ARGUS: IO EXPLORER MISSION
University of Leicester Department of Physics & Astronomy
Doc No: SEEDS-RSH-2015-0909 Issue: 1 Rev: 3 Date 21 September 2015
Name Signature Date
Lakshay Aggarwal Prepared by Lakshay, Azhen 21September 2015 Azhen Jarjes
Checked by Dr. Nigel Bannister
Approved by Dr. Richard Ambrosi
ARGUS: Io Explorer Mission i
DISTRIBUTION
Name Organisation Azhen Mziry University of Leicester Lakshay Aggarwal University of Leicester
ARGUS: Io Explorer Mission i
CHANGE LOG
Date Issue Revision Section Reason for change 3 July 2015 1 1 PDR File New document 21 August 2015 1 2 CDR File Revised Document with RIDS 21 September 1 3 Final Camera Document revised (Order, Paraphrasing, CDR 2015 Ready Copy presentation)
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TABLE OF CONTENTS
1. INTRODUCTION ...... 1 1.1 SCOPE ...... 1 1.2 APPLICABLE DOCUMENTS ...... 1 1.3 REFERENCE DOCUMENTS ...... 1 1.4 WORK BREAKDOWN ...... 3 2. IO – THE MOST ACTIVE BODY IN SOLAR SYSTEM ...... 4 2.1 OVERVIEW: ...... 4 2.2 THE INTERIOR OF IO: ...... 4 2.3 THE SURFACE OF IO: ...... 5 2.4 THE ATMOSPHERE OF IO: ...... 8 3. FINAL DESIGN REVIEW APPROACH ...... 10 3.1 SCIENCE MATRIX: ...... 10 3.2 REQUIREMENT MATRIX: ...... 11 4. MISSION DESIGN: ...... 16 4.1 REFERENCE MISSION ...... 16 ANALYSIS: ...... 17 4.2 SCIENCE ORBIT: ...... 18 4.2.1 Implementing in SPENVIS: ...... 18 4.2.2 Preliminary Study: ...... 20 4.3 SUMMARY: ...... 23 5. LAUNCH VEHICLE: ...... 24 5.1 PRELIMINARY STUDY ...... 24 5.2 ATLAS V 541: ...... 25 6. PROPULSION: ...... 27 6.1 REQUIREMENTS: ...... 27 6.2 GENERAL OPTIONS: ...... 28 6.3 ANALYSIS: ...... 28 7. COMMAND AND DATA HANDLING ...... 33 7.1 SCIENCE PAYLOAD DATA CALCULATION ...... 33 7.2 DATA GENERATION & DUTY CYCLES ...... 35 7.3 COMPRESSION TECHNIQUES ...... 38 7.4 DIGITAL MODULATION TECHNIQUE ...... 39 8. COMMUNICATIONS ...... 42 8.1 RF (RADIO FREQUENCY) MODE SELECTION ...... 43 8.2 GROUND STATION AND DOWNLINK ...... 43 8.3 DATA CONSTRAINTS ...... 46 8.4 LINK BUDGET ...... 47 9. RADIATION ENVIRONMENT AND MAGNETIC FIELD: ...... 53 9.1 OVERVIEW: ...... 53 9.2 THE VAN ALLEN BELTS: ...... 53 9.3 JUPITER’S RADIATION ENVIRONMENT: ...... 53 9.3.1 Inner magnetosphere: ...... 54 ARGUS: Io Explorer mission iii
9.3.2 Middle magnetosphere: ...... 54 9.3.3 Exploring the magnetic field of Jupiter: ...... 54 9.4 ANALYTICAL METHODS: ...... 56 9.4.1 Energetic Particle Radiation: ...... 56 9.4.2 Trapped Particles: ...... 56 9.4.3 Worst Case Fluence: ...... 57 10. TOTAL IONISING DOSE AND SHIELDING: ...... 59 11. WEIGHT ...... 60 11.1 INSTRUMENT SPECIFICATIONS ...... 60 11.2 SYSTEM WEIGHT DATA COLLECTION ...... 62 12. POWER SOURCE ...... 63 13. IMPACT OF OTHER MISSIONS ON ARGUS ORBITER ...... 68
LIST OF TABLES
Table 1: Contribution ...... 3 Table 2: Mountain types ...... 7 Table 3: Science Objectives ...... 10 Table 4: Requirement Matrix ...... 12 Table 5:Atlas V 541 Specifications ...... 26 Table 6:Orbiter ΔV Budget ...... 27 Table 7:The properties of R-42DM ...... 29 Table 8:the properties of AMBRTM (Aeroj)...... 32 Table 9: Frames ESA Standards ...... 33 Table 10: Io’s Case ...... 35 Table 11: Argus’s Case ...... 37 Table 12 Compression ...... 41 Table 13 Duty Cycle and overall Data ...... 46 Table 14: Link Budjet analysis ...... 48 Table 15: weight criteria ...... 60 Table 16 Overall Mass ...... 62 Table 17 Power System ...... 66 Table 18 Mission Comparison ...... 68 ARGUS: Io Explorer mission iv
TABLE OF FIGURES Figure 1: represents enhanced pictures of IO 4 Figure 2: Layers of Io’s interior (NASA.gov). 5 Figure 3: represents Loki Patera Figure 4 7 Figure 5: Io’s upper atmosphere, with different colours indicating the different element emissions (Wikipedia). 9 Figure 6: One year local optimal mission database graph (trajectory browser) 16 Figure 7: Database Reference Trajectory 17 Figure 8:The integral electron and proton flux for three Jovian models with energy less than 1MeV. 20 Figure 9:The radiation environment for different orbits and inclinations 21 Figure 10: Worst case integral fluence for different perijove angles in relation to Jupiter’s equator. 22 Figure 11:Characteristic energy graphs (Atlas V) 24 Figure 12: Characteristic energy graphs (Arian 5) 25 Figure 13:Thruster R-42DM (Aerojet) 30 Figure 14:HIPAT TM thruster (Aerojet Rocketdyne) 32 Figure 15: represents Io’s position at different times during 1 complete orbit 36 Figure 16: represents Argus’s position at different times during 1 complete orbit 38 Figure 17: represents schematic view of coder 39 Figure 18: represents Eb/N0 vs. BER 40 Figure 19: represents Data Flow 42 Figure 20: represents NASA's DSN 44 Figure 21: represents ESA's ESTRACK 45 Figure 22: represents Key Link Parameters 47 Figure 23: represents Transmitted Power vs. Link for 70 m and 34 m (X band) 50 Figure 24: represents Transmitted Power vs. Link for 34 m Ka 50 Figure 25:represents Transmitted Power vs. Link for 34 m Ka 51 Figure 26:Predicted average trapped electron and proton flux for a complete ARGUS mission 57 Figure 27:The average integral flux above 1 MeV for the ARGUS mission 58
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Figure 28: Dose in silicon target as a function of spherical aluminium shielding 59 Figure 29:represents RTG heritage 65
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1. INTRODUCTION
1.1 SCOPE
This report is Final Review Document for MSc. Space Exploration Systems. This report is a critical phase A study which covers major topics towards designing a space system for exploration of Io.
1.2 APPLICABLE DOCUMENTS
The following are applicable documents
1. Link Budget Excel File (Link Analysis for Antenna Sizing and Power Constraints)
2. True Anomaly Positioning Excel File (Positioning of IO and Argus at a given time)
1.3 REFERENCE DOCUMENTS
The following are reference documents:
A COMPARISON OF DIGITAL MODULATION METHODS FOR SMALL SATELLITE DATA LINKS, Daniel J. Mulally, Vice President, and Don K. Lefevre, President, Cynetics Corporation Jet Dr.Rapid City, SD 57709
Bagenal F, Dowling TE, McKinnon WB. Jupiter: the planet, satellites, and magnetosphere. Cambridge: Cambridge University Press; 2004.
BER performance of OFDM-BPSK,-QPSK,- QAM over AWGN channel using forward Error correcting code, Vineet Sharma, Anuraj Shrivastav, Anjana Jain, Alok Panday, International Journal of Engineering Research and Applications (IJERA) ISSN: 2248, May-Jun 2012
BER Vs Eb/N0 for QPSK modulation over AWGN, Gaussian Waves,Digital Modulation, Mathuranthan, October 2010
Connerney, J. M. Acuna , N. Ness and T (1998). Satoh New models of Jupiter's magnetic field constrained by the Io flux tube footprint, J. Geophys. Res., 103, pp. 11929-11940.
De Angelis, G., Clowdsley, M. S., Nealy, J. E., Tripathi, R. K., & Wilson, J. W. (2004). Radiation analysis for manned missions to the Jupiter system.Advances in Space Research, 34(6), 1395-1403.
Divine, N., & Garrett, H. B. (1983). Charged particle distributions in Jupiter's magnetosphere. Journal of Geophysical Research: Space Physics (1978–2012), 88(A9), 6889-6903.
Dr. Les Deutsch. "NASA’s Deep Space Network: Big Antennas with a Big Job", July 2009.
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G.Schubert, J. D. Anderson, T. Spohn, W. B. Mckinnon 2007. The interior of Io in Jupiter: the planet satellites and magnetosphere (pp. 286-291). Cambridge university press.
Heynderickx D, Quaghebeur B, Wera J, Daly EJ, Evans HDR. New radiation environment and effects models in ESA's space environment information system (SPENVIS). IEEE; 2003.
J.L. Barth, C.S. Dyer, and E.G. Stassinopoulos (2003) “Space, Atmospheric, and Terrestrial Radiation Environments,” IEEE Trans. Nucl. Sci., vol. 50, no. 3, pp. 466-482.
Juno Tele Communications, Ryan Mukai, David Henson, Anthony Mittskus, Jim Taylor, Monika Danos, October 2012, Design and Performance Series, JPL, NASA
Khurana, K. K., Kivelson, M. G., Vasyliunas, V. M., Krupp, N., Woch, J., Lagg, A., & Kurth, W. S. (2004). The configuration of Jupiter’s magnetosphere.Jupiter: The planet, satellites and magnetosphere, 1, 593-616.
Lellouch, E. (2005). Io’s atmosphere and surface-atmosphere interactions.Space science reviews, 116(1-2), 211-224.
Moore, W. B., Schubert, G., Anderson, J. D., & Spencer, J. R. (2007). The interior of Io. In Io After Galileo (pp. 89-108). Springer Berlin Heidelberg.
NASA Trajectory Browser: http://trajbrowser.arc.nasa.gov/ (accessed June-July 2014)
Planetary Science Decadal Survey Io Observer, Mission Concept Study, Elizabeth Turtle, NASA HQ POC: Curt Niebur, May 2010
Schneider, N. M., & Bagenal, F. (2007). Io’s neutral clouds, plasma torus, magnetospheric interaction. In Io After Galileo (pp. 265-286). Springer Berlin Heidelberg.
SPENVIS – Space Environment Information System: http://www.spenvis.oma.be/spenvis/ (Accessed june- sep. 2015)
TEC-EES & SRE-PAP (2013). “JUICE Environmental Specification”. Issue 5, Revision 0
United Launch Alliance March2010. “Atlas V Launch Services User’s Guide”, Revision 11
Wertz. J and Larson. W, Space Mission Analysis and Design, 3rd ed. 2003
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1.4 WORK BREAKDOWN
Table below outlines the contribution of each member to the executive summary by highlighting the sections which each member completed. Table 1: Contribution
Chapter Author
Io, The most active Body in Solar System The Icy Giant Section 2,1: Lakshay Aggarwal
Section 2.2/ Azhen Jarjes
Section 2.3/ Lakshay Aggarwal
Section 2.4/ Azhen Jarjes
Final Design Approach (Science Matrix, Requirement Lakshay Aggarwal Matrix)
Mission Design Azhen Jarjes
Launch Vehicle Azhen Jarjes
Propulsion Azhen Jarjes
Power Source (Estimation and Study), Impact of RTG’s Lakshay Aggarwal
Command and Data Handling Lakshay Aggarwal
Communication Lakshay Aggarwal
Radiation Environment and Magnetic field Azhen Jarjes
TID and Shielding Azhen Jarjes
Weight Lakshay Aggarwal
Impact of other missions on Argus Orbiter Lakshay Aggarwal
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2. IO – THE MOST ACTIVE BODY IN SOLAR SYSTEM
2.1 OVERVIEW:
Io, one of the major Jovian moon is about the same size and density as Earth's Moon. Io is the most volcanically active body known in the Solar System meaning, volcanic eruptions on Io are so common and massive that the entire surface could be buried under 100 meters of material every 1 million years. An important common feature i.e. impact craters, are not present on Io because of the frequent volcanic eruptions on the surface. For example, some areas on Io appear to be hot red and points towards recent explosive eruptions and volcanic plumes. Figure below represents the most prominent red oval surrounds the volcano Pele.
Figure 1: represents enhanced pictures of IO
2.2 THE INTERIOR OF IO:
Io is primarily composed of silicate rock and iron, which was discovered from gravity data collected by previous missions. Io is the closest planet to Earth in term of composition than the other planets in the solar system. The density of Io is very high, to the extent that it is the densest moon in our solar system, and is about 3,527.5 kg /cm3. However, the measurements obtained from the Voyager and Galileo space probes resulted in different findings of Io’s mass, radius, particularly regarding the composition of the interior of the moon. Io’s core is considered to be about 20% of its entire
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mass, and the radius of the core depends on the composition and its amount. If the core consists of iron only then the radius is between 350 and 650 km, whereas if it contains a mix of iron and sulphur then the radius is between 90 and 550 km (Schubert et al., 2004). The core is surrounded by a silicate rock shell that extends to the surface (Moor et al., 2007).
Figure 2: Layers of Io’s interior (NASA.gov).
2.3 THE SURFACE OF IO:
The surface of Io has three common features:
Mountains A similar feature to earth i.e. mountains of Io are uneven and isolated. They are separated by the plains. Together the mountains and plains cover about 2% of the Io’s surface. Some of the known mountains on Io are even 100 km long with a relief of about 9 km. Although the mountains are mantled in sulphur, the mountains do not appear to be of volcanic origin and are thought be older
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than both the plains and volcanoes. Regular erosional processes on the surface has modified the mountains too. This mountainous presence on Io indicates towards a rigid lithosphere, possibly up to 30 km thick.
Plains About 40% of Io is covered by plains with low relief and light and dark areas. The plains are probably layers of pyroclastic material erupted from volcanoes and possibly lava flows of different compositions or ages. Other plains contain plateaus with smooth tops and escarpments from 150 to 1700 m high which points towards the presence of erosion process.
Volcanoes Only about 5 percent of Io is covered by volcanic vents. Although there were 500-700 identified volcanic centres but, the energy (enormous rate) has been released at only about four centres in the last decade. Carr (1997) reported that 356 calderas had been identified in the Voyager and Galileo coverage. The most common type of vent on Io is Paterae (low-profile volcanic shields). Their flows can cover large areas and reach lengths of 700 km. These long production ranges indicates towards high eruption rates and/or low viscosity material. Figure 3 below is the comparison of 1979 Voyager 1 image of Loki Patera with Galileo images taken in 1996. The patera is at the centre of the images. A dark fissure is just above and right of the patera. Voyager observed an eruption from this fissure in 1979.
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Figure 3: represents Loki Patera Figure 4
Explosive eruptions have been observed on Io and there is indirect evidence for effusive eruptions. Galileo images below (Figure 4) showing two volcanic plumes on Io.
Two types of eruption plumes have been observed: Prometheus-type and Pele-type.
Table 2: Mountain types Eruptions Prometheus-type Pele-type Plume heights 50-120 km up to 300 km Plume character optically optically thick, dark jets thin (transparent) Deposits bright halos, 200-600 km diameter dark halos, 1000-1500 km in diameter Eruption Velocities about 500 m/s up to 1000 m/s Duration months to years days to months Location common near equator restricted longitudes Temp. about 450 K about 600 K of associated ‘hot spots’
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Earth-based monitoring of thermal emissions on Io have been interpreted as eruptions of surface lava flows.
2.4 THE ATMOSPHERE OF IO:
The last 15 years has witnessed a dramatic improvement in our knowledge of Io’s atmosphere, with numerous data obtained from UV, IR wavelength, and millimetres, and there is a wealth of new advanced models describing the chemical composition, vertical and horizontal structure, and plasma interaction, along with temperature mapping (Lellouch, 2005). Io’s atmosphere mainly contains sulphur dioxide (SO2) and other minor composites such as sulphur (SO), sodium chloride (NaCl), oxygen (O), and Sulphur (S). The density of Io’s atmosphere is not constant, and it changes by time, location, and volcanic activity with a range of 3.3×10-5 to 3×10−4 Pascal. In addition, peak pressures have been observed at the volcanic plumes with a range of 5×10−4 to 40×10−4 Pa (Lellouch, 1992). Furthermore, Io’s pressure is lower in the night side than the dayside. Volcanisms are the essential source of Io’s atmosphere, which carries special properties regarding the moon’s composition, structure, and maintenance. The volcanic plumes also represent the main source of SO2, since they 4 eject about 10 kg of SO2 every second into Io’s atmosphere, although most of the SO2 returns to the moon’s surface (Moullet et al., 2010). Jupiter’s magnetosphere releases gases from its atmosphere, which escape to either the cloud that surrounds Io or Io’s plasma torus, which is “a ring of ionized particles that shares Io’s orbit” and turnover the magnetosphere of Jupiter. The removal material from the atmosphere is about a tonne for every second (Schneider and Bagena, 2007).
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Figure 5: Io’s upper atmosphere, with different colours indicating the different element emissions (Wikipedia).
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3. FINAL DESIGN REVIEW APPROACH
The Jovian moon Io, has the most active volcanic surface in the solar system where large scale geological processes can be studied in detail which would be helpful to provide an insight about different planetary evolution processes. The orbital resonance of Io is 2:1, with the next moon Europa and an orbital resonance of 4:1 with Ganymede. Due to this resonance setting, Io maintains an orbital eccentricity of 0.0041 which results in a massive tidal heating in its interior. In addition to these, Jupiter’s magnetosphere is strongly coupled with Io which results in high radiation levels around Io. Also, leading a significant mass flux from the moon’s atmosphere.
3.1 SCIENCE MATRIX:
Table 3: Science Objectives Science Objective # Science Investigation
Primary Objectives Determine the IA1 Determine the bulk elemental abundances present in the atmospheric structure atmosphere of Io. and composition of Io IA2 Mapping 3-D velocity distributions of electrons and individual ion species IS1 Determine the bulk elemental abundances present in the interior of Io. Determine the IS2 Determine Io’s interior structure, e.g., whether it has a magma surface, internal ocean. structure and IS3 Determine the magnitude, spatial distribution, temporal composition of Io variability, and dissipation mechanisms of Io’s tidal heating. IS4 Understand the processes that form mountains and Paterae on Io.
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Determine the active IV1 Eruption mechanisms for Io’s lavas and plumes and their volcanic processes of implications for volcanic processes on Earth (similar heat Io. flow) Secondary Objectives Determine the IM1 Precision Mapping of the magnetic Field (if present) and its magnetic fields of Io variability Understand how Io IJ1 Understand Io’s surface chemistry, atmosphere, and affects the Jovian ionosphere, the dominant mechanisms of mass loss. system and IJ2 contributions to the Understand Plasma Torus torus.
3.2 REQUIREMENT MATRIX:
Represented below is the proposed list of instruments along with their science contribution, characteristics and their heritage.
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Table 4: Requirement Matrix Model Instrument Science Contribution Characteristics
Pixels Per Image – 1) Eruption mechanisms for Io’s lavas and 125000 plumes and their implications for volcanic Samples Per Pixel – 1 processes on Earth (similar heat flow) Bits Per Sample - 8 >3 band passes (2- Thermal Mapper 2) Determine the magnitude, spatial distribution, 20µm) (Mars Odyssey temporal variability, and dissipation mechanisms Themis) of Io’s tidal heating. 5 visual bands: 0.425, 0.540, 0.654, 0.749, 3) Understand Io’s surface chemistry, 0.860 microns atmosphere, and ionosphere, the dominant mechanisms of mass loss. 10 infrared bands: 6.78 (used twice), 7.93, 4) Understand Plasma Torus 8.56, 9.35 ,10.21, 11.04, 11.79, 12.57, 14.88 microns
1) Determine the bulk elemental abundances present in the atmosphere of Io Pixel per Image – 3072 Samples per pixels – 1 Ultraviolet 2) Eruption mechanisms for Io’s lavas and Bit per sample – 12 Spectrometer plumes and their implications for volcanic (JUICE) processes on Earth (similar heat flow) Wavelength range 55- 210 nm with spectral
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3) Understand Io’s surface chemistry, resolution of <0.6 nm atmosphere, and ionosphere, the dominant mechanisms of mass loss.
1) Determine Io’s interior structure, e.g., whether Pixel per Image – it has a magma ocean, and implications for the 1048576 coupled orbital-thermal evolution of Io and Samples per pixels – 1 Europa Bit per sample – 12 Narrow Angle Colour Filter – 12 Camera & Wide 2) Understand the processes that form mountains FOV ~ 1 Angle Camera and Paterae on Io <10m/pix at 500 Km (IVO, New <5Km/pix at 105 Km Horizons) 3) Eruption mechanisms for Io’s lavas and plumes and their implications for volcanic processes on Earth (similar heat flow) Pixel per Image – 4194304 4) Understand Io’s surface chemistry, Samples per pixels – 1 atmosphere, and ionosphere, the dominant Bit per sample – 14 mechanisms of mass loss (WAC) Colour Filter – 14 FOV ~ 10.5
Gamma Ray 1) Determine the bulk elemental abundances Estimated Data rate – Spectrometer present in the interior of Io. 1.75E+02 bps (Messenger)
Dual Technique 1) Precision Mapping of the magnetic Field (if Number of Channels - 3 Magnetometer present) and its variability (Cassini)
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1) Mapping 3-D velocity distributions of electrons and individual ion species Thermal Electron and Plasma Particle ion detection ~ 3eV – Detector (JUICE) 2) Understand Io’s surface chemistry, 20keV atmosphere, and ionosphere, the dominant mechanisms of mass loss. 1) Understand Io’s surface chemistry, Neutral Mass atmosphere, and ionosphere, the dominant Samples per pixels – 1 Spectrometer mechanisms of mass loss. Number of Channels – (Rosetta, IVO) 1 2) Understand Plasma Torus FOV ~ 8
A study was done in order to identify the major features which would be required by the proposed set of instruments in order to perform accurate science when in orbit. Proposed set of science instruments above would be able to perform:
1) Repeated< 100 m resolution multicolour imaging of wide areas
2) Topographic mapping (laser or stereo), 2 m relative precision
3) Spectroscopy with < 1 km spatial resolution would provide compositional constraints on fresh lavas, temperature information i.e. composition and eruption style
4) 10, 20 micron thermal mapping, 10 km resolution (Measures heat flow, total lava output)
5) <0.6 micron UV spectroscopy, 20 km resolution, for detailed spatial mapping of atmosphere and plumes
6) Plasma instruments capable of mapping 3-D velocity distributions of electrons and individual ion species
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7) Neutral mass Spectrometers for close perigee flybys
8) GRS system with a broad spectrum in order to differentiate between wide ranges of elements
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4. MISSION DESIGN:
4.1 REFERENCE MISSION
ARGUS spacecraft will use Earth’s gravity in order to travel through space, and will arrive at Jupiter approximately five years after launch. The spacecraft will conduct more than 30 flybys of Io. The assistance from Earth’s gravity decreases the ΔV required and the required fuel to arrive at Jupiter.
In order to study the feasibility of a flyby exploration of Io, the available trajectories for a one-way mission to Jupiter was funded using the Mission Design Centre Trajectory Browser. This database indicates that one favourable launch window will be available in 2027, which will provide the extra boost needed to reach Jupiter. The trajectory uses Earth’s gravity after one deep space manoeuver. The one year optimum trajectories are shown in Figure 6. The colours in this picture specify the requirement ΔV of the trajectory.
Figure 6: One year local optimal mission database graph (trajectory browser)
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The most favourable trajectory is shown in Figure 7. The duration of the route is approximately five years, and the mission will commence in 2027 and arrive at Jupiter at the end of 2031.
Figure 7: Database Reference Trajectory
The trajectory launch commences from the lower Earth orbit at an altitude of 200 km, with an energy launch equal to 24 km2/s2. The required velocity to change the orbit from circular to hyperbolic is equal to 4.3 km/s. After one year travelling in a deep space manoeuver, there will be a change in the spacecraft’s velocity of about 513 m/s, which is required to bring the spacecraft into earth and to use Earth’s gravity assist to boost the spacecraft towards Jupiter.
ANALYSIS:
The drawback of choosing this trajectory is the low altitude of the arrival orbit (200 km), which is inappropriate for a number of reasons. Firstly, Jupiter has a strong magnetic field and radiation environment, which is known to be stronger at Jupiter’s lower altitudes than the higher altitudes (Bagenal et al., 2004). Secondly, the minimum perijove radius needed for the ARGUS mission is 5 RJ, which is the Io orbiter radius. For this reason, some changes have been made to the last part of
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the trajectory, in order to alter the circular orbit from an altitude of 200 km to an elliptical orbit with a perijove of 5 RJ and different apojove distance; this will be discussed in more detail in Section 4.2.
In order to find the requirement propulsion for the orbit, some calculations have been made working backwards from the trajectory browser database value. The arrival velocity of the orbit can be calculated by adding the velocity of the circular orbit to the arrival ΔV given from the database (223 m/s). This gives a value of 60.33 km/s, which can be used to determine easily the energy C3 of approximately 27 km2/s2. Then, the periapsis velocity of any altitude can be calculated, which allows for a determination of the required ΔV by subtracting this with the desired velocity of the final orbit periapsis.
4.2 SCIENCE ORBIT:
The strength of Jupiter’s radiation environment means that Jovian radiation has to be considered when choosing science orbits. Therefore, the Space Environment Information System (SPENVIS) has been used in order to find the orbit with the least radiation for ARGUS, which will result in better data.
4.2.1 Implementing in SPENVIS:
SPENVIS is the abbreviation of the European Space Agency’s (ESA) Space Environment Information System, which has information on radiation environments. The models in the SPENVIS have been prepared to determine radiation sources and effects, the total ionising dose (TID), and several other packages discussed by D. Heynderickx et al. (2003). The radiation sources that are provided in the SPENVIS are trapped protons and electrons, solar energy protons, and cosmic rays and solar energetic ions (Heynderickx et al., 2003).
Jovian Modelling:
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To test a spacecraft’s reliability in the Jovian magnetosphere, the trapped particle models in the SPENVIS are used. There are several trapped particle models for Jupiter available in the SPENVIS that are organised by data provided in situ by orbiting and flyby missions (Connerney et al., 1998).
1- Divine and Garrett model (D&G83):
This model is the oldest model, and its measurements were collected by the Pioneer and Voyager Jupiter flybys in combination with based observations. The coordinate electron range in this model is from the surface of the planet to the tail of magnetosphere, whereas the coordinate proton range is from the surface to 12 RJ. The energy range is higher than 0.6 and 0.06MeV (Divine & Garrett, 2012).
2- The GIRE model:
The Galileo Interim Radiation Electron model (GIRE) is the successful model from the Galileo mission. The coordinate range is 8-16 RJ, with an energy range of 0.5-30MeV for electrons (see Figure 8) (Divine & Garrett, 2012).
3- The Salammbo model:
The Salammbo model is a theoretical adaptation for Jupiter, developed from a Salammbo model used for Earth by ONERA/DESP in 1998. It is applied to both electrons and protons of energies from 1 to 600 MeV and 1MeV to 1GeV, respectively (see Figure 8), for a limited radiation distance from the surface to 9.5 RJ for electrons and 6 RJ for protons (inside Europa orbit) (Heynderickx et al., 2003).
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Average Integral Flux of Trapped electron and proton
for different Jovian Models 1.00E+08 Proton_D&G87 1.00E+07 Electron_D&G87 1.00E+06 /MeV/S)
2 Salammbo_Pro 1.00E+05 30 cm 1.00E+04 Salammbo-ELE 1.00E+03 600 GIRE-ELE 1.00E+02 MeV (/ MeV
1 1.00E+01 > 1.00E+00 2000 1.00E-01 1.00E-02 1.00E-03 1000
Integral Flux Flux Integral 1.00E-04 2000 1.00E-05 0 500 1000 1500 2000 Energy (MeV)
Figure 8:The integral electron and proton flux for three Jovian models with energy less than 1MeV.
4.2.2 Preliminary Study:
4.2.2.1 Different apojove distance and inclination:
A preliminary study of the science orbit was undertaken based around the radiation environment of Jupiter. The detailed explanation of the study is discussed in the appendix. The first estimation for the science orbit was orbiting Io in order to obtain the best data. However, due to the high radiation environment between Io and Jupiter, it was found that this orbit is not appropriate because the spacecraft will require strong shields that will increase its mass. Therefore, the Io flyby was selected, as it is an elliptical orbit with a perijove radius equal to the Io radius from Jupiter. As such, during this part of the orbit the spacecraft will face Io. Jupiter’s radiation models obtained by the SPENVIS were used to determine the best apojove and inclination in terms of radiation, and models used for this were Salammbo+ D&G87. Figure 9 shows the different apojove and inclination options and also the scenarios of an Io orbit or flyby.
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compare between the average electron flux in orbit of Io and a flyby elliptical orbit of Io. and using Ganymede, Callisto radial distance for apojove, and different inclination.
MeV MeV 1.00E+18 1 1.00E+17 1.00E+16 1.00E+15 5Rj
) 2 1.00E+14 15Rj_0 1.00E+13 15Rj_45 (#/cm 1.00E+12 26Rj_0 1.00E+11 1.00E+10 26Rj_45 1.00E+09 1.00E+08
Total mission electron fluence> electron mission Total 0 5 10 15 20 25 30 35 Energy MeV
Figure 9:The radiation environment for different orbits and inclinations
The 5RJ (Io orbiter radius), indicated by the darker blue line in Figure 9, indicates the environment of the spacecraft when orbiting Io. The four other lines are the orbits that flyby Io with different Jupiter radii for apojove and different inclination. The total mission electron flux is more than 1013 #/cm/s for 5 RJ. However, for the Io flyby the radiation environment is 10 times less when choosing an apojove equal to the Ganymede orbit (15 RJ) and 40 times less when choosing an apojove radius similar to that of the Callisto orbit (26 RJ), and thus this orbit was chosen for the ARGUS science orbit. In addition, Figure 9 also shows the two flyby orbits with different inclination degrees (0 and 45), and it is clear from the curve that orbiting Io at a 45 degree inclination with Jupiter declines the environment radiation by a factor of 2. To sum up, the science orbit for ARGUS is elliptical, with a perijove distance similar to the Io orbit distance from Jupiter and an apojove distance equal to the Callisto orbit distance (26 RJ), with a 45 degree inclination with Jupiter.
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4.2.2.2 Perijove argument (w):
Other studies have endeavoured to find the best science orbit and environment with the least radiation. Changing the perijove argument angle from an equatorial to polar by 15 degree each time gives us different results, and as such SPENVIS was used to compare the total mission electronic fluence for a different perijove angle (w) (see Fig. 10). Because Io is approximately in the equatorial range of Jupiter, therefore the meeting point of Io and ARGUS is different for each orbit. For example, when the perijove is in the equatorial region or has an inclination of 15 degrees, Io and ARGUS will meet in the perijove of the orbit. However, for orbits with different perijove angle inclinations (e.g., 45, 60, 75, and 90), the meeting point will be in the semi-parameter.
Integral fluence for different perijove inclination for worst (12,42,100) hrs 3.5E+13
) 2 3E+13 (/cm 2.5E+13 MeV MeV 1 > 2E+13 12 hr Fluence 1.5E+13
Fluence 42 hr Fluence 1E+13 100 hr Fluence ntegral ntegral i 5E+12
0 0 20 40 60 80 100 Different perijove inclination
Figure 10: Worst case integral fluence for different perijove angles in relation to Jupiter’s equator.
Figure 10 illustrates the worst hours for the ARGUS mission for different perijove arguments (12, 24, and 100 hours). It is clear from the figure that a 15-degree inclination orbit provides the lowest radiation environment, followed by zero degree. Conversely, the orbits with more than a 15-degree
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inclination (w) are in a very high radiation environment in comparison with 0-15 degrees. Consequently, the most appropriate orbit in terms of radiation is a 15-degree inclination, but because Io is in the equatorial region of Jupiter, then the orbit with w=0 has been chosen for the ARGUS mission.
4.3 SUMMARY:
ARGUS will orbit Jupiter equatorially with a perijove equal to the orbital radius of Io to Jupiter, an apojove of 26 RJ (the orbital radial distance of Callisto), and with a 45-degree inclination. ARGUS will do 33 flybys of Io over the course of two years.
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5. LAUNCH VEHICLE:
5.1 PRELIMINARY STUDY
In order to select the proper launch vehicle for the mission to Io, it is crucial to know the mass of the spacecraft; this was first calculated as 3355 kg. According to plots of mass vs. C3, most of the considered launch vehicles are capable of carrying the required mass at the desired C3. However, according to our chosen trajectory obtained from the browser, the C3 is equal to 24 km2/s2, and therefore appropriate launch vehicles will be Atlas V 541 and Ariane 5 fregate, both of which could carry more than 3400 kg of payload mass for the chosen C3 (see Figs. 11-12).
Figure 11:Characteristic energy graphs (Atlas V)
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Figure 12: Characteristic energy graphs (Arian 5)
After comparing both the selected vehicles, it was decided that Atlas V 541 will be more appropriate for the ARGUS mission because of the use of radioactive power. The USA has launch pads capable of launching nuclear powered systems and also has established risk assessment procedures, since it has previously used nuclear power in missions such as Horizon and Cassini. In addition, Atlas V 541 is cheaper than Ariane 5 fregate.
5.2 ATLAS V 541:
Operated by United Launch Alliance, Atlas V 541 is part of the Atlas rocket family, which have been used previously and successively as launch vehicles. Atlas V 541 has launched the Mars laboratory spacecraft in 2011. Atlas V 541 contains four solid rocket boosters that are operated at time zero, followed by a common core booster and upper stage Centaur, which produces several burns in order to ensure that the spacecraft arrives at a different orbit. Furthermore, it has a payload
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firing of 5.4 m long. The current engines used in Atlas V 541 are the RD-180, which powers the first stage by using liquid oxygen and kerosene fuel, and RL10, which uses liquid hydrogen and liquid oxygen to power the Centaur. More information about Atlas V 541 is shown in Table 5.
Table 5:Atlas V 541 Specifications Atlas V 541
Height 62.2m
Diameter 3.81m
Launch Mass 540,300kg
Stage 1 Atlas Common Core Booster
Boosters Four
Span 6.9m
Stage 2 Centaur
Launch Cost $164M
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6. PROPULSION:
6.1 REQUIREMENTS:
The first step of analysing propulsion requirements was discussed during the mission analysis part of this project, while the first orbiting requirement was fixed as it is provided by NASA’s database. It is important to note that the deep space manoeuvre remains the same and the moon tour provision was assumed by taking 10% of the required primary burn and adding a margin of 5% to it. One can determine the total ΔV requirement easily by adding this to the propulsive manoeuvres and required primary burn ΔV. This yields a value more than 3 km/s, as shown in Table 6 below.
Table 6:Orbiter ΔV Budget Orbiter ΔV Requirement Table
ΔV Margin 5%
Arrival Periapsis Velocity 25099 m/s
Desired Orbit Periapsis 22221 m/s Velocity
Required Primary Burn 2878 m/s ΔV
Moon Tour Provision 301.8 m/s
DSM TCM 513 m/s
Total ΔV Required 3692.8 m/s
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The electric propulsion system is not appropriate because this amount of ΔV needs a high thrust to weight ratio, and the thrust for electric propulsion is low (Sutton, 2001). However, a large amount of chemical propulsion is more appropriate, and provides us with four different options: solid, liquid, gas, and hybrid thrust.
6.2 GENERAL OPTIONS:
Solid propellants are easy to store and handle, and the size can be maintained by using a solid propellant due to its high density. Furthermore, the simplicity and low cost make such a propellant a good option. However, for the ARGUS mission this propellant is not appropriate because it has a low specific impulse relative to liquid propellant. Gas propellants are cheap and can be stored for a long time, but their pressure decreases over time. Another cheap type of propellant is the hybrid, which is safer and can be controlled easily but is nevertheless very complex. Therefore, the most appropriate type of chemical propellant for the ARGUS mission is the liquid propellant, which has a high specific impulse, is easy to control, and has low pressure propellant tanks. Furthermore, several liquid oxidisers have higher specific impulses than their solid counterparts and are also cheaper (ESA, 2015).
After choosing liquid chemical propulsion, a preliminary study was conducted for several thrusters, and the shared properties between these thrusters are that they are bipropellant and use the same fuel monomethylhydrazine (MMH) and dinitrogen tetroxide (MON-3). This fuel has been used extensively in space missions due to its advantages of transferring heat properties, and because its temperature of MMH is better than hydrazine. Moreover it self-ignites directly after the fuel meets the oxidizer (Sutton, 2001).
6.3 ANALYSIS:
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A preliminary study was conducted on a number of thrusters to find the most appropriate one for the ARGUS mission, and in this report only the thrusters that were selected to be the best will be discussed here, and others will be discussed in the appendix with more calculation details.
One of the thrusters that was appropriate for ARGUS is the Aerojet R-42DM which is a bipropellant engine that uses hydrazine and nitrogen tetroxide as propellants. More properties of this engine are shown in Table 7.
Table 7:The properties of R-42DM R-42DM Properties
Thrust 890 N
Chamber Pressure 140 psi (0.96 Mpa)
Expansion Ratio 200:1
Specific Impulse 327 s
Steady State Firing 1000 s
Inlet pressure range 370-200 psi, (2.55-1.37) Mpa
Mass 7.3 kg
In the launch vehicle part of the report, it was considered that the maximum payload mass should not be more than 3.4 tonnes, and therefore the selection of the thruster was on the lowest mass of propellant and the propulsion system, as on the higher thrust. Hence, the rocket equation was used in order to calculate the mass of the propellant. Using the exhaust velocity in terms of specific impulse (327 in this case) and Earth’s gravity constant, the ΔV requirement was discussed in the
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previous part of this section (3692 m/s) and the dry mass was calculated to be 940 kg, which includes the mass of the motor and 5% margin of the propellant remaining in the tanks after burn. It also includes 10% of the dry mass for heaters that remain in the tanks at an appropriate heat, and the total mass of the tanks which will be discussed later in this section. After calculating the required mass of the fuel, which is approximately 2 tonnes, it is divided between MMH and MON-3 in nominal 1:1. Then, the volume of the propellant was easily determined using the density of the fuel to get 0.84 m3 for MMH and 0.7 m3 for MON-3.
Figure 13:Thruster R-42DM (Aerojet)
After that, the fuel volume was used in order to find the mass of the pressurant by using an equation from Sutton and Biblarz (2001), which can be seen in the appendix. The pressure of the pressurant tank is 3000 psi, which is taken also from Sutton and Biblarz (2001). The gas selected for the pressurant is helium and its physical properties needed for an equation were found from the LENNTECH website, by applying propellant volume and the helium properties with the ratio of pressurant and propellant tank pressure in the equation, which yields a mass of pressurant gas equal to 30kg and a volume of 0.9 m3. It is worth noting that the pressure requirement for the propellant
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tanks was taken to be 20% more than the chamber pressure of the thruster, which yields to 1.15 Mpa. From pressurant volume, the radius of the pressurant tank can be easily calculated.
However, to calculate the volume of the tanks, the assumption of the shape of the tanks was as follows, the propellant tanks were assumed to be cylindrical in order to decrease the mass of the tank and the pressurant tank to be spherical, which is believed to decrease the stress. Aluminium was selected for tanks due to its low density and its good structure properties, since it has a tensile yield stress of 460 Mpa. However, adding a margin of 20% for safety results in approximately 307 Mpa. This value is representing the maximum range for hoop stress in tanks, therefore by rearranging the hoop stress equation, the reliable thickness for the cylindrical tanks could be easily found. The outer radius of the tanks was assumed to be 30 cm, and the length of the tank 100 cm for the MMH tank and 70 cm for MON-3. This was chosen after several attempts at different lengths and based on the overall size of the spacecraft. It is worth noting that the mass will decrease if length is increased and the radius and thickness are decreased. The pressurant tank mass was calculated in the same way but the radius fixed from pressurant gas volume was calculated to be ~ 37 cm. The stress of the spherical tanks is half of the hoop stress (Wertz & Larson, 1992; see the appendix for the equations).
As such, the required thickness was 0.25 cm for the MMH and MON-3 tanks and 1.25 cm for the pressurant tank. Hence, the total mass of MMH was determined to be about 16.4 kg, MON-3 to be about 11.48 kg, and the pressurant tank to be about 77 kg. These including a 30% margin to account other component like valves and pips. Therefore, the total mass for the propulsion system is ~ 173 kg.
The same thing was done for AMBRTM and the results was as follow, the propulsion system mass is 229 kg, and the fuel mass is 1.9 tonnes. However, the wet mass of the spacecraft of AMBRTM is slightly less than the R-42 DM because less fuel were required, and that due to the high specific impulse of AMBRTM, but the thrust is lower than R-42 DM (see table 8). Therefor the thruster that was selected for ARGUS is R-42 DM.
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Table 8:the properties of AMBRTM (Aeroj) HiPATTM Properties
Thrust 445 N
Chamber Pressure 137 psi
Expansion Ratio 300: 1, 375:1
Specific Impulse 320 s
Steady State Firing 1800 s
Mass 300:1= 5.2 kg Figure 14:HIPAT TM thruster (Aerojet Rocketdyne) Inlet pressure range 400- 100 psi
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7. COMMAND AND DATA HANDLING
There is a need for message format standards in order to ensure the compatibility of the spacecraft tele command decoding and telemetry encoding systems with the control ground stations and the ground data handling and processing systems.
Tele command message is organized in frames preceded by a group of bits for acquisition of synchronism. Each frame consists of words of several bits. The length of the frame depends upon the applied slandered. In ESA, frame consists of 96 bits.
Table 9: Frames ESA Standards Address Mode Mode First First Second Second Third Third synchronisation selection selection data data data data data data word 4 bits repeated word word word word word word 16 bits 4 bits 12 bits repeated 12 bits repeated 12 bits repeated 12 bits 12 bits 12 bits 16 Bits 4 Bits 4Bits 12Bits 12Bits 12Bits 12Bits 12Bits 12Bits
Telemetry message is organized in frames, and a group of frames constitutes a format. Each frame consists of words and starts with a synchronising code; the first frame contains a format identification word. In ESA standard, the format consists of 16 frames and each frame consists of 48 words.
7.1 SCIENCE PAYLOAD DATA CALCULATION
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
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Thermal Mapper UV Spectrometer Pixels per Image – 125000 Pixels per image – 3072 Samples per pixels – 1 Samples per pixel – 1 Bits per sample – 8 Bits per sample – 12 Estimated Data rate – 1.01E+06 Estimated Data rate – 3.69E+04
NAC (Narrow Angle Camera) WAC (Wide Angle Camera) Pixel per Image – 1048576 Pixel per Image – 4194304 Samples per pixels – 1 Samples per pixels – 1 Bit per sample – 12 Bit per sample – 14 Colour Filter – 12 Colour Filter – 14 Estimated Data rate – 1.51E+08 Estimated Data rate – 8.2E+08
Plasma Instrument GRS (Gamma Ray Spectrometer) Instruments – 5 Estimated Data rate – 1.75E+02 Sample per instrument - 1 Bits per sample - 12 Operation - 60 Estimated Data rate - 3.60E+03
Magnetometer Estimated Data Rate – 4.50E+02
Neutral Mass Spectrometer Number of Channels – 1 Estimated Data Rate – 1.5E+06
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7.2 DATA GENERATION & DUTY CYCLES
In order to find out maximum possible data per orbit, excel spreadsheet along with GMAT was used to calculate the time during which the spacecraft would actually be in an encounter with Io.
As Argus uses a very high apogee elliptic orbit around Callisto and IO’s orbit is a circular orbit. It was estimated that
Time taken by Argus to complete one orbit ~ 546 hours
Time taken by Io to complete one orbit ~ 42 hours
This implies that
Number of Argus’s Orbit = X (Number of Io’s Orbit) = 13 Io Orbits
In order to find a point of contact, it was assumed that the science orbit insertion was performed at a Zero orbit plane which resulted in the position of Io and Argus at given intervals of time.
Mean Anomaly (M) = nT
√ And T is variable time
Case 1 – Io circular orbit, e= 0.0041
Table 10: Io’s Case Time (s) Mean Motion (n) Mean Anomaly Eccentric True Anomaly Anomaly 10000 4.11E-05 0.410964 0.41266 0.41435 20000 4.11E-05 0.821927 0.82531 0.8287
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30000 4.11E-05 1.232891 1.23797 1.24305 40000 4.11E-05 1.643854 1.65062 1.6574 50000 4.11E-05 2.054818 2.06328 2.07175 60000 4.11E-05 2.465781 2.47593 2.4861 70000 4.11E-05 2.876745 2.88858 2.90045 80000 4.11E-05 3.287708 3.30124 3.31479 90000 4.11E-05 3.698672 3.71389 3.72319 100000 4.11E-05 4.109635 4.12654 4.14348 110000 4.11E-05 4.520599 4.53999 4.55782 120000 4.11E-05 4.931562 4.95184 4.97216 130000 4.11E-05 5.342526 5.36449 5.3865 140000 4.11E-05 5.753489 5.77714 5.80083 150000 4.11E-05 6.164453 6.18978 6.21516 152841.6 4.11E-05 6.281232 6.30704 6.3329
True Anomaly vs Time 7
6
5
4
3
True True Anomaly 2
1
0 0 20000 40000 60000 80000 100000 120000 140000 160000 180000 Time (s)
Figure 15: represents Io’s position at different times during 1 complete orbit
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Case 2 – Argus elliptic orbit, e = 0.693
Table 11: Argus’s Case Time (s) Mean Motion Mean Anomaly Eccentric True Anomaly (n) Anomaly 10000 9.04E-06 0.09037 0.29446 0.69149 100000 9.04E-06 0.9037 2.94171 6.90127 200000 9.04E-06 1.8074 5.87257 13.73664 300000 9.04E-06 2.7111 8.75067 20.37213 400000 9.04E-06 3.6148 11.58137 26.79083 500000 9.04E-06 4.5185 40.3835 33.01118 600000 9.04E-06 5.4222 17.08664 38.86321 700000 9.04E-06 6.3259 19.74227 44.45284 800000 9.04E-06 7.2296 22.29074 49.65577 900000 9.04E-06 8.1333 24.75888 54.53595 1100000 9.04E-06 9.9407 29.48024 63.41812 1200000 9.04E-06 10.8444 31.61313 67.23201 1300000 9.04E-06 11.7481 33.77235 70.967 1400000 9.04E-06 12.6518 36.08785 74.83269 1500000 9.04E-06 13.5555 38.10018 78.07665 1600000 9.04E-06 14.4592 40.04822 81.11659 1700000 9.04E-06 15.3629 41.9354 83.96943 1800000 9.04E-06 16.2666 43.76505 86.65088 1969065 9.04E-06 17.79444 46.69414 90.77615
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True Anomaly vs Time 120
100
80
60
True True Anomaly 40
20
0 0 500000 1000000 1500000 2000000 2500000 Time (s)
Figure 16: represents Argus’s position at different times during 1 complete orbit
The results from the parametric study implies that the time Argus orbiter would spend during a close Io flyby its high apogee orbit would be around 60-70 hours.
7.3 COMPRESSION TECHNIQUES
Data compression is the process of encoding information for reducing data volume through use of specific encoding schemes
Lossless Compression - These algorithms are reversible and exploit statistical redundancy in such a way as to represent the original data more concisely without error.
Standardization through CCSDS - Selects and specifies performant, resource efficient algorithms, supports proposed algorithm via documentation, test data, and code.
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Lossless Compression - CCSDS 121.0-B-1 algorithm for lossless (RICE) data compression, Generic, lossless, low complexity
Image Compression - CCSDS 122.0-B-1 algorithm for lossless & lossy image compression, Low complexity and memory requirements, fixed / floating point. The compression technique described in this Recommended Standard can be used to produce both lossy and lossless compression. An alternative lossless compression technique, which has lower complexity but is not specifically tailored for imagery, has been previously adopted by CCSDS.
Figure 17: represents schematic view of coder
7.4 DIGITAL MODULATION TECHNIQUE
The selection of a good modulation scheme for spacecraft data link should involve careful consideration of several factors. Bit-error-rate (BER), initial cost, power consumption, circuit complexity, channel linearity, reliability, and bandwidth must be considered and weighed in the selection process. There are various schemes applicable i.e. Frequency-shift keying (FSK), bi- phase-shift keying (BPSK), quadrature-phase-shift keying (QPSK), minimum-shift keying (MSK).
In terms of power efficiency, BPSK, QPSK, and MSK are theoretically optimal. FSK is 3 dB worse if coherent detection is used and about 4 dB worse if non-coherent detection is used.
BPSK and QPSK have side lobe spectral properties that would likely cause adjacent channel interference. This requires that the signal be filtered to reduce the side lobe strength.
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In terms of complexity, FSK is more complex but it is relatively simple when compared with PSK and MSK especially if non-coherent demodulation is used.
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. QPSK modulation scheme will be used for the mission. QPSK has a 2-bits per Hz channel capacity limit which gives it an edge over BPSK modulation in the same bit-error probability area.
Figure 18: represents Eb/N0 vs. BER
Figure 18 above stats that the Eb/N0 for 10^-5 (BER) is approximately 9.6.
Eb/N0 requirements can be reduced further by using forward error correction techniques. For the final parametric study, an Eb/N0 of 4.5 (10^-5 BER) is assumed as an implementation of Viterbi decoding technique.
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Table 12 Compression Instrument Raw Data ICER (imagers Rice Viterbi Final Rate only) algorithm
Thermal Mapper 1.01E+06 2.02E+05 1.44E+05 2.89E+05 4.33E+05
Ultraviolet Spectrometer 3.69E+04 3.69E+04 2.64E+04 5.27E+04 5.27E+04
Narrow Angle Camera 1.51E+08 3.02E+07 2.16E+07 4.31E+07 4.31E+07
Wide Angle Camera 8.20E+08 1.64E+08 1.17E+08 2.34E+08 4.31E+07
Plasma Particle Detector 3.60E+03 3.60E+03 2.57E+03 5.14E+03 7.71E+03
Gamma Ray 1.75E+02 1.75E+02 1.25E+02 2.50E+02 3.75E+02 Spectrometer
Magnetometer 4.50E+02 4.50E+02 3.21E+02 6.43E+02 6.43E+02
Neutral Mass 1.50E+06 1.50E+06 1.07E+06 2.14E+06 3.21E+06 Spectrometer
Total 9.74E+08 1.96E+08 1.40E+08 2.80E+08 9.00E+07
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8. COMMUNICATIONS
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, tracking and command (TT&C) deals with
1) Receiving control signals from ground to initiate manoeuvers and to change the state or mode of operation. 2) Transmitting results of measurements, information concerning satellite operation of equipment and verification of the execution of commands to the ground 3) Enabling measurement of ground-satellite distance and possibly the radial velocity, in order to permit location of the satellite and determination of the orbit parameter.
Telemetry and tele command links are low bit rate links, a few kilobits per second. This differs for scientific telemetry for which the data rates are much greater, typically a few megabits per second.
Figure 19: represents Data Flow
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8.1 RF (RADIO FREQUENCY) MODE SELECTION
Mode Selection is the initial most steps in designing communications system. There were two possible candidates for the design i.e. RF and Optical. Factors such as Link Reliability, Antenna mass, Antenna power consumption, TRL, Data rate capability plays an important role.
The reason behind selecting RF communication system for this mission was its design maturity and heritage which have been proven with many previous deep space missions. On the other hand, Optical communications suffer from heavy attenuation in the atmosphere resulting in link loss during bad weather. Despite of being a high power and high mass option, TRL for RF systems is much better than Optical systems because of flight heritage.
8.2 GROUND STATION AND DOWNLINK
The RF communications system will require ground stations to receive, demodulate and identify the data sent from the spacecraft. There are two possible options which can be used for downlink:
1. Deep Space Network (DSN), NASA
2. ESTRACK network, ESA
The Deep Space Network (DSN) is a world-wide network of large antennas and communication facilities, located in the United States (California), Spain (Madrid), and Australia (Canberra), that supports interplanetary spacecraft missions. DSN currently consists of three deep-space communications facilities placed approximately 120 degrees apart around the Earth.
1) The Goldstone Deep Space Communications Complex, outside Barstow, California 2) The Madrid Deep Space Communication Complex, 60 kilometres (37 mi) west of Madrid, Spain 3) The Canberra Deep Space Communication Complex (CDSCC) in the Australian Capital Territory, 40 kilometres (25 mi) southwest of Canberra, Australia near the Tidbinbilla Nature Reserve
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Figure 20: represents NASA's DSN
ESTRACK Network, ESA is a quite similar setup as DSN stations and support various ESA spacecraft and facilitate communications between ground operators and scientific probes such as XMM-Newton and Mars Express. The network consists of ten ESA-owned stations and four stations run cooperatively with other organisations. New Norcia Station (Australia), Perth Station (Australia), Redu Station (Belgium), Kourou Station (French Guayana), Cebreros Station (Spain), Maspalomas Station (Gran Canaria, Spain), Villafranca Station (Spain), Kiruna Station (Sweden), Santa Maria (Azores, Portugal), Malargüe Station, Argentina are some of the ESTRACK stations.
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Figure 21: represents ESA's ESTRACK
As stated earlier, both systems have many similarities and can be considered comparable in areas such as cost and availability. DSN holds an edge over ESTRACK in terms of its coverage and antenna sizes. The equally spaced DSN (120 degrees) are ideal for 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. ESTRACK mainly has 35 m dishes in comparison with DSN’s 70m dishes which allows a greater link with Ka band availability for moth uplink and downlink scenarios which makes DSN a more suitable option over ESTRACK. The parametric link study below represents link requirements and earth network selection.
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8.3 DATA CONSTRAINTS
Table 13 Duty Cycle and overall Data Instrument Raw Data Rate After Compression Hours Active Data Generated
Thermal Mapper 1.01E+06 4.33E+05 70 3.03E+07
Ultraviolet 3.69E+04 5.27E+04 273 1.44E+07 Spectrometer
Narrow Angle 1.51E+08 4.31E+07 70 3.02E+09 Camera
Wide Angle Camera 8.20E+08 4.31E+07 70 3.02E+09
Plasma Particle 3.60E+03 7.71E+03 273 2.10E+06 Detector
Gamma Ray 1.75E+02 3.75E+02 70 2.63E+04 Spectrometer
Magnetometer 4.50E+02 6.43E+02 546 3.51E+05
Neutral Mass 1.50E+06 3.21E+06 273 8.76E+08 Spectrometer
Total 6.96E+09
As discussed in the previous section, the zero orbit plane was selected in order to calculate true anomalies for the IO encounters. Instruments like Magnetometer would be switched on during the whole orbit whereas Plasma Particle Detector, Neutral Mass Spectrometer and Ultraviolet Spectrometer would be switched on during half of the orbit time in order to map the Jupiter IO readings before entering into long apogee phase. Instruments specifically designed for Io’s imagery
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like Thermal Mapper, Narrow and Wide angle Camera would be switched on during the close perigee encounters. Total Maximum Data generated was fond out to be around 7 GB. An extra 30% contingency margin is also applied and the link calculations done were based on 9.5 GB data per orbit.
8.4 LINK BUDGET
A link budget is accounting of all of the gains and losses from the transmitter, through the space path to the receiver in a telecommunication system. It accounts for the attenuation of the transmitted signal due to propagation, as well as the antenna gains and miscellaneous losses.
Where P is the transmitter power, is the line loss, is the transmitter antenna gain, is the free space path loss, is the atmospheric path loss, is the receiver antenna gain, K is Boltzmann's constant, Ts is the system noise temperature and R is the data rate.
In other words, Link is defined as the relationship between data rates, antenna size, propogation path length and transmitter power. This relationship is known as the link equation as presented above.
Figure 22: represents Key Link Parameters
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The output of the link budget after all parameters have been chosen is Eb/N0 which is defined as the ratio of received energy-per-bit to noise density.
The equation above can also be written in the decibel output form i.e.
Where P, , , are in dB, Ts is in Kelvin and R (Data Rate) is in bps.
An excel sheet was produced in order to study different antenna sizes with different transmitted powers. A few estimations were made from the JUNO telecommunications guide book. A parametric study was conducted on X and Ka band with 34 m and 70 m earth network dishes. The central frequencies used for X and Ka bands were 10 and 33.5 GHz respectively. Study conducted was based on different antenna sizes and power along with required Eb/No. Table Below represents an achieved Eb/No of 4.5 for different bands and requirements.
Table 14: Link Budjet analysis Parameters Ka (34m) X (70m)
Data Rate (bps) 70000 (Orbit time * Data) 70000
Range R (in Km) 8.15E+08 (Orbit Distance) 8.15E+08
Speed of light, c 3.00E+08 (Constant) 3.00E+08
Frequency (Ka band Central) (in 3.35E+10 (Central Frequencies) 1.00E+10 Hz)
Transmitter Power (In W) 26 Assumed (For Required 65 Link)
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Transmitter Power (In db) 14.14973 18.129
Antenna Size (HGA) (in m) 2 Assumed 2
Atmospheric Attenuation (in db) -1.09 (Juno Mission -1.09 Reference)
Line Loss (in db) -0.25 (Juno Mission -0.25 Reference)
Efficiency 0.55 0.55
Transmitter Gain (in Hz) 2.70E+05 2.41E+04
Transmitter Gain (in db) 5.43E+01 4.38E+01
Space Path loss (in Hz) 1.31E+30 1.17E+29
Space Path loss (in db) -3.01E+02 -2.91E+02
Antenna Size (DSN) (in m) 34 (DSN/ESTRACK) 70
Efficiency 0.6 0.6
Receiver Gain (in Hz) 8.53E+07 3.22E+07
Receiver Gain (in db) 7.93E+01 7.50E+01
Noise (in K) 114.43 (Sky, Planetary, 114.43 Galactic, Solar, Internal)
Link Calculation (Eb/N0) 4.84E+00 4.59E+00
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Power vs Link (X band 34 and 70 m) 7
5
3
1 70 m
Eb/No 34 m -1 0 50 100 150 200 250 300
-3
-5 Power (W)
Figure 23: represents Transmitted Power vs. Link for 70 m and 34 m (X band)
Power vs Link (Ka band 34m) 6.00E+00
5.00E+00
4.00E+00
3.00E+00 Eb/No 2.00E+00
1.00E+00
0.00E+00 5 10 15 20 25 30 Power (W)
Figure 24: represents Transmitted Power vs. Link for 34 m Ka
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Size vs Eb/No
10
9
8
7
Eb/No 6
5
4
3 1.5 1.7 1.9 2.1 2.3 2.5 2.7 2.9 3.1 3.3 3.5 Size
Figure 25:represents Transmitted Power vs. Link for 34 m Ka
Summary
In order to achieve the required Eb/No i.e. 4.5, the parametric study above considered X and Ka bands with their central frequencies. As the maximum data output was found out to be 9.5 GB after compression techniques, the data rate of 70 Kbps was constrained. With a constrained data rate and data volume, total transfer time was calculated to be around 37 hours. Argus’s instruments would be functional for 60-70 hours per orbit i.e. during Io’s encounters at Orbit’s periapsis which would also be a window for the data transfer as Earth’s visibility point of view. Argus’s Antenna size was also constrained to 2 m in order to reduce the overall weight. Results indicated that a 70 m X band (Earth) would be able to produce the required link at 65 Watts, whereas 34 m X band (Earth) required a very high power of 275 watts in order to achieve the same link. Ka band was also
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effective as the required link was achieved at a 26 watt 34 m antenna. During the PDR, use of DSN was suggested in order to access 70 m dishes. Final studies suggested that DSN would be a viable option in terms of power requirements and ESTRACK 34 m could be used as a backup.
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9. RADIATION ENVIRONMENT AND MAGNETIC FIELD:
9.1 OVERVIEW:
Space radiation has a negative impact on the spacecraft and electronics, and therefore it is important to have knowledge about the radiation that a spacecraft will be exposed to during the mission. The three most important space radiations are trapped charged particles, galactic cosmic rays, and solar particle events (Barth et al., 2003). For the ARGUS mission, during parking orbit the primary radiation environment will be trapped particles from Earth’s radiation belts. In addition, during the deep space manoeuvre, and the transfer phase that occurs before arriving at Jupiter’s magnetic field, the dominant radiation environment will be a solar particle event. Finally, during the orbiting mission around Jupiter the primary radiation will be from trapped particles. However, Jupiter’s magnetic field acts as a shield for the spacecraft during this latter phase, thus protecting it from the other radiation sources. Therefore, during the ARGUS mission the two most profound radiation sources will be from Earth and Jupiter’s magnetic fields. Trapped radiation belts are regions where energetic particles such as electrons and protons are concentrated, and thus such belts are produced on both sides of a planet. The two radiation belts that ARGUS will experience are those of Earth and Jupiter.
9.2 THE VAN ALLEN BELTS:
The Van Allen belts are two belts surrounding Earth on both sides, which are produced by the interaction between the charged particles and Earth’s magnetic field. On each side of Earth there is an inner and an outer belt. The inner belt generally contains trapped electrons (300 to 1200 km), while the outer radiation belt contains proton and electron particles (Barth et al., 2003).
9.3 JUPITER’S RADIATION ENVIRONMENT:
Jupiter has a very strong magnetic field, which is the largest in our solar system. The strength of this magnetic field at high altitudes is less than at low altitudes because some energy is absorbed by the
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interaction with the solar wind and also by the rings and moons that are located in the Jovian magnetosphere. Jupiter’s magnetosphere is different from those of all other planets because it derives most of its plasma from the satellite moon Io and not from the solar wind. The Jovian magnetosphere is divided into three areas: the inner magnetosphere, which extends to a distance of ten Jupiter radii, the middle magnetosphere, which is located approximately between 10 RJ and 40 RJ, and the outer magnetosphere, which is located from 40 RJ to the magnetopause (Khurana et al., 2004). In this section, the inner and middle magnetospheres will be discussed as they are the regions through which ARGUS will travel.
9.3.1 Inner magnetosphere:
The inner magnetosphere produces plasma for the Jovian magnetosphere. It is known that the plasma torus of Io, which consists of several million tonnes, is located in this region approximately between a radial distance of 5.2RJ and 10RJ. The plasma torus is spread slowly by the assistance of “instabilities feeding on the centrifugal force” (Khurana et al., 2004). The plasma in this region is less than 0.2 and it has a minimum effect on the magnetic field because the internal magnetic field is very strong, as such, the temperature of the torus plasma is low (Khurana et al., 2004).
9.3.2 Middle magnetosphere:
In the middle magnetosphere, the plasma’s co-rotation with the Jovian magnetosphere gradually decreases, and the co-rotation of the magnetosphere via radial current produces aurorae in the Jupiter ionosphere. This occurs by “accelerating electrons into the ionosphere” (Khurana et al., 2004). Currently, the azimuth are extreme near to the equatorial plane of Jupiter, in the thin current sheet, and this causes the production of magnetic field perturbations, which are equal to the inner magnetosphere with a distance of more than 20 RJ (Khurana et al., 2004).
9.3.3 Exploring the magnetic field of Jupiter:
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Six different spacecrafts have been used to explore the magnetosphere of Jupiter, excluding Cassini, which in December 2000 only skimmed the dusk boundary region of Jupiter.
Pioneer 10: In December 1973, Pioneer 10 passed by Jupiter on a trajectory with a low inclination; the trajectory had a distance of less than 10 RJ. The mission confirmed the presence of a large Jovian magnetosphere (Khurana et al., 2004).
Pioneer 11: In December 1974, one year after Pioneer 10, Pioneer 11 arrived at Jupiter with a high inclination and a radial distance of 1.6 RJ. Because of its trajectory features (low altitude and high inclination), the spacecraft was able to provide a clear description of Jupiter’s internal field (Khurana et al., 2004).
Voyagers 1 and 2: In March and July of 1979, Voyagers 1 and 2 arrived at Jupiter with radial distances of 4.9 RJ and 10.1 RJ, and approximately equatorial trajectories. A dense plasma torus was discovered by Voyager 1, whilst Voyager 2 discovered Jupiter’s equatorial sheet current (Khurana et al., 2004).
Ulysses: In February 1992, Ulysses made the closest Jupiter flyby, at an approach of 6.3RJ of Jupiter’s radius. This spacecraft discovered the “dusk high latitude region of Jupiter” (Khurana et al., 2004).
Galileo: In December 1995, Galileo arrived at Jupiter with the goal of exploring the planet’s magneto tail. The spacecraft was able to complete its studies in the dusk of Jupiter as well as the post-non-sectors of the magnetosphere (Khurana et al., 2004).
Therefore, these six spacecrafts were able to map completely the low altitude of Jupiter’s magnetosphere.
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9.4 ANALYTICAL METHODS:
9.4.1 Energetic Particle Radiation:
In general, the energetic radiation environment consists of magnetically trapped charged particles, solar protons, and galactic cosmic rays. However, it is the penetrating particles that pose the main problems, which include upsets to electronics, payload interference, degradation and damage to components and solar cells, and deep dielectric charging. In order to study the effect on the ARGUS spacecraft, the radiation environment needs to be assessed carefully. This chapter presents the predicted radiation environment for ARGUS.
For ARGUS, by far the main contribution to the radiation dose will be from the radiation belts, primarily from high energy electrons.
9.4.2 Trapped Particles:
As discussed above, the trapped charged particles are mainly electron and proton particles. The environment model used in this part is the mean D&G83 and Salammbo model. Figure 26 shows the predicted fluxes for electron and proton integral flux for the whole ARGUS mission.
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ARGUS mission Average electron and proton Flux 1.00E+09 1.00E+08 1.00E+07 1.00E+06
1.00E+05 /s] 2 1.00E+04 1.00E+03 proton 1.00E+02 electron 1.00E+01
Integral Flux[/cm Integral 1.00E+00 1.00E-01 0 200 400 600 800 1000 1200 1.00E-02 1.00E-03 1.00E-04 Energy MeV
Figure 26:Predicted average trapped electron and proton flux for a complete ARGUS mission
9.4.3 Worst Case Fluence:
The worst case fluence of the mission is shown in Figure 27. The Salammbo and Divine & Garrett models have been used in the SPENVIS to calculate the worst hours for science orbit with the nearest point to Jupiter (5 RJ).
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Figure 27:The average integral flux above 1 MeV for the ARGUS mission
From the spectrum above, the worst case was determined to find the worst hours for the ARGUS mission (12, 42, and 100 hours) and then was used as a normalisation for the electron and proton spectrum to find the total ionising dose in the worst hours (for more information, see the appendix).
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10. TOTAL IONISING DOSE AND SHIELDING:
Dose depth curve can identify the ionising dose environment. The dose depth offers dose as a function of thickness of shielding. For the Argus mission, Giant 4 has been used in the SPENVIS to determine the ionising dose. The reference geometrical configuration used for the dose-depth curve is a hollow spherical silicon target with a hollow aluminium sphere. Figure 28 shows the total ionising radiation dose in the silicon target, with a thickness of 10 μm as a function of spherical aluminium shielding with different thickness.
Total Ionizing Dose for Electron and Proton particles 7.00E+05
6.00E+05
5.00E+05
4.00E+05
3.00E+05 electron proton 2.00E+05 Total Total Dose (rad) 1.00E+05
0.00E+00 0 5 10 15 20 25 -1.00E+05 thickness (mm)
Figure 28: Dose in silicon target as a function of spherical aluminium shielding
The figure above illustrates the total ionising dose for electrons and protons as a function of shielding thickness, and the bars in the curves indicate the error percentage. It is worth noting that the Radiation Design Facture RDF was not included. The shielding thickness of 22 mm (TID) was calculated to be approximately ~3.6×105. However to achieve 78 Krad of the total dose, 3 cm were found to be the ARGUS shielding thickness (see appendix for more details).
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11. WEIGHT
11.1 INSTRUMENT SPECIFICATIONS
Table 15: weight criteria Instrument Mass TRL Margin Mass Power Margin Power Name (Kg) (Kg) (W) (Harness (W) with with and margin margin TRL) UV 5.2 6-7 25 6.5 8.5 30 11.05 Spectrometer
Thermal 11.2 6-7 25 14 14 30 18.2 Mapper
Gamma Ray 9.2 5 30 11.96 16.5 35 22.275 Spectrometer
Narrow Angle 8.6 6-7 25 10.75 15 30 19.50 Camera (NAC)
Wide Angle 4.8 5 30 7.5 12.8 35 17.28 Camera (WAC)
Plasma 10 6-7 25 12.5 15 30 19.5 Particle Detector
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Instrument
Neutral Mass 10.3 6-7 25 12.875 32 30 41.6 Spectrometer
Dual 3 6-7 25 3.75 3.10 30 4.03 Technique Magnetometer
TRL Margins above are applied with respect to ESA standards, TRL margin 25 for mass and 30 for power are used for the instruments which have been used for deep space missions but not the similar kind of environment (Io) whereas 35 for mass and 35 for power margin have been used for the instruments below TRL 5.
Total Mass (without TRL margin) – 62.3 Kg
Total Mass (with TRL margin) – 79.835 Kg
Total Power (without TRL and Harness margin) - 116.9 W
Total Power (with TRL and Harness margin) – 153.435 W (Peak Power for instruments)
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11.2 SYSTEM WEIGHT DATA COLLECTION
Table 16 Overall Mass System Type Mass with margin (Kg)
Instruments 80 Spacecraft Bus ( Standard Nuclear 225 Powered) Telecommunications 70 Command and Data Handling 30
Attitude Control 25 Power (MMRTG & RHU) 180 Propulsion 173 TOTAL (including system margin) 973 Dry Propellant 1655 Shielding Mass 345 TOTAL Wet Mass 3350
A system margin of 20% has been used for the estimations.
Weight TRL – Individual Margin (Instruments) + Subsystem Margin – 45% - 50 % avg
For Power estimations it was found that peak power requirement if all the instruments are switched on would be about 458 W including a system margin of 25%. Also with the use of duty cycles, the power consumption would be around 400 Watts which justifies the use of 4 MMRTG’s.
Power TRL – Individual Margin (Instruments) + Subsystem Margin = 50% avg
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12. POWER SOURCE
Option 1
Solar Power
Solar power is one of the available option for Argus Orbiter and would also have heritage (Juno/Juice) by the time Argus gets launched. The spacecraft requires up to 290-300 W of power after taking an account of the calculated instrument duty cycles. A calculation was done in order to calculate the area of the solar panels required for the mission which is presented below
Where r = 5.2 AU (Jupiter) and 5.45 AU (Io), Assuming 5.45 AU for the calculation.
L (solar luminosity) = 3.83*10^26 W
P (power required) = 300 W, it becomes clear that the solar panel has a minimum required area of
A= 47.11
Assuming an annual degradation of 1.03 up to eight years, Required Area (A) = 84
Although solar panels have been considered for JUICE and JUNO, Argus mission emphasize just on IO which has the worst radiation in solar system. Shielding costs will be significantly higher and TRL estimations will be larger in comparison with other power options.
Option 2
MMRTG –
It 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 and only requires a 20% contingency margin
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Where t = 8 years, t1/2 = 86.4 years (P 238)
Each MMRTG has a mass of 43 kilograms and provide 125 Watts. In order to meet the power requirements, 3 MMRTG’s will be required
MMRTG degrades by 1.58 % average per year. Using the equation below
Option 3
RTG (GPHS) 238-Pu –
Production development of these NASA RTG’s was halted mainly due to the high cost and shortage of 238-Pu. New Horizons was the last mission which uses GPHS 238-Pu and used spare parts from previous missions and a combination of old and new fuel. It had a mass of 55.9 kilograms, which required about 8.1 kilograms of fuel to produce about 258 Watts at launch.
Both Solar and RTG powered missions were examined as part of this study before the selection for the final phase.
Cost and TRL - Solar powered missions were slightly less costly and would have heritage (i.e. Juno) in time for this launch opportunity. MMRTG’s too have heritage from Pioneer, Voyager, Ulysses, Galileo, Cassini and New Horizons. As Argus Orbiter uses a deep space manoeuvre high apogee orbit, MMRTG’s were found to be the better option for the mission even with a slight extra cost.
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Figure 29:represents RTG heritage
Safety - Plutonium, the active ingredient in most radioactive thermal generators (RTGs) is a toxic heavy metal. Plutonium (required amount) is sealed inside a hard, radiation-proof shell. The shell is designed to survive all conceivable accidents, so even in the unlikely event a launch goes wrong, none of the radioactive particles will escape. Because MMRTGs use radioactive decay, each launch of any vehicle with an MMRTG requires Presidential approval. Given that they are the only reasonable way to power satellites beyond the orbit of Mars (as solar panels stop becoming effective), MMRTGs are necessary. RTG's cannot explode like a nuclear weapon. Nuclear weapons are made of high-grade uranium and have to be arranged very carefully to go into fission. RTGs, at
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best, can produce a warm fizzle.
Also, USA has launch pads capable of launching radioactive powered systems along with the risk assessment procedures. Hence, MMRTG’s are safe.
Example – While launching Cassini, an extra safety measure that was procured by NASA was the potential release of plutonium into the environment from the baselined RTG’s and RHU’s in any unlikely event of severe launch or inadvertent entry into earth’s atmosphere during a flyby.
Launch Vehicle – As this is a MMRTG powered mission, Atlas Launch vehicle is selected which satisfies both the weight and safety criteria. MMRTG’s can only be launched with American launch pads which would make this a NASA based mission.
Power Data Collection
Table 17 Power System System Power (W) Power inc Margin (W) – worst case
Structure and Mechanism (Std. N/A N/A Bus)
Instruments 116.9 146.9
Power Source 20 25
Command & Data (Ref) 30 37.5
Propulsion (Ref) 27 33.75
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Thermal Control 35 43.5
Telecommunications 26 32.5
Attitude Control (Ref) 39 48.75
Total (dry) 293.9 367.9
Table above represents
Total power required without margin – 293.9 watts
Total power required with margin if all the instruments are running (worst case) – 367.9
Using the duty cycles it was found that total power required would be 290-300 Watts and the power source selected was MMRTG’s based on the study and trade-offs presented above
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13. IMPACT OF OTHER MISSIONS ON ARGUS ORBITER
Table 18 Mission Comparison Parameters Juno (NASA) Juice (ESA) IVO (NASA) Argus (Leicester)
Mission Launch 2011, August 5 2022 2021, May 29 2027, Jan 27
Science Orbit 2016, July 2030(J) - 2026, Feb 2031, Nov 22 Insertion 2033(G)
IO Flybys Possible None 6-8 13-16 Encounter
1) Microwave 1) Janus Camera 1) Narrow 1) Thermal Mapper Radiometer System Angle Camera 2) Ultraviolet 2) Infrared 2) Imaging 2) Wide Angle Spectrometer Aurora Mapper Spectrometer Camera 3) Narrow Angle 3) 3)Ultraviolet 3) Thermal Camera Magnetometer Spectrograph Mapper 4) Wide Angle 4) Gravity 4) Wave 4) Particle Camera Science instrument Environment 5) Plasma Instrumentation 5) Aurora 5) Laser Package (Ion, Instrument Distribution Altimeter Plasma) 6) Magnetometer Experiment 6) Radar 5) Dual 7) Gamma Ray 6) Energetic 7) Fluxgate Spectrometer Particle Detector Magnetometer Magnetometer 8) Neutral Mass 7) Radio & 8) Radio and Spectrometer Plasma Sensor Plasma Wave 8) Ultraviolet Investigator
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Spectrograph 9) Doppler 9) JunoCam Experiment 10)Geophysics and Gravity instrument
Jupiter Science Ganymede, Io Study Io Study Major with possible Callisto and Objectives moon science Europa Study study
Jupiter orbit, Radiation Aluminium Thick Aluminium Radiation using titanium Shielding, Shielding and Shielding Environment vault (1 cm Possible Juno like vault thick) for Ganymede in order to instruments Lander (RSA) minimize the radiation
1) Juno has limited imaging capabilities, but it could provide monitoring of Io's volcanic activity using its near-infrared spectrometer, the Jupiter Infrared Aurora Mapper (JIRAM).
2) Juice mission is focused on studying three of the Jupiter’s Galilean moons i.e. Ganymede, Callisto and Europa and would not consider Io’s exploration. A possible Ganymede lander Laplace- P is currently being evaluated by the Russian Space Research Institute which could be a part of Juice mission.
3) Another Discovery class mission known as IVO Explorer has been proposed by NASA (Expected arrival date, mid 2020’s). If selected, IVO would use high-inclination orbits of Jupiter to flyby Io at least six times, perhaps more if an extended mission got approved considering the radiation effects on structure and instrumentation. The main goals of this proposed mission include
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measuring Io's volcanic eruption temperatures, determining the composition of Io's lavas, sampling its volcanic plumes through mass spectrometry, and mapping Io's internal structure using electromagnetic induction sounding.
As this would be the first mission solely concentrated on Io’s exploration, the radiation effect results alongside with other data collected would be useful in order to make necessary changes in the Argus Explorer. Although the orbit for IVO would be a shorter one in comparison with Argus as Argus uses a high apogee Callisto orbit which would limit the radiation effects up to some extent which could result in more possible flybys. Also, IVO will consist a range of instruments like NAC, WAC, and Thermal Mapper. Argus Orbiter is designed for a detailed longer exploration of IO with higher number of flybys and longer mission and hence would be able to produce a data with broad range and time period. As Io is a quite active surface, a longer mission duration is a must.
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