IRIS Mars Communications Network - Michael Hoppermichaelhopper.net/IRIS Mars Communications...

IRIS Mars Communications Network

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Cover Image Art: Created by UCF Team. 3D Model developed by Jon Bell. Space Environment created by Michael Hopper.

University Of Central Florida

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Acknowledgements We’d like to thank everyone involved in the design of this Satellite Network. Special thanks to our Advisor from SSPI, Mr. Edward Ashford, and UCF professor, Mr. Michael Young, for their help. We’d also like to thank Cody Pike and Andrew Haines for their help in mission design.


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Abstract Mars, one of the most talked about topics of the 21st century has become humans’ next pioneering adventure in its utmost effort to becoming a space faring civilization in securing the survival of its species. Although the establishment of this colony on the red planet will not be accomplishable until 2040, it is crucial that a reliable communication infrastructure should be in place when settlers finally arrive. The purpose of this paper is to propose a feasible and reliable satellite communications network design that can provide an overall 98% continuous communications between Mars and Earth, both providing support for robotic exploration during its early years of study, colonization and early development on Mars by 2040. The design was created in response to the SSPI initiated student research and engineering competition: Satellites Around Mars. Furthermore the mission will be funded and developed by multiple nations which will also contribute terrestrial communications infrastructure to maintain the link with Mars.


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Preface Hundreds of thousands of spectators have gathered the area, taking in the beautiful morning as the sun rises over the swamps of Cape Canaveral. You could still see stars in the sky, slowly fading away as sunlight overwhelmed each one of them, effortlessly. It’s just after 7:00 in the morning of June 18th, 2020. There is a cool breeze flowing off the Atlantic’s cool waters, as children play tirelessly from the excitement of what is to come. The engineers in the Atlas Control Center run around frantically as they make final preparations for an important mission: the start of a frontier to Mars. These engineers have worked intensely over the last 24 months as they have designed, and built a critical component to our eventual colonization of Humanity on Mars. Today, they are launching the start of the much anticipated IRIS Mars Communications Network. IRIS I would be placed in orbit directly over the planned colony, individually providing connections back with Earth over 30% of each day. Soon, this would be joined by another two satellites to increase the margin to nearly 100%. With the final preparations made on the launch pad, workers begin evacuating, every one of them both excited and nervous at the same time. A few stragglers stay behind to verify their work, before running to catch up with the rest. It’s all up to launch control now. The Launch Manager begins the final launch poll: Launch Manager: “This is the NLM with the IRIS I final launch readiness poll. NASA CE?” - Chief Engineer: “Go!” Launch Manager: “NASA MM?” - Mission Manager: “Go.” Launch Manager: “SMA?” - Safety and Mission Manager: “SMA is ‘go’.” Launch Manager: “SMD?” - Spacecraft Mission Director: “Go.” Launch Manager: “NAM?” - NASA Advisory Manager: “NASA Advisory Team is ready.” Launch Manager: “Copy that. NASA Team is ready to Proceed.” With that, the Launch Manager begins the final countdown. Spacecraft power is initiated, fuel tanks are pressurized, and the launch control system is enabled.

Launch Manager: “This is Atlas Mission Control at T minus 10 … 9 … 8 … 7 … 6 … 5 … 4 … 3” Sparks appear below the Atlas V rocket as the Main Engine is ignited. A great flame appears before a large cloud of steam quickly consumes it along with the rest of the launch pad.

Launch Manager: “… 2 … 1 … 0 … And we have liftoff of the Atlas V rocket with the IRIS I payload of the IRIS Mars Communications Network. Extending satellite communications in anticipation of future colonies on the surface of Mars.”

The Atlas V rocket quickly accelerates into the sky as the large crowd of spectators marvel over how such a large object is accelerating away from them so effortlessly, yet with so much power. They can not only hear the powerful engine beneath the rocket, but feel it in their bones. An experience they would never soon forget. The rocket accelerates. Faster and faster it goes. The onboard satellite enduring many times the strength of gravity, that of which no human could endure. Reaching the upper limits of the Earth’s atmosphere, it begins to adjust its path for parking orbit insertion that of which would be used to slingshot itself away from Earth toward its destination: Mars.


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The giant flame behind the rocket begins to subside as the main engine consumes the last of its fuel. Its first stage was nearing completion. An explosive charge goes off, freeing the booster of grip from the rest of the spacecraft as gravity takes over. Pulling to the surface, the booster eventually succumbs to the new dominant force and begins its descent back down to the surface. Free of the lost weight, a second engine kicks in to finish the job. It would take just over an hour to make a complete revolution around the Earth. Having achieved maximum velocity, the second booster completes its burn and releases the satellite. IRIS I is now moving at nearly 22,000mph. Even at this velocity, it will still take over 7 months to reach Mars. The satellite adjusts its orientation to allow for communication through a low gain antenna, and prepares for the long journey ahead.


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Table of Contents Acronyms ........................................................................................................................................ 9

Mission .......................................................................................................................................... 10

Robotic exploration (mission 1) ................................................................................................ 11

Human colonization (mission 2) ................................................................................................ 12

Business plan ................................................................................................................................. 13

Funding ...................................................................................................................................... 13

Cost Estimating Relationships ................................................................................................... 14

Learning Curve........................................................................................................................... 14

Total Cost ................................................................................................................................... 14

Launch Vehicle .............................................................................................................................. 15

Heritage (Atlas V) ...................................................................................................................... 15

Fairing, fairing dimensions and spacecraft integration ............................................................ 15

Parking orbits ............................................................................................................................ 15

Orbital Paths ................................................................................................................................. 16

Basic Space Dynamics Theory ................................................................................................... 16

Overview ................................................................................................................................... 17

Phase I .................................................................................................................................... 18

Phase II ................................................................................................................................... 20

Phase III .................................................................................................................................. 22

Phase IV ................................................................................................................................. 22

First Order Approximation and Future Modeling ..................................................................... 23

Structure ....................................................................................................................................... 24

Chassis ....................................................................................................................................... 24

Materials Selection .................................................................................................................... 25

Static Analysis ............................................................................................................................ 25

Propulsions ................................................................................................................................ 26

Attitude Control Thrusters ........................................................................................................ 30

Thermal Control ............................................................................................................................ 33

Heat Transfer Theory ................................................................................................................ 33

Multilayer Insulation ................................................................................................................. 35


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Optical Solar reflectors .............................................................................................................. 36

Ammonia piping thermal - Control System / Thermo-Electric Control .................................... 37

Electrical Power system ................................................................................................................ 38

Solar Panels ............................................................................................................................... 38

Chosen Solar Panel Characteristics ........................................................................................... 40

Solar Array Retention Mechanism ............................................................................................ 42

Solar Array Drive Mechanism .................................................................................................... 44

Batteries .................................................................................................................................... 45

Satellite Subsystems .................................................................................................................. 47

Guidance Navigation and Control ................................................................................................. 48

Computer .................................................................................................................................. 48

Star Tracker ............................................................................................................................... 48

Sun Sensor ................................................................................................................................. 48

Mars Tracker ............................................................................................................................. 49

IMU ............................................................................................................................................ 49

Reaction Wheels ........................................................................................................................ 49

Communications System .............................................................................................................. 50

Satellite Constellation Arrangement ......................................................................................... 50

Communication Windows ......................................................................................................... 51

System Capabilities ................................................................................................................... 54

Antennas ................................................................................................................................... 54

Satellite-to-Earth High Gain Antenna (HGA1) ....................................................................... 55

Satellite-to-Earth Low Gain Antenna (LGA) ........................................................................... 55

Satellite-to-Satellite High Gain Antenna (HGA2 & HGA3) ..................................................... 56

Satellite-to-Mars Medium Gain Antenna (MGA) .................................................................. 56

Earth Ground Station(s) ............................................................................................................ 58

Deep Space Network ............................................................................................................. 58

Communications Subsystem Components ............................................................................... 59

Command & Data Handling ................................................................................................... 61

TWTAs .................................................................................................................................... 61

X and Ka Band Transponders (SDST) ..................................................................................... 63


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Electra UHF Transponder (EUT) ............................................................................................. 63

BandPass Filters (BPF)............................................................................................................ 64

Couplers (CP) ......................................................................................................................... 64

Diplexers (DX) ........................................................................................................................ 64

Isolators (IS) ........................................................................................................................... 64

Switches (S) ............................................................................................................................ 64

Multiplexers (MUX) ............................................................................................................... 65

Ultra-Stable Oscillators (USO) ............................................................................................... 65

Link Budgets .............................................................................................................................. 66

Signal Generation ...................................................................................................................... 67

Polarization ............................................................................................................................ 67

Modulation ............................................................................................................................ 67

Coding .................................................................................................................................... 68

Duplexing ............................................................................................................................... 68

Conclusion ..................................................................................................................................... 69

Appendix ....................................................................................................................................... 70

Solar Panel Calculations ............................................................................................................ 75

Ion Propulsion Station Keeping ................................................................................................. 77

Station Keeping ......................................................................................................................... 80

Reference Documents ............................................................................................................... 83

References .................................................................................................................................... 96

Additional References ................................................................................................................... 97


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Acronyms AMO – Areostationary Mars Orbit

BPF – Band Pass Filter CG – Center of Gravity CP – Coupler DX – Diplexer ETC – Electric Thermal Controller EUT – Electra UHF Transponder GEO – Geostationary Earth Orbit HGA – High Gain Antenna IS – Isolator/Circular Wave Guide LEO – Low Earth Orbit LGA – Low Gain Antenna MEO – Mid Earth Orbit MGA – Medium Gain Antenna MOC – Mars Orbit Capture MOI – Mars Orbit Insertion MRO – Mars Reconnaissance Orbiter MUX – Multiplexer S – Switch SDST – Small Deep Space Transponder

SOI – Sphere of Influence TCS – Thermal Control System

TOF – Time of Flight TWTA – Travelling Wave Tube Amplifier USO – Ultra Stable Oscillator VCHP – Variable Conductance Heat Pipe


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Mission The goal of this mission is to create a satellite communications system that is able to support exploration, colonization and early development of Mars, including both surface to space communications and Mars to earth communications. The first part of the Mission involves initial exploration of Mars to be conducted by robots in which they will assess environmental conditions, availability of resources and potential colonization sites. It is expected that this part of the mission is to last approximately 10 years during which the satellite will provide a minimum of 6 hours of communications with Earth per robotic lander or about 25% of the sidereal day. After the robotic mission, a dispatch of habitat units assembled in Earth orbit will be flown robotically to Mars, then landed. Once operation of the habitats are confirmed to be working in ideal conditions, a human fleet will travel from Earth to Mars and settle in the location the habitats where placed in. However, before the human colonists reach the surface of Mars the communication system will provide connectivity just about 98% percent of the sidereal day. This part of the mission will begin 20 years from the introduction of the first communication satellite.


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Robotic exploration (mission 1)

The beginning ten years of the mission involves proving communication from Mars to Earth for early robotic exploration. Moreover, the mission requires an overall 25% connectivity with earth, thus assessing the problem a proposed medium size satellite with an estimated wet mass of 1200 kg will be place in a areostationary orbit at around 17,000 km altitude. Named after the messenger Greek goddess, IRIS-1 will be the first generation satellite to be placed in orbit around Mars not only proving high bit rate communications for early robotic exploration but also establishing itself as the first of the constellation. Positioned just above Meridiani Planum, located just about 2 degrees south of Mars equator the satellite will be stationed at 0ºN -13.5ºW [Fig.1] This location was chosen to due to its high geological scientific value, geographical advantage, and low orbit station keeping of around 10 m/s yearly burn. IRIS-1 will also be design to last approximately 20 years.

Figure 1: Meridiani Planum

IRIS-1 will have the most advance deep space communication subsystem ever attempted, featuring a large 3.5 meter high gain antenna for communication with Earth, two .5 meter antennas for communication with satellite constellation and a .5 meter antenna for Mars surface communication. Furthermore, IRIS-1 is able to transmit a variety of communication frequencies including UHF, X-band and Ka-band at rates up to 24 Mbps at its farthest distance of 400 million kilometers between Earth and Mars. However, at closer distances bit rates of 30 Mbps or higher can be accomplish. In addition, IRIS-1 will communicate with any of the 34 meter Deep Space Network Antennas on earth this will allow for 6 or more hours of communications a day. A detailed communication overview will be discussed later on this paper. Please refer to the Communications section for further information on communication capabilities and parameters.


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Human colonization (mission 2)

Human colonization on Mars will begin 20 years from the establishment of IRIS-1 over Meridiani Planum, and thus an overall 98% connectivity with earth is crucial for mission success once the colonist arrive. Consequently two more similar satellites IRIS-2 ( 0˚N -132.5˚W) and IRIS-3 (0˚N 107.5˚E) will join IRIS-1 to ultimately compose a communication constellation that can cover a large area of the planet while in areostationary orbit Fig [2]. Furthermore, it was essential that both IRIS-2 and IRIS-3 be placed in fairly stable parking orbits, while optimizing line of view of one another. Thus their station keeping requires roughly 25 m/s yearly burn. Though, their station keeping yearly burns are fairly higher than IRIS-1 it is important to understand that these satellites beginning of life (BOL) is somewhat 2 or 3 years before the colonist arrive thus ensuring fuel for their 20 year mission life.

Assuming the colonists settle within the vicinity of Meridiani Planum it is crucial to ensure almost constant communication with Earth for mission success, accordingly the IRIS constellation will utilize relaying information to one another when not in line of view of Earth. This will be done in one of two ways. The first way is a direct communication with earth, this will entitle IRIS-1 being in line of sight of earth thus the communication signal will go from Mars surface to Satellite then to Earth DSN. The second way will be during the time IRIS-1 is not in the line of sight of Earth, hence during this time IRIS-1 will then relay the signal over to either of the two other satellites who do have line of sight. This process will ensure 98% communication between the planets allowing for a 2% error due to black outs and white outs to interferences of the sun. Yet, in order to maintain almost continuous communication IRIS-1 is schedule to be replace by IRIS-4 due to the circumstance that IRIS-1 will be reaching end of life (EOL) once colonist begin to arrive to Mars.

Figure 2: IRIS Communication Constellation


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Business plan

Funding

Due to the scope of the mission proposed in this paper, an international effort to fund this project would definitely be necessary in order to make a stable and practical communications relay with our neighboring planet, Mars.

Cost Component Parameter, X

(Unit)

Input Data

X

RDT&E CER'

(FY00$K)

RDT&E CER'

(FY15$K)

Parameter, Y

(Unit)

Input

Data Y

TFU CER'

(FY00$K)

TFU CER'

(FY15$K)

1.

Communications

Comm.

Subsystem weight (kg)

$

394.00

$

139,200.20

$

190,147.47

comm.

Subsystem wt. (kg)

$

394.00

$

55,160.00

$

75,348.56

2. Spacecraft spacecraft dry

wt. (kg)

$

749.30

$

75,679.30

$

103,377.92

spacecraft dry

wt. (kg)

$

749.30

$

32,219.90

$

44,012.38

2.1 Structure structure wt. (kg) $

322.31 $

18,956.88 $

25,895.10 structure wt.

(kg) $

322.31 $

4,222.26 $

5,767.61

2.2.1 Thermal thermal wt. (kg) $

43.90

$

4,349.72

$

5,941.72

thermal wt.

(kg)

$

43.90

$

733.45

$

1,001.90

2.2.2 Thermal spacecraft wt. + payload wt. (kg)

$ 1,143.00

$ 8,453.44

$ 11,547.40

N/A N/A N/A N/A

2.3.1 Electrical

Pwer System EPS wt. kg

$

232.17

$

14,557.06

$

19,884.94 EPS wt.

$

232.17

$

7,150.36

$

9,767.40

2.3.2 Electrical

Power System BOL Power

$

1,836.00

$

26,809.40

$

36,621.64 N/A N/A N/A N/A

2.4 Telemetry,

Tracking, and

Command

TT&C/ DH wt.

(kg)

$

23.93

$

6,106.21

$

8,341.09

TT&C/DH

wt. (kg)

$

23.93

$

3,854.90

$

5,265.79

2.5 Attitude

Determination Weight. (kg)

$

57.24

$

15,503.99

$

21,178.45

ACDS wt.

(kg)

$

57.24

$

6,801.43

$

9,290.75

3.1 Integration,

Assembly & Test

spacecraft + payload total

RDT&E cost

(FY00$K)

$

214,879.50

$

47,188.09

$

64,458.93

spacecraft bus wt.

payload wt.

(kg)

$

214,879.50

$

2,234,746.80

$

3,052,664.13

4. Program Level

spacecraft +

payload total

RDT&E cost (FY00$K)

$

214,879.50

$

59,882.00

$

81,798.81

spacecraft +

payload total

recurring cost (FY00$K)

$ 87,379.

90

$

29,796.55

$

40,702.08

5. Ground Support Equipment (GSE)

spacecraft +

payload total RDT&E cost

(FY00$K)

$ 214,879.50

$ 24,545.62

$ 33,529.31

N/A N/A N/A N/A

6. Launch &

Orbital Operations

Support (LOOS)

N/A N/A N/A N/A

spacecraft

bus + payload

wt. (kg)

$

1,143.3

0

$

5,602.17

$

7,652.56

Total $

602,722.79

$ 3,251,473.17


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Cost Estimating Relationships

Using a recommended cost estimating relationship (CER) outlined in a reference book for space mission analysis and design (1), the chart above was populated with excel and the recommended functions for a preliminary approximation. As can be extrapolated from the data, the research, design, testing, and experimentation phase of the IRIS mission was approximated to cost $602,722,790, an acceptable amount due to the first-of-a-kind nature of the space mission; which requires state of the art systems in: solar array system to counteract the reduced solar flux, propulsion system to stabilize station keeping, and communications system to counter the increased free space path loss for communications system. These innovations necessary to push the boundaries of what feasible technology can accomplish is further reflected in the theoretical first unit (FTU) cost estimating relationship. The CER yielded an approximate FTU of $3.251 billion dollars; however, costs are expected to go down for the following two units since they will be built in rapid succession after this thus yielding the benefits of the learning curve.

Learning Curve

The learning curve is a technique for accounting the productivity improvements accompanying the manufacture of a larger quantity of units. It can be represented by a mathematical model that includes cost reductions due to economies of scale, set up time, and human learning as the number of units increased. The total production cost for 𝑁 amount of units can be approximated as:

𝑃𝑟𝑜𝑑𝑢𝑐𝑡𝑖𝑜𝑛 𝐶𝑜𝑠𝑡 = 𝑇𝐹𝑈 ∙ 𝐿 Where:

𝐿 ≡ 𝑁𝐵

𝐵 ≡ 1 −ln((100%) /𝑆)

ln 2

Since the amount of units produced is less than 10; a recommended 𝐵 of 0.926 and 𝑆 = 95% was used. This yielded the following cost for units after the FTU:

Unit Number Cost of Unit

1 $ 3,251,473.17

2 $ 3,010,864.16

3 $ 2,788,060.21

Total Cost

Adding the cost of all the units along with the RDT&E calculated from the previous section, the total cost for the IRIS communication mission is approximated to be $9,653,120.33.


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Launch Vehicle

Heritage (Atlas V)

The launching vehicle for IRIS will be the Atlas V Heritage rocket, which was responsible for many successful launches. Such successful launches include Juno, MSL, MRO, MAVEN, and New Horizons. Given its extremely high success rate, the Atlas V has proven to be a reliable launch vehicle based on its previous missions. It is estimated that the cost to launch an Atlas V would be around $223 million and can vary significantly depending on the size and weight of the payload.

Fairing, fairing dimensions and spacecraft integration

A 5-m Short Fairing with a cylindrical cargo space of 5888.2 mm high with a diameter of 4572 mm. The Atlas V 5-m short, medium and long payload fairings were developed along with the increased launch vehicle performance to accommodate evolving spacecraft requirements. The 5-m PLF is composed of a composite structure made from sandwich panels with carbon fiber face sheets and a vented aluminum honeycomb core.

Figure 3: Fairing Dimensions

Parking orbits

Based on the data on the Atlas V below, any of the Atlas V models will suffice in providing IRIS with a LEO (Low Earth Orbit) at a distance of 6578.1 km.


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Orbital Paths

Basic Space Dynamics Theory

The theory behind the calculations for this section of the paper have been derived from simplified orbital mechanics equations with the following assumptions: two dimensional plane, coplanar, two body problem, and instantaneous velocity change. For calculating the velocity of an orbit at a distance from a large body, the velocity term is isolated from the kinetic energy equation in a simplified energy balanced assumed to consist of kinetic energy and gravitational potential energy:

𝐾𝑖𝑛𝑒𝑡𝑖𝑐 𝐸𝑛𝑒𝑟𝑔𝑦 + 𝑃𝑜𝑡𝑒𝑛𝑡𝑖𝑎𝑙 𝐸𝑛𝑒𝑟𝑔𝑦 = 𝑇𝑜𝑡𝑎𝑙 𝐸𝑛𝑒𝑟𝑔𝑦 The left-hand side above equation when written in terms of mass, velocity, and the gravitational

constant 𝐺 = 6.674 × 10 − 11 𝑁 ⋅ 𝑚2/𝑘𝑔2. The above equation yields:

1

2𝑚𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡𝑣2 +

𝐺𝑚𝑒𝑎𝑟𝑡ℎ𝑚𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡

𝑟2= 𝑇𝑜𝑡𝑎𝑙 𝐸𝑛𝑒𝑟𝑔𝑦

Where 𝑟 represents the distance between the centers of gravity of the bodies, 𝑚𝑒𝑎𝑟𝑡ℎ is the mass of the earth, 𝑚𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡 is the mass of the spacecraft, and 𝑣 is the scalar magnitude of the

spacecraft’s velocity vector. If the above equation of energy balance is divided by the mass of the spacecraft and the term 𝐺𝑚𝑒𝑎𝑟𝑡ℎ is replaced by the gravitational parameter,

𝜇 = 𝐺𝑚𝑒𝑎𝑟𝑡ℎ The following equation is derived that can relate the spacecraft’s velocity with the altitude of the spacecraft and the semi major axis of an elliptical orbit 𝑎.

𝑣2

2−

𝜇

𝑟= −

𝜇

2𝑎

With a bit of algebraic manipulation the magnitude of the velocity of the spacecraft can be rewritten as:

𝑣 = √2𝜇

𝑟−

𝜇

𝑎


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The period of an orbit can be calculated as:

𝑃 = 2𝜋√𝑎3

𝜇

As for a plane change necessary to align the orbital planes of two planets, the following equation for inclination change of a circular orbit can be used.

∆𝑣𝑖 = 2𝑣 sin (∆𝑖

2)

Where ∆𝑖 the change in inclination is angle, and 𝑣 is the orbital velocity of the orbiting body. The Equation for an Oberth Escape velocity is a function of parking velocity:

𝑉𝑒𝑠𝑐𝑎𝑝𝑒 = 212 ∙ 𝑉𝑝𝑎𝑟𝑘𝑖𝑛𝑔

The oberth escape velocity is used in conjunction with the heliocentric infinity velocity in order to get the required injection velocity:

𝑉𝑖𝑛𝑗𝑒𝑐𝑡𝑖𝑜𝑛 = √ 𝑉𝑖𝑛𝑗𝑒𝑐𝑡𝑖𝑜𝑛2 + 𝑉𝑖𝑛𝑗𝑒𝑐𝑡𝑖𝑜𝑛

2

Overall, the methodology for sending and placing a spacecraft in an orbit around mars requires several orbital maneuvers, most notably launch insertion, hohmann transfers, oberth escape maneuver, and planetary capture.

Overview

The means to achieve orbital insertion of the satellite communications spacecraft into mars areostationary orbit was divided into 4 main phases that highlight major events in terms of burn maneuvers: launch vehicle insertion into circular mars coplanar parking orbit, oberth earth escape maneuver into heliocentric mars hohmann transfer, mars rendezvous planetary capture burn, and a two-burn hohmann transfer from Martian sphere of influence (SOI) into the lower altitude areostationary orbit. Each phase has a critical goal and purpose in order to ensure the overall success of proper placement of the satellite in an optimal position. Certain elements of the mission profile have been purposefully selected with the intention of duplicating proven methodology and features of the MRO orbital insertion mission. Each satellite will be inserted into its proper orbit in the same sequence of maneuvers, the only thing differentiating each satellite from another is time of launch in order to achieve insertion into the proper longitude on mars. Below, the different phases are described in further detail with velocity budget and timing derived from a simplified two body coplanar, conic sections first order approximation calculation attached in the appendix section near the end of this paper.


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Phase I

The main purpose of this phase is to relocate the satellite from a stationary point on the surface of the earth into a purposeful parking orbit that was implemented in order to allow a place in the mission for attitude and altitude corrections before the critical large burn maneuver during the following phase of the mission. This will be achieved using the first stage of the Atlas V launch vehicle launched from Kennedy Space Center in Cape Canaveral.

Figure 4: Atlas V launch from Cape Canaveral [1]

Throughout the placement of the spacecraft into the parking orbit, the first stage of the Atlas V is still attached to the second stage and is used for multiple impulse burns. Taking into consideration the angular velocity of earth and assuming a launch from near the equator, the Atlas V was calculated to require 7553 m/s of 𝛥𝑉 for a 2 impulse burn maneuver in order to place the satellite into a circular orbit of 185 km. This circular parking orbit is assumed to be coplanar with the 1.85 degree plane change between earth’s heliocentric orbit and mar’s heliocentric orbit.

Figure 5: Simulated picture of Atlas V on ascent towards Parking Orbit


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During ascent, the fairings protecting the spacecraft from the turbulent winds will break apart after a sufficient height has been reached and is no longer necessary.

Figure 6: Simulated Atlas V with Fairings Ejected in stable orbit


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Phase II

This phase is perhaps the most critical and demanding phase of the mission with the main sequence of actions consisting of: alignment of spacecraft into insertion plane, single impulse burn into oberth escape trajectory, stage separation of the centaur stage from the main spacecraft, deployment of both solar panels, and testing of communication systems along with other system checks.

Figure 7: Separation of the Centaur stage from the Atlas V first stage (top and bottom)

Once the satellite was inserted into the parking orbit, the orientation and position will be closely monitored in order to align IRIS with the mars transfer plane for the following step. During this


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step certain correctional burns may be implemented in order to ensure the operational orbit meets the proper tolerances. Mars heliocentric planes a 1.85° relative to Earth’s heliocentric

plane. The amount of delta V need for a plane change would be 251.6 𝑚

𝑠 and would be supplied

by the centaur stage of the launch vehicle.

Figure 8: Parking Orbit of IRIS spacecraft before Oberth Insertion

After the preplanned conditions have been met for alignment, a single impulse burn will be implemented in order to begin the oberth escape maneuver. This single burn will output

3613.192 𝑚

𝑠 of ∆𝑣. With the spacecraft finally achieving the required kinetic energy to escape

the earth’s sphere of influence and on its way to mars, the centaur stage is no longer needed.

Figure 9: Centaur stage performing Oberth Escape Maneuver before Separation Maneuver


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Sequentially, the centaur will separate after all the fuel is used up; at which point, the Centaur will perform a maneuver to move away from the spacecraft in order to prevent collision at a further point in the mission, otherwise the upper stage would follow the satellite all the way to mars. [2] After the safety of the spacecraft is confirmed, the deployment sequence of the solar panels at the sides of the spacecraft will begin, described in better detail in the solar array retention mechanism (SRAM) section of the report. After the folded up panels reach their maximum extension, the solar cells will be gimbaled into the orientation with max solar irradiance. In order to begin the next step of the phase, the spacecraft will allocate electrical power to critical subsystems and the onboard battery. Once enough power is generated to charge the battery the next step of the phase can begin, testing of hardware performance and function. Critical systems such as the computer, telecommunications transponder, and gimbal systems will be tested in order to inquire whether all pieces of the spacecraft survived the violent launch into deep space. Besides routine system checks all that is left for the spacecraft to do is coast its way to mars.

Phase III

During this phase, the spacecraft will perform a planetary capture, which roughly translates into a single impulse burn during entry of the Martian SOI. The reason why this burn is important is that upon entering the region occupied by mar’s heliocentric orbit, the spacecraft will have a slower velocity due the fact that it is coming in from an elliptical transfer orbit. The valence velocity necessary for the spacecraft to “catch up” with the heliocentric orbital velocity of mars will need to be properly compensated in order to avoid a hyperbolic flyby. The single impulse burn necessary for planetary capture has been calculated to be 2664.291 𝑚/𝑠 . This burn is approximated to be instantaneous and will guarantee that the spacecraft stays in a stable orbit around mars for further orbital maneuvers. It is necessary to have this operational orbit as a transitional mechanism into the next phase of the mission.

Phase IV

This is the phase that is comprised of the final stages of the orbital insertion mission before a stable areostationary orbit is attained. The sequence of main events is as follows: proper alignment of spacecraft in mars longitudinal coordinate system, first burn into transitional elliptical orbit, second burn at periapsis, and final maneuver of orientation and positioning of satellite. When the spacecraft reaches the sphere of influence of mars and achieves a stable circular orbit at the edge of that sphere, this will provide a time in the mission in order to calibrate the point


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of entry of the spacecraft into the next step of the phase in order to achieve the desired longitudinal coordinate at the areostationary orbit. This would also be the time where the spacecraft would make any orbital corrections in terms of inclination or roundness of orbit. Unsurprisingly, the next step in the mission would be to actually perform the first burn in order to achieve the elliptical transfer orbit for the first stage of the hohmann transfer. The required

∆𝑣 would be 207.468 𝑚

𝑠 in the decelerating direction in order to lower the perigee to 17000 km.

In order for the burn to be considered instantaneous relative to the orbital period of the transfer orbit is… After the burn is completed using the primary onboard hydrazine thruster, the spacecraft will coast to the desired periapsis height of 17000 km for areostationary orbit. When the spacecraft reaches the areostationary height, a second burn in the decelerating direction is performed in order to lower the apoapsis and circularize the final orbit of the satellite.

The second burn of 626.193 𝑚

𝑠 will be performed and is assumed to be instantaneous.

After the main burns are completed, all that is left to do for the spacecraft is to calibrate and align the spacecraft into the proper orbital elements for a stable areostationary orbit. This is the final steps for the orbital insertion of the mars satellite, and all that is left to do afterwards is routine station keeping and attitude corrections throughout the mission life.

First Order Approximation and Future Modeling

Naturally, the numbers reflected in the preceding section overviewing the main sequences of the orbital insertion of the spacecraft are not accurate enough to ensure success of the mission due to the amount of assumptions and simplifications used. Furthermore, during the event of proper funding and realistic deadlines, a more precise computer simulated model, generated from a program such as systems toolkit, would be desired to more accurately determine launch windows, timing sequence, and arrival times.


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Structure

Chassis

The main goal in designing the chassis for the satellite was that the size had to be somewhat minimalistic, yet fully functional and spacious for various components. The main design feature of the chassis stems from the current designs of the MAVEN (Mars Atmosphere and Volatile EvolutioN Mission) satellite. MAVEN was designed to have a cube chassis along with an "X-shaped" internal skeleton. This provides a strong structure, yet spacious areas for its components. Seeing how this is an efficient design, its idea was adopted for the design of IRIS with dimensions of 2.5m-by-2.5m-by-2.75m. The main antenna, measuring 3.5m in diameter, is mounted on an armature assembly that allows the antenna to rotate with respect to the chassis and have a constant view of Earth. Two additional 0.5m diameter antennas that communicate with surrounding satellites are mounted on the side panels that will be attached to the chassis to protect the internal components. One 0.4m diameter antenna is mounted at the base of the chassis and has a clear view of Mars via a cut-out section which allows the antenna to have a lower profile. Finally, the middle section of the chassis will serve as a fuel tank.


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Materials Selection

After conducting research about modern day satellite designs, it was found that many current satellites utilize lightweight aluminum with honeycomb structures. For IRIS, 5052 Honeycomb Aluminum with a mass-density of 4.5 lb/ft3 was chosen for the main body. With this material, the weight of the chassis is estimated to be 198.35kg. [3]

Static Analysis

A static analysis was conducted using Solidworks 2014. Unfortunately, Solidworks does not have the 5052 Honeycomb Aluminum in its database, so a temporary material property was created instead. The results from the test are displayed below, however, they should not be read as actual real-life results. The test was conducted using 10g's of gravitational force, which produced an average stress of 95.69kPa on the chassis. A more proper CAD design with actual connections will be required in order to display realistic high stress areas. However, this simple design can illustrate the simple average stresses that can accumulate.


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Propulsions

The main propulsion system on the IRIS series spacecraft will execute Mars orbit capture (MOC) and Mars orbit insertion MOI. Thus, it is critical that the main engine be capable of performing the overall ΔV of 2,664 m/s for capture and 833.7 m/s for orbit insertion. Subsequently giving an overall 3,497 m/s ΔV burn total for both burn maneuvers. Likewise, by knowing the overall

estimated Wet mass of the satellite at 1200 kg and burn time being an overall .01% of the overall time of flight (TOF) at just about 2.8 hour. Thus by knowing the above parameters it was calculated [Eq. 1] that the required thrust be a minimum of 419.724 N.

Therefore, the engine chosen to carry out any significant burn maneuvers will be a 4.30 kg Bi-Propellant Apogee motor by Airbus defense and Space, model # S400-15 [Figure 10]. Furthermore, the Apogee engine is capable of producing an incredible 425 N nominal thrust with an Isp of 321s and nozzle ratio of 300:1. This nozzle ratio was chosen due to its increase in Isp thus a decrease in fuel consumption of the satellite during its main propulsion maneuvers. The engine uses a single coil solenoid, and single seat valve that requires 35 watts to actuate. This allows for the engine to be restarted and turned off when needed. In addition the engine is held by a thrust frame with 4 right angle titanium (6Al/4V) brackets with 16 titanium bolts which hold the engine in place [Figure 11], as well as a nozzle heat shield for adequate heat transfer. Moreover, the propellant of choice that was selected for the mission was a mixed ratio of the fuel unsymmetrical dimethylhydrazine (UDMH) and nitrogen tetroxide as the oxidizer (NTO). This fuel UDMH was selected not only due to its storability properties but also to its low cost and low density, ultimately allowing a decrease on the satellites fuel mass. [4]

Eq. 1


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Figure 11 - Thrust frame

It was crucial that the satellite needed to be as light as possible thus it was decided that two bladder propellant tanks with a blown-down configuration in which the propellant is loaded in a locked up configuration with specified gas was chosen due to the fact that with this mode additional gas pressurant vessels could be avoided. This would ultimately reduce mass, volume, system complexity and the need for any pumps. Some other benefits that would help in the overall mission life are that of the elimination of gas/propellant mixing which overall assures predictable thrust conditions when thrusters need to be reignited. However, because the tanks are pre-pressurized their overall life expectancy is low due to loss of pressure over time. Therefore the main fuel tanks will only be used for MOC and MOI, and any remaining fuel will be used as soon as possible. Also, in order to maintain inertial properties of the Satellite, the fuel

Bracket x 4 Bolts x16 Mass

.04m W x .488m L x .005m T .01m D x .025m L x.01m T .44343 kg

Figure 10 - Main Engine


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tanks for UDMH and NTO would be placed at the center of the satellite or its center of gravity (CG) [Figure 12]. [5], [6]

Propellant is a commodity in space and so it is for the most part the most important for mission BOL and EOL. Subsequently, the satellite is equip with the ample amount of propellant needed for all required burn maneuvers for MOC, and MOI as well as parking orbit station keeping which will be discussed in the next section. It was particularly crucial that the satellite’s mass of propellant be calculated to estimate the overall wet mass of the satellite thus the mass of NTO, and UDMH were found by tacking the overall estimated wet mass of the satellite then dividing it by e raised to the power of the total change in velocity for orbit maneuvers, over the Isp of the engine times the force of gravity of earth [Eq. 2]. Once the mass of the fuel was found which approximately 400 kg it was then partially divided into a 1:1 ratio of fuel to oxidizer mixture. In addition the masses where then divided by their densities giving in an overall volume of the specified propellant tank size. [Figure 13] give an approximation of the mass of the fuel and oxidizer, as well as their required tank sizes and mass [Figure 14]. Nevertheless, understand that these are industry standard propellant tanks which fall in the range of the volume necessary for the total propellant mass and so their mass might not accurately represent the total wet mass of the satellite due to the fact that the tanks would be costume made for the satellite specifications. [7]

Eq. 2

Figure 12 - Fuel Tank Configuration


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Figure 14 - UDMH Fuel tank and NTO Oxidizer tank

Another main factor to the propulsion system is the propellant pipes and their mass in correlation to the placement of the fuel tanks, main propulsion system, and ability to handle large amounts of pressure. Thus the material of choice for the pipes was chosen to be titanium 6AI/4V with a density of 4430 kg/m^3. Furthermore, the relative dimension of the pipes were picked to have a 13mm outer diameter and 9.8mm inner diameter. With a total length of titanium piping of 3.97m the estimated weight of the propellant piping will be 4.031kg. Although, they are relative estimates of mass, it is important to calculate them to perfectly map out a complete mass budget of the space craft bus. Moreover, by knowing the location of the pipes it helps mitigate the probability that they could potentially freeze leading to mission failure. [Figure 15]

Propellant Density Mass Volume

UDMH 793.00 kg/m^3 200 kg 252.207 L

NTO 1450 kg/m3 200 kg 137.931 L

Propellant Propellant Tank Size Tank Dimension Propellant Tank Mass

UDMH 252.207 L d=.650m h=.700 21 kg

NTO 176 L d=.6m h=.896 13.5 kg

Figure 13 - Fuel Mass and Fuel Tank Mass


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Figure 15 - Propellant Tanks/Oxidizer and Fuel Lines

Attitude Control Thrusters

The attitude control thrusters selected for the mission will not only perform the required 10 m/s station keeping yearly burn for IRIS-1, but they will also be capable of providing IRIS-2 and IRIS-3 with their 25 m/s station keeping requirements for the year. However, it needed to be taken into account that because of the 20 year life of the mission for each satellite, conventional hydrazine would prove difficult for storage over that period of time. This was partially because the bladder propellant tanks would lose pressure over time, likewise regular pressurized tanks would degrade, and introduce complex and heavy subsystems that would prove inefficient. Consequently, the decision was made to use ion propulsion systems for orbit station keeping, as well as minor attitude control adjustments. Correspondingly the ion propulsion system will also help unload any excessive torque that the reaction wheels accumulate over time. The way the system will work is that during the eclipsing time of the satellite the ion propulsion systems will fire and perform its monthly station keeping burn maneuver. Thus in order to properly calculate the amount of xenon propellant mass, burn time, and thrust required for station keeping it was particularly important to know how long the TOF of the satellite would be during shade. Thus, by using the solar power calculations for total time during shading from the solar array system section above, it was calculated that the satellite would be there for 29,530 seconds, approximately 8.2 hours. This is advantageous in the way that during this period of time the communications subsystem of the satellite which uses the most power consumption during its hours of operation, would actually not be operational during the shading part of the orbit since it would lose connectivity with earth station antennas. Therefore, by taking advantage of the

UDMH Fuel Line NTO Oxidizer Line


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situation the reserved 1.2kw of power would be rather used for orbit station keeping. However, ion propulsions system though efficient they do lack in thrust, so it was necessary to find the thrust required to maintain both the 10 m/s station keeping for IRIS-1 and the 25 m/s burn for IRIS 2 and IRIS 3. Using [Eq. 1] derived in the propulsion section it can then be calculated the least amount of burn thrust for either of the two station keeping orbits. Subsequently for the 10 m/s yearly burn for IRIS-1 it would require a minimum thrust of around .012 N. This will entitle a minimum of 24 burns, at around .4166 m/s each for about 28000 seconds or around 7.7 hours every week for a year. Therefore, that over 20 year’s lifetime of the satellite will equivalent to 480 burns. However, IRIS-2 and IRIS-3 are placed at somewhat unstable orbits thus their required station keeping burn is approximately 25 m/s. In order to accomplish its station keeping, both IRIS-2 and IRIS-3 must perform 1 burn a week, every week for the next 20 years. This will equivalent to around 1040 burns each burn being .384 m/s for and estimated time of 28000 seconds or 7.7 hours. In order for all requirements to be met for the orbital station keeping for both the 10 m/s and 25 m/s yearly burns, the ion propulsion thruster of choice was the RIT 10 EVO by Airbus Defense and Space [Figure 16].

Figure 16 - Attitude and control system


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This thruster is cable of delivering 15mN with an efficiency Isp of 3000s with an nominal input power of 435W. Subsequently, it was then calculated that IRIS-1 being in the most stable location for orbit stationing would require around .543 kg of xenon fuel, but IRIS-2 and IRIS-3, whom which are in very unstable orbit stations require roughly 1.085 kg of xenon fuel. Extra fuel will be added for any other type attitude control, thus an estimated extra. .1kg of xenon would suffice any other type of maneuver the satellite must do. Take a look at [Figure 17] for specifications on propellant tanks, and propellant mass.

Propellant Mass Propellant Tank Size Propellant Tank Mass

IRIS-1 Xenon .643 kg 2 x 58L 17 kg

IRIS2,3 Xenon 1.185 kg 2 x 104L 12.8 kg

Figure 17 - Propellant and Propellant Tank

The ion thrusters will be arranged in blocks of 5 one facing north, south, west, east, and one pointing up on the z axis, this cluster of thrusters will then be placed on the north wall and south wall of the satellite allowing redundancy, and complete 3 axis control, in addition the propellant tanks will be placed at the center of the face of the wall of the satellite one on the north and one of the south thus, giving the satellite great inertial moments and adequate center of gravity. All calculations for station keeping can be found in the appendix of this document.


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Thermal Control

Heat Transfer Theory

The thermal control system (TCS) is a critical system that generates the most optimal working conditions for the other major systems in a spacecraft. Some failure modes mitigated by the TCS include but is not limited to: material unreliability, onboard propellant freezing, thermal cycling damage, sensitive instrument misalignment, and meet instrument temperature requirements. A typical spacecraft design incorporates multiple components with different operating temperature ranges (OTR) and survival temperature range (STR) as is shown with an example sheet of a hypothetical design [8]:

Figure 18 - OTR and STR chart of Theoretical spacecraft

There are three fundamental forms of heat transfer: conduction, convection, and radiation. Due to the vacuum of space, conventional convection heat dissipation systems are obsolete due to the lack of pressurized atmospheric gases present to be a working fluid. That is why methods of conduction and radiation heat control are employed heavily in spacecraft. Conduction is the most efficient method of heat addition into cold hardware, while radiation is widely used as the best method for heat rejection in a vacuum environment. More specifically, heat sinks with large areas that radiate heat through infrared and other electromagnetic radiation is the main form of passive heat control. To begin width, the fundamental equation for conduction is listed below for 3 basic geometries:

𝑞 = 𝑘𝐴

∆𝑥(𝑇1 − 𝑇2) (Rectangular)

𝑞 =2𝑘𝜋𝐿(𝑇1−𝑇2)

ln (𝐷𝑜𝐷𝑖

) (Cylindrical)

𝑞 =4𝑘𝜋𝑅𝑖𝑅𝑜(𝑇1−𝑇2)

ln (𝑅𝑜−𝑅𝑖) (Spherical)


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These equations describe the heat flux that travels from static particle to particle in the described geometry. The next equation can be described as the dynamic version of conduction and introduces an important parameter ℎ, heat transfer coefficient.

𝑞 = ℎ ∙ 𝐴 ∙ ∆T (Convection) This describes the transfer of heat through means of enthalpy and kinetic energy change, a crucial concept in order to understand how to input work from a heat reservoir to a cold reservoir through phase change and flow; the main mechanisms of ammonia piping thermal control system. As for the third form of heat transfer, it is the way microwaves heat water and heat sinks reject heat in space:

𝑞 = 𝜀𝜎𝑇4 (Radiation) Where 𝜀 is the emissivity, 𝜎 is the Stefan-Boltzmann’s constant, and T is the temperature of the radiating material in kelvin. On top of material properties being exploited for heat transfer properties, there exists an entire field devoted to active feedback control theory in order to more precisely and uniformly regulate temperature throughout the spacecraft body.


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Multilayer Insulation

Multilayer insulation (MLI), along with their single layer counterparts, are among the most heavily used thermal control elements; this passive part of the TCS serves two main roles: prevent excessive heat loss from onboard components and to shield against excessive heating from environmental fluxes or rocket plumes [9]. MLI blankets are typically used to protect internal propellant tanks, propellant lines, solid rocket motors, and cryogenic dewars. MLI are only beneficial in vacuum environments due to the fact that MLI blankets are not very effective in the presence of gas. The selected MLI for the IRIS series spacecraft is AAErotherm S10-150, the specifications sheet is found in the appendix section of this report. The MLI blanket is made up of a 10 layers of 6 𝜇𝑚 thick polyester film aluminized on both sides, polyester non-woven fleece spacer and 350 angstrom aluminized foils. Below is a diagram with the layer characteristics of the selected MLI blanket [10]:

Figure 19: Layer Characteristics of AAErotherm S10-150 MLI blanket


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Optical Solar reflectors

Another passive system other than MLI blankets is the optical solar reflector (OSR), which are sheets of reflective material (usually a metal) laid on top and the side of the spacecraft. More specifically, an OSR is used for thermal control of spacecraft on the solar-facing sides of the satellites by shielding the thermally conducting material of the satellite bus from the high energy radiation particles ejected from the sun.

Figure 20 - Sample picture of OSR on spacecraft [11]

The selected OSR for the IRIS space mission is the QIOPTIQ plain OSR (CMO) glass type. The CMO

has a density of 2.54𝑔

𝑐𝑚3, which when applied to the 40 𝑚2 of exposed surface on the satellite

at a thickness of 0.10 𝑚𝑚 [12], yields an added weight of only 10.16 𝑘𝑔.


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Ammonia piping thermal - Control System / Thermo-Electric Control

The IRIS spacecraft generation is a design satellite that needs to last for around 20 years, thus its thermal control is a very important aspect to the spacecraft’s subsystem, not only does the space craft have a very well build passive thermal control system talked about previously but it also has a very active one. This was done in order to ensure the survivability of all the components in the satellite. One which it will be discussed about thoroughly would be the electronic thermal controller. The TAYCO Solid-state controller was the heater of choice for the satellites due to the fact that it had a solid state device thus no moving parts that could potentially break. The reliability and life expectancy of the solid state controller is like none other, thus why they are the industries favorite. The system employs a temperature sensor that can be located either at the controller or at a remote location, as desired. Moreover, another advantage of solid state controller is that extremely tight dead bands with are possible for very precise temperature control. The TYCO solid state controller also has unavailable heritage, ranging from the ISS all the way to military applications. IRIS over all carries about 70 solid state controllers to keep a numerous amount of systems alive, such as the CPU, IMU, ion thrusters, UHF, SDST, batteries, gimbals, propellant… etc., it was made sure that each system was acquired with one if needed. In addition ammonium heat pipes were also added to the space craft in order to radiate heat away from high temperature emitting systems. Thus an ammonium heat pipe will be implemented inside the satellite against one of the walls, then the high energy radiating system would be placed on top consequently helping the system cool.


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Electrical Power system

Solar Panels

When a spacecraft is sent on a long duration mission as is planned for the IRIS spacecraft, a source of electrical power generation is absolutely necessary in order to have a successful mission. Without a method of regenerating the onboard batteries, the size of the battery or duration of the mission would be unfeasible. There are two primary approaches of generating power onboard a deep space application spacecraft: solar panels and radioisotope thermoelectric generator (RTG). Usually a solar array system (SAS) is reserved for mission where solar irradiance is readily available and a more consistent power generation at beginning of life (BOL) and end of life (EOL) is desired. Whereas, an RTG system is reserved for missions that go farther outwards from the center of the solar system where the amount of solar irradiance is so low that not even solar concentrators can sustain onboard electronics.

Figure 21: GaAs Solar Cell Array on ISS (right), diagram of RTG (left) (6) (4)

Of the two options, the former was selected due to several advantages it holds over the latter; more specifically, a solar array system is more cost effective, is readily available, and is easier to manufacture. More specifically a gallium arsenide (GaAs) array was selected over its silicon crystalline counterpart due to its higher efficiency and more stable output. Due to the main source of power in an RTG being the radioactive material, plutonium; gaining the fuel source for this system is restrictive in both the financial and the political sense. With a price tag of $4.4 million for a kg of plutonium [8], the price of the most expensive GaAs arrays is dwarfed in comparison to the amount of radioactive fuel source needed for this specific mission.


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Figure 22: New Horizons Spacecraft with RTG

Aside from budgetary concerns, acquiring the materials for each system is presents its own challenge. For the solar arrays, GaAs presents a toxic hazard to both humans and the environment [10], as such, proper handling and manufacturing of the system is costly due to costs associated with the facility of manufacture and qualifications of trained personnel, with a bit amount of paperwork needed to acquire said materials; however, the amount of paperwork and procurement process for GaAs pales into comparison with the foreign relations policy associated with attaining plutonium, the only way to acquire plutonium is through the extremely limited and regulated supply kept by the United States Department of Energy [13] or to get it abroad from Russian post-cold war stockpiles, who currently has an embargo places against the U.S. [14]. For this reason, the use of gallium arsenide is widely accepted as the industry standard.

Figure 23: Worker performing black light testing on GaAs SAS in clean room (16)


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In terms of degradation over time, RTG are intrinsically tied to the half-life of plutonium due to its operational mechanism, and as such provides a far more stable power output than a GaAs array; an advantage which is highly outweighed by its disadvantages. As for the options of solar panels that are space rated, of the two (GaAs and silicon crystalline), GaAs has a lower degradation from BOL to EOL when exposed to a radiation heavy environment [9]. Several other advantages of GaAs are as follows: Output power 30% higher with same exposure to sun, 20% higher radiation reliability under same operating conditions, lower coefficient of efficiency change with temperature, and over 20% higher efficiency in energy conversion when compared to Si solar cells. Due to the advantages outlined in this section, GaAs solar panels have been selected as the primary source of power generation and is discussed in more detail in the following section.

Chosen Solar Panel Characteristics

The selected PV cell is comprised of a triple junction gallium arsenide system which yields an optimal efficiency of 29.5% for solar energy conversion. For the purposes of this mission the percentage loss of power generation from beginning of life to end of life has been approximated to be 42.7% over a 20 year span, with an assumed efficiency loss of 2.75% per year. The total weight of the cells by themselves is 19.773 Kg. Below is a graphic depicting the chemical composition of the different layer substrates that compromise the selected photovoltaic cell.

Figure 24: Schematic of layers for GaAs cell (9)

The triple junction GaAs battery is the most efficient solar cell available from Emcore [12]. With an areostationary altitude of 17000 km, the total orbit period of the satellite around mars was verified by calculation to be the same amount of time as a Martian day, 24.556 hours. Taking into consideration altitude and the radius of mars, the angular radius comprising the section of shading for the spacecraft is approximated to be 0.167 rads, converting the angular units to time using the period of areostationary orbit yields an eclipse period of 8.203 hours. This leaves a time


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exposed to sun of about 16.353 hours; in other words, this is the total amount of time the solar

panels can generate power each day. The solar irradiance of the sun was taken to be 586 𝑊

𝑚2

[15]. From the power consumption chart listed in the appendix, the minimum power required to run the critical systems and account for nominal use of active systems is 815 𝑊; since EOL power draw is only 57.3% the amount of BOL power draw, an efficiency compensation was used to recalculate the required BOL power draw in order to guarantee system operation at EOL, this required power draw was approximated to be 1,424 W. Taking into account the fact that the solar panels do not generate power during shading, and using the ratio of orbital period over sun exposure period, to provide further compensation, the total power draw capability of the solar array system must be 2,138 W in order to guarantee system operation both during shading and EOL. The equation for power draw of a solar cell in orbit around a planet can be approximated to be:

𝑃𝑑𝑟𝑎𝑤 = 𝐻𝑜 ∙ 𝜂𝑠𝑝 ∙ 𝐴 ∙ cos (𝜃𝑖)

Where 𝐻𝑜 is the solar flux of 586 𝑊

𝑚2 at mars distance from the sun, 𝐴 is the area of the solar

cells, 𝜃𝑖 is the incident angle of the plane parallel with the front face of the panel with respect to the flux vector of radiation from the sun, and 𝜂𝑠𝑝is efficiency of solar panels listed at 29.5% in the

specifications sheet [12]. In order to find the total area needed for the solar panels, the above equation needed to be rearranged to solve for 𝐴. An average incident angle of 45 degrees was arbitrarily selected as a stationary value and used in a rearranged form in the following equation for total power draw:

𝐴 = 𝑃𝑑𝑟𝑎𝑤

𝐻𝑜 ∙ 𝜂𝑠𝑝 ∙ cos (𝜃𝑖)

In this form of the equation the total area needed for the solar cells to produce the required

power is 23.539 𝑚2. With the specified surface density of 84 𝑔

𝑐𝑚2, the total mass of the solar

arrays came out to be 19.773 𝑘𝑔.


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Solar Array Retention Mechanism

Immediately following the oberth escape maneuver of the IRIS series spacecraft, the deployment of the spacecraft will prove to be a critical stage in the mission, since failure would most certainly mean failure of the overall mission. That is why a reliable and precise solar array retention mechanism (SARM) will be needed in order to expand the solar cells properly; furthermore, the same system that performs this initial action will also be tasked with guaranteeing that the collapsible solar array remains in an extended position at all times. This amplifies the necessity of selecting a dependable system impervious to long-term fatigue. IRIS will employ a system, manufactured by Alcatel Space that has been tested and proven on the COSMO-SkyMed satellite system launched in June 8 2007 [16]. A schematic of the SARM onboard the COSMO-Skymed is provided below:

Figure 25: Schematic of Main Components of SRAM

The SRAM is designed to allow a continuous control of deployment, be simple, reliable, lightweight, and has capacity [17]. There are 6 points of hinges where the SRAM will be placed in order to fully extend both 3-segment solar panels on the IRIS spacecraft. A crude schematic of stowed phase and initial deployment of one of the panel arrays is shown below.


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Figure 26: Solar Panel in stowed position (left) and initial deployment (right)

Below is the corresponding schematic of the SRAM in stowed and deployed configurations:

Figure 27: Schematic of stowed (top) and deployed (bottom) SRAM


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Solar Array Drive Mechanism

Aside from the initial deployment and extension maintenance operations for the solar arrays, the solar arrays must also be rotated in order to be pointed at the sun in the most efficient way. That is the main function of the solar array drive mechanism (SADM). This mechanism coupled with the system computer, and Airbus BASS7 sun tracker system [18], described in the Guidance Navigation and Control section of this document, will comprise of the active feedback system that senses and adjusts panel orientation in order to optimize the power generation of the SAS. For the IRIS series spacecraft, the Surrey Solar Array Drive Mechanism was chosen due to its capabilities, low power consumption, and heritage onboard the NigeriaSat-2 (the Antenna Pointing Mechanism) and GIOVE-A (slip ring) [19]. A picture of the selected Surrey SADM is shown below with more detailed engineering schematics located in the appendix:

Figure 28: Picture of SADM for Rotation of Panels (left) Engineering Isometric View (right)

The selected SADM is a very capable system with the single-axis model capable of ± 0.45 °𝑠⁄

maximum rotation speed and ± 0.07 °𝑠⁄ nominal speed, more than enough to match the

angular speed of the dynamic solar incidence angle. Furthermore, there is a large operational temperature range of 30°𝐶 and 60°𝐶 and nonoperational temperature of −40°𝐶 to 70°𝐶 ; meaning, that orientation correction can occur in the range of the former and system integrity will be maintained in the range of the latter. The wide temperature ranges allow for a larger target temperature for the thermal control system (TCS) which in turn allows the active feedback control to regulate the temperature more efficiently, which translates to more operational time and design lifetime. Although the specifications sheet for the Surrey SADM states that the intended design lifetime is 6 years [19], upon procurement, the manufacturer will be asked to make the appropriate modifications to extend the lifetime to a minimum of 20 years for the IRIS series spacecraft.


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Batteries

The choice of batteries for this mission is made of lithium ion cells from Saft Batteries, model VES 180 [20], due to its high energy density of 583,783 J/Kg relative to the other Saft battery models as is shown in the figure below:

Figure 29: Comparison Chart of Various Saft Li-Ion models

With a standard depth of discharge of 75% the calculated minimum energy capacity needed from the battery system is 124.6 𝑀𝐽 with 213.438 𝐾𝑔 of lithium ion needed to achieve this; however, each cell provides 648,000 J and weighs 1.11 𝐾𝑔. Therefore in order to take into consideration the step size nature of battery selection and achieve the desired energy capacity, 193 battery cells are required yielding a total weight of 214.23𝐾𝑔 and provides 125.1 𝑀𝐽 when arranged properly. The Power draw necessary to compensate for eclipse time and EOL efficiency loss is 2,138 𝑊 when in the sun. The Batteries will be placed near the center of the spacecraft in order to minimize the offset of center of gravity of the spacecraft and to mitigate moment arms in order to induce attitude stability.


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Figure 30: Possible Multiple Cell Arrangements (18)

Due to the relatively high energy application needed for the IRIS series space mission, a custom order scaled version of the proposed battery that has been extensively tested would be needed. The small 1.11 𝑘𝑔 cells were used in calculations as a means of attaining a first order approximation of battery size, mass and capability; the reality is that the Saft VES 180 modular lithium ion battery is restricted to being arranged to a maximum of 12 cells parallel and 24 cells in series, suitable for LEO but not sufficient for GEO or AMO applications. Conversely, there is a mention of design that provide capabilities of 3 𝑘𝑊 to 25 𝑘𝑊 in the Saft brochure [20].

Figure 31: Diagram of Heritage Flights and Missions for Saft


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Satellite Subsystems

The following chart breaks down the power consumption of each subsystem onboard each satellite.

Hardware Power (W) Mass (Kg)

Structure

Chassis Passive 322.31

Propulsion

Main propulsion system + thrust frame 35 4.74343

Propellant tanks N/A 34.5

Propellant N/A 400

Propellant fuel pipes N/A 4.031

Thermal Control

Thermal Blanket N/A 6.4

Optical solar Reflectors N/A 25.4

Electronic Thermal Controllers 70 2.1

NH3 Heat Pipes 1.05 10

Electrical power system

Solar panels passive 16.87

Solar cell cover glass N/A 14.9

Solar array retention mechanism X 6 N/A 3

Solar Panel Gimbals X2 7.2 12

Batteries active 185.4

Guidance Navigation and Control

Computer (2x) 75 6

Star Tracker (2x) 9 11

Sun Sensor (2x) passive 0.13

Mars Tracker (2x) 6.5 6.8

IMU 8 6.8

Reaction Wheels(x4) 90 14.8

(10x) Attitude ion control thrusters 435 18

Xenon Propellant Tank N/A 0.643

Xenon Propellant Tank N/A 17

Communications

Communication HGA-1 gimbal 3.9 3.3

HGA-1 Ka-band TWTA sat-earth 97.3 1.495

HGA-2 Ka-band TWTA sat-sat 1.05 1.495

HGA-3 Ka-band TWTA sat-mars 0.883 1.495

HGA-1 X-band TWTA sat-earth 152 0.965

HGA-2 X-band TWTA sat-sat 14.5 0.965

HGA-3 X-band TWTA sat-mars 12.8 0.965

(2x) Electra UHF 33.5 3

(4x) SDST 78 12.8

1130.683 1149.30743

695

<- Ion thrusters only used when not in view of Earth


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Guidance Navigation and Control

Computer

The computer system will be used to perform all important tasks onboard the satellite from running the communications subsystem, to controlling thrusters and reaction wheels. Using the following sensors as input, the computer is able to maintain station and attitude.

Star Tracker

A star tracker from Ball Aerospace, model CT-602 will be used for attitude control. The CT-602 has a mass of about 5.5 kg and a power consumption of 8-9 W. Its purpose will be to sense the orientation of the satellite via a programmed set of star clusters. If the star tracker senses a change in the viewing angle of the cluster, the satellite will change its orientation to properly view it. The star tracker will also provide information on pitch and roll for the satellite.

Figure 32: Star Tracker

Sun Sensor

A passive sun sensor from Airbus Defense & Space, model BASS7, which has a mass of 65 grams, a field of view of +/- 15 degrees, and a sun detection of +/- 90 degrees. This will provide the solar panels with the direction of sunlight so that the panels will always be facing the sun.

Figure 33: Sun Sensor


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Mars Tracker

IRIS will use a reprogrammed STD15 Earth sensor to track Mars from Sodern. The "Mars sensor" will have an altitude range between 15,000 – 140,000 km, a mass of 3.4 kg, and will consume 6.5 W. This sensor will allow the bottom antenna to have a constant view of Mars for communications for the ground crew.

Figure 34: Mars Tracker

IMU

An IMU-HG1700-H62 from Novatel will be used for navigation. This IMU features a ring-laser gyro for precise measurement, has a mass of 3.4 kg, and a typical power consumption of 8 W. The IMU will play a crucial role in attitude correction. It will provide information on pitch, roll, and yaw for movement.

Figure 35: IMU

Reaction Wheels

For attitude correction, IRIS will utilize the RSI 12 Momentum and Reaction Wheels from Rockwell Collins. They will provide an angular momentum storage of 12 Nms with a nominal torque of 75 mNm. The mass and max power are rated to be 4.85 kg and 90 W. A total of four reaction wheels will be present and will be assembled in a “tetrahedral” configuration, as shown below. This configuration will provide a means of redundancy in the event that one fails.

Figure 36: Tetrahedral Configuration


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Communications System The IRIS communications system is designed to provide fast transmissions, providing reliability and long life without excessive use of onboard resources.

Satellite Constellation Arrangement

The design of our system incorporates the use of three medium sized satellites positioned at roughly 120° intervals with one (IRIS I) placed directly above Meridiani Planum in equatorial areostationary orbit, and two others (IRIS II and IRIS III) placed ±120° about the equator from that location. Due to the varying surface features such as mountain ranges and the ellipsoidal shape of Mars, station keeping will require a substantial amount of fuel if not placed at, or near, defined locations about the Martian Equator.

The figure below visualizes the annual ∆V required along longitudinal stations necessary for station keeping.

Figure 37 - Annual ∆V for longitude station


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Because there are four stable points for areostationary orbit and there is a need to position only three satellites, it becomes necessary to offset at least two from the stable points. To minimize the fuel required for station keeping, while ensuring that each of them are around 120° apart, the satellites will be placed at the following locations:

IRIS I: 0°N, -13.5°W

IRIS II: 0°N, -132.5°W IRIS III: 0°N, 107.5°E

IRIS I will be placed at the most stable location of the three as it will require the most reliability due to its direct communication over the vicinity of the proposed colony, Meridiani Planum, which is located at 0.2°N, -2.5°E. For a more in-depth analysis of station keeping at each satellite location, please refer to the Propulsions section of this paper.

Communication Windows

During the beginning early stages of the mission, the communication windows with Earth’s Deep Space Network will be limited. The mission requirements only involve providing about 6 hours of communication a day. Therefore, by placing an areostationary satellite around mars could increase the overall communication window. An STK scenario was run in order to simulate the access times the satellite would have with Deep Space Network 34 meter high gain antennas. Furthermore, the access report was done individually on every facility which included DSS 25, DSS 26 both in Goldstone California, DSS 34 in Canberra Spain, and DSS 25 near Madrid. The Figure below shows all the locations of the Deep Space Network which were simulated with STK.


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During the simulation each individual facility was simulated to find an overall communication access with that specific location. In the following pages some of the simulated results will be displayed and discussed in detailed to how long each facility was able to talk to the satellite or was in line of view. However, take into consideration that free space loss, signal to noise ratio, and other losses were not taken into consideration.

The first simulation shown above displays the access IRIS 1 had with DSS-55 which came to a total of 44946.143 seconds or 12.485 hours with a mean duration of 6.24 hours. However, the report shows that there was a time when the satellite only had a30 min connectivity. Thus it’s important to keep on striving for almost 100% communication reliability between Mars, and Earth. In addition many more simulations were ran on other DSN facilities and it was found that some others had worst communication consistency than the first facility simulated had. Once the robotics face of the mission has passed within the first 10 years, IRIS must prepare for its ultimate task. Not only will IRIS become the first ever communication constellation orbiting Mars, but it will also provide a way of communications for the future settlers of mars. Therefore, it was crucial that all three satellites be able to communicate with earth at any given time. Figure 1 and 2 shows the communication infrastructure that would be set in place for the colonist.


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The above illustration displays the communication system that will be in place during Martian colonization, it can be seen here that there are 3 areostationary satellites which at all times are able to have communication with Earth when in line of sight. Subsequently, by increasing the number of satellites it then increases the amount of time one or more satellites could talk with earth, ultimately increasing transfer rates of date from the Martians. Nevertheless, there is no getting away from nature and the phenomena of black outs and white outs due to the suns solar flares. This usually occurs every two years and only hinders communications for a few days. Fig_-- displays this phenomenon where there is no getting away.


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System Capabilities

Each satellite will be equipped with features to further improve the success of the IRIS satellite communications. To allow for increased data transmissions to Earth, each satellite will have the ability to communicate with adjacent satellites to allow for easy data passing. Once all three satellites are placed in orbit, communication with Earth will always be accessible. When a colonist or a ground vehicle is in need of transmitting to Earth while on the far side of Mars, transmissions would simply be uplinked to the satellite overhead, passed to an adjacent satellite with line-of-sight to Earth, and transmitted. Similarly, a received transmission from Earth can be directed toward a ground station or vehicle on the far side of Mars using the same methodology. Through the use of encoded messages attached to each transmission, satellites can determine how to direct a message appropriately. The IRIS communications system will also be uniquely capable of reconfiguring waveguides and data paths to switch between frequency bands on all antennas. This not only works as a redundancy factor in onboard communication systems, but also allows for better penetration through noise on its way to Earth. When conditions are ideal, the main high gain antenna can switch to transmissions on the higher frequencies of Ka-Band. Typically, Ka-Band is more susceptible to noise, but when conditions allow, it would allow for much faster data transfer rates.

Antennas

To allow the IRIS satellite communications system to provide the excellent data transfer rates necessary for this network, a total of five antennas will be needed. This includes two antennas for communication with Earth (a large main dish antenna, and a backup low gain antenna), two opposite facing dish antennas for message passing between satellites, and one facing the Martian surface.

Figure 38 - IRIS antenna Layout


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Satellite-to-Earth High Gain Antenna (HGA1)

The largest of the onboard antennas will be the 3.5m diameter high gain parabolic antenna to be used for transmitting and receiving data with the Deep Space Network on Earth. Providing 47.9 dBi of gain in X-Band mode and 59.5 dBi of gain in Ka-Band, this antenna will be capable of transmitting up to 24Mbps in ideal conditions at its farthest distance from Earth, and potentially higher than 30Mbps when Earth and Mars are closest.

Figure 39 - 3.5m High Gain Antenna schematic

While using a 32GHz Ka-Band frequency, the beamwidth would be around 0.31°, whereas an X-Band frequency of 8.5GHz would be around 1.18°. Because the expected beamwidths are so small (particularly Ka-Band), a long lasting high precision gimbal system will be necessary to ensure accurate aim and reliability.

Satellite-to-Earth Low Gain Antenna (LGA)

Each IRIS satellite will also be equipped with a Low Gain Antenna. Typically used as backup antennas, the primary purpose of the low gain antenna will be to communicate with Earth when no other options are available. With a beamwidth of around 10°, much of the signal will be lost to space, but will still allow communication at low rates. This antenna would primarily be used while in transit to Mars, but could also be used in the case of extreme circumstances while in commission.

Low Gain Antennas provide very low

data rates, but are a basic necessity of any satellite.


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Satellite-to-Satellite High Gain Antenna (HGA2 & HGA3)

To make inter-satellite communications possible, there needs to be an antenna mounted on either side of each IRIS satellite. While this antenna is considered high gain, the beamwidth is as large as 8.25° while using X-Band (2.19° in Ka-Band). Because the bandwidth will translate to a fairly large coverage area, aiming will not require the same precision as the main HGA, allowing both to be mounted in a stationary position on the satellite.

Figure 40 - 0.5m antenna to be used for inter-satellite communications

Providing 42.5 dBi of gain in Ka-Band and 31.3 dBi of gain in X-Band, each IRIS satellite will be capable of transmitting at 24Mbps or higher.

Satellite-to-Mars Medium Gain Antenna (MGA)

From approximately 17,000km above the surface of Mars, IRIS satellites will have the capability of communicating over X-Band, Ka-Band, or over UHF frequencies. Depending on the frequency used, the data rate can be very fast or very slow. The break-down for each frequency is listed below:

Band Frequency Dish Gain Bitrate Beamwidth

Ka 30.5 GHz 40 dBi 24 Mbps 2.87°

X 8 GHz 28.6 dBi 24 Mbps 11°

UHF 405 MHz 2.7 dBi 160 Kbps >180° Figure 41 - Satellite-to-Mars link budget


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The antenna consists of a 0.4m diameter, 0.1m depth paraboloid antenna with the signal reflector positioned at 0.1m from the back end of the dish.

Figure 42 - 0.4m diameter antenna to be used for communications with the surface of Mars.

Due to the nature of differing frequencies on a dish antenna, there will be a variety of ground cover from the satellites based on frequency band.

Figure 43 - Area of coverage on Mars for each frequency band

As you can see in Figure 42, Ka-Band use will primarily be restricted to communication with areas within or near the vicinity of the proposed colony location, while X-Band can be used along with the other two satellites to gain nearly full coverage of the planet. This antenna is also capable of UHF, giving it a very low gain, but also allowing backwards compatibility with rovers currently on the surface of Mars.


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Earth Ground Station(s)

To allow for constant communication, the rotation of the Earth must be taken into consideration. As the Earth rotates, any individual ground station would quickly lose line-of-sight with IRIS satellites. To resolve this, NASA’s Deep Space Network (DSN) could be used. If funding allows, it would be more cost effective in the long-run to build an independent set of antennas, based on similar specifications, dedicated to the use of communication with Mars.

Deep Space Network

Figure 44: DSN Antenna

NASA’s Deep Space Network is a set of ground stations and antennas used for communications with interplanetary satellites setup at roughly 120° longitude intervals around Earth. With cryogenically cooled hardware, DSN is capable of decoding very low received transmission signals, as low as -156dBm with few bit errors [21], allowing effective communication with spacecraft from billions of kilometers away. DSN has three locations, each of which are located in either desert or mild climates. Because of this, weather will rarely be a factor in degradation of signal strength. Attenuation of the signal due to instances of inclement weather will generally result in around 0.2dB loss, which may require switching to X-Band, but will not pose any major issues.

Figure 45: DSN Locations: Goldstone, California – Madrid, Spain – Canberra, Australia

IRIS would require use of the 34 meter diameter antennas of each location, including DSS 25 & DSS 26 in Goldstone, DSS 34 in Canberra, and DSS 55 near Madrid. Use of the Deep Space Network


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currently costs approximately $1057/hr. The table below lists the communications specifications of the 34m antennas.

Antenna 34-m BWG

Band S X Ka

Tx Freq. Range [MHz] 2025-2120 7145-7235 34315-34415

Tx Gain [dBi] 56.25 67.09 79.52

Tx Beamwidth [degrees] 0.263 0.077 0.016

Tx Power [dBm] 73 73 59

Rx Freq. Range [MHz] 2200-2300 8400-8500 31800-32300

Rx Gain [dBi] 56.84 68.24 79.00

Rx Beamwidth [degrees] 0.242 0.066 0.017

Figure 46: 34m DSN wireless communications specifications [22]

Communications Subsystem Components

The communications subsystem is the most critical system onboard each satellite. Because it is necessary to ensure that there is a reliable connection at all times, regardless of distance or weather, the communications subsystem has been designed to be backwards compatible with existing Mars rovers along with having the capability of adapting to real-time conditions. Figure 47 on the following page contains a breakdown of the communications subsystem. As can be seen from the diagram, there are redundant systems for each required component, shared between waveguide paths through a series of switches, couplers, and multiplexors. This ability to redirect waveguides allows us to customize data transmission rates through diplexing, left and right circular polarization, and frequency band switching.


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Figure 47: IRIS Telecommunications Subsystem Block Diagram


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Command & Data Handling

The Command and Data Handling system consists of two computer processors, the RAD 5545, developed by BAE Systems, to act as a redundancy and to allow for fault detection caused by interference from radiation. The processors themselves contain four RAD5500 cores, providing us with the computing power necessary to run four transponders at a time, in addition to all other necessary tasks required by the satellite. Because the RAD 5545 is still currently in development, alternate radiation hardened processors may be used.

Figure 48: Example BAE Systems computer assembly [23]

Within the communications subsystem, the computer will be responsible for switching waveguide switches into the appropriate positions to allow for transmitting or receiving signals, determining the correct waveguide paths, managing coding blocks, and interfacing with the small deep space transponders.

TWTAs

The Traveling Wave Tube Amplifier strengthens the signals generated by the transponders. Each IRIS satellite will contain three Ka-Band TWTAs and three X-Band TWTAs, each of which are rated for circularly polarized frequencies. Because each of the antennas are capable of transmitting or receiving Ka-Band and X-Band transmissions, switches within the system must be set to route the appropriate TWTAs to the correct antenna. As you can see in Figure 48, TWTA2 of each set will be capable of transmitting through any of the antennas. Because of the power requirements of each antenna vary, and since each TWTA has a set power input, TWTA2 will either need to be set to the power requirements of HGA1, or there will need to be an additional switch to adjust power input when necessary. All TWTAs will be mounted in an enclosed panel on the outside of the satellite chassis to minimize noise generation through heat dissipation and isolation from onboard electronics.


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Figure 49: TWTAs to be used on IRIS satellites. Produced by L3 Communications. [24]


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X and Ka Band Transponders (SDST)

IRIS satellites will have a total of 6 transponders onboard. To generate X-Band and Ka-Band frequencies, four Small Deep Space Transponders (SDSTs) will be used [25], each to be shared between the five antennas. The SDSTs will act as the bridge between frequency wave signals and digital data. Each one will be capable of receiving or transmitting at any time.

Figure 50: Small Deep Space Transponder

Electra UHF Transponder (EUT)

Electra UHF Transponders will also be used for communication with ground vehicles on the surface of Mars. These transponders are self-contained and do not require amplifiers. Two will be onboard for added redundancy.

Figure 51: Electra UHF Transponder


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BandPass Filters (BPF)

Coaxial bandpass filters (BPF 1-4) are used to filter out interference from the received frequencies. BPF5 is a waveguide bandpass filter which is necessary to have in order for the low gain antenna to filter out harmonics generated at the transmit frequency.

Couplers (CP)

Couplers are necessary to allow switching between TWTAs and to provide a 90° phase shift (time delay) for use in generating the circular polarization.

Diplexers (DX)

Diplexers control the direction of frequency signals as they are transmitted, or are received. This is required to have on each antenna to ensure that received transmissions are directed into the transponder and not back into the TWTA.

Isolators (IS)

Isolators protect the TWTAs in the case of a waveguide short with received transmission signals.

Switches (S)

Having several switches onboard the IRIS satellite communications subsystem will be necessary to control and allow the necessary customization of frequency bands for each antenna. Mitigation of risk of failure in switches can be achieved by opting to use reliable switches that will be left in an operational position, rather than between positions, in the case of failure.

Figure 52: Example SPDT waveguide switch [26]

Different variations of switches will be required. S2 and S3 are coaxial DPST circulator switches to allow switching receiver inputs for the SDSTs. S1 and S4 – S8 are all SPDT switches that allow switching of inputs or outputs.


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Multiplexers (MUX)

Waveguide multiplexers are necessary to select the input TWTA for each antenna. These work as multiple input waveguide switches.

Figure 53: Ka-Band Multiplexer – Bosch [27]

Ultra-Stable Oscillators (USO)

To ensure that each transponder delivers the correct frequency, each transponder is connected to an onboard Ultra-Stable Oscillator. These USOs provide the clock signals to ensure that everything onboard runs with the correct timing rate. The system includes a redundant USO in order to mitigate risks and raise the probability that there is at least one operational system throughout the satellite’s mission life.

Figure 54: MicroSemi Ovenized Master Oscillator [28]


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Link Budgets

All link budgets are calculated based on a worst case scenario, assuming an SNR of 2 and a distance of 400 million km from Earth. Each satellite is 30,000km apart, and in orbit at 17,000km above the surface. The calculated Data Rate represents the Shannon-Hartley Limit.

X-Band Satellite to Earth (DSN 34m)

Distance: 400,000,000km Parabolic Dish Diameter: 3.5m Dish Efficiency: 0.65 SNR: 2 Bandwidth: 2 MHz Frequency: 8.44 GHz Data Rate: 3.17 Mbps Receiver Gain (DSN): 68.24dBi Transmitter Gain (Satellite): 47.959 dBi Free Space Loss: 283.02 dBi Temperature: 31.33 K Transmit Power: 83.45W Received Power: -117.618 dBm

Ka-Band Satellite to Earth (DSN 34m)

Distance: 400,000,000km Parabolic Dish Diameter: 3.5m Dish Efficiency: 0.65 SNR: 2 Bandwidth: 15 MHz Frequency: 32 GHz Data Rate: 23.77 Mbps Receiver Gain (DSN): 79 dBi Transmitter Gain (Satellite): 59.5 dBi Free Space Loss: 294.59 dBi Temperature: 31.33 K Transmit Power: 52.54W Received Power: -108.867 dBm

X-Band Satellite to Satellite

Distance: 30,000km Parabolic Dish Diameter: 0.5m Dish Efficiency: 0.65 SNR: 2 Bandwidth: 15 MHz Frequency: 8.5 GHz Data Rate: 23.77 Mbps Receiver Gain: 31.1 dBi Transmitter Gain: 31.1 dBi Free Space Loss: 200.5 dBi Temperature: 280 K Transmit Power: 7.97W Received Power: -99.355 dBm

Ka-Band Satellite to Satellite


X-Band Satellite to Mars Surface

Distance: 17,000km Parabolic Dish Diameter: 0.4m Dish Efficiency: 0.65 SNR: 2 Bandwidth: 15 MHz Frequency: 8 GHz Data Rate: 23.77 Mbps Receiver Gain: 28.64 dBi Transmitter Gain: 28.64 dBi Free Space Loss: 195.11 dBi Temperature: 280 K Transmit Power: 7.06W Received Power: -99.355 dBm

Ka-Band Satellite to Mars Surface



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Signal Generation

Generating the most effective signal is optimal; at times, transmissions between Earth and Mars can reach a maximum distance of 400,000,000 km. The Free Space Path Loss associated with these distances can lead to signal power loss and lead to a higher SNR. To help with this, various techniques have been implemented that can improve the signal as well as methods of signal reconstruction. Data packets will be transmitted using time-division multiplexing.

Polarization

Rather than using the typical linear radio waves, circular polarization will be used. This not only for adds stability to the transmission of the radio waves, but also allows communication devices on the ground to be positioned in various orientations while still being able to receive signals.

Figure 55 - A circularly Polarized radio wave

With reconfiguration of hardware, IRIS satellites could support transmissions in both Left Circular Polarization (LCP) and Right Circular Polarization (RCP) on the same bandwidth that would nearly double the transmission rate.

Modulation

Modulation is the method used to represent data with radio waves. Depending on the strength of the signal, increasingly complex modulation methods can be used, leading to an increase in data rates. The following methods of modulation are expected to potentially be used in communications. BPSK – Binary Phase Shift Keying Transmits 1 bit per symbol (slowest) QPSK – Quadrature Phase Shift Keying Transmits 2 bits per symbol (average) QAM16 – Quadrature Amplitude Modulation

(4x4 bit constellation diagram) Transmits 4 bits per symbol (fastest)

Figure 56 - Illustration of QAM16 modulation


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Coding

Coding is the method of adding parity bits to add additional information to help reconstruct corrupted data on the receiving end of a transmission. Because there will only be one channel of communication at any given moment, Turbo Coding is the optimal method of coding which provides a transmission rate near the Shannon-Hartley limit. Depending on the amount of signal noise, various bit ratio protocols may be used. When very high noise is present, a transmission protocol of ½ ratio would be employed; this translates to transmitting one parity bit within every two bits (doubling the amount of data). However, when less noise is present, the coding rate goes down to 1/4 ratio or lower allows for higher data bandwidth.

Duplexing

Duplexing can be restricting based on the method used. To half-duplex transmissions, only one transponder is needed at a time for each antenna. Because of limitations on the amount of hardware allowed onboard the IRIS satellites, it will not be feasible to perform full-duplexing on all antennas, thus restricting communications to half-duplexing. On the other hand, given that an antenna (either receiving or transmitting) is not in use, full duplexing could be used in order to accommodate situations where high traffic is expected.


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Conclusion In conclusion, the IRIS Mars Communications Network will provide the best communications transmission rates to date, while integrating ground breaking communications techniques. It will employ top-of-the-line satellite subsystems that will provide each IRIS satellite with a 20+ year expected lifespan. Furthermore, IRIS will provide the benefit of keeping human colonists in touch with loved ones even at a distance of 400 million kilometers away.


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Appendix Mars Orbital Maneuvers – MathCad Calculations

Hohman Transfer Phase 2 Looking for V.infinity entry into hohman transfer from the orbit of the earth to the orbit of Mars Time required to reach orbit of Mars

Sun and mars data

mars velocity around the sun

radius of mars

Earth Parameters

one AU is earths distance from the earth to the sun

earths velocity around the sun

radius of earth

S 1.3271020

m

3

s2

AU 1.49610

11m

Mars

M 4.2971013

m

3

s2

VM 24140m

s

rM 3.393106m

aM 1.524AU

aM 2.28 1011

m

E 3.9861014

m

3

s2

aE AU 1.496 1011

m

VE 29790m

s

rE 6.378106m


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Transfer trajectory

Transfer Velocity at Perihelion

at

aE aM 2

at 1.888 1011

m

Et

S

2at

3.514 108

m

2

s2

Vtp 2 Et

S

aE

3.273 104

m

s

Ht aE Vtp 4.896 1015

m

2

s

.2081

Vta

Ht

aM

2.148 104

m

s

Vinf Vtp VE 2.939 103

m

s

TOF 2at

3

S

4.474 107

s

TOF

20.709yr


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Rotating earth to parking orbit delta v

Transfer Ellipse

Inclination change at Parking Orbit

hparking 185km

earth 7.2921159105

rad

s

Rparking rE hparking 6.563 106

mVearth earth rE 465.091

m

s

ra2 6378.1km 185km 6.563 106

mrp2 rE 6.378 10

6 m

a2

rp2 ra2

26.471 10

6 m

Vparking

E

Rparking

7.793 103

m

s

Vp2

2E

rp2

E

a2

7.962 103

m

s

Va2

2E

ra2

E

a2

7.737 103

m

s

Vpark1 Vp2 Vearth 7.497 103

m

s

Vpark2 Vparking Va2 55.994m

s

Vparktotal Vpark1 Vpark2 7.553 103

m

s

i 1.85

180

Vinclination 2Vparking sini

2

Vinclination 251.621m

s


73 | P a g e S e p t e m b e r 2 0 1 5

1 Impulse Oberth Escape Transfer Maneuver

Planetary Capture

Finding Δ V going from Mars SOI to Areostationary Orbit

Elliptical Transfer Orbit

Vesc 2.5

Vparking 11021.29m

s

Vinj Vesc2

Vinf2

11406.421m

s

Vinj Vinj Vparking 3613.192m

s

Vtransfer Vparktotal Vinj 11165.877m

s

Vplanetarytransfer VM Vta 2664.291m

s

Vcapture Vtransfer Vplanetarytransfer 13830.169m

s

rSOImars 5.781108

m rareo 17000km

ra4 rSOImars rp4 rareo

VSOI

M

rSOImars

272.635m

s

a4

rSOImars rareo 2

2.975 108

m

Va4

2M

ra4

M

a4

65.167m

s

Vp4

2M

rp4

M

a4

2216.051m

s


74 | P a g e

Total Δ V Needed for Mission

Total Δ V Needed on Spacecraft

Total Δ V Supplied by Launch Vehicle

Vtrans1 VSOI Va4 207.468m

s

Vareo

M

rareo

1589.858m

s

Vtrans2 Vp4 Vareo 626.193m

s

VMOI Vtrans1 Vtrans2 833.661m

s

Vtotal Vcapture VMOI Vinclination 14915.451m

s

Vspacecraft Vplanetarytransfer VMOI 3497.953m

s

Vlaunchvehicle Vtotal Vspacecraft 11417.499m

s


75 | P a g e S e p t e m b e r 2 0 1 5

Solar Panel Calculations m m^3/s^2

m

Mars Angular Radius

rads

Time of Eclipse

Time Exposed to Sun

Calculating Energy Density of Sun with Planck's Blackbody Formula

λ = ?

rmars 3390000 mars 42828109

sp 0.295

h 17000000 a rmars h Pmin 700 Watts

asinrmars

rmars h

0.167

Pmars 2 a3

mars

Pmars 8.84 104

P 2 Pmars

P 2.953 104

Psun Pmars P

Psun 5.887 104

h 6.631034

Tsun 5800

c 299800 k 1.381023

Hsun2 h

5

e

h c

k Tsun1

Ho

Rsun

dmars

2

Hsun Hsun

Ho 590


76 | P a g e

Efficiency loss from BOL to EOL

Calculating Power Required

Calculating Area using power

Average incident angle of 45 degrees

m^2 Area of Solar Panels

Mass of Solar Panels

kg Mass of Solar Panels

Battery Calculations

EOL 1 .0275( )20

EOL 0.573

Po

Pmin

EOL

Po 1.223 103

Pdraw Po

Pmars

Psun

Pdraw 1.836 103

Watts

i 45

APdraw

Ho sp cos i

A 20.081

.84

msp A

msp 16.868

E Po Pmars

583783E 1.081 10

8

MassbatE

185.138 kg


77 | P a g e S e p t e m b e r 2 0 1 5

Ion Propulsion Station Keeping

Ion propulsion system RIT 10 EVO

Fuel: xenon

Isp: 3000s

power: 435 W

Burn schedule and burn time: 2 weeks every month, for one year is 24 times 20 years 480 burns

Total fuel used

v 1 .4166m

s t 28000s m1 800kg

Fthrust1

v 1

t

m1

Fthrust1 0.012N

Isp 30000.s

g0 9.81m

s2

mf1

m1

e

v 1

Isp g0

mf1 799.999kg

m 1 m1 mf1

m 1 0.001kg


78 | P a g e

ORIGIN 0

Remainingmass m0

800

i 0 1 1040

mi 1

mi

e

v 1

Isp g0

mi

i 0 1040for

m

Fuelused m0

800

i 0 1 1040

mi 1

mi

e

v 1

Isp g0

mi

mi

mi 1

mi

i 0 1040for

m


79 | P a g e S e p t e m b e r 2 0 1 5

30 weeks remaining fuel per week in kg 30 weeks fuel used per week in kg

Fuelused

0

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

29

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

-31.132·10

...

Remainingmass

0

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

800

799.999

799.998

799.997

799.995

799.994

799.993

799.992

799.991

799.99

799.989

799.988

799.986

799.985

799.984

799.983

799.982

799.981

799.98

799.978

799.977

799.976

799.975

799.974

799.973

799.972

799.971

799.969

...

Remainingmass480

799.457 kg


80 | P a g e

Station Keeping

Ion propulsion system RIT 10 EVO

. Fuel: xenon

Isp: 3000s

power: 435 W

Burn schedule and burn time: every week every month for 52 weeks for a year equals 1040 burn maneuvers

Total fuel used

v 1 .384m

s t 28000s m1 800kg

Fthrust1

v 1

t

m1

Fthrust1 0.011N

Isp 30000.s

g0 9.81m

s2

mf1

m1

e

v 1

Isp g0

mf1 799.999kg

m 1 m1 mf1

m 1 0.001kg


81 | P a g e S e p t e m b e r 2 0 1 5

ORIGIN 0

Remainingmass m0

800

i 0 1 1040

mi 1

mi

e

v 1

Isp g0

mi

i 0 1040for

m

Fuelused m0

800

i 0 1 1040

mi 1

mi

e

v 1

Isp g0

mi

mi

mi 1

mi

i 0 1040for

m


82 | P a g e

30 weeks remaining fuel per week in kg 30 weeks fuel used per week in kg

Fuelused

0

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

29

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

-31.044·10

...

Remainingmass

0

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

800

799.999

799.998

799.997

799.996

799.995

799.994

799.993

799.992

799.991

799.99

799.989

799.987

799.986

799.985

799.984

799.983

799.982

799.981

799.98

799.979

799.978

799.977

799.976

799.975

799.974

799.973

799.972

...

Remainingmass1040

798.915 kg


83 | P a g e S e p t e m b e r 2 0 1 5

Reference Documents


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89 | P a g e S e p t e m b e r 2 0 1 5


90 | P a g e

\


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92 | P a g e


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96 | P a g e

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2006 IEEE Aerospace Conference. Web. (7) Shin, Dong K. "201 Frequency and Channel Assignments." JPL, 15 Dec. 2014. Web. (8) Young, Michael J. "Understanding Decibels and Their Use in Radio Systems." Web. (9) Young, Michael J. "Planning a Microwave Radio Link." Web. (10) Kuhn, J. Darcey. "7.1 Antenna Sizing and Communication Issues." Web. (11) Silva, Juan J., and Pilar Romero. "Optimal Longitudes Determination for the Station Keeping of

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Phys Org. 18 Feb. 2015. Web. (17) "Digital Transmission: Carrier-to-Noise Ratio, Signal-to-Noise Ratio, and Modulation Error Ratio." (2012).

Web. (18) "Small Deep-Space Transponder (SDST) - Reliable X-Band and Ka-Band Deep Space Transmission." Web. (19) "Ion Propulsion Systems." Space Propulsion. Airbus Defence & Space. Web. (20) "Coarse Si-Axis Sun Sensor." GNC Sensors & Actuators. Airbus Defence & Space. Web. (21) "CT-602 Star Tracker." Ball Aerospace & Technologies Corp. Web. (22) "Mars Reconnaissance Orbiter Mission Overview." NASA JPL. Web. (23) "Earth to Mars Calendar." Clowder. Web. (24) "Stage II Centaur Upper Stage Booster - Mars Reconnaissance Orbiter." NASA JPL. Web. (25) Marshall, Joseph R. "SpaceWire Satellite Usage." (2013). Web. (26) Rennels, D.a. "Architectures for Fault-tolerant Spacecraft Computers." Proceedings of the IEEE 66.10

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