Attitude and Orbit Control of Small Satellites and Constellations with FEEP Thrusters Salvo Marcuccio*, Stefano Giannelli§, Mariano Andrenucci^ Centrospazio, Via Gherardesca 5, 56014 Ospedaletto, Pisa, Italy Tel. +39 50 985072 - Fax +39 50 974094 - E-mail:
[email protected]
Abstract
L
OW thrust electric propulsion systems find a promising application field on small-to-medium sized spacecraft, with mass ranging from a few hundred to about 1000 kg and orbital altitudes of 700 km or higher, like those envisaged for most of the proposed LEO telecommunication constellations. To evaluate the performance of FEEP for such missions, a study was carried out on a typical attitude and orbit maintenance case, using current models for perturbation torques and forces. It was assumed that three-axes attitude control and drag compensation be fully performed by means of FEEP thrusters for a mission duration of 5 years, under a maximum thrust constraint of 1 mN per thruster. A reference case (900 kg spacecraft in 800 km LEO) was analyzed in detail, and a parametric performance analysis was carried out for the spacecraft mass range 100 - 1000 kg and orbital altitude range 400 800 km. As a result, the use of FEEP to replace momentum and reaction wheels and cold gas or hydrazine thrusters was found to be very attractive for satellites of mass in excess of 400 kg in orbits higher than 400 km about. Outside this mass and altitude ranges, the use of FEEP may still lead to significant mass savings, but the AOCS configuration must be carefully studied on a case by case basis. For all the cases studied, the mass of a full, 16-thrusters propulsion system, including thrusters, neutralizers, propellant, and redundant electronics, is less than 45 kilograms.
Introduction Electric thrusters have been traditionally associated with two main application domains, that is, in the near term, NSSK of telecommunication satellites and, in the far fu* Project manager, Centrospazio. § Undergraduate student, University of Pisa. ^ Director, Centrospazio; Professor, Aerospace Dept., University of Pisa. Copyright 1997 by the Electric Rocket Propulsion Society. All rights reserved.
ture, interplanetary primary propulsion. This picture is rapidly changing with the recent trend towards the use of smaller spacecraft. Electric propulsion (EP) may play a key role in several classes of missions, spanning from scientific drag-free satellites, to earth observation and communication microspacecraft, to formation flying and constellations. In many cases, the mass savings made possible by the use of EP instead of chemical or cold gas propulsion is a mission enabling issue. Of course, this new scenario makes a new approach to EP systems necessary. The case of FEEP is a typical example of the shift in perspective that is occurring for some EP applications. During the 80’s, many efforts were dedicated to increasing thrust, in spite of the evident drawbacks associated with the high specific impulse of this thruster. Those attempts had to face the high power-tothrust ratio of FEEP (about 60 W/mN, vs. 35 W/mN of ion engines and 20 W/mN of SPT), and development was eventually abandoned. In the 90’s, the low thrust capabilities of FEEP have been rediscovered, thanks to the scientific community demand for modulable micronewton thrusters, and the development efforts have gained new momentum - this time, in the right direction. Today, the envisaged application domain of FEEP is the 1 µN - 1 mN range, were power-to-thrust is not a key factor, but low system mass and low propellant consumption are. The basic technology of the 80’s has been fully exploited, and much research is underway to refine the thruster design at engineering level and to scale down the subsystems (neutralizer and power conditioning and control electronics). For many commercial small satellites applications, FEEP is an attractive choice for such tasks as attitude control and orbit maintenance. Scientific missions may benefit of the accurate thrust control capabilities and high bandwidth of FEEP for drag-free control and formation flying1. The purpose of this paper is to show that small satellite attitude and orbit control can be performed with a FEEP system in an efficient way, meeting the spacecraft constraints (power, mass, pointing requirements). To this end, a simple study was carried out using the available FEEP performance and system data. The study included the following steps:
The whole procedure was then repeated for different reference satellites, with mass ranging from 100 to 1000 kg, and for different orbital altitudes in the 400 - 800 km range.
FEEP Thruster Model FEEP is presently developed by Centrospazio under ESA funding. The power and control electronics (FEEP Electronics Unit - FEU) is developed by LABEN. Much experimental data on FEEP emitter performance is available, especially in the 1 - 100 µN thrust range. Reference performance data for 5 mm slit length emitters are presented in Ref. 2. For the purpose of this study, it is assumed that the thruster be operated with cesium propellant, albeit other liquid metals or alloys may be used instead. The electrical efficiency of FEEP is about 98 %, vs. much lower values of all other electric engines (from x % for ion engines to 17 % for PPTs). This is due to absence of thermal losses, as ionization occurs in the liquid phase, and to the favorable ion extraction geometry, which minimizes ion impingement losses. FEEP has no moving parts, no pressurized vessels, no propellant feed lines. Measured minimum impulse bit, which is a key issue for fine point-
l = 7 cm 6 cm 5 cm
Thrust (mN)
1.6 1.5 1.4 1.3 1.2 1.1 1.0 0.9 0.8 0.7 0.6 0.5 0.4 0.3 0.2 0.1 0 3.5
4 cm 3 cm 2 cm 1 cm 0.5 cm 4
4.5
5
5.5
6
6.5
7
7.5
Emitter voltage (kV)
Fig. 1 - Thruster model: thrust vs. emitter voltage at constant accelerator voltage
8
ing, is as low as 5.10-9 N.s, vs. 1.10-4 N.s of PPTs3. The FEEP emitter is easily scalable to different thrust levels. With good approximation, thrust is a linear function of the length, l, of the ion emitting slit, at least for the slit lengths considered in this work, at equal total voltage. For the purposes of this work, it was assumed a maximum allowable thrust level of 1 mN. This is by no means an intrinsic limit of the FEEP concept, as emitter length is limited by manufacturing difficulties only; in fact, emitters with slit lengths up to l = 15 cm, corresponding to maximum thrust of about 3 mN, have been successfully operated in the past. Moreover, emitters may be easily stacked. However, the 1 mN ceiling was imposed to put an upper limit to power consumption, and to take into account the actual design of the FEU prototype. Emitter performance was extrapolated from l = 5 mm to different slit lenghts assuming constant linear thrust density at equal total voltage. Fig. 1 and fig. 2 show the thrust vs. voltage and power vs. voltage curves, respectively, for different slit lenghts. Thrust includes the (slight) reduction due to plume divergency. The accelerator voltage was fixed at Va = -3 kV; however, different distribution of total voltage between the electrodes would have negligible effect on the thruster performance, at least for our present purposes. Although accurate enough for this study, this model has a number of limitations. Power consumption includes the power conveyed by emitter current and the power loss due to ion current drained at the accelerator; it does not include neutralizer power, nor emitter heater power (needed to keep the emitter at 35 °C). However, these contributions are within 10 % of the power estimated using this model. Throughout this study, no use was made of the thrust modulation capability of FEEP. According to the very simple control strategy envisaged (see below), it was assumed to use the thruster at a constant operating point, which will be specified in the following for each case.
90 85 80 75 70 65 60 55 50 45 40 35 30 25 20 15 10 5 0
l = 7 cm 6 cm
Power consumption (W)
• setting up a thruster reference model; • choosing a reference mission; • setting up an analytical model of spacecraft orbital and attitude dynamics, using established models for perturbing forces and torques; • simulating the spacecraft dynamics, assuming a simple control law; • once verified that the propulsion system and control law chosen satisfied the mission requirements, evaluating thrusting frequency and duration and propellant consumption.
5 cm 4 cm 3 cm 2 cm 1 cm 0.5 cm
3.5
4
4.5
5
5.5
6
6.5
7
7.5
Emitter voltage (kV)
Fig. 2 - Thruster model: power vs. emitter voltage at constant accelerator voltage
8
A reference spacecraft: SAT 900 Mass: Main body dimensions: Solar panels: Moments of inertia:
900 kg 2.25 × 2.25 × 2.25 m3 2 × 4.5 m2 each Iroll = 2110 kg m2 Ipitch = 840 kg m2 Iyaw = 2160 kg m2 Orbit: 800 km, sun-synchronous (inclination i = 98.5 deg) Lifetime: 5 years Installed power (average on lifetime): 1.8 kW
Our reference spacecraft, SAT 900, is a medium-sized satellite in a sun-synchronous orbit. Several types of payloads make use of such vehicles: communications, remote sensing, scientific. It is not essential to our purpose to specify the payload of SAT 900; however, the attitude control requirements are typical of most communication payloads. They may assumed also to be not too dissimilar from the mutual pointing requirement of spacecraft in constellations.
Attitude control requirements Roll: Pitch: Yaw:
The geometry of SAT 900 is simplified enough to make calculations easy, yet still retains sufficient resemblance to real spacecraft of its class. Main parameters were chosen as average values from statistical data on satellite mass and power budgets4. SAT 900 features a cubic main body and two solar panel wings, hinged on opposite faces of the cube. Conservative assumptions were made whenever appropriate along this study. Reference attitude is with the plane of the wings parallel to the ram direction, in a such a way as to minimize drag and to get a significant fraction of incident sunlight. Use of NiH 2 batteries is assumed to cope with eclipse periods (35 minutes out of 101 minutes of orbital period). The characteristics of SAT 900 are listed in Table 1.
Yaw
5
0.2 deg 0.2 deg 0.4 deg Table 1 - SAT 900 data and requirements
Orbit Maintenance Compensation of residual atmospheric drag on LEO satellites is usually needed to avoid cumulative orbit degradation. The burden this poses upon the propulsion system may eventually result in a serious life limiting factor, whereas the payload may still retain significant survival margins. Therefore, most spacecraft may greatly enhance their operating lifetime by performing drag compensation. On a constellation spacecraft the drag compensation issue may become important even on a short time frame, as it may be needed to avoid orbit drifting in order to maintain the formation geometry accurately. In this case, drag is to be compensated on a more frequent basis. In general, another important orbit maintenance issue is raised by effect of the Earth oblateness. However, in a sun-synchronous orbit like that of SAT 900, orbital plane drift due to J 2 is a deliberately exploited effect. Cancelation of J2 effects may be achieved with out-of-plane thrusting, in a similar way as what is done in-plane for drag compensation. In its 900 km orbit, SAT 900 flies at 7.5 km/s against the atmosphere, whose average density is ρm = 7.24.10-14 11
14
Roll
4
7 10
13 2
Pitch
15
ROLL
8
6
SET 1
Roll
12
-
(7,8)
(3,4)
9
Yaw Pitch
+
3
1
SET 2
(9,10) (11,12)
16
Fig. 3 - SAT 900 layout, left; thrusters layout, right
PITCH + (2,5) (4,7)
(1,6) (3,8)
YAW +
-
(2,6)
(1,5)
(9,12) (10,11) (13,15) (14,16) (14,15) (13,16)
kg/m3. As a result, the spacecraft experiences an orbitaveraged drag Dm = 105 µN, computed considering a standard atmosphere model at moderate solar activity. If not counteracted, this continuous braking would result in an altitude loss of 35 km in 5 years. This may be acceptable for single-spacecraft missions, but will not be tolerable within constellations. Attitude Control This study was carried out assuming that the initial condition for the attitude control task of the FEEP system is nominal attitude, i.e. that attitude acquisition has been already performed. The maximum values of the perturbing torques included in the attitude dynamics model are as follows: • • • •
gravity gradient: 15 µN .m solar radiation pressure: 78 µN .m magnetic dipole: 43 µN .m atmospheric drag: 21 µN .m
Thus, in the most unfavorable case in which all disturbances act in the same direction with maximum magnitude, the resulting perturbing torque would be 157 µN .m.
Propulsion System for SAT 900 As of writing, much effort is underway to further reduce the mass of the subsystems, to improve the sealed container operation and to reduce the neutralizer power consumption. However, the FEEP system for SAT 900 was sized according to the present state of development of the system5, which is summarized below. This engineering qualification model has undergone functional tests at Centrospazio, is scheduled for an extended endurance test
Fig. 4 -Thruster/container assembly
at ESTEC in late 1997 and will be flown on a Space Shuttle GAS canister in 1999. FEEP subsystems The FEEP thruster includes the following subsystems: • emitter/container assembly • neutralizer • power and control electronics. Emitter/container assembly The thruster container is an integrated unit dedicated to hosting the FEEP emitter and providing all necessary mechanical and electrical interfaces. It is made of a sealed cylindrical reservoir with an openable lid, containing the emitter and the accelerator, the relevant high voltage isolators, electrical feedthroughs, and the thermal monitoring and control system. During the thruster preflight preparation phase, the emitter, filled with cesium, is sealed in the container under inert atmosphere. Cover is released, in orbit by means of a small paraffin actuator. A picture of the thruster/container assembly during thermal vacuum test is shown in fig. 4. Neutralizer The engineering qualification model is equipped with a simple thermionic neutralizer, consisting of a hot filament. This solution is not optimal from a system point of view, as the electrical power needed to keep the filament at the high temperature required for electron emission is quite high. In fact, this drawback is quite tolerable, due to the very low overall thruster power level. However, this compromise solution will be superseded by a different kind of neutralizer, presently under development at Centrospazio. This new device, based on field emission from arrays of microtips made with silicon micromachining techniques, operates at room temperature with a total power consumption per unit current of about 0.12 W/mA , that is about 2% of thruster power consumption. While mating with the cesium emitter will need some additional effort, the basic technology is well understood and already widely used in several technological applications (e.g., flat displays). Power and control electronics The power and control electronics (FEU - FEEP Electronics Unit) is developed by LABEN (Milan, Italy). In its present configuration, the system supports the operation of up to 4 thrusters, including full redundancy. Power and control functions are concentrated in a single box, 250 × 165 × 200 mm3, which is a plug-and-play, self standing unit. The FEU is interfaced to the satellite power bus and to the onboard computer (via a standard RS422 or MIL 1553B interface) from which it receives the thrust
command and, accordingly, provides the required voltages to the FEEP emitters, accelerators and neutralizers. Thrust is monitored and controlled sensing the current drained by the motor. Thermal control is also performed to maintain the propellant in the proper temperature range. In addition, the FEU provides telemetry data to the onboard computer. The system includes extended built-in tests providing real-time status of the propulsion system to the user, enabling onboard fault detection functions.
Satellite dynamics simulation The equations of orbital and attitude dynamics were numerically solved using a standard fourth order Runge-Kutta method. The drag compensation philosophy is aimed at minimizing the thrust-on time while keeping altitude loss within a narrow band. This is achieved by firing the emitters at full thrust (4 × 1 mN) for short durations, ending up in the firing frequency shown in table 3. Maximum computed deviation from reference altitude is 0.5 m.
Thruster locations The FEEP system for SAT 900 consists of 16 thrusters, driven by 4 FEU’s. Thrusters are arranged in pairs at each vertex of the spacecraft main body, as shown in fig. 3. Four neutralizer are placed on the edges of the main body, halfway between adjacent pairs of thrusters. This arrangement was chosen to provide full thrusting authority along all axes (except pitch axis, that is along the cross-track direction) with a high degree of redundancy. Thrusters are in planes perpendicular to the pitch axis, in such a way as to minimize the risk of plume impingement on the solar panels. Orbit maintenance is performed by tangential thrusting, using the thrusters located on the anti-ram side of the satellite (thrusters 1, 2, 13 and 16). All thrusters are fired together, producing a total thrust of 4 mN. Attitude control is performed with pairs of thrusters. Maximum torque along each axis is 4.5 mN.m.
Item
Thruster assembly Neutralizer FEU
Unit mass (kg)
Quantity
Total (kg)
0.4 0.1 6.6
16 4 4
6.4 0.4 26.4
Total dry mass (16 thrusters): 33.2 kg
Table 2 - FEEP System Mass Breakdown
Average drag acceleration: Altitude loss per orbit: Altitude loss in 5 years: Total ∆V to compensate for altitude loss: Thrust-on time: Time between firings: Total thrusting time: Propellant consumption (per thruster):
Table 3 - Orbit Maintenance
Attitude control strategy is also kept very simple. Angular deviation is allowed to build up until it crosses a predefined deadband (0.1 deg for roll and pitch, 0.2 deg for yaw); then, thrusters are fired at maximum thrust, decreasing and then reversing angular velocity, until deviation returns within deadband. This law, which is essentially of the bang-bang type, was found to produce a smoother control than a proportional one. Of course, this strategy is far from optimal; more sophisticated schemes may be devised which take into account the thrust modulation capability of FEEP. In consideration of the first approach character of this study, this control law was selected as a conservative case. Nevertheless, the law chosen has the favorable property of considerably reducing the thrust-on time with respect to proportional schemes, resulting in a reduced total thruster lifetime requirement. In fact, cumulative thrusting time was a major concern in designing of the propulsion system. To ensure minimum total thrust-on time, it was chosen to maximize the control torque by firing together the two available thruster pairs per axis and direction (e.g., torque on the positive roll axis is produced
ROLL + Thrusters Thrust-on time (s) Firing frequency (s-1) Total propellant consumption (kg)
PITCH -
(7,8 / 9,10) (3,4 / 11,12) 95 96 4.10-3 4.10-3 1.17 1.24
- 1.16.10-7 m/s2 1.365 m 35 km 18.5 m/s 5s 173 s 1154 h 40 g
+
YAW -
(4,7 / 9,12) (3,8 / 10,11) 30 36 9.10-3 8.10-3 0.81 0.88
Table 4 - SAT 900 Attitude Control
+
-
(2,6 /13,15) (1,5 / 14,16) 90 80 4.10-3 4.10-3 0.97 0.92
Roll angle (deg) Pitch angle (deg)
Time (s)
Yaw angle (deg)
Time (s)
Time (s)
Fig. 5 - Attitude angles vs. time
by thrusters 7, 8, 9 and 10). In this way, full 4.5 mN.m torque authority is exploited. Power consumption associated with this operation is less than 300 W. Cumulative thrust-on times for a 5 years mission are of the order of 15000 hours. A limitation of this scheme is that attitude control must be performed one axis at a time, due to the power consumption requirement. However, as the thrust-on durations are quite short for all axes with respect to the repetition periods, it is possible to devise a firing sequence such as to avoid having more than four thrusters firing at the same time. SAT 900 attitude control parameters are summarized in table 4. Figure 5 shows the computed behavior of controlled roll, pitch and yaw angles. All angles stay well within the specified boundaries, and attitude chatter is near absent.
Parametric analysis The spacecraft model setup for SAT 900 was simplified and adapted for a parametric analysis of orbit maintenance and attitude control over a wide range of satellite masses and orbital altitudes. The main assumptions for this study are: • orbits are sun-synchronous and circular;
• available power is 2 watts per kilogram of total satellite mass; • thrust level is chosen according to a realistic evaluation of the fraction of power available for the propulsion system as a function of satellite mass; • inertial properties are computed scaling the configuration of SAT 900, i.e. assuming the same configuration (two solar panel wings hinged on a cubic body). Orbital and attitude disturbance were computed taking into account the combined effect of all forces and forces on the scaled geometries. Orbital altitudes of 400 km, 500 km, 600 km, 700 km and 800 km were considered, with satellite mass ranging from 100 kg to 1000 kg. The attitude control requirement was set to 0.1 deg, all axes, for satellites up to 700 kg. This requirement, which is more demanding than what was required for SAT 900, can be met within the specified maximum power constraint as long as the satellite’s inertia is small enough; for larger satellites, the 0.2 deg value of SAT 900 was retained. Of course, larger satellites make possible thruster placement choices different than the one of SAT 900; for example, placing thrusters at tip of solar arrays, where the arm is much larger, would result in increased torque at equal power consumption. Therefore, the results of this study are much conservative. Table 5 summarizes the characteristics of the spacecraft considered.
Spacecraft mass(kg)
100 200 300 400 500 600 700 800 900 1000
Power (W)
Thrust per emitter (mN)
200 400 600 800 1000 1200 1400 1600 1800 2000
Control torque (mN.m)
0.4 0.5 0.6 0.7 0.8 1.0 1.0 1.0 1.0 1.0
Attitude requirement (deg)
0.865 1.363 1.872 2.404 2.960 3.931 4.139 4.327 4.500 4.661
0.1 0.1 0.1 0.1 0.1 0.1 0.1 0.2 0.2 0.2
Table 5 - Parametric study spaceraft data
The total thrusting time per orbit required to perform orbit maintenance is shown in table 6. Satellites of 300 kg or more in a 400 km orbit (grayed-out in table 6) would require a thrusting time longer than orbital period (101 min); therefore, FEEP would not be suited to that operation, unless more thrust could be made available. In that case, obviously, power consumption would become unacceptably high for such small satellites. On all other spacecraft, especially those at higher altitudes, orbit maintenance is easily performed. Propellant consumption for orbit maintenance (fig. 6), which is less than 4 kg in all cases, drops to below 1 kg for orbits at or above 600 km, irrespective of mass. Attitude control is performed very satisfactorily on all spacecraft considered. Propellant consumption, shown in fig. 7, doesn’t vary much with altitude. In most demanding cases, i.e. for larger spacecraft, total propellant consumption for 5 years attitude control is 8 kg, that is 500 grams per thruster. Total propellant mass for attitude control and orbit maintenance is shown in fig. 8.
Spacecraft mass(kg)
400
100 200 300 400 500 600 700 800 900 1000
58.8 86.2 102.9 114.3 122.3 115.5 132.9 150.2 167.3 184.3
Fig. 9 shows a comparison between the FEEP system mass, including all hardware and propellant, and a traditional system with four reaction wheels and six hydrazine thrusters for desaturation. Traditional system mass was assumed to be 0.1 kg per kg of satellite mass. The advantages of FEEP for masses exceeding 400 kg are evident. For larger satellites, several tens of kilograms may be saved. It shall be noted that this analysis does not forget any “hidden” equipment mass: that is, no extra power generation plant is required, as the satellite model was sized taking into account realistic mass budget for all subsystems, including a power plant (panels, batteries, regulator) tailored for supporting both FEEP and all onboard systems, payload included. In addition to the large mass savings, the all-thruster system has also a number of other advantages: • thruster operation is completely vibration-free, due to the absence of moving parts and to the very small impulse bit;
Orbital altitude (km) 500 600
13.5 19.8 23.6 26.2 28.1 26.5 30.5 34.5 38.4 42.3
4.1 6.1 7.2 8.1 8.6 8.1 9.4 10.6 11.8 13.0
700
800
1.7 2.5 2.9 3.3 3.5 3.3 3.8 4.3 4.8 5.3
0.9 1.3 1.6 1.8 1.9 1.8 2.1 2.4 2.6 2.9
Table 6 - Orbit maintenance: total firing time per orbit (minutes)
4
10 400 km
3.5
400 km 2.5
500 km
2
600 km
1.5
700 km 800 km
1
Propellant mass (kg)
Propellant mass (kg)
500 km
8
3
600 km 700 km
6
800 km 4
2
0.5 0
0
100
200
300
400
500
600
700
800
900
1000
100
300
400
500
600
700
800
900
1000
Fig. 7 - Attitude control propellant consumption
Fig. 6 - Orbit maintenance propellant consumption 110
25
400 km
15
400 km
100
500 km
90
600 km
80
700 km
70
700 km
60
800 km
50
Reaction wheels & desat. thrusters
AOCS Mass (kg)
20 Propellant mass (kg)
200
Spacecraft mass (kg)
Spacecraft m ass (kg)
800 km 10
500 km 600 km
40 30
5
20 10
0 100
200
300
400
500
600
700
800
900
1000
Spacecraft m ass (kg)
100
200
300
400
500
600
700
800
900
1000
Spacecraft m ass (kg)
Fig. 8 - Total propellant consumption
Fig. 9 -AOCS mass comparison: FEEP vs. reaction wheels
• occasional slewing maneuvers can be easily performed in very short time, at the only expense of an additional power consumption for a short time. This may turn out to be quite useful on communication spacecraft in constellations, e.g. for re-configuration maneuvers following in-orbit replacement of spent satellites; • unlike wheel-based systems, FEEP provides orbit maintenance capabilities.
Investigations underway at Centrospazio include the elaboration of a control strategy to exploit thrust modulation under a maximum instantaneous power consumption constraint to reduce total thrusting time.
Conclusions This simple, conservative simulation study leads to the following conclusions: • an all-thrusters, no wheels attitude and orbit maintenance system based on FEEP is feasible with the presently available technology; • mass savings resulting from a FEEP AOCS in place of traditional systems are most attractive for satellites exceeding 400 kg. Mass reductions of 30 - 60 kg are possible for satellites of 700 - 1000 kg; • application on small satellites is strongly dependent on the specific configuration of the spacecraft.
References 1. H. Klotz, H. Strauch, W. Wolfsberger, S. Marcuccio, C. Speake, “Drag-Free, Attitude and Orbit Control for LISA”, Proc. ESA/ESTEC 3rd International Symposium on Spacecraft Guidance, Navigation and Control, ESA SP-381, ESTEC, Noordwijk, The Netherlands, 1996. 2. Marcuccio, S., Genovese, A., Andrenucci, M., “Experimental Performance of FEEP Microthrusters”, Proceedings of the 3rd International Symposium on Space Propulsion, Beijing, China, 1997. 3. Cassady, R. J., et al., “Pulsed Plasma Thruster Systems for Spacecraft Attitude Control”, Proc. 10th AIAA/USU Conference on Small Satellites, 1996. 4. Wertz, J.R., Larson, W.J., Space Missions Analysis and Design, Kluwer, 1992. 5. Marcuccio, S., Genovese, A., Andrenucci, M., “FEEP Thrusters: Development Status and Prospects”, Proceedings of the 2nd European Spacecraft Propulsion Conference, ESA SP398, ESTEC, Noordwijk, The Netherlands, 1997.
