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Flexible variable-specific-impulse electric propulsion systems for planetary missions
Acta Astronautica 59 (2006) 931 – 945 www.elsevier.com/locate/actaastro
Flexible variable-specific-impulse electric propulsion systems for planetary missions E. Chestaa,∗ , D. Estubliera , B. Fallisa , E. Gengembrea , J. Gonzalez del Amoa , N. Kutufaa , D. Nicolinia , G. Saccocciaa , L. Casalinob , P. Dumazertc , M. Prioulc , J.P. Boeufd , A. Bouchoulee , N. Wallacef , G. Nocig , M. Bertih , M. Saverdih , L. Biagionih , S. Marcuccioh , A. Cadioui , F. Darnoni , L. Joliveti a European Space Agency, ESTEC, TOS-MPE, Keplerlaan 1, Noordwijk aan Zee, The Netherlands b Politecnico di Torino, Dipartimento di Energetica, Corso Duca degli Abruzzi 24, Torino, Italy c SNECMA Moteurs, Service Propulsion Plasmique, Villaroche Nord, Moissy-Cramayel, France d CNRS-CPAT, Université P. Sabatier, 118 Route de Narbonne, Toulouse, France e GREMI Laboratory, Université d’Orléans & CNRS, 45067 Orléans Cedex 2, France f QINETIQ, Space Department, Y72 Building, Farnborough, Hampshire, UK g LABEN, Divisione Proel Tecnologie, Via Albert Einstein 35, Campi Bisenzio, Firenze, Italy h ALTA S.p.A., Via A. Gherardesca 5, Ospedaletto, Pisa, Italy i CNES, Toulouse Space Centre, 18 Avenue Edouard Belin, 31401 Toulouse Cedex 9, France
Available online 3 October 2005
Abstract Variable-specific-impulse electric propulsion systems can provide important advantages and cost reductions to any scale of planetary mission. The first part of this paper describes their potential benefits, with special emphasis on mission analysis aspects. The second part identifies the physical limitations of traditional electric propulsion technologies and presents some preliminary development results achieved within the frame of ongoing European activities. © 2005 Published by Elsevier Ltd.
1. Introduction ∗ Corresponding author.
E-mail addresses: [email protected] (E. Chesta), [email protected] (L. Casalino), [email protected] (M. Prioul), [email protected] (J.P. Boeuf), [email protected] (A. Bouchoule), [email protected] (N. Wallace), [email protected] (G. Noci), [email protected] (L. Biagioni), [email protected] (L. Jolivet). 0094-5765/$ - see front matter © 2005 Published by Elsevier Ltd. doi:10.1016/j.actaastro.2005.07.053
Electric propulsion systems based on traditional technologies, hall effect thrusters (HETs) or gridded ion engines (GIEs), are well known to substantially enhance (and often to enable) scientific missions with demanding performance requirements. The benefits usually consist in increased payload capability (or decreased launch cost), enhanced mission flexibility
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(including launch windows enlargement and reduced trip time for long-distance targets), and extended S/C operating life. Even more consistent benefits can be achieved using homogeneous or modular variable-specificimpulse electric propulsion systems. In fact, in this case a single-staged full-electric spacecraft could effectively perform different mission phases, even though based on conflicting performance requirements: high-specific impulse (Isp ) during interplanetary cruise (heliocentric segment) and high thrust (T ) for Earth escape (geocentric segment from a negative C3 launch) and for final orbit acquisition around the target planet (planetary segment). The mission would therefore result considerably simplified at system level, and it would be possible to share most of the propulsion subsystem equipments with consistent cost reductions and enhanced reliability. Advanced simulation tools have also shown the possibility to achieve minimum propellant trajectories by adopting variable-specific-impulse strategies. In the heliocentric segment this can be achieved either independently from the available power (it is usually convenient to increase Isp as the S/C mass diminishes because of propellant consumption, or to adjust Isp and T to allow for optimal gravity assist encounters), or as a function of the available power when solar energy is used (typically, Isp grows together with the available power for inner solar system missions, while for missions to the outer solar system Isp is usually kept quasi-constant at its initial value while T decreases). In the geocentric or planetary segment, the general rule is that Isp is better decreased at the apsides of the orbit, when a higher thrust can be employed more efficiently. Additional benefits in terms of recurring costs reduction can be achieved with development programs based on a “modular approach”, at equipment level (aimed at employing similar subcomponents) or at subsystem level (whole propulsive block with shared equipment). The ideal situation would be of course to have an entirely homogeneous system, based on a single type of thruster technology capable of a wide range of performances. The higher Isp capabilities that are necessary to successfully perform the interplanetary part of the cruise are also likely to find convenient application for attitude control and station keeping of GEO telecom satellites and Earth observation missions, and the higher T necessary to reach
the escape orbit during the geocentric segment can be used in a similar way for orbit raising and quick transit through Van Allen belts. The European Space Agency and CNES have initiated several activities to quantify and overcome the performance limitations of conventional EP systems. In particular, the possibility to profitably employ readily available concurrent technologies on the same electric propulsion hybrid platform is currently under assessment for short-term applications. At the same time, two activities for the development of a double-stage hall-effect thruster (DSHET) prototype are running, and the concept selection phase has already been successfully achieved. If these new thrusters will prove effective, they will be the ideal candidates for both scientific and commercial missions on the medium term. For the long term, new concepts and improvements are currently under consideration for possible future research and development programs.
2. Mission benefits of variable-Isp electric propulsion systems In typical electric propulsion systems, the power of the electric source is proportional to the specific impulse and to the thrust level (if constant thruster efficiency is assumed). A high Isp reduces the propellant requirement, whereas a high T reduces the trip time. When a constant-Isp operation is considered, the mission designer selects the specific impulse and the thrust level to obtain the best compromise between mission payload and time length. Optimization methods are used to search for minimum-propellant trajectories. An increasing number of revolutions around the main body reduces the propellant consumption and a time constraint is necessary to make the optimization problem meaningful. In this case, the mission designer is able to obtain the optimal value of the specific impulse to accomplish the trajectory in the assigned time, once the available power is given; moreover, the optimal power level can be determined when the specific mass of the electric source is assigned. A thruster that can exploit a wide range of values for the specific impulse and the thrust could be used both for the escape from Earth and the interplanetary portion of the flight. However, a smaller range of operational Isp is also useful to improve performance, as
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it allows a better management of the available power during the interplanetary trajectory. Most of the benefit is obtained by increasing Isp as the spacecraft mass diminishes during the flight; the optimal Isp is inversely proportional to the spacecraft mass in the absence of gravitational losses. However, the control law is modified by the presence of this kind of loss; Isp is reduced when the spacecraft is in positions where a large thrust level is advisable (for instance, at the apsides of the orbit), whereas it is increased when the thrust cannot be usefully exploited (for instance, in correspondence to the coast arcs of a constant-Isp trajectory). A variablespecific-impulse engine, without bounds on the attainable values of Isp , is always switched on, and the specific impulse continuously varied. A time constraint is also required in this case to avoid a meaningless solution with very high specific impulse and infinite trip time. Coast arcs again appear when the constraint is applied to the thrust-on time instead of constraining the overall mission time or when bounds on the specific impulse are imposed. An indirect optimization method is used to obtain the optimal trajectories with either constant or variable specific impulse, to evaluate the benefit of variable specific impulse during the heliocentric leg of interplanetary trajectories. A point-mass spacecraft with variable mass m is considered. The patched conic approximation is adopted and the time spent inside the planets’ sphere of influence is neglected; therefore, the equations of motion are integrated only in the heliocentric reference frame. The state equations are: dr/dt = v,
(1)
dv/dt = g + T/m,
(2)
dm/dt = −T /c,
(3)
where r and v are the spacecraft position and velocity vectors, g is the gravitational acceleration (an inversesquare gravity field is assumed throughout), T is the engine thrust and c =g0 Isp is the exhaust velocity. The thruster exploits solar electric propulsion. The available power Pa is related to the nominal power at 1 AU (P1 ) by means of a function, which depends on the mission (in general, Pa is proportional to the inverse square distance from the Sun). Variables are normalized using the radius of the Earth’s orbit, its corresponding circular velocity and the ini-
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tial mass of the spacecraft as reference values. Results are presented in non-dimensional form and can be applied to any level of S/C initial mass. The reference units for thrust and power correspond to 5.93 N and 176.6 kW per each 1000 kg of S/C initial mass. A power penalty coefficient = / = 10, which corresponds to = mg /P1 = 40 kg/kW for a thruster with constant = 0.7, has been assumed to compute the payload mass mu = mf − P1 / = mf − P1 . In order to allow for direct comparison of variable and constant Isp performances, simpler boundary conditions than in actual missions are assumed. The spacecraft always leaves the Earth’s sphere of influence on a parabola, i.e., with zero hyperbolic excess velocity. The procedure could however provide the optimal value of the hyperbolic excess velocity, once the characteristics of the engine used to escape from Earth were assigned. The arrival at the target planet with zero hyperbolic excess velocity is considered, as it appears to be the most general condition for the rendezvous. The position of the arrival planet is instead left free, which means that the spacecraft is inserted into the same orbit of the planet around the sun, but the target planet is in a different position. Thus, mission performance is not affected by the phase angle between the Earth and the target planet. This “time-free” trajectory can be flown every year departing on the same day, and the rendezvous mission can be easily obtained in those years when the target position is favorable. The payload mu is the performance index which is maximized; a constraint on the trip time is also imposed. Details on the optimization method can be found in Ref. [1]. The trajectory is controlled by the thrust magnitude and direction. The desired thrust level T = 2P /c is obtained by selecting the exhaust velocity and the engine input power taking operational constraints into account. The input power is always limited by the availability of solar power, while the exhaust velocity may be bounded or, in the limit case, constant. The boundary value problem, which arises from the application of the theory of optimal control, is solved by a procedure developed at the Politecnico di Torino and based on Newton’s method. Numerical results for missions to Venus, Mercury and Mars are presented in Figs. 1–3. Variable- and constant-Isp trajectories are compared in terms of payload and average exhaust velocity
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0.42
1.5 cm mu
0.021
1.4
0.022
0.023
0.024
4
Fig. 1. (a and b) Mission to Venus: performances for different trip times (465 and 581 days, respectively).
cm = g0 Ism , where the average specific impulse Ism is computed as the ratio of the total impulse to the consumed propellant. The trip time has been fixed at = 8 (about 465 days) in Fig. 1a, = 12 (698 days) in Fig. 1b, and = 10 (581 days) in Figs. 2 and 3. Figs. 2b and 3b show the control and thrust history for missions with nominal power P1 = 0.022 and 0.013, the latter corresponding to 3.4 kW for a 1500kg S/C of the Mars Exobiology mission class [2]. It is well known that an optimal value of the power exists for each trip time. Observing Figs. 1a and b it can be seen that the performance improves for the same value of the available power as the mission time length grows, because the same total impulse can be provided with lower thrust and, conversely, with higher specific
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Fig. 2. (a and b) Mission to Mercury: performance and control history for P1 = 0.022.
impulse and lower propellant consumption (the position of the maximum payload shifts to the left when grows, and the optimal power is lower in the case of variable specific impulse). Another very important feature is that the benefit of variable-Isp trajectories over constant-Isp trajectories is especially significant for shorter trip times. In this case the larger freedom coming from the additional control capability allows the designer to obtain a larger payload by reducing the losses which are related to the magnitude and direction of the thrust, whereas the average specific impulse is almost the same for constant and variableIsp thrusters. For instance, the specific impulse is reduced and higher thrust can be used where it is more beneficial for the reduction of gravitational losses. Fig. 2b shows that the larger power that is available in a
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945 1.4
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Fig. 3. (a and b) Mission to Mars: performance and control history for P1 = 0.013.
mission to inner planets is mainly used to increase the specific impulse, whereas the thrust level is almost constant (similar plots are obtained for the Venus mission). The maximum payload is again obtained with a reduced power with respect to the constant-Isp case. Variable- and constant-Isp trajectories use the excess power that comes from the proximity of the Sun in a quite different manner: in the first case, the specific impulse can be increased, whereas the optimal thrust is almost constant; on the contrary, when c is constant, the excess power can only increase the thrust; coast arcs appear where the thrust would be inefficiently used. The optimization procedure can also deal with planetary flybys, which can be usefully exploited to reduce the V requirement. During the
route to Mercury, Venus can provide for instance a remarkable gravity assist. Low Isp is exploited in this case at departure in order to permit the large thrust necessary to encounter Venus in the most favorable position. Also in the case of a mission to Mars a higher payload mass is obtainable with a variable Isp strategy (Fig. 3a). In this case the spacecraft moves towards an external planet and, during the variableIsp operation, the reduction of the available power mainly affects the thrust level, whereas the exhaust velocity is kept almost constant (Fig. 3b). The average specific impulse is however larger for the variable-Isp thruster, which can compensate for the lower average thrust by properly throttling the engine when required. The specific impulse used in these calculations ranges over quite a large span, and is allowed to assume values at the limits of currently available technology (c = 2 corresponds to about Isp = 6000 s). It is possible however to use the program to compute and compare the benefits obtainable in case of unbounded Isp , bounded Isp , dual Isp and constant Isp . In Fig. 4 the results are obtained in the case of a direct mission to Mercury. It is quite interesting to notice how the
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variable-Isp thruster permits a 10% mean reduction of the propellant consumption. This value is slightly lower (about 9%) if the unavoidable upper and lower limits for the Isp are imposed at reasonable levels, but it can double in special cases (like for long missions with gravity assist and low total propellant mass). It has to be recognized that up to 80% of the achieved propellant mass savings could be obtained using dualmode thrusters, considerably simpler to develop and qualify.
3. Performance limits of traditional EP technologies Typical control parameters, for EP thrusters, are the inlet propellant mass flow rate and the applied voltage. For power-limited operations, these two parameters cannot be varied independently. In fact, neglecting losses due to incomplete propellant ionization or to beam divergence, the thrust power and the electrical power can be written as Pt = 21 mc ˙ 2, P=
(4)
Isp T g 0 Pt q = = IV = mV ˙ , 2 M
(5)
where I is the beam current, the thruster efficiency, and V the discharge voltage in the case of HETs, the beam voltage for GIEs and the emitter voltage for FEEPs. The exhaust velocity is in all three cases: c=
2
q V. M
(6)
Specific impulse and thrust-to-power ratio will then show reversed proportional behavior with respect to the square root of the applied voltage: c Isp = = g0 mc ˙ T = = P P
2q √ V, g02 M
(7)
2m 1 √ . q V
(8)
HETs, GIEs and FEEPs present optimal efficiency while operating around nominal points with relatively low, high and very high V , respectively. For this reason, HETs are usually considered for relatively high T applications, while GIEs and FEEPs for relatively high and very high Isp applications. It is a worthwhile exercise to identify what are the physical boundaries (mostly related to the ion beam extraction process) for thrust and specific impulse encountered when operating these three different technologies at variable power levels, and if possible what are the alternative control parameters that can permit to operate them within an extended range of performances. 3.1. Hall-effect thrusters In the case of Hall-effect thrusters, a recent activity performed by the GREMI Laboratory (Université d’Orléans) and supported by ESA in the frame of the Aurora programme has addressed the topic of “Dual mode high power SPT operation” [3]. This work has shown that the ability of single-stage HETs to operate in dual-mode increases with the power level. In fact, the minimum value of propellant flow rate in the thruster channel is the one that assures a neutral ionization time (roughly inversely proportional to the electron density and hence to the propellant flow rate) significantly shorter than the transit time of the neutral atoms in the ionization layer. This limit, especially at low power levels, translates directly in a constraining upper boundary for the discharge voltage (for the reachable specific impulse). When the thruster is designed to operate at high power levels this problem is mitigated, since the mass flow density for effective ionization is inversely proportional to the channel width. The maximal propellant flow rate value is on the contrary only limited by the maximal thermal fluxes transferable to the thruster structure for a given design (less restrictive constraint), or to the maximum magnetic field strength obtainable when up-scaling the power [4]. Experimental tests performed in collaboration with RIAME MAI (Moscow) on a 4–5 kW laboratory model SPT 140 have proven that a HET designed for discharge voltages lower than 500 V can operate during short periods in a high specific impulse mode with relatively good efficiency: the discharge voltage was increased up to 1000 V, experiencing
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0.65
2.31 mg/s 2.77 mg/s 3.23 mg/s 3.70 mg/s 4.16 mg/s 4.62 mg/s 5.08 mg/s 5.55 mg/s 6.01 mg/s 6.47 mg/s P = 2400 W P = 2000 W P = 1600 W P = 1200 W
Discharge efficiency
0.6
0.55
0.5
0.45
0.4 200
300
400
500
600
700
800
900
1000
Discharge voltage (V) Fig. 5. SNECMA PPS䉸 -1350-G: efficiency versus discharge voltage for variable flow rate/power [5].
only minor problems related to the non-optimized design of the thruster at high power levels (overheating of the coils with consequent magnetic excitation saturation, stochastic electrical breakdowns). Similar successful results have also been obtained by the Groupement de Recherche CNRS/CNES/SNECMA/ ONERA n.2232 “Propulsion à Plasma pour Systèmes Spatiaux” on a SNECMA PPS䉸 -1350-G engineering model [5], whose performances were explored in a wide range of operating points (Fig. 5), reaching Isp levels above 3000 s with good efficiency (as expected, at constant power levels the efficiency decreases when the discharge voltage is increased). Another interesting input from [4] is that both flexibility and performances increase when the thruster unit power (i.e. size) increases. A major concern is obviously related to the possible impact of similar dual-mode operations on thruster lifetime. The trends in terms of thruster power mentioned above suggest that this problem could also be reduced at higher thruster size and power. The assessment of this critical point will be hopefully the object of future dedicated activities.
3.2. Gridded ion engines In the case of gridded ion engines (GIEs), the most stringent limitation concerns the maximum value of propellant flow rate, restricted by space-charge effects related to the selected thruster optics geometry. In fact, Child’s law fixes a fundamental limit on the current I that can be drawn across a gap distance d between two plates of surface A by a given potential difference V : 40 2q V 3/2 A I =A = 2 V 3/2 , (9) 2 9 M d d where 0 is the permittivity of free space and is Child’s constant. This formula is applicable to GIEs ion extraction optics, but in this case the effective acceleration length d must take into account also the screen grid thickness and the variable plasma sheath bowed shape. For increasing values of plasma density (due to increased mass flow rate or more negative accelerator voltage), the plasma sheath pushes closer to the screen grid and becomes flatter, causing direct impingement of the under-focused beam on the
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230 mN
Measured Accel Grid Current (mA)
25
20
15
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1850V, Y=1.627X + 1.371X + 0.318, IB limit = 3.4A 2
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5
1500V, Y=2.755X - 1.522X + 1.797, IB limit = 2.75A
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2
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Fig. 6. QinetiQ T6 experimental characterization and identification of maximum beam current [6].
accelerator grid. This is called sometimes the “perveance limit” (the perveance being the ratio I /V 3/2 ), and, for a given thruster design, it is the main limiting factor to the achievable maximal thrust and to the employment of large power-saving negative accelerator voltages (a part from considerations related to higher
energy charge-exchange ions impingement when no deceleration grid is present). An experimental characterization of this limit for a T6 thruster by QinetiQ, part of the “Bepi Colombo Technology Demonstration Activity” [6], is presented in Fig. 6. The maximum beam current for each beam
Specific Impulse [s]
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7000 6800 6600 6400 6200 6000 5800 5600 5400 5200 5000 4800 4600 4400 4200 4000
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7.0
7.5
8.0
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10.0
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Fig. 7. FEEP specific impulse vs total voltage at different values of accelerator voltage (Courtesy: ALTA).
voltage is limited by the onset of direct impingement (accelerator grid current). The reciprocal phenomenon occurs for decreasing values of plasma densities (very high beam voltages and low mass flow rate): in this case the plasma sheath retreats from the screen grid and becomes more concave, such that the trajectories of some ions cause them to crossover and impinge on the accelerator grid. This “crossover limit” is less restrictive than the perveance limit (the impingement current increase with beam voltage is smoother), and can easily be overcome if very high specific impulse is desired by adopting thruster geometries with larger grid separation.
for electron quantum tunneling [7]:
3.3. Field-emission propulsion
where Vs is the extrapolated starting voltage and C3 a constant taking into account again the geometry of the problem, the temperature of the propellant and the average number of emission sites. In both cases, the emitted current and the total applied voltage are intrinsically correlated and cannot be changed independently. The emitter temperature is controllable, but usually it is maintained at the minimum value necessary to keep the propellant liquid. An interesting feature of this kind of thrusters is the possibility to employ very negative accelerator voltages (to operate at very low ratios of the net-to-total accelerating voltage). This is possible because the emission sites
FEEP thrusters can operate only when the electric field applied between the accelerator and the emitter electrode is sufficiently strong to counteract the surface tension of the liquid metal propellant and initiate ion field emission. For this reason, they are traditionally considered only for very high Isp and low T applications. For a given thruster geometry, the mass flow rate cannot be used as a control parameter. In fact, experimental studies have demonstrated that, in low current regimes, the emission law is exponential and closely related to the Fowler–Nordheim equation
I = C1 AV 2 e−C2
3/2
/V
.
(10)
is here the propellant work function; the constants C1 and C2 depend on the thruster geometry and on the Fermi energy measured from the bottom of the conduction band. In high current regimes, the emitter obeys a law determined by space-charge limits [8]: I = C3 [(V /Vs )3/2 − 1],
(11)
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are discretely localized, the liquid meniscus shape is only locally and microscopically affected by the applied electric field (no appreciable changes in the acceleration length are produced), and the chargeexchange ions production rate is much lower (the neutral vapor density is proportional to the emitter temperature, relatively low if propellant with low melting point like cesium are used). If the accelerator negative voltage (instead of the positive emitter voltage) is increased, the extracted current increases according to (11) with reduced energetic cost: P = I V e = C4 (Ve + |Va |)3/2 Ve .
(12)
Of course this involves consequently reduced Isp operation (Fig. 6). As for GIEs, the maximum achievable Isp is only power limited and can be increased for a given power level by employing higher emitter voltage and larger electrodes separation. The maximum achievable T is strictly related to the emitter geometry and the number of emission sites. Based on the specific characteristics of this technology, the possibility to develop high-current prototypes for future highpower applications deserves accurate consideration (Fig. 7).
4. Development of hybrid EP systems (short term) While HET and GIE technologies have complementary performances and very distinct technical characteristics at thruster level (as shown in previous paragraph), on the contrary they present important similarities at system level. Their possible synergies can be profitably employed to solve requirement conflicts on planetary scientific missions and telecommunication satellites (e.g. @Bus), and are currently under assessment in the frame of the fast-track GSP activity “hybrid (coupled plasma/ion) electric propulsion system feasibility study”. The first goal of this activity is to perform an exhaustive preliminary assessment at system level between HET and GIE technologies, aimed at identifying all possible commonalities and incompatibilities. The second goal is to execute a schematic design of two complete hybrid electric propulsion systems
(HEPSs), one for a GEO Telecom application and the other for an interplanetary scientific mission, using if possible the same hybrid electric propulsion module (HEPM) inclusive of one HET, one GIE and the relevant sub-system components. This approach is relatively innovative, since past planetary mission studies involving combinations of different EP technologies, mainly conducted by NASA [10], usually considered S/C composed of jettisonable independent homogeneous propulsive stages. This alternative solution is viable and provides some mission advantages related to a lighter S/C after separation of the HET stage, but it is globally less attractive than the hybrid platform because the higher thrust jettisoned engines cannot be used to optimize the interplanetary trajectory or final planetary orbit acquisition (an additional chemical propulsive stage would be necessary); in addition, the S/C would be more complex and failure risks and costs would be higher. The commonality assessment first iteration has provided some interesting preliminary results, showing that a modular approach at subsystem component level can generate important cost savings, and that a flexible HEPM composed of hybrid common elements is achievable (modular approach at propulsive system level). An example of interesting solution for a hybrid Power Processing Unit is presented in Fig. 8 (courtesy QinetiQ). Electric propulsion systems based on flexible hybrid modules operating well-known thruster technologies in their ideal performance envelopes are likely to prove very competitive on the satellite market in the near term. If the results of this study are positive, next step could be a consequent experimental demonstration requiring the development of specific hardware to breadboard level.
5. Development of double-stage Hall effect thrusters (medium term) The highest benefits, in terms of cost and mass savings per mission, can be achieved when the S/C is equipped with a homogeneous variable-specificimpulse propulsion system supporting a variable number of identical thrusters (instead of hybrid modules as considered earlier). As demonstrated earlier, EP
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1
Heater/ Magnet 2
Anode/ Keeper/ Magnet
3
4 5 6
Accel. Grid
7
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8
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9
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Beam/ Discharge
Fig. 8. Potential hybrid PPU concept (QinetiQ).
systems based on currently available technology are unable to operate efficiently for long time in highly variable-Isp mode. In the medium term, it is possible to envisage the development and qualification of new thruster prototypes directly or indirectly derived from the existing ones, thus not requiring complex technology break-throughs. An extremely promising concept in this field is the DSHET, designed to achieve at least a partial decoupling of what are called ionization and acceleration regions in traditional stationary plasma thrusters. The expected performance improvements, with respect to traditional HETs, concern mainly an enhanced ionization process in the first stage and a finer control of the acceleration phase, with consequent benefits on the efficiency and on the beam divergence, which can be dramatically reduced if the magnetic field is optimized and capable to focus the ions in the acceleration stage (with consequent increase of thruster life capabilities and decrease of near-wall interactions and discharge noise). The prac-
tical realization of this separation can be obtained in very different ways. The ionization process is closely related to the electron current in the ionization region; in order to reduce the power dissipated by the electrons while crossing the acceleration stage, two solutions are possible [11]: the plasma source can be designed to require the lowest possible electronic current, and/or part of the electrons can be injected directly into the ionization zone. Two parallel assessments recently performed by independent industrial consortia under ESA sponsorship have identified two very promising concepts. 5.1. DSHET based on Galathea confinement The first concept, developed by SNECMA [12] in collaboration with CPAT, QinetiQ and ALT A, is based on the Russian [13] prototype engine SPTMAG [14], and is schematically represented in Fig. 9. The ionization stage is conceived in order to
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E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945 Inductive ionization stage
‘ Classical ’ acceleration stage
Typical ion trajectory
anode Xe
Myxina
Conducting wall
Magnetic field lines
Fig. 9. Schematic of SNECMA SPT-MAG thruster [11].
maximize the electron residence time and to optimize ion injection into the acceleration stage. The magnetic field topology can be classified as a (half) Galathea magnetic trap, based on a single current-carrying conductor completely imbedded in the plasma (mixina). Thanks to this configuration, it is possible to create a magnetic well in front of the acceleration stage inlet. In addition, a higher voltage is applied to the mixina surface and to the ionization chamber walls with respect to the anode of the acceleration stage, thus creating a potential well where ions oscillate in their motion towards the zero-B region. The imaginary magnetic line passing through the zero-B point and intercepting the anode is also the minimum equipotential line, it is called “separatrix”, and its downstream portion can be considered as the border between ionization and acceleration stage. The thruster functional analysis can be performed using a quasi-neutral hybrid model where electrons are described as a fluid and ions and neutrals as macro-particles [16]. An example of simulated electric potential distribution and ion trajectories obtained with this model is presented in Fig. 10a and b. The design of a 3 kW prototype is currently in progress (Fig. 11). The possibility to include an emitting internal electrode (hollow cathode) in the ionization stage has been carefully assessed and finally rejected [17]. The prototype will be manufactured
Fig. 10. (a,b) SPT-MAG internal modelling [16].
Fig. 11. 3D external representation of SNECMA 3 kW SPT-MAG concept [15].
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
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Fig. 12. Schematic representation of LABEN-ALTA DSHET concept [18].
by SNECMA and fully tested in ALTA facilities next year. 5.2. DSHET based on double-peaked magnetic field and semi-active intermediate electrode The second concept, developed by LABEN/Proel in collaboration with ALTA, is schematically presented in Fig. 12. The separation between ionization and acceleration stage is achieved in this case through a doublepeaked magnetic field [18]. The specific topography of the magnetic field in the two stages will be independently controlled and experimentally optimized, but it is probable that a configuration with reversed radial polarity will be finally adopted in order to assure a complete decoupling of the two processes. The intermediate electrode is a semiactive thermionic device (axis-symmetric impregnated tungsten coil or multiple-dispenser), whose emission capabilities increase if its working function is lowered. In this way, part of the electronic current necessary to the ionization process is directly injected into the ionization stage, increasing current utilization and reducing power dissipation and electron losses to the wall. A possible magnetic topography has been simulated with Quickfield 4.3 and is reported in Fig. 13a, with axial and radial magnetic field components in the middle of the channel enlarged in Fig. 13b [19]. A 3D
representation of the thruster internal configuration is presented in Fig. 14. This prototype will be manufactured by LABEN/Proel and also tested in ALTA facilities next year.
6. Future research (long term) On the long term, the possibility to upscale at very high power levels the technologies presented in previous paragraphs will be considered, together with the analysis of alternative and more innovative solutions. Some original ideas have emerged during the concept and technology assessment phase of the activities presented. These concepts span from double-stage thrusters with high-frequency (capacitive or inductive or helycon) plasma sources, to multi-staged or segmented thrusters with channels presenting axially variable characteristics (dimensions, materials and electric potentials), passing from cusped magnetic field confinement in the ionization stage to magnetic nozzles for the acceleration stage, including also intermediate heating stages based on ion cyclotron or electron cyclotron resonance. The potential performances of these concepts are not easily predictable without dedicated experimental activities, which are relatively expensive and hardly achievable in parallel.
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Fig. 13. (a and b) LABEN-ALTA DSHET [19]: magnetic induction lines distribution and magnetic flux density.
7. Conclusions The utility and cost-effectiveness of variable-Isp electric propulsion systems for planetary missions have been discussed and demonstrated. The physical limitations of traditional EP technologies have been shown. Reliable solutions, in the short term, can be provided by hybrid modular systems operat-
ing well-known HET and GIE technologies in their ideal performance envelopes (only system optimization and modularity issues have to be solved). In the medium-term double-stage Hall-effect thrusters, if they prove capable to reach their theoretical performances, could be extensively employed. For the long term it would be useful to define a clear technology road-map capable of focusing common European
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
Fig. 14. 3D cut of LABEN-ALTA DSHET concept [19].
efforts on the most promising innovative variable-Isp concepts. References [1] L. Casalino, G. Colasurdo, Optimization of variable-specificimpulse interplanetary trajectories, AIAA-2002-4897. [2] E. Chesta, G. Janin, J. Gonzalez del Amo, G. Saccoccia, A full solar electric propulsion concept for Mars Exobiology, AIAA-2001-4319.
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[3] GREMI Laboratory, Evaluation of dual mode high power SPT’s, Final Report, ESA Contract 16218/02/NL/PA (November 2002). [4] ALTA, Challenges of very high-power Hall-thrusters, Final Presentation, ESA Contract 16223/02/NL/PA (December 2002). [5] P. Dumazert, F. Marchandise, M. Prioul, L. Jolivet, PPS1350-G Qualification status May 2003, AIAA-2003-4549. [6] QINETIQ, Bepi-Colombo TDA, Phase1 Test Report, ESA Contract 16333/02/NL/PA (August 2003). [7] ALTA, Development and supply of a FEEP system for Microscope, FEEP Performance and technology overview, Technical Note ALTA/MS/TN-05, ESA Contract 15.231/01/NL/PA (May 2003). [8] Institut für Allgemeine Elektrotechnik und Elektronik, FEEP spectroscopic investigations on slit emitters, Final Report, ESA Contract 5051/82/NL/PB (December 1985). [10] S. Oleson, Mission advantages of constant power, variable Isp electrostatic thrusters, AIAA-2000-3413. [11] SNECMA, Development of a DSHET, Concept and Technology Assessment Technical Note, WP2000, ESA Contract 16723/02/NL/CP (February 2003). [12] O. Secheresse, A. Bougrova, A. Morozov, Patent No. 03 08384, 9 July 2003. [13] A.I. Morozov, A.I. Bugrova, A.D. Desiatskov, V.K. Kharchevnikov, M. Prioul, L. Jolivet, Research on two-stage engine SPT-MAG, IEPC-2003-290. [14] CNES Contract 712/CNES/01/8590/00. [15] SNECMA, Development of a DSHET, Synthesis technical note of the concept choice, WP 3200, ESA Contract 16723/02/NL/CP (July 2003). [16] CPAT, Development of a DSHET, Hybrid simulation of DSHET, Research Report WP3000, ESA Contract 16723/02/NL/CP (May 2003). [17] SNECMA, Development of a DSHET, Possible effect of introducing an internal cathode in the ionisation stage of SPTMAG type DSHET, WP 3200, ESA Contract 16723/02/NL/CP (June 2003). [18] ALTA, Development of a DSHET, Concept and Technology Assessment Technical Note, WP2000, ESA Contract 16724/02/NL/CP (February 2003). [19] ALTA, Development of a DSHET, DSHET Preliminary design, WP4100, ESA Contract 16724/02/NL/CP (August 2003).
Flexible variable-specific-impulse electric propulsion systems for planetary missions E. Chestaa,∗ , D. Estubliera , B. Fallisa , E. Gengembrea , J. Gonzalez del Amoa , N. Kutufaa , D. Nicolinia , G. Saccocciaa , L. Casalinob , P. Dumazertc , M. Prioulc , J.P. Boeufd , A. Bouchoulee , N. Wallacef , G. Nocig , M. Bertih , M. Saverdih , L. Biagionih , S. Marcuccioh , A. Cadioui , F. Darnoni , L. Joliveti a European Space Agency, ESTEC, TOS-MPE, Keplerlaan 1, Noordwijk aan Zee, The Netherlands b Politecnico di Torino, Dipartimento di Energetica, Corso Duca degli Abruzzi 24, Torino, Italy c SNECMA Moteurs, Service Propulsion Plasmique, Villaroche Nord, Moissy-Cramayel, France d CNRS-CPAT, Université P. Sabatier, 118 Route de Narbonne, Toulouse, France e GREMI Laboratory, Université d’Orléans & CNRS, 45067 Orléans Cedex 2, France f QINETIQ, Space Department, Y72 Building, Farnborough, Hampshire, UK g LABEN, Divisione Proel Tecnologie, Via Albert Einstein 35, Campi Bisenzio, Firenze, Italy h ALTA S.p.A., Via A. Gherardesca 5, Ospedaletto, Pisa, Italy i CNES, Toulouse Space Centre, 18 Avenue Edouard Belin, 31401 Toulouse Cedex 9, France
Available online 3 October 2005
Abstract Variable-specific-impulse electric propulsion systems can provide important advantages and cost reductions to any scale of planetary mission. The first part of this paper describes their potential benefits, with special emphasis on mission analysis aspects. The second part identifies the physical limitations of traditional electric propulsion technologies and presents some preliminary development results achieved within the frame of ongoing European activities. © 2005 Published by Elsevier Ltd.
1. Introduction ∗ Corresponding author.
E-mail addresses: [email protected] (E. Chesta), [email protected] (L. Casalino), [email protected] (M. Prioul), [email protected] (J.P. Boeuf), [email protected] (A. Bouchoule), [email protected] (N. Wallace), [email protected] (G. Noci), [email protected] (L. Biagioni), [email protected] (L. Jolivet). 0094-5765/$ - see front matter © 2005 Published by Elsevier Ltd. doi:10.1016/j.actaastro.2005.07.053
Electric propulsion systems based on traditional technologies, hall effect thrusters (HETs) or gridded ion engines (GIEs), are well known to substantially enhance (and often to enable) scientific missions with demanding performance requirements. The benefits usually consist in increased payload capability (or decreased launch cost), enhanced mission flexibility
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(including launch windows enlargement and reduced trip time for long-distance targets), and extended S/C operating life. Even more consistent benefits can be achieved using homogeneous or modular variable-specificimpulse electric propulsion systems. In fact, in this case a single-staged full-electric spacecraft could effectively perform different mission phases, even though based on conflicting performance requirements: high-specific impulse (Isp ) during interplanetary cruise (heliocentric segment) and high thrust (T ) for Earth escape (geocentric segment from a negative C3 launch) and for final orbit acquisition around the target planet (planetary segment). The mission would therefore result considerably simplified at system level, and it would be possible to share most of the propulsion subsystem equipments with consistent cost reductions and enhanced reliability. Advanced simulation tools have also shown the possibility to achieve minimum propellant trajectories by adopting variable-specific-impulse strategies. In the heliocentric segment this can be achieved either independently from the available power (it is usually convenient to increase Isp as the S/C mass diminishes because of propellant consumption, or to adjust Isp and T to allow for optimal gravity assist encounters), or as a function of the available power when solar energy is used (typically, Isp grows together with the available power for inner solar system missions, while for missions to the outer solar system Isp is usually kept quasi-constant at its initial value while T decreases). In the geocentric or planetary segment, the general rule is that Isp is better decreased at the apsides of the orbit, when a higher thrust can be employed more efficiently. Additional benefits in terms of recurring costs reduction can be achieved with development programs based on a “modular approach”, at equipment level (aimed at employing similar subcomponents) or at subsystem level (whole propulsive block with shared equipment). The ideal situation would be of course to have an entirely homogeneous system, based on a single type of thruster technology capable of a wide range of performances. The higher Isp capabilities that are necessary to successfully perform the interplanetary part of the cruise are also likely to find convenient application for attitude control and station keeping of GEO telecom satellites and Earth observation missions, and the higher T necessary to reach
the escape orbit during the geocentric segment can be used in a similar way for orbit raising and quick transit through Van Allen belts. The European Space Agency and CNES have initiated several activities to quantify and overcome the performance limitations of conventional EP systems. In particular, the possibility to profitably employ readily available concurrent technologies on the same electric propulsion hybrid platform is currently under assessment for short-term applications. At the same time, two activities for the development of a double-stage hall-effect thruster (DSHET) prototype are running, and the concept selection phase has already been successfully achieved. If these new thrusters will prove effective, they will be the ideal candidates for both scientific and commercial missions on the medium term. For the long term, new concepts and improvements are currently under consideration for possible future research and development programs.
2. Mission benefits of variable-Isp electric propulsion systems In typical electric propulsion systems, the power of the electric source is proportional to the specific impulse and to the thrust level (if constant thruster efficiency is assumed). A high Isp reduces the propellant requirement, whereas a high T reduces the trip time. When a constant-Isp operation is considered, the mission designer selects the specific impulse and the thrust level to obtain the best compromise between mission payload and time length. Optimization methods are used to search for minimum-propellant trajectories. An increasing number of revolutions around the main body reduces the propellant consumption and a time constraint is necessary to make the optimization problem meaningful. In this case, the mission designer is able to obtain the optimal value of the specific impulse to accomplish the trajectory in the assigned time, once the available power is given; moreover, the optimal power level can be determined when the specific mass of the electric source is assigned. A thruster that can exploit a wide range of values for the specific impulse and the thrust could be used both for the escape from Earth and the interplanetary portion of the flight. However, a smaller range of operational Isp is also useful to improve performance, as
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
it allows a better management of the available power during the interplanetary trajectory. Most of the benefit is obtained by increasing Isp as the spacecraft mass diminishes during the flight; the optimal Isp is inversely proportional to the spacecraft mass in the absence of gravitational losses. However, the control law is modified by the presence of this kind of loss; Isp is reduced when the spacecraft is in positions where a large thrust level is advisable (for instance, at the apsides of the orbit), whereas it is increased when the thrust cannot be usefully exploited (for instance, in correspondence to the coast arcs of a constant-Isp trajectory). A variablespecific-impulse engine, without bounds on the attainable values of Isp , is always switched on, and the specific impulse continuously varied. A time constraint is also required in this case to avoid a meaningless solution with very high specific impulse and infinite trip time. Coast arcs again appear when the constraint is applied to the thrust-on time instead of constraining the overall mission time or when bounds on the specific impulse are imposed. An indirect optimization method is used to obtain the optimal trajectories with either constant or variable specific impulse, to evaluate the benefit of variable specific impulse during the heliocentric leg of interplanetary trajectories. A point-mass spacecraft with variable mass m is considered. The patched conic approximation is adopted and the time spent inside the planets’ sphere of influence is neglected; therefore, the equations of motion are integrated only in the heliocentric reference frame. The state equations are: dr/dt = v,
(1)
dv/dt = g + T/m,
(2)
dm/dt = −T /c,
(3)
where r and v are the spacecraft position and velocity vectors, g is the gravitational acceleration (an inversesquare gravity field is assumed throughout), T is the engine thrust and c =g0 Isp is the exhaust velocity. The thruster exploits solar electric propulsion. The available power Pa is related to the nominal power at 1 AU (P1 ) by means of a function, which depends on the mission (in general, Pa is proportional to the inverse square distance from the Sun). Variables are normalized using the radius of the Earth’s orbit, its corresponding circular velocity and the ini-
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tial mass of the spacecraft as reference values. Results are presented in non-dimensional form and can be applied to any level of S/C initial mass. The reference units for thrust and power correspond to 5.93 N and 176.6 kW per each 1000 kg of S/C initial mass. A power penalty coefficient = / = 10, which corresponds to = mg /P1 = 40 kg/kW for a thruster with constant = 0.7, has been assumed to compute the payload mass mu = mf − P1 / = mf − P1 . In order to allow for direct comparison of variable and constant Isp performances, simpler boundary conditions than in actual missions are assumed. The spacecraft always leaves the Earth’s sphere of influence on a parabola, i.e., with zero hyperbolic excess velocity. The procedure could however provide the optimal value of the hyperbolic excess velocity, once the characteristics of the engine used to escape from Earth were assigned. The arrival at the target planet with zero hyperbolic excess velocity is considered, as it appears to be the most general condition for the rendezvous. The position of the arrival planet is instead left free, which means that the spacecraft is inserted into the same orbit of the planet around the sun, but the target planet is in a different position. Thus, mission performance is not affected by the phase angle between the Earth and the target planet. This “time-free” trajectory can be flown every year departing on the same day, and the rendezvous mission can be easily obtained in those years when the target position is favorable. The payload mu is the performance index which is maximized; a constraint on the trip time is also imposed. Details on the optimization method can be found in Ref. [1]. The trajectory is controlled by the thrust magnitude and direction. The desired thrust level T = 2P /c is obtained by selecting the exhaust velocity and the engine input power taking operational constraints into account. The input power is always limited by the availability of solar power, while the exhaust velocity may be bounded or, in the limit case, constant. The boundary value problem, which arises from the application of the theory of optimal control, is solved by a procedure developed at the Politecnico di Torino and based on Newton’s method. Numerical results for missions to Venus, Mercury and Mars are presented in Figs. 1–3. Variable- and constant-Isp trajectories are compared in terms of payload and average exhaust velocity
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
mu
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0.012
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0.42
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0.021
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0.023
0.024
4
Fig. 1. (a and b) Mission to Venus: performances for different trip times (465 and 581 days, respectively).
cm = g0 Ism , where the average specific impulse Ism is computed as the ratio of the total impulse to the consumed propellant. The trip time has been fixed at = 8 (about 465 days) in Fig. 1a, = 12 (698 days) in Fig. 1b, and = 10 (581 days) in Figs. 2 and 3. Figs. 2b and 3b show the control and thrust history for missions with nominal power P1 = 0.022 and 0.013, the latter corresponding to 3.4 kW for a 1500kg S/C of the Mars Exobiology mission class [2]. It is well known that an optimal value of the power exists for each trip time. Observing Figs. 1a and b it can be seen that the performance improves for the same value of the available power as the mission time length grows, because the same total impulse can be provided with lower thrust and, conversely, with higher specific
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(b)
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Fig. 2. (a and b) Mission to Mercury: performance and control history for P1 = 0.022.
impulse and lower propellant consumption (the position of the maximum payload shifts to the left when grows, and the optimal power is lower in the case of variable specific impulse). Another very important feature is that the benefit of variable-Isp trajectories over constant-Isp trajectories is especially significant for shorter trip times. In this case the larger freedom coming from the additional control capability allows the designer to obtain a larger payload by reducing the losses which are related to the magnitude and direction of the thrust, whereas the average specific impulse is almost the same for constant and variableIsp thrusters. For instance, the specific impulse is reduced and higher thrust can be used where it is more beneficial for the reduction of gravitational losses. Fig. 2b shows that the larger power that is available in a
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945 1.4
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2
4
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Fig. 3. (a and b) Mission to Mars: performance and control history for P1 = 0.013.
mission to inner planets is mainly used to increase the specific impulse, whereas the thrust level is almost constant (similar plots are obtained for the Venus mission). The maximum payload is again obtained with a reduced power with respect to the constant-Isp case. Variable- and constant-Isp trajectories use the excess power that comes from the proximity of the Sun in a quite different manner: in the first case, the specific impulse can be increased, whereas the optimal thrust is almost constant; on the contrary, when c is constant, the excess power can only increase the thrust; coast arcs appear where the thrust would be inefficiently used. The optimization procedure can also deal with planetary flybys, which can be usefully exploited to reduce the V requirement. During the
route to Mercury, Venus can provide for instance a remarkable gravity assist. Low Isp is exploited in this case at departure in order to permit the large thrust necessary to encounter Venus in the most favorable position. Also in the case of a mission to Mars a higher payload mass is obtainable with a variable Isp strategy (Fig. 3a). In this case the spacecraft moves towards an external planet and, during the variableIsp operation, the reduction of the available power mainly affects the thrust level, whereas the exhaust velocity is kept almost constant (Fig. 3b). The average specific impulse is however larger for the variable-Isp thruster, which can compensate for the lower average thrust by properly throttling the engine when required. The specific impulse used in these calculations ranges over quite a large span, and is allowed to assume values at the limits of currently available technology (c = 2 corresponds to about Isp = 6000 s). It is possible however to use the program to compute and compare the benefits obtainable in case of unbounded Isp , bounded Isp , dual Isp and constant Isp . In Fig. 4 the results are obtained in the case of a direct mission to Mercury. It is quite interesting to notice how the
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variable-Isp thruster permits a 10% mean reduction of the propellant consumption. This value is slightly lower (about 9%) if the unavoidable upper and lower limits for the Isp are imposed at reasonable levels, but it can double in special cases (like for long missions with gravity assist and low total propellant mass). It has to be recognized that up to 80% of the achieved propellant mass savings could be obtained using dualmode thrusters, considerably simpler to develop and qualify.
3. Performance limits of traditional EP technologies Typical control parameters, for EP thrusters, are the inlet propellant mass flow rate and the applied voltage. For power-limited operations, these two parameters cannot be varied independently. In fact, neglecting losses due to incomplete propellant ionization or to beam divergence, the thrust power and the electrical power can be written as Pt = 21 mc ˙ 2, P=
(4)
Isp T g 0 Pt q = = IV = mV ˙ , 2 M
(5)
where I is the beam current, the thruster efficiency, and V the discharge voltage in the case of HETs, the beam voltage for GIEs and the emitter voltage for FEEPs. The exhaust velocity is in all three cases: c=
2
q V. M
(6)
Specific impulse and thrust-to-power ratio will then show reversed proportional behavior with respect to the square root of the applied voltage: c Isp = = g0 mc ˙ T = = P P
2q √ V, g02 M
(7)
2m 1 √ . q V
(8)
HETs, GIEs and FEEPs present optimal efficiency while operating around nominal points with relatively low, high and very high V , respectively. For this reason, HETs are usually considered for relatively high T applications, while GIEs and FEEPs for relatively high and very high Isp applications. It is a worthwhile exercise to identify what are the physical boundaries (mostly related to the ion beam extraction process) for thrust and specific impulse encountered when operating these three different technologies at variable power levels, and if possible what are the alternative control parameters that can permit to operate them within an extended range of performances. 3.1. Hall-effect thrusters In the case of Hall-effect thrusters, a recent activity performed by the GREMI Laboratory (Université d’Orléans) and supported by ESA in the frame of the Aurora programme has addressed the topic of “Dual mode high power SPT operation” [3]. This work has shown that the ability of single-stage HETs to operate in dual-mode increases with the power level. In fact, the minimum value of propellant flow rate in the thruster channel is the one that assures a neutral ionization time (roughly inversely proportional to the electron density and hence to the propellant flow rate) significantly shorter than the transit time of the neutral atoms in the ionization layer. This limit, especially at low power levels, translates directly in a constraining upper boundary for the discharge voltage (for the reachable specific impulse). When the thruster is designed to operate at high power levels this problem is mitigated, since the mass flow density for effective ionization is inversely proportional to the channel width. The maximal propellant flow rate value is on the contrary only limited by the maximal thermal fluxes transferable to the thruster structure for a given design (less restrictive constraint), or to the maximum magnetic field strength obtainable when up-scaling the power [4]. Experimental tests performed in collaboration with RIAME MAI (Moscow) on a 4–5 kW laboratory model SPT 140 have proven that a HET designed for discharge voltages lower than 500 V can operate during short periods in a high specific impulse mode with relatively good efficiency: the discharge voltage was increased up to 1000 V, experiencing
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
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0.65
2.31 mg/s 2.77 mg/s 3.23 mg/s 3.70 mg/s 4.16 mg/s 4.62 mg/s 5.08 mg/s 5.55 mg/s 6.01 mg/s 6.47 mg/s P = 2400 W P = 2000 W P = 1600 W P = 1200 W
Discharge efficiency
0.6
0.55
0.5
0.45
0.4 200
300
400
500
600
700
800
900
1000
Discharge voltage (V) Fig. 5. SNECMA PPS䉸 -1350-G: efficiency versus discharge voltage for variable flow rate/power [5].
only minor problems related to the non-optimized design of the thruster at high power levels (overheating of the coils with consequent magnetic excitation saturation, stochastic electrical breakdowns). Similar successful results have also been obtained by the Groupement de Recherche CNRS/CNES/SNECMA/ ONERA n.2232 “Propulsion à Plasma pour Systèmes Spatiaux” on a SNECMA PPS䉸 -1350-G engineering model [5], whose performances were explored in a wide range of operating points (Fig. 5), reaching Isp levels above 3000 s with good efficiency (as expected, at constant power levels the efficiency decreases when the discharge voltage is increased). Another interesting input from [4] is that both flexibility and performances increase when the thruster unit power (i.e. size) increases. A major concern is obviously related to the possible impact of similar dual-mode operations on thruster lifetime. The trends in terms of thruster power mentioned above suggest that this problem could also be reduced at higher thruster size and power. The assessment of this critical point will be hopefully the object of future dedicated activities.
3.2. Gridded ion engines In the case of gridded ion engines (GIEs), the most stringent limitation concerns the maximum value of propellant flow rate, restricted by space-charge effects related to the selected thruster optics geometry. In fact, Child’s law fixes a fundamental limit on the current I that can be drawn across a gap distance d between two plates of surface A by a given potential difference V : 40 2q V 3/2 A I =A = 2 V 3/2 , (9) 2 9 M d d where 0 is the permittivity of free space and is Child’s constant. This formula is applicable to GIEs ion extraction optics, but in this case the effective acceleration length d must take into account also the screen grid thickness and the variable plasma sheath bowed shape. For increasing values of plasma density (due to increased mass flow rate or more negative accelerator voltage), the plasma sheath pushes closer to the screen grid and becomes flatter, causing direct impingement of the under-focused beam on the
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230 mN
Measured Accel Grid Current (mA)
25
20
15
10 2
1850V, Y=1.627X + 1.371X + 0.318, IB limit = 3.4A 2
1700V, Y=1.695X + 1.304X + 0.265, IB limit = 3.1A 2
5
1500V, Y=2.755X - 1.522X + 1.797, IB limit = 2.75A
50 mN
2
1200V, Y=6.531X - 10.858X + 7.4, IB limit = 1.85A 0 0.5
1.0
1.5
2.0
2.5
3.0
3.5
Beam Current (Amps)
8000
7000
Toral Power (Watts)
6000
1850V Beam Voltage 1700V Beam Voltage 1500V Beam Voltage 1200V Beam Voltage
230 mN
200 mN
5000 175 mN 4000 150 mN 3000
125 mN 100 mN
2000
1000
0 1.0E-03
50 mN
1.5E-03
2.0E-03
2.5E-03
3.0E-03
3.5E-03
4.0E-03
4.5E-03
5.0E-03
Total Mass Flow Rate (g s-1)
Fig. 6. QinetiQ T6 experimental characterization and identification of maximum beam current [6].
accelerator grid. This is called sometimes the “perveance limit” (the perveance being the ratio I /V 3/2 ), and, for a given thruster design, it is the main limiting factor to the achievable maximal thrust and to the employment of large power-saving negative accelerator voltages (a part from considerations related to higher
energy charge-exchange ions impingement when no deceleration grid is present). An experimental characterization of this limit for a T6 thruster by QinetiQ, part of the “Bepi Colombo Technology Demonstration Activity” [6], is presented in Fig. 6. The maximum beam current for each beam
Specific Impulse [s]
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
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7000 6800 6600 6400 6200 6000 5800 5600 5400 5200 5000 4800 4600 4400 4200 4000
Va-2kV Va-1kV
7.0
7.5
8.0
8.5
9.0
9.5
10.0
Total Voltage [kV]
Fig. 7. FEEP specific impulse vs total voltage at different values of accelerator voltage (Courtesy: ALTA).
voltage is limited by the onset of direct impingement (accelerator grid current). The reciprocal phenomenon occurs for decreasing values of plasma densities (very high beam voltages and low mass flow rate): in this case the plasma sheath retreats from the screen grid and becomes more concave, such that the trajectories of some ions cause them to crossover and impinge on the accelerator grid. This “crossover limit” is less restrictive than the perveance limit (the impingement current increase with beam voltage is smoother), and can easily be overcome if very high specific impulse is desired by adopting thruster geometries with larger grid separation.
for electron quantum tunneling [7]:
3.3. Field-emission propulsion
where Vs is the extrapolated starting voltage and C3 a constant taking into account again the geometry of the problem, the temperature of the propellant and the average number of emission sites. In both cases, the emitted current and the total applied voltage are intrinsically correlated and cannot be changed independently. The emitter temperature is controllable, but usually it is maintained at the minimum value necessary to keep the propellant liquid. An interesting feature of this kind of thrusters is the possibility to employ very negative accelerator voltages (to operate at very low ratios of the net-to-total accelerating voltage). This is possible because the emission sites
FEEP thrusters can operate only when the electric field applied between the accelerator and the emitter electrode is sufficiently strong to counteract the surface tension of the liquid metal propellant and initiate ion field emission. For this reason, they are traditionally considered only for very high Isp and low T applications. For a given thruster geometry, the mass flow rate cannot be used as a control parameter. In fact, experimental studies have demonstrated that, in low current regimes, the emission law is exponential and closely related to the Fowler–Nordheim equation
I = C1 AV 2 e−C2
3/2
/V
.
(10)
is here the propellant work function; the constants C1 and C2 depend on the thruster geometry and on the Fermi energy measured from the bottom of the conduction band. In high current regimes, the emitter obeys a law determined by space-charge limits [8]: I = C3 [(V /Vs )3/2 − 1],
(11)
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are discretely localized, the liquid meniscus shape is only locally and microscopically affected by the applied electric field (no appreciable changes in the acceleration length are produced), and the chargeexchange ions production rate is much lower (the neutral vapor density is proportional to the emitter temperature, relatively low if propellant with low melting point like cesium are used). If the accelerator negative voltage (instead of the positive emitter voltage) is increased, the extracted current increases according to (11) with reduced energetic cost: P = I V e = C4 (Ve + |Va |)3/2 Ve .
(12)
Of course this involves consequently reduced Isp operation (Fig. 6). As for GIEs, the maximum achievable Isp is only power limited and can be increased for a given power level by employing higher emitter voltage and larger electrodes separation. The maximum achievable T is strictly related to the emitter geometry and the number of emission sites. Based on the specific characteristics of this technology, the possibility to develop high-current prototypes for future highpower applications deserves accurate consideration (Fig. 7).
4. Development of hybrid EP systems (short term) While HET and GIE technologies have complementary performances and very distinct technical characteristics at thruster level (as shown in previous paragraph), on the contrary they present important similarities at system level. Their possible synergies can be profitably employed to solve requirement conflicts on planetary scientific missions and telecommunication satellites (e.g. @Bus), and are currently under assessment in the frame of the fast-track GSP activity “hybrid (coupled plasma/ion) electric propulsion system feasibility study”. The first goal of this activity is to perform an exhaustive preliminary assessment at system level between HET and GIE technologies, aimed at identifying all possible commonalities and incompatibilities. The second goal is to execute a schematic design of two complete hybrid electric propulsion systems
(HEPSs), one for a GEO Telecom application and the other for an interplanetary scientific mission, using if possible the same hybrid electric propulsion module (HEPM) inclusive of one HET, one GIE and the relevant sub-system components. This approach is relatively innovative, since past planetary mission studies involving combinations of different EP technologies, mainly conducted by NASA [10], usually considered S/C composed of jettisonable independent homogeneous propulsive stages. This alternative solution is viable and provides some mission advantages related to a lighter S/C after separation of the HET stage, but it is globally less attractive than the hybrid platform because the higher thrust jettisoned engines cannot be used to optimize the interplanetary trajectory or final planetary orbit acquisition (an additional chemical propulsive stage would be necessary); in addition, the S/C would be more complex and failure risks and costs would be higher. The commonality assessment first iteration has provided some interesting preliminary results, showing that a modular approach at subsystem component level can generate important cost savings, and that a flexible HEPM composed of hybrid common elements is achievable (modular approach at propulsive system level). An example of interesting solution for a hybrid Power Processing Unit is presented in Fig. 8 (courtesy QinetiQ). Electric propulsion systems based on flexible hybrid modules operating well-known thruster technologies in their ideal performance envelopes are likely to prove very competitive on the satellite market in the near term. If the results of this study are positive, next step could be a consequent experimental demonstration requiring the development of specific hardware to breadboard level.
5. Development of double-stage Hall effect thrusters (medium term) The highest benefits, in terms of cost and mass savings per mission, can be achieved when the S/C is equipped with a homogeneous variable-specificimpulse propulsion system supporting a variable number of identical thrusters (instead of hybrid modules as considered earlier). As demonstrated earlier, EP
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
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1
Heater/ Magnet 2
Anode/ Keeper/ Magnet
3
4 5 6
Accel. Grid
7
FILTER UNIT
8
Heater/ Keeper
9
10
Beam/ Discharge
Fig. 8. Potential hybrid PPU concept (QinetiQ).
systems based on currently available technology are unable to operate efficiently for long time in highly variable-Isp mode. In the medium term, it is possible to envisage the development and qualification of new thruster prototypes directly or indirectly derived from the existing ones, thus not requiring complex technology break-throughs. An extremely promising concept in this field is the DSHET, designed to achieve at least a partial decoupling of what are called ionization and acceleration regions in traditional stationary plasma thrusters. The expected performance improvements, with respect to traditional HETs, concern mainly an enhanced ionization process in the first stage and a finer control of the acceleration phase, with consequent benefits on the efficiency and on the beam divergence, which can be dramatically reduced if the magnetic field is optimized and capable to focus the ions in the acceleration stage (with consequent increase of thruster life capabilities and decrease of near-wall interactions and discharge noise). The prac-
tical realization of this separation can be obtained in very different ways. The ionization process is closely related to the electron current in the ionization region; in order to reduce the power dissipated by the electrons while crossing the acceleration stage, two solutions are possible [11]: the plasma source can be designed to require the lowest possible electronic current, and/or part of the electrons can be injected directly into the ionization zone. Two parallel assessments recently performed by independent industrial consortia under ESA sponsorship have identified two very promising concepts. 5.1. DSHET based on Galathea confinement The first concept, developed by SNECMA [12] in collaboration with CPAT, QinetiQ and ALT A, is based on the Russian [13] prototype engine SPTMAG [14], and is schematically represented in Fig. 9. The ionization stage is conceived in order to
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E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945 Inductive ionization stage
‘ Classical ’ acceleration stage
Typical ion trajectory
anode Xe
Myxina
Conducting wall
Magnetic field lines
Fig. 9. Schematic of SNECMA SPT-MAG thruster [11].
maximize the electron residence time and to optimize ion injection into the acceleration stage. The magnetic field topology can be classified as a (half) Galathea magnetic trap, based on a single current-carrying conductor completely imbedded in the plasma (mixina). Thanks to this configuration, it is possible to create a magnetic well in front of the acceleration stage inlet. In addition, a higher voltage is applied to the mixina surface and to the ionization chamber walls with respect to the anode of the acceleration stage, thus creating a potential well where ions oscillate in their motion towards the zero-B region. The imaginary magnetic line passing through the zero-B point and intercepting the anode is also the minimum equipotential line, it is called “separatrix”, and its downstream portion can be considered as the border between ionization and acceleration stage. The thruster functional analysis can be performed using a quasi-neutral hybrid model where electrons are described as a fluid and ions and neutrals as macro-particles [16]. An example of simulated electric potential distribution and ion trajectories obtained with this model is presented in Fig. 10a and b. The design of a 3 kW prototype is currently in progress (Fig. 11). The possibility to include an emitting internal electrode (hollow cathode) in the ionization stage has been carefully assessed and finally rejected [17]. The prototype will be manufactured
Fig. 10. (a,b) SPT-MAG internal modelling [16].
Fig. 11. 3D external representation of SNECMA 3 kW SPT-MAG concept [15].
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
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Fig. 12. Schematic representation of LABEN-ALTA DSHET concept [18].
by SNECMA and fully tested in ALTA facilities next year. 5.2. DSHET based on double-peaked magnetic field and semi-active intermediate electrode The second concept, developed by LABEN/Proel in collaboration with ALTA, is schematically presented in Fig. 12. The separation between ionization and acceleration stage is achieved in this case through a doublepeaked magnetic field [18]. The specific topography of the magnetic field in the two stages will be independently controlled and experimentally optimized, but it is probable that a configuration with reversed radial polarity will be finally adopted in order to assure a complete decoupling of the two processes. The intermediate electrode is a semiactive thermionic device (axis-symmetric impregnated tungsten coil or multiple-dispenser), whose emission capabilities increase if its working function is lowered. In this way, part of the electronic current necessary to the ionization process is directly injected into the ionization stage, increasing current utilization and reducing power dissipation and electron losses to the wall. A possible magnetic topography has been simulated with Quickfield 4.3 and is reported in Fig. 13a, with axial and radial magnetic field components in the middle of the channel enlarged in Fig. 13b [19]. A 3D
representation of the thruster internal configuration is presented in Fig. 14. This prototype will be manufactured by LABEN/Proel and also tested in ALTA facilities next year.
6. Future research (long term) On the long term, the possibility to upscale at very high power levels the technologies presented in previous paragraphs will be considered, together with the analysis of alternative and more innovative solutions. Some original ideas have emerged during the concept and technology assessment phase of the activities presented. These concepts span from double-stage thrusters with high-frequency (capacitive or inductive or helycon) plasma sources, to multi-staged or segmented thrusters with channels presenting axially variable characteristics (dimensions, materials and electric potentials), passing from cusped magnetic field confinement in the ionization stage to magnetic nozzles for the acceleration stage, including also intermediate heating stages based on ion cyclotron or electron cyclotron resonance. The potential performances of these concepts are not easily predictable without dedicated experimental activities, which are relatively expensive and hardly achievable in parallel.
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Fig. 13. (a and b) LABEN-ALTA DSHET [19]: magnetic induction lines distribution and magnetic flux density.
7. Conclusions The utility and cost-effectiveness of variable-Isp electric propulsion systems for planetary missions have been discussed and demonstrated. The physical limitations of traditional EP technologies have been shown. Reliable solutions, in the short term, can be provided by hybrid modular systems operat-
ing well-known HET and GIE technologies in their ideal performance envelopes (only system optimization and modularity issues have to be solved). In the medium-term double-stage Hall-effect thrusters, if they prove capable to reach their theoretical performances, could be extensively employed. For the long term it would be useful to define a clear technology road-map capable of focusing common European
E. Chesta et al. / Acta Astronautica 59 (2006) 931 – 945
Fig. 14. 3D cut of LABEN-ALTA DSHET concept [19].
efforts on the most promising innovative variable-Isp concepts. References [1] L. Casalino, G. Colasurdo, Optimization of variable-specificimpulse interplanetary trajectories, AIAA-2002-4897. [2] E. Chesta, G. Janin, J. Gonzalez del Amo, G. Saccoccia, A full solar electric propulsion concept for Mars Exobiology, AIAA-2001-4319.
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[3] GREMI Laboratory, Evaluation of dual mode high power SPT’s, Final Report, ESA Contract 16218/02/NL/PA (November 2002). [4] ALTA, Challenges of very high-power Hall-thrusters, Final Presentation, ESA Contract 16223/02/NL/PA (December 2002). [5] P. Dumazert, F. Marchandise, M. Prioul, L. Jolivet, PPS1350-G Qualification status May 2003, AIAA-2003-4549. [6] QINETIQ, Bepi-Colombo TDA, Phase1 Test Report, ESA Contract 16333/02/NL/PA (August 2003). [7] ALTA, Development and supply of a FEEP system for Microscope, FEEP Performance and technology overview, Technical Note ALTA/MS/TN-05, ESA Contract 15.231/01/NL/PA (May 2003). [8] Institut für Allgemeine Elektrotechnik und Elektronik, FEEP spectroscopic investigations on slit emitters, Final Report, ESA Contract 5051/82/NL/PB (December 1985). [10] S. Oleson, Mission advantages of constant power, variable Isp electrostatic thrusters, AIAA-2000-3413. [11] SNECMA, Development of a DSHET, Concept and Technology Assessment Technical Note, WP2000, ESA Contract 16723/02/NL/CP (February 2003). [12] O. Secheresse, A. Bougrova, A. Morozov, Patent No. 03 08384, 9 July 2003. [13] A.I. Morozov, A.I. Bugrova, A.D. Desiatskov, V.K. Kharchevnikov, M. Prioul, L. Jolivet, Research on two-stage engine SPT-MAG, IEPC-2003-290. [14] CNES Contract 712/CNES/01/8590/00. [15] SNECMA, Development of a DSHET, Synthesis technical note of the concept choice, WP 3200, ESA Contract 16723/02/NL/CP (July 2003). [16] CPAT, Development of a DSHET, Hybrid simulation of DSHET, Research Report WP3000, ESA Contract 16723/02/NL/CP (May 2003). [17] SNECMA, Development of a DSHET, Possible effect of introducing an internal cathode in the ionisation stage of SPTMAG type DSHET, WP 3200, ESA Contract 16723/02/NL/CP (June 2003). [18] ALTA, Development of a DSHET, Concept and Technology Assessment Technical Note, WP2000, ESA Contract 16724/02/NL/CP (February 2003). [19] ALTA, Development of a DSHET, DSHET Preliminary design, WP4100, ESA Contract 16724/02/NL/CP (August 2003).