23 Fi g Fi g T h D a ta r o p in t in c pl a fr o 0 g . 10. Power prof i g . 11. Mass and c o h e extreme exa m a wn” class missi rg ets with a sin gl p ens up the trade t erest tar g ets are c lination, eccentr a nnin g could all o o m several secon d i le for the Dawn m o st comparison f o m ple of a multip l on is the conce p l e launch. The t h space of achieva b not necessaril y c icit y , period, etc o w missions to v d ar y near-b y tar g m ission (Oh, 200 7 o r the Dawn mis s l e tar g et missio n p t of travelin g t o h rou g hput poten t b le multi-rendez v o-located to allo w . Sufficient thro u v isit multiple hi g g ets. Aerospace Te c 7 ). s ion (Reference). n would be a S u o and stoppin g a t ial of both NEX T v ous options. U n w for short trans f ug hput capabilit g h interest tar g et s c hnologies Advanc u per-Dawn. A “ S a t several hi g h i n T and HiVHAC g n fortunatel y , mo s f ers due to varia n ty and creative m s and g ain infor m ements S uper- n terest g reatl y s t hi g h n ces in m ission m ation Low-thrust Propulsion Technologies, Mission Design, and Application 231 The targets for a hypothetical “Super-Dawn” mission were chosen from a list of high interest targets formulated by the scientific community. Based on preliminary analysis of throughput requirements and delivered mass, a single spacecraft, with only a 5-kW array, could be used to rendezvous with four high interest near-Earth targets shown in table 1. The final delivered mass is comparable to the Dawn spacecraft. The “Super-Dawn” mission illustrates the tremendous potential of electric propulsion for these types of missions. Studies have looked at using a single spacecraft for tours of near-Earth objects, main-belt asteroids, and even Jupiter Trojans. Sample return missions are multi-body missions because they need to return to Earth. Sample return missions are often considered high priority because of the higher fidelity science that can be performed terrestrially. Mars sample return was under investigation for many years, but the large costs of such a mission has deterred its implementation. Regolith from Phobos and Deimos are of high scientific value. The mission options offer significantly lower cost with minimal technology development required. Segment Target Start Mass, kg Propellant Required, kg End Mass, kg 1 Nereus 1650 309 1341 2 1993 BD3 1341 52 1289 3 Belenus 1289 44 1245 4 1996 FG3 1245 456 789 Table 1. Table of ΔV for a “Super-Dawn” type mission. Two concepts for a Phobos and Deimos sample return mission were evaluated using solar electric propulsion: a single spacecraft to both moons or twin spacecraft capable of returning samples from either moon. The small bodies of Phobos and Deimos, with small gravity fields (especially Deimos), make electric propulsion rendezvous and sample return missions attractive. Electric propulsion systems can be used for the transfer to Mars, and then to spiral into an orbit around the moons. Chemical systems cannot easily leverage the Oberth effect for the sample return mission from Mars‘ moons because of the higher altitude orbit requirement. So while the mission can be completed, it comes at a large mass penalty. Figure 12 illustrates the benefits of using electric propulsion for a Phobos and Deimos sample return mission. Results show significant savings for using electric propulsion for Phobos and Deimos sample return missions. The baseline case uses a NEXT thruster with one operating thruster, and a spare system for redundancy (1+1). A Delta II class launch vehicle is capable of delivering enough mass for a sample return from both targets. For electric propulsion, the transfer between Phobos and Deimos has minimal mass implications. The mass and technology requirements could potentially fit within the Mars Scout cost cap. Using an Evolved Expendable Launch Vehicle (EELV), twin electric propulsion vehicles can be sent for a low-risk approach of collecting samples from Phobos and Deimos independently. However, the use of an EELV enables a chemical solution for a sample return mission. Going to a single moon chemically remains a significant challenge and results in a spacecraft that is greater than 70 percent propellant; a mass fraction more typical of a launch vehicle stage. Launching a single chemically propelled spacecraft to retrieve samples from both moons requires staging events adding risk and complexity. Aerospace Technologies Advancements 232 Fig. 12. Comparison of required launch mass for chemical and EP Mars’ moons missions. The use of electric propulsion was studied for various comet surface sample return (CSSR) missions. The results are highly dependant on the targets of interest. Electric propulsion compares favorably with chemical alternatives resulting in either higher performance or reduced trip times. Studies for Temple 1 (Woo et al., 2006) determined the SOA NSTAR thruster to be inadequate due to its propellant throughput capability. The mission required the use of a NEXT thruster. Studies for the comet Wirtanen (Witzberger, 2006) were conducted and determined that the NSTAR could not deliver positive payload while both the NEXT and HiVHAC thrusters can complete the mission with sufficient margin. The largest benefit is that electric propulsion enables a wide range of targets that cannot be reached using chemical propulsion systems. In 2008, NASA GRC completed a mission design study for a multiple near-Earth asteroid sample return mission (Oleson et al., 2009). The results indicated that it is feasible to use electric propulsion to collect multiple samples from two distinct targets in very different orbits. An Earth fly-by was performed after leaving the primary target and before arriving at the second to releae the sample return capsule for a lower risk mission and mass savings to the secondary target. This mission was not feasible using chemical propulsion. The conceptual spacecraft for the multi-asteroid sample return mission is shown in figure 13. 4.2 Inclined targets Other missions enabled by electric propulsion are missions to highly inclined targets. There are several Earth crossing targets that are thought to be old and inactive comets. These asteroids typically have inclined orbits. The ∆V requirement for a plane change is a function of the spacecraft velocity and angle of the plane change as shown in equation 1. With the Earth’s heliocentric orbital speed near 30 km/s, a simple plane change of even 30 degrees will require a ∆V of at least 15 km/s to perform a fly-by, following equation 4. Lo w Fi g A n E a in c hi g lo w 14 th e Fi g w -thrust Propulsio n g . 13. NEARER s p n example missi o a rth-crossin g bo d c lination of 64º. B g h launch veloci t w er velocit y . Fi gu km/s. The Tsio l e mission is com p g . 14. Optimal ch e n Technologies, Mi s p acecraft photo. o n to an inclined dy with a semi- m B ecause of the in c ty so that the sp a u re 14 illustrates t l kovsk y ’s mass f r p letel y infeasible e mical tra j ector y s sion Design, and A tar g et would be m a j or axis of 1. 2 c lination chan g e, a cecraft can perf o t he chemical tra n r action is onl y o n with an y launch to Tantalus. A pplication to the asteroid T 2 9 AU, an ecce n the optimal che m o rm the plane ch a n sfer which requi n the order of o n vehicle usin g ch e T antalus. Tantal u n tricit y of 0.3, a m ical transfer req u a n g e at hi g h AU res a post-launc h n e percent dr y m e mical propulsio n 233 u s is an a nd an u ires a with a h ∆V of ass, so n . Aerospace Technologies Advancements 234 The electric propulsion transfer to Tantalus is also a challenging mission. The low-thrust transfer is over 30 km/s over 4.5 years, but can still deliver over 800 kg of dry mass on a rendezvous mission using an Atlas V. The mission would require two NEXT thrusters, and would not be viable with the NSTAR or Hall thruster based propulsion system. Rather than going to high AU to perform the plane change, the low-thrust transfer gradually performs the plan change through several revolutions. Figure 15 illustrates the low-thrust transfer to Tantalus. Because of the advantages of electric propulsion, efficient use of propellant and low-thrust trajectory options, scientists can plan missions to high interest targets previously unattainable. Fig. 15. Optimal low-thrust trajectory to Tantalus. 4.3 Radioisotope electric propulsion Another area of interest pushing the limits of propulsion technology is the use of a radioisotope power source with an electric propulsion thruster. This achieves high post launch ∆V on deep-space missions with limited solar power. Radioisotope electric propulsion systems (REPS) have significant potential for deep-space rendezvous that is not possible using conventional propulsion options. One example of mission that can benifit from REPS is a Centaur orbiter. The Centaurs are of significant scientific interest, and recommended by the Decadal Survey Primitive Bodies Panel as a New Frontiers mission for reconnaissance of the Trojans and Centaurs. The original recommendation was for a flyby of a Jupiter Trojan and Centaur. While a flyby mission can use imaging, imaging spectroscopy, and radio science for a glimpse at these objects, a REP mission provides an opportunity to orbit and potentially land on a Centaur. This greatly increases the science return. An exhaustive search of Centair obiter missions concluded that a wide range of Trojan flybys with Centaur Rendezvous missions are pracitical with near-term electric propulsion technology and a Stirling radioisotope generator (Dankanich & Oleson, 2008). With near-term technology, flyby missions may no longer be scientifically acceptable. Investigations are continuing using the enabling combination of electric propuslion and radioisotope power systems. On-going and recent studies include multi-Trojan landers, Kuiper-belt object rendezvous, Titan-to-Enceldaus Lo w tr a sh o Fi g 5. T h sa t al t m o be t co m A us e tr e p o tr e co m Se v st a to co m ap si g ex p ti m K e p o in w -thrust Propulsio n a nsfer, and Nep t o wn in fi g ure 16. g . 16. Centaur or b Nea r -Earth ap h ere are currentl y t ellites have mul t itude, two from o on. The vast ma j t ween low-Eart h m mercial satellit e con j unction of d e of electric prop e nds of satellite p o wer technolo gy , e nds of both i n m munication sat e v eral studies w e a rtin g at LEO. A levera g e the la u m mercial launc h o g ee altitudes o f g nificantl y reduc e p osure to the r a m es and ∆V re q e nned y Space Ce n o wer-to-mass rati fi g ures 18 and n Technologies, Mi s t une Orbiters. A b iter spacecraft d e plication of el e y over 200 satell i tiple thrusters. O NASA, one fro m j orit y of mission h orbit (LEO) a e s with GEO ope r d evelopments g r e ulsion for Earth- o p ower-to-mass r a and the broad u n creasin g power - e llites are shown e re performed t o stron g deterrent u nch vehicle to h vehicles is to a f 185 km and 3 5 e the orbit trans a diation environ m q uirements fro m n ter (KSC)) wer e os (Dankanich & 19 respectivel y. s sion Design, and A A conceptual sp a e si g n. e ctric propulsi i tes usin g electri c O nl y four spacec r m JAXA to a nea r pull of electric p r a nd g eos y nchr o r ational orbits. e atl y increased t h o rbit transportat i a tios, required s p u se of electric pr o - to-mass ratios in fi g ure 17 (B ye o evaluate the u is the ver y lon g an eccentric o r g eos y nchronou s 5 ,786 km respec t fer time while a m ent within the m equatorial lau n e evaluated and c & Woodcock, 200 6 . Above 3 W/k g A pplication a cecraft desi g n o on c propulsion an d r aft flew be y on d r -Earth ob j ect, a n r opulsion is for a p o nous Earth or b h e practicalit y a n i on. Ke y amon g t p acecraft mass, a o pulsion s y stem s and spacecraft e rs & Dankanich, u se of electric p transfer times. O r bit. The primar y transfer orbit ( G t ivel y . Launchin g a lso reducin g ris k Van Allen belts n ch sites (Baiko n c haracterized as a 6 ). Results for K S g , the accelerati o f a Centuar or b d the ma j orit y o f d g eos y nchronou n d one from ESA p plication in the b it (GEO,) spec i n d expectations f t hese advanceme n a dvancements i n s . An illustration mass for com m 2008). p ropulsion for t r O ne near-term o p y desi g n capabi G TO) with peri ge g directl y to GT O k of orbital deb r . Anal y ses for t r n ur, Kourou, a n a function of spa S C launches are s on is hi g h eno u 235 b iter is f those s orbit to the re g ion i ficall y f or the n ts are n space of the m ercial r ansfer p tion is lit y of e e and O will r is and r ansfer n d the cecraft s hown ug h to Aerospace Technologies Advancements 236 minimize gravity losses of the transfer. The low-thrust ∆Vs are approximately 800 m/s more than a chemical GEO insertion. Fig. 17. Trends of commercial satellite beginning-of-life (BOL) P/M ratio and average mass. Fig. 18. GTO-to-GEO transfer times as a function of spacecraft specific power. Low-thrust Propulsion Technologies, Mission Design, and Application 237 Fig. 19. Required ΔV from GTO-to-GEO as a function of spacecraft specific power. The GTO-to-GEO transfer time and ∆V is dependant on the launch site, or initial starting inclination. Figure 20 illustrates the penalty of launch at inclined launch sites and the benefit of near-equatorial launches. Fig. 20. Effect of starting inclination on transfer time and ΔV from GTO-to-GEO. Aerospace Technologies Advancements 238 There were 32 commercial communication satellites launched in 2005 and 2006 as provided by the Union of Concerned Scientists database. These specific satellites were evaluated for potential to use an integrated electric propulsion system with a specific impulse of 1000 seconds, 1500 seconds, and 2100 seconds. Integrated electric propulsion systems assume the use of 95% of the onboard solar array power of the spacecraft as launched. Using electric propulsion for the GEO insertion has significant mass benefits. Typically this is evaluated as a method to leverage the launch vehicle performance to deliver the greatest possible mass. Another perspective is to evaluate the potential for existing launch vehicles to meet the demands of the COMSAT market. Figure 21 illustrates that currently launch vehicles with GTO drop mass capabilities in excess of 7,500kg are required for a complete market capture. However, using electric propulsion, a launch vehicle with a drop mass capability of 5,500kg can have complete market capture. A low cost launcher with a capability to deliver 3,500kg to 5,500kg can create a paradigm shift in the commercial launch market. This assumes the commercial entity is willing to endure the long transfer time, ranging from 66–238 days, depending on the spacecraft power-to-mass ratio and EP thruster selected. Fig. 21. Capture fraction as a function of GTO drop mass for various propulsion options. 6. Conclusion Electric propulsion technology is widely used today, and multiple thrusters exist for primary electric propulsion application. NASA and the U.S. commercial market developed several thrusters suitable for primary electric propulsion on full scale spacecraft. The Low-thrust Propulsion Technologies, Mission Design, and Application 239 technology drivers for new electric propulsion thrusters include: ability to use available power (i.e. high maximum power with large throttle range), increased total throughput capability, and lower cost systems and integration. The optimal specific impulse is limited by thrust required to minimize propulsive inefficiencies and available power. Due to power constraints, the optimal specific impulse is typically less than 5,000s and closer to 2,000s for near-Earth application. Electric propulsion is an enabling technology for a large suite of interplanetary missions. Several targets are infeasible with advanced chemical propulsion technologies, while practical with today’s electric propulsion options. Electric propulsion is well suited for missions with very high post-launch ∆Vs including multi-target missions, sample return missions, deep-space rendezvous, and highly inclined targets. Electric propulsion has tremendous capability to impact the commercial launch market by leveraging on-board available power. Today’s commercial satellites have mass-to-power ratios for practical GTO-to-GEO low-thrust transfer. As available power and performance demand continues to rise, electric propulsion technologies will continue to supplant chemical alternatives for a wide range of missions. The technology will continue to focus on developing lower cost propulsion systems with higher power and longer lifetime capabilities. 7. References Brophy, J. R. (2007). Propellant Throughput Capability of the Dawn ion Thrusters, IEPC- 2007-279, 30th International Electric Propulsion Conference, Florence, Italy, September 2007. Brophy, J., Rayman, M. D., & Pavri, B. (2008). Dawn: An Ion-propellanted Jounrey to the Beginning of the Solar System, IEEE Aerospace Conference, Big Sky, MT, March 2008. Byers, D., & Dankanich, J. W. (2008). Geosynchronous-Earth-Orbit Communication Satellite Deliveries with Integrated Electric Propulsion. Journal of Power and Propulsion, Vol. 24, No. 6, November–December 2008, pp 1369–1375. Dankanich, J. W. & Oleson, S. R. (2008). Radioisotope Electric Propulsion (REP) Centaur Orbiter Mission Design, 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Hartford, CT, July 2008. Dankanich, J. W., & Woodcock, G. R. (2007). Electric Propulsion Performance from GEO- Transfer to Geosynchronous Orbits, International Electric Propulsion Conference, Florence, Italy, September 2007. Kamhawi, H., Manzella, D., Pinero, L., & Mathers, A. (2009). Overview of the High Voltage Hall Accelerator Project, AIAA 2009-5282, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Denver, CO, August 2009. Manzella, D. (2007). Low Cost Electric Propulsion Thruster for Deep Space Robotic Missions, 2007 NASA Science Technology Conference, University of Maryland, MD, June 2007. Oh, D. (2007). Evaluation of Solar Electric Propulsion Technologies for Discovery-Class Missions. Journal of Spacecraft and Rockets, Vol. 44, No. 2., March-April 2007, pp 399– 411. [...]... within 180 minutes data latency requirement in order to be ingested by the operational weather forecast model (Fong et al., 2009b) 4 FORMOSAT-3 satellite design Figure 3 illustrates the FORMOSAT-3 satellite designed by Orbital Science Corporation in a deployed configuration and its major components The FORMOSAT-3 satellite avionic block 246 Aerospace Technologies Advancements diagram is shown in Figure... diameter, 18 cm in height 80 0 km altitude, circular 72o 52.5o apart ~ 81 W orbit average S-band uplink (32 kbps) and downlink (2 Mbps) 1,600 ~ 2,400 soundings per day 15 minutes to 3 hours 5 years April 15, 2006 Table 2 The FORMOSAT-3 mission characteristics 3.2 System architecture Figure 2 shows the FORMOSAT-3 system architecture After two years’ in orbit operations, starting from mid-April 20 08, the... cyclone/typhoon/hurricane forecasts (Kuo et al., 2004; Anthes et al., 20 08) The mission results have shown that the RO data from FORMOSAT-3 are of better quality than those from previous missions and penetrate much further down into the troposphere, mission results could be referred to Liou et al (2007), Anthes et al (20 08) , Fong et al (2008a, 2008b, 2008c & 2009a), and Huang et al (2009) In the near future, other... et al., 19 78; Yakovlev et al., 1996) 1 http://en.wikipedia.org/wiki/Occultation [cited 1 July 2009] 242 Aerospace Technologies Advancements 2 GNSS radio occultation mission After GNSS becomes operational, substantial and significant progress has been made in the science and technology of ground-based and space-based GNSS atmospheric remote sensing over the past decade (Davis et al., 1 985 ) The ground-based... neutral atmosphere, and electron density in the ionosphere with global coverage (Anthes et al., 2000 & 20 08; Liou et al., 2006a, 2006b, & 2007; Fong et al., 2008a & 2009a) In this chapter the FORMOSAT-3/COSMIC mission was referred to as the FORMOSAT-3 mission for simplicity 2 244 Aerospace Technologies Advancements The retrieved RO weather data are being assimilated into the NWP models by many major weather... IEEE Transactions on Geoscience and Remote Sensing, Vol 46, No.11, Nov 20 08, pp 3 380 -3394 doi:10.1109/TGRS.20 08. 2005203 Fong, C.-J.; Huang, C.-Y.; Chu, V.; Yen, N.; Kuo, Y.-H.; Liou, Y.-A.; & Chi, S (2008c) Mission results from FORMOSAT-3/COSMIC constellation system AIAA Journal of Spacecraft and Rockets, Vol 45, No 6, Nov.-Dec 20 08, pp 1293-1302 doi:10.2514/1.34427 Fong, C.-J.; Yen, N L.; Chu, C.-H.;... Applications to the Asia-Pacific Region Liou, K.-N., Chou, M.-D (eds), pp 4 58- 483 , World Scientific Publishing, Singapore Kuo, Y.-H.; Liu, H.; Ma, Z & Guo, Y.-R (2008b) The impact of FORMOSAT-3/COSMIC GPS radio occultation Proceedings of 4th Asian Space Conference and 20 08 FORMOSAT-3/COSMIC International Workshop, Taipei, Taiwan, 1-3 Oct 20 08, NSPO, Hsinchu, Taiwan Kliore, A J.; Cain, D L.; Levy, G S.; Eschleman,... frequency fluctuations as observed in radio occultation experiments on the satellite-to-satellite link, Journal of Communications Technology and Electronics, Vol 41, No 11, pp 9939 98, Nov 1996 2 58 Aerospace Technologies Advancements Yen, N L & Fong, C.-J., ed (2009) FORMOSAT-3 Evaluation Report and Follow-on Mission Plan, NSPO-RPT-0047_0000, 10 May 2009, National Space Organization (NSPO), Hsinchu,... numerical results on the nozzle ablation it suffice to put β = 0.2 and to obtain values of vm,t and vm,ex by regression 264 Aerospace Technologies Advancements Fig 2 Sketch of a cross-section of an idealized geometry of the multi-segment RSRMV rocket and an example of the design curves (8) for the head section 2.4 Model of the burning-though of a hole To complete the model of the case breach fault for the... C.; Schreiner, W S.; Sokolovskiy, S V.; Syndergaard, S ; Thompson, D C.; Trenberth, K E.; Wee, T K.; Yen, N L.; & Zeng, Z (20 08) The COSMIC/FORMOSAT-3 Mission: Early Results Bulletin of the American Meteorological Society, Vol 89 , No.3, Mar 20 08, pp 313-333 doi:10.1175/BAMS -89 -3-313 Bonnedal, M (2009) RUAG GNSS Receivers and Antennas, Proceedings of Global Navigation Satellite System Radio Occultation . Occultation Receiver (GOX) Power TBB TIP +28V from MIU Payload Bus TO MIU TICE TIP Sensor Assembly (TSA) +28V from MI U +28V from MIU To PCM Downlink Uplink Cont Loop Cosine Sensors [7 -8] Ant. +28V from BCR +28V to MIU Fwd. Occultation Receiver (GOX) Power TBB TIP +28V from MIU Payload Bus TO MIU TICE TIP Sensor Assembly (TSA) +28V from MI U +28V from MIU To PCM Downlink Uplink Cont Loop Cosine Sensors [7 -8] Ant. +28V from BCR +28V to MIU Fwd. ug h to Aerospace Technologies Advancements 236 minimize gravity losses of the transfer. The low-thrust ∆Vs are approximately 80 0 m/s more than a chemical GEO insertion. Fig. 17.