Title: Executive Summary: Mission To Uranus Classification: Unclassified Issue: 1 Date: 15/09/2014
Contact: University of Leicester Department of Physics & Astronomy Space Research Centre University Road Leicester LE1 7RH UK
Chris Greenaway, Jack Boughton, Prepared by: Date: 15/09/2014 Nils Dittel
Reviewed by: N/A Date: N/A
Approved by: N/A Date: N/A
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DOCUMENT CHANGE DETAILS Issue Date Page Description Of Change Comment 1 15/09/2014 N/A – First issue
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TABLE OF CONTENTS 1.1 Reference Documents ...... 1
1.2 Work Breakdown ...... 5
2 The Icy Giant...... 6
2.1 Science Objectives and Justification ...... 6
2.1.1 The atmosphere ...... 6
2.1.2 The interior of Uranus ...... 7
2.1.3 The Rings of Uranus ...... 7
2.1.4 The unusual magnetic field: ...... 8
2.2 Icy Giant Matrix ...... 11
3 The Uranian System ...... 13
3.1 Science Objectives and Justification ...... 13
3.1.1 The Natural Satellites ...... 13
3.1.2 Natural Satellite Observation Priority Classification ...... 17
3.2 The Uranian System Matrix ...... 18
4 Model payload ...... 19
4.1 IR mapping Spectrometer ...... 20
4.2 Radio science experiment ...... 20
4.3 Microwave Radiometer ...... 21
4.4 UV Spectrometer ...... 22
4.5 Narrow Angle Camera ...... 23
4.6 Wide Angle Camera ...... 24
4.7 Magnetometer ...... 24
4.8 Plasma Particle Detector ...... 25
5 Mission Design ...... 27
5.1 Reference Mission ...... 27
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5.2 Analysis ...... 28
6 Launch Vehicle ...... 32
6.1 Overview & Preliminary Selection ...... 32
6.2 Characteristic Energy Analysis...... 33
6.3 Characteristic Energy Analysis Results ...... 36
6.4 Summary ...... 37
7 Propulsion ...... 38
7.1 Requirements ...... 38
7.2 General Architecture Options ...... 38
7.3 Analysis ...... 40
8 Attitude Control System ...... 44
8.1 Requirements and Attitude Determination ...... 44
8.2 Reaction Wheel(s)...... 44
8.3 Thrusters ...... 46
8.4 Overview ...... 48
9 Command and Data Handling ...... 49
9.1 Telecommand message standards ...... 49
9.2 Instrument Processing Requirements ...... 50
9.3 compression ...... 51
9.4 ICER image Compression...... 52
9.5 Digital Modulation and forward Error Correction ...... 53
9.6 Processor and data storage ...... 54
10 Communications ...... 56
10.1 Mission Communications overview ...... 56
10.2 Communication Subsystem Requirements and Constraints ...... 57
10.2.1 Telemetry Requirements ...... 57
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10.2.2 Tracking Requirements ...... 57
10.2.3 Command Requirements ...... 57
10.2.4 Fault Protection and Reliability Requirements ...... 58
10.3 RF Vs Optical Trade-Off ...... 58
10.4 Ground stations and Downlink Window ...... 59
10.41 Ground station network choice ...... 59
10.42 Deep Space Network Coverage ...... 60
10.43 Spacecraft orbit...... 60
10.5 Data Generation ...... 61
10.6 Typical Science Orbit Breakdowns ...... 62
10.7 Link Budget ...... 65
10.71 Introduction ...... 65
10.72 The Link Equation ...... 65
10.73 Detailed Link Budget Results ...... 66
10.74 Summary ...... 67
10.8 Antenna size - Pointing Error: parametric study ...... 67
10.9 X-band Vs Ka-band architecture ...... 68
10.10 Subsystem Architecture ...... 69
10.11 Mission Communication Phases ...... 70
10.12 Launch and Initial Acquisition ...... 70
10.13 Cruise ...... 70
10.14 Uranus arrival and Science Orbits ...... 70
11 Radiation ...... 71
11.1 Radiation Sources ...... 71
11.2 Damage to digital systems ...... 71
11.3 Mission environment ...... 72
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12 Mass and Power ...... 73
12.1 Preliminary Design System Requirements ...... 73
12.2 Power and Mass Tables ...... 74
12.3 Methodology of Mass Reduction ...... 76
12.3.1 Trade Off ...... 77
12.4 Power and ASRG 241-Am Subsystem ...... 79
12.4.1 Power reduction ...... 84
12.4.2 Worst case scenario for the Power Subsystem ...... 85
12.5 Reference Orbit ...... 85
12.6 Battery ...... 86
13 Thermal ...... 87
13.1 The Case Scenario ...... 87
13.2 Earth Orbit ...... 88
13.2.1 Analysis of the Results ...... 88
13.3 Transfer Phase ...... 90
13.4 Uranus Orbit ...... 91
13.5 Alternative Thermal Power Requirements ...... 93
13.6 Thermal Subsystem Summary ...... 93
14 Architecuture ...... 95
TABLES
Table 1 - Work Breakdown by author ...... 5
Table 2 - The expected surface composition of Uranus’ moons [based on the book Astrobiology, Future Perspective by P. Ehrenfreund (2005)] ...... 13
Table 3 - Names, physical and orbital characteristics of the regular satellites of Uranus. [Values are based on Table 1.8 of Schmude (2008)] ...... 13
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Table 4 - The Natural Satellites Science Objective Traceability Matrix [table based on (Greeley & Dougerty, 2010)] ...... 18
Table 5 – Model Payload ...... 19
Table 6 – IR Mapping Spectrometer Bands ...... 20
Table 7 - Orbiter ΔV Requirement Table ...... 38
Table 8 - R-42DM Thruster Properties ...... 40
Table 9 - XLR-132 Thruster Properties ...... 42
Table 10 - Attitude Determination Breakdown ...... 44
Table 11 - Reaction Wheel Properties ...... 46
Table 12 - MR-50T Properties ...... 47
Table 13 - Attitude Control System Breakdown ...... 48
Table 14 – Science Payload data calculations ...... 50
Table 15 - Data volumes as ICER, Rice and Viterbi techniques are applied ...... 54
Table 16 - RF Vs Optical Trade-Off Summary ...... 58
Table 17 - DSN Locations ...... 60
Table 18 - Science Phases Data Generation ...... 61
Table 19 - Option 2 Data rates ...... 62
Table 20 - Option 3 Data Rates ...... 63
Table 21 - Option 4 Data Rates ...... 63
Table 22 - Option 5 Data Rates ...... 64
Table 23 - Downlink Ka-band (34-m DSN, S/C HGA) ...... 66
Table 24 - Link Budget Summary ...... 67
Table 25 - Communications Sub-System Architecture ...... 69
Table 26 - Model Payload characteristics for the Uranus Orbiter (Hubbard, 2010; Dougherty, et al., 2011; Spencer, 2010)...... 74
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Table 27 - Estimated Equipment and Subsystem characteristics for the Uranus Orbiter (Hubbard, 2010; Dougherty, et al., 2011; Spencer, 2010)...... 75
Table 28 - Mass Reduction Possibilities...... 78
Table 29 - Mass differences between the trade-off options...... 78
Table 30 - Power Source Selection ...... 80
Table 31 - Driving Factors for Trade-off ...... 81
Table 32 - Power source trade-off results ...... 81
Table 33 - Power Operation Scenarios (Hubbard, 2010)...... 82
Table 34 - Instruments Power Operation Scenarios (Hubbard, 2010)...... 83
Table 35 - Science Case Power Operation Scenarios (Hubbard, 2010)...... 83
Table 36 - Reference Orbit Power Requirements ...... 86
FIGURES Figure 1 - Pressure and temperature profile of the Uranian atmosphere, credit (Lindal et al, 1987) ...... 6
Figure 2 - Interior model of Uranus, credit (Uranus pathfinder, 2011) ...... 7
Figure 3 - The rings and moons of Uranus (Wikipedia) ...... 8
Figure 4 – Uranus’ Magnetosphere, credit (Balogh 2010) ...... 9
Figure 5 - The Largest 5 moons of Uranus (Darling, 2014) ...... 14
Figure 6 - ”Phebus inner view” (Chassefiere & Maria, 2010) ...... 23
Figure 7 - Uranus Transfer Database Graph ...... 27
Figure 8 - Database Reference Trajectory ...... 28
Figure 9 - GMAT Approximation of Transfer ...... 28
Figure 10 - Science Orbit ...... 30
Figure 11 - Perturbations due to Uranus' Moons...... 31
Figure 12 - Falcon Heavy Render, Credit: SpaceX Website ...... 32
Figure 13 - Soyuz Fregat, Credit: ESA Website ...... 33
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Figure 14 - Soyuz Fregat Characteristic Energy ...... 34
Figure 15 - Characteristic Energy Graphs ...... 36
Figure 16 - Pressure-fed Rocket Systems, (Wertz & Larson, 1992) ...... 39
Figure 17 - R42DM, Credit: Aerojet ...... 41
Figure 18 - XLR-132, Credit: Rocketdyne ...... 42
Figure 19 - MW1000 Reaction Wheel, ...... 46
Figure 20 - MR-50T, Credit: Aerojet ...... 47
Figure 21 - Schematic of a typical source coder, [CCSDS 121.0-B-2]...... 52
Figure 22 - A sequence of images showing gradual increase in image quality under progressive compression as more data is received. [A. Kiely et al. 2003] ...... 52
Figure 23 - QPSK constellation diagram (1) ...... 53
Figure 24 - QPSK BER Vs Eb/N0 performance (1) ...... 54
Figure 25 – RAD750 processor, BAE Systems ...... 55
Figure 26 - Data Flow ...... 56
Figure 27 - Arial view of the Canberra DSN complex (Credit: NASA) ...... 59
Figure 28 - DSN Coverage, NASA ...... 60
Figure 29 - Pointing Offset Vs Eb/N0 for varying dish diameter ...... 68
Figure 30 - Antenna dish and Spacecraft Configuration ...... 87
Figure 31 - Transient Thermal Analysis for Dish during the Earth Orbit. Temperature decreases from 300K to 281.76K...... 89
Figure 32 - Transient Thermal Analysis for Spacecraft during the Earth Orbit. Temperature varies between 300K and 300.63K...... 89
Figure 33 - Transient Thermal Analysis for Dish during the Transfer phase. Temperature decreases from 281.76K down to 173.76K...... 90
Figure 34 - Transient Thermal Analysis for Spacecraft during the Transfer phase. Temperature decreases from 300.63K down to 283K...... 91
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Figure 35 - Transient Thermal Analysis for Antenna-dish during the Uranus Orbit. Temperature decreases from 173.2K to 167.96K...... 92
Figure 36 - Transient Thermal Analysis for Spacecraft during the Uranus Orbit. Temperature decreases from 283K to 274K...... 92
Figure 37 - Alternative Antenna Dish and Spacecraft Configuration ...... 93
Figure 38 – CAD model representing basic architecture ...... 95
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1.1 REFERENCE DOCUMENTS The following documents are referenced for supporting information and are referred to as RDx in the text:
RD Title /Author Document number 1. Science Requirement Matrix.pdf Chris Greenaway, Jack Boughton, Nils Dittel
Alexander, A. (1965) “The Planet Uranus”, Faber and Faber Ambrosi, R., 2014. 7051 Lecture Notes, Leicester: s.n. Angelopoulos, V. (2008). "The THEMIS Mission." Space Science Reviews 141(1-4): 5-34. Arridge, C., 2013. The Science Case for an Orbital Mission to Uranus, London: s.n. Arridge, C. S.; Agnor, C. B.; André, N. (2011). “Uranus Pathfinder: exploring the origin and evolution of Ice Giant planets” London: Springer. Balogh, A. (2010). "Planetary Magnetic Field Measurements: Missions and Instrumentation." Springer Science + Business Media. Behannon, K. W., et al. (1987). "The magnetotail of Uranus." Journal of Geophysical Research: Space Physics 92(A13): 15354-15366. Bennett, G. L., 2006. Mission of Daring: The General-Purpose Heat Source Radioisotope Thermoelectric Generator. California: 4th International Energy Conversion Engineering Conference and Exhibit (IECEC).
Bennett, J.; Donahue, M.; Schneider, N.; Voit, M. (2010). “Astronomie” Munich: Pearson Studium Bin Mazlan, M. A. I., 2009. Development of Lithium Ion Power System. s.l.:s.n.
Bly, V. & Di Pietro, D., 2007. Enceladus Flagship Mission Concept Study. Goddard: NASA.
Brown, M., 2009. Trojan Tour Decadal Study, s.l.: NASA .
CCSDS, lossless data compression standards, [CCSDS 121.0-B-2] May 2012 Chassefiere, E.; Maria, J.L. (2010) “Phebus: A double ultraviolet spectrometer to observe Mercury's exosphere”, Paris: E;sevier Darling, D., 2014. The Worlds of David Darling. [Online] Available at: http://www.google.de/imgres?imgurl=http%3A%2F%2Fwww.daviddarling.info%2Fimages%2FUranus_moons.jpg&im grefurl=http%3A%2F%2Fwww.daviddarling.info%2Fencyclopedia%2FU%2FUranusmoons.html&h=340&w=400&tb nid=3W8j3R9kxx0nyM%3A&zoom=1&docid=Tld_qRFRWDR5mM&ei=dbvDU8 [Accessed 14 07 2014]. Deep space network link design handbook, NASA, Document number 810-005 Delamere. A et al, MRO high resolution imaging science experiment (hirise): instrument Demircioglu. E, Nefes. M, Reliability-based TT&C subsystem design methodology for complex spacecraft missions, 2008 Development, 2003
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Divine, N., and H. B. Garrett (1983), Charged particle distributions in Jupiter's magnetosphere, J. Geophys. Res., 88(A9), 6889–6903 Dougherty, M.; Grasset, O.; Bunce, E. (2001) “Juice Yellow Book”. E. Miller et al. The visual and infrared mapping spectrometer for Cassini. JPL Ehrenfreund, P.; Irvine, W.; Owen, T. (2005) “Astrobiology: Future Perspective” New York: Kluwer Academic Publisher Error bounds for convolutional codes and an asymptotical optimum decoding algorithm, Viterbi. A.J, 1967 Esposito, L. W. (2002). "Planetary rings". Reports On Progress In Physics 65 (12): 1741–1783. Europa-Study-Team, 2012. Europa Study 2012 report. California: NASA.
Karkoshka, Erich (2001). "Photometric Modeling of the Epsilon Ring of Uranus and Its Spacing of Particles". Icarus 151 (1): 78–83. Fieseler. P. D, Shadan M. Ardalan, Dr. Robb A. Frederickson, The Radiation Effects on Galileo Spacecraft Systems at Jupiter, JPL 2002 Fuqin Xiong, Digital Modulation Techniques, Artech House, May 2006 Gilmore, D. G., 2002. Spacecraft Thermal Control Handbook. 1st ed. Virginia: AIAA.
Greeley, R.; Dougherty, M. (2010). “Europa Jupiter System Mission (EJSM)”. s.l.:Nasa & ESA. Guillot, T. (2005). “The interiors of giant planets: models and outstanding question”. Ann. Rev. Earth Planet. Sci. 33, 493–530 Hamme, H.B. et al. (2009). “The Dark Spot in the atmosphere of Uranus in 2006: Discovery, description, and dynamical simulations” Icarus 201, 257-271 Hanley, T.R. (2008). “The Microwave Opacity of Ammonia and Water Vapour: Application to Remote Sensing of the Atmosphere of Jupiter” ProQuest, page 21. Hubbard, W. B., 2010. Ice Giants Decadal Study by. arizona: National Aeronotics and Space Administration .
Irwin, P. (2003) “Giant Planets of our solar system”, Springer-praxis Jankowski, G. D., (1988). “Solid-State Ice Volcanism on the Satellites of Uranus”. Cornell: Science Karttunen, H.; Kröger, P.; Oja, H. (2003). “Fundamental Astronomy”. Helsinki: Springer
Kiely. A and Klimesh. M, The ICER Progressive Wavelet Image Compressor, IPN progress report 2003 Klaus. J. D, Radio Astronomy, Cygnus-Quasar 1966 Kusnierkiewicza, D. Y., Hersmana, C. B., Guoa, Y. & Kubotaa, S., 2005. Adescription of the Pluto-bound NewHorizons spacecraft. Laurel: PERGAMON. Larson, W. J. & Wertz, J. R., 2005. Space Mission Analysis and Design. California: The Space Tecnhonlogy Library. Lindal, G.F. et al. (1987). “The Atmosphere of Uranus' Results of Radio Occultation Measurements with Voyager 2” Journal of geophysical research, vol. 92, no. A13, pages 14,987-15,001
Macgregor. E, Reid. S, Low-Noise Systems in the Deep Space Network, JPL, 2008 Maral and Bousquet, Satellite communication systems, fifth ed. 2009
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Martin, J.W. 1986. Liquid Propellant Management in Space Vehicles. Quest Magazine (TRW, Inc) 9 (1) Martinez, I., 2014. SPACECRAFT THERMAL MODELLING AND TESTING. s.l.:s.n.
Mc Bride, N.; Gilmour, I. (2003). “An Introduction to the Solar System”, Cambridge: The Open University McFadden, J. P., et al. (2008). "The THEMIS ESA Plasma Instrument and In-flight Calibration." Space Science Reviews 141(1-4): 277-302. Mongoose-V Architecture description, Synova website NASA, 2013. Advanced Stirling Radioisotope. s.l.:National Aeronautics and Space Administration.
Ness, N. F., et al. (1986). "Magnetic Fields at Uranus." Science 233(4759): 85-89. NIMS instrument Parameters, http://jumpy.igpp.ucla.edu/~nims/inst/insttab1.html, date accessed 17/08/2014 Norton, J., n.d. Advanced Technology Oscilator for the Pluto Flyby Mission. The Johns Hopkins University Applied Physics Laboratory.
Palumbo, P.; Jaumann, R. (2014) “Janus: the visible camera on board the ESA JUICE mission to the Jovian system”. Berlin: 45th Lunar and Planetary Science Conference Patel, M. R., 2004. Spacecraft Power Systems. Boca Raton: CRC PRESS.
Peale, S. J., “Origin and Evolution of the Satellites”. s.l.:s.n. RAD750 Specification document, BAE systems product catalogue Rowe, D. M. & Abelson, R. D., 2006. Thermoelectrics Handbook, Macro to Nano. Boca Raton: Taylor & Francis .
Schaefer, E. D., Bailey, V. L. & Ercol, C. J., 2008. Spacecraft Packaging. Jon Hopkinson APLTechnical Digest: s.n.
Schmude, R. W. (2008). “Uranus, Neptune, and Pluto and how to observe them”. New York: Springer Showalter, Mark R.; Lissauer, Jack J. (2006-02-17). "The Second Ring-Moon System of Uranus: Discovery and Dynamics". Science 311 (5763): 973–977 Smith, B et al. 1986, Voyager 2 in the Uranian System: Imaging Science Results". Science 233 (4759): 43–64. Spencer, J., 2010. Planetary Science Decadal Survey Titan Saturn System Mission. California: National Aeronautics and Space Administration.
Spilker. L. J, The Cassini-Huygens Mission, JPL october 1997 Sromovsky, L. A.; Fry, P. M. (2005). "Dynamics of cloud features on Uranus". Icarus 179 (2): 459–484. Stanley, S. and G. Glatzmaier (2010). "Dynamo Models for Planets Other Than Earth." Space Science Reviews 152(1- 4): 617-649. Stern, S. A.; Slater, D. C.; Scherrer, J. (2006). “ALICE: The Ultraviolet Imaging Spectrograph aboard the New Horizons Pluto-Kuiper Belt Mission”. Boulder: s.n Stone. E. C, The Voyager 2 encounter with Uranus, Journal of geophysical research 1987 Taylor. J et al, Cassini orbiter Telecommunications, DECANSO January 2002 Tittemore, C.W. (1990). “Tidal Evolution of the Uranian Satellites”. Tucson: Icarus Townsend, G.E. (1962). "Design Guide to Orbital Flight", Chapter 6. Martin-Marietta Verheylewege, A.; Karateki, O.; Noyelles, B. (2014). “Coupled orbital-thermal evolution of Miranda”. Namur: Icarus
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Wallace. G. K, The JPEG still picture compression standard, IEEE Transactions on consumer electronics 1991 Wertz. J and Larson. W, Space Mission Analysis and Design, 3rd ed. 2003 Williams, H. & Bannister, N., 2014. Lecture notes: Systems Engineering and Spacecraft Systems. Leicester: s.n.
Williams, H. R., Ambrosi, R. M. & P, B. N., 2012. A conceptual spacecraft radioisotpe thermoelectric and heating unit (RTHU). Internalional Journal of Energy Research.
Zeilik, M.; Gregory, S. A. (1998). “Introductory Astronomy & Astrophysics”. s.l.:Brooks/Cole Zupella, P.; Corso, A.J. (2012). “Optical subsystems calibration and derived radiometric instrument response of the Phebus spectrometer on board of the BepiColombo Mission”, Padova: Jinst
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1.2 WORK BREAKDOWN Table 1 outlines the contribution of each member to the executive summary by highlighting the sections which each member completed. Sections 2 and 4 were a combined effort of the group, and in these cases the individual sub-sections are provided with the author.
Table 1 - Work Breakdown by author Chapter Contributors Comments
2 – The Icy Giant Boughton, J.; Greenaway, C. Section 2.1.1 – Greenaway, C. Section 2.1.2 – Greenaway, C. Section 2.1.3 - Greenaway, C. Section 2.1.3 – Boughton, J. Section 2.2 – Boughton, J.; Greenaway, C.
3 – The Uranian System Dittel, N.
4 – Model Payload Boughton, J.; Dittel, N.; Section 4.1 – Greenaway, C. Greenaway, C. Section 4.2 – Greenaway, C. Section 4.3 – Greenaway, C. Section 4.4 – Dittel, N Section 4.5 – Dittel, N Section 4.6 – Dittel, N Section 4.7 – Boughton, J. Section 4.8 – Boughton, J.
5 – Mission Design Boughton, J.
6 – Launch Vehicle Boughton, J.
7 - Propulsion Boughton, J.
8 - ACS Boughton, J.
9 –Command and Data Handling Greenaway, C.
10 - Communications Greenaway, C.
11 - Radiation Greenaway, C.
12 – Mass and Power Dittel, N.
13 - Thermal Dittel, N.
14 - Architecture Boughton, J.
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2 THE ICY GIANT
2.1 SCIENCE OBJECTIVES AND JUSTIFICATION
2.1.1 The atmosphere When attempting to understand the evolutionary mechanisms and formation processes of Uranus and other Ice Giants, accurate knowledge of bulk composition and internal structure is incredibly important. A major goal of this mission will therefore be to provide the scientific community with unparalleled compositional data in both the atmosphere and deep interior of Uranus. When looking at how these processes differ from those of the Gas Giants we can gain a unique look into the development of the early solar system (Guillot, T, 2005). The compositional data that this mission will return will also help interpret telescopic observations of exoplanets, in particular their ice to rock ratio (Arridge et al 2011). The processes of energy and material transport around Ice Giants are another area where our knowledge can be greatly improved. As a consequence of this the mission will study the weather dynamics of Uranus’s atmosphere in great detail, including both small scale convective events and planet wide circulation. This can also be connected to the Uranian system by providing us with a greater understanding of how the atmosphere of an ice giant affects its nearby environment, while also revealing the differences between the Gas and Ice Giants (Arridge et al 2011). The atmosphere of Uranus is known to be comprised of primarily hydrogen and helium, with hydrocarbons and “ices”, such as water ammonia or methane at depth. Much of our knowledge has roots in the voyager 2 flyby, which to date is the only craft to study it in detail. As shown in figure 1 temperature and pressure in the upper atmosphere is extremely low and so there are few heavy gases in the stratosphere.
The three main layers of the atmosphere are the troposphere, between -300 and 50km altitude (negative altitudes are locations below the nominal surface at 1 bar), the stratosphere between 50 and 4000km and the exosphere that starts at 4000km and extends for many Uranian radii. Four cloud layers are thought to exist in the troposphere. Methane clouds, Hydrogen sulphide and ammonia clouds, ammonium hydrosulphide clouds and water clouds. Only the first two have been directly observed.
Figure 1 - Pressure and temperature profile of the Uranian
atmosphere, credit (Lindal et al, 1987)
The atmosphere is known to have a very unique weather system that is heavily influence by both the tilt of Uranus’s axis and the lack of an internal heat source. The axial tilt causes large climate variations by season and the lack of an internal heat source means that atmospheric activity is limited in comparison to Neptune. Despite the limited activity cloud tracking has revealed zonal winds in the upper troposphere and retrograde winds at the equator (Sromovsky et al. 2005).
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2.1.2 The interior of Uranus Completing an accepted model for the interior of Uranus clearly is of great importance for any Uranian mission. There are numerous techniques available to help us understand the internal composition of Uranus, which is well below the range of remote sensing instruments. Of these, some of the most common are analysis of gravitational field data, estimation of a planets moment of inertia and analysis of magnetic fields (Irwin, 2003). The small radii of the Ice Giant planets, in comparison to the Gas Giants, tells us that although we observe hydrogen and helium as the bulk of the molecules in the observable atmosphere, the density of the planets would be far too low if they were the predominant elements in the planet as a whole (Alexander, 1965). It was suggested that the appropriate density could be found if Uranus consisted of a Hydrogen atmosphere enveloping a rocky core, however this would have a moment of inertia far smaller than observed (Irwin, 2003).
Currently, the interior of Uranus is known to exist of mainly “ices” such as H20, NH3 and CH4, along with rocks, H2 and He. The best observations put an upper limit on the ice/rock mass fraction at 85% however there are many conflicting models. Most models predict a rocky core, however there are no theories which agree with all observations and there are large gaps in our knowledge. One such model is shown in figure 2, which is consistent with magnetic field and gravity measurements however does not match the low luminosity of Uranus (Arridge et al 2011).
Figure 2 - Interior model of Uranus, credit (Uranus pathfinder, 2011)
An accurate and universally accepted model for the interior of Uranus will place strong conditions on theories of the evolution of the Uranian system but also the solar system as a whole. Furthermore, this knowledge can be applied to our understanding of Neptune and also other ice Giant exoplanets. This mission will therefore have numerous science goals based around bettering our understanding in this area.
2.1.3 The Rings of Uranus Thirteen rings have been directly observed around Uranus and they have varying properties from wide and dusty to narrow and dense. The radius of the closest ring is approximately 38,000km and the largest approximately 98,000km. They are designated 1986U2R/ζ, 6, 5, 4, α, β, η, γ, δ, λ, ε, ν and μ (Showalter M. 2006).The composition of the rings is largely unknown due to voyager 2 not having the ability to view them in the Infrared, however they are hypothesised to contain large amounts of water ice and darkened radiation processed organics, with the latter explaining their low albedo. This mission will aim to accurately determine the composition of Uranus’s rings, which will place important constraints on any models of planetary system evolution. Overall however, the ring system is thought to contain little dust, with the majority of bodies being 0.2-20m in diameter. The relatively small amount of dust has been put down to aerodynamic drag from the extended Uranian exosphere-corona.
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The dynamics of the rings, their subsequent interactions with the moons and magnetic field are another area in which this mission will provide valuable information. Since the voyager 2 flyby in 1986 numerous changes have been observed and so the rings must be inherently unstable. The knowledge of what causes the rings to migrate and change when they interact with the nearby environment are of huge importance when attempting to build an accurate model of planetary evolution. It was previously assumed that every narrow ring had a pair of shepherd moons, however voyager observed that only one ring has such a pair, Cordelia and Ophelia around the ε ring. A large physics problem is currently unanswered due to this, as the rings would quickly spread out radially without such a mechanism to anchor them. The mission will initially focus on the narrow main rings, particularly the ε ring, δ ring and γ ring. The ε ring is the brightest and densest of all rings thought to be responsible for two thirds of all light reflected by the system [Smith, B et al. 1986]. The γ ring is geometrically similar to the ε ring except has no observed shepherd moons. Identifying the method of confinement for this ring will therefore be an interesting science goal of the mission. The δ ring is another key target for investigation due to Voyagers 2’s imaging experiment failing to resolve it.
Figure 3 - The rings and moons of Uranus (Wikipedia)
2.1.4 The unusual magnetic field: The shape of any planetary magnetosphere is defined by the relative orientation of the magnetic dipole axis, the planetary spin axis and the ecliptic (and therefore the direction of the solar wind). The only data on the magnetic field of Uranus stems from the Voyager 2 flyby in 1986, the results of which are detailed by (Ness, Acuna et al. 1986). This single flyby revealed a lot about the curious nature of Uranus’ magnetic field, with the best fit to the data indicating a magnetic dipole axes offset from the rotation axis by 60° and offset spatially by 0.3 RU (1 RU = 25,559km), as shown in figure 1. The maximum field strength detected during the flyby was 413nT, near the closest approach of 4.19 RU; the magnetic dipole moment is estimated to be approximately 50 times that of Earth. The aforementioned tilt means that the contours of the magnetopause and bow shock vary with time as the planet and field rotate, with a period estimated to be 17.29±0.1 hours.
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Figure 4 – Uranus’ Magnetosphere, credit (Balogh 2010) Despite this peculiarity the magnetic axis is always close enough to being orthogonal to the ecliptic that a bow shock and a fully developed magnetic tail with similar geometries to that observed at Earth is maintained. While similar to the magnetotail of Earth there are some major differences, most notably the rotation about the longitudinal axis as described, but also the resultant twisting of the tail with a helical pitch found to be 5.5° ±3.0° (Behannon, Lepping et al. 1987), though this effect does not seriously result in major deformation of the overall shape. Specifically, the bow shock is thought to be located at ~23 RU and the magnetopause at ~18 RU with the tail extending some 390RU or more.
The magnetic field of Uranus is not thought to be as heavily dominated by the dipolar magnetic field as it is for more familiar fields like those of Jupiter or Earth. The explanations put forward for this are summarised by (Stanley and Glatzmaier 2010), but none of the models currently existing are fully supported by observation. The exact origin of the magnetic field is a difficult question, but most explanations centre on the motion of salty and/or metallic liquid motion close to the core, though some theories suggest that such motion may be occurring closer to the surface.
The principal science objective regarding the unusual magnetic field of Uranus is to map the field in detail over a long period. While direct measurements of the field strength itself are of course a crucial element of this the interaction with the solar wind is a major factor in the shaping of said field As such the electron and proton fluxes, which are the main constituent parts of the solar wind would be enormously beneficial in determining the shape of the field as well as in identifying regions of plasma trapped by the magnetosphere.
In all other observed magnetospheres plasma co-rotation and convection interfere with each other, but during the time of the Voyager flyby the co-rotation and convection directions were orthogonal and so were not interacting. Co-rotating plasma is that which is close to the planet and rotating with it while plasma convection describes the motion of plasma perpendicular to both the direction of the magnetic field (out of the ecliptic for Earth) and to the electric field imposed by the solar wind interaction, which is from dawn to dusk, resulting in a sunward motion. Again like Earth, it is believed that Uranus has a partial ring of hot plasma surrounding it, at a radius of approximately 5 RU (Herbert, 2009; Bridge et al., 1986), with plasmas in the tens of keV.
Identifying the location and intensity of these regions of trapped plasma is a core element of the second science goal related to Uranus’ magnetic field: to characterise the interactions between the magnetic field and the Uranian satellites, both the moons and the rings. While the vast majority of plasma in the Uranus system is made up of ionised Hydrogen it is also expected that some neutral particles would also be present, especially in regions close to the moons. This is due
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to a combination of sputtering effects as a result of impacts of other plasmas on the moon surfaces, though also potentially as a result of sublimation of surface ices.
As described in the preceding section it is believed that the relatively dark material evident on the rings and moon surfaces is due to this trapped plasma; demonstrating that the intensity of these regions relates to this darkening effect would go a long way towards proving this. Further information would also be needed regarding the composition of the rings and moon surfaces to show whether or not those materials support the darkening theory.
Accurately mapping Uranus’ magnetic field and documenting any changes over time would be useful in determining how the field is generated and hence give clues to the internal structure of the planet. Due to the very short duration of the Voyager 2 flyby the only information currently available regarding the magnetic field and its effects comes from a single slice of the orbit. While for most planets this would be fairly representative of the situation for an entire orbit the curious alignment of Uranus’ rotation axis means that different stages may exhibit different behaviour of the magnetic field, or more specifically the boundary where it interacts with the solar wind.
The ideal would be to have an operational spacecraft providing magnetic field measurements constantly throughout at least a quarter of a Uranus year to provide information of how the magnetic field changes with time as the solar pointing changes, but that would mean a somewhat unfeasible 21 year mission lifespan after transit period. While that would be somewhat foolish to plan for it would be sensible to choose a launch window which would deliver the spacecraft to Uranus at a time when the rotation axis of the planet is close to perpendicular to the solar facing vector; the opposite state to that during the Voyager 2 flyby.
Further characterisation of Uranus’ magnetic field would as has been stated give clues about the internal structure of the planet, but also the solar system’s other icy giant Neptune which shares many of the peculiarities of magnetic field. Additionally, many of the extrasolar planets thus far detected are of similar size and so potentially share other properties with these giants so a better understanding of Uranus may well extend out to benefit our understanding of more distant worlds.
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2.2 ICY GIANT MATRIX Theme Science Objective Science Investigation Determine the UA.1 Determine the bulk elemental abundances present in the atmosphere of Uranus atmospheric structure and composition of UA.2 Determine the mixing ratio and abundances of elements at different layers within the atmosphere Uranus UA.3 Determine the temperature profile of the atmosphere from 1 to 200 bar UA.4 Determine the pressure profile of the atmosphere from 1 to 200 bar Investigate UD.1 Investigate and the speeds and locations of Uranus's zonal winds atmospheric Zonal UD.2 Track methane clouds throughout the troposphere of Uranus winds and weather dynamics. UD.3 Investigate the methods of material and energy transport in the atmosphere UD.4 Locate and monitor weather events in the atmosphere such as storms UD.5 Investigate wind speeds as a function of latitude UD.6 Determine the thermodynamics of atmospheric phenomena The Icy Giant Icy The Determine the UI.1 Determine the bulk elemental abundances present in the interior of Uranus internal structure and UI.2 Determine the structure of Uranus's interior composition of Uranus UI.3 Investigate the existance of a rocky core Determine the UR.1 Determine the bulk composition of the rings composition and structure of the rings UR.2 Determine particle density and sizes in the rings
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Determine how the UB.1 Examine the process of the ε ring shepherding and search for other shepherd moons rings behave UB.2 Determine the migration patterns of the rings and changes since the voyager 2 flyby dynamically UB.3 Investigate the nature of ring confinement at Uranus Precision Mapping of UM.1 Characterisation of variations in the magnetic field environment on seasonal timescales the Magnetic Field and Interaction with the UM.2 Characterisation of variations in the magnetic field environment on diurnal timescales Solar Environment UM.3 Determination of solar interaction with magnetic field
The Icy Giant Icy The Investigate UMU.1 Characterisation of distrubances in magnetic field due to Uranian satellites Interaction of the UMU.2 Determination of impact of Uranian satellites on trapped plasma regions Magnetic Field with UMU.3 Investigation of magnetic field interaction with Uranus' rings the Uranian System
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3 THE URANIAN SYSTEM
3.1 SCIENCE OBJECTIVES AND JUSTIFICATION
3.1.1 The Natural Satellites Uranus has approximately 27 natural satellites (2008), which are subdivided into three categories. Firstly, the regular satellites which are Uranus’ five largest moons (Miranda, Ariel, Umbriel, Titania and Oberon). Secondly, thirteen collision fragments (such as Puck, Portia). Finally nine captured objects, including Sycorax and Caliban (Schmude, 2008). Most of these satellites are made of ice, rocks and carbonaceous chrondite. As a result of the Uranian satellites’ large distance to the sun, the ice they contain is likely to possess a large proportion of ammoniac, methane and water (Schmude, 2008; Bennett, et al., 2010). The interiors of Uranus’ larger, regular satellites are not fully understood. While they may simply have a homogeneous composition of carbonaceous chondrite, ice and rock, they could also possess a differentiated composition of these three layers of material (Schmude, 2008; Bennett, et al., 2010).
Table 2 - The expected surface composition of Uranus’ moons [based on the book Astrobiology, Future Perspective by P. Ehrenfreund (2005)]
H2O CH CO2 OH C NH3-hydrate HC
Ariel
Umbriel
Titania
Oberon
Miranda
May have Have
Table 3 - Names, physical and orbital characteristics of the regular satellites of Uranus. [Values are based on Table 1.8 of Schmude (2008)] Name Radius Mass Density Average Distance Orbital Period Inclination (km) (1018 kg) (g/cm3) to Uranus (km) (days)
Titania 789 3480 1.7 436300 8.706 0.1
Oberon 761 3020 1.6 583500 13.463 0.1
Umbriel 585 1290 1.5 266000 4.144 0.1
Ariel 579 1270 1.6 190900 2.52 0.0
Miranda 236 66 1.2 129900 1.413 1.413
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The five largest of Uranus' moons are: Miranda, Ariel, Umbriel, Titania and Oberon. They are moving along Uranus’ “equatorial plane and revolve in the same direction as the planet rotates” (Zeilik & Gregory, 1998). The study of each of these natural satellites could contribute towards a better understanding of smaller solar system bodies, which would lead to an improved insight into the origin of life in our solar system (Arridge, et al., 2011). As these moons are protected from the solar wind by Uranus’ magnetosphere, a different space weathering process acts on these bodies compared to those which are exposed to the solar wind (Arridge, et al., 2011). This means that the magnetospheric as well as tidal interactions have most likely played a significant role in the evolution of the Uranian moons. For example, the tidal heating during periodic passages by orbital resonances could have caused internal melting for some of the Uranian moons (Arridge, et al., 2011).
At present, not much is known about these five natural satellites. We only possess some incomplete, low resolution Voyager 2 images of the southern hemisphere, which provide some clues regarding tectonic processes and geological evolution. Two moons, Oberon and Titania, are of particular interest as they may contain liquid water between their rocky core and their outer ice shell (Arridge, et al., 2011; Arridge, 2013).
A new mission to Uranus could increase our understanding of the tectonic processes and geological evolutions of these moons. This would enable a more in-depth knowledge of the crater size frequency distribution, allowing for more precise age dating of the surfaces. This current study would also lead to a better understanding of the impactors on the surface of the icy moons. Due to the uncertainty regarding the moons’ bulk composition, information relating to their origin is limited (Arridge, et al., 2011; Arridge, 2013). There are many unresolved questions concerning these five major moons as well as their geological history and their previous activity. For example, Ariel displays volcanic and tectonic activity, whereas Umbriel is mainly covered by craters (Bennett, et al., 2010) (Karttunen, et al., 2003). One current question is: why are Ariel and Umbriel so different from each other despite their similar size (Bennett, et al., 2010) (Karttunen, et al., 2003)? Similarly to Ariel and Umbriel, Titania and Oberon also have very different geological activity even though they are of a similar size. Evidently, in order to understand these differences, a new mission to Uranus is urgently required. In the following sections I will outline what we have learned about the moons of Uranus since Voyager 2, and demonstrate where further study would be required (Bennett, et al., 2010).
Figure 5 - The Largest 5 moons of Uranus (Darling, 2014)
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Titania Uranus largest regular satellite is Titania, with a radius of ~789km (Schmude, 2008; Arridge, et al., 2011). Voyager 2 images indicate that Titania has a surface which is heavily cratered (simple and complex), containing fault regions (some of which are 1500km long and 2 to 5km high). It could be argued therefore that Titania is a satellite which has been geologically more active in the past compared to Oberon which is of a similar size (Bennett, et al., 2010; Mc Bride & Gilmour, 2003). On the basis of Voyager 2 images, no lava floodings have been detected inside the fault regions, but cryovolcanic activity is very likely to have occurred in the past (Mc Bride & Gilmour, 2003). The Voyager 2 images were not high quality, each pixel is 3.4km across. This means that craters and features with a diameter smaller than 7km are not detectable. However, the images detected canyons, scarps and faults (Schmude, 2008). The crater size frequency distribution of Titania suggests that its surface is younger than that of Oberon and Umbriel. The surface of Titania is most likely composed of carbon dioxide ice and water ice.
Currently experts are yet to determine whether Titania has an atmosphere. Consequently, this mission could lead to further indication of whether or not Titantia does possess an atmosphere. The absence or the presence of Titania’s atmosphere could place constraints on Titania’s surface composition (Schmude, 2008). The biggest question concerning Titania is whether it contains a liquid-water ocean beneath its outer ice shells. This subsurface ocean could have been created during previous melting events (Arridge, et al., 2011). “The strongly inclined magnetic dipole moment of Uranus with respect to its spin axis generates time-variable fields near the moons at their synodic rotation periods. These fields will produce induction magnetic fields, which are diagnostic of the moons interior, particularly with respect to the possible salty liquid sub-surface oceans in Titania and Oberon” (Arridge, et al., 2011). “The tilt of the magnetic equatorial plane with respect to the ring plane also means that magnetospheric plasma and energetic particle irradiation of the moons varies strongly in time, peaking periodically as the moons cross the magnetic equator” (Arridge, et al., 2011).
Oberon
Uranus second largest regular satellite is Oberon, with a radius of 761km (Schmude, 2008; Arridge, et al., 2011). Oberon and Titania are very similar in size but Titania is geologically more active than Oberon, for unknown reasons(Bennett, et al., 2010). The surface of Oberon is similar to the surface of Umbriel, even though its surface is not as dark (Mc Bride & Gilmour, 2003). From previous Voyager 2 images we know that Oberon’s surface is saturated by shallow craters of simple and complex origins. The shallowness of these craters may have been caused by the icy nature of Oberon’s surface (Schmude, 2008). Furthermore Oberon’s surface displays bright patches, which are believed to be impact craters from a more recent origin (Mc Bride & Gilmour, 2003). It is likely that there has been some strong geological activity on Oberon in the past (Schmude, 2008). Therefore, it is possible that previous melting events could have caused “liquid water oceans beneath” Oberon’s outer icy shell (Arridge, et al., 2011). Previous Voyager 2 images of the surface of Oberon displayed 6 km within 1 pixel, which means that features smaller than 12km cannot be identified. If the same picture resolution would have been applied on Miranda, some of Miranda features would not have been detected, such as its canyons (Schmude, 2008). We know now that Oberon has some larger canyons which are older than those found on the surface of Titania. This suggests that Oberon’s geological activity stopped much earlier than that of Titania (Schmude, 2008). Previous observations revealed water ice on the surface of Oberon but failed to detect carbon dioxide ice (Schmude, 2008).
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Umbriel
Umbriel is the third largest of Uranus' moons and has an unusually dark, heavily cratered surface without any signs of geological activity (Karttunen, et al., 2003) (Mc Bride & Gilmour, 2003). The most remarkable feature on Umbriel is a crater which is 100km in diameter. This crater “is lighter in colour because the impact has penetrated, and excavated material from the subsurface layers, which appear to be of a somewhat different composition from the surface material” (Mc Bride & Gilmour, 2003). It is thought that after Umbriel's surface solidification and formation, it was struck by meteoroids. Since these impacts, parts of Umbriel’s surface underwent resurfacing, which covered some of these craters but not all. Since this partial resurfacing process occurred, further meteoroids impacted on Umbriel, causing further craters to form. Furthermore, Umbriel underwent expansions causing canyons to form across its surface. Previous spectroscopic studies detected carbon dioxide and water ice across Umbriel’s surface (Schmude, 2008).
Ariel
Ariel is the fourth largest of Uranus' moons and has a radius of 579km (Schmude, 2008). Previous Voyager 2 images were able to identify features which are larger than 2km (Schmude, 2008). These images display canyons (Schmude, 2008), endogenic resurfacing (Arridge, et al., 2011), impact craters (Zeilik & Gregory, 1998), large fractures (Zeilik & Gregory, 1998) , valleys (Zeilik & Gregory, 1998), scarps (Schmude, 2008) and faults (Schmude, 2008). Surface images of Ariel have a better resolution than the pictures for Oberon, Umbriel and Titania (Schmude, 2008). The craters found on Ariel's surface are both simple and complex craters with some displaying signs of bright ejecta (Schmude, 2008). Some of Ariel crater’s ejecta blankets are very bright with albedos up to 0.55 (Schmude, 2008). Some of Ariel’s craters, which are between 8-12km in diameter, are quite shallow. This could be due to Ariel’s icy nature or due to its “recent geological activity” (Schmude, 2008).
Previous crater density studies on Ariel reveal that it’s “surface has a wide range of ages” (Schmude, 2008). From Ariel’s crater density study and surface study, astronomers have determined that early in the history of Ariel an endogenic resurfacing occurred, after which the moon expanded further. This endogenic resurfacing is “associated with tectonic systems”, which could have involved cryovolcanic processes (Schmude, 2008) (Arridge, et al., 2011). During the expansion, canyons, scarps and faults were created. The surface continued to be bombarded after the expansion, leaving craters on its surface (Schmude, 2008). Previous studies regarding Ariel suggest that, in the past, “tidally induced melting” was the trigger which provoked Ariel’s resurfacing (Arridge, et al., 2011). Furthermore, Ariel displays “trough-like fractures that break up the impact-scared surface. Many of these fractures will have been subsequently filled by icy lavas, although it is unlikely that Ariel has suffered any major cryovolcanism in the recent past” (Mc Bride & Gilmour, 2003). As can be seen in Table 2, the surface composition of Ariel is carbon dioxide ice and water ice (Schmude, 2008).
Miranda
Miranda is the innermost regular satellite of Uranus, and it is also the smallest, with its 236km in radius (Schmude, 2008) (Zeilik & Gregory, 1998). It is also the moon which is the most peculiar and least understood. Considering its size, we would expect Miranda to be an icy moon covered by craters (Bennett, et al., 2010). However, the images received from Voyager 2 show an extremely exotic and rugged terrain, where multiple geological formations are mixed together. “The surface of Miranda contains four types of terrain, which are the corona, heavily cratered terrain, lightly cratered terrain and terrain dominated by faults, scarps and canyons” (Schmude, 2008). This is rather unique for our solar system, as usually these surface features are found on other moons and planets individually (Mc Bride & Gilmour, 2003). In addition, Miranda is formed in an unusual V-shape, which could have been caused by “vast collisions that
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broke the moon apart; some pieces may have later settled down, inside out” (Karttunen, et al., 2003). Miranda displays processes of endogenic resurfacing at specific areas which can be linked to tectonic systems that could involve cryovolcanic processes (Arridge, et al., 2011). Furthermore these processes seen on Miranda seem to be ongoing (Arridge, 2013). Other areas of Miranda reveal impact craters, which could mean that these surfaces are older. Some of these older craters are covered with ‘snow’ or dust which is believed to be the result of some previous strong cryovolcanic eruptions (Mc Bride & Gilmour, 2003). One unresolved question concerning Miranda is: Why is Miranda so much more geologically active than other moons, e.g. Saturn’s moon Mimas, which has a diameter similar to that of Miranda (Bennett, et al., 2010)?
Crystalline H2O has been identified on the surface of Miranda and further studies suggest that ammonia hydrate could also be found (Schmude, 2008) (Ehrenfreund, et al., 2005). “Miranda has an unusually high inclination (I = 4.338°), and its surface reveals signs of past endogenic activity. Investigations regarding the dynamic aspects of its orbital evolution suggest probable resonant processes, in particular with Umbriel, as an explanation for the present high inclination of Miranda. The tidal heating induced by gravitational interactions can lead to the rise of eccentricities and consequently to the increased dissipation of energy inside the satellite and higher internal temperatures” (Verheylewege, et al., 2014).
3.1.2 Natural Satellite Observation Priority Classification Comparing the natural satellites with one another is essential due to possible difficulties in observing each natural satellite individually. A flyby of Titania and Oberon is preferred “due to the possible presence of internal oceans” (Arridge, et al., 2011). Due to the stronger geological activity on Titania in the past, Titania appears to enhance our studies concerning the evolution and origin of natural satellites to a great extent than Oberon (Bennett, 2006). The third most scientifically interesting natural satellite in the Uranian system is Miranda, as it is the most peculiar and least understood moon. Previous surface images of Voyager 2 show that Miranda has an extremely exotic and rugged terrain, where multiple geological formations are combined. This is rather unique for our solar system, as usually those geological formations are found on other moons and planets individually (Mc Bride & Gilmour, 2003; Schmude, 2008). The forth most scientifically interesting surface is Ariel, as it displays volcanic and tectonic activity while Umbriel is mainly covered by craters (Bennett, et al., 2010) (Karttunen, et al., 2003).
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3.2 THE URANIAN SYSTEM MATRIX Table 4 - The Natural Satellites Science Objective Traceability Matrix [table based on (Greeley & Dougerty, 2010)] Science Natural Satellites Science Investigation Reference Objective Search for MA1 TA1 OA1 AA1 UA1 Investigate the magnitude, nature and distribution of tidal dissipation on the moon. (Arridge, et al., 2011) (Tittmore & C, 1990) present and MA2 TA2 OA2 AA2 UA2 Investigate the moon’s cryovolcanism processes and composition. (Jankowski & G, 1988) past activity (active dynamic MA3 TA3 OA3 AA3 UA3 Determine the moon’s rotation state. (Tittmore & C, 1990) processes) on MA4 TA4 OA4 AA4 UA4 Investigate the moon’s interior. (Schmude, 2008; Bennett, et al., 2010) the moon: How (Schmude, 2008; Bennett, et al., 2010; Arridge, MA5 TA5 OA5 AA5 UA5 Understand the moon’s evolution and origin. did its surface et al., 2011; Mc Bride & Gilmour, 2003) features form? (Schmude, 2008; Bennett, et al., 2010; Arridge, MA6 TA6 OA6 AA6 UA6 Determine regional and global surface ages. et al., 2011; Mc Bride & Gilmour, 2003) Observe processes of deposition and erosion and their influence on the moon’s (Schmude, 2008; Bennett, et al., 2010) MA7 TA7 OA7 AA7 UA7 physical surface properties. Determine the characteristics and formation mechanisms of the moon’s surface (Schmude, 2008; Bennett, et al., 2010; Arridge, MA8 TA8 OA8 AA8 UA8 features (craters, corona, faults, scarps and canyons). et al., 2011; Mc Bride & Gilmour, 2003) Characterise MB1 TB1 OB1 AB1 UB1 Investigate the structure of the moon’s icy shell, with emphasis on its properties. (Ehrenfreund, et al., 2005) the moon’s icy Correlate subsurface structures and surface features to investigate interior and near- (Schmude, 2008; Bennett, et al., 2010; Arridge, shell MB2 TB2 OB2 AB2 UB2 surface processes. et al., 2011; Mc Bride & Gilmour, 2003) Determine the Observe inorganic and organic surface chemistry, as well as distribution and (Schmude, 2008; Ehrenfreund, et al., 2005) ME1 TE1 OE1 AE1 UE1 evolution and abundance of materials. distribution of Relate the properties and composition of the non-water-ice materials to the (Schmude, 2008; Ehrenfreund, et al., 2005) ME2 TE2 OE2 AE2 UE2 the moons’ distribution across the surface. surface (Schmude, 2008; Ehrenfreund, et al., 2005) materials, as ME3 TE3 OE3 AE3 UE3 Determine the moon’s volatile content. well as its global (Schmude, 2008; Ehrenfreund, et al., 2005) ME4 TE4 OE4 AE4 UE4 Investigate the nature and origin of exogenic materials. composition. TF1 OF1 Determine the phases and amplitudes of the gravitational tides (Greeley & Dougerty, 2010; Peale, n.d.) Analyse the induced magnetic fields, which give clues to the moon’s interior. (Arridge, et al., 2011) Subsurface TF2 OF2 Investigate the time-dependent variation of energetic particles and Oceans. magnetospheric plasma irradiation, during the magnetic equator passage. Investigate the rocky mantle, rock-ocean, core compensation and interface of (Schmude, 2008) TF3 OF3 the icy shell.
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4 MODEL PAYLOAD Table 5 – Model Payload Model Instrument Science Contribution Characteristics IR Mapping Composition of the atmosphere, rings and λ =0.5-5 µm Spectrometer moon surfaces. Tracking of cloud features Spectral sampling <5nm and weather events in the troposphere. IFOV ~0.25mrad Temperature mapping. FOV ~3.5o HgCdTe CMOS detector Operating temperature <100K Radio Science Temperature and pressure profiles of the experiment atmosphere. Combined with other USO Allen deviation: instruments can be used to find 1 s - < 10-13 atmospheric composition and mixing 100 s - < 10-13 ratios. 1000 s - < 10-13 Particle sizes and dynamics in the Uranian ring and moon system. 1 way Doppler in Ka band Microwave Composition of the deep atmosphere and Measurements taken in multiple wavelengths ranging Radiometer water content. Temperature measurements from 0.6 to 10GHz using multiple Radiometers with centre frequencies at 0.6, 1.25, 2.6, 5.2, and deeper than that of the IR spectrometer and an accurate vertical profile of water 10.0GHz. content Weighting functions allow for probing of different depths within the atmosphere
Longitudinal footprint of ~10o Plasma Particle Characterisation of solar interaction with Thermal electron and ion detection range: Detector magnetic field. ~3eV-20keV Determination of trapped plasma regions Field of view including entire nadir/beam plane and interaction of such with Uranus’ satellites. UV Spectrometer Composition, magnetosphere, atmosphere Extreme UV, Far UV, Middle UV Spectrometer of Uranus. Chemistry, evolution, origin, λ ~50nm - ~390nm composition and structure of the moons. IFOV: ~0.01mrad FOV: ~ 2° Narrow Angle Camera Geological Processes of the moons. 12 colours filters, Uranus’ cloud dynamics and properties. FOV: ~ 0.3° IFOV: ~ 0.005 mrad <10m/px at 500km <5km/px at 106 km 1024 * 1024 sensors Wide Angle Camera Geological Processes of the moons. 12 colour filters. Uranus’ cloud dynamics and properties. 1024 * 1024 sensor FOV: 10.5° IFOV: 0.179mrad Magnetometer Characterisation of Uranus’ unusual Ranges up to ±20µT, resolution 4pT in smallest magnetic field. range (0 through 1-200nT) Provide details of temporal variation in the Continuous operation. field and verification of internal structure Knowledge of sensor orientation to within 2 models. arcminutes.
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4.1 IR MAPPING SPECTROMETER Science goals and measurements: The instrument will provide spectral imaging of the atmosphere of Uranus, its rings and natural satellites in the visible to thermal IR wavelength range. The ions below: Composition and structure of the Uranian atmosphere, with emphasis on the detection of hydrocarbons and ices (water, ammonia, methane ect.). Cloud tracking and identification of weather events in the troposphere, particularly methane clouds. Compositional measurements of the Rings of Uranus with emphasis on the detection of dark radiation processed organics to explain their low albedo Surface compositional measurements of the satellites of Uranus Performance requirements and concept: Comparable instruments are LEISA on board New Horizons and the VIRHIS instrument on JUICE. LEISA has a spectral range 0.4-2.5 µm and is designed to detect ices on and around Pluto, while VIRHIS is the prime instrument for determining the surface composition of the satellites and rings of Jupiter and has a spectral range 0.4-5.2 µm, with emphasis on organics. An orbital distance of 2Ru and IFOV of 0.25mrad would yield a pixel size of ~3km on the atmosphere of Uranus with ~480 pixels at a total FOV of 3.5 degrees. This is suitable for cloud tracking and the identification of weather events at greater orbital distances, with the largest discovered being the “dark spot” of Uranus that was measured to be around 3000km across (Hamme et al, 2009). A spectral sampling of <5nm is consistent with the flight heritage of similar instruments. The table below gives the diagnostic bands of some important elements/bonds relevant to Uranus which can be detected with the spectral range 0.5-5 µm.
Table 6 – IR Mapping Spectrometer Bands Element Diagnostic bands (µm) H2O Crystalline: 1.04, 1.25, 1.5, 1.65, 2.05, 3.0 Amorphous: 1.04, 1.25, 1.5, 2.0, 3.0 Continuum at 3.6 can be used as an indicator for grain sizes. C-H bonds 1.73, 3.40 C=N bonds 2.42, 4.35, 4.90 CO2 2.02, 2.70, 2.78, 4.25 Tholins 4.57, Visible Slopes Hydrated Minerals 1.40, 1.95
The instrument will use an Offner configuration which will allow for a compact design because of the concentric nature of the layout. The detector will be a HgCdTe (mercury-cadmium-telluride) CMOS multiplexer with amplifiers optimised for speed and noise. The choice of a HgCdTe detector will allow for fine tuning of the band gap and cut-off wavelength, with HgCdTe detectors routinely grown to cut-off wavelengths ranging from 1.7 to 16.5 µm (OPN, June 2008). The cut-off wavelength is changed by varying the mercury-cadmium ratio which follows the form Hg1-xCdxTe (Rogalski, 2005). A choice of a CMOS detector will also reduce the power requirements in comparison to a CCD detector. The majority of observations will be taken at nadir pointing. To reach the acceptable levels of noise the detector will be cooled down to temperatures of <100K, as is standard on other instruments of this type.
4.2 RADIO SCIENCE EXPERIMENT Science goals and measurements: The radio science experiment uses the spacecraft’s high gain antenna and ultra-stable oscillator to produce high stability signals for radio science. The scientific observables are the changes in phase,
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frequency and amplitude as the signal passed through the atmosphere, ionosphere and populations of particles. The goals of the radio science experiment are: Radio occultation’s of the atmosphere of Uranus to obtain temperature, pressure profiles and zonal wind speeds Compositional data of the atmosphere by radio occultation Determination of ring structure to an accuracy of ~1km and the size of ring constituents with an accuracy of ~10m using radio occultation. Performance requirements and concept: Comparable experiments have flown on Cassini, which used a USO to perform radio occultations on the rings of Saturn, and the radio science experiment on JUICE, which will perform both atmospheric and ring occultations. The USO will be of the Quartz crystal resonator design and will create a coherent signal in the Ka radiofrequency band that will be detected by the DSN ground stations for analysis. The main factors involved in the quality of occultation’s are the amplitude and phase stability of the radio signal, and the available signal to noise ratio (Kliore et al, 2002). The signal to noise ratio depends strongly on both the ring opening angle and the antenna pointing, while the amplitude and phase stability are dependent on the stability of the USO. The stability of the USO is affected by the environment including magnetic field, ionizing radiation and acoustic noise vibration. For the atmospheric science, dual frequency signals are important in allowing observers to discriminate between the effects on the signals properties caused by the ionosphere and neutral atmosphere. TO provide data on zonal wind speeds at different latitudes and altitudes, a sufficient amount of occultations are needed at well-spaced latitudes. The voyager 2 spacecraft, which performed the only previous radio occultation on the rings of Uranus, had a USO with -12 a phase stability Allen Deviation σy = 5 x 10 at 1s integration. This limited the resolution to ΔR > 250m (Marouf et al, -13 1986) . The recent Cassini mission to Saturn had a USO stability of σy = 2 x 10 which yielded a resolution limit ΔR > 10m (Kliore et al, 2002), similar to the sizes of particles thought to exist in the rings of uranus.. The radio occultations using a modern USO will therefore provide far greater scientific returns than were possible during the voyager flyby in 1986. For a ring feature of optical depth τ, the signal to noise ratio is reduced because of attenuation from the free space value by the factor exp[ -τ / sin(B) ], where B is the ring opening angle (Kliore et al, 2002). As an example a ring feature of optical depth 1.5, such as parts of Uranus’s ε ring, will have a measurement SNR a factor of 600 higher when viewed at a ring opening angle of 20 degrees than 8 degrees. The choice of orbit for ring occultation is therefore extremely important in maximising SNR and should be examined closely during mission profiling. The spacecraft high gain antenna should be continuously directed towards earth while occultations are being made to further increase the SNR.
4.3 MICROWAVE RADIOMETER Science goals and measurements: The microwave radiometer instrument will probe the deep atmosphere of Uranus. Its science goals are: To provide temperature profiles of the atmosphere down to pressures of +100 bar To measure the abundance of water and ammonia in the atmosphere Measure the vertical profile of water in the atmosphere Performance requirements and concept: A comparable instrument is the JUNO microwave radiometer that will measure the temperature profile and water abundance in the Jovian atmosphere. To obtain temperature and compositional profiles of the atmosphere, multiple wavelengths ranging from 0.6 to 10GHz and weighting functions for each band will be used. The microwave radiometer will use the strong absorption of ammonia and water to determine their abundance (Hanley, 2008). Separate receivers with centre frequencies of 0.6, 1.25, 2.6, 5.2, and 10.0GHz will be used to give almost evenly spaced weighting functions down through the atmosphere. Multiple antenna will be needed which will create a challenge for the instrument engineers when attempting to keep mass low and to direct the microwaves. Each antenna will have a waveguide slot array in its design, much like the instrument flown on JUNO. This is especially true for the large antennas required for the longer wavelengths. The antennas must be designed to function in extremely low temperatures as they are located on the outside of the spacecraft.
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A longitudinal footprint of ~100 will provide the option for good overall coverage of the planet.
4.4 UV SPECTROMETER Scientific goals and measurements: The Ultraviolet Spectrometer instrument will deliver a wide variety of spectral, temporal and spatial observations of Miranda, Ariel, Umbriel, Titania and Oberon, Uranus’ rings, Uranus’ atmosphere, and of course, of Uranus itself. Observations of the Uranian satellites with this instrument will allow us to achieve a better understanding of their evolution, origin, subsurface structure and surface features. Furthermore the UV instrument will allow us to investigate the structure of the moons’ icy shells, with emphasis on their properties. The observation of inorganic and organic chemistry, as well as of the distribution and abundance of materials, will be supported by this UV instrument. Finally it will reveal properties and compositions of surface features and “their distribution to geology” (Dougherty, et al., 2011). Observation of the Uranian rings will allow us to determine their nature, process and bulk composition, and the possibility to detect more shepherd moons (Dougherty, et al., 2011).
The instrument will perform the following tasks: •Monitor surface changes and cryovolcanic activity on Miranda, Ariel, Titania. •Investigate interaction between Uranus’ moons and Uranus’ magnetospheric dynamics and magnetosphere. •Observe surface inorganic and organic chemistry, as well as distribution and abundance of materials. • “Relate compositions and properties and their distribution to geology” (Dougherty, et al., 2011). •Investigate the structure of the moon’s icy shell, with emphasis on its properties. •Correlate subsurface structure and surface features to investigate interior and near-surface processes. •Understand the moon’s evolution and origin. •Observe the surface reflectance of the moons to characterise the surfaces and to map materials which are not water-ice. •Determine their nature, process, bulk composition of the Uranus ring system. •Search for shepherd moons. •Investigate the surface composition variation of Uranus which is linked to its interaction with plasma (Dougherty, et al., 2011). •Observe Uranus aurora and upper atmosphere, upper atmospheric changes and structure (Dougherty, et al., 2011).
Performance requirements and concepts: The selected instrument is similar to Phebus on Bepi Colombo. It covers the majority of the frequencies required to detect the majority of the UV emissions on the Uranian systems. Therefore, this instrument fulfils a wide variety of spectral, temporal and spatial resolution requirements (Dougherty, et al., 2011).
Possible instrument concept: The entire instrument consists of an optical scheme and an electronic unit (“data acquisition, processing and buffering electronics and the power, command and data interface with the spacecraft systems”) (Dougherty, et al., 2011). The optical scheme can be split into two parts (See Figure 6). Firstly, the spectrometer part and secondly the collecting part. The UV imaging spectrometer part, is a double spectrometer whose spectral range is between the Extreme Ultraviolet (50-155nm) and the Far Ultraviolet (145-320nm). Furthermore “a specific near ultra-violet (NUV) detector is implemented” which can detect elements such as Ca and K (400-420nm) (Chassefiere & Maria, 2010). Although this instrument is not capable of detecting short wavelengths, ions such as Mg+, S+ and C+ can still be detected by the Extreme Ultraviolet and Far Ultraviolet instrument (Dougherty, et al., 2011; Stern, et al., 2006; Chassefiere & Maria, 2010). The collecting part consists of three subparts, which are the primary mirror, the entrance slit, and the straylight rejection baffle (Dougherty, et al., 2011; Stern, et al., 2006; Chassefiere & Maria, 2010). “The mirror is a Silicon Carbide (SiC) off-axis parabola”, with a 170mm focal length and a 50°inclination angle (Stern, et al., 2006). The SiC was selected “for its efficiency performance in the whole 55-315nm spectral range and for its mechanical and thermal properties […]. The nominal surface roughness has been specified as 0.5nm RMS in order to minimize the straylight inside the instrument” (Stern, et al., 2006).
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As is shown in Figure 6, the mirror is located between the “baffle exit and accommodated inside a scanner rotating mechanism” (Chassefiere & Maria, 2010).The incoming light is directed towards the “entrance slit of an imaging spectrograph with a reflective holographic diffraction grating” (Dougherty, et al., 2011). “Two holographic gratings share the same mechanical mount and are accommodated in front of the slit” (Chassefiere & Maria, 2010). “The grating disperses the radiation onto the focal plane, where an UV-sensitive microchannel plate detector records the spectrum” (Dougherty, et al., 2011). For Extreme Ultraviolet, a mean grove density of 2700 grooves/mm, and for Far Ultraviolet of 1600 grooves/mm are provided with this instrument (Dougherty, et al., 2011; Chassefiere & Maria, 2010).
Figure 6 - ”Phebus inner view” (Chassefiere & Maria, 2010)
4.5 NARROW ANGLE CAMERA Scientific goals and measurements: The narrow angle camera, also known as NAC, provides high resolution images of Uranus, its moons and its rings. High resolution images of selected moons will improve our understanding of the composition and ages of their surfaces. The previous occurrence of tectonic and cryovolcanic processes on the moons’ surfaces can also be analysed by these high resolution surface images. The analysis of the surfaces may demonstrate the transportation of material from an irregular satellite to a regular satellite, by collisions. As only the southern hemispheres of the moons have been imaged in low to middle resolution, these new images will improve our knowledge of these icy moons. Furthermore this instrument will result in a better insight into Uranus’ rings, allowing us to determine their nature, particle densities and sizes. The narrow angle camera will also indicate the interaction between the rings, moons and magnetic field of Uranus. Finally this instrument will aid us to understand Uranus’ weather systems (cloud dynamics and properties) by making use of cloud tracking (Dougherty, et al., 2011; Arridge, et al., 2011; Arridge, 2013).
The instrument will perform the following tasks: •”Detailed characterisation of the morphology” of the surface of Uranus’ moons (Dougherty, et al., 2011). •Investigation of the magnitude, nature and distribution of tidal dissipation of the moons. •Geologic, astrometic, morphologic and geodetic observation of the moons (Miranda, Ariel, Titania). •Monitoring of cryovolcanic activity and other surface changes on Miranda, Ariel, Titania and Umbriel. •Determination of the moon’s rotation state. •Imaging of Uranus with middle resolution, with the purpose of studying the cloud dynamics and properties. •Study Uranus aurora. •Detailed high imaging resolution of the surface of Uranus satellites and rings. •Monitoring of Uranus’ rings (Dougherty, et al., 2011; Arridge, et al., 2011; Arridge, 2013). The selected narrow angle camera is similar to the New Horizon’s camera Lorri, which is currently used for the New Horizon’s mission as well as the Janus (‘Jovis, Amorum ac Natorum Undique Scrutator’) camera on Juice (Palumbo & Jaumann, 2014).
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4.6 WIDE ANGLE CAMERA Scientific goals and measurements: The Wide Angle Camera will provide images of Uranus and its moons and addresses the goals of meteorology, geology, geophysics and geodesy. The Wide Angle Camera’s main tasks are listed below. •Imaging of Uranus for the purpose of studying cloud dynamics and properties. •Study Uranus’ aurora. •Monitor atmospheric Zonal winds. •Monitor weather events on Uranus. •Mapping of Uranus’ rings and moons and supports the analysis of their migration changes since Voyager 2. •Supports a better understanding of particle sizes and densities of the rings. •Monitor the interaction between the rings and the natural satellites. •Monitor the rings’ dynamical behaviour. •Determine the moon’s rotation state. •Monitor the interaction between the Uranian System with Uranus’ Magnetic Field. •Search for present and past activity (active dynamic processes) on Uranus natural satellites, how did their surface feature form (Greeley & Dougerty, 2010)?
Performance requirements: The Wide Angle Camera has to capture images at extremely “low solar illumination levels at more than” 19.2Au away from the sun (Dougherty, et al., 2011; Arridge, et al., 2011; Greeley & Dougerty, 2010).The selected Wide Angle Camera is similar to the WAC camera MDIS (The Mercury Dual Imaging System) on Messenger (Dougherty, et al., 2011; Arridge, et al., 2011; Greeley & Dougerty, 2010).
4.7 MAGNETOMETER The instrument shall provide measurements of magnetic field strength and fluctuations in the field of Uranus to better define the properties of asymmetric planetary magnetic fields. The science goals for the instrument are as follows: Characterise Uranus’ asymmetric magnetic field with implications for verification of internal structure models Provide details of the temporal variation in the field as a result of the unusual planetary rotation axis The magnetometer instrument will also provide context for plasma experiment results, particularly with regard to identifying regions of plasma trapped by the magnetic field. Performance requirements and concept. The Voyager 2 instrumentation provided field strength measurements in the 8- 200,000nT range between two systems. The highest field strength detected at Uranus (413nT), was measured using the low field magnetometer in the 5th dynamic range mode (±710nT), which had a quantisation uncertainty of ±173pT with an RMS noise level of 6pT (Behannon et al., 1977). An appropriate comparison would be to those instruments used on the THEMIS mission, as summarised by (Angelopoulos 2008). For that mission a large emphasis was placed on detection of fluctuations within the field for the purpose of analysing space weather, but this is not a driving requirement in this case, with high resolution in a smaller range being more appropriate. The requirements for the magnetometer instrument for this mission are a range of up to 20µT, 100µT is the maximum field strength predicted at Uranus’ surface (Herbert, 2009), which corresponds to a probable maximum of ~12.5µT at 2 RU as magnetic field strength drops off as a cube of distance, so having a maximum range of 25µT gives allows for measurements to be confidently made at orbital radius as low as ~1.58RU. The smallest measurement the magnetometer will be called upon to make is of the minimum predicted magnetic field strength at ~18RU (at the magnetopause). This can be calculated simply using the cubic relation with the minimum surface value of 7700nT (Herbert, 2009) to obtain a value of ~1.32nT. The required resolution of the instrument is therefore 4pT or better; allowing this measurement to be made with an error of no more than 0.3% while incorporating a comfortable margin. This resolution requirement is consistent with similar systems such as that proposed for the JUICE mission (Dougherty et al., 2001) and the THEMIS fluxgate magnetometer.
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A pair of magnetometers should be flown, each sharing a boom, with one placed at the end and another halfway along; this allows for accurate assessment of field readings resulting from spacecraft electronics and subsequently reducing the effects of these through processing. The sensors would most appropriately be tri-axial fluxgate sensors, with the accompanying electronics within the spacecraft. The length of the boom would be driven by the magnetic cleanliness requirements, appropriate values for which are a spacecraft DC field of <2nT and AC field of 0.1nT rms up to 64Hz (field strengths as a result of spacecraft electronics as measured by the outboard sensors). Other requirements. The primary science goal regarding determination of variations in the field as a function of time makes long-term measurements desirable, with constant operation throughout the lifetime of the spacecraft being the ideal scenario. This is fortunately entirely viable, with the probably power and telemetry cost of magnetometer operations being very low. The most impactful requirement is that the arrival time at Uranus be when the spin axis of the planet is at least 45° out of line with the solar wind, preferably while moving towards an orthogonal state. That would ideally mean Uranus insertion dates in the period 2036-2046 (where 2046 sees the orthogonal state). This is required to provide information on the temporal variation of the magnetic field on a seasonal timescale, to provide a significant difference between the output data from this mission and from the historical data from Voyager 2. There are no pointing requirements of magnetometer operation, but knowledge of spacecraft attitude and positioning of the instrument of the spacecraft should be known to within 1 arcminute so that sensor orientation would be known to within 2 arcminutes during flight; this is important as the output data is vectorial. The requirement of knowing sensor orientation makes boom rigidity very important, with as much precision as possible in the joints allowing it to unfold.
4.8 PLASMA PARTICLE DETECTOR The instrument shall provide measurements of plasma particle density, energy and velocity in order to meet the following main science goals: Characterisation of Uranus’ magnetosphere with particular reference to the interaction with the solar wind Determination of plasma dynamics throughout Uranus’ magnetosphere and variations over time Characterisation of the interaction between trapped plasmas and the satellites of Uranus Performance requirements and concept. The plasma particle detection instrumentation is required to detect both low energy (a few eV to ~20keV) plasma as well as particles with energy in the tens of keV. The plasma to be detected is mostly composed of ionised Hydrogen so the focus is on proton and electron detection, with a lesser requirement being the detection of less energetic neutrals. To detect the ions two instruments a required: a thermal ion detector (TID) and higher energy ion detector (HEID). This setup is mirrored by the THEMIS plasma detection packages: ESA (Electrostatic Analyser, 2.96kg, 1.7W) described by (McFadden, Carlson et al. 2008) and SST (Solid State Telescope, 1.43kg, 1.2W), which detect low and high energy ions and electrons respectively. Detection of neutral plasmas requires a pair of additional detectors mirroring the Energetic Neutral Analyser (ENA) and ion & neutral gas mass spectrometer (INM) from the JUICE mission. The energetic neutral analyser characterises neutrals in the range of 10eV to 10keV, which allows for imaging of the plasma environment close to Uranus’ moons, specifically allowing for the determination of particle surface release processes and patterns of ion precipitation. The neutral gas mass spectrometer will allow for characterisation of the elements which make up the neutral plasma around Uranus and requires a mass resolution of M/M > 1000. Characterisation of cold plasma (fractions of eV) is required to provide the spacecraft potential in order to interpret low energy electron and ion measurements, and for sensor calibration (Dougherty et al., 2001), which can be achieved using a Langmuir probe (LP) (effectively just a simple electrode outside the craft with electric potential between it and the main hull). Such a device would also provide useful data on plasma density. To properly characterise the plasma environment around Uranus the field of view of the instrument must be as large as possible. To achieve a 360° field of view in the nadir/beam plane a single instrument with a much more restricted FOV could be used on a spin-stabilised spacecraft, using the rotation of the craft to track across the desired field. Unfortunately the pointing requirements of other instruments necessitate a different solution. To minimise complexity two instruments, with port and starboard pointing, could be used, each with a FOV of at least 181.15°, assuming that the separation of the sensors is 2m or less; this ensures a dead zone along the beam axis of 200m or less centred on the
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spacecraft. The primary requirement is to detect plasma flux so angular resolution of the detector is not critical; up to 20° is acceptable. Other requirements. The plasma particle detector must run continuously alongside the magnetometer instrument to fully characterise the magnetic field and interaction with the solar wind. The field of view for the instrument must include the entire nadir/beam plane, and a conductive surface for the spacecraft is required due to sensitivity of the low energy measurements to the potential of the spacecraft. In order to characterise the interaction of trapped plasma regions with the Uranian satellites an orbital tour route is required which allows for low passes through the wake of the moons (periapsis altitude <1 body radius).
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5 MISSION DESIGN
5.1 REFERENCE MISSION The goal of the mission orbit analysis section of this project was to provide a reference trajectory and science orbit which could feasibly be achieved by current or in development systems to deliver a science payload to Uranus and remain in orbit for around five years. An additional driver was the transit duration from the use of ASRG power sources; these have a lifetime of approximately 17 years, so to provide a five year mission the transit period must be less than 12 years, preferably 11 or less to allow some margin.
Initial efforts towards determining viable mission trajectories were made into creating bespoke software for finding potential gravity assist sequences and ΔV requirements. Due to time constraints however it proved that this approach was not practical so a NASA database of mission trajectories was searched to find an appropriate route. This database only accounted for single gravity assists using Jupiter, which severely limits the launch opportunities which are shown below in figure 7, with the colours indicating the ΔV requirement of that trajectory.
Figure 7 - Uranus Transfer Database Graph The optimum route from this database is shown in figure 8, which has a launch date in 2034, which fortuitously leads to an arrival date in 2044; a total transit duration of ~10.1 years, very close to the ideal arrival time in terms of magnetic field science and keeping within the bounds imposed by the power source. While more efficient trajectories, likely using a greater number of gravity assists including multiple Venus/Earth passes, probably exist the time involved in finding them is prohibitive. Additionally, while more efficient routes exist they are all almost certainly substantially slower than this trajectory, making them unsuitable for this mission. For comparison, a direct Hohmann-type transfer to Uranus from Earth would take approximately 16 years, far longer than is practical for this mission even if the ΔV requirements of doing so were feasible.
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Figure 8 - Database Reference Trajectory
5.2 ANALYSIS The data provided by the database derives from a trajectory finder which uses pure two-body patched conic approximations, without accounting for solar radiation pressure or other effects. It also assumes departure from a 200km parking orbit and orbit stabilisation at Uranus with a periapsis of 200km and a C3 of zero – so an apoapsis effectively on Uranus’ sphere of influence boundary. A basic version of this database trajectory was built in GMAT (NASA’s open source General Mission Analysis Tool) which is shown in figure 9 in order to show the basic feasibility of the route, but not for performing detailed analysis.
Figure 9 - GMAT Approximation of Transfer
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The approach taken for this was to use a forward-shooting method; setting initial conditions then allowing a differential corrector to modify the magnitude of the initial burn (in V, N and B directions) to leave a low Earth parking orbit (200km altitude, in the plane of the ecliptic). This manoeuvre was modelled as an impulsive burn as the error inherent in this when compared to a finite burn is minimal in this case, given that it would be performed by the launch vehicle upper stage and so would have a high thrust to weight ratio. The integrators used were PrinceDormand78 (for interplanetary space) and RangeKutta89 for the most part elsewhere. Note that the arrival date is more than a year ahead of the database value; this is due to some elements of the trajectory having been manually iterated to reach an approximation of the trajectory. A closer match to the database would require a more sophisticated and time-consuming analysis, likely using a multiple-shooting method, which is discussed further in the appendix. The most significant divergence from the database trajectory lies at the end, with the Uranus orbit stabilisation manoeuvre. The low, 200km, altitude assumed for the stabilisation is inappropriate in reality due to the presence of the Uranian ring system. The rings are known to extend out to around 3.5 Uranus radii with infrequent but dense bands of material and it is believed that large quantities of dust particles extend from there down almost to the cloud-tops (Esposito, 2002). As such the minimum periapsis for the mission needs to be above the main ring zone, with a radius of 4 RU selected to provide margin. To determine the orbiter propulsion requirements in terms of the orbital stabilisation manoeuvre at Uranus, which is the driver for that sub-system and as such much of the overall spacecraft architecture, calculations were made working backwards from the optimal NASA database value. Given that the ΔV for the stabilisation burn from the database value is for a specific given orbit one can simply determine the orbital entry velocity at periapsis as the orbital speed for the specified orbit plus the ΔV for the stabilisation manoeuvre (1.07km/s). This yields a value of ~21.3 km/s, which 2 2 translates to an arrival C3 of approximately 46.8km /s . From this energy value the periapsis velocity of any altitude arrival can be determined, allowing one to find the magnitude of stabilisation burn by simply determining the required change in velocity between this and the periapsis velocity of the desired final orbit. The science goals of this mission make a much lower orbital apoapsis than assumed by the NASA database desirable in order to provide more frequent opportunities to study the moons and to provide magnetic field data close to the planet on a regular basis. It is also desirable that the argument of periapsis place the periapsis in eclipse, so that the shortest period possible is spent outside of line-of-sight with the Earth (and by extension the Sun given that from Uranus the angular distance between the two is fairly small). While it would be possible to use an orbit with the semi-major axis in alignment with the local horizontal of Uranus’ orbit and the spacecraft in full view at all times this would require extensive trajectory correction manoeuvres leading to a higher total mission ΔV requirement for very little real gain. A balance also must be maintained between low-altitude science phases and higher-altitude periods during which communications downlink to Earth would be scheduled; due to the pointing requirements of the communications system it would be largely impossible to achieve this during data acquisition except with instruments such as the particle detectors and magnetometer which do not need specific facings.
Following analysis of a number of different apoapsis altitudes a radius of 28 RU was selected to provide a good balance while also not imposing too great a requirement upon the propulsion system for the orbit stabilisation manoeuvre. To inject onto this 4 by 28 RU orbit this stabilisation burn would be approximately 2.7 km/s, which while high is not outside the realms of possibility. A visualisation of this orbit is shown in figure 10, with the periapsis behind the planet with regards to Earth for the reasons discussed above. The image shows the mission orbit in red with the orbits of Uranus’ moons shown in beige (with the rings not shown, though they share the plane of the moons, which is in line with Uranus’ equator), with the plane of the ecliptic represented by the blue grid and the yellow line showing the direction of the Sun. It is possible that more information regarding the properties of the ring system could come to light between the time of this study and the proposed approximate launch date, but this is by no means a certainty as such data would be extremely difficult to gather without sending a vehicle to explore the Uranian system. If data showing that it is possible to perform the stabilisation manoeuvre at a lower altitude the ΔV saving would be significant: for a scenario in which the primary stabilisation burn is performed at a radius of 1.2 RU and then a burn performed at apoapsis to raise the periapsis of the final orbit back up to 4 RU there would be an overall saving of 512 m/s.
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Figure 10 - Science Orbit If it transpired that the science orbit itself could use a lower orbital periapsis the saving would be even greater, while reducing the time spent per orbit in eclipse, thereby increasing the proportion of the orbital period during which Earth communication is possible. Injecting onto a 1.2 by 28 RU orbit would incur a stabilisation burn of ~1.57 km/s, which is a saving of ~1.1 km/s, which could allow for an even more comprehensive science payload, a larger provision for a more complete tour of the moons or even the use of a lighter launch vehicle. As such any data allowing one to accurately the risky of operating inside the radius of Uranus’ main rings is highly desirable, though it is assumed for the purposes of this report that doing so is too unsafe and so the 4 by 28 RU orbit is used as reference.
The mission also requires access to Uranus’ moons, all of which come within ~4-5 RU of the orbiter at some point during the mission assuming that the initial science orbit is maintained, with closer views of the innermost principle moons, Miranda and Ariel, occurring regularly. This by itself would provide a great deal of opportunity for gaining valuable science, but further provision should be made for on-orbit manoeuvres to grant closer passes where possible (which would largely be achieved by modifying the argument of periapsis). A detailed study of such opportunities is outside the scope of this project but it is expected that several close passes of at least a few of the major moons could be achieved with a relatively modest ΔV budget. It is also hoped that moon flybys could provide gravity assists reducing the need for major manoeuvres using the main engine, though a detailed analysis of the potential for this is again something outside the scope of this document. Some work was done however in determining whether this would be at all possible, most significantly using repeated GMAT simulations of the science orbit using multiple gravitational bodies; Uranus and the principle moons. The previously shown science orbit in figure 10 was generated neglecting the mass of the moons, in order to provide a clean visualisation of the orbital form. Figure 11 shows the same orbit over a period of nearly 43 days with all of the moons modelled as point masses.
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Figure 11 - Perturbations due to Uranus' Moons Note that the degree of orbital perturbation is significant even over a relatively short timeframe. As such it is reasonable to assume that with some degree of trajectory correction these perturbations could be used to adjust the orbit to provide closer passes of at least some of the moons. Modelling the moons, and indeed Uranus itself, as point masses is not ideal, but at this time there is no reliable information on the internal composition to the degree of providing an accurate gravitational field map, so is the best approximation which can be made for now. The perturbations due to the moons result even without manoeuvring, while never entering the gravitational spheres of influence of any of the moons. A thorough analysis of the ΔV requirements to perform a moon tour was outside the scope of this project due to the time it would take to perform such analysis comprehensively. In order to reach the outer two main moons, Titania and Oberon, small changes to apoapsis altitude and minimal adjustments to the argument of periapsis could be made. It would not be economical to perform major changes regularly, with plane changes particularly problematic; while adjusting the apoapsis from 28 RU to 23 RU to match Oberon would take only ~132m/s of ΔV it would take ~4.3km/s to align the semi-major axis with the equator (and thereby the moon orbits). This is clearly not feasible, so the long-term moon tour strategy must be one of minor manoeuvres made sparingly. At the end of the mission it would be desirable to reduce the periapsis altitude to provide data close to Uranus itself; particularly valuable for magnetometer readings to provide clues about the internal structure of the planet. To drop the periapsis down to the cloud-tops from the initial science orbit would require a burn of ~678 m/s. While a less aggressive burn would still bring the spacecraft low enough to ensure eventual orbital decay several hundred m/s of ΔV would still be required. This somewhat high requirement makes the use of moon gravity to perform this final manoeuvre the most feasible method.
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6 LAUNCH VEHICLE
6.1 OVERVIEW & PRELIMINARY SELECTION Early in the project lifecycle a preliminary study of launch vehicles was undertaken based around the Earth escape orbit properties given in the Ariane 5 user manual, which was at the time the proposed launcher. This study is discussed further in the appendix, but the conclusion of it was that the Ariane 5 is not capable of launching this mission, at least in a standard configuration. From there it was decided that future efforts would concentrate on heavier launch vehicles of which only two are going to be flight-ready in the near future: NASA SLS and the SpaceX Falcon Heavy (figure 12).
The NASA Space Launch System was shown to be vastly over specified for the task even in the lightest configuration, as well as being somewhat prohibitively expensive with estimated launch costs running in excess of one billion US dollars. As such, the focus was put upon the SpaceX Falcon Heavy, for which the estimated launch cost is of order $85 million, though it is likely that the cost would be higher for a fully disposable launch which would certainly be required for this mission. Both the Falcon Heavy and the SLS will launch from the Kennedy Space Centre, with the Falcon using pad 39A; recently leased to SpaceX for 20 years, which was the launch site for many of the Apollo missions including Apollo 11, as well as the first and last Space Shuttle launches.
The Falcon Heavy design is based upon the currently in operation Falcon 9, which at the time of writing has had 12 successful launches with all primary payloads reaching the intended destination orbit (though in one case a secondary payload was lost due to booster engine failure). One unique feature of the Falcon 9 (and heavy) is the reusability of the core stage; instead of a disposable system the core stage is designed to safely descend and perform a powered vertical landing onto a set of foldaway landing legs. This functionality has not yet been fully demonstrated, though recent flights have
successfully performed a powered descent to the ocean.
The Falcon 9 has two main stages: the core stage and an upper stage booster, both of which use liquid oxygen and RP-1 (rocket-grade kerosene) as propellant. The core stage uses 9 Merlin-1D engines, theoretically allowing for successful launch even in a multiple engine failure scenario, while the upper stage is a more conventional single- engine (the Merlin Vacuum Engine) approach. The Falcon Heavy shares the same core and upper stages as the Falcon 9, with the notable addition of two large boosters, which are themselves effectively the same as the core stage, with similar fuel quantities and the same 9-engine setup (which are arranged with a central engine and 8 more surrounding it in a circle, though the original version of the Falcon 9 launched with the engines configured in a 3 by 3 grid).
Because of the commonality of systems with the already-flown Falcon 9 it is not unreasonable to presume that the Falcon Heavy will also share in this success. It is worth noting that the two booster stages on the Falcon Heavy feed fuel into the central core stage while attached, allowing all 27 engines to fire at launch while leaving nearly Figure 12 - Falcon Heavy Render, a full tank for the main core at the time of booster separation, maximising efficiency. Credit: SpaceX Website Clearly, retaining fuel in each stage for a powered descent has a drastic impact on the
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potential payload, but SpaceX have indicated that for any particularly demanding missions disposable cores would be considered an option.
To surpass the work performed in the preliminary study a graph of characteristic energy versus deliverable mass was needed to assess the capability of the vehicle to launch interplanetary missions. The characteristic energy of an orbit is equal to the square of the hyperbolic excess velocity, which is the velocity of an object in a
hyperbolic orbit at infinity. As such, the characteristic energy (C3) is an excellent
measure of the energy requirement of a given escape orbit, and a C3 versus deliverable mass graph is a standard way of demonstrating the capability of a launch vehicle. Currently there has been no such curve published for the Falcon Heavy, which is understandable given that it is yet to fly.
6.2 CHARACTERISTIC ENERGY ANALYSIS
Due to the lack of official, published C3 data on the Falcon Heavy analysis of it, and a number of other vehicles, was undertaken to produce an approximation largely from first principles. The starting point for this analysis was with the Soyuz-Fregat (figure 13); while this vehicle is of a much lighter class and is not as modern as the
Falcon series it provided a C3 curve as part of the user manual, allowing for a solid check of the analytic method used before applying the same method to other vehicles with fewer known properties.
The Soyuz-Fregat has a 100% success rate for the 46 launches in its lifetime, and has been used extensively to launch manned vehicles to the International Space Station. In the cargo configuration which is of interest the Soyuz-FG can be thought of as a four-stage vehicle, with the distinctive Soyuz four liquid boosters followed by a large core stage and then a smaller upper booster all of which use liquid oxygen and RP-1 as fuel. These stages are then followed by another small upper stage, the Fregat, which uses Nitrogen tetroxide and UDMH (unsymmetrical dimethylhydrazine).
To create the C3 curve for the Soyuz the total ΔV of the vehicle needed to be calculated, using Tsiolkovsky’s rocket equation sequentially for each stage. The details of this process are discussed more extensively in the appendix, but it is worth noting that the ΔV of the vehicle is highly dependent upon the payload mass, with this effect being particularly pronounced for the upper stages where the payload makes up a larger proportion of the dry mass of the stage. The Soyuz-FG user manual is very comprehensive and provides complete data regarding the wet and dry mass of each stage, and the masses of interstage and fairing elements, which made analysis of the vehicle ΔV relatively easy.
Once the total ΔV of the vehicle is known the requirement for achieving Earth orbit can be subtracted, which is the orbital velocity required as well as aerodynamic and gravity losses inherent in the ascent. The orbital velocity can simply be calculated using standard equations (though this is more explicitly detailed in the appendix), while to estimate the other losses a parametric equation originally presented in Figure 13 - Soyuz Fregat, Credit: ESA Website
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(Townsend, 1962). This parametric was based upon launch vehicle configurations of the early 1960s, which were mostly based on ballistic missile technology that generally uses much more aggressive ascent profiles than modern launchers. The equation depends primarily upon the total ascent time of the launch, which for the older launchers the parametric derived from is much shorter than that generally exhibited by more modern launchers, which generally favour a steep ascent through the high loss regime of the lower atmosphere followed by a gradual climb on a lower thrust but highly efficient upper stage.
Due to the difference in launch vehicle architectures since the creation of the parametric it was expected that an additional loss factor would need to be applied, in order to calibrate the results of the analysis with the user manual data. This did indeed turn out to be the case, and the additional loss factor used was 790m/s, on top of the 2.78km/s loss factor calculated using the parametric. This additional factor represents approximately 5.6% of the total ΔV of the vehicle using a nominal 1 tonne payload, where the total ΔV is ~14.1 km/s, leaving ~3.97km/s after reaching a 200km stable orbit.
Given the calculated on-orbit ΔV determining the possible C3 is a relatively simple matter of determining the hyperbolic excess following a burn using all of said ΔV in low Earth orbit. The result of this analysis, repeated for a range of payload masses, is shown in figure 14 as the traditional C3 versus deliverable mass curve. Both the model data derived from this analysis methodology and the original data taken from the user manual are shown for comparison purposes. 2 2 Note that the small divergence for C3 over 24 km /s is due to a change transition between sub-orbital and direct injection ascent profiles which are not accounted for by the analysis method.
In most respects though it is clear that the analysis delivers a remarkably close match to the published data for the system. An alternative analytic method was briefly considered but disregarded as it produced results which produce a much less convincing match, which is discussed in the appendix.
Model Data Original Data 1700 Soyuz Fregat C3 vs Delivered Mass 1600 1500 1400 1300 1200 1100 1000 900 800
Delivered Mass (kg) 700 600 500 400 300 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40 42 44 46 2 2 C3 (km /s )
Figure 14 - Soyuz Fregat Characteristic Energy Following the fairly successful outcome of the analysis performed for the Soyuz-Fregat analysis the same was also performed for a number of other launch vehicles as well as the SpaceX Falcon Heavy. The Ariane 5 was examined despite the earlier study showing it to be unlikely to give sufficient performance in the stock configuration. As such it
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was decided to examine a few hypothetical configurations, most importantly a version using an additional stage with properties the same as for the aforementioned Fregat upper.
Study was also made of the potential performance of a version of the upcoming Ariane 5 ME (Mid-life Evolution) which features a new upper stage engine, with an assumption of an increase in the propellant mass for that stage. While this configuration would provide a significant increase in deliverable mass it is not presented here as the results are contingent upon a large number of assumptions which would need verification.
The Titan 3E was also examined as it is notable in that both Voyager missions were launched on such vehicles, which used trajectories involving a direct transfer to Jupiter as this mission does, albeit with a more aggressive initial burn 2 2 requiring a C3 of 105.5 km /s for Voyager 1, which had the larger requirement of the two (REF).
Finally the Falcon Heavy was given the same treatment, calibrated using the deliverable payload to Mars of 13,200 kg, quoted from the SpaceX website. It was assumed that the transfer to Mars for this payload mass would be an ideal 2 2 Hohmann transfer requiring a C3 of 8.68 km /s for the escape orbit; this value was calculated using a standard patched conic approximation. The details of the analyses of these vehicles are discussed further in the appendix. The characteristic energy curves for all of the above mentioned vehicles are compiled into a single figure (figure 15).
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6.3 CHARACTERISTIC ENERGY ANALYSIS RESULTS Falcon Heavy Soyuz FG Titan 3E Comparison C vs Delivered Mass Ariane 5 Ariane 5 Fregat Mission C3 3 Falcon Heavy Quoted Performance Voyager 1 Voyager 2 Ariane 5 User Manual Data 16000
14000
12000
10000
8000
6000 Delivered Payload(kg) 4000
2000
0 86.3 0 20 40 60 80 100 120 140 2 2 C3 (km /s )
Figure 15 - Characteristic Energy Graphs
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6.4 SUMMARY The comparison graph shown in figure 15 clearly demonstrates that the Falcon Heavy is a much more capable launcher for a mission of this type, delivering a payload to Uranus using the chosen trajectory of approximately 3.28 tonnes based upon this analysis. While the stock Ariane 5 (red) is demonstrably not capable of directly launching onto the required escape orbit in a default configuration the hypothetical version which includes a Fregat upper would be capable of delivering a payload ~2.43 tonnes onto the desired orbit.
A preliminary estimate of the required wet mass of the orbiter was made of ~3.13 tonnes. This assumes the use of an engine with a specific impulse of 317s with the ΔV requirement derived from the orbital stabilisation manoeuvre required for delivery onto the desired science orbit with additional provision of 300 m/s for TCMs and manoeuvring in- situ.
Using this estimate even the Ariane 5 with a modest additional stage would be insufficient, as would the theoretical version of the Ariane 5 ME (not shown). The possibility remains, however, that a launcher based on the Ariane 5 ME with an additional upper stage would be able to provide sufficient performance for this mission. At this time this possibility has not been explicitly examined however, as the analytical method here used relies upon a real-world calibration point, and such a configuration is too far removed from reality for any such calibration to reliably take place. While the potential of such a launcher is discussed further in the appendix it would require a much more focused study to accurately determine whether such an approach is feasible. As such, the SpaceX Falcon Heavy is the assumed launch vehicle for the mission, with the possibility of using an additional small upper even on this heavier launcher if necessary.
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7 PROPULSION
7.1 REQUIREMENTS The first stage of the analysis of the orbiter propulsion system was refinement of the requirements generated for the system during the mission orbit analysis phase of the project. While the primary circularisation burn requirement remains the same it is worth noting that the Jupiter TCM and moon tour provision were separated, though the ΔV requirement for those also remains unchanged. It was decided however to place a 5% margin on the ΔV requirement to account for unforeseen TCMs and hopefully to allow for additional manoeuvring to arrange moon flybys. This led to a total ΔV requirement of a little more than 3 km/s, as broken down in table 7.
Orbiter ΔV Requirement Table
ΔV Margin 5.00%
Arrival Periapsis Velocity 12687 m/s
Desired Orbit Periapsis Velocity 9998 m/s
Required Primary Burn ΔV 2690 m/s
Moon Tour Provision 250 m/s
Jupiter TCM 44 m/s
Total ΔV Required 3132 m/s
Table 7 - Orbiter ΔV Requirement Table The ΔV requirement is very high and most of it is accounted for by a single, time-sensitive, burn to stabilise the Uranus orbit. As such, a high thrust to weight ratio is a requirement in order to complete the burn in a matter of no more than a few hours. This requirement alone effectively precludes the use of an electric propulsion system, which would require a prohibitive amount of power to produce the necessary thrust. This is especially true given the distance from the Sun; with an orbit of 16 AU the solar flux at Uranus is almost negligible (0.3% that at Earth), requiring any solar array to power even a modest electric propulsion system to cover hundreds of square metres.
While a nuclear power system would be able to power an electric propulsion system the thrust to weight ratio requirement would rule out all but very large nuclear reactors, which would drive costs up dramatically as mission complexity, orbiter mass and hardware development requirements mount. As such, it was deemed that a traditional chemical approach would be more suitable, which aligns with the assumption made at the beginning of the project.
7.2 GENERAL ARCHITECTURE OPTIONS Following the confirmation that a chemical approach would be most appropriate analysis was undertaken of a number of different thruster options. One area of commonality between all of the thrusters is the propellant; monomethylhydrazine and dinitrogen tetroxide (or more commonly MON-3, which is 97% dinitrogen tetroxide with the remaining 3% being Nitric acid, which generally improves propellant performance). This is a commonly used propellant, having been a staple of spaceflight since the beginning, seeing use on everything from Apollo to Soyuz and the Space Shuttle (or a very similar mix, in some case unsymmetrical dimethylhydrazine is used, though this is generally avoided where possible due to the toxicity of the material).
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The other most commonly used propellants are a liquid oxygen/liquid hydrogen mix, and liquid oxygen/kerosene. Both of those choices of course use one or more cryogenic components which makes long-term storage prohibitively difficult, hence those propellants are used almost exclusively for launch vehicles where propellant boil-off effects are minimal. Additionally, those propellants require ignition systems, whether those take the form of an electrical starter or even a smaller rocket motor in the case of very large systems; this generally means that the number of engine firings is severely limited if more than one is even possible. Using a hypergolic fuel such as MMH/N204 removes this requirement as mixing the two propellant leads to self-ignition. Clearly this near-unlimited restart capability is a tremendously valuable capability, especially for long-term station keeping and other tasks in need of repeated firings.
There are two major groups within the considered thrusters: pressure-fed and pump-fed. Pressure-fed systems are generally used for smaller systems as the reduced mechanical complexity saves mass in many cases. The propellant tanks for a pressure-fed system must be kept at a higher pressure however, specifically a value somewhat higher than the chamber pressure of the engine so for larger systems a pump-fed system is generally lighter. The general process of sizing the various components of a propulsion system beyond the thruster itself is detailed in other sources, with both (Wertz & Larson, 1992) and (Sutton & Biblarz, 2001) used as reference for the analysis performed for this project. By far the most massive of the other elements of a propulsion system are the propellant tanks, especially for a pressure-fed system.
For pressure-fed systems a pressurant gas will be required in addition to the propellants in order to maintain the internal pressure in the fuel tanks. This pressurant must be inert, not condense and be insoluble in the propellant (Sutton & Biblarz, 2001), which makes noble gases a natural choice, though Nitrogen is also used in many cases, though it will dissolve in nitrogen tetroxide or liquid oxygen so Helium is more typically used. There are broadly two different approaches to using a pressurant gas: regulated and blowdown. A regulated system uses a separate high-pressure tank for the pressurant, feeding into the propellant tank through valves, while a blowdown system keeps all of the pressurant in the main tank. These two options are shown in figure 16. The gas and propellant in both cases is kept separate principally by surface tension forces, though also by the force exerted by the thruster during a burn.
Figure 16 - Pressure-fed Rocket Systems, (Wertz & Larson, 1992) The advantage of a regulated system is that the main propellant tank does not need to be kept at a dramatically higher pressure than the chamber pressure of the engine, though the pressurant must be kept at a very high pressure externally. In a blowdown system however the initial pressure of the tank must be set such that the final pressure once most of the propellant has been expelled is slightly above the thruster chamber pressure. This makes a regulated system dramatically better for relatively large pressure-fed propulsion systems, as the mass implicit in using large high- pressure tanks for a blowdown system is significant, as will be demonstrated later.
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While in a pressure-fed system the pressurant gas largely ensures that the propellant is kept in place to be accessed by the fuel lines an unregulated tank, as is used in a pump-fed system, does not do that, so an alternative solution for this is required to prevent unwanted gases or vapours being injected into the engine in place of liquid. There are several strategies commonly employed to circumvent this issue, either by applying a force to simulate gravity to settle out the liquid or a positive expulsion system, which use mechanical systems to force liquid into the feed lines, exploiting surface tension forces (Martin, 1986).
A pump-fed system does not require a separate pressurant, with surplus volume in the propellant tanks largely occupied by fuel vapours. The pressure in the tanks is largely irrelevant, though is typically between 10 and 50 psi (Sutton & Biblarz, 2001); for the purposes of this report the value of 50 psi has been used to remain conservative. This is still a much lower pressure than the typical chamber pressures of thrusters used in pressure-fed systems (~140-200 psi), hence the generally lower mass for systems requiring large quantities of propellant, despite the increase in thruster assembly mass, which can be up to around ten times the mass of the pressure-fed counterparts.
7.3 ANALYSIS A fairly large number of thrusters were analysed to a greater or lesser extent as part of this project, but in the interests of expediency here will be presented only two representative of pump and pressure-fed options of the appropriate class. Further discussions of the other thrusters which were examined are included in the appropriate appendix. It should be noted that no good examples of pump-fed thrusters could be found of the necessary size which are currently in production, but those examined were developed to a high level so are a good indication of feasibility.
The pressure-fed option here discussed is the Aerojet R-42DM, which as of 2011 was at TRL6 awaiting formal qualification and the final flight design, though by all accounts the final performance of the engine should not deviate significantly for the values taken from the Aerojet data sheets. As has already been mentioned all of the thrusters evaluated use hydrazine and nitrogen tetroxide as propellant, in this case using the MON-3 mix of N204 and Nitric acid. Table 8 lists the most important properties of this particular thruster. This description relates to the system using a regulated pressure system, with a blowdown version briefly considered further on.
R-42DM Properties
Thrust 890 N
Chamber Pressure 140 psi
Expansion Ratio 200:1
Specific Impulse 327 s
Steady State Firing 1000 s
Mass 7.3 kg
Table 8 - R-42DM Thruster Properties In order to calculate the masses of the required propellant and pressurant tanks the primary consideration is a comparison of the pressure the tank is required to maintain with the tensile strength of the material being used. For these purposes the assumption was made that the tanks would be constructed of aluminium, which offers a good mix of relatively low density and good structural properties. For the propellant tanks the pressure requirement was taken to be 20% higher than the chamber pressure of the engine, so in this case 168 psi (~1.15 MPa). It should be noted that while pounds per square inch values are liberally quoted here for all calculations SI units (Pa) were used rather than imperial; the use of psi is purely due to this being convention in most of the texts examined during this project.
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The pressure of the pressurant tank uses a nominal value of 3000 psi taken from (Wertz & Larson, 1992), with Helium the gas used. In order to calculate the required mass of pressurant an equation was used from (Sutton & Biblarz, 2001) , which can be found in the appendix, for which the principle components are the physical properties of the pressurant gas, the volume of propellant and the ratio of pressurant and propellant tank pressures. For this system this leads to a pressurant mass requirement of ~6.2kg of Helium, which occupies a volume of ~0.19 m3 at the required pressure.
Figure 17 - R42DM, Credit: Aerojet The propellant mass requirement was found using the rocket equation, using the specific impulse of the specific thruster, which is 327s in this case, starting from the ΔV requirement previously discussed and assuming the dry mass of the spacecraft to be exactly 1200kg. Additionally a 5% margin on the propellant quantity was added for conservatism, with a further 5% representing propellant remaining unused in the tank, as expelling 100% of the propellant is in practical terms impossible. This leads to a requirement of ~2.2 tonnes of propellant split between hydrazine and MON-3 in a nominal 1:1 mixing ratio. This requires internal tank volumes of ~0.6 m3 for the MMH tank and ~1.48 m3 for the MON-3, which includes an additional 10% to account for ullage gas volume.
As the propellant tanks occupy a significant portion of the overall spacecraft volume cylindrical tanks were assumed in order to utilise the space most effectively, while a spherical form was assumed for the pressurant tank as that reduces stress which is particularly important for such a high-pressure vessel. The tensile yield strength of the tank material, aluminium, was taken from (Wertz & Larson, 1992) as 460 MPa, though an additional margin of 20% was used for safety, so a value of ~307 MPa was used. This was then compared with the hoop stress on the sides of the cylindrical tanks which is the dominant factor, in order to find the required tank thickness for a given tank radius. Specifically, the radius used was 0.47m, chosen in order to effectively use the volume within the spacecraft while keeping tank thickness requirements reasonable. The same comparison was then performed for the pressurant tank, though the radius for that was simply defined by the required volume, with the caveat that of course the equation for spherical stress which is effectively half the equivalent hoop stress. The equations used for this can be found in the appendix.
The required tank thicknesses calculated as indicated above are ~1.8 mm for both propellant tanks and ~1.2 cm for the pressurant tank. These values relate to total tank masses of ~43 kg for the MMH tank, ~92 kg for the MON-3 tank and ~72 kg for the pressurant tank, all of which include a 30% margin to account for mounting hardware, valves, pipes and similar. Including a 20% sub-system margin this leads to a total dry mass of the propulsion system of ~263 kg for this thruster using a regulated pressure feed system.
For comparison purposes a few amendments were made to crudely approximate a blowdown system; removal of the pressurant tank and an increase in the propellant tank pressures by a factor of four, which is a value derived from average values from (Wertz & Larson, 1992). Note that to accurately determine the sizing of a blowdown system the volumes of the propellant tanks would in fact be larger to accommodate the pressurant gas, which would increase the mass further. Using these values and calculating everything in the same manner as for the regulated system above results in a total dry mass of the propulsion system of ~676 kg using the same thruster, so it is clear that for a system of this size a blowdown approach is inappropriate.
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While the regulated pressure option is inarguably better than the blowdown 263 kg is still a very large proportion of the 1200 kg dry mass of the spacecraft and far above the approximate value of 140 kg estimated for the propulsion system before the commencement of this more detailed analysis. Of particular concern is the spacecraft wet mass requirement of ~3.4 tonnes, which is more than the deliverable mass estimate generated for the SpaceX Falcon Heavy in the previous section. It is also important to note that the thrust to weight ratio yielded by one thruster requires a circularisation burn time of approximately 1.8 hours, more than six times maximum steady state burn time for the thruster, so an array of thrusters would be required which would further drive up the mass.
Investigation of pump-fed systems yielded much more favourable results, however. The best option considered was the Rocketdyne XLR-132, which was developed to a high level in the 1990s but never put into production, presumably due to lack of demand for motors of this size. The most pertinent properties of the thruster are listed below in table 9.
XLR-132 Properties
Thrust 16.68 kN
Chamber Pressure 1490 psi
Expansion Ratio 400:1
Specific Impulse 340 s
Steady State Firing 4000 s
Mass 54 kg
Table 9 - XLR-132 Thruster Properties Clearly, this motor is a dramatic step above the previously examined R-42DM, with almost 19 times the thrust and a chamber pressure more than 10 times greater. The dramatic increase in mass is largely due to the addition of the turbopump to supply fuel. The treatment for this motor was near-identical to that of the R-42DM, yielding a propellant mass requirement of ~2.07 tonnes, with tank masses of ~12kg for the hydrazine and ~25.3kg for the dinitrogen tetroxide. As there is no need for an additional pressurant tank in a pump-fed system this leads to a total propulsion system dry mass of a little under 110 kg including the sub-system margin.
Figure 18 - XLR-132, Credit: Rocketdyne This estimate assumes the use of standard unregulated tanks with no provision for positive expulsion systems (such as bellows or pistons, which would increase the mass). As such, ullage motors would be required to apply a small force to compel the propellant to enter the fuel lines prior to firing of the main engine. The mass of these is not included within
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the propulsion sub-system, with the intent to include them as part of the attitude control system which already includes thrusters. This is discussed further later in the report.
Due to the immense saving in terms of sub-system mass for the pump-fed system, a factor of 2.37, it was decided that a formal trade-off between the two options would be frivolous. This is especially true given that the pump-fed systems require total wet mass of the orbiter to be under the performance estimate of the Falcon Heavy which is not the case for any of the pressure-fed systems evaluated. It should also be noted that due to the much higher thrust to weight ratio resultant of the use of the XLR-132 the stabilisation burn could be completed in just 5.6 minutes, much less than the maximum steady state operation time for the motor.
Indeed, the one significant issue with the XLR-132 is that it is not a production engine, nor could any off-the-shelf motor of comparable performance be found. As such, the conclusion of this propulsion analysis is that an engine with similar properties would be required to allow a mission of this type to proceed without major changes to the mission profile. In particular, the engine must be pump-fed to keep tank mass down, and exhibit a specific impulse upwards of 330s in order to keep the orbiter wet mass requirement within the bounds of what the SpaceX Falcon Heavy can likely deliver.
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8 ATTITUDE CONTROL SYSTEM
8.1 REQUIREMENTS AND ATTITUDE DETERMINATION The principal operational requirements of the attitude control system are twofold: pointing accuracy and knowledge of that pointing. The former is imposed by the communications sub-system, a value of ±0.05°, which is discussed in chapter 10. The pointing knowledge requirement is imposed by the magnetometer; to within two arc minutes (±0.033°). Further, many of the instruments included in the mission payload have specific pointing requirements, mostly for nadir facings.
The pointing accuracy requirement is significant, and requires a three-axis stabilised zero-momentum system (Wertz & Larson, 1992) with a vibration-isolated platform for the communications system, which may require minor articulation. To achieve three-axis stabilisation an approach using both reaction wheels and thrusters is favourable to maximise the degree of control. The thrusters can be used both for large slew manoeuvres as well as de-saturation of the reaction wheels. In Earth orbit magnetic torque rods are sometime used for reaction wheel desaturation, but with so little known about Uranus and its magnetic field one cannot be sure that this would be possible at Uranus, or at least that it may well be unreliable, given the unusual nature of the field. As such, the thruster system is crucial for long-term attitude control.
In order to meet the pointing knowledge requirements a star mapper is required, which is a system with a built-in catalogue of known star positions, which are then scanned for and found to provide very accurate pointing knowledge. In addition to this a high-grade sun sensor should be installed to use as a backup. The placement of the star mapper is not of particularly high importance assuming that a wide range of possible tracking stars are provided in the database it uses. The sun sensor should be placed such that it has an unobstructed view of the Sun for as much of each orbit as possible; the front of the spacecraft is normally a favourable solution. Table 10 provides a mass and power breakdown of these sensor systems with no margin applied (this is instead applied for the entire ACS sub-system later on).
Attitude Determination Sub-system Breakdown
Mass Max Power
Star Mapper 7kg 20W
Sun Sensor 2kg 3W
Total 9kg 23W
Table 10 - Attitude Determination Breakdown
8.2 REACTION WHEEL(S) To provide stabilisation in three axes three mutually orthogonal reaction wheels are required; one for each axis. To mitigate the effects of the failure of a single wheel it is customary to include a fourth mounted such that a component of the torque it applies lies within each axis. Doing so ensures that with a single failure three axis stabilisation using just the reaction wheel is still possible, with the other wheels applying torque to counter the effect of the unwanted components of the torque supplied by the fourth wheel (except of course in the circumstance that it is the three main wheels which remain operational).
There are two main properties of the reaction wheels to define: the torque and the maximum momentum storage. The torque requirement can be driven either by the need to dominate the maximum on-orbit disturbances or to provide actuation for slewing the spacecraft. The principal sources of orbital disturbances are the planetary gravity gradient, solar radiation pressure, magnetic field torques and, for low-altitude orbits, aerodynamic drag. The science orbit for this mission keeps the spacecraft well clear of Uranus’ atmosphere so the last can be entirely disregarded. While
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aerodynamic effects would be in evidence at Earth the attitude control authority would at that point be with the upper stage of the launcher, so is not considered.
The effects of solar radiation pressure applying torque are effectively negligible at Uranus, given the distance to the Sun. The equations used to calculate this and the other disturbance torques are detailed in the appendix, but come from (Wertz & Larson, 1992). At Uranus’ orbit the flux density of the solar wind is approximately 4.1 W/m2, just 0.3% that at Earth. Assuming a spacecraft surface area of 9 m2, a nominal reflectance of 0.6 and centre of mass/centre of solar pressure deviation of 0.3 m this yields a maximum disturbance torque due to radiation pressure of 5.9x10-8 Nm at Uranus. The same calculation for Jupiter orbit (solar flux density of ~56 W/m2) gives a maximum disturbance torque of ~8.1x10-7 Nm.
The gravitational disturbance torque is a function of the moments of inertia of the spacecraft in z and y axes (the largest and smallest axes if different), which were simply approximated to rods of the same mass as the complete spacecraft in those axes, given that the final form of the spacecraft is not completely certain. The maximum disturbance torques due to gravitation for Uranus and Jupiter are 5.99x10-7 and 2.41x10-8 Nm respectively (the latter being relatively small due to the high altitude of the closest approach). The Uranus values assume a maximum of 1 degree deviation from local vertical for the z axis during closest approach, which is appropriate as this section of the orbit would be used for science using nadir-pointing instruments, with the maximum for a worst-case 45 degree deviation being 1.72x10-5 Nm.
The magnetic field disturbance torques are similarly low, with maximum values of 7.62x10-7 and 5.28x10-8 Nm for Uranus and Jupiter respectively. These values assume a residual dipole for the vehicle of 1 Am2, an average value taken from (Wertz & Larson, 1992). Assuming an absolute worst-case disturbance regime the maximum external torque acting on the spacecraft is therefore 2.7x10-5 Nm for Uranus and 1.06x10-6 Nm for Jupiter, which applies a 20% margin to the sum of all three disturbance sources discussed.
Considering that the magnitude of the disturbance torque is low it is no surprise that slewing requirements are the main driver for the reaction wheel torque. To determine this a slew rate of 30° in ten minutes was used, assuming maximum acceleration throughout. To provide this slew rate for the fully-fuelled orbiter (assumed wet mass of 3.3 tonnes) requires a torque of ~22 mNm, ~8 mNm for the dry mass of the spacecraft, more than 28,000 times the max disturbance torque.
The second major property of a reaction wheel, the maximum momentum storage, is mostly dependent upon the disturbance torque however; while the slew torque is much larger those events occur infrequently, while the disturbance torques are applied to some extent throughout the whole duration of every orbit. By far the largest of the disturbance torques is due to the gravity gradient; this results in a sinusoidal disturbance over each orbit. Using an equation to approximate this, taken from (Wertz & Larson, 1992) and using the rms average of a sinusoidal function (0.637), yields an approximate storage requirement of 0.335 Nms for the Uranus science orbit – this using under the assumption that low deviation from local vertical around periapsis would be allowed, though a significant margin should be provided by the reaction wheels to accommodate relatively short periods spent with worst-case pointing at these times, without the need to use thrusters for attitude control.
In order to meet these requirements a number of different off-the-shelf options were investigated, with the MSCI MW1000 (figure 19) found to provide sufficient performance for the task. This option provides a significant (36%) margin on the maximum torque requirement while providing more than 3.2 times the momentum storage (thereby allowing for less than ideal orientations to be maintained throughout the orbit). This solution is also advantageous in that the mass and power requirements are relatively low, as seen in the specification shown in table 11 (with values taken from the MSCI website).
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MSCI MW1000 Properties
Minimum Power per Wheel 2 W
Maximum Power per wheel 9 W
Torque 30 mNm
Angular Momentum Cap 1.1 Nms
Mass 1.4 kg
Max Power 36 W Figure 19 - MW1000 Reaction Wheel, Total Mass 5.6 kg Credit: MSCI Website
Table 11 - Reaction Wheel Properties 8.3 THRUSTERS The role of the thruster system is to act as actuators for large slew manoeuvres such as re-orientation of the spacecraft for communications downlink periods, and to de-saturate the reaction wheels. Additionally, provision needed to be made for additional thrusters to be used as ullage motors for the propulsion system, which would also allow for very small orbital correction manoeuvres to be made with these thrusters instead of the main engine. This could be advantageous as the main engine would introduce vibrations and slight changes in attitude, which would be problematic for certain instruments such as the magnetometer which have very high pointing knowledge requirements.
To determine the slew requirements in terms of thruster force equations from (Wertz & Larson, 1992) were again used and detailed fully in the appendix. The main determining factors are the thruster moment arm, the duty cycle and the desired slew rate. A slew rate of 30° in one minute, with thrusters firing for 10% of that time, was used: this is higher than would realistically be required in almost any circumstance, thereby providing a degree of margin.
For the moment arm values of 1.5 m for pitch and yaw, and 1 m for roll were used; this assumes placement of the thrusters at one end of the spacecraft, most likely on small booms in the same fashion as the Cassini orbiter. Specifically, eight thrusters are called for to affect attitude changes: two on each corner aligned with pitch and yaw axes. This allows for any desired attitude change to be made by firing two thrusters simultaneously. A second redundant set of thrusters should be flown mounted next to the first to take over in the event of failure or potentially to be fired alongside the other set to perform very rapid attitude changes, though this is an unlikely scenario. Additionally, one extra thruster should be mounted per block facing the rear of the spacecraft to act as ullage motors as previously mentioned; therefore calling for a total of 20 thrusters.
The thrust required to perform the described 30° slew in one minute is 7.2 N in the pitch and yaw axes and 10.8 N in the roll, to perform said manoeuvre with the spacecraft fully-fuelled. To determine the required thruster force for
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momentum dumping operations it was assumed that one-second pulses would be used to remove the maximum calculated storage; this leads to a requirement of 0.34 N, so is not a driving requirement. Following this the total number of pulses required was calculated assuming two large manoeuvres consisting of six seconds of thruster firing per orbit, plus one second of thruster time per reaction wheel each orbit, for a total of 1834 pulses per year, 9170 throughout a five year mission. This pulse requirement can then be used, alongside the calculated requirements for thruster force, to provide the total impulse which the ACS thrusters must provide, which was calculated to be ~8570 Ns.
For a relatively small rocket system such as this a pressure-fed setup is generally favourable, with the small total impulse required making a lower efficiency monopropellant system favourable as it decreases overall system mass for such small systems. A number of thrusters were examined from the range which Aerojet produce, which are representative in terms of performance of the chemical thrusters of this type currently available. Electric ACS thrusters were rejected on the grounds of power limitations for the same fundamental reasons as for the main engine; power requirements that would not be feasible to meet. All of the thrusters examined use hydrazine as a monopropellant which is a standard choice for systems of this type.
The Aerojet MR-50T (figure 20) was identified as a viable thruster to use, offering significant margin on thrust while retaining a relatively low mass as can be seen from the properties of the thruster listed in table 12. The catalyst used is
S4O5 and the thruster has a long, proven flight heritage, having first flown in 1974 and used for attitude control on the Viking and Voyager missions among others. A newer thruster in the same class has been developed by the same company: the MR-106L, which uses a S4O5/LCH-202 catalyst and provides marginally improved performance, though it has a less impressive flight heritage, but is still flight proven. The older MR-50T is used here as a base, providing some additional margin given that newer alternatives do offer some incremental improvement. The thruster is rated to provide 26,000 pulses, several times what would be required for the mission.
MR-50T Properties
Thrust 22 N
Specific Impulse 225 s
Individual Thruster Mass 0.68 kg
Propellant Mass 29.36 kg
Propellant Tank Mass 2.1 kg
Figure 20 - MR-50T, Credit: Aerojet Pressurant Mass (Helium) 0.07 kg
Pressurant Tank Mass 0.59 kg
Valve Power 25.3 W
Valve Heater Power 1.96 W
Catalyst Bed Heater Power 3.27 W
Total Power (2 Thrusters) 61 W
Total Mass 54.28 kg
Table 12 - MR-50T Properties
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The mass of propellant required was found simply using the total impulse requirement and the specific impulse of the motor. The propellant tank mass, required mass of pressurant and mass of that tank could then be found in exactly the same way as described for the pressure-fed system in the propulsion system section, with the caveat that both propellant and pressurant tanks were assumed to be spherical, which is appropriate given the small size of said tanks. The chamber pressure of the thruster, upon which the propellant tank pressure requirement is based, is 84 psi (0.69 MPa). It was assumed that the mass of any thruster mountings would be covered by the already substantial margin on the mass of the propellant and pressurant tanks.
The power requirement for each thruster is somewhat deceptive; while to power the valve fully is indeed quite a large value this would only be required for a few seconds each day, with the heaters only operating at full power shortly before this to prepare for firing. This brief but high power requirement makes the use of batteries to provide power to the thrusters desirable.
8.4 OVERVIEW The attitude control system as described is capable of meeting the requirements imposed upon it by other systems, while using proven technologies and architecture. Redundancy is provided in all areas of operation to provide confidence that operations could be maintained even in the event of multiple failures. Table 13 shows a breakdown of the mass and power of the various components which make up the system.
Attitude Control System Breakdown
Attitude Determination 9 kg 23 W
Reaction Wheels 5.6 kg 36 W
Thrusters 55 kg 61 W
Total 83 kg 84 W
Table 13 - Attitude Control System Breakdown The total power requirement indicates the value required for attitude determination to be operational with two thrusters operating at maximum. This is an unlikely scenario, and as previously mentioned the thrusters would only draw this for a few seconds each day. As such a value of approximately 60W would be more realistic maximum to be drawn over any period of time, representing the power for attitude determination to operate at maximum alongside the reaction wheels, though even this is unrealistic as to draw that much power the reaction wheels would all have to be operating at maximum saturation which should be avoided if at all possible. It is still worth considering 60W as an approximate maximum to be expected in order to cover all eventualities however, though such a period would be something of an emergency where one would not expect many instruments to also be operating.
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9 COMMAND AND DATA HANDLING The command and data handling sub-system includes all housekeeping, data processing and formatting while also controlling data traffic and time management on the spacecraft [Maral and Bousquet, 2009]. These functions include but are not limited to:
Decoding, validation and execution of commands, either instantaneously or at a later time Management of traffic between sub-systems and data timing/synchronisation Dating of events and measurements Compression, coding and formatting of telemetry data Monitoring and analysis of spacecraft diagnostic parameters and appropriate decision making Data storage and retrieval
In early space missions’ on-board data handling (OBDH) was often considered part of tracking, telemetry and command (TT&C), however it is now thought of as a standalone sub-system. This is because of the development of powerful flight computers and the expanded functions of the command and data handling sub-system, while TT&C is concerned only with the communications link between the spacecraft and the ground stations
9.1 TELECOMMAND MESSAGE STANDARDS The telecommand message standards arise due to the need for compatibility between the spacecraft and the ground control stations. It also helps stop the execution of erroneous or faulty commands which can cause serious damage to the spacecraft. The command message frame in the ESA standard consists of 96 bits and is as follows [Maral and Bousquet, 2009]:
Address Mode First data Second Third Mode First data Second Third synchronisation selection word data word data word selection word data word data word word repeated repeated repeated repeated 4 bits 12 bits 12 bits 12 bits 16 bits 4 bits 12 bits 12 bits 12 bits
The first word of 16 bits is an address and synchronisation word that specifies the destination decoder. This is then followed by a mode selection word which indicates the type of command. The command is then transmitted in three words of 12 bits each which are repeated for reliability reasons. If the repeated words are found to be different from the original, then the command will be terminated. These commands are not always carried out instantly but can be stored and activated at a later time.
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9.2 INSTRUMENT PROCESSING REQUIREMENTS
A large portion of the on-board data handling will be dedicated to the compression, coding and formatting of telemetry data. The amount of processing power required for each instrument is calculated by first estimating the maximum raw data output, which can be done from instrument parameters.
Table 14 – Science Payload data calculations 1. IR Spectrometer 5. UV Spectrometer Pixels per Image 230400 Pixels per Image 3072 Samples per Pixel 1 Samples per Pixel 1 Bits per sample 12 Bits per sample 12 Estimated Data rate [picture] 2.76E+06 Estimated Data rate [bps] 3.69E+04
6. Narrow angle 2. Microwave Radiometer Camera 5 Different frequencies 5 Pixels per Image 1048576 Samples per scan 128 Samples per Pixel 1 8 bits per sample 12 12 colour filters 12 Scan length [s] 0.5 Bits per sample 12 Estimated Data rate [bps] 1.54E+04 Estimated Data rate [bps] 1.51E+08
7. Particle Detector 3. Wide angle Camera package Pixels per Image 1048576 5 Distinct instruments 5 Samples per Pixel 1 1 Sample per instrument 1 6 colour filters 12 8 bits per sample 12 Bits per sample 12 60Hz Operation 60 Estimated Data rate [Picture] 1.51E+08 Estimated Data rate [bps] 3.60E+03
2. Magnetometer 3 Samples 3 120Hz Operation 120 8 bits per sample 8 2 Magnetometers 2 Estimated Data rate [bps] 5.76E+03
1. The IR spectrometer uses a 480*480 pixel detector which measures 1 sample per pixel, the intensity. Each sample will be comprised of 12 bits. This is consistent with similar instruments such as VIMS flown on Cassini [E. Miller et al. JPL] and NIMS on the Galileo orbiter [NIMS instrument Parameter Document].
2. The microwave radiometer measures energy at 5 different frequencies and takes 128 samples per frequency, per scan. Each sample in the scan has 12 bits and the scan lasts for 0.5 seconds.
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3. The wide angle camera has a 1024*1024 pixel detector and measures one sample per pixel, the intensity. The design consists of 12 colour filters that used together build up a colour image. The camera uses 12 bits per sample, such as found in the HiRise camera [Alan Delamere et al, 2003] giving an amount of data in a single multicolour frame as 75.5Mbit.
4. The magnetometer operates at 120Hz and measures a vector (x, y , z) which translates to 3 samples. Each sample has 8 bits and there will be 2 magnetometers running at all times
5. The UV spectrometer is a non-imaging spectrometer and has 3072 pixels, with 1 sample per pixel and 12 bits per sample.
6. The Narrow angle camera has the same raw data output as the Wide angle camera.
7. There are 5 instruments in the particle detector package measuring different particle flux’. Each has one sample and there will be 12 bits per sample. The detector package operates at 120Hz.
As expected the imagers have the largest data generation, however the magnetometer and particle detector package will probably produce larger totals, due to being run continually throughout the mission.
9.3 COMPRESSION Data compression will be applied according to the consultive committee for space data systems (CCSDS) standards [CCSDS 121.0-B-2]. This mission will employ lossless data compression techniques, which allows for moderate data rate reduction with no additional distortion to the original data. Data compression has benefits including but not limited to:
Reducing the total data rate for a given transmission speed Decreasing the required channel bandwidth
A lossless source encoder is made up of two separate parts, a pre-processor and an adaptive entropy encoder. The initial inputs to the source encoder are comprised of a block of J n-bit samples, as shown below.
A reversible function is applied by the pre-processor which changes the data samples x to a preferred source which can be written:
The ideal output of the pre-processor has specific properties: the {δi} is statistically independent and identically distributed the preferred probability, pm, that any sample δi will take on integer value m is a non-increasing function of value m, for m= 0, 1, ... (2n–1).
The CCDSD standards state that in general, the best lossless pre-processors meet these requirements and in doing so produce the lowest entropy data stream. This is defined as the smallest average number of bits which can be used to describe each sample. The adaptive entropy encoder takes the output of the pre-processor and creates uniquely decipherable, variable length codewords corresponding to each block of samples input from the pre-processor. For each block of J samples the coder selects the coding option that gives the highest compression ratio and attaches an identifier to the first block that the decoder can identify. The CCSDS standards for lossless compression use a Rice algorithm to code the data stream. The Rice algorithm will provide a compression ratio of root 2 [CCSDS 121.0-B-2].
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Figure 21 - Schematic of a typical source coder, [CCSDS 121.0-B-2]
9.4 ICER IMAGE COMPRESSION
As well as the lossless compression of the data stream, data intensive imagers may require further lossy, or lossless, compression in order to fit within the telemetry budget if large numbers of images are taken. The three imagers will use the ICER progressive wavelet image compressor, which can achieve high compression ratios with minimal loss of image quality (A. Kiely et al. 2003). ICER is specially designed to meet the needs of deep space applications while achieving state-of-the-art compression effectiveness. It operates under a progressive data compression scheme, meaning that as more of the data stream is received the rebuilt image increases in quality. This is illustrated in figure 22.
Figure 22 - A sequence of images showing gradual increase in image quality under progressive compression as more data is received. [A. Kiely et al. 2003]
Progressive compression therefore provides a convenient method of meeting a constraint on compressed data without having to guess at a preferred level of image quality. Using ICER it is possible to view a small portion of the data from a compressed image and then decide if it is interesting enough to warrant further transmission, maximising the science return for a given data rate. This could feasibly be completed in a single orbit downlink phase and so not require a huge increase of data storage on the spacecraft. ICER has been used extensively by NASA’s camera heavy Mars rovers and proven to be an effective method of compression, competitive with other forms such as JPEG 2000 (A. Kiely et al. 2003). The Data volumes before and after nominal ICER compression are shown in table 15.
Despite the high quality of the compression technology, data loss is inevitable and so for every ten frames transmitted to Earth, one will be have no lossy compression applied. This will help ensure science return with a minimal increase in data volume. It will further provide a reference for the amount of image quality lost during ICER compression.
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9.5 DIGITAL MODULATION AND FORWARD ERROR CORRECTION There are three primary criteria that are considered when choosing a digital modulation scheme, bandwidth efficiency, power efficiency and system complexity. Bandwidth efficiency is the number of bits-per-second per hertz that can be transmitted by the system, the power efficiency is defined as the required Eb/N0 for a specific bit-error probability (Usually 10-5) and system complexity involves the technical difficulties of the electronic system (Xiong, Fuqin. 2006).The digital modulation scheme that will be used is QPSK, Quad-riphased phase shift keying, also sometimes known as 4-PSK. QPSK encodes two bits per symbol and uses four points on the constellation diagram, shown in figure 23, which represent the four possible carrier phase shifts.
Figure 23 - QPSK constellation diagram (1) 01 is represented by a 45o phase shift, 00 is represented by 135o, 10 is represented by 225o and 11 is represented by 315o. The channel capacity limit for a QPSK modulated carrier is 2-bits per Hz, so it has the advantage of better bandwidth efficiency over BPSK modulation, while maintaining the same bit-error probability (Xiong, Fuqin. 2006). -5 Both figure 24 and [Wertz and Larson] put the Eb/N0 required for a BER of 10 at 9.6 for the QPSK and two higher order PSK modulation schemes.
Eb/N0 requirements can be reduced further by using forward error correction techniques. Extra bits, known as parity bits are added into the stream at the transmitter which allow for correction of bit errors at the receiver. The disadvantage of such techniques is the obviously increased data rate. The communications system will employ the convolutional coding with Viterbi decoding technique. This is implemented by creating and transmitting two bits for each bit of data, which therefore means the data rate is one half of the transmitted bit rate [Viterbi, 1967]. The received coded stream is compared with possible sequences and the possible sequence that most closely matches is chosen. Viterbi decoding -5 reduces the Eb/N0 required for a BER of 10 to 4.4.
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Figure 24 - QPSK BER Vs Eb/N0 performance (1) The total amount of data that will be transmitted by the communications sub-system will be a function of the raw data output, Rice compression, ICER compression and Viterbi forward error correction. This is summarised in table 15 which displays the changes in size as each compression or forward error correction technique is applied from left to right, with the Viterbi column displaying the final data values.
Table 15 - Data volumes as ICER, Rice and Viterbi techniques are applied Raw Data rate (bps, ICER Instrument unless indicated) (Imagers only) Rice algorithm Viterbi IR Spectrometer 1.84E+06 3.69E+05 2.61E+05 5.21E+05 Microwave Radiometer 1.02E+04 1.02E+04 7.24E+03 1.45E+04 Wide angle Camera 7.55E+07 1.51E+07 1.07E+07 2.14E+07 UV Spectrometer 3.69E+04 3.69E+04 2.61E+04 5.21E+04 Magnetometer 5.76E+03 5.76E+03 4.07E+03 8.15E+03 Radio Science Experiment 0 0 0 0 Narrow angle Camera 1.51E+08 3.02E+07 2.14E+07 4.27E+07 Particle Detector package 2.40E+03 2.40E+03 1.70E+03 3.39E+03 TOTAL 2.28E+08 4.57E+07 3.23E+07 6.47E+07
9.6 PROCESSOR AND DATA STORAGE
The processing power in the command and data handling sub-system will be provided by the RAD750 radiation hardened single board computer. The RAD750 is a single core 110 to 200 MHz processor which can achieve over 266 MIPS [RAD750 Specification document]. The CPU can survive 200,000 to 1,000,000 rads and temperatures ranging from -55 to 125 degrees Celsius. The RAD750 has extensive flight heritage on a variety of spacecraft including Mars Reconnaissance Orbiter, Kepler Space telescope and Juno, launched in 2005, 2009 and 2011 respectively. The RAD750 computer will handle the identification, trafficking and compression of the instrument data. The processing power will also be used to support other parts of the spacecraft such as thermal control and attitude and orbit control algorithms. For reliability all components in the command and data handling sub-system will be dual redundant and if one fails the spacecraft will automatically switch to the other.
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Figure 25 – RAD750 processor, BAE Systems
The spacecraft will carry two solid state recorders with a 2 Gigabyte storage capacity each. These will be rad-hard solid state recorders using SD-RAM. Power constraints of the mission dictate that most instruments cannot be fully powered on at the same time as the communications system is sending telemetry data to Earth. The solid state recorders will therefore be used to store the science data until such a time when downlink to Earth is established. This on board storage is the same size as that which was flown on the Cassini orbiter [Linda J. Spilker, The Cassini-Huygens Mission]. In the event of this filling up data will need to be discarded and techniques can be used to try to minimise the science loss. The nature of this will depend on the specific science data being stored at the time, an example being the deletion of every tenth magnetometer value.
There are many pre-existing rad-hard solid state recorders and so the TRL for this component is high. The chosen part for this mission should have a total dose limit >100Krad and have a single event upset per-day rate of <1E-10 in earth geo. This will be a substantial margin on the expected radiation levels and length of the mission and so minimise any negative radiation effects.
(1) http://en.wikipedia.org/wiki/Phase-shift_keying
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10 COMMUNICATIONS
10.1 MISSION COMMUNICATIONS OVERVIEW
The uplink and downlink communications between the ground stations and spacecraft can be divided into three main categories: Telemetry, Tracking and Command (TT&C).
Telemetry comprises all automated monitoring of the spacecraft and can include thousands of individual functions that are checked for temperature, voltages and pressure, to give three examples. It is therefore extremely important in determining if all spacecraft systems are functioning correctly and any subsequent troubleshooting if they are not. Telemetry also encapsulates all the science data sent to earth. This includes everything measured by the instruments in the scientific payload and so can vary hugely in size depending on which instruments are in operation. This will be the bulk of downlink data once the science phase of the mission begins. Tracking provides the mission operators with updated spacecraft ephemeris and location, while Command allows elements of the spacecraft to be controlled by mission operators, including propulsion, ACS and instruments. The mission data can be split into two areas: uplink TT&C and the orbiter – earth TT&C downlink.
Figure 26 - Data Flow
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10.2 COMMUNICATION SUBSYSTEM REQUIREMENTS AND CONSTRAINTS As stated in section 2.1, the communications subsystem provides tracking, telemetry and command functions for the orbiter. It also forms the major component of the radio science experiment and so the system shall therefore:
Transmit carriers for radio science Receive carriers for radio science Transmit carriers for downlink TT&C Receive carriers for uplink TT&C
These requirements can be expanded by looking at the functions of telemetry, tracking and command individually and the constraints that are imposed by them.
10.2.1 Telemetry Requirements The communications subsystem in sync with the command and data handling sub-system will generate and modulate the downlink telemetry carriers that must be compatible with the ground station network. The major factor specifying the telemetry requirements of the mission is the data rate required for maximum science return, which will dictate the required transmitter power and antenna size. In practice, the transmitter power usually scales linearly with data rate for a given antenna size. The telemetry system will also transmit the outputs from many sensors within the spacecraft and transmit this data to Earth. For a large mission such as this there will be hundreds of sensors located in critical areas of the spacecraft. For example, to measure current in electrical components, pressure in the fuel tanks and the temperature of various sub-systems. Although there are hundreds of separate sensors, once in space, these parameters are expected to vary slowly and housekeeping will be a fraction of the total data rate. The science data will constitute the largest portion of downlink data. Depending on which instruments are in use the amount of data will fluctuate dramatically and so on-board data storage will be required to keep any data that cannot be transmitted to earth immediately. This is especially true for data intensive imagers such as the wide/narrow angle camera and the Infra-red mapping spectrometer.
10.2.2 Tracking Requirements
The communication tracking system can use a number of techniques to determine the location or current orbit of the spacecraft at any given time. Velocity and acceleration sensors can be monitored to establish the change in orbit from the last known position and the Doppler shift of the carrier can be measured to determine the rate at which the range is changing. Measurement of exact distance can be achieved by timing how long a pulse, or sequence of pulses, takes to return from the spacecraft, while the ground antenna azimuth and elevation can give the angular location. Further information can be found by analysis of images of Uranus, its satellites and star fields, returned in telemetry. Image quality for optical navigation requires a bit error rate (BER) no larger than 10-5 [J. Taylor et al. 2002].
10.2.3 Command Requirements The communication system will receive and demodulate command waveforms sent to the orbiter by the ground network. The command structure is vital to the success of any mission and is used by the controllers on the ground to make manual changes. These could include attitude and orbit control corrections, instrument or subsystem operations and maintenance. The command process must include safeguards against accidental operation and ensure a high probability that all commands will be received error free.
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10.2.4 Fault Protection and Reliability Requirements As with other subsystems, the integrity of the mission necessitates that there are no single faults that would result in the permanent loss of a critical function. The critical functions for the communications subsystem are the ability to transmit telemetry, command capability and two way coherent carrier tracking. This will usually need to be implemented by having dual-redundant critical components for example exciters and RF power amplifiers (E. Demircioglu, 2008).
10.3 RF VS OPTICAL TRADE-OFF
The first major design decision is the choice of communications system. Two candidates have been selected as possibilities, RF communication and optical communication. The parameters that will be examined in the trade-off are:
Link Reliability Antenna mass Antenna power consumption TRL Data rate capability
RF communication systems have been used extensively on previous missions and have been proven with many missions boasting above 99% link reliability [J. Taylor et al. 2002]. The technology is also mature and most designs are built upon flight heritage. Optical communications on the other hand have not been widely used in the deep space environment. Optical communications suffer from heavy attenuation in the atmosphere resulting in link loss during bad weather. Due to the low beam divergence of optical communication there are far greater pointing constraints on the S/C than RF. The better system for link reliability is therefore RF, as both link establishment and link maintenance are easier. The optical system has both mass and power advantages because it does not require a large high gain parabolic antenna that an RF system needs. The low divergence of the beam means that less power is required. The TRL of an RF system is obviously higher than that of an optical system because of the long flight heritage of RF communication systems in deep space and the lack of flight heritage for optical systems. This will have a direct impact on both the cost and development time. Optical systems have the capability for larger data rates than RF systems because of the higher frequencies involved and can reach many Mb/sec. This will not be such an important factor in our mission because the instrument suit we have selected is not so data heavy as to require data return at such speeds.
Table 16 - RF Vs Optical Trade-Off Summary
Link Antenna Antenna Data Reliability Mass Power TRL Rate Total Weight 0.3 0.1 0.1 0.2 0.2 1 Optical -1 1 1 -1 1 -0.1 RF 1 -1 -1 1 -1 0.1
The trade-off is summarised in table 16 which applies weighting factors to the parameters discussed above. The output of the trade-off is that an RF communications system will be used, which is consistent with all previous deep space missions.
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10.4 GROUND STATIONS AND DOWNLINK WINDOW
10.41 GROUND STATION NETWORK CHOICE The RF communications system will require ground stations to receive, demodulate and identify the data sent from the S/C. The three options that have been identified are:
1. Co-operation with NASA and use of the Deep Space Network (DSN) 2. Use of the European Space Agency’s ESTRACK network 3. Construction and operation of a dedicated network, designed around our specific mission
Despite the obvious advantages of having a dedicated network of antennas, the phenomenal initial cost quickly render it unfeasible for this mission. The parameters that are used considered when deciding between options 1 and 2 are: Cost Availability System sophistication Coverage
Both systems have many similarities and can be considered comparable in areas such as cost and availability. The advantage of the DSN comes from its coverage and antenna sizes. The DSN complexes are situated equally spaced around the earth providing constant coverage for deep space missions. They also have centralised complexes, whereas the ESTRACK network is far more spread out, meaning the possibility of frequent movement between sites. The operation of 70-m antennas in the DSN allows for greater link strength and can be invaluable at distances such as Uranus. In addition all DSN complexes have both uplink and downlink Ka band capabilities, whereas only certain ESTRACK sites do. These factors lead to NASA’s DSN being the most suitable choice for the mission
Figure 27 - Arial view of the Canberra DSN complex (Credit: NASA)
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10.42 DEEP SPACE NETWORK COVERAGE Due to the large distances involved, tracking a spacecraft in deep space requires far fewer ground stations than one that is in a low Earth orbit (LEO). Uranus will be visible for extended periods of time from large portions of the Earth’s surface. The DSN has three complexes situated in California, Madrid and Canberra approximately 120 degrees apart around the earth.
Table 17 - DSN Locations
DSN Latitude Longitude Altitude (m) Goldstone CA 33.59⁰N 116.848 ⁰W 1063 Madrid Spain 40.1⁰N 3.685⁰W 670 Canberra Australia 35.52⁰S 149.008⁰E 1036
Figure 28 - DSN Coverage, NASA
10.43 SPACECRAFT ORBIT The specific orbit of the spacecraft around Uranus is important in the context of the communications sub-system as it may pass behind the planet as viewed from earth. If this was a substantial period of time there may be consequences for the available telemetry budget. This will not be an issue for this mission as the spacecraft will adopt a highly eccentric polar orbit in the reference frame of Uranus, meaning there will be almost continual line of sight to Earth. The specifics of the orbit are discussed in section 5.
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10.5 DATA GENERATION A large constraint of the mission is the available downlink telemetry rate as mentioned in section 2.2.1. In order for all generated data to be returned the mission will use on board data storage and employ distinct science phases, with particular instruments focusing on a specific science area. Of the entire scientific payload, the largest challenge for the telemetry downlink will be the imagers, with the narrow and wide angle cameras producing 42.7 Mbit images each after compression. This will limit the imagers to a specific number of images per science cycle. The different science modes are outlined in table 3, which displays the status ON/OFF of the instruments and the maximum data rate in bps or bits-per-frame, for the imagers.
Table 18 - Science Phases Data Generation Option 1 Option 2 Option 3 Option 4 Option 5 Zonal Winds Uranus Uranus All Instruments and Weather Structure, Rings Magnetic Field Atmosphere Dynamics and Moons
IR Mapping ON 7.82E+05 ON 7.82E+05 ON 7.82E+05 ON 7.82E+05 ON 7.82E+05 Spectrometer Radio Science ON 0.00E+00 ON 0.00E+00 ON 0.00E+00 ON 0.00E+00 OFF 0.00E+00 experiment Microwave ON 2.17E+04 ON 2.17E+04 ON 2.17E+04 OFF 2.17E+04 OFF 2.17E+04 Radiometer Plasma Particle ON 5.09E+03 OFF 5.09E+03 ON 5.09E+03 ON 5.09E+03 ON 5.09E+03 Detector ON 5.21E+04 ON 5.21E+04 OFF 5.21E+04 ON 5.21E+04 OFF 5.21E+04 UV Spectrometer Narrow Angle ON 4.27E+07 ON 4.27E+07 OFF 4.27E+07 ON 4.27E+07 ON 4.27E+07 Camera Wide Angle ON 4.27E+07 ON 4.27E+07 OFF 4.27E+07 ON 4.27E+07 ON 4.27E+07 Camera ON 8.15E+03 ON 8.15E+03 ON 8.15E+03 ON 8.15E+03 ON 8.15E+03 Magnetometer
Total data 8.63E+07 8.63E+07 8.17E+05 8.63E+07 8.62E+07 generation
It is highly unlikely that all instruments will be operating simultaneously due to power and data constraints and so option 1 will not be looked at in further detail. Options 2-5 need to be broken down further by indicating the maximum length of time an instrument can operate, or the maximum number of images that an instrument can take during each orbit, with the driving parameter being the available downlink per orbit.
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10.6 TYPICAL SCIENCE ORBIT BREAKDOWNS
Each option is divided into different configurations that correspond to a single reference orbit (A, B, C… ect.). These decisions are made based on both power consumption and data generation. The two instruments that will run continually throughout the mission are the magnetometer and particle detector package and so will constitute the bulk of the data. Numbers in red indicate that the number represents the number of frames taken and not the length of time the instrument is collecting data per reference orbit.
Table 19 - Option 2 Data rates
Option 2
Uranus Atmosphere
Time Time Time A active B C active active bps (hr) bps bps
OFF 0 0 ON 7.82E+05 400 OFF 0 0 IR Mapping Spectrometer OFF 0 0 OFF 0 0 ON 0 20 Radio Science experiment
OFF 0 0 ON 2.17E+04 4 OFF 0 0 Microwave Radiometer
ON 5.09E+03 106.056 ON 5.09E+03 106.056 ON 5.09E+03 106.056 Plasma Particle Detector OFF 0 0 ON 5.21E+04 4 OFF 0 0 UV Spectrometer ON 4.27E+07 20 OFF 0 0 OFF 0 0 Narrow Angle Camera ON 4.27E+07 20 OFF 0 0 OFF 0 0 Wide Angle Camera ON 8.15E+03 106.056 ON 8.15E+03 106.056 ON 8.15E+03 106.056 Magnetometer 6.43E+0 Total data 85428876 6.76E+09 869096.46 13237.039 5.05E+09 9 generation
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Table 20 - Option 3 Data Rates
Option 3
Zonal Winds and Weather
Time active A Time active (hr) B bps bps (hr) OFF 0 0 ON 7.82E+05 400 IR Mapping Spectrometer OFF 0 0 ON 0 0 Radio Science experiment OFF 0 0 ON 2.17E+04 4 Microwave Radiometer ON 5.09E+03 106.056 ON 5.09E+03 106.056 Plasma Particle Detector OFF 0 0 OFF 0 0 UV Spectrometer ON 4.27E+07 20 OFF 0 0 Narrow Angle Camera ON 4.27E+07 20 OFF 0 0 Wide Angle Camera ON 8.15E+03 106.056 ON 8.15E+03 106.056 Magnetometer 85428876 6.76E+09 816962.89 5.68E+09 Total data generation
Table 21 - Option 4 Data Rates Option 4
Uranus Structure, Rings and Moons
Time Time Time A B active C active active (hr) bps bps (hr) bps (hr) OFF 0 0 ON 7.82E+05 400 OFF 0 0 IR Mapping Spectrometer OFF 0 0 OFF 0 0 ON 0 10 Radio Science experiment OFF 0 0 OFF 0 0 OFF 0 0 Microwave Radiometer ON 5.09E+03 106.056 ON 5.09E+03 106.056 ON 5.09E+03 106.056 Plasma Particle Detector OFF 0 0 ON 5.21E+04 4 OFF 0 0 UV Spectrometer ON 4.27E+07 20 OFF 0 0 OFF 0 0 Narrow Angle Camera ON 4.27E+07 20 OFF 0 0 OFF 0 0 Wide Angle Camera ON 8.15E+03 106.056 ON 8.15E+03 106.056 ON 8.15E+03 106.056 Magnetometer 85428876 6.76E+09 847374.14 6.12E+09 13237.039 5.05E+09 Total data generation
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Table 22 - Option 5 Data Rates Option 5
Magnetic Field
Time active A bps (hr) ON 7.82E+05 400 IR Mapping Spectrometer OFF 0 0 Radio Science experiment OFF 0 0 Microwave Radiometer ON 5.09E+03 106.056 Plasma Particle Detector OFF 0 0 UV Spectrometer ON 4.27E+07 20 Narrow Angle Camera ON 4.27E+07 20 Wide Angle Camera ON 8.15E+03 106.056 Magnetometer 86210880 7.08E+09 Total data generation
The largest amount of data per orbit taken from tables 19-22 is 7.08Gbit with all the compression techniques discussed above applied. The communications system must therefore be designed to transmit at least 7.08Gbit of data, with margin, every orbit. If the communications system cannot meet this requirement data will have to be discarded or science operations reduced. A further small margin will need to be present to account for the housekeeping data, although this will be small in comparison to the total.
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10.7 LINK BUDGET
10.71 INTRODUCTION The relationship between transmitted power, data rates, path lengths, frequency and antenna sizes is defined by the link budget. The link budget uses these parameters to output the signal to noise (SNR) ratio of the communications subsystem and allows trade-offs to be made [Wertz and Larson, 2003].
The output of the link budget after all parameters have been chosen is ⁄ , which is defined as the ratio of received energy-per-bit to noise density. This then allows us to find the bit error rate, or BER, of the link depending on the chosen modulation scheme.
10.72 THE LINK EQUATION The standard equation used in the link budget is
(1)
Where P is the transmitter power, Ll is the line loss, Gt is the transmitter antenna gain, Ls is the free space path loss, La is the atmospheric path loss, Gr is the receiver antenna gain, K is Boltzmann's constant, Ts is the system noise temperature and R is the data rate.
It is often easier to express equation (1) in decibels (dB) as the standard link equation is a product of successive terms.
Taking each term in dBs, which is 10log10 of the number, equation (1) becomes
(2)
Where ⁄ , Ll, Gt, Ls, La, Lpr and Gr are in decibels, P is in decibel-Watts (dBW), R is in bps Ts is in Kelvin (K) and
10log10K = -228.6 dBW/(Hz*K).
The link equation will be used to calculate ⁄ for the radio frequency bands S, X and Ka, with centre frequencies of
3, 10 and 33.5 GHz respectively. Each term in equation (2) is explained in detail in the appendix attached to this document, along with all relevant equations used in the calculations. The results were created in Excel and the workbook is also described in the appendix.
⁄ values were obtained for S, X and Ka frequency bands using both the 34 and 70-m DSN antennas.
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10.73 DETAILED LINK BUDGET RESULTS
Table 23 - Downlink Ka-band (34-m DSN, S/C HGA)
TELEMETRY DOWN-LINK PARAMETER INPUTS Up/Down-Link Two-Way RF Band Ka Transmit Antenna 3m HGA Telecom Link DSN 34-m Bandwidth 30000 Hz Data Rate 45000 bps
Operations Mode Nominal Mission Phase Uranus Science Orbit Attitude Pointing Earth Pointed, 0.05 degree max pointing error Range 2.60+09 Km
Link Parameter Unit Value
TRANSMITTER PARAMETERS S/C Transmitter Power dBW 18.45 S/C Line Loss dB -1.5 S/C Antenna Gain dBi 57.90 S/C Pointing Loss dB -0.69
PATH PARAMETERS Space Loss dB -311.25 Atmospheric Attenuation dB -0.92
RECIEVER PARAMETERS DSN Antenna Gain dBi 80.04
NOISE SOURCES Sky Noise K 54.75 Planetary Noise K 0.29 Galactic Noise K 0 Solar Noise K 0 Receiver Internal Noise K 31 Total K 86.046
CARRIER PERFORMANCE Carrier Loop SNR (CNR) dB 51.28
TELEMETRY PERFORMANCE Available Eb/N0 dB 4.75
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10.74 SUMMARY Table 24 summarises the results and also includes the link budget calculations for S-band and X-band, which can be found in detail in the appendix. S-band is not suitable for this mission as the available Eb/N0 is too low and will not -5 provide the value of 4.4 needed for a bit error rate of 10 . The Eb/N0 of X-band is only acceptable if a 70-m ground station antenna is used, which could limit link availability. Ka-band however has a suitable Eb/N0 when using 34-m antennas with a margin of 0.35dB which identifies it as the best option after the link budget calculations. Currently Ka- band is not compatible with the 70-m antennas, although the DSN goes through procedural upgrades and so these may be available by mission launch. Using a telemetry rate of 45Kbps and a 53 hour downlink time the data transmission per orbit is 8.61Gbit. This provides a margin of 1.53Gbit which will account for additional housekeeping data and any other rise in science data. The orbit of the spacecraft is likely to change during the mission and subsequently the total data generation per orbit will fluctuate. Any increase in orbit size, and therefore data per orbit, can be accounted for both by the margin shown on the reference orbits above and by increasing the allocated telemetry downlink window. Other interesting outputs of the link budget worth mentioning are that Ka-band has the largest path loss followed by X- band and then S-band. X-band has the lowest system noise temperature and Ka-band has the largest parabolic antenna gain.
Table 24 - Link Budget Summary
Results- Summary Eb/N0 Variations S X Ka -12.48 -1.64 4.75 34m Antenna -4.72 5.98 NA 70m Antenna
10.8 ANTENNA SIZE - POINTING ERROR: PARAMETRIC STUDY During the link budget calculations a trade-off was undertaken to determine the worst case pointing constraints that would be required during telemetry transmission. Essentially this trade-off determines the mass, power and complexity of the high precision ACS system which will be used. The variables are the available Eb/N0 provided by the architecture of the communications sub-system and the pointing accuracy.
In the trade off 4 different antenna diameters were used (2, 2.5, 3 and 3.5m) with Eb/N0 plotted against increasing pointing accuracy for each in X-band and Ka-band. S-band was not calculated as it had already been ruled out as an option. The transmission power remained at 70W. Varying antenna diameters produce different sized beamwidths, which affects the amount of pointing loss the signal experiences for a given pointing offset and carrier frequency. Pointing offset was increased in steps of 0.01 degrees from 0 to 0.5 degrees. The results are shown in figure 29.
Clearly, as the antenna diameter increases so does the starting Eb/N0 with Ka-band seeing the largest rise. In addition to this, the sensitivity of Ka-band to pointing error rises significantly with a larger dish diameter. X-band is far less affected by the changes in dish size, however would require other changes in the sub-system, such as input power, to achieve an adequate Eb/N0. If this could be achieved an X-band system would allow for greater pointing error and so a less complicated ACS system. The pointing error at which the frequency band with the largest Eb/N0 switches from Ka- band to X-band moves from 0.26 degrees with a 2m antenna to 0.15 degrees with a 3.5m antenna and so does not change dramatically. Then examining the point at which the Eb/N0 of Ka-band falls below 4.4, it is found that this moves from 0.06 degrees with a 3m antenna to 0.08 degrees with a 3.5m antenna. This small increase in pointing error will not be sufficient to justify the mass penalty of a 0.5m diameter increase.
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Figure 29 - Pointing Offset Vs Eb/N0 for varying dish diameter
The conclusion of this study is that without dramatically increasing the power of the system a 3m Ka-band design will be the most suitable, however it will require high pointing accuracy. There will therefore be a considerable pointing constraint of 0.05 degrees on the ACS system. A 3.5m would give marginally better performance between 0 and 0.05 degrees pointing offset, however this will have a large mass penalty and does not justify the small performance increase. It would also suffer far greater signal losses if the S/C could not achieve the desired accuracy.
10.9 X-BAND VS KA-BAND ARCHITECTURE In addition to the link budget, differences can be found in the sub-system architecture required for carrier generation. In this area Ka-band also excels in comparison to X-band. Due to the higher frequency the overall size of the components needed to produce the carrier are reduced and this will benefit both the volume requirements and mass requirements of the system. The amount of power used is another area where Ka-band carrier generation is preferential and small, however not insignificant savings are made.
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10.10 SUBSYSTEM ARCHITECTURE
Table 25 - Communications Sub-System Architecture Architecture Description Comments High Gain Antenna 3m Parabolic dish
Low Gain Antenna Dual omnidirectional Antennas Positioned on opposite sides of the S/C to maximise coverage Telemetry Radiofrequency band Ka-band 33.5GHz
Radio-science frequency band Ka-band 33.5GHz
Maximum Data Rate 45kbps
Downlink time per orbit Minimum of 53 hours
Transmitter power 70W
Maximum allowable pointing Error 0.01 Degrees When transmitting at maximum data rate from a Uranus science orbit Modulation scheme QPSK modulation with Viterbi Viterbi results in a doubling of the bit forward error correction stream for a given amount of data
Compression Rice algorithm and ICER Compression by root 2 and 5 image compression respectively Bandwidth 30kHz
Maximum data downlink per Uranus 8.61Gbit orbit Downlink Margin 1.53Gbit Comfortably accommodates for the addition of housekeeping data on top of the science data Ground Stations NASA Deep Space Network
The communications sub-system will use Ka-band carriers for both the radio science and telemetry needs of the spacecraft. This is possible because the spacecraft will not transmit telemetry at the same time as transmitting carriers for radio science. Due to this only one radiofrequency band is required which is favourable because a single system has power, mass and complexity advantages over one with multiple radiofrequency bands. The exciters, TWTAs and electronics will all be dual redundant in order to maximise the reliability of the system and meet the fault protection requirements. The only non-redundant component will be the HGA although this is considered highly reliable. The 3m HGA will need to be accurately earth pointed with a minimum offset of 0.05 degrees during transmission to provide the necessary signal to noise ratios, as is discussed in section 6.5.
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10.11 MISSION COMMUNICATION PHASES
The mission communications can be divided into distinct phases, Launch and Earth orbit, Cruise and Uranus arrival and science orbits. Each phase has particular operating environments and requirements.
10.12 LAUNCH AND INITIAL ACQUISITION After launch the communication subsystem will be turned on and initial checks will be required to check all functions are operating as expected. This will involve the first acquisition of the DSN uplink which relies on the following.
The spacecraft must be in view of a transmitting DSN antenna The spacecraft receiver must be functioning correctly The spacecraft transmitter must be functioning correctly The spacecraft ACS must have correctly oriented the spacecraft
During Earth escape and early cruise the S/C will use the HGA as a sun-shield, rendering it unavailable for communication. The S/C will therefore use a LGA for initial acquisition and communication, which has the benefit of reduced pointing constraints. The short distances involved at this stage of the mission and the low data rates being transferred mean there is no significant reduction in link quality when using the LGA over the HGA.
10.13 CRUISE The main events during the cruise phase are Start of Jupiter transfer Jupiter gravity assist No major science is planned for the cruise phase and so the data rates are expected to be low and limited to navigation and housekeeping. The major functions for the communication system will therefore be navigation via Doppler and ranging data and the transmission of low rates of command and telemetry. When required the HGA can be sun-pointed.
10.14 URANUS ARRIVAL AND SCIENCE ORBITS The main events of this mission phase are.
Uranus arrival (SOI) Science Orbit Moon tour
This phase will see the largest amount of data return and will require the use of on-board solid state recorders during the periods where science data is being generated but the communications system is turned off. The spacecraft will need to be accurately earth pointed during downlink in order to maintain the necessary signal-to-noise ratio.
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11 RADIATION As with all deep space missions careful analysis of the radiation environment the spacecraft will experience needs to be considered during design. Radiation hardening is the act of making the electrical components resistant to the damages that can be caused by ionising radiation. Primary mechanisms that occur are Ionization effects and Lattice displacement [Wertz and Larson].
Lattice displacement is lasting damage that occurs when a particle changes the arrangement of the crystal lattice. Ionization effects contribute to the total ionizing dose (TID) of the system, degrading performance over time and causing errors or failures in their performance. Specific errors that occur in digital systems will be discussed below.
11.1 RADIATION SOURCES
Van Allen Radiation Belts are trapped particle regions surrounding the earth, caused by the interactions of particles and the Earth’s magnetosphere. Exact conditions can vary wildly depending on the sun and magnetosphere at a given time. Other planets with magnetic fields, which includes Uranus, also have trapped particle regions, varying in strength and size.
Cosmic Rays are an unavoidable flux of approximately 85% protons, 14% alpha particles and 1% heavier ions, combined with gamma and x-ray radiation that come from all directions in space.
Solar Particles are ejected from the sun and made up of large numbers of high energy protons and heavier ions
11.2 DAMAGE TO DIGITAL SYSTEMS
There are many possible effects of high energy particles impacting on digital systems but only three of the most common will be discussed here.
Single event upsets are caused by an ion interacting with an electronic chip and causing a bit to change state. In modern chips this should not cause lasting damage to the device.
Single event latchups happen when a heavy ion or high energy proton pass through the component causing it to ‘short’. This can occur in chips that have a parasitic PNPN structure. A single event latchup requires the device to be power cycled. If the effect is large enough it is possible for permanent damage to be inflicted and the component may fail.
Single event transient is an unwanted signal travelling through the circuit due to the build-up of charge from ionisation events.
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11.3 MISSION ENVIRONMENT The two major radiation environments that the spacecraft will be exposed to during the mission are found at Jupiter, during the gravity assist, and at Uranus. Jupiter has an incredibly harsh radiation environment due to its large and powerful magnetosphere (Divine. N, 1983). Any spacecraft that passes through this radiation will require substantial shielding of internal electronics or risk electronic degradation and catastrophic failures (Fieseler, P, 2002). This will not however be a driving factor in this mission because of the large altitude of the Jupiter gravity assist, currently 25.86 Jupiter radii. The radiation environment at Uranus is far less intense than that of Jupiter but will provide the majority of total mission dose due to the time spent in orbit. When Voyager 2 visited Uranus it found radiation belts of a similar intensity to those at Earth (E. C. Stone, 1987). Standard models of the Van Allen belts can be obtained from the National Space Data Science Centre. These are titled AE8MIN and AE8MAX for minimum and maximum electron fluxes and AP8MIN, AP8MAX for Protons. The energy of these particles will not necessitate significant additional shielding to be added to the spacecraft on top of its main structure. Particularly vulnerable components such as detectors and electronic boards should still be of concern and either radiation hardened or housed within the spacecraft at locations that provide the maximum possible protection. As with Earth orbiting satellites conductive layers should be used in order to mitigate the risk from local surface charging effects.
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12 MASS AND POWER It needs to be underlined that due to the strong linkage between the mass and power requirements for the Spacecraft, both are included in this section.
The Power Subsystem study was conducted in two stages, which are:
1. Before the Preliminary Design Review 2. After the Preliminary Design Review.
The difference between these two stages is due to new System Requirements which resulted in a limitation of the power and mass of the spacecraft. Before the Preliminary Design Review the Spacecraft dry mass was estimated as 1421.67Kg, where the mass of the Orbiter was 1061.1Kg (excluding systems margin) and the mass of the Probe was 123.66Kg (excluding systems margin). Adding these 2 values and including a systems margin results in a total dry mass of 1421.67Kg.
Furthermore the total power consumption when all the Instruments and Subsystems are running simultaneously was estimated to be 662.01W, while the development of individual case scenarios predicted that the power requirements vary between 279.10W and 389.22W.
12.1 PRELIMINARY DESIGN SYSTEM REQUIREMENTS Following the PDR, new system requirements were created such as:
ID-UO-00100 The Orbiter total dry mass shall be below 1200Kg. Justification: The Spacecraft should be launched by the Falcon Heavy. Comments: Total mass excluding system margin is 1000Kg.
ID-UO-00200 The radioisotope power sources (RPS) shall be ASRG (Advanced Stirling Radioisotope Generator) 241-Am. Justification: Current RPS in development by ESA.
ID-UO-00300 Up to four ASRG 241-Am shall be permitted for this mission. Justification: Mass and cost constraints.
ID-UO-00400 Each ASRG 241-Am shall produce up to 100We. Justification: Current ASRG 241-Am under development by ESA.
ID-UO-00500 The Spacecraft shall have an operation lifetime of up to 17 years after the ASRGs 241-Am systems are turned on. Justification: The Sterling pistons have a lifetime of up to 17 years.
To summarise these systems requirements: ID-UO-00100 to ID-UO-00500, limit the total spacecraft mass, the power source and the total power available.
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12.2 POWER AND MASS TABLES Table 26 - Model Payload characteristics for the Uranus Orbiter (Hubbard, 2010; Dougherty, et al., 2011; Spencer, 2010). Model Instrument Mass Instrument Power Other Instrument Information Instrument Mass Shielding Shielded Margin Total Mass Power Harness Margin Total Power Volume (mm) Max Data Heritage TRL Source/ Values (Kg) Margin Mass (TRL) Kg consumption Margin (TRL) consumption Rate (kbps) based on: (Kg) (W) (W) Instruments 60.5 16.25% 72.02 20% 84.40 98.3 5% 20% 122.88 16000 IR Mapping 10 20% 12 20% 14.40 7 5% 20% 8.75 400*400*300 ~250 Juice 7-9 (Arridge, et al., Spectrometer 2011) (Dougherty, et al., 2011) (Spencer, 2010) Radio Science 0.5 5% 0.53 20% 0.63 1 5% 20% 1.25 50*50*50 NA – Carrier Juice, 7-9 (Dougherty, et al., experiment only Cassini 2011) (Arridge, et al., 2011) (Spencer, 2010) Microwave 20 20% 24 20% 28.80 30 5% 20% 37.5 150*150*150 ~2 JUNO, 7 (Arridge, et al., Radiometer (Electronics) Juice 2011) (Dougherty, Receiver et al., 2011) antennas 200*200*200 Plasma 10 20% 12 20% 14.40 15 5% 20% 18.75 350*400*250 5-50 THEMIS, 7-9 (Dougherty, et al., Particle JUICE 2011) Detector (Angelopoulos, 2008) UV 6.5 20% 7.8 20% 9.36 16 5% 20% 20 300*300*200 30 Phebus on 7 (Norton, n.d.; Spectrometer (no baffle) Bepi Spencer, 2010; Colombo Greeley & Dougerty, 2010) Narrow Angle 8.6 20% 10.32 20% 12.38 15 5% 20% 18.75 500*200*200 12600 Lorri, New 7 (Hubbard, 2010) Camera Average Horizons (Dougherty, et al., 1300 2011) (Arridge, et al., 2011) (Arridge, 2013) (Palumbo & Jaumann, 2014) (Greeley & Dougerty, 2010) Wide Angle 1.5 20% 1.8 20% 2.16 12.3 5% 20% 15.37 157*117*104 3000 MDIS, 7 (Palumbo & Camera Messenger Jaumann, 2014) (Chassefiere & Maria, 2010) Magnetometer 3.4 5% 3.57 20% 4.28 2 5% 20% 2.5 100*60*60 7-70 THEMIS 7-9 (Angelopoulos, (sensor, *2) 2008) 160*160*120 (electronics box)
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The total orbiter instrument mass is 84.40Kg. In addition to the TRL (contingency) margin, a system margin of 20% needs to be included. Therefore the total mass of the orbiter’s instruments is ~101.28Kg (Dougherty, et al., 2011). The power consumption of the instruments could go up to 122.88W if all instruments run simultaneously.
Table 27 - Estimated Equipment and Subsystem characteristics for the Uranus Orbiter (Hubbard, 2010; Dougherty, et al., 2011; Spencer, 2010).
Subsystem Mass Subsystem Power Other Subsystem Information
Power Total Power Source/ Additional Shielded Margin Total Mass Harness Margin Volume Subsystems Mass (kg) consumption consumption TRL Values Shielding mass (Kg) (TRL) (Kg) Margin (TRL) (mm) (W) (W) based on:
Total 751.31 1.63% 761.14 23% 965.31 367.8 5% 20% 459.75 Instruments 60.5 16.25% 70.33 20% 84.40 98.3 5% 20% 122.88 7
Structure & (Hubbard, 165 0% 165 20% 198 - - - 7 Mechanism 2010) (Hubbard, Propulsion 91.22 0% 91.22 20% 109.46 76 5% 20% 95 7-8 2010) (Hubbard, 2010) Power 19 0% 19 20% 22.8 20 5% 20% 25 7 (Spencer, 2010) (Williams, et al., 2012) 4 ASRG 241 Am 200 0% 200 50% 300 - - - 5-6 (Ambrosi, 2014) (Hubbard, 2010) Thermal control 39.35 0% 34.38 20% 47.22 - - - 7 (Williams, et al., 2012) 3m dish (Hubbard, Communication 54 0% 54 20% 64.8 70 5% 20% 87.5 7 (40Kg) 2010)
Command & Data (Hubbard, 14.65 0% 14.65 20% 17.58 58 5% 20% 72.5 7 Handling 2010)
Control, (Hubbard, Navigation, 68.88 0% 68.88 20% 82.66 45.5 5% 20% 56.88 8 2010) Guidance (Hubbard, Harness 32 0% 32 20% 38.4 - - - 7 2010)
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Table 27 highlights that the Spacecraft dry mass is 965.31Kg. Including a system margin of 20% the total dry mass of the Spacecraft is 1158.4Kg (Dougherty, et al., 2011). A limit of 1200Kg dry mass was given in systems requirements: ID-UO-00100. If all the subsystems and instruments are turned on simultaneously the total power consumption could go up to 459.75W if no precautions are taken. Therefore, different power scenarios are developed as demonstrated in Table 33, Table 34 and Table 35. Maximum power produced by four ASRG-241 is 320W as will be demonstrated in section 12.4.
Table 26 and Table 27 in this Power Section represent the expected power and mass requirements of each subsystem and payload. These tables highlight the mass differences between the unshielded mass, the shielded mass and the shielded mass including a contingency margin. The tables show the percentage differences between these three values. The percentage differences for additional shielding varies between 0% (no additional shielding), 5% (very low additional shielding required if any) and 20% (large additional shielding mass required). On top of the additional shielding mass, a contingency margin has been included. The values for this margin vary between 20% and 50%.1 20% was selected for instruments or subsystems whose TRL (Technology Readiness Level) are above, or equal to 7. All values below 7 have a margin of 50%. Furthermore after finalising the values of each subsystem, an additional system mass margin of 20% is included (Dougherty, et al., 2011). Table 26 and Table 27 also highlight the power requirements of the subsystems and instruments. These values are separated again into instrument or subsystem values which include a 20% margin and a 5% harness margin. Furthermore these tables also demonstrate volume, maximum data rate, heritage and sources of the information (Greeley & Dougerty, 2010) (Hubbard, 2010) (Spencer, 2010).
Due to the limited resources and current stage of this study, exact values for some of the subsystems in Table 27 cannot be determined at this stage. Consequently, worst case values from previous studies, such as the decadal study survey ‘Titan Saturn System Mission’, the ‘Ice Giants Decadal Study’ or the ‘Uranus Pathfinder’ were used (Greeley & Dougerty, 2010) (Hubbard, 2010) (Spencer, 2010) (Arridge, et al., 2011).
The final mass value within Table 27 provides us with the dry mass of the Spacecraft. Following the Preliminary Design Review, the TRL values of the instruments and subsystems have been lowered in consideration of the fact that there has been no previous mission to Uranus except Voyager 2.
The final dry mass of the Spacecraft is 1158.4Kg (including systems margin), which is just below the maximum mass limit for the Spacecraft. The methods used to reduce the mass of the Spacecraft are demonstrated in section 12.3.
12.3 METHODOLOGY OF MASS REDUCTION Following the Preliminary Design Review, new requirements for the mission were set, which are given in section 12.1. These resulted in the request to minimise the Spacecraft mass by 221Kg.
Furthermore due to the change from the ASRG 238-Pu to ASRG 241-Am, further mass reduction needed to be considered. Before the PDR it was considered to include up to five (4 +1 spare) ASRG-238Pu in our mass estimates. Each ASRG-238Pu was considered to have a shielded mass of approximately 20.2Kg, which resulted in a total mass of 181.8Kg including TRL and system margins (Bly & Di Pietro, 2007).
Due to new requirements ID-UO-00200 to ID-UO-00400, up to four ASRG-241Am are needed to be selected for our orbiter. Each ASRG-241 requires a shielded mass of 50Kg, which results in a total mass of 360Kg (including TRL and system margins). The mass difference between these two RPS systems is 178.9Kg and therefore the spacecraft requires a total mass reduction of 399.9Kg. Furthermore the detailed study of the Control, Navigation and Guidance subsystem required 32.81Kg more than originally expected. This results in a necessary mass reduction of 432.7Kg (including margins).
1 Margin values based on module PA7412, Session 8: Technical Budgets
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12.3.1 Trade Off To reduce the mass of the Spacecraft, the following factors have been identified which could lead to the mass reduction: 1. Reinvestigate the subsystem values, as most subsystems have not been studied in detail before the PDR, most values selected need to remain worst case scenario values based on previous studies unless studied in more depth. Justification: a. Mass and power requirement for Structure and Mechanism, Propulsion, Power, Thermal Control, Command and Data Handling, Control Navigation Guidance Harness were all originally based on worst case values of previous studies.
2. Reducing the dish size for the communication subsystem from 3.5m down to 3m. Justification: a. The reduction of the size of the dish could safe 10Kg, while minimal power increase occurs (excluding margins).
3. Reducing the probe which would reduce the total dry mass of the spacecraft by 123.66Kg. Justification: a. Due to strong mass restrain, the reduction of the probe becomes necessary as most other subsystems of the spacecraft had to remain worst case scenario values until further study can predict the requested mass and power requirements of them. b. The payload of the probe has a low TRL (5), which leads to a higher mass margin.
4. Reducing the mass and power requirements of the plasma particle detector. Justification: a. The plasma particle detector has a mass before PDR of 31.45Kg which can be reduced to 14.4Kg, while its high power requirement of 60W can be cut down to 18W. Comment: Remove the mass spectrometer.
5. Reducing the values for Command and Data Handling from 55.3Kg down to 32.1Kg (including margins). Justification: a. Previous values were based on Planetary Science Decal Study Survey: Titan Saturn System Mission. Apart from this survey two further surveys to Uranus were considered, which both required lower mass for this subsystem. These two surveys to Uranus are more similar to this proposed mission, dependent on mass, payload, power and environment and therefore present a more suitable comparison.
6. Adapting the value for the Thermal Subsystem Justification: a. The in-depth study of the Thermal Subsystem, lead to the conclusion that only 39.38Kg are required instead of the previously considered 82Kg (both values excluding margins). b. Power and mass reduction due to RTHU. Comment: Values include mass for MLI, heaters, radiators, heat pipes, Louvers, RTHU for both the spacecraft and the payload.
7. Adapting the value for the Power Subsystem Justification: a. The in-depth study of the Power Subsystem leads to the conclusion that only 19Kg are required instead of the previously considered 88.5Kg (both values excluding margins). Comment: Values include mass for Battery, Power Distribution Unit and Power System Electronics.
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8. Adapting the value for the Propulsion Subsystem Justification: a. The in-depth study of the Propulsion Subsystem leads to the conclusion that only 91.22Kg are required instead of the previously considered 154Kg (both values excluding margins). b. Previous values were based on a spacecraft exceeding a dry mass of 2000Kg (Planetary Science Decal Study Survey: Titan Saturn System Mission).
9. Remove the Wide Angle Camera from the payload. Justification: a. The Wide Angle Camera may not be necessary as it fulfils the same science objectives as the Narrow Angle Camera. Removing this camera from the payload could save 2.59Kg.
10. Reducing the mass for the Structure of the Spacecraft Justification a. Reduce the Probe Separation System and other miscellaneous Probe related components, as this saves 20Kg (excluding margins).
A summary of these trade off options is given in Table 28.
Table 28 - Mass Reduction Possibilities. Possibility to Reduce CASE A Dish Size CASE B Power Subsystem CASE C Probe CASE D Plasma Particle Detector CASE E Command and Data handling Subsystem CASE F Thermal subsystem CASE G The Wide Angle Camera CASE H Propulsion Subsystem CASE I Structure
It becomes obvious by comparing these trade-off options that they contain different categories: those which can definitely be reduced (CASE B, C, E, F, H and I) and those which could be removed if it is unavoidable (CASE A, D and G).
Table 29 - Mass differences between the trade-off options. Initial mass in kg Suggested new mass in kg Mass decrease in kg CASE A 92.16 77.76 14.4 CASE B 127.44 27.36 100.08 CASE C 148.39 0 148.39 CASE D 31.45 14.4 17.05 CASE E 55.30 21.10 34.2 CASE F 118.08 49.65 56.88 68.43 61.2 CASE G 2.59 0 2.59 CASE H 221.76 131.36 90.40 CASE I 266.4 237.6 28.8 Total mass 497.11 decrease
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If all of these suggested mass reduction methods would be taken into account, the final mass of the spacecraft would fulfil requirement ID-UO-00100, leaving a flexibility of 64.4Kg. This could enable CASE A or D and G to remain within the spacecraft’s payload. In Table 8 and Table 9 as well as in the Thermal Analysis in Section 13, it is assumed that the Plasma Particle Detector and the Antenna-Dish have been downsized. Ideally, the full size Plasma Particle Detector and the Wide Angle Camera remain in the Spacecraft. The greatest disadvantage of the large Plasma Particle Detector is its large power requirements, which need to be reduced in a worst case scenario, see section 12.4.2.
12.4 POWER AND ASRG 241-AM SUBSYSTEM At the beginning of the analysis of the power and the ASRG subsystems multiple power sources were considered (see Table Table 30) as well as Batteries, Power distribution Units and the Power System Electronics. As previously mentioned four ASRG 241-Am have been selected as power source and the Power distribution Unit (PDU) and a Power System Electronics (PSE) (including a Shunt Regulator Unit (SRU)) are still included in the subsystems. The Spacecraft makes use of a regulated bus, where the bus voltage of the spacecraft is controlled during the entire mission. A bus voltage of 30V ±2V has been selected. Excess power created by the four ASRG 241-Am is stored within the battery or dumped as heat by shunt dissipaters. The power distribution unit consists of fault protection, cabling and switches, to turn the power for spacecraft loads on and off (Europa-Study-Team, 2012) (Patel, 2004).
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Table 30 - Power Source Selection Power Source Notes ASRG 241-Am ASRG 241-Am are currently under development by ESA. The acronym ASRG stands for Advanced Stirling Radioisotope Generator, and the fuel for this ASRG is Americium (241Am). Americium is considered by ESA to be a more ideal fuel as it is available in Europe at a lower
price than Plutonium. Americium has a high half-life: = 432 years and a specific power between 2-2.2W/Kg. An ASRG 241-Am contains two Stirling engines; each Stirling engine
converts 250Wt with 20% efficiency into 50W. The disadvantage of ASRG compared to RTGs is the 17 year lifetime of the Stirling engines. Typically, the total power production is 80W including a 20% margin (100W excluding margin) for each ASRG 241-Am. Its power loss is 1% every 10 years. An ASRG 241-Am producing 100W (excluding margins) requires a fuel mass of 6Kg. After 17 years of operations, four ASRG 241-Am still produce 312W. This can be determined by applying the following equation:
where P0 is the power produced at the beginning of the mission and t is the mission duration. From the specific power of Americium it can be determined that each ASRG-241 requires a mass of between 45.46Kg and 50Kg in order to produce 100W (excluding margins). As ASRG have never been flown before, a TRL below 7 is expected for this device for this mission, which results in a contingency mass margin of 50%. Therefore the total mass of four ASRG 241-Am is 360Kg (including margins) (Ambrosi, 2014; Williams, et al., 2012). ASRG 238-Pu Conversely, ASRG 238-Pu was under development by the NASA. This development has been paused over the previous years due to its high development costs. The ASRG 238-Pu was assumed to have a mass of 20.2Kg (excluding margins) and to have a half-life of 88 years. For this mission, up to 5ASRG-Pu would have been required. Each ASRG 238-Pu produced less than 143W at the start of the mission and only 120W at the end (excluding margins). Each ASRG 238-Pu required a fuel mass of 1.2Kg. Due to the shortage of 238-Pu in Europe and in the United States as well as its high development costs, the application of Plutonium should be avoided for this mission (NASA, 2013; Bly & Di Pietro, 2007). RTG (MMRTG) 238-Pu The only current alternative RPS option to ASRG 241-Am is the RTG (MMRTG) 238-Pu, which stands for Multi-Mission-Radioisotope Thermoelectric Generator. The MMRTG 238-Pu, is currently under development and has been flown in the past. Therefore it has a higher TRL than ASRG 241-Am. It requires 4.5Kg Plutonium as a fuel and produces up to 125W at the beginning of life (mass of 44Kg without margins). After 17 years this system only produces 95W, which is less compared to ASRG 241-Am. Its main advantage is its higher TRL and therefore only requires a 20% contingency margin and does not require a spare RPS (Rowe & Abelson, 2006). RTG (GPHS) 238-Pu RTG (GPHS) 238-Pu, is a previous NASA RTG, whose development was halted mainly due to the high cost and shortage of 238-Pu. The last mission using the GPHS 238-Pu was New Horizons, which already used spare parts from previous missions and a combination of old and new fuel. It typically had a mass of 55.9Kg, while the fuel required 8.1Kg alone and produced up to 285W at launch (Bennett, 2006; Rowe & Abelson, 2006). Solar Arrays As an alternative to RPS, solar arrays could be investigated. The spacecraft requires up to 300W 17 years after launch. By applying the following equation:
where ( ), , r= 19.2 AU, P= ~300W, it becomes clear that the solar panel has a minimum area of : A =649.7m2. Taking annual degradation of 1.03 of up to 17 years into account, this leads to a final solar panel area of A=1100m2, which is unachievable. Other factors which would need to be included are the decrease of solar panel efficiency due to low temperature and intensity as well as a low TRL value. There is a low TRL for such a solar arrays due to the surface area, low temperature and solar intensity environment (Arridge, et al., 2011; Arridge, 2013).
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Table 30 represents all the suggested power sources which were considered to power the Orbiter and notes relevant to each of these power sources. It is noteworthy that ASRG 238-Pu and RTG (GPHS) 238-Pu are both currently not in development and that Solar Array would require a too large surface area and mass to be viable for this mission. Therefore RTG (MMRTG) 238- Pu and ASRG 241-Am are the only viable options. Table 32 represents a trade-off for all of the power sources while Table 31 highlights the driving factors for the trade-off. For this study we considered only 7 driving factors which are listed as the following: power source lifetime (does the power source still produce more than 90W after 17years?), fuel and material cost (Pu is very difficult to get access to and very expensive), mass (does the component weight more than 50Kg?), volume, European Union/ESA production, reliability (based on TRL values, in what stage of development is the component) and activity (is the component currently still in development). Each of these factors are given a certain weight to contribute to the final result. Table 31 - Driving Factors for Trade-off Parameters Units Weighting factor Lifetime Years 0.10 Fuel/material cost Euro 0.15 Mass Kg 0.15 Volume m3 0.15 EU/ESA – production 0.25 Reliability (TRL) 0.05 Activity 0.15 Total 1
Table 32 - Power source trade-off results Low Fuel Mass Low Lifetime EU/ESA- Reliability, Recent Results or material <50Kg Volume ~17a production TRL >7 Activity cost W >90W ASRG 1 0 1 1 1 0 (TRL 5) 1 0.8 241-Am ASRG 0 1 1 1 0 0(TRL5) 0 0.4 238-Pu RTG 0 1 1 1 0 1 (TRL 7) 1 0.6 (MMRTG) 238-Pu RTG 0 0 1 1 0 1 (TRL8) 0 0.3 (GPHS) 238-Pu Solar 1 0 0 1 1 0 1 0.65 Arrays
From Table 32, it becomes obvious that ASRG 241-Am is a much more suitable power source for our particular mission compared to RTG (MMRTG) 238-Pu due to its longer lasting higher power production. Furthermore the difficulty to access 238-Pu at the current stages makes the ASRG 241-Am the more ideal power source for the orbiter.
Following the PDR, four ASRG 241-Am were selected for this mission, producing 320W (including 20% margin). Before the PDR, five ASRG 238-Pu were selected mainly due to their high power production and lower mass requirements (181.18Kg less than four ASRG 241-Am). In the case that ASRG-241-Am are not available for this mission, MMRTG are the only viable alternative.
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Table 33 - Power Operation Scenarios (Hubbard, 2010).
Subsystem Launch Cruise TCM UOI Communication IDLE Propulsion OFF 0 OFF 0 ON 75% 57 ON 65% 49.4 OFF 0 OFF 0 Power ON 20 ON 20 ON 20 ON 20 ON 20 ON 20 Communication ON 70 ON 70 ON 40 ON 40 ON 70 OFF 15 Command ON 58 ON 58 ON 58 ON 58 ON 58 ON 58 Control ON 45.5 ON 45.5 ON 45.5 ON 45.5 ON 45.5 ON 45.5 Instruments OFF 0 ON 2 ON 2 ON 2 ON 17 OFF 0 Power 193.5 195.5 222.5 214.9 210.5 138.5 Harness 5% 203.18 205.28 233.63 225.65 221 145.4 Total Power +20% Margin 243.81 246.33 280.35 270.77 265.23 174.5 [in W]
TCM: Trajectory correction manoeuvre. UOI: Uranus Orbit insertion.
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Table 34 - Instruments Power Operation Scenarios (Hubbard, 2010).
Option 1 Option 2 Option 3 Option 4 Option 5 All Uranus Atmosphere Zonal Winds and Uranus Structure, Magnetic Field Instruments Weather Dynamics Rings and Moons A B A B A B IR Mapping Spectrometer ON 7 ON 7 ON 7 ON 7 ON 7 ON 7 ON 7 ON 7 Radio Science experiment ON 1 ON 1 ON 1 ON 1 ON 1 ON 1 ON 1 ON 1 Microwave Radiometer ON 30 ON 30 ON 30 ON 30 OFF 0 OFF 0 OFF 0 OFF 0 Plasma Particle Detector ON 15 OFF 0 OFF 0 ON 15 ON 15 ON 15 ON 15 ON 15 UV Spectrometer ON 16 ON 16 ON 16 OFF 0 ON 16 ON 16 OFF 0 OFF 0 Narrow Angle Camera ON 15 ON 15 OFF 0 OFF 0 ON 15 OFF 0 ON 15 OFF 0 Wide Angle Camera ON 12.3 OFF 0 ON 12 OFF 0 OFF 0 ON 12.3 OFF 0 ON 12.3 Magnetometer ON 2 ON 2 ON 2 ON 2 ON 2 ON 2 ON 2 ON 2 Total Power [in W] 98.3 71 68 55 56 53.3 40 37.3
Table 35 - Science Case Power Operation Scenarios (Hubbard, 2010).
Subsystem Science Option 1 Option 2A Option 2B Option 3 Option 4 A Option 4 B Option 5 A Option 5 B Power ON 20 20 20 20 20 20 20 20 Communication OFF 15 15 15 15 15 15 15 15 Command ON 58 58 58 58 58 58 58 58 Control ON 45.5 45.5 45.5 45.5 45.5 45.5 45.5 45.5 Instruments ON 98.3 71 68.3 55 56 53.3 40 37.3 Power 236.8 209.5 206.8 193.5 194.5 191.8 178.5 175.8 Harness 5% 248.64 219.96 217.14 203.18 204.23 201.39 187.43 184.59 Total Power including 20% Margin 298.37 263.97 260.57 243.81 245.07 241.67 224.91 221.51
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As not all instruments and subsystems run simultaneously, different operational scenarios were developed to determine the spacecraft’s maximum power level. Therefore Table 33, Table 34 and Table 35 were created. Table 34 is an instrument power table based on tables in the Juice Yellow Book and the Ice Giants decadal study document (Hubbard, 2010) (Dougherty, et al., 2011). Option 1 highlights the rare point where all instruments are used at the same time. Options 2 - 5 are based on the instruments required to analyse specific scientific question categories after the developed traceability matrix for Uranus.
Table 35 determines the total spacecraft power required for each individual science case. Table 33 highlights the power required for different phases of the spacecraft such as the power requirement during the launch or the cruise. From Table 26 and Table 27 the power requirements, without margins, of each subsystem and instrument have been selected to develop Table 33, Table 34 and Table 35. The additional harness and contingency margins are included in Table 33 and Table 35 in the bottom two rows (Greeley & Dougerty, 2010) (Hubbard, 2010) (Spencer, 2010). Table 33 and Table 35 demonstrate that the final power requirement varies between 174.51W and 298.37W.
12.4.1 Power reduction Following team discussion after the PDR, the power requirements of the communication subsystem can be reduced from 99W down to 70W (excluding contingency margin), which is based on the team’s study made for the Uranus orbiter communication subsystem.
Furthermore, the reduction of the power of Plasma particle detector which originally required 50W, was decreased to 15W (excluding margin). This strongly reduced the spacecraft’s power requirements. As was mentioned in section 4.8, the Plasma Particle Detector contains multiple instruments within its configuration which detect: ions (Thermal ion detector, higher energy ion detector), neutral plasma (Energetic Neutral Analyser and Ion & neutral gas mass spectrometer) and cold plasma.
A full size plasma particle detector is ideal to have in this spacecraft’s payload. Due to the high power requirement of 62.5W (including margins), this instrument becomes critical. In the case that all four ASRG-241-Am are operational, this instrument can be fully functional and would require up to 70% of the payload’s power depending on the science scenarios. But in the case that one ASRG-241 is not functional, only 240W is available for all the subsystems and instruments. When this low power is provided, only the Plasma Particle Detector could be turned on. Therefore the power required for the plasma particle detector has to be reduced. Reduction is possible by turning the mass spectrometer off as this reduces the power required for the instrument to 18.9W. Alternatively it was considered to replace the plasma particle detector in the mass trade-off with a smaller plasma particle detector (14.4Kg instead of 31.45Kg) which requires 18.9W (excluding the mass spectrometer). Both changes have a strong impact on the power requirements of the spacecraft and demonstrate that four ASRG 241-Am will be sufficient.
The instruments and subsystem power scenario Table 33, Table 34 and Table 35 have undergone slight changes since the PDR, such as adapting the communication values or separating some power scenarios in two sections. Section A represents the use of a Narrow Angle Camera while the Wide Angle Camera is turned off, while section B highlights the use of the Wide Angle Camera while the Narrow Angle Camera is turned off, as this reduces the power requirements for the individual science scenarios.
After analysing the Thermal subsystem of the Spacecraft, it was determined that the power required to maintain the Spacecraft within temperature limits can be reduced due to the application of RTHU-241-Am by 30W.
Within the science scenario the maximum power requirements vary between 221.51W and 263.97W (excluding scenario 1 which is when all the instruments are running at the same time). The maximum power operation scenarios for the different stages of the mission vary between 174.51W and 280.35W. This shows that four ASRG 241-Am are
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sufficient when they are combined with a powerful battery. The power requirements have been significantly reduced since the PDR, where the estimated power requirements scenarios varied between 279W and 385W.
12.4.2 Worst case scenario for the Power Subsystem In a worst case scenario, one of the four ASRG 241-Am may not be functional. This could result in the three remaining, functional ASRG 241-Am supplying the power to the entire spacecraft. If such an eventuality occurs, a battery must be included within the spacecraft. This worst case scenario is possible as ASRGs have never been tested before in space application.
As four ASRG 241-Am provide 320W (400W total), three ASRG 241-Am only produce 240W including margins (300W total).
For this worst case scenario of only three operational ASRG 241-Am, the 20% included power margin could technically be ignored. This is due to other precautions which have been considered in the spacecraft’s mass and power budget, such as:
a. 20% power margin included in each subsystem level. b. 5% harness margin included for each subsystem level. c. 50% mass contingency margin for the ASRG-241-Am d. 20% systems margin for the ASRG-241-Am e. We assumed that an ASRG has a mass of 50Kg and that the Specific Power of Americium varies between: 2- 2.2W/Kg. As a precaution, we selected 2W/Kg. However, if the actual Specific Power is 2.2W/Kg the ASRG could produce 88W (110W total). f. The power scenarios provided are worst case scenarios for all the different stages of the mission or different science scenarios. This is due to the fact that in each of these science scenarios, the chance of all the instruments running simultaneously is unlikely.
If the ASRG power margin is not ignored, the application of the battery becomes crucial, see section 12.6.
12.5 REFERENCE ORBIT To be able to determine the additional power requirements of the Spacecraft during Uranus orbit, sample orbit durations need to be determined. The orbit of the Spacecraft is a polar orbit (inclination ~ 96°), where the Perigee is assumed to 2 5 be at 4RU and Apogee at 28RU. Therefore the orbit duration takes 6.71 10 s (7.77days). Furthermore we know that the average distance (in RU) from the natural satellites to Uranus are as follows:
Oberon 23.09373 RU
Titania 17.26786 RU
Umbriel 10.52773 RU
Ariel 7.555429 RU
Miranda 5.141175 RU All of the orbits of the moons are in an equatorial orbit. As the Spacecraft and the Moons are in different orbits, the approximate intersection point of the two orbits needs to be determined. Knowing the Perigee and Apogee of the selected orbit, allows the semi-parameter to be determined. This is the approximate point where the Spacecraft and the equatorial orbit cross. This is achieved when the Spacecraft is at a distance of 7RU from Uranus. Therefore, between
4Ru and 13.5 RU the Ring and Moons can be analysed and the spacecraft is most likely to apply Science Option 4 from
2 Possible Alternative Orbits are: Perigee 4RU and Apogee at 18RU, or Perigee at 1.1RU and Apogee at
18RU
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Table 35. Furthermore it can be assumed that half of the orbit time the Spacecraft is in Communication mode. To further simplify the problem, it is assumed that the spacecraft’s remaining time is split equally between the three remaining Science Options. Alternatively the Spacecraft could remain in one science scenario option and communication option for an entire orbit. Depending on the additional power requirements for that mode, the balance between the science scenario options (Options 2, 3, 4 require power) and communication option (power available) has to be altered. The power available from batteries in one orbit is 612W (see Appendix).
Table 36 - Reference Orbit Power Requirements Uranus Zonal Uranus Magnetic Field Communication Structure, Winds and Atmosphere Rings and Weather Moons Dynamics
Example Orbit in RU 4-13.5 13.5-17.4 17.4-20.4 20.4-22.9 20.4-28 Percentage in each mode 20% 10% 10% 10% 50% Time in hours 37.28 18.64 18.64 18.64 93.21 Power required/h -3.37 -3.81 -22.269 16.791 12.57 Total Power -125.61 -71.03 -415.13 313.01 1171.62 required/delivered
12.6 BATTERY In the event that one ASRG-241-Am fails, the remaining three ASRG must provide enough power for the spacecraft, with the help of batteries. Due to the lack of power for the spacecraft, the power consumption of the Communication, Cruise and Launch phase power scenario needs to be downsized by reducing the power of the communication subsystem from 70W to 40W (excluding margins). Also, the power requirements of the Plasma Particle Detector need to be reduced from 50W to 15.12W (excluding margins). Alternatively the Spacecraft might need to be placed into the IDLE mode for a certain time period of the orbit around Uranus, to recharge its batteries (See Appendix).
Assuming that the Spacecraft is for 1 hour in TCM (Technical Correction Manoeuvre) mode, 40.35W will be required from the batteries.
Furthermore, if we assume that the Spacecraft is for 1 hour in the UOI (Uranus Orbit insertion) stage, 30.77W will be required from the batteries.
Finally in an orbit around Uranus, 612W (Table 36) are required from the battery (See Appendix).
Furthermore, if we assume a 90% depth of discharge in the Uranus Orbit, a total battery capacity above 680W is required. Ideal batteries for this mission are Li-ion batteries due to their excellent depth of discharge characteristics and their operation within the temperature range of -50°C and 20°C (Williams & Bannister, 2014). The bus voltage of the Spacecraft is 30V and the current is 22.7A. The battery has a capacity of 22.64Ah (679.2 Wh), which results in a minimum battery mass of 7.55Kg. A 10Kg mass has been allocated to the spacecraft for the battery (Williams & Bannister, 2014). A Lithium Battery has an average cell voltage of 3.6V. Consequently, 9 cells are required (Bin Mazlan, 2009). We know that this battery has an energy efficiency of 78% and therefore enough power is available to recharge the batteries in an Orbit (See Appendix).
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13 THERMAL In this section the thermal design of the spacecraft and the antenna-dish are discussed. It is assumed that the electronics, batteries, spacecraft-structure, propulsion system are included in the spacecraft and have to be kept ideally between 0°C and 40°C. Furthermore the individual instruments must be kept between specific operation temperature ranges. Therefore, they have their individual heaters or cryogenic systems for which additional mass-budget is considered (Williams & Bannister, 2014; Larson & Wertz, 2005).
Every component in a spacecraft has its thermal limits and is damaged if the environmental temperatures are too cold or too hot. Therefore, a detailed study needs to be conducted to prevent a spacecraft and its components from becoming too cold or too hot during each phase of the spacecraft’s mission (Martinez, 2014).
For our particular spacecraft, three thermal models have to be completed. These are the spacecraft’s orbit around Earth, around Uranus and during the flight from Earth to Uranus.
13.1 THE CASE SCENARIO Two bodies are included in these thermal models. One circular disc which represents the Antenna-dish, with a radius of 1.5m and a cube (Spacecraft) ~ 2m in length. The centre of both objects is separated by 1.2m, where a hollow cylinder connects these two objects. The Antenna dish has a mass of ~ 40Kg and is painted white on both sides. Its emissivity is 0.853 and its absorptivity is 0.252 (Larson & Wertz, 2005). Gilmore et al. (2002) argue that a combination of Aluminium and Kapton is never used to cover the external side of an antenna-dish as “a conductive aluminium layer is not transparent to radio-frequency (RF) energy. Any material used in the path of an antenna beam must be close to 100% RF transparent so as not to attenuate the signal” Figure 30 - Antenna dish and Spacecraft (Gilmore, 2002). One example of 100% RF transparent Configuration material is white paint.
The Spacecraft has a mass between ~ 1200Kg (dry mass) and 3000Kg (wet mass) and is made out of Aluminium and Kapton, with an emissivity of 0.03 and absorptivity of 0.12 (Gilmore, 2002). Spacecraft structures are commonly made out of Aluminium and Kapton.
During these thermal models, the spacecraft uses the Antenna-dish as a sunshield. This is because temperature variations for the spacecraft should be minimised (Martinez, 2014). The Antenna-dish side facing the sun is called the external side, which is referred to as 'e' in Figure 30. The side facing the Spacecraft is the internal side and referred to as 'i'.
The surface material of the Antenna-dish could be different on the internal and external side. Generally white paint is applied on the external (solar exposed) side and Aluminium (ε=0.05 α=0.17) with Kapton (ε=0.83 α=0.7) on the internal side. However this has little influence in this case study.
The maximum temperature limit for the Antenna dish is 398K, while the minimum temperature limit is 158K (273K±125K). The absolute temperature limits are 273K±150K, therefore a 20% margin has been included (Hubbard, 2010).
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The highest temperature limit for the spacecraft is 315K (42°C), while the minimum temperature limit is 223K (-50°C) (Hubbard, 2010). The spacecraft must ideally be kept within the temperature range of 273K to 315K to maximise other subsystems' and instruments' performance. To maintain the spacecraft within this temperature range,
195WT (234WT including margin) needs to be supplied by multiple RTHU. Typically the mass for one RTHU ranges between 0.9Kg and 2.86Kg, where 1Kg produces ~ 25WT (Williams, et al., 2012).
13.2 EARTH ORBIT
During the Earth Orbit, it is assumed that the spacecraft orbits around earth in an equatorial orbit at 200km altitude. By applying Kepler’s Third Law, an orbit duration of 5.3 103s (88.34min) was determined. Furthermore the eclipse time is given by (Martinez, 2014):
√ Eq. 13.1
3 TEclipse = 0.423TOrbit = 2.2 10 s.
For this study, it is assumed that the spacecraft completes two orbits around Earth before leaving Earth orbit to enter the Transfer stage from Earth to Uranus. The starting temperature for the Earth's orbit is assumed to be 300K for both bodies.
During both the Earth Orbit and the Uranus Orbit, a complex Thermal balance equation is applied for both the Antenna and the Spacecraft. This is explained in more depth in the Appendix where the meaning of each term is explained.
To determine the Antenna Temperature variation over time the Equations: Eq. 13.2 and Eq. 13.3 are applied (Williams & Bannister, 2014; Martinez, 2014):
Eq. 13.2 ( ) ( ) ( )
[ ( ) ( ) ( )]
Eq. 13.3
To determine the Spacecraft temperature variation over time the following equation is applied instead of Eq. 13.2.
( ) ( ) ( ) Eq. 13.4
( ) ( )
13.2.1 Analysis of the Results Within this study it is assumed that the Spacecraft and the Antenna dish are launched with a starting temperature of 300K. The assembly completes two orbits around Earth before entering the transfer phase. During those two orbits, the Antenna dish's average temperature decreases to 281.77K, while the Spacecraft average temperature increases to 300.72K.
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450
400
350 300 300 300 281.76 250 281.76
200 Temperature Temperature in K 150
100 0 2000 4000 6000 8000 10000 12000 Time in seconds Cold Limit Hot Limit Analysis Minimum Maximum
Figure 31 - Transient Thermal Analysis for Dish during the Earth Orbit. Temperature decreases from 300K to 281.76K. Temperature of the Spacecraft 320
300
280
260
240
Temperature Temperature in K 220
200 0 2000 4000 6000 8000 10000 12000 Time in seconds Cold Limit Hot Limit Analysis
Figure 32 - Transient Thermal Analysis for Spacecraft during the Earth Orbit. Temperature varies between 300K and 300.63K. Both Figure 31 and Figure 32 demonstrate that the Antenna dish and the Spacecraft are kept within the expected Thermal Limits and no strong temperature variations are observed.
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13.3 TRANSFER PHASE In this study, a 10 year long transfer phase from Earth to Uranus was estimated and no planetary flybys have been included during this phase.
For the transfer phase, the Antenna temperature variation over time equation has been reduced from Eq. 13.2 to the following equation (Williams & Bannister, 2014; Martinez, 2014) :
Eq. 13.5 ( )
[ ( ) ( ) ( )]
To determine the Spacecraft temperature variation over time Eq. 13.4 has been reduced to:
( ) ( ) ( ) Eq. 13.6
In Eq. 13.5, the value for distance ‘d’ constantly increases, which causes the direct solar radiation to decrease continuously. At Earth orbit the direct solar radiation contributed up to 2565W, while at Uranus Orbit the contribution of this factor is only 8W. This has a significant impact on the temperature change of the Antenna dish.
Figure 33 shows, the temperature of the Antenna dish decreases from 281.76K to 173.03K. Note that 281.76K is the final value of the Earth Orbit Phase and the starting value of the transfer phase for the Antenna dish. Similarly, Figure 34 demonstrates the temperature variation of the Spacecraft during the transfer phase. From the end of the Earth Orbit to the entry of the Uranus Orbit, the Spacecraft’s temperature decreases from 300.63K to 282.57K.
Temperature of the Antenna-Dish
390
340
281.76 281.76 290
240 Temperature Temperature in K 190 173.03 173.03 140 0.00E+00 1.00E+08 2.00E+08 3.00E+08 Time in seconds Cold limit Hot Limit Analysis Maximum Minimum
Figure 33 - Transient Thermal Analysis for Dish during the Transfer phase. Temperature decreases from 281.76K down to 173.76K.
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Temperature of the Spacecraft 320 315
310 305 300.63 300 295 290 285
Temperature Temperature in K 280 282.57 275 270 1.00E+00 1.00E+08 2.00E+08 3.00E+08 Time in seconds Cold Limit Hot Limit Analysis Minimum Maximum
Figure 34 - Transient Thermal Analysis for Spacecraft during the Transfer phase. Temperature decreases from 300.63K down to 283K.
13.4 URANUS ORBIT With regard to the Uranus Orbit, it is assumed that the Spacecraft orbits around Uranus in an elliptical polar orbit. 5 Perigee is assumed to be at 4RU and Apogee at 18RU. Therefore, the Orbit duration takes 3.85 10 s (4.44days). By 4 applying Eq. 13.1 the eclipse time can be determined: TEclipse = 0.064TOrbit = 2.464 10 s. In this instance an Apogee value of 18RU was selected as the time the spacecraft is in the eclipse is larger and therefore the final temperatures for the Antenna-dish and the Spacecraft are lower.3 We assume that the Spacecraft will remain in a Uranus Orbit for a maximum period of 7 years. To determine the temperature variation for both the Antenna-dish and the Spacecraft the same equations are applied as for the Earth Orbits. Figure 35 demonstrates a small temperature decrease of the Antenna-dish from 173.2K to 167.96K. Similarly, the temperature of the Spacecraft decreases from 283K to 274K.
3 In an Orbit where the Perigee is 4RU and the Apogee is 28RU, the final temperature for the Antenna dish is 173.3K and for the Spacecraft is 274.2K.
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Temperature of the Antenna-Dish
420 370 320
270 220 170 120 Temperature in in Kelvin Temperature 0.00E+00 5.00E+07 1.00E+08 1.50E+08 2.00E+08 Time in seconds
Cold Limit Hot Limit Analysis
Figure 35 - Transient Thermal Analysis for Antenna-dish during the Uranus Orbit. Temperature decreases from 173.2K to 167.96K. Temperature of Spacecraft
320
310
300
290
280
Temperature Temperature in Kelvin 270 0.00E+00 5.00E+07 1.00E+08 1.50E+08 2.00E+08 Time in seconds
Cold Limit Hot Limit Analysis
Figure 36 - Transient Thermal Analysis for Spacecraft during the Uranus Orbit. Temperature decreases from 283K to 274K.
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13.5 ALTERNATIVE THERMAL POWER REQUIREMENTS The previously mentioned Spacecraft box has a side length of 2m. Alternatively, it could be that the side length of the spacecraft is considered to be 1m, which would lead to the same or similar results as previously shown if the power dissipated is adapted. With the application of 60Wt the spacecraft temperature after 17 years could be at 274K and the temperature of the dish could be 283K. These are the same results as for the 2m side length of the spacecraft.
Figure 37 - Alternative Antenna Dish and Spacecraft Configuration
13.6 THERMAL SUBSYSTEM SUMMARY The thermal control for the Spacecraft and the Antenna dish are an important concern for this particular mission. Due to the strong thermal environmental variations between Earth and Uranus, the correct selection of surface materials and the power required to maintain the spacecraft’s temperature are essential. To simplify the problem, the antenna-dish is assumed to be continuously applied as a sunshield, leaving the remainder of the spacecraft in the shadow (Dougherty, et al., 2011).
This analysis demonstrates that the three stages of the mission are achievable for both the Antenna-dish and the Spacecraft, while remaining within their temperature limits. This was achieved without adapting the size and the mass of both the Antenna (40Kg) dish and the Spacecraft (between 1200Kg [dry mass]-3000Kg [wet mass]). In the entire calculation it has been assumed that the Antenna-dish is painted white on both sides. Its emissivity is 0.853 and its absorptivity is 0.252 (White Enamel) (Larson & Wertz, 2005). Meanwhile, the spacecraft is made out of Aluminium and Kapton, with emissivity of 0.03 and absorptivity of 0.12 (Gilmore, 2002). To maintain the internal spacecraft’s temperature above 273K (over a 17 years time period), 195WTh are required, which are provided by RTHU-241-Am. The RTHU-241-Am have been selected to replace and to minimise the use of electrical heaters. Within the mass and power budget a 20% margin has been included on top of the spacecraft’s thermal mass subsystem requirements. This could lead to up to 234WTh being supplied to the spacecraft. This could result in a final spacecraft temperature of 287K at the end of the mission lifetime.
The RTHU-241-Am achieve a thermal power output of up to 25Wt/Kg, therefore 7.8Kg (9.36Kg including margin) are required to ensure that the spacecraft’s temperature remains within its limits. RTHU-241-Am would be ideal for the mission as they provide both thermal spot heating and electrical power at once, while combining the functionality of typical Radioisotope Heater Units and RPS (Williams, et al., 2012). Up to 9We power could be obtained by these 241 RTHU-241-Am ( Am2O3) which could support the spacecraft’s power requirements (Williams, et al., 2012). In the 241 case that RTHU-241-Am ( Am2O3) are unavailable, alternative RHU may be applied. Therefore the electrical power produced from these RTHU are not included in the Power subsystem section. Previous missions, such as Cassini or New Horizons, made use of RTG’s waste heat to contribute towards the thermal control of a spacecraft. This will be avoided in this mission to Uranus mainly because the “high surface temperature and thermal power of these units is difficult to distribute without causing excessive local heating” (Williams, et al., 2012).
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Apart from the RTHU-Am-241 (7.8Kg), six louvers (1Kg each) will be used to get rid of excess heat and keep the spacecraft at a temperature between 273K and 313K. Ideally, the Spacecraft will use the ‘thermal bottle’ approach which aims to minimise thermal loss by using the RTHU- 241-Am thermal power and the electronics waste heat to maintain the spacecraft in an operational temperature for its components (Schaefer, et al., 2008).
It is estimated that the multilayer isolation required for the Spacecraft has a mass of up to 15Kg. Furthermore two heat pipes are used, which require a total mass of 10Kg (Hubbard, 2010). While the thermal bottle approach aims to minimise thermal loss, the heat pipes are applied to distribute the heat throughout the spacecraft (Brown, 2009). Due to the application of multiple RTHU-Am-241, heaters might no longer be required, which could save up to 7.8Kg. Furthermore, as a precaution it should be assumed that some components of the instruments selected for this missions require additional thermal control as they are located on the exterior (Magnetometer boom) of the spacecraft, or need to be cooled (Infrared detector). The power required for the instrument's thermal control is assumed to consume up to 30W which are provided in the worst case by heaters, thermostats or radiators with a mass of up to 3Kg. These 30Wt are ideally provided by RTHU-241-Am which require a mass of up to 1.2Kg. Furthermore it could be assumed that the thermal control of the spacecraft instruments is included in the margins of the individual instruments (between 44% and 80%). (Gilmore, 2002; Larson & Wertz, 2005; Dougherty, et al., 2011; Hubbard, 2010; Williams, et al., 2012; Kusnierkiewicza, et al., 2005).
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14 ARCHITECUTURE To illustrate the approximate form of the spacecraft and general placement of many of the most important elements a simple CAD model was built using Siemen’s NX. Most of the elements are represented purely by encompassing box volumes in order to provide an indication of size and placement. Figure 38 shows this model, labelled to clarify the positions of the modelled systems. Note that only one of the ASRG power units is labelled, with the other three in corresponding corners of the spacecraft, including one hidden by the propellant tanks.
Some elements were not included in this model, notably the attitude control system, which would principally consist of two small spherical tanks placed close to the centre of mass, as well as thruster elements at the base of the spacecraft. Also, the only part of the communications system shown is the high gain antenna, and no power cables of fuel pipes are included. These elements were omitted due to the limited use in modelling such small elements compared to the time required to do so, and to provide a clearer image.
Figure 38 – CAD model representing basic architecture
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