[American Institute of Aeronautics and Astronautics SpaceOps 2006 Conference - Rome, Italy ()]...

American Institute of Aeronautics and Astronautics

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Design of a Jupiter Europa Orbiter Using Emerging Radioisotope Power System Technologies

S. Ravindran* and T. Bowling† Space Research Centre, School of Engineering, Cranfield University, Cranfield, Bedford MK43 0AL, UK

This study analyses a single Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) unit,

and two Stirling Radioisotope Generator (SRG) units for powering Jupiter Europa Orbiter. The RTG power system will independently produce an uninterrupted power supply, as opposed to a solar powered orbiter, as studied by Astrium for ESA, which depends on distance from Sun, eclipse and Jupiter’s occultation. The orbiter’s thermal control system uses excess heat from the RTG to maintain the spacecraft’s temperature using a heat transfer loop. The additional heat loads received by the spacecraft during Venus and Earth flybys are removed with this heat transfer loop through the RTG radiators. The MMRTG provides a continuous power output of 110 W with a 13% reduction in mass compared to a solar powered equivalent. The total radiation dose from the MMRTG is about 5 krad behind 10 mm aluminium shielding, which is much lower than the expected mission natural radiation dose. The electromagnetic interference produced by the MMRTG magnetic field strength is only 17 nT which is less than the NASA standard requirement of 25 nT at one meter distance. The power system with two SRG units has an overall mass saving of 9% compared to the solar cell equivalent and requires only four General Purpose Heat Sources (GPHS) to produce 220 W of electrical power. It emits less radiation and has a weaker magnetic field than the MMRTG. Results show that a Jupiter Europa Orbiter spacecraft designed with single MMRTG unit or with two SRG units as power source would be the best choice for the proposed ESA Jovian Minisat Exploration mission.

Nomenclature I = radiation intensity after passing through the absorber I0 = initial radiation intensity µ = linear absorption coefficient x = thickness of the absorber

I. Introduction The exploration of Jupiter’s moon Europa with solar cell powered spacecraft has been studied by the European Space Agency. However results suggest that a Radioisotope Thermoelectric Generator (RTG) power system would be more appropriate. The objective of this study is to design an RTG powered orbiter to survey Europa for the presence of liquid water which could indicate the existence of life now or in the past. ESA’s Jovian Minisat Explorer mission scenario in Fig. 1 foresees two small spacecraft, the Jupiter Relay Spacecraft (JRS) in a highly elliptical orbit with perijove of 907948 km and apojove of 1880240 km around Jupiter outside the high radiation zones and the Jupiter Europa Orbiter (JEO) orbiting Europa at 200 km polar orbit. The entire Jovian minisat mission lifetime is about 8 years. The JRS acts as a relay between the JEO and Earth. The JRS will carry all subsystems that are not directly required for the Europa observation mission. JRS will carry a complete communication system to facilitate the link between Earth and the JEO, data processing and data storage units for both JEO science and JRS science. The relay satellite will also carry a small highly integrated scientific payload suite dedicated for the study of Europa’s plasma environment and atmospheric composition, and Jupiter’s dust,

* Postgraduate Student, Space Research Centre, School of Engineering, Cranfield University/[email protected] † Supervisor, Space Research Centre, School of Engineering, Cranfield University/[email protected]

SpaceOps 2006 Conference AIAA 2006-5586

Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


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plasma environment as well as imaging instrumentations. The Europa orbiter will include a highly integrated remote sensing payload suite and a communication system for telemetry and telecommand with the JRS and Earth. The JEO will carry a penetrator analyser of less than 15 kg mass to analyse composition of surface material of Europa. The

envisaged dry mass limit for JRS is about 600 kg and JEO is about 400 kg.

The ESA’s baseline spacecraft has been designed based on a solar power system due to radioisotope powered system launch safety constraints and lack of existing hardware in Europe [1]. However the low solar fluxes near Jupiter, frequent eclipses, occultation, and moreover present technology difficulty in producing GaAs solar cells to operate efficiently in low intensity and low temperature (LILT) environment has led to the redesign of the spacecraft with a RTG as power source. The RTG technology will not only surpass the solar cells on power production but it will also reduce the mass of the total spacecraft and its complexity. The objective of this study is to redesign the JEO spacecraft after careful review of Europan radiation environment, JEO orbit and payload requirements. Also the aim is to define and conceptually design the new RTG powered spacecraft, spacecraft subsystems, configuration and integration of spacecraft, communication architecture, and the JEO

critical technology requirements. It also introduces new concepts in RTG technology namely the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) and the Stirling Radioisotope Generator (SRG) for the JEO mission and their advantages. Several trade-off studies were performed and the best possible scenarios were selected for each subsystem. The relatively high radiation environment of Jupiter especially around Europa and the new Europan environmental committee proposal on cleanliness requirement and planetary protection guidelines by the Committee on Space Research (COSPAR) makes the design of JEO mission more challenging.

II. Orbital Analysis The preferred orbit for science is a near circular polar orbit at an altitude around or less than 200 km. The JEO operational time period is mainly governed by the strong orbit perturbations caused by Jupiter, radiation dose and limited propellant availability. Hence it will have highly limited operational period. The present proposed orbiter has a lifespan of 60 days of operational phase [2]. The JEO with MMRTG as power source should also be disposed of after completion of the science phase into Jupiter’s atmosphere. The JEO orbital altitude and inclination has been selected to achieve maximum coverage of the moon however previous study [3] shows that the orbit is relatively unstable in an inclination range between 50o to 125o degrees. The polar orbit might also be unstable and hence a more stable orbit could be considered by lowering the inclination. One of the less risky high area coverage inclinations is 75o degrees. The worst case scenario suggests having the orbit around 45o inclination which covers only around half the area of Europa; however the stability of the orbiter at this inclination encourages considering this orbit in case other orbits prove to be more unstable.

III. Payload System Selection The science instruments are selected from the strawman payload [2] to achieve the science objective. The payload instruments and their science objectives are listed in Table 7. The science observation covers mapping the thickness of the ice layer, structure determination, topography, surface reflectivity, geology, surface composition, surface temperature, tidal processes, magnetic field, electron environment and Jovian radiation environment. The science packages need to be confined into small spaces and should operate with lower power resource compare to previous missions. This objective is achieved by using a Highly Integrated Payload Suite (HIPS) [2]; with this mass and power could be saved considerably and shielding requirements also minimise due to clustering of instruments together. However accuracy, integration and programming require new research and development in order to achieve better science objectives using HIPS. The limited power availability divides the science observation into two distinct

Figure 1. ESA Jovian Mini Satellites.


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modes; one is without the Ground Penetrating Radar (GPR) operating while other science instruments operate and the other mode is only GPR operating while other instruments are quiescent. HIPS also requires a considerable amount of autonomy as the distance between Earth and JEO during science operation is too far to allow controlling from the ground station. The HIPS autonomy needs to address short initiation time for instruments, long operation time allocation and repeated switch-off and on according to power availability and science requirements. The science data will be stored in a 50 Giga bit RAM over a period of 10.6 days and will be transmitted to JRS while it is in the nearest distance to JEO. The RAM needs to be radiation hardened up to 1 Mrad and requires considerable amount of research and development, as does the miniaturisation of the payload with less power. The radar deployment mechanism needs to be developed with the need to deploy it after 7.5 years from its stowed position. The preliminary study shows that the HIPS requirements give immense challenges to be overcome for the JEO mission to be successful. The HIPS concept was proposed by ESA on similar projects [4].

IV. Isothermal Analysis The operational and working temperature of spacecraft electronics is in the range of -100 C to +500 C [5]. All instrument hardware located internally, externally and in the bus shall be in the range of -200 C to +500 C [5]. Throughout the mission the propulsion subsystem needs to be maintained within the temperature range of +50 C and

+500 C [5]. The thermal system for the JEO mission needs to address both high heat loads during Earth, Venus flyby and cold conditions during the deep space transit and science operation. The simple isothermal model described in Fig. 2 considers solar flux, Jupiter and Europa albedo and IR and spacecraft electronics heat output. The high heat load during Earth and Venus flyby could be minimised by facing the parabolic antenna toward the Sun and hence minimising the exposed area on the spacecraft. The excess heat inside the spacecraft will be rapidly absorbed by the thermal loop shown in Fig. 3 and dissipated through the MMRTG radiators. The thermal equilibrium inside the spacecraft is achieved by keeping the inside temperature within 283 K to 293 K throughout the mission.

The cold case scenario dominates the entire mission more than the hot case scenario. The best option to keep the spacecraft warm enough is to use the excess heat available from the MMRTG. A portion of this excess heat will be used to transfer about 147 W of thermal energy into the spacecraft in total during the coldest possible condition. The input thermal energy will be adjusted according to the equilibrium condition prevailing inside the spacecraft. The thermal control system for JEO would use loop heat pipe system moving around the spacecraft and dissipating heat during cold conditions and absorbing and radiating heat during hot conditions. The passive control system will use multi layer insulation (MLI) inside the spacecraft and goldised kapton paint coating on the external surface. The system proposed for this mission is not new and it is similar to a Mars Rover concept [6] which has proved to be working on the present Rover mission.

Figure 2. Simple JEO Thermal Model, considering Qenvironmental (Solar flux, Albedo & IR of Europa and Jupiter), Qrtg (MMRTG heat load) and Qelectronics (spacecraft electronics heat output)

Figure 3. Thermal control system for JEO


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V. Power System The power system for the JEO mission is mainly governed by distance from the Sun, the rotational period of Europa and thermal heat requirement. Since the solar flux is in the range of 50 W/m2 a solar concentrator principle was proposed for the baseline mission [1]. The system theoretical study shows that a considerable amount of solar power could be collected using solar panels with concentrators at JEO during end of life (EOL); (15.8 W/m2 with a panel weighing about 3.9 W/kg [1]). However this technology has to be developed for GaAs type solar panel in low intensity low temperature (LILT) environment where the Sun light availability is also limited by occultation of Jupiter. Both are considered to be major hurdles in successful implementation of solar cell technology in the vicinity of Europa. Possibly the best way of producing sufficient power for the spacecraft and for thermal heating is to use MMRTG technology.

Table 1 Power Level Estimates for the JEO Spacecraft Subsystem Duty Cycle Mode 1

LaunchMode 2 Orbit

insertion / correction

Mode 3 Science without

GPR

Mode 4 Science

only GPR

Mode 5 Science

and Telecom without

GPR

Mode 6 Telecom

only

Heritage

W W W W W WPayload 7.3 7.3 25.9 28.8 25.9 7.3 Ground penetrating radar On Demand 2.5 2.5 2.5 24.0 2.5 2.5 [2]Stereo Camera On Demand 0.2 0.2 1.2 0.2 1.2 0.2 [2]Visible - Near Infrared spectrometer On Demand 0.2 0.2 1.2 0.2 1.2 0.2 [2]Europa radiometer On Demand 0.4 0.4 2.4 0.4 2.4 0.4 [2]Magnetometer On Demand 0.3 0.3 0.6 0.3 0.6 0.3 [2]Europa Laser Altimeter On Demand 2.0 2.0 6.0 2.0 6.0 2.0 [2]Standard Radiation Monitor On Demand 0.4 0.4 2.4 0.4 2.4 0.4 [2]Europa X-ray spectrometer On Demand 1.0 1.0 6.0 1.0 6.0 1.0 [2]Europa Gamma and Neutron Spectrometer On Demand 0.3 0.3 3.6 0.3 3.6 0.3

[2]

AOCS 14.0 30.0 20.0 20.0 20.0 20.0 Star Trackers Always on 4.0 4.0 4.0 4.0 4.0 4.0 [10] Bepi

ColomboInertial Measurement Unit On demand 10.0 10.0 [1]Reaction wheels Always on 16.0 16.0 16.0 16.0 16.0 [1]

CMDS 24.0 24.0 24.0 24.0 24.0 24.0 Processor Always on 24.0 24.0 24.0 24.0 24.0 24.0 [1]

Power 7.0 7.0 7.0 7.0 7.0 7.0 MMRTGPower conditioning 7.0 7.0 7.0 7.0 7.0 7.0 [10]

Telecom 1.3 1.3 1.3 1.3 1.3 1.3 UHF Transiver (Lander) Always on 1.3 1.3 1.3 1.3 1.3 1.3 [10]

Thermal 12.5 12.5 12.5 12.5 12.5 12.5 Heaters Always on 5.0 5.0 5.0 5.0 5.0 5.0 [10]Propulsion Tank Heaters Always on 5.0 5.0 5.0 5.0 5.0 5.0 [10]Propulsion Line Heaters Always on 2.5 2.5 2.5 2.5 2.5 2.5 [10]

Propulsion 2.3 2.3 2.3 2.3 2.3 2.3 HP transducer Always on 0.7 0.7 0.7 0.7 0.7 0.7 [10]LP transducer Always on 1.7 1.7 1.7 1.7 1.7 1.7 [10]

Continous Power MMRTG (1) 68.4 84.4 93.0 95.9 93.0 74.4

Telecom (on demand) 15.6 15.6 - - 24.0 24.0 X-band SSPA 1 hr / day 15.6 15.6 Ka-band SSPA 6 hr / 10.6 days 24.0 24.0

Thrusters ignition (on demand) - 24.0 - - - - Thruster - I (22N - 4 nos) 2 hr 12.0 [10]Thruster - II (22N - 4 nos) 2 hr 12.0 [10]

Contingency power requirment on demand 20.0 20.0 20.0 Payload additional power 3 hr / day 10.0 10.0 10.0 Other instruments additional power 3 hr / day 10.0 10.0 10.0

Intermittent Power Battery (2) 15.6 39.6 20.0 20.0 44.0 24.0

Total Power (1)+(2) 84.0 124.0 113.0 115.9 137.0 98.4


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The trade-off analysis performed between the baseline solar powered system and that with MMRTG and SRG show an overall mass saving of about 13% and 9% respectively. The selection of radioisotope with less radiation compared to plutonium dioxide is not an attractive solution as the radiation levels are similar to that of plutonium and they do not show any considerable advantage over the former [8].

The present power level estimate (Table 1 and Fig. 4) shows six distinctly different modes of operation for the available power resource. The most favourable option after analysis is to have one unit of MMRTG for continuous power supply of 96.3 W (EOL) and one set of 2 x 10 Ah Li-Ion batteries for intermittent power demands. This system is selected after careful trade-off between using 2 units of SRG to meet the power demand and MMRTG combined with battery power system. The results show that MMRTG is the lowest mass option.

The alternate two SRG units system mass is about 75 kg compared to 60 kg of MMRTG battery combined power system. But the advantage of higher power resources of SRG is more comfortable and makes it a formidable substitute for the present proposed MMRTG option. The first SRG mission should carry a spare unit to become space qualified and JEO mission could fulfil the challenge as well as utilise the excess power produced. The two SRG units power system uses total of only four General Purpose Heat Source GPHS units to produce 2 x 110 W power compared to the MMRTG which uses 8 GPHS modules to produce only 110 W of power. The radiation levels of two SRG units are less than one MMRTG unit due to the employment of a total 4 GPHS modules.

0

10

20

30

40

50

60

70

80

90

100

110

120

130

140

150

Mode 1 - Launch Mode 2 - Orbitinsertion /correction

Mode 3 - Sciencewithout GPR

Mode 4 - Scienceonly GPR

Mode 5 - Scienceand Telecomwithout GPR

Mode 6 - Telecomonly

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Figure 4. Operating Modes and Power Level Estimates for JEO Spacecraft (Battery Charged during low

Power Demand)

VI. MMRTG / SRG and Radiation The MMRTG technology is the newest generation of radioisotope power system proposed by NASA. It uses the same technology used in a RTG but with a lower number of GPHS (8) to produce 110 W compared to conventional 285 W RTG (18). It also produces 1878 W heat which could be utilised for keeping the spacecraft within the operating temperature. The new system is also safer in case of unfortunate accident. MMRTG converts power using thermocouples. The one MMRTG unit on the JEO will emit beta, gamma and neutron radiation. The various total ionising dose (TID) presented in Fig. 5 have been calculated based on NASA dose rate estimation for GPHS units [9] which uses 5 GPHS modules. The document suggests it could be used to size different types of GPHS units by scaling. Shielding is aluminium and the radiation environments are based on a proposed Advanced Radioisotope Power Source (ARPS) with five General Purpose Heat Source (GPHS) modules repackaged from a Cassini RTG. The proposed power subsystem has one unit MMRTG each containing eight GPHS modules; therefore, the flux, fluence, and dose rate data presented in NASA’s Outer Planets Program, Environmental Requirements [9] must be


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scaled by a factor of 8/5. The Lambert’s law for linear attenuation states that the approximate amount of radiation that pass through the slab decrease exponentially with the increase in thickness, the intensity ‘I’ is found to decrease exponentially as the thickness ‘x’ of the absorbing material increases according to the equation (1),

xeII µ−= 0 (1)

1.00E+00

1.00E+01

1.00E+02

1.00E+03

1.00E+04

1.00E+05

1.00E+06

100 1000 10000

� ��

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Figure 5. Radiation Dose (Lifetime 8 years) Generated by Single unit of MMRTG / SRG versus Radial Distance from GPHS

Table 2 MMRTG / SRG Radiation dose (8 year mission)

Radioisotope Power DistanceSystem from each WithoutSingle MMRTG unit and RTG ShieldingTwin SRG unit Source (from Figure 5) 2.5 mm 4 mm 8 mm 10 mmMMRTG Total 300 mm 11900 9276 7989 5363 4394

SRG unit - I 300 mm 2930 2284 1967 1320 1082SRG unit - II 300 mm 2930 2284 1967 1320 1082SRG Total 5860 4568 3934 2640 2164

Using Lambert's equation (1)

Radiation Dose rad (Si)With Aluminium Shielding

The MMRTG total ionising dose for the JEO mission without shielding is about 12 krad. The TID behind 2.5 mm thick aluminium shielding is about 9.3 krad and this result is comparable with similar NASA baseline missions proposed in Standard RPS Mission Concepts and Application which is about 10 krad per MMRTG [10]. The SRG is another new technology for producing power. Stirling cycle devices have been flown into space as cooling system before but not as power source. The requirement for space qualification is to carry a redundant unit for the first unit to be flown on a space mission. SRG converts the heat from a GPHS module into reciprocating motion; with a linear alternator this reciprocating motion is converted into power. The unshielded radiation dose for a single unit of SRG is in the range of 3 krad and 2.3 krad behind 2.5 mm aluminium shielding for the proposed alternate option for JEO mission. The lower radiation makes SRG a good alternate power system for this mission.

VII. Radiation The entire Jovian minisat mission lifetime is about 8 years and it is totally controlled by the harsh radiation environment of Jupiter and Europa. The envisaged total radiation dose behind 10 mm aluminium shielding is in the


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order of 775 krad. The current strategy is to use radiation hardened electronics and shielding to protect the sensitive instruments. However with present technology we can achieve only 300 krad to 500 krad of radiation tolerance [7]. The alternate approach to the above problem in case it is not feasible to achieve the tolerance level of 1 Mrad for some instruments, is to provide additional shielding at increased mass. The proposed MMRTG mass savings could be utilised effectively to achieve the radiation level within the maximum wet mass allocated for the JEO spacecraft. The mass saving of using Tantalum instead of Aluminium is considerably higher for shielding thicknesses more than 4 mm and Tantalum should be used for higher required thicknesses [2]. The radiation shielding material mass allocated is about 27 kg for avionics box shielding and spot shielding of various instruments which are sensitive to radiation. Also the radiation would be attenuated effectively by placing the sensitive instruments within the propulsion unit where there is additional shielding thickness from the structure around it.

Table 3 JEO Total dose (8 year period)

Source 4 mm shielding

8 mm shielding

10 mm shielding

Jupiter tour 3170 krad 805 krad 350 krad 60 days around Europa 2100 krad 720 krad 420 krad Radiation from single MMRTG 8 krad 6 krad 5 krad Total 5278 krad 1531 krad 775 krad

VIII. Communication System The selected communication frequencies between JEO and Earth are 7 GHz (X-band) for telecommand (TC) and 32 GHz (Ka-band) for science. The communication frequencies between JEO and JRS are 8 GHz (X-band) for telecommand and 32 GHz (Ka-band) for science. The data communication system has been analysed for direct communication between Earth and JEO and between JEO and JRS to download science data and telecommand. The combination of different types and sizes of antenna have been considered and suitable optimum antenna diameters were obtained, a High Gain Antenna (HGA) diameter of 1.5 m and Medium Gain Antenna (MGA) diameter of 0.12 m for the JEO mission. This is the same as the ESA baseline proposed antenna sizes. The downlink data rates govern the entire communication system design. The communication between JEO and Earth is very much limited. The results show that 1.5 m HGA antenna can download up to 25 kbps directly to Earth if Deep Space Antenna (DSA) is pointed accurately to receive data from the JEO; however the limited window opening for communication means we can receive little science data with this link. But this link could be used to forward TC for the JEO mission. The JEO MGA direct communication with Earth DSA can transmit only 2 kbps of data which could be utilised for TC in case JEO MGA is used. The communication schematic between JEO and JRS is shown in Fig. 1. The spacecraft are in 3:1 resonant orbit which facilitates optimal communication window between them every 10.6 days. The HGA to HGA transmission data rate could vary between 20 Mbps at the nearest distance (235934 km) and 0.1 Mbps at the farthest distance (2573712 km). Between JEO MGA to JRS HGA it would be in the range of 6 Mbps and 50 kbps for the same distances. Between JEO MGA to JRS MGA it would be in the range of 70 kbps and 6 kbps for the same distances. However these data rates will rapidly deteriorate when the pointing accuracy reduces. The communication between JRS to Earth is achieved mostly through the HGA communicating with one of the 35 m Deep Space Antennas. The communication data rate between JRS HGA to DSA is about 87 kbps at the nearest distance (6x108 km) and 34 kbps at the farthest distance (9.6x108 km). The 34 kbps date rate is slightly higher than the ESA’s baseline rate of 30 kbps. The JRS MGA communicating with the DSA in case of JRS HGA failure is in the range of 6.9 kbps and 2.7 kbps for the same distances. The communication system has been designed from the data rate requirement and the overall mass of the subsystem is about 25 kg.

IX. Spacecraft Design The spacecraft is a box type structure with high strength and integrity; it is based on a similar structure to Mars Express [11]. It also makes the integration of JEO and JRS composite spacecraft easier. Figures 6 and 7 illustrate the JEO MMRTG spacecraft configuration. The propulsion system based on the propellant requirement gives results similar to the baseline results [1]. The system accommodates two 78 litre spherical Hydrazine tanks (EADS OST 31/0), one 85 litre N2O4 Oxidiser tank, one 0.31 m diameter helium pressure tank and 4 pairs of gimballed 22 N thrusters. The propulsion system mass is about 40 kg and the propellant mass is about 254 kg. The MMRTG has been located diagonally to minimise the radiation and direct heat load from it. The MMRTG location configuration


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is similar to that of the Pluto New Horizon mission [12]. The spacecraft attitude control and navigation propellant estimate is about 7 kg and the AOCS module mass is about 8 kg. The peak power level of 22.3 W is within the proposed power level of 23 W. The highly integrated avionics package consists of on board computer and other control devices assembled into shielded compact slots. The mass memory storage device needs to store about 40 Gigabit of science data and a 50 Gigabit of solid state memory system has been selected. The placement of thrusters next to the parabolic antenna is to facilitate attaching the JRS spacecraft with JEO as well as testing during the deep space transit. The avionics box and MMRTG are placed as far apart as possible to minimise any additional radiation. The payload instruments are placed in the best possible position for nadir pointing.

Figure 6. Jupiter Europa Orbiter spacecraft

660

mm

1552 mm

1552 mm

1200 mm

1010 mm


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Figure 7. JEO Ground or Ice penetrating radar deployed

X. Spacecraft Configuration with Two SRG units Power System The possible configuration of the JEO spacecraft with two SRG units is shown in Fig 8. The configuration is the same as that of the MMRTG system except that two SRG units have been located on the sides instead of one MMRTG unit. As the intermittent power could be supplied from the second SRG unit the battery system is not required in this configuration. The overall envelope size in launch configuration is 2.9 m (L) x 2.1 m (W) x 1.92 m (H).

Figure 8. JEO powered with twin SRG units

XI. JEO Spacecraft Trade-off Analysis The mass allocations of selected subsystems were evaluated for the most suitable radioisotope power system, as presented in Table 4.


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Table 4 JEO Spacecraft Trade-off Analysis Solar Power Vs Proposed RTG Technology

Proposed Power Subsystem Solar Power RTG I

1x285 WMMRTG II

1x110 WSRG III

2x110 W ARPS 9 W/kg

Technology Status JEO-Baseline [ESA]

Existing (only oneunit available)

Jeo-Proposed Under Development 2008

Spacecraft subsystemPower (conditioned) 106 80 60 75 30AOCS 8 8 8 8 8Propulsion 40 40 40 40 40CMDS 26 26 26 26 26Communications 25 25 25 25 25Structure 73 73 73 73 73Thermal 6 6 6 6 6Radiation shielding 27 27 27 27 27

Sub-total 311 285 265 280 235Margin 20% 62 57 53 56 47

JEO PLATFORM (as above) 373 342 318 336 282JEO SCIENCE INSTRUMENT 30 30 30 30 30JEO PENETRATOR 15 15 15 15 15

JEO Dry Mass 418 387 363 381 327

JEO PROPELLANT 291 270 254 266 229

Total - JEO Wet Mass 709 657 617 647 556

Saving in mass 7% 13% 9% 22%

Mass in kg

Future Technology 2010

I) GPHS RTG - General Purpose Heat Source RTG, Used in Galileo, Ulysses, Cassini, Pluto New Horizons.

Available power for each unit 285 W Beginning of Life (BOL); present inventory shows that there is only one unit available [10].

II) MMRTG - Multi-Mission RTG underdevelopment, Available power from each unit is 110 W (BOL) and 96.3 W at End of Mission (EOM) and thermal power of 2000 W (BOL) and 1878 W (EOM) [10].

III) SRG – Stirling Radioisotope Generator underdevelopment, it has not been flown into space before for power generation. It needs to be space qualified; to become space qualified the first SRG mission should carry a spare unit as well. Available electric power from each unit 115.8 W (BOL) and 101.6 W (EOM) and thermal power of 500 W (BOL) and 469 W (EOM) [10].

The results show that the JEO spacecraft with solar power system wet mass would be about 709 kg compared to the proposed MMRTG equivalent system which is about 617 kg only. The single unit MMRTG Jupiter Europa Orbiter provides continuous power output of 96.3 W at the End of Mission (EOM) with a mass saving of 13% compared to its solar power equivalent. The 8 units GPHS based power system produces uninterrupted power as opposed to a solar powered orbiter which depends on distance from Sun, solar flux, eclipse and Jupiter’s occultation. The innovative loop heat transfer thermal system utilises the excess heat from the MMRTG to keep the spacecraft within the required temperature level and it eliminates most of the electrical based thermal system mass from the spacecraft. The radiation dose from the MMRTG on the sensitive electronics is about 5 krad behind 10 mm aluminium shielding which is much lower than the natural radiation dose. This implies the initial system design of radiation hardened instruments with 1 Mrad tolerance is sufficient to withstand additional radiation dose from the MMRTG. The other alternate RPS power system with two SRG units overall mass saving is 9% compared to solar powered equivalent.


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XII. The Challenges for JEO The challenges of a Jupiter Europa orbiter are enormous and the most important requirement is the radiation hardened system (in the order of 1 Mrad) to enable operation in the harsh environment.

• The payload instruments should be developed with low power and mass, miniaturisation without compromising the accuracy and development of a highly integrated payload suite (HIPS).

• The spacecraft requires a high level of autonomy during all the mission phases, commissioning, operation and disposal of spacecraft.

• Thermal system design needs low mass and a high efficiency heat loop pipe system. • The entire mission must be developed in agreement with Committee on Space Research (COSPAR) Europa

environmental protection requirements on material selection, operational requirements and disposal of spacecraft after completion of mission.

The other challenges are listed in Table 5.

Table 5 Summary of Research and Development Area Challenges Requirement Target Technology Requirement USA Europe

RadiationRadiation hardening Protection against

> 3 Mrad1 Mrad Radiation hardened electronics up

to the level of 1 MradX-2000 Space component steering

board initiated programmes(Atmel AT 697 LEON beingdeveloped)

Power Generation MMRTG 110 W power per

unit (45 kg)110 W GPHS, Enclosure, Radiator MMRTG None

SRG 110 W power perunit (34 kg)

110 W SRG /ARPS

Facilities are available

Small RPS 10 to 20 watts 10 - 20 W RPS Small Research groups RHU (Radioisotope Heating Unit) 0.01 watts and

above0.01 W > RPS Small Research groups

ThermalMMRTG / SRG Heat exchangerfluid loop

within 6 kgincluding heaters

6 kg Reduction of mass and increase inefficiency

RTG Research facilities available

Venus flyby heat rejection efficient fluid loop Integration with MMRTG / SRG RTG New

PayloadHighly integrated payload suite Technology

development inThermal, Mechanical, pointing, accuracy, programming andintegration

Before assembly

Development of technology andnew programming techniques

X-2000 Programmes initiated

Deployment of radar Deployment mechanism

Before assembly

Able to deploy after 7.5 years N/A Programmes available

CommunicationsX and Ka band high data ratetransponder

3 Mbps 3 Mbps High data rate transponder JPL Programmes available

Ka band SSPA 3.5 W RF 30%efficiency

3.5 W low power high efficiency JPL Programmes available

CMDS (Command and Data Management System)Radiation hardening > 3 Mrad 1 Mrad Radiation tolerance X 2000 Programmes initiatedStorage RAM > 3 Mrad 1 Mrad Radiation tolerance X 2000 Programmes initiated

AOCS (Attitude and Orbit Control System)Accurate pointing of instruments 5 mrad 5 mrad High accuracy instrument X 2000 Programmes availableNew horizon sensor edge detection New detection

systemNew New type of detection system New New

AutonomyRobust autonomous system New New Highly autonomous X 2000 NewHigh radiation induced safe modes New New New N/A NewSoftware development New New New N/A NewOptical navigation New New New N/A NewManoeuvre scheduling New New New N/A NewSun sensor for lower solar flux New New New N/A New

Planetary protectionCleanliness COSPAR requirement New standards N/A NewInflight decontamination due toradiation

High radiationexperiments

New New standards N/A New

Penetrator ImpactDevelopment of material andsubsystem capable of withstandingvery high impact shocks (~10,000to 100,000 g)

New New New N/A New

Miniaturised Attitude control system New New New N/A New

Material compatibilityCompatibility of selected materialbehaviour for extreme radiationenvironment

New New New N/A New

Cleanliness requirement New New New N/A NewPlanetary protection confirmation COSPAR New New N/A New

Programmes


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XIII. Conclusion This Jupiter Europa Orbiter spacecraft design study demonstrates that an MMRTG is the most reliable alternative power source to the ESA baseline solar powered system. The total wet mass requirement for the MMRTG system is considerably lower than the baseline system. The one unit MMRTG could be replaced with two SRG units to have less radiation impact to the entire mission at a small increase in mass. The SRG powered system also gives additional power to operate all of the science instruments together unlike one MMRTG.

• The overall mass savings on JEO wet mass using a single MMRTG unit is 13% compared to the baseline solar powered spacecraft.

• The overall mass savings on JEO wet mass using two SRG units is 9% compared to the baseline solar powered spacecraft.

• The MMRTG or SRG total radiation dose and magnetic field are much lower than the Jovian natural radiation dose and magnetic field and don’t pose a threat to the mission. The stray magnetic field strength of MMRTG is only 17 nT, which is less than the requirement of 25 nT, at one meter distance [10].

• The MMRTG and SRG are the next generation of multi-mission radioisotope power systems expected to be available by 2009. But the baseline solar powered system needs technologies such as GaAs LILT to be developed to operate in Jovian environment.

• The additional power available from the redundant SRG unit could be well utilised for the additional power required to operate all science instruments. It is evident that in the event that the SRG become space qualified before the JEO mission, then a single SRG unit along with a battery powered system is sufficient enough to accomplish the mission objectives.

• The excess heat produced from the MMRTG or SRG will be absorbed by the innovative heat loop system and utilised inside the spacecraft where heating is required. The heat load produced during the Venus gravity assist flyby can be rapidly removed from the spacecraft using the same heat loop system to dissipate through the radiators.

• Power and mass is saved by the innovative thermal management system as it does not employ additional electrical heating system.

• The 3:1 resonant orbit of JEO and JRS facilitates communication between them without any gimballed antenna mechanism.

• The HIPS approach and strawman payload require modest mass and power resource. The miniaturisation reduces the harness as well as shielding material mass.

• The spacecraft configuration and placement of the MMRTG diagonally reduces the radiation dose without the need for an additional mounting platform.

• The thrusters are placed on the parabolic antenna platform and not obstruct the mating surface for the composite spacecraft. This also allows us to test the spacecraft’s thrusters during the deep space transit and these could be utilised for minor correction during navigation in addition to JRS thrusters.

• Tantalum radiation shielding for thickness more than 4 mm can reduce the shielding mass. • The complete autonomous system reduces the necessary contact time of the spacecraft from ground. • The selection of integrated thrusters for spacecraft propulsion and AOCS control eliminates separate

thrusters for each subsystem. The alternate RPS power system with two SRG units would require only four GPHS modules in total to produce 2 x 110 W (BOL) of electrical power compare to eight GPHS modules in the MMRTG to produce 1 x 110 W (BOL). The reduced use of GPHS modules emits less radiation and magnetic interference. The science instruments can operate with a single SRG in a similar way to MMRTG science mode phases. In addition to this the power of the redundant SRG unit could be utilised to achieve more science. This interesting technological development is particularly important in terms of environmental safety and handling of Plutonium. The SRG with less Plutonium at reduced cost is a significant development in using a SRG power system for future European space missions. Table 6 and 7 summarises the JEO mission with the MMRTG as the power source. This study concludes that a Jupiter Europa Orbiter spacecraft designed with single MMRTG unit or two SRG units as the power source would be the optimum choice for the proposed ESA Jovian Minisat Exploration mission.


13

Table 6 JEO Mass and Power Summary

Spacecraft Components Quantity Dimension m MMRTG

1x110 WSGR

2x110 W Idle Power

WAverage Power

WSpacecraft subsystem

Power (conditioned) 60 75 7 17.0MMRTG 110 We 1 0.66 (L) x 0.64 dia 45 SRG 110 We 2 1.04 x 0.38 x 0.29 68 Battery Li-ion 400 Whr 1 0.15 x 0.15 x 0.15 8 0 10.0Power conditioning & Misc Hardware 7 7 7.0

AOCS 8 8 14 30.0IMU 1 0.12x0.12x0.1 0.5 0.5 8.0Sun sensor 1 0.13x0.062x0.05 0.35 0.35 2.0Star tracker 1 0.12x0.12x0.11 2.4 2.4 4.0Horizon Sensor 1 0.2 x 0.13 1.1 1.1 4.0Reaction wheel 4 0.12 dia x 0.12 0.8 0.8 12.0Misc. 2.85 2.85

Propulsion 40 40 2.3 2.3EADS OST 31/0 - Hydrazine tanks 2 dia 0.58 12.8 12.8 0.0EADS OST 31/0 - MON3 tank 1 dia 0.62 7 7 22 N Bipropellant Thruster Model S22 - 02 8 0.055 dia x 0.212 length 5.2 5.2 Titanium Helium Pressure tank (2 kg of Helium) 1 dia 0.31 7.4 7.4 Piping, contol system and Misc components 7.6 7.6 2.3

CMDS 26 26 24 24.0

Communications 25 25 15.6 39.6X / Ka-Band HGA 1.5m diameter dish 1 1.5 x 1.4 depth 7.5 7.5 15.6X / Ka-Band MGA 0.12m horn 1 0.12 x 0.300 2.5 2.5 24.0Small Deep Space Transponder 1 0.174 x 0.134 x 0.141 2.95 2.95 InclKa- Band SSPA (KAPA) 1 0.142 x 0.152 0.66 0.66 InclDiplexer 1 - 0.2 0.2 InclDAM 2 - 0.8 0.8 InclWTS 2 - 0.78 0.78 InclMisc. hardware 9.61 9.61 Incl

Structure 73 73

Thermal 6 6 12.5

Radiation shielding 27 27

Sub-total 265 280Margin 20% 53 56 Incl

JEO PLATFORM (as above) 318 336

JEO SCIENCE INSTRUMENT [ESA] 30 30 7.3 43.7

Ground Penetrating Radar (ELRR / GPR) 1Stowed: 1.34 x 0.47 x 0.3Deployed: 10 x 2

4 4 2.524.0

Stereo Camera (EUSCam) 1 0.2 x 0.2 x 0.2 0.6 0.6 0.2 0.7Visible / Near IR spectrometer (EUVN-IMS) 1 0.095 x 0.122 x 0.217 1.8 1.8 0.2 1.0Radiometer (EuRad) 1 0.06 x 0.1 x 0.2 1.6 1.6 0.4 1.0Altimeter (EuLat) 1 0.1 x 0.1 x 0.3 2 2 2.0 6.0Magnetometer (EUMAG) 1 0.1 x 0.05 x 0.1 0.7 0.7 0.3 0.5Gamma and Neutron Spectrometer (EuGS) 1 dia 0.11 3.6 3.6 1.0 1.0UV camera (EuUVcam) 1 0.04 x 0.04 x 0.1 0.7 0.7 0.3 1.0Radiation monitor (EuREM) 1 0.1 x 0.05 x 0.1 0.5 0.5 0.4 1.0Boom 1 2.00 0.4 0.4 - DPU + CPS 1 - 2 2 3.4Shielding (20%) 4.7 4.7 - Structure 2.5 2.5 - Misc 1.8 1.8 4.2Margin (10%) 3.1 3.1

JEO PENETRATOR 15 15

JEO Dry Mass 363 381

JEO PROPELLANT 254 266

Total - JEO Wet Mass 617 647 70.2 169.1

Mass kg

JEO Equipment List


14

Table 7 JEO with MMRTG - Mission Summary

Science objectives [ESA]Determine the presence or absence of a subsurface ocean (including mapping of the ice thickness).Measure the global topography and tidal effects at Europa.Characterise the global geology and surface composition of Europa.Observe Europa’s magnetic field.Measure the radiation environment around Europa.

Payload [ESA] Ground (or) Ice penetrating radar (ELRR) Mapping thickness of ice layer, structure determination, topography, surface reflectivity

Stereo Camera (EuSCam) Topography, geology and surface compositionVisible – Near Infrared spectrometer (EuVN-IMS) Topography, geology and surface compositionEuropa radiometer (EuRad) Measuring Europa's surface temperatureEuropa Laser Altimeter (EuLat) Topographical mapping, study of tidal processesMagnetometer (EuMAG) Measuring Europa's magnetic fieldEuropa Gamma and Neutron Spectrometer (EuGS) Surface compositionUV camera (EuUVcam) Measuring the electron environment of EuropaStandard Radiation Monitor (EuREM) Analysis of Jovian radiation environment

Transfer [ESA] Soyuz Fregat 2-1b launch from Kourou2 Spacecraft composite transfer to Jupiter via a Venus-Earth-Earth GATransfer duration ~ 6 yearsAfter Jupiter Orbiter Insertion the S/C separate and perform a tour of the Jovian systemJEO will achieve orbit around Europa in 545 daysJRS will achieve a highly elliptical orbit around Jupiter (~ 20 inclination w.r.t. equator) 449 days

Operational Orbit [ESA]

Mission Lifetime [ESA] 6+1.5 years until Europa orbit insertion~ 60 days of science operations

Spacecraft Details

Stabilisation 3 axisOrientation Nadir / JRSSize overall (mm) 3200 (length) x 2300 (width) x 1920 (height)Mass Mass figures include 5-20% component margin (depending on maturity) and 20% system

margin.

Payload (kg) 30Dry (kg) 363Wet (kg) 617 Margin w.r.t. launcher (kg) 39

Radiation ~ 775 krad (10mm Al shielding)Power (incl. margins) (W) 169TM band Ka, XAntenna 1.5 m HGA, 0.12 m MGAData storage (Gbit) 50Payload power (W) 25Payload data rate (kbps) 40

JEO Mission Summary

Primary Objectives

200 km circular polar orbit, period = 2.3 hours

�

References 1Renard, P., Koeck, C., Kemble, S., Atzei, A., and Falkner, P., “System Concepts and Enabling Technologies for an ESA Low-Cost Mission to Jupiter / Europa,” Proceedings of 55th International Astronautical Congress, Vancouver, Canada, 2004, IAC-04-Q.2.a.02, pp. 5-8.

2Atzei, A., and Falkner, P., “Study overview of the Jovian Minisat Explorer TRS,” An ESA Technology Reference Study SCI-AP/2004/TN-085/AA, issue 3, revision 2, 22nd March 2005, pp. 10, 22, 24. 3Scheeres, D.J., and Guman, M.D., “Stability Analysis of the Europa Orbiter,” AAS/AIAA Spaceflight Mechanics Meeting, Clearwater, Florida, Paper AAS 00-154, 2000, pp.11. 4Falkner, P., “Science Future Programme Technologies,” Workshop on Spacecraft Data Systems, ESTEC, Noordwijk, The Netherlands, 5-7th May 2003, pp. 13-23. 5NASA, “Europa Orbiter Mission and Project Description, The Outer Planet Program Announcement of Opportunity,” April 1999. URL: http://centauri.larc.nasa.gov/outerplanets/Europa_MPD.pdf [cited 26th June 2005], pp. 35


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6Abelson, R.D., (ed.), “Enabling Exploration with Small Radioisotope Power Systems,” NASA Office of Space Science, Jet Propulsion Laboratory, California Institute of Technology, JPL Pub 04-10, September 2004, chap. 2 pp. 23. 7Niebur, C., and Balint, T.S., “Europa Surface Science Package Feasibility Assessment,” Outer Planets Advanced Studies, Jet Propulsion Laboratory, California Institute of Technology, JPL D-30050, 22nd September 2004, pp. 8. 8Ulrich, P.B., “Final Environmental Impact Statement for Cassini Mission, Alternatives, Including the Proposed Action,” Solar System Exploration Division, Office of Space Science, NASA Headquarters, Washington DC, June 1995, chap. 2 pp. 51-58. 9NASA, “Outer Planets Program, Environmental Requirements,” August 1999. URL: http://centauri.larc.nasa.gov/outerplanets/Envir_rqts.pdf [cited 30th July 2005] pp. 25-32. 10Abelson, R.D. (ed.), “Expanding Frontiers with Standard Radioisotope Power Systems,” NASA Science Mission Directorate, Jet Propulsion Laboratory, California Institute of Technology, JPL D-28902, PP-266 0332, 12th January 2005, chap. 2 pp.72, chap. 3 pp.2, 5, 7. 11Schmidt, R., Credland, J.D., Chicarro, A., and Moulinier, Ph., “ESA’s Mars Express Mission – Europe on Its Way to Mars,” ESA Bulletin 98, June 1999. 12Lindstrom, K., “Draft Environmental Impact Statement for The New Horizons Mission,” Science Mission Directorate, NASA, Washington, DC, February 2005, chap. 2 pp. 5.

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