The status of electric propulsion development and applications in Russia

Available online at www.sciencedirect.com

Acta Astronautica 54 (2003) 25 – 37

www.elsevier.com/locate/actaastro

The status of electric propulsion development and applications in Russia
A.S. Koroteeva , O.A. Gorshkova;∗ , V.N. Akimova , A.A. Sinitsina , V.M. Murashkob , B.A. Arkhipovb , V.N. Vinogradovb a Keldysh b EDB

Research Center, Onezhskaya 8, Moscow 125438, Russia “Fakel”, Moskovskiy pr., 181, Kaliningrad 236001, Russia

Received 29 April 2002; received in revised form 30 August 2002; accepted 9 October 2002

Abstract Modern state of development and application of electric propulsion (EP) in Russia is analyzed in this report. The characteristics of 3elds of advisable EPS application are presented. The power-to-weight and economic e5ectiveness of applying EP systems of di5erent types for solving perspective tasks, including spacecraft transportation to operating orbits, its station keeping in the orbit and in the structure of satellite constellations, operation and stabilization and taking away from the orbit after the end of the lifetime are studied in this paper. c 2003 Elsevier Ltd. All rights reserved. 

1. Introduction Development of space-rocket technology is accompanied by wide electric propulsion (EP) application and tasks, realized with it. Long experience has proved high technical–economical e5ectiveness of EP application to realize orbit adjustment of spacecraft (SC) of di5erent purposes, and 3rst of all, geostationary SC. EP has practically no alternative for systems of




Based an paper IAF-01-S, E4.02 presented at the 52nd International Astronautical Congress, 1–5 October 2001, Toulouse, France. ∗ Corresponding author. Tel.: +7-095-456-6465; fax: +7-095456-8228. E-mail addresses: [email protected] (A.S. Koroteev), [email protected] (O.A. Gorshkov), [email protected] (V.M. Murashko).

orbit adjustment of geostationary communication spacecraft with on-orbit life (OL) of 15 years and above. Last years are marked by the beginning of EP application to realize the transport operations both in near-earth space and in interplanetary missions (transfer of SC based on American platform BS-702 to geostationary orbit (GEO) by ion thrusters, interplanetary SC “Deep Space-1,” Russian project “Phobos-Grunt” and other projects [1]). The perspectives of wide EP application for providing transport operations in near-earth space are caused not only by high speci3c impulse of EP. They are also determined by trends of space systems development (3rst of all, space communication systems) and the stated present preconditions for realization of unconventional ways of SC transfer to the operation orbit,

c 2003 Elsevier Ltd. All rights reserved. 0094-5765/03/$ - see front matter
doi:10.1016/S0094-5765(02)00279-5

26

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

using EP: • Increase of mass and OL of geostationary SC, and wide application of EP in their structure (satellite platforms with mass at GEO of 4 –5 tons and above with OL of 15 years and above are developed). • Noticeable increase of solar batteries speci3c characteristics and increase of available power of SC up to 4 –5 kW=t (electrical power of power supply systems at geostationary SC of new generation runs up to 15 –20 kW). • Development of multisatellite low-orbital communication systems on the base of small spacecraft of new generation. • Wide application in the world practice of SC transfer to high operation orbits with onboard “apogee” propulsion system (PS) based on low thrust liquid or solid rocket engines (LRE or SRE). • Basing of high ballistic e5ectiveness of SC transfer to GEO with EPS, using schemes with high-elliptical mediate orbits. • Wide range of developed EP with di5erent power of individual thrust modules. EPS application allows increase of SC mass transferred to GEO by launch vehicles of medium and heavy classes in 1.5 –2.0 times [2–4]. Besides, there is a possibility to transfer small SC to GEO by LV of light class (for example, Russian SC “Dialog” and “Ruslan-MM” with mass of 500 kg, developed by Khrunichev State Scienti3c and Production Space Center and Scienti3c and Production Association for Machine Building (NPO MASH), accordingly; LV “Rokot” and “Strela” [5,6]). EP can be widely applied to SC of multi-satellite low orbital and medium-altitude communication systems (orbit raising, maintaining of orbital formation structure and deorbiting after OL is 3nished) and also to low-orbit remote-sensing SC (aerodynamic drag compensation). In the near-term outlook, basing on nuclear power plants with power of 100 –150 kW together with EPS of the same power, it will be possible to transfer the SC with mass up to 10 tons and high level of available power to geostationary and geosynchronous orbits by heavy class LV. Such SC will realize power-consuming tasks like high resolution radar probing.

In further outlook without EP it will be impossible to realize large-scale projects in near-earth and interplanetary space, connected with solution of global energy and ecological problems, large-scale Moon developing, realization of the program of manned missions to Mars. Such projects are based on powerful solar or nuclear power plants of megawatt level and it is supposed to apply EP to solve power-consuming transportation tasks. Particularly, in present considered concepts of Mars expedition complex, based on EP systems (EPS) the required power of the power plant is 6 –15 MW. The primary EPS of the complex can be developed as the assembly of thrust modules (plasma or ion) with module power of 50 –100 kW and speci3c impulse up to 8000 s. So, EP, having passed the long way of development and taken the strong stands in space-rocket technology, can (and should) be the main thruster type, which provides transport operations in interplanetary and near-earth space. At the same time, EP application in transport operations in near-earth space, what provides high energy-ballistic e5ectiveness (SC mass increase transferred to operating orbit), is connected with noticeable time increase for transport operation realization. In this connection, it is necessary to conduct complex analysis of technical–economical e5ectiveness of EP application in transport operations. The estimations and analysis results of ballistic and economical e5ectiveness of EP application to transfer geostationary communication, low-orbit communication or remote sensing SC to operating orbits are presented below. 2. Developed thrusters and plans of their application Development and application of Hall-e5ect thrusters (HET) is the basis of the program of work in Russia in the 3eld of EPS. During the last time the range of e5ective operation of this class thrusters has been noticeably increased and is from tens of watts to tens of kilowatts. Some of developed thrusters with their manufacturers are presented in Table 1. The state of development and perspectives of their application are reported in [7].

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

27

Table 1 Russian HET Power (kW) ¡ 0:5

0.5 –5.0

5.0 –10.0

10.0 –50.0

DB “Fakel” and RIAME MAI

SPT-25 SPT-35 SPT-50

SPT-70 SPT-100 PPS-1350a SPT-140

SPT-200

SPT-290

Keldysh Research Center

KM-37 KM-45

X-85M ROS-99 ROS-2000

T-160E

TsNIIMASH

D-27 D-38

D-55 D-80

D-100 II D-150

a Together

TM-50

with SNECMA.

HET of medium power (0.5 –5:0 kW) of DB “Fakel” and Keldysh Research Center (KeRC) are considered in detail in this paper. Thrusters of this class are the most studied. Thrusters, regularly used in space, and thrusters, developed for the concrete tasks, belong to this power range. Thrusters of DB “Fakel” SPT-70 and SPT-100 are manufactured serially and are used in space at some SC. SPT-70 with nominal thrust 40 mN has been used in space since 1982 for East–West station keeping of geostationary satellites. Already 15 spacecraft with 68 thrusters onboard have been put into orbit. Total operating time of these thrusters in space is more than 4500 h. Since 1994, 48 SPT-100 with nominal power of 1:35 kW have been operated in space at Russian geostationary SC. EPS, based on SPT-100, provide SC positioning in the orbit, and also East–West and North–South (E–W & N–S) station keeping. Total operating time of thrusters SPT-100 in space is more than 7500 h [8,9]. Further expansion of application of EPS based on SPT-70 and SPT-100 at Russian SC is planned. Particularly, several light SC are developed. So, SPT-70 will be used at geostationary communication SC “Yamal-200” of Rocket and Space Corporation “Energiya” and “Express-1000” of Scienti3c and Production Association of Applied Mechanics (NPO PM) for orbit adjustment and SC stabilization and

attitude control. E–W& N–S station keeping will be realized. Thrusters SPT-100 are planned to be applied for light SC transfer (“Dialog” of Khrunichev State Scienti3c & Production Space Center and “Ruslan-MM” of NPO MASH) from low intermediate orbit to geostationary one with further orbit adjustment. Besides, SPT-100 will be used on remote-sensing SC “Yachta-D33” for orbit adjustment and SC attitude control. The application of thrusters with anode layer on small SC for interplanetary missions is also analyzed. The thruster SPT-100 has passed quali3cation tests for correspondence to the western standards and in the nearest 1–3 years it is supposed to use it on foreign geostationary SC “STENTOR,” “ASTRA-1K,” “TELSTAR-8,” “INTELSAT-10,” “INMARSAT-4,” etc. The thruster ROS-99 with nominal power level of 1:35 kW was developed in Keldysh Research Center. The output parameters of this thruster are close to parameters of the thruster SPT-100: thrust—78 mN, speci3c impulse—1663 s, eOciency—47%. The thruster has passed 3300 h lifetime test, the predicted lifetime is 6500 h. State of development is quali3cation model. One of the directions to improve thrusters of this class is noticeable increase of their speci3c impulse. The reason is the increase of total impulse of propulsion systems of SC, developed on the basis of new satellite platforms.

28

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

Particularly, estimations show, that application of thrusters with speci3c impulse up to 3000 s in order to solve the tasks of orbit adjustment of advanced geostationary communication SC (“Express-2000” type of NPO PM), will allow halving the required propellant store. At that it is necessary to provide 80 mN of thrust at power of 1900 W that corresponds to thruster eOciency of 62%. According to NPO PM requirements the thruster X-85M was developed in KeRC. The thruster was tested in the power range between 750 and 2000 W. One of the thruster’ operating modes at total power 2010 W is: thrust—85 mN, speci3c impulse with all losses—3100 s and total eOciency—64%. The expansion of tasks, solved by PS on individual spacecraft, leads to the expediency of developing multimode thrusters. Multimode EPS should admit control in the wide range of power, thrust and speci3c impulse during the Right, providing the greatest e5ectiveness of every task realization. The quali3cation of the thruster SPT-140 is carried out in DB “Fakel” according to Integrated High Royo5 Rocket Propulsion Technology (IHPRPT). The special requirement to SPT-140 is to be multimode, i.e. to operate in wide power range from 2 up to 4:5 kW at the eOciency about 55% and lifetime of 7200 h [10]. The thruster T-120/ROS-2000, designed for tests on new satellite platform Eurostar-3000 (ASTRIUM) is developed at KeRC. The thruster has three nominal operation modes, corresponding to 1:35=2:0=2:5 kW power levels. Particularly, at power 2 kW the thruster

provides: thrust—111 mN, speci3c impulse—1850 s and total eOciency—50%. The state of development is engineering model. The thruster T-160E with nominal power level of 4:5 kW was developed at KeRC for the application in the structure of adjusting PS of heavy geostationary SC and for their transfers. In 1999 the thruster passed the tests in NASA Glenn. According to the information, presented by NASA, at nominal mode the thruster provided thrust of 290 mN, speci3c impulse of 1793 s and eOciency of 57%. The thruster passed the full scope of development for Right tests in the structure of SC “Express—M”, developed by NPO PM. 3. Energy–mass and economical eectiveness of SC transfer to geostationary orbit The combined scheme of SC transfer to GEO is considered: the spacecraft is put to low parking orbit by LV and then transferred to some intermediate orbit by chemical booster (CB). SC transfer from intermediate orbit to GEO is realized by another booster (“apogee” PS) with EPS. The intermediate orbit parameters are optimized according to the minimum criteria of characteristic EPS velocity (permanent work on illuminated trajectory areas). To transfer the SC to GEO from the spaceport Baikonur according to combined scheme (CB + EPS), the optimal parameters of intermediate orbit depending on characteristic velocity, picked up by CB, are presented in Fig. 1.

Fig. 1. Parameters of optimal intermediate orbit vs. CB characteristic velocity.

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

29

Fig. 2. EPS characteristic velocity vs. CB characteristic velocity.

Fig. 3. Thrust eOciency vs. speci3c impulse.

Dependence of minimal characteristic velocity, picked up by “ apogee” booster with EPS, on characteristic CB velocity at optimal parameters of intermediate orbits is presented in Fig. 2 (the spaceport Baikonur). At determining SC mass transferred to GEO the approximate model of EPS mass estimation was used. It considers: mass of EP thrust modules (accepted thruster speci3c mass is 1:5 kg=kW, reservation coeOcient is 2), mass of power processing and control system (accepted speci3c mass is 5 kg=kW, the e5ectiveness of power processing system is assumed to be 90%), propellant mass (Xe) with storage system (accepted speci3c mass of tanks with Xe is  = 0:15; the propellant stock for control and guarantee reserve was supposed to be 15% from nominal propellant con-

sumption); EPS construction mass (speci3c construction mass is assumed 10% from total EPS mass). The accepted general dependence of thrust e5ectiveness of primary EPS T on speci3c impulse Isp is presented in Fig. 3. Ballistic e5ectiveness of EPS application in inter-orbital transportation systems is determined by two interconnected parameters—payload mass, transferred to operating orbit and transfer duration. The increase of transfer duration allows increasing SC mass, put to operation orbit. Dependencies of SC mass in GEO, transferred by LV “Aurora” (the spaceport Baikonur, mass in parking orbit Mo = 11:75 t, initial SC available power in GEO is NU = 5 kW=t), on transfer duration are presented in Fig. 4 for the following cases: without taking solar

30

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

Fig. 4. SC mass vs. transfer time considering degradation (SB based on p-Si, SB based on GaAs) and without degradation; Aurora LV.

arrays (SA) degradation into account and with considering the degradation of SA on the base of gallium arsenide and silicon photo-converters. At nominal transfer duration Ttr = 150–200 days and 3xed SC mass, SB degradation leads to Ttr increase on ∼ 20–25% for silicon SA and on ∼ 5–10% for gallium arsenide SA (protection glass thickness of SA was 300 m). At estimation of economic e5ectiveness of EPS application for transportation operations di5erent economic activities can be used. Widely applied criteria of LV economic e5ectivetr ness is speci3c cost of transfer to operating orbit Csp (transfer expenses per SC unit mass). In case of EPS application in the structure of inter-orbital transportatr tion systems to estimate Csp the following factors are to be considered additionally to LV launching cost (or rather to freight cost including insurance cost): • Additional expenses to primary apogee EPS, including propellant cost; • Additional expenses for ground control complex, which realizes the navigation and control during the long section of transfer with low thrust; • Additional expenses for modernization of onboard SC control system, connected with its functions complication; • Additional expenses, connected with the necessity to increase power (and cost) of SA because of their degradation in Earth radiation belts.

The inRuence of transfer duration on speci3c cost of transport operation (speci3c transfer cost tr ) is determined (beside the foregoing additional Csp expenses) by the capital value (the scale of bank interest) and by the cost of the spacecraft itself. If the capital value is zero, then the delay of SC put into operation, connected with transfer duration, does not lead to the additional losses and, accordingly, does not increase speci3c transfer cost. The expenses are esteemed (recalculated) at the moment of transport operation completion. The estimations of calculated speci3c transfer cost tr Csp were made depending on transfer duration Ttr at constant available SC power in GEO NU = 5 kW=t and considering SA radiation degradation (it was supposed to use gallium arsenide SA). The estimations were made at the example of SC transfer to GEO from Baikonur spaceport by LV “Aurora” with booster “Corvet” and apogee EPS. The following basic values of economic activities were accepted in calculations [11,12]: • Speci3c cost of satellite platform (not considering the cost of apogee EPS) Csp SC = 44; 000 $=kg. • EPS speci3c cost (not including the propellant storage system and the propellant itself) Csp EPS = 30; 000 $=kg. • Speci3c cost of propellant storage system, including the propellant itself Csp Xe = 2000 $=kg.

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

31

tr vs. transfer duration (Aurora LV, Baikonur). Fig. 5. Transfer speci3c cost Csp

• Speci3c expenses for ground control complex Csp contr = 7; 000; 000 $=year. • Speci3c cost of solar arrays (GaAs) Csp SA = 1; 500; 000 $=kW. • Speci3c mass of solar arrays SA = 20 kg=kW. • Launch cost (freight) of LV “Aurora” CLV = $50; 000; 000. • Capital value (annual bank interest) E = 15%. Figure 5 illustrates the inRuence of transfer duration tr on speci3c transfer cost Csp . In spite of the fact that SC mass increases with increase of transfer time, speci3c tr transfer cost Csp does not decrease monotonously, but reaches the minimum at Ttr 200 days. The presence tr is determined by the expenses for of minimum Csp SC manufacturing at non-zero capital value. For the tr optimal transfer duration Csp decreases for about 30% in comparison with SC transfer without apogee EPS using (direct SC transfer to GEO by chemical booster “Corvet”). Concerning geostationary communication SC it is advisable to estimate the economical e5ectiveness of “apogee” EPS using the index of speci3c cost of transp transp equivalent transponder Csp (Csp = C =ntransp , where C —total expenses at the moment of SC transfer to GEO with taking into account launch cost of LV, SC cost, “apogee” EPS cost and other expenses, considering capital value; ntransp —number of equivalent transponders). The index of speci3c transfer cost tr Csp (k$/kg) characterizes economical e5ectiveness of launch vehicles (in regarded case, EPS inRuence is considered) and the index of speci3c cost of

transp characterizes commerequivalent transponder Csp cial e5ectiveness of geostationary communication SC together with regarded scheme of its transfer, using EPS. At estimation of the number of equivalent transponders it was assumed that payload relative mass (transponder complex) is constant and is 30% of SC mass in GEO. transp The dependence of transponder speci3c cost Csp (M$/transp.) on SC transfer duration to GEO by launching system LV “Aurora” + CB “Corvet” + EPS (considering expenses for LV, spacecraft and foregoing basic values of the main economic indexes) is presented in Fig. 6. It is obvious, that the value transp Csp has the minimum at transfer duration about 200 days. For the optimal transfer duration the speci3c cost of transponder is reduced approximately for 14% in comparison with transfer without “apogee” EPS usage. The inRuence of regarded economic indexes on pretransp sented speci3c cost of equivalent transponders Csp was analyzed. Some results are presented in Figs. 7–9. It is necessary to note the weak inRuence of EPS speci3c cost (including propellant) and expenses for ground control complex on transponder speci3c cost. Besidescc, from results, presented in Figs. 7–9, it is seen that at wide variation of main regarded economic indexes, the optimal transfer duration changes slightly near Ttr 200 days (with the exception of capital value E change case). tr Beside of criteria of transfer speci3c cost Csp transp and speci3c cost of equivalent transponder Csp it

32

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

Fig. 6. Speci3c cost of transponder vs. transfer time (Aurora LV, Baikonur).

Fig. 7. InRuence of SC speci3c cost on transponder speci3c cost.

Fig. 8. InRuence of capital value on transponder speci3c cost.

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

33

Fig. 9. InRuence of EPS speci3c cost on transponder speci3c cost.

Fig. 10. NPV vs. operation time.

is reasonable to use the most general criterion of net present value (NPV), widely applied at estimating the investment aptitude of projects, and which allows to compare the income level of the project with the average one in the market. Net present value includes the expenses for the SC, its transfer and the income from its operation during the whole term of OL (for telecommunication SC the annual income is determined by the number of transponders and annual rent cost of the

transponder [13]). While estimating NPV, the income from SC operation was calculated by the moment of SC launch. The dependence of NPV on the duration of SC operation for LV “Aurora” at di5erent transfer duration with EPS is presented in Fig. 10. To compare, the change of NPV at SC transfer with LV “Proton” and CB “Briz-M” (without EPS) is also presented in this 3gure. It is seen that SC transfer to GEO by medium class LV “Aurora” with apogee EPS usage provides practically the same NPV level during Ttr = 180 days,

34

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

Fig. 11. NPV vs. transfer time (inRuence of transponder rent cost on NPV).

in comparison with SC transfer by LV “Proton” without apogee EPS. Figure 11 illustrates the inRuence of the cost of annual transponder rent on NPV. It is obvious, that according to NPV criteria the optimal values of transfer duration increase and reach ∼ 300 days. At the same time, the quickest NPV increase is till Ttr 200 days. At further Ttr increase NPV rising slows down. For example, for transponder rent cost Dtransp = 1; 000; 000 $=year the Ttr increase till 200 days leads to NPV increase in three times. The further Ttr increase from 200 days till 300 days leads to the additional NPV increase for 10%. Figure 12 illustrates the sensitivity of three regarded transp tr criteria of the economic e5ectiveness (Csp ; Csp , NPV) to the change for 1% of regarded determining parameters at transfer time Ttr = 180 days. It is clear, that the launch cost, the SC cost and the capital value have the greatest inRuence on regarded criteria. The optimization of speci3c impulse of primary EPS was realized according to criteria of SC mass tr in GEO and transfer speci3c cost Csp . SC mass change (without apogee EPS) as a function of EP speci3c impulse is presented in Fig. 13 for the variant of using gallium arsenide SA, (the analogous calculations are presented for silicon SA). The application of gallium arsenide SA allows to increase the SC mass in GEO for 8–9% in comparison with silicon SA at similar transfer duration in the range of Ttr = 180–360 days. The optimal values of speci3c impulse depend on transfer duration opt opt 2000 s for Ttr = 180 days and Isp 2500 s and are Isp

Fig. 12. InRuence of determining parameters on economic e5ectiveness.

for Ttr = 360 days. A very regular optimum is realized. tr The dependence of transfer speci3c cost Csp (k$=kg) on speci3c impulse of apogee EPS is presented in Fig. 14. It is seen, that optimal values of speci3c impulse tr are little bit higher according to criteria Csp , than these values according to MSC criteria. For example, for opt Ttr = 180 days − Isp 2100 s and for Ttr = 360 days − opt Isp 3000 s. It is necessary to note, that for thrusters, adjusted orbit inclination of heavy geostationary SC with high level of available power the optimal value of speci3c impulse is ∼ 3000 s and above. In this

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

35

Fig. 13. SC mass vs. speci3c impulse (SA based on GaAs).

Fig. 14. Transfer speci3c cost vs. speci3c impulse.

Fig. 15. Advantage in SC payload mass vs. operating orbit altitude (initial orbit altitude H0 = 500 km, SC available power 2 W=kg).

36

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37

connection, it is more reasonable to develop the special EP modi3cations for combined multifunctional EPS, used both at the area of SC transfer to GEO (Isp = 2000 s) and in operating orbit during its regular operation (Isp = 3000 s). 4. Eectiveness of EP application for low-orbit and medium-altitude spacecraft The EPS application for inter-orbital transportation is e5ective not only for SC transfer to high operating orbits (3rst of all, to GEO). The noticeable e5ect gives the EPS application for SC transfer to low- and medium-altitude orbits, including small spacecraft. For example, many SC of multisatellite low-orbit communication systems or remote-sensing systems have the operating orbit from 700 km till 1500 km and above (Iridium, Globalstar and others). LV realizes individual or group transfers of such SC to low orbits, from which they are transferred to higher operating orbit by their onboard PS with low thrust LRE or electric-heating thrusters. EP application instead of low thrust LRE provides noticeable advantage in payload mass. The dependencies of payload mass increase on operating orbit altitude are presented in Fig. 15 for two levels of EP speci3c impulse (1350 and 1500 s) at available power of small SC NU = 2 kW=t. Characteristic velocity losses for SC transfer, orbit keeping during OL and for deorbiting after OL is 3nished were taken into account. For example, at initial orbit altitude H0 =500 km and operating orbit altitude Hf = 1500 km the payload mass increase, in case of EPS usage, is from 10% to 20% of SC mass at EPS speci3c impulse 1350 and 1500 s, accordingly. At spacecraft mass ∼ 1000 kg the payload mass increase is VM = 100–200 kg, accordingly. EPS required power for the spacecraft with mass 500 – 1000 kg is NEPS = 1–2 kW. Such EPS can be formed not only on the base of thrust modules of kilowatt level, but also basing on EP bunches with unitary power of hundreds watts. 5. Conclusion Electric propulsion, presently used mainly for orbit adjustment of geostationary SC, can become the main

thruster type in the nearest future, providing transportation missions in interplanetary and near-earth space (3rst of all, for SC transfer to GEO). The examples of SC transfer to GEO by medium class LV “Aurora” show high ballistic (energy–mass) and economic e5ectiveness of EP use in the structure of “apogee” propulsion system. At transfer duration 180 days it is possible to double SC mass in GEO. To estimate economic e5ectiveness of EPS use for transfer to GEO three types of economic indexes were studied: speci3c transfer cost, speci3c transponder cost (applying to communication SC), net reduced value (applying to communication SC). For all studied criteria of economic e5ectiveness the rational transfer duration has similar values ∼ 200 days. At this the rational value of EP speci3c impulse (as for energy– mass as for economic e5ectiveness) is ∼ 2000 s. The main activities of Russian EP developers are concentrated on HET as the most suitable thrusters for tasks of orbit adjustment of SC for di5erent purpose and for providing transportation missions in near-earth space. References [1] V.S. Avduevsky, E.L. Akim, G.B. E3mov, T.M. Eneev, M.Ya. Marov, S.D. Kulikov, O.V. Popkov, M.S. Konstantinov, G.A. Popov, Space vehicle of new generation for solar system study, Paper IAF-98-Q.2.06, Melbourne, Australia, September/October 1998. [2] S.R. Oleson, R.M. Myers, C.A. Kluever, J.P. Riehl, F.M. Curran, Advanced propulsion for geostationary orbit insertion and North–South station keeping, Journal of Spacecraft and Rocket 34 (1997) 22–28. [3] S.R. Oleson, Mission advantages of constant power, variable Isp electrostatic thrusters, Paper AIAA-2000-3413, 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, USA, July 2000. [4] S.R. Oleson, R.M. Myers, Launch vehicle and power level impacts on electric GEO insertion, Paper AIAA 96-2978, 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Lake Buena Vista, USA, July 1996. [5] A.E. Buravin, M.N. Grishko, V.G. Ivashin, Realization of satellite systems using small sized spacecraft on geostationary orbit, The Third International Conference on Satellite Communications, Moscow, Russia, September 1998. [6] Spacecraft systems of communication and broadcasting, Supplement No. 2 to the Annual 1999/2000. [7] O.A. Gorshkov, A.S. Koroteev, B.A. Arkhipov, V.N. Murashko, N.A. An3mov, V.I. Lukyashenko, V. Kim, G.A. Popov, Overview of Russian activities in electric propulsions, Paper AIAA-2001-3229, 37th AIAA/ASME/SAE/ASEE Joint

A.S. Koroteev et al. / Acta Astronautica 54 (2003) 25 – 37 Propulsion Conference and Exhibit, Salt Lake City, USA, July 2001. [8] B. Arkhipov, R. Gnizdor, K. Kozubsky, M. Day, et al., Extending the range of SPT operation: development statua of 300 and 4500 W thrusters, Paper AIAA-96-2708, 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Lake Buena Vista, USA, July 1996. [9] K. Kozubsky, N. Maslennikov, S. Pridannikov, A. Rumiantzev, Study of long operation capacity plasma thruster SPT-100 at power 3500 W, Paper AIAA-99-120, 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Los Angeles, USA, 1999. [10] R. Gnizdor, K. Kozubsky, N. Maslennikov, V. Murashko, S. Pridannikov, V. Kim, Performance and quali3cation

37

status of multimode stationary plasma thruster SPT-140, Paper IEPC-99-090, 26th International Electric Propulsion Conference, Kitakyushu, Japan, October 1999. [11] Yu.G. Milov, State of work and perspectives of spacecraft development to provide communication and television, The Third International Conference on Satellite Communications, Moscow, Russia, September 1998. [12] T.M. Miller, B.G. Seaworth, Mission factor a5ecting cost optimization of solar electric orbital transfer vehicles, Paper IEPC-93-202, 23rd International Electric Propulsion Conference, Seattle, USA, September 1993. [13] Yu.M. Gornostaev, V.V. Sokolov, L.M. Nevdiaev, Perspective satellite communication systems, Moscow, Russia, 2000.