ON THE CONCEPT OF ALL ELECTRIC PROPULSION SPACECRAFT

全文:

Introduction. The appearance of the new concept of “All electric spacecraft” has caused considerable controversy among specialists about the benefits and drawbacks of this concept of constructing spacecraft propulsion subsystem. The substantive content of these discussions is often hampered by discrepancies in under- standing of the term in contrast to the clear-cut concept of “Propulsion subsystem”. Therefore there is a need to raise two interdependent issues and respond to them. The issues are the following: a) clarifying the content of the concept of “All electric propulsion spacecraft”; b) assessing the advantages and disadvantages arising from the use of this concept while designing and operat- ing GEO spacecraft. The article is devoted the consid- eration of these issues. Clarifying the content of the concept of “All elec- tric propulsion spacecraft”. We can translate the Eng- lish term “All electric propulsion spacecraft” into Rus- sian, but the most accurate translation according to the meaning would be “spacecraft with electro jet propulsion subsystem solving any tasks”. A simplified version close to word-for-word translation and the most widespread is “All electric propulsion spacecraft”. Let’s consider the meaning of this term that comprises a certain ap- proach to the construction of propulsion subsystem of GEO spacecraft. In considering the issue, it should be noted that his- torically, there are two different technical schools dealing with the concept of launching spacecraft into geostation- ary orbit. The Western School (that some American and European companies adhered to from the very beginning) involves initial launching of a satellite into elliptical geo- transfer orbit with the further raising into geostationary orbit with the help of the own apogee engine unit of a satellite. The engine unit of a satellite is called “apogee” because it produces master impulse at orbit apogee. Spe- cialists used only two-component systems with rather high thrust (hundreds of pound feet in order of magni- tude) that allowed to perform apogee maneuvers at an optimal point of orbit and to obtain quite acceptable char- acteristics of engines in terms of economic efficiency (specific impulse). The main advantage of this approach was the speed of the execution of manoeuvres. Rather high engine thrust allowed to perform a manoeuvre during the limited period of time - not more than several days. That reduced to the minimum the dose obtained by a sat- ellite while passing through radiation belts and allowed to put a satellite into operation as quickly as possible. One more important competitive advantage of the approach was the possibility to use launch vehicles of different types. It helped select the most convenient options with regard to the price and other points. Thrusters used re- maining fuel after the completion of manoeuvres for the orbit correction during operational lifetime of the spacecraft and for the creation of control moments while main- taining satellite orientation in space. While implementing the concept, it is necessary to place in spacecraft the propulsion subsystem with massive tanks of oxidizing agent and fuel, with inflation system and other necessary attributes; in this, the mass of the fuelled propulsion subsystem unit is 40-50 % of the launch mass of spacecraft. In contrast to the Western Technical School, the So- viet (Russian) school comprised the launch into geosta- tionary orbit on the so-called “direct” scheme: using a separate Booster that has some properties of spacecraft but it is actually a part of a launching vehicle. To some extent, a Booster may be considered as an upper stage of a launch vehicle. In such a conception there is no apogee engine unit as a component of spacecraft. Orbit correction and the creation of control moments are carried out during the spacecraft lifetime by on-board subsystems on the basis of low thrust, both two-component and single- component electric propulsion thrusters [1]. In this, in the latter case, developers can significantly reduce the mass of a fuel due to the high economical efficiency of electric propulsion that are generally used for correcting the orbit of spacecraft. The evolution of the western approach was that spe- cialists began to apply electric propulsion, in particular, ion propulsion, for correcting orbits of geostationary spacecraft in north-south direction (the correction of in- clination of the orbit or north-south station keeping). In this, the apogee engine unit in spacecraft was main- tained, the correction of the longitude (West-East) and the performance of control moments continued to be carried out by the two-component liquid-fuelled low thrusters on the rest of the fuel of the apogee engine unit. In particular, the spacecraft on 601HP platform base with the use of the XIPS-13 thrusters were developed by Boeing [2]. The development of the concept comprised the application of more powerful ion thrusters XIPS-25 in spacecraft on Boeing 702 base platform not only for the correction of orbital inclination but for the partial per- formance of the maneuver of orbit raising as well [2; 3]. Four ion thrusters were installed on the revolving plat- forms that provided thrust in the direction of orbital speed in the mode of raising, and in the mode of orbital correc- tion - in the directions “North-South” and “West-East” (fig. 1). In this, in the mode of orbital raising the engines operated at maximum power 4.2 kW, and in the mode of orbital correction - on half-power 2.1 kW. As the engines were installed on the moving platforms (drives), during the operation it was possible to generate control moments in order to unload momentum wheels of orientation sys- tems. It allowed to considerably save the fuel for the atti- tude control thrusters that were on board the spacecraft together with the apogee engine unit. Fig. 1. The location of orbit control thrusters on board spacecraft built on Boeing 702SP platform base Рис. 1. Размещение двигателей коррекции на КА, построенном на базе платформы Boeing 702SP The next considerable evolutionary step in the con- struction of propulsion system of GEO spacecraft was a complete rejection of the apogee propulsion. In this ver- sion being implemented in the spacecraft created on Boe- ing 702SP platform base, the electrical propulsion subsys- tem took up the challenge of orbital raising [4]. It allowed to significantly decrease the mass of the fuelled propul- sion subsystem and the satellite in whole. This, in turn, gave the possibility to perform pair launches using one launch vehicle that allowed to considerably reduce the cost of the injection of spacecraft into orbit. But it caused the increase of the period of raising operation to 7-9 months owing to fundamentally low thrust of an electrical propulsion. However, in the opinion of western specialists, the advantages of the concept outweigh the disadvantages. With its help, only for the European operator EUTELSAT three spacecraft were manufactured - EUTELSAT 115 West B" EUTELSAT 117 West B (both on Boeing 702SP platform base), EUTELSAT 172B (on EUROSTAR-3000 base) [5]. These satellites were put into operation. The duration of orbital raising was from 4 to 6 months. On the satellite called EUTELSAT 172B the Russian-made thrusters called SPD-140D with a power of 4.5 kW were first used. The interest in the satellites of this type is increasing in the world. We can note that the evolutionary development of the concept of the apogee engine unit with higher significance of electric propulsion in solving the problem of orbital raising led to the qualitative leap - total rejection of the chemically fuelled apogee propulsion; it generated the concept of “all electric propulsion spacecraft”. In western specialists' understanding, the lack of the chemically fu- elled apogee propulsion is an essential feature of this con- cept. There is, however, an additional issue concerning atti- tude control thrusters. If correction thrusters are installed on the drives that allow to decline a thrust vector from the line passing through the centre of spacecraft masses, there is every reason to impose the task of creating control moments to the correction engines as well and to com- pletely refuse to use attitude control thrusters. In this case the concept of all electric propulsion spacecraft will be carried to its logical conclusion, i. e. in this case space- craft will not have any other thrusters apart from electric propulsion thrusters, and they will solve two different functional tasks: moving the center of spacecraft’s mass of and control its angular position. However, the following circumstance hinders such ul- timate realization of the “all electric propulsion space- craft” conception: there are so called initial attitude modes of spacecraft and modes of ensuring survivability when the use of attitude control system with momentum wheels is difficult or impossible. Attitude control thrusters are necessary in this case. However, since impeded power balance is possible for the current modes, the use of elec- tric propulsion is difficult due to their high energy con- sumption. Therefore, an additional independent engine subsystem providing the creation of control moments ar- ranged along the three axes of spacecraft is necessary for such cases. The choice of the thruster type for such a sub- system depends on the preferences of designers and can be quite broad - from simple nozzles on a cold gas or heated nozzles to mono-fuel or two-component low- thrusters. Generally when speaking about the advantages of the concept of “all electric propulsion spacecraft” this issue is neglected and it is considered to be of secondary impor- tance and not worthy of special attention. It is interesting to note that if this concept is interpreted in this way, all spacecraft developed by JSC “ISS” starting from 1982 and equipped with electrically orbit correction propulsion subsystems and mono-fuel attitude control subsystems can be considered as “all electric propulsion spacecraft”. A. Vnukov and his co-authors repeatedly pointed to it in publications [6]. The question is, how to interpret the concept of “all electric propulsion spacecraft” correctly and unequivocally so that the specialists of both Russian and Western Technical Schools understand it equally? The following approach is possible: if we take into account the fact that the term “all electric propulsion spacecraft” was invented in the Western Technical School, the use of the interpretation of the term in the form in which it was formed in the Western specialists' view will be logical, i. e. it is logical to mean by this term a kind of spacecraft with an electrically propulsion subsystem without chemically fuelled apogee propulsion. The issue concerning the auxiliary attitude control propul- sion is not considered. This approach implies the rejection of the ultimate interpretation of the term “all electric propulsion spacecraft”, which excludes the presence of any other engine subsystems on board, except the electri- cal propulsion subsystem. If we agree with the proposed definition, the meaning of the term or the concept of “all electric propulsion spacecraft” will be unambiguous for everyone. In addi- tion, there will be freedom in choosing the type of auxil- iary attitude control propulsion subsystem, which in any case is necessary on board spacecraft and the fact of its existence excludes the possibility of applying the ultimate interpretation of the “all electric propulsion spacecraft” conception Evaluating the advantages and disadvantages of the versions of the concept of “all electric propulsion spacecraft”. Let us suppose that the proposed interpreta- tion of the concept “all electric propulsion spacecraft” is adopted. It is possible to modify it, in this case we give the auxiliary attitude control subsystem a little more ad- vanced functions, i. e. we assign to it the tasks of creating control moments not only in some separate operating modes of a satellite, but during its entire service life as well. This approach, in particular, has been applied to all spacecraft developed by JSC “ISS” [6; 7]. We have evaluated the advantages and disadvantages of “all elec- tric propulsion spacecraft” concept in comparison with the extended version applied on the products of JSC “ISS”, which have been taken as a base. We have taken the following baseline data for the assessment: - the need for the total impulse for various tasks with respect to the general stock are: orbit raising and orbit correction - 98 %, orientation during the service life in- cluding the period of the initial modes and the period of the modes of ensuring survivability - 2 %; - the base scheme includes the propulsion subsystem of orbit raising and correction of the orbit and the attitude control propulsion subsystem; - the orbit control and orbit raising propulsion subsys- tem consists of 6 perspective plasma engines of KM-75 type [8], a modified version of power processing unit (PPU) suitable for powering two thrusters simultaneously, a xenon feed unit and a xenon storage unit developed by JSC “ISS” [9]; - the attitude control propulsion consists of 8 thruster units with mono-fuel K50-10 thrusters [10], storage and feed unit. In all the versions in which we do not use mono-fuel thrusters, cold xenon nozzles powered from the common xenon tank create control moments in the initial modes and modes of ensuring survivability. In the version of hard-mounted correction thrusters, we have increased their number to ten to create control moments along three axes, taking into account the thrust- ers of orbit raising. In the version of the correction thrust- ers installed on rotary devices (drivers), the thrusters are used both in the mode of orbit raising and in the orbit correction mode with simultaneous creation of control moments. We have reduced the number of engines to the lowest possible - four. We have taken into account the mass of the rods, drives and its control units. The main criterion for evaluating the versions is the total mass of fueled propulsion subsystems. Additional criteria are operational reliability and easy control. The mass estimate for different design versions of the propul- sion subsystems is presented in table. The results presented in the table show that the base design version of the satellite propulsion system (consist- ing of two independent subsystems: orbit control and or- bit raising subsystems based on plasma thrusters and atti- tude control subsystem based on a monopropellant thrust- ers) has a significant advantage over other versions. We obtained similar results when carrying out project evalua- tions for other types of plasma engines and other types of PPU. This result suggests that at the present stage generat- ing of control moments for the orientation of spacecraft by plasma orbit control thrusters is possible but it is irra- tional, since it requires a significant increase in mass in comparison with the version of two independent propul- sion subsystems. The results of design estimates are un- expected as, at first sight, a system based on single, uni- versal, very economical and reduced to a minimum num- ber of thrusters must have the lowest mass. However, many calculations that we have carried out taking into account the attendant factors, the available data on the masses of the blocks of propulsion subsystems and proto- types of new equipment suggest otherwise. The reason is that in order to implement the versions using the electrical propulsion in creating control moments, it is necessary to increase the number of thrusters or install them on drives and rods, which together with their control units have a sufficiently large mass. In addition, it is necessary to take into account the mass of the auxiliary cold gas-reactive subsystem with the reserve of the fuel mass. We note that papers [11; 12] demonstrate that the version with fixed orbit control thrusters on the levers relative to the center of mass proved to be lighter than the base version, but this result was received without taking into account the xenon mass consumption for orbit raising and disturbing torques compensation at this stage. We should note that the pessimistic estimates of the increase in mass for the implementation of versions dif- ferent from the basic one are characteristic for the current level of perfection of the design of the equipment being used. But, if some lighter devices of controlling thrust vector appear, we can significantly reduce the mass of the driven version. The total mass of geostationary satellite propulsion subsystems for the different design versions Design version Difference from the base scheme, kg The scheme with ten fixed correction units based on KM-75 using modified PPU +32 The scheme with four correction units based on KM-75 on four rods with uniaxial drives using modified PPU +71 In order to comprehensively evaluate the feasibility of applying this or that concept, we have considered other aspects of the versions of engine subsystem constructions. Operational reliability of the system is one of the most important issues. It is necessary to compare the evaluation of reliability of the system with the combination of func- tions of orbit and attitude control with the reliability of the original scheme with separate propulsion subsystems of ones. We have determined the reliability of the original sys- tem of RINIT, by the following expression: RINIT = RCP · RACP, (1) where RCP - reliability of the of correction propulsion; RACP - reliability of the attitude control propulsion. The reliability of the system with the combination of the functions R' is similarly defined by the expression: R' = RAUX · RUNI, (2) where RAUX - reliability of the auxiliary propulsion for operation in the initial modes and modes of ensuring sur- vivability; RUNI - reliability of the universal propulsion subsystem of orbit and attitude control The reliability of the auxiliary propulsion subsystem in the first approximation can be comparable (equal to or slightly higher) with the reliability of the attitude control propulsion subsystem in the original version. Thus, the reliability of the system with the combination of functions depends on the second component - the reliability of the universal propulsion subsystem. Obviously, its reliability is lower than the reliability of the orbit control subsystem in the original version, since the composition of the sub- system and the structural scheme of reliability are more complicated. Most likely, the reliability of a system with the combination of functions is comparable or lower than the reliability of the original scheme with two independ- ent engine subsystems. To confirm this assumption, we have obtained a nu- merical estimate of the reliability for a specific version of the propulsion subsystem. It is possible to do it, for ex- ample, for the scheme with the fixed correction thrusters on the body of the spacecraft considered in [11]. This paper considers only the problem of orbit correction with the simultaneous creation of control moments without orbit raising. We have chosen the diagonal placement of 4 thrusters (fig. 2) as a base scheme of the orbit control pro- pulsion subsystem. Fig. 3 shows the structural scheme of reliability for this design of the propulsion subsystem. Fig. 2. Coordinate system and nominal thrust directions of orbit correction thrusters in the base scheme Рис. 2. Система координат и номинальные направления выдачи тяги двигателей коррекции орбиты в базовой схеме Подпись: RSFS Подпись: RPPU Подпись: RTU-3 Подпись: RTU-1 Подпись: RTU-1 Inclination correction Longitude correction +Y Longitude correction -Y Fig. 3. Propulsion subsystem reliability scheme for the base version: RSFS - reliability of xenon storage and feed system (storage unit and feed unit); RPPU - reliability of power processing unit; RTU-3 - equivalent reliability of a scheme fragment for the correction of orbital inclination at triple reservation of the thruster units; RTU-1 - equivalent reliability of a scheme fragment for longitude orbit correction in plus or minus directions at single reservation of the thruster units Рис. 3. Структурная схема надежности подсистемы коррекции для базового варианта: РСХП - ВБР системы хранения и подачи ксенона (блок хранения и блок подачи ксенона); РСПУ - ВБР системы преобразования и управления (СПУ); РДК-3 - эквивалентная ВБР участка схемы для коррек- ции наклонения при трехкратном резервировании двигателей коррекции; РДК-1 - эквивалентная ВБР участков схемы для коррекции долготы в направлении «плюс» или «минус» при однократном резервировании (дублировании) двигателей коррекции Fig. 4. The location of orbit control thrusters using hard attaching on the SC structure Рис. 4. Размещение двигателей коррекции в схеме с неподвижным закреплением двигателей на корпусе КА Подпись: RSFS Подпись: RPPU Подпись: RPPU Подпись: RTU-7 РДК- Inclination correction Longitude correction +Y Longitude correction -Y Подпись: RTU+X Подпись: RTU-X Подпись: RTU+Y Подпись: RTU-Y Подпись: RTU+Z Подпись: RTU-Z Control moments +МХ ..-МZ Fig. 5. The scheme of reliability of propulsion subsystem for the version with the combination of functions: RSFS - reliability of xenon storage and feed system (storage unit and feed unit); RPPU - reliability of power process- ing unit; RTU-7 - equivalent reliability of a scheme fragment for the correction of orbital inclination at sevenfold res- ervation of the thruster units; RTU-3 - equivalent reliability of a scheme fragment for longitude orbit correction in plus or minus directions at triple reservation of the thruster units; RTU+X…-Z - equivalent reliability of scheme frag- ments for control moment creation at series-parallel reservation of thrusters Рис. 5. Структурная схема надежности двигательной подсистемы с совмещением функций: РСХП - ВБР системы хранения и подачи ксенона; РСПУ - ВБР СПУ; РДК-7 - эквивалентная ВБР участка схемы для коррекции наклонения при семикратном резервировании двигателей коррекции; РДК-3 - эквивалентная ВБР участков схемы для коррекции долготы в направлении «плюс» или «минус» при трехкратном резервировании двигателей коррекции; РДК±X…Z - эквивалентные ВБР участков схемы при последовательно- параллельном резервировании двигателей коррекции для создания управляющих моментов The article also represents the diagonal placement of eight engines in pairs with small (approximately 1°) de- viations of the thrust lines relative to the center of mass to create control moments (fig. 4). Engines work in pairs to create “pure” control mo- ments of a certain sign on a certain axis, while a symmet- rically located pair duplicates the selected one. Due to the lack of estimations of the reliability probability of the PPU for powering two engines, we have applied two simi- lar PPU with the ability to power one thruster. The esti- mates of the reliability for such a device are known. If we take into account our assumptions, the structural scheme of the reliability of the propulsion subsystem with the combination of functions will have the form shown in fig. 5. Obviously, this scheme looks more complicated than the one shown in fig. 3. The calculation using the above-mentioned structural schemes taking into account the known values of the reli- ability of the constituent elements shows RINIT ≈ 0.9775 for the initial scheme, and R' ≈ 0.9769 for the scheme with the function combination, i. e. the reliability of the original version with independent propulsion subsystems is actually slightly higher, although the differ- ence is observed only in the third decimal place. For the version with the installation of orbit control thrusters on the drives, the structural scheme of reliability obviously must be different and it must take into account the reliability of the drives including mechanical compo- nents, electronic control units and flexible units for xenon feed. Probably the total reliability for such a scheme will roughly correspond to the version with fixed installation of the thrusters, that is, slightly lower than the original scheme contains. The above-mentioned formal conclusions about the reduced reliability of the propulsion subsystem with the combination of functions are illustrated by using a simple example: if we assume that the electrical propulsion sub- system is completely or partially out of order, the problem of control moments creating is either not solved at all, or it is solved extremely limitedly by the auxiliary subsys- tem on a cold gas. We should note that the inoperability of electrical pro- pulsion can occur both as a result of internal and external causes. One of the main external reasons is the limitation of power consumption. So, in case of the failure in the orientation of spacecraft, it is actual to create control mo- ments for its reconstruction. But in the non-orientable state, the orientation of solar cells is also disturbed, so the generation of electric power is reduced. The chemical battery capacity is limited. If in this case we use electrical propulsion, which are fairly powerful consumers of en- ergy, for orientation, the problem of ensuring the surviv- ability of spacecraft becomes very critical. In the case of the presence of two independent propulsion subsystems, even in case of disorientation, the thrusters creating con- trol moments remain operational, since the energy neces- sary for their operation is enclosed in fuel itself. Thus, to ensure the operational reliability and surviv- ability of spacecraft, the concept of two independent pro- pulsion subsystems (attitude control and orbit control) is preferable. With a certain arrangement of the attitude con- trol thrusters, they can also be used to produce an orbit corrective impulse. In particular, certain SC of JSC “ISS” after the end of their service life in the presence of a re- sidual fuel were taken to the disposal orbit by the orienta- tion thrusters. Thus, to ensure flexibility in application, the concept of two independent propulsion subsystems is useful as well. In this case, the algorithms for their use are com- pletely separated from each other. That is, the program of ignition of orbit control thrusters is completely unrelated to the algorithm of attitude control thruster’s ignition for momentum wheels unloading. Each correction session consists of one ignition of a thruster that is selected to output an impulse in a certain direction. If correction thrusters are used to create the control moments, the algo- rithms influence each other, therefore, for example, if the thrusters are fixed, the number of their ignitions in the correction session can increase to 6-8 [13; 14]. This com- plicates planning of corrections, as well as it puts addi- tional demands on the resource of the thrusters according to the number of ignitions. Conclusion. We have examined the essence of the concept of “all electric propulsion spacecraft”. We can offer the following definition of this term: a spacecraft without an apogee propulsion system on chemical fuel, equipped with an electrical propulsion subsystem. This definition does not include the presence of any auxiliary propulsion subsystem onboard the spacecraft to provide initial orientation modes after separation from the upper stage and modes of ensuring survivability. We have shown that without such an auxiliary subsystem of orien- tation, the construction of the propulsion system of “all electric propulsion spacecraft” is impossible, therefore in its “pure form” with only electrically reactive thrusters this concept cannot be realized. We have evaluated the advantages and disadvantages of the versions of the of “all electric propulsion space- craft” propulsion subsystems in comparison with the ver- sion of two independent subsystems - for orbit control and attitude control of spacecraft. We have demonstrated that with the existing level of mass perfection of equip- ment, the concept of two independent propulsion subsys- tems has the advantage over other versions by total mass, operational reliability and ensuring survivability of the spacecraft, as well as for flexibility in application. The estimates of the mass of competing versions should be clarified in the development of more advanced thrust vec- tor control devices.

作者简介

Yu. Yermoshkin

JSC “Academician M. F. Reshetnev “Information Satellite Systems”

Email: erm@iss-reshetnev.ru
52, Lenina Str., Zheleznogorsk, Krasnoyarsk Region, 662972, Russian Federation

D. Volkov

JSC “Academician M. F. Reshetnev “Information Satellite Systems”

52, Lenina Str., Zheleznogorsk, Krasnoyarsk Region, 662972, Russian Federation

E. Yakimov

JSC “Academician M. F. Reshetnev “Information Satellite Systems”

52, Lenina Str., Zheleznogorsk, Krasnoyarsk Region, 662972, Russian Federation

参考

补充文件