. The link between C&DH and Telecommunications providesdirectional information from the OBC to the antenna for pointing accuracy and allows a direct datastorage link for the commands once they are uplinked. C&DH also connects directly to the payload

instruments to provide command information directly to the payload. Finally, a link exists betweenADCS and the thruster to provide a quicker reaction to a need to change the pointing of the CubeSats

Figure 18: Pugh Evaluation


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Figure 19: Imager and Diver Functional Diagram with both payloads

Figure 20: SWARM Imager Model

8.1.1 3D Model

The 3D models are shown in figures 20 and 21.

8.2 Spacecraft Subsystems

8.2.1 Structure

The ISIS 3U CubeSat Structure was chosen due to the modular design, high TRL and aluminummaterial. The ISIS structure mass has passed and is expected to be flown in the next 12 months. Thisalso has to do with the absorptivity and how the type of material has to be a particular type of ofaluminum grade structures. The absorptivity of the aluminum structure allows for less damage due to

radiation. Additionally, the thermal range of the structure is from −40◦

− 80◦

C which is ample for thismission. The modular design allows for a simpler model design due to the availability of space withcertain components as seen in Figure 20. In terms of loading, because the Imager’s heaviest componentis the star tracker, the star tracker has been placed near the middle of the CubeSat to prevent excessiveloading on the weak points located near the edges of the inner and outer cube. For the Diver, themass spectrometer has been placed in this position. Additionally, the volume of the mass spectrometeris large and placing the spectrometer in its own cube eliminates vibrational responses from othercomponents that might affect the spectrometer’s performance.


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Figure 21: SWARM Imager Model

8.2.2 Electrical Power Subsystem

The EPS board is responsible for controlling the power subsystem as a whole by providing an interfacefor each of the components to interact. As seen in Figure 22, the EPS board is connected to the solararrays, the Power Distribution Module (PDM), and the On Board computing unit which includesthe Command and Data Handling. The board monitors the output of the solar arrays to confirm

the appropriate amount of power is being used to power the spacecraft. Moreover, the solar panelsneed to be maintained at the appropriate temperature to maintain the overall spacecraft’s power atthe peak operation power point. Additionally, the power is restricted to and from the batteries toensure sufficient energy is being stored. The EPS board used is a Clyde Space Power Management andDistribution (PMAD) module. This board provides three main advantages:

• The Clyde Space PMAD has a robust flight heritage. The PMAD was used on the ANDESITE 3UConfiguration from Boston University and AOSAT from Arizona State University. Clyde Spacehas also developed a 3U specific flight package that can handle the solar array configuration andincludes built in power tracking in addition to over-current and battery under voltage protection

• The PDM distributes power and has the ability to switch the power users on and off to preventa component from overheating. Each circuit connected to the module is protected because thepower is moved along specific power buses which are connectors that supply the power and datato each component.

• SWARM will utilize lithium ion batteries because of the high energy density relative to othersources. Unfortunately, as SWARM moves further from the Sun into extremely cold temperaturesthe batteries become unstable and can break down. Therefore, SWARM will utilize damperheaters to heat the batteries to prevent damage. Additionally, the batteries have a finite, up to2,000, number of charges. To maximize the battery life, the depth of discharge will never exceed20%. Unfortunately, SWARM cannot add more batteries to offset the redundancy requirements

because of limits to the mass and volume of the CubeSats.


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Part names Operating Temperatures Survival TemperaturesC1U nanocamera 0 60 -40 85

Star Tracker -20 40 -40 85S band transceiver -40 60 -60 80

Table 9: SWARM Thermal Analysis

(a) EPS Overall Layout (b) EPS Physical Layout

Figure 22: EPS Functional Diagrams

8.2.3 Thermal Management

For the thermal management process, the initial analysis is focused towards defining the optimal andsurvival temperatures for CubeSats. The average temperature of the spacecraft was computed byassuming the spacecraft external surfaces are the boundary conditions. From the boundary conditionsaverage power consumption and heating conditions were used to apply the law of thermodynamicsFrom the first law of thermodynamics, the heat energy coming into an isolated system must be equalto the heat energy leaving the system. The heat energy inside the CubeSat would be based on the solarconstant, the area of the solar panels, and the absorptivity of the CubeSat. The heat energy outside ofthe CubeSat would be based on the heat from the Sun’s radiation, the emissivity of the surface, and thegeometry planes. The thermal management process was analyzed by assuming the spacecraft was asingle thermal boundary. The Stefan Boltz law treats the radiation from the spacecraft as a black bodywhere the emissive power is proportional to temperature to the fourth power. This law assumes the black body radiance will be radiating energy in all directions isotropically. The temperatures calculatedfrom the Stefan-Boltz equation provide the initial state of constant temperatures throughout the CubeSatFigure /refminandmax lists the minimum and maximum survival and operating temperatures for eachcomponent to perform a thorough thermnal budget seen in Section 9.


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Figure 23: SWARM Thermal Analysis

8.2.4 Propulsion

It was determined that propulsion systems would be required in order to place the Imager and DiverCubeSats in their respective science orbits. Because of the Didymos binary system’s low gravity, theSWARM CubeSats will require very little ∆V to perform orbital corrections, as will be further discussedin Section 9.1.5. The key drivers behind the propulsion system were mass, volume, power, and risk, allof which were to be minimized. Chemical propulsion systems were ruled out from the beginning dueto the excessive ∆V provided, the high risk imposed upon AIM by the SWARM CubeSats, and the largeamount of propellant required. Likewise, electric propulsion systems such as ion thrusters produced

by Busek Co Inc. and Aerojet Rocketdyne were relegated due to relatively large power requirements onthe order of tens of Watts.

The remaining technologies under consideration were thus cold-gas systems and the nascent microelectrospray propulsion systems. A decision was made to select a cold-gas propulsion system based onthe following arguments:

• SWARM mission development phase would be too short (5 years) for adequate maturation ofCubeSat-ready micro-electrospray technologies because current developments remain in theconceptual phase.

• Cold-gas propulsion systems are inherently simpler in terms of operation compared to thecomplex microelectromechanical systems (MEMS) required by electrospray thrusters.

• Cold-gas propulsion systems have been flight-tested on CubeSat architectures [17].


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The cold-gas thruster chosen for the SWARM CubeSats is the Nanosatellite Micropropulsion Systemfrom CubeSatShop. Figure 10 below details the specifications and features of this propulsion device.

Table 10: Specifications for the CubeSatShop Nanosatellite Micropropulsion System

Nominal Thrust 100 uN to 10mN

Thrust Control 1 to 100%, 1% resolution

Impulse Duration 2ms to unlimitedSpecific Impulse 50s - 100s (warm gas option)

Pointing Resolution 0.1 arcsecMinimum total system mass 300 gram

Maximum required power 2 WattSupply Voltage 12 Volt

Operational Temperature -20◦C to +70◦C

A 1/3 U configuration was chosen for the propulsion system because of the minimal ∆V requirements

In addition to meeting the mass, power, and temperature constraints (See Section 9.1, the device is alsocapable of functioning as an attitude control system.

8.2.5 Attitude Determination & Control

The attitude determination and control system focuses on meeting the science and mission requirementswhich consists of ensuring pointing is known to within four degrees of freedom and making sure thespacecraft angular and translational state vector is known at all times. The required instruments willinclude a sensor and a control actuator. Due to IMU and gyroscope lack of absolute measurementsother instruments would be required for precision. However, due to volume constraints, star trackerswas determined to be the best device to consume volume on both CubeSats while still meeting the

mission requirements. The Miniature Integrated Star Tracker (MIST) was chosen as it poses a pointingknowledge within 4 degrees of the registration area of interest and enables determination of theCubeSat’s state vector at all times. Figure 11 provides MIST specifications qualified for SWARM’smission duration.

Table 11: Miniature Integrated Star Tracker (MIST)

Attitude Performance

Attitude Knowledge Error 30 arcsec ( 1 sigma)Update Rate 10 Hz

Slew Tolerance (no degradation) 1deg/secSlew Tolerance (w/degradation) 5deg/secTime to First Star ID < 1 sec

Mission PerformanceMission Life 2 yrs

Power Consumption < 5 WSEE Mitigation SEL, SEFI and SEU

MIST will primarily serve as the attitude determination system and secondary for the attitudecontrol system since the Nanosatellite Micropropulsion System can also provide attitude control.


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8.2.6 Telecommunications

Analyzing Table 12, the Clyde Space S band transceiver is the best option for SWARM’s purposes because this component is easily integrated with the solar array and the Power Module. Althoughthe transmit power is high and the temperature range is limited compared to other options, the hightransmit frequency will allow SWARM to downlink information at an appropriate rate. SWARM willalso be utilizing

Table 12: Transceiver

Manufacturer Clyde Space ISIS Analog Devices Inc.Temperature Range (◦C) -25 - 61 -30 - 58 -40 - 85

Mass (g)


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systems to prevent problems from occurring. An analog to digital converter (ADC) converts the voltagesto a signal that the computer can understand. The computer monitors the results from thermistorsthat measure temperature within the spacecraft, voltages that are excessively high across any givencomponent, the amount of power going to certain components, and the voltage generated by the solarpanels. Additionally, the processor is connected to the flash memory in order to store data sent fromthe payload onto a temporary storage device if downlink capabilities are not readily available. Finallythe processor connects to a data bus in order to tell the spacecraft when to uplink or downlink.

In terms of software, the primary responsibility will be to confirm the release of the antenna tomaintain appropriate communications. In addition, the software will warn SWARM when data is

being prepared to be taken, uplinked, or downlinked. Without this communication, data could be lostwhen transmitted. To prevent this from happening, SWARM’s software interface is a set from ClydeSpace that has a TRL of 8 and has been used on multiple CubeSat flights to prevent any errors from asoftware standpoint.

Figure 25: SWARM C&DH Subsystem Diagram

8.2.8 Payload Accommodations

If extra space is opened up for the launch such that the solar panels can be integrated outside of theCubeSats instead of inside, then space will open up inside the CubeSat such that more componentscould be added without restrictions in terms of volume. Additionally, SWARM will utilize a flightadapter developed by Moog Inc. that will reduce the amount of vibrations felt by SWARM during theflight since the larger AIM satellite can withstand greater vibrational loads. Additionally, SWARM willutilize propulsive adapter rings so that the payload can travel beyond the launch vehicle insertion orbitThis will allow SWARM to more quickly enter our orbit following deployment without waiting forthe power and thruster to insert SWARM. Another option being considered is to use Wafer adapterssuch as the Cubestack in order to launch SWARM as a potential tertiary payload in the case that otherpayloads are considered for the Soyuz launch. These adapters can also mount to 24 inch surfaces likein the larger ESPA Grande adapter in the case that the space for the payload increases.


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9. System Engineering

9.1 Technical Resource Budgets with Descriptions

9.1.1 Mass Budget & TRL’s

Analyzing the mass budgets between the Diver and the Imager, both have evenly distributed budgetswhere no one component is overriding the mass constraint as seen in Figure 26. Looking at figures eand f, the margin of the mass and volume for both CubeSats is positive. Although the margin for theDiver volume is small, the contingency is built into each component so that even with a small marginthe volume should still be large enough. Additionally, the margin value is mostly due to the massspectrometer. This component has a large given volume, but there are other commercial vendors thamay be able to provide another spectrometer with similar specifications besides Goddard Space FlightCenter. Another important note is that the transceiver is lighter for the Diver than the Imager becausethe Diver needs more power in order to downlink more data. As a result, the Diver has a larger modeof the Clyde Space transceiver which includes more subcomponents, and as a result, has a greater massFinally, the solar panels are the majority of the volume for the Imager because more power is neededThe solar panels will be inside of the CubeSat during the flight to Didymoon. After release, the solar

panels will unfold and a little more volume will be available with Imager.

(a) Imager Mass and Volume Budget (b) Diver Mass and Volume Budget

Figure 26: Mass and Volume Budgets

Table 13: Technology Readiness Levels

Object TRL Heritage AvailabilityMass Spectrometer 6 Dellingr Goddard Space Flight Center

Thermal Imager 3 none FLIR Systems, Inc.Nanocamera 7 PlanetLabs Innovative Solution In Space

Cold gas thruster 5 SloshSat Innovative Solution In SpaceStar Tracker 7 CSTB3 (Boeing) Blue Canyon Tech

Transceiver 6 DUSTIE (Virginia Tech) Clyde SpaceOn-board Computer 8 numerous Innovative Solution In Space

Solar Panels 7 ELFIN (Helio1, UCLA) Clyde SpaceEPS + batteries 6 InflateSAIL Clyde Space

Structure 7 ALICE Innovation Solutions in Space


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9.1.2 Power Budget

As set by current CubeSat standards and requirements, the power must be off for launch. Followingthe ejection from the fairing, SWARM’s batteries must power components until detumbling is completeand the solar panels can be deployed to begin providing power. As seen in Figure 12, the batteries cansupply idle power for up to 3 hours so that in the event that multiple pieces of hardware are in useduring this time, the power supply will still be sufficient. Another requirement for CubeSats is that thehardware is completely enclosed in a skeleton. As a result, a charging method has been established forthe batteries. A Universal Serial Bus (USB) cord and adapter can be inserted into the PDM to allow the

battery to charge up to a full 30 Watt-Hours (WH) to begin detumbling and deploy the solar panels.As noted in Figure 27, the maximum power requirement of the Imager and the Diver during

utilization of the thermal imager or mass spectrometer is about 29.7 W and 25.0 W, respectively. On theother hand, for uplinking and downlinking processes, the nominal power needed to do so is about 24.1W and 20.8 W. Therefore, to find the total energy needed for every hour to downlink data, multiplythe power by the time. With a downlink time with AIM of close to 4 hours out every 4.5 hour periodthe total energy that the spacecraft will need during downlink is about 116 W-hr. Secondly, the totalenergy that must be delivered by the batteries each orbit is around 3.49 W-hr after taking into accountthe power necessary for the payload subsystems. This is calculated from performing a calculations

based on the eclipse of the solar panels. If the solar panels are in eclipse for 10 minutes of every orbitduring this time, the batteries can store up to 3.49 W-hr. To power the spacecraft for 10 minutes onlyuses about 0.5 W-hr assuming the spacecraft only reaches nominal power mode during eclipse. Thenassuming that the batteries are only 90% efficient to be conservative (even though the ideal efficiency is98%), the total amount of energy drained from the batteries each downlink is around 3.87 W-hoursThe solar panels must replace this energy when the thermal imager or mass spectrometer are notrunning. Again assuming that there is about 10 minutes of charging in any given orbit for chargingthe batteries, then the spacecraft must supply 2.75 W to the ADCS system (Star tracker) and 0.03 W tothe transceiver, which leaves 6.2 W available to charge the batteries. Assuming a charging efficiencyof 60 %, the batteries can gain 3.24 W, which will charge about 70% of each battery. Therefore, morework needs to be done on assuring the batteries can provide full power during eclipse from the SunThe degradation of the batteries is found using equation 2 where D is the degradation per year of the batteries and L is the satellite lifetime in years.

Ld = (1− D)L (2)

The Clyde-Space 3U solar panels have an initial voltage of 15 V, and during cruise at -10 ◦C, thepanels can provide about 21.15 V. As SWARM’s heliocentric orbit radius increases, the panels can onlyprovide 25.13 V. In terms of power, this corresponds to about 32.75 W of power available after leavingEarth, and after maintaining cruise, an increase to about 36.18 W available. The panels also have anefficiency of approximately 28.3% that’s taken into account in these calculations. The necessary area ofsolar panels is shown in the calculation below with the following requirements: 1) Providing 32.75 W

during daylight and eclipse, 2) Able to withstand eclipse duration of 100 minutes, 3) Have a designlifetime of about 3 months. The following equations were used to calculate the power requirementslisted in Figure 28. The significant variable in the table is Psa . Psa is the amount of power the solar panelscan provide for the CubeSat. This value will confirm that the area of the solar panels is appropriateto provide an abundant amount of power for the components in the maximum power configurationUsing this approach, the required solar panel size for each CubeSat is about 0.32 m2. Consequently, 3Usolar panels from Clyde Space were chosen to satisy the maximum power requirements. A descriptionof the variables is given in the nomenclature section.


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(a) Imager Power Budget

(b) Diver Power Budget

Figure 27: Power Budget


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Psa = PeT e/X e + PdT d/X d

T d(3)

PEOL = PBOL Ld (4)

Asa = Psa/PEOL (5)

Figure 28: SWARM Power Variables

As noted in Figure 29 below, the ability of the solar arrays to produce power depends on the ambienttemperature. These models were created from data taken from the Lithium Thionyl Chloride batteriesModels created for other batteries with similar trendlines were also included. For a conservativeanalysis, an additional contingency of 10% was incorporated under the assumption that the actual

battery won’t produce the full amount of power calculated for this model. According to the modelat lower temperatures the solar cells operate more efficiently and produce higher peak voltages. Asa result, heat needs to be dissipated from the cells for efficient operation. Additionally, there is anoptimal point at which for a given temperature, the power is the highest. This is known as the peakpower point, and its value is about 3.7 A for the 0 degree Celsius case and 3.96 A for the -25 degreeCelsius case. As a result of this data, SWARM will utilize a front mounted two unit (2U) panel and twosingle deployed 2U panels on the sides. This will provide sufficient power for the maximum peak caseand allow the batteries to store energy during non-peak times.

Figure 29: SWARM Solar Cell Output As A Function of Temperature [18]

9.1.3 Telecommunications Link Budget

The driving requirement for the Telecommunication system is the necessity of a data rate near10 Mbps to communicate data to AIM. Additionally, the data must be reliable and require littlepower consumption to prevent overheating of batteries and loss of power from short circuits. Thetelecommunication system contains the following components:


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• Antenna: Both CubeSats have a Clyde Space CS-CPUT-STX-02 S band patch antenna optimizedfor low mass and small size requirements. The patch, monopole antenna will be fixed to the sideof the CubeSat so that it acts like a flat, thin printed circuit board with an antenna embeddedinside it. A monopole is a single radiating wire with low gain and approximately sphericaradiation profile. This is favorable for SWARM because the complexity of the relatively lowcomplexity and pointing requirements of each CubeSat.

• Transceiver: The Syrlink S band transceiver is composed of a transmitter and receiver circuitModulation is unnecessary for SWARM’s purposes because phase or frequency shifting isprimarily used for long range signal transmissions where high bandwidth and efficiency arecritical to successful communications. As a result, this design is robust and flexible because theimplementation is straightforward. As a backup, a secondary radio operating on the 2.4 GHz S

band is being considered to even further solidify SWARM’s downlink rate capability throughoutthe mission. This backup transceiver provides data rates up to 2 Mbps.

Looking at the Telecommunications budget outlined in Figure 30, the total available normalizedsignal to noise ratio (Eb/No) is positive which signifies a positive link margin. As a result, SWARMhas a greater predicted Eb/No than what is required. The orange values in the tables are values that

were manually varied in order to comply with the telecommunications requirements which state thatSWARM must maintain a 10 Mbps data rate downlink to AIM and have constant uplink or downlinkavailable with AIM. The uplink and downlink frequencies are both 2.3 GHz as a result of using an S

band transceiver. The chosen antenna for the SWARM CubeSats has a 55% efficiency and 7.6 degree2

coverage area. The transmit power requirement is a result of the transceiver choice. As a resultthe equivalent isotropic radiated power (EIRP) is about 22.4 dBW. EIRP represents the power of thetransceiver minus the losses from the output transmission and other hardware involved between thetransmitter and antenna, plus the gain. In other words, this is the amount of power that an isotropicradiator with no losses would need in order to radiate with a power flux density similar to the beampeak of the antenna. The major losses are listed under the Propagation Range section of each tableThe atmospheric losses are slightly lower than normal because Didymoon is about 1.1 AU from theSun such that the radiation environment may be slightly less disturbed due to the lack of magnetic andelectrical fields interacting with electrons. The line loss is also minuscule because the distance betweenthe transmit and receive antennas is less than 1 km as both SWARM and AIM are within 100 km ofthe binary asteroid system. Equation 6 below shows the free space loss. This loss is the largest factorin terms of reduction in power flux density resulting from geometric dispersion of electromagneticradiation. In Equation 7, C represents the total power at the receiver amplifier. Pr is the actual receivedpower. Using equation 8, the total received power comes to -78.4 dBm. Using this value to calculate thereceived Eb/No the result comes to 0.65 dB as noted in table 30 above. Finally, the total available Eb/Nois so large for both scenarios because the losses due to the atmosphere or line losses are relatively low

because of the mission. This allows for a relatively efficient number of users per channel so that data

can be successfully sent from SWARM to AIM without having interruption from channel switching bythe OBC.

Ls = (4π r

λ )2 (6)

C = EIRP + Gr − Ls − Latm − Lin − LTx (7)

Pr = PtGtGr(4π d/λ)2

(8)

EbN o

= PtGtGr

(4π d/λ)2kT sRsin2( β) =

PrN

sin2( β) (9)


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(a) Uplink Budget (b) Downlink Budget

(c) Final Characteristics

Figure 30: Telecommunications Budget


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In the case that SWARM needs to downlink information to Earth because AIM has a rupturein communications, a solid state power amplifier (SSPA) will be included to amplify power. SSPAhas a relatively low mass and volume compared to a traveling wave tube amplifier, and its powerlevels require 5 W or less at certain operational states. SWARM’s antenna was chosen knowing that arelatively small gain is able to accomplish the requirement of downlinking data at 10 Mbps to AIMThe smaller gain of the Clyde Space antenna allows for a full 3 dB beamwidth of 7 degrees and theoption of a conical beam. Additionally, this antenna can be used to reduce the amount of transmitter

power required to support emergency communications in the event that either the Imager of Diverspacecraft necessitates immediate assistance. Finally, an S band transceiver is the ideal choice for thismission because AIM has an antenna with a high gain while SWARM’s is low. Therefore, the link is RFfrequency independent to a first order. The decision of the transceiver is then driven by mass, cost, andavailability where S band provides the best options when compared with higher bands like X, Ku, andKa bands.

9.1.4 Thermal Energy Balance

A thermal budget was performed using the process outlined in [18]. The first step is calculatingdifferent parts of the mission profile as listed in Table 14 below.

Type of orbit Elliptical (e=0.7)Inclination 23 ◦

Design lifetime 3 monthsAltitude 500 m

Orientation Spinning rectangular prism (assumption)Spin Axis Perpendicular to orbit plane

Power Dissipation Goal 3 W with a 6 W peak every 20 minutesTemperature Requirements for Internal Average -15◦ - 20 ◦ C

Configuration l=10 cm, w = 10 cm, h = 30.4 cm | 25 % covered with solar cells

Table 14: Thermal Budget Parameters

The next step is to compute the thermal heating environment for each heat node at different pointsin the mission profile. This involved identifying the minimum, maximum, and average heat flux fromeach component at different time periods and averaging over the time and area of the node. Then theabsorbed environmental heat can be found on each object or surface using equation 10 below whereα is the solar source and σ is the infrared source from the IR spectrometer. The other variables aredetailed in the nomenclature section. Next, the total IR dissipation is found by summing all externalsurfaces of the spacecraft’s geometry using equation 11. Subsequently, a steady state heat balanceequation is used assuming the satellite is an isothermal object including the total power dissipated fromTable 14. The equation is 12. Finally, the temperatures for the maximum and minimum heat loads arefound using the maximum and minimum surface properties along with the internal heat dissipation

and then solving for T from equation 13. Tables 15 and 16 shows the thermal balance for the maximumand minimum thermal cases on each surface. The key to note is that the final thermal value for thecold and hot case are within the range of temperatures survivable by each component noted above.

Qenv = αS[ AP + RAR] + IRAIR (10)n

∑ n=1

n An (11)

Qenv + Qin = σ T 4

n

∑ n=1

n An (12)


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Q = ASσ T 4 (13)

Surface Num Descrip. Area (m2) Inc.

SolarInc.Albedo

Inc.Earth IR

α Absorbed

Solar + alb

AbsorbedDidymoon IR

Elec.Power

QinRad Area/abs Area

Temp

1 Zenith-Y .03 101.3 0.85 0.9 56.76 0 0 69.2 1 12 Nadir -Y .03 7.2 20.9 53.1 0.85 0.9 23.9 62.3 0.0 40.4 1 -13 Sun-P .03 0.09 18.2 20.4 0.85 0.9 3.2 14.2 0.0 19.8 1 -4 A-Sun-P .03 0.09 18.1 20.3 0.85 0.9 3.2 14.2 0.0 19.8 1 -5 Ram-R .03 99.3 7.6 17.9 0.85 0.9 56.74 14.1 0 69.1 1 1

6 A-Ram-R .03 99.3 7.6 17.9 0.85 0.9 56.74 14.1 0 69.1 1 1Total .18 307.28 93 129.6 0.85 0.9 200.54 118.9 3 287.4 6 -7

Table 15: Thermal budget for Cold Case (500 m, β = 0)

Surface Num Descrip. Area (m2) Inc.

SolarInc.Albedo

Inc.Earth IR

α Absorbed

Solar + alb

AbsorbedDidymoon IR

Elec.Power

QinRad Area/abs Area

Temp

1 Zenith-Y .03 42.3 0.85 0.9 77.5 0 0 77.5 1 15.2 Nadir -Y .03 12.4 25.1 63.6 0.85 0.9 45.8 79.3 0.0 89.5 1 -203 Sun-P .03 267.4 6.8 14.5 0.85 0.9 273.2 14.9 0.0 285.6 1 624 A-Sun-P .03 0.0 1.5 15.1 0.85 0.9 0.83 14.7 0.0 12.7 1 -245 Ram-R .03 38.8 3.9 14.7 0.85 0.9 34.1 7.5 0 42.6 1 -5.4

6 Ram-R .03 38.8 3.9 14.7 0.85 0.9 34.1 7.5 0 42.6 1 -5.4Total .18 399.7 41.2 122.6 0.85 0.9 465.53 124.0 3 550.5 6 21.

Table 16: Thermal budget for Hot Case (500 m, β = 40)

9.1.5 ∆V Bud

CubeSat Constellation Mission Design

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