Patent Grant
8468794
Two general types of electric propulsion thrusters used in space include ion thrusters and Hall-effect thrusters.
Both ion thrusters and Hall-effect thrusters are ‘electrostatic’ electric propulsion devices used for spacecraft propulsion that create thrust by accelerating ions. The thrust created is very small compared to conventional chemical rockets, but a very high specific impulse, or high exhaust velocity, which reduces the propellant requirements for missions is obtained. This high ‘propellant efficiency’ is achieved through the very frugal propellant consumption of the electric propulsion system. They do however require large amounts of power; typically 1 kWe per 0.030-0.040 Newtons thrust for ion thrusters, and 1 kWe per 0.050-0.080 Newtons thrust for Hall-effect thrusters.
Ion thrusters and Hall-effect thrusters both generate a beam of ions (electrically charged atoms or molecules) to create thrust in accordance with Newton's third law. The method of accelerating the ions varies, but all designs take advantage of the charge/mass ratio of the ions. This ratio means that relatively small potential differences can create very high exhaust velocities. This reduces the amount of reaction mass or fuel required, but increases the amount of specific power required compared to chemical rockets. Electric propulsion thrusters are therefore able to achieve extremely high specific impulses.
The drawback of the low thrust is low spacecraft acceleration because the mass of current electric power units is directly correlated with the amount of power required. This low thrust makes electric propulsion unsuited for launching spacecraft into orbit, but they are ideal for in-space propulsion applications.
Gridded electrostatic ion thrusters commonly utilize xenon gas. This gas has no charge and is ionized by bombarding it with energetic electrons. These electrons can be provided from an electron source as a hot cathode filament, or more typically a hollow cathode assembly (HCA), which are then accelerated in the electrical field of the cathode fall to the anode (Kaufman type ion thruster).
The positively charged ions are extracted by an extraction system consisting of 2 or 3 multi-aperture grids. After entering the grid system via the plasma sheath the ions are accelerated due to the potential difference between the first and second grid (named screen and accelerator grid) to the final ion energy of typically 1-2 keV, thereby generating the thrust. Typical ion velocities are in the range of 20,000-50,000 m/s, and higher for some energetic mission applications.
In spacecraft propulsion, a Hall-effect thruster also accelerates ions by an electric field. Hall-effect thrusters trap electrons in a radial magnetic field and then use the electrons to ionize propellant, efficiently accelerate the ions to produce thrust, and neutralize the ions in the plume.
The essential working principle of the Hall-effect thruster is that it uses an electrostatic potential to accelerate ions up to high speeds but does so without the application of a gridded extraction system used in ion thrusters. In a Hall-effect thruster the attractive negative charge is provided by an electron plasma at the open end of the thruster instead of a grid. A radial magnetic field of a few tens of milli-Tesla is used to confine the electrons, where the combination of the magnetic field and an attraction to the anode upstream surface force a fast circulating electron current around the axis of the thruster and only a slow axial drift towards the anode occurs.
A propellant, such as xenon gas is fed through the anode, which has numerous small holes in it to act as a gas distributor. As the neutral xenon atoms diffuse into the channel of the thruster, they are ionized by collisions with high energy circulating electrons.
The xenon ions are then accelerated by the electric field between the anode and the cathode. The ions quickly reach speeds of around 15,000 m/s for a specific impulse of 1,500 seconds (15 kN·s/kg). Upon exiting however, the ions pull an equal number of electrons with them, creating a plume with no net charge. The axial magnetic field is designed to be strong enough to substantially deflect the low-mass electrons, but not the high-mass ions which have a much larger gyroradius and are hardly impeded. About 30% of the discharge current is an electron current which does not produce thrust, which limits the energetic efficiency of the thruster; the other 70% of the current is in the ions. The ionization efficiency of the thruster is thus around 90%.
The magnetic field thus ensures that the discharge power predominately goes into accelerating the xenon propellant and not the electrons, and the thruster turns out to be reasonably efficient.
Because of the counter-flowing electron and ion currents in the Hall-effect thruster channel, a greater ion flux can be achieved as compared to that of the ion thruster—thereby yielding higher thrust-to-power than ion thrusters. Ion thrusters however are capable of achieving higher exhaust velocities with higher overall thrust efficiencies.
It would be desirable in the propulsion field to provide an electric propulsion device that delivers the performance capabilities of high thrust-to-power devices (such as Hall-effect thrusters) and the high specific impulse of high total-impulse devices (such as electrostatic gridded-ion thrusters).
One limitation is that no single Electric Propulsion Thruster (EPT) exists which can operate over the full specific impulse range of interest and do so with the combined characteristics of both Hall-effect and ion thrusters, such as high Thrust-to-Power (T/P).
The design attributes of the Hall-effect thruster and the ion thrusters specific to their ion-acceleration systems: the Hall-effect thruster utilizes backstreaming electrons accelerated from an external cathode toward an anode upstream of a radial-geometry magnetic field within an azimuthally-symmetric channel to generate a plasma and create counter-flowing accelerated ion current; and the ion thruster utilizes a closely-spaced multi-aperture electrodes (electrostatic ‘ion optics’) with a large applied E-field to focus and accelerate ions from a discharge plasma to form a (space-charge-limited)˜mono-energetic beam.
The accompanying drawings, which are incorporated in and constitute a part of the specification, illustrate various example systems, methods, and so on that illustrate various example embodiments of aspects of the invention. It will be appreciated that the illustrated element boundaries (e.g., boxes, groups of boxes, or other shapes) in the figures represent one example of the boundaries. One of ordinary skill in the art will appreciate that one element may be designed as multiple elements or that multiple elements may be designed as one element. An element shown as an internal component of another element may be implemented as an external component and vice versa. Elements may not be drawn to scale and in some instances, cross-hatching is not shown to improve clarity.
As used herein, the relevant design attributes of a Hall-effect and ion thrusters are their respective ion-acceleration systems: an azimuthally-symmetric channel with axial E-field and radial B-field for the Hall-effect thruster yielding closed-drift electrons to generate a plasma and create counter-flowing accelerated ion current; and closely-spaced multi-aperture electrodes or electrostatic ion optics with a large applied E-field to focus and accelerate ions from a discharge plasma to form a space-charge-limited mono-energetic beam for the ion thruster.
The plasma production and acceleration mechanisms of the Hall thruster are closely-coupled and are intimately connected to the geometric construction of the thruster discharge geometry.
On the other hand ion thrusters have de-coupled plasma production and acceleration mechanisms. As such, the thruster discharge geometry can be constructed in a variety of fashions without compromising the operational integrity of the ion thruster—so long as the acceleration mechanism, the electrode geometry, is maintained. An ion thruster discharge chamber may take a number of geometries—cylindrical, an oblate-spheroid, rectangular-box, etc. So long as there is a high degree of azimuthal symmetry to the ion thruster discharge geometry, a magnetic circuit can be designed to contain the discharge plasma that will yield a high discharge electrical efficiency. This is particularly true when the thruster is operated at high plasma densities. Maintaining high discharge electrical efficiency is an important consideration when trying to improve thrust to power (T/P)-ratio and overall thruster efficiency.
In a first embodiment illustrated in exploded and partially cut away
It is appreciated that the ion optics 120 are shown “exploded.” from the exhaust annulus 130. The ion optics 120 may be configured as a set of parallel annular electrodes having an outer radius 180 substantially conforming to an outside edge of the ion thruster 100. The ion optics further are defined by an inner radius 182 selected to ensure the exhaust annulus is covered by the ion optics 120. A distance or span 190 is defined between the outer radius 180 and the inner radius 182. Moreover, ion optics 120 include closely-spaced apertures 122, usually circular, through the thickness which are aligned between the electrodes. In one example, on the outer most electrode, the apertures are 0.075″ diameter, with 0.093″ center-to-center spacing in a hexagonal array, so the electrode has a very high open area fraction. Using the example, across any 1-inch span there are ˜10 apertures.
With reference now to
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An anode 450 may be operatively disposed within and electrically insulated from the discharge chamber 410. In the illustrated embodiment, the anode 450 is electrically isolated and disposed along the interior surface 442, closed end 444 and outermost surface 446. Such an anode 450 provides additional surface area when compared to an anode disposed along only a single surface, for example the outer surface. As more completely described below, the added surface area improves electrical characteristics of the engine. Ion thruster 400 is further illustrated with a discharge cathode assembly 460 to generate a discharge plasma, the cathode assembly including a cathode 462 and propellant feed 464. Rare earth permanent magnets 470 are embedded within or formed integrally with the interior surface 442 and outermost surface 446 to establish a boundary magnetic circuit, of ring-cusp, or line-cusp geometry, to contain and control the energetic electrons emitted from the cathode assembly 460 to create an efficient plasma. Although only illustrated partially on a single side of the chamber 410, it is understood that the magnets 470 line the surfaces 442, 446. At least one ion plenum 480 is disposed toward the exhaust annulus 430 to provide a second, and primary, propellant feed.
One advantage of the illustrated ion thruster 400, as compared to ion thrusters of conventional configuration (cylindrical discharge with spherically-domed circular ion optics), includes the annular discharge chamber being able to provide for efficient packaging by providing a central position for mounting the neutralizer cathode assembly (NCA) 486 within the annulus. This reduces the outer profile of the engine and eliminates the need for a cantilevered-outboard NCA employed on conventional ion thrusters.
Another advantage of the illustrated ion thruster 400 includes the annular-geometry ion optics allowing for scaling of ion thrusters to very high power by permitting very-large beam areas with relatively small electrode spans, and relatively small span-to-gap ratios. This reduces the manufacturing, and the mechanical and thermal stability issues inherent with attempting to increase the beam area via increasing the diameter of spherically-domed ion optics as on conventional cylindrical ion thrusters.
Another advantage of the illustrated ion thruster 400 includes the annular-geometry ion optics allowing for the application of flat ion optics electrodes. Flat electrodes improve thrust to power (T/P)-ratios and efficiencies as compared to conventional ion thrusters by substantially reducing or eliminating the off-axis beam vectoring of ions which occurs with spherically-domed ion optics electrodes.
Another advantage of the illustrated ion thruster 400 includes the annular-shaped discharge chamber providing an opportunity to increase the effective anode-surface area for electron-collection as compared to a conventional cylindrically-shaped ion thrusters of equivalent beam area. This allows the illustrated engine to operate at the full-capability of the ion optics, and not have its maximum input power level limited by the available anode surface area. This increase in anode surface area allows the engine to operate at higher discharge currents and therefore higher beam currents and input power levels than a conventional ion thruster of equivalent beam area for a given specific impulse. Alternatively, for the same input power, the increased anode surface area associated with this annular geometry allows the engine to include a smaller outside diameter than a conventional ion thruster.
For example, the NEXT ion thruster (nominal 40 cm beam diameter) has a beam area of approximately 1257 cm2, and a discharge chamber anode area of approximately 3334 cm2. An engine according to the teachings here permit a smaller diameter (comparable to that of the 30 cm diameter NSTAR ion thruster) yielding a comparable discharge chamber anode area of approximately 3336 cm2, using an outside annular discharge chamber diameter of only 31 cm and an inside annular discharge chamber diameter of 6 cm (inside which a neutralizer cathode assembly may be contained). As detailed in Table 1, an annular discharge chamber engine may support an annular beam area of about 727 cm2 which is sufficient to support operation of the engine at the full-power operating power of the larger NEXT thruster—3.52 A beam current and 6.86 kWe. Such an engine would have a much reduced optics span, and optics span-to-gap ratio as compared to the NEXT ion thruster.
In operation an annular thruster channels propellant flow through the discharge cathode assembly (DCA) 460, the ion plenum 480, and the central common NCA 486. As discussed more completely below, an Ion Anode Power Supply, an Ion Beam Power Supply and an Ion Accelerator Electrode Power Supply are all energized. Operating range would be typically 2000-5000 seconds specific impulse.
With reference now to
Referring now to
The annular ion thruster 600 is further illustrated with a discharge cathode assembly 660 to generate a discharge plasma, the cathode assembly including a cathode 662 and propellant feed 664. Permanent magnets 670 are embedded within the outer wall 622 and interior wall 626. As above, the magnets are only illustrated partially, it is understood that the magnets 670 may be annularly disposed within the walls 622, 626, or discreet magnets may be spaced around the discharge chamber 620 and/or spaced between the exhaust annulus 640 and the end 624. Additionally, at least one ion plenum 680 is disposed near the exhaust annulus 640 side of the chamber 620 to provide a second propellant feed. In operation, the annular ion thruster 600 provides thrust from the exhaust annulus 630 as indicated by the arrow 690 although it is appreciated that thrust is provided from all or substantially all of the exhaust annulus 640.
With continued reference to
In one embodiment, a dual electric propulsion machine provides for efficient packaging to minimize overall engine diameter.
In one embodiment of a dual electric propulsion machine using a Hall-effect thruster component 610 as the second propulsion engine, maintains the design integrity of the Hall-effect thruster to ensure its performance characteristics of high T/P-ratio at low specific impulse;
A centrally-located neutralizer cathode assembly (NCA) 698 may provide dual functionality as both the discharge plasma generation and beam neutralization for the Hall-effect component 610, and beam neutralization for the ion thruster component 600. Additionally, a centrally located NCA 698 may reduce the outer profile of the engine and eliminate the need for a cantilevered-outboard NCA (not shown) dedicated for the operation of the ion thruster.
The annular-geometry ion optics 630 allow for very-large beam areas while creating very small electrode spans, and very small span-to-gap ratios. This permits use of flat ion optics electrodes. Flat electrodes yield improved T/P-ratios and efficiencies as compared to Current ion thrusters. This is because conventional ion thrusters are cylindrical in geometry, requiring spherically-domed ion optics electrodes to ensure both adequate stiffness for launch vibration, and thermo-mechanical stability under thermal loads during operation to maintain a uniform, controlled inter-electrode gap over very-large spans. The domed electrodes however result in thrust-losses associated with beamlets directed off-axis. Reducing thrust-losses associated with off-axis beam vectoring by using flat electrodes result in an improvement in overall efficiency for a given input power as more completely described below.
The annular-shaped discharge chamber 620 of the ion component 600 of the dual electric propulsion machine increases the effective anode-surface area for electron-collection as compared to a conventional cylindrically-shaped ion thruster of equivalent beam area. This allows the ion thruster component 600 to operate utilizing the full-perveance capability of the ion optics 630, and not have its maximum input power level limited by the available anode surface area. This increase in anode surface area allows the ion component to operate at higher discharge currents and therefore higher beam currents and input power levels than current ion thrusters of equivalent beam area for a given specific impulse. This increase in input power capability will be more completely described below.
The increase in operable surface area due to expanded surface available in the annulus also creates more radiative surface, permitting operation at higher discharge power levels as compared to a comparable beam area thruster of conventional construction. This is expected to maintain acceptable temperature margins on critical components such as the rare-Earth permanent magnets.
Geometric differences in ion optics electrodes 630, and discharge chamber anode surface areas as compared to conventional ion thrusters are documented in Tables 2 and 3 respectively. For purposes of illustration, the center core, or second thruster of the dual thruster engine is assumed to be a NASA GRC 300M Hall-Effect thruster in this embodiment. The 300M is a 20 kW-class laboratory electric propulsion thruster with an external diameter of 15⅜″ (39 cm).
Dual thruster engines with an ion thruster beam area equivalent to a conventional cylindrical ion thruster are listed in Table 2. As noted the dual thruster engine approach allows for a dramatic reduction in optics span, and span-to-gap ratio for a given beam area. Specifically, in one example, a 4-6× reduction in span and span-to-gap ratios. This comparatively small span is expected to result in a first-mode natural frequency high enough to allow for the use of flat electrodes. This in combination with the application of a long-life, low thermal coefficient-of-expansion material such as pyrolytic-graphite for the electrodes will provide an optimal flat-electrode design solution which is expected to eliminate the thrust-losses inherent with a domed electrode geometry used in conventional ion thrusters.
The differences in anode surface areas between conventional electric propulsion thrusters (Ion thrusters: NSTAR and NEXT are partial-conic; 50 cm is cylindrical) and the annular-portion (ion component) of the dual thruster engine having an equivalent beam area are documented in Table 3 (assuming a core of a 300M Hall-Effect thruster component). As noted, the dual thruster engine yields a much larger anode surface area for an equivalent beam area; e.g. −2.4× increase in area as compared to the NEXT ion thruster. It should be noted that these calculations are based on the increase in geometric surface area of the anode. For ring-cusp magnetic circuit plasma discharges the actual effective anode surface area is the sum of the magnetic cusp lineal areas. While estimated, the surface area ratios documented in Table 3 should be reasonably accurate.
The comparatively-large surface area of the dual thruster engine anode compared to either beam area alone or anode area of conventional EPT's of the same or similar beam areas allows for operation at higher discharge currents. This enables operation at much higher power levels, thereby taking full-advantage of the current-extraction capability (perveance) of the ion optics.
Several other design features and attributes of embodiments of the dual thruster-engine are noted here, with continued reference to
The Hall-effect component 610 of the dual thruster engine may be of conventional construction, using a solenoid electromagnet (not shown) to create the appropriately-shaped radial magnetic field at the exit plane of the channel. Use of a centrally-located cathode 698, and implementation of the extensible-channel concept (to enhance life time) employed on the NASA GRC HiVHAC low power Hall-effect thruster may be used in the dual thruster.
A ring-cusp magnetic circuit is shown within the annular ion discharge chamber 620, created by rows of alternating-polarity rare-Earth permanent magnets 670 attached or embedded within the surfaces of the annular discharge chamber surfaces 622, 626. Magnetic shielding 676 may be added between the exterior of the Hall-effect discharge 678 and the ion exhaust annulus 640 to separate the magnetic circuits of the two discharges. Alternately, it may be possible to use the fringe-magnetic field created by the Hall-effect discharge solenoid magnet—which would naturally penetrate into the annular ion discharge chamber 620 without shielding—and shape it by appropriate application of magnetic materials within the walls of the annular ion discharge chamber 620 to generate and control the discharge plasma in this zone. This would be with the potential benefit of reduced mass by elimination of the rare-Earth permanent magnets 670.
The annular discharge chamber 620 of the ion thruster component 600 may include a conventional discharge cathode assembly (DCA) 662 to generate the discharge plasma.
Typical ion thrusters require 3 separate propellant feeds; a Hall-effect thruster 2 propellant feeds. In one embodiment a dual thruster uses a total of 4 separate propellant feeds: one each for the ion discharge 680 and Hall discharge 694; one for the ion DCA 664; and one for the central common NCA 698.
The annular discharge chamber 620 of the ion thruster component 600 may use a ‘reverse-feed’ plenum which may inject the propellant from the ion optics-end of the discharge backwards, giving it an initial axial velocity component resulting in an increased neutral atom residence time and improved propellant efficiency.
The dimensions of the channel width and depth of the Hall-effect component 610 would be defined by the intended operating power level(s).
The annular area of the ion optics 630 would be established by the intended operating power level(s), which would then establish the electrode span (and discharge channel width) and overall dual thruster outside diameter.
The depth of the annular ion component discharge chamber 620 may include a tradeoff of electrical and propellant efficiencies, maximum desired input power, and overall dual thruster mass. It may be, for example, most effective mechanically and magnetically to match the channel depths of the ion thruster component 600 and the Hall-effect component 610.
With reference now to
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One additional advantage to this approach is to eliminate the asymmetry in the ion thruster component 830 created by a singular discharge cathode assembly within an annular-shaped discharge chamber 836. This asymmetry could cause a (minor) asymmetry in the plasma at the plane of the ion optics 840, and also result in a (minor) asymmetry in the magnetic circuit (‘a hole’) in the area of the DCA resulting in an increase in discharge electrical losses.
In the same manner that the discharge plasma is generated for the Hall-effect component, namely electron-back-streaming of current from the NCA 810 to the Hall-effect anode 846, it is expected that a discharge plasma for the ion thruster component 830 may be generated similarly. In the embodiment illustrated in
By appropriate shaping of the magnetic field in the passage (e.g. —axial B-field component across the radial slots) and de-energizing the Hall-effect solenoid (not shown) creating the downstream radial B-field component in the Hall channel, it should be possible to back-stream electrons 860 from the NCA 810, through the radial slots 850 toward the ion discharge anode 870 and in the process deplete their energy and use them efficiently to generate the ion thruster discharge plasma. The location of electron current collection and hence which discharge is operated (Hall-effect or ion) could be controlled by the switches identified in
With reference now to
With reference now to
The illustrated embodiment includes an annular ion anode 1070 of potentially greater surface area permitting higher currents and desirable features as described above. In the embodiment shown, annular ion anode 1070 lines but is insulated from an exterior surface 1082, a bottom 1084 of the cylindrical section, and an interior surface 1086.
As described in the embodiments above, a dual thruster is anticipated to have certain operational and performance characteristics. For example, a dual thruster can be operated in multiple ‘modes.’ Referring now to
As another example, the dual thruster can be operated in an ‘ion mode’ 1120 where the propellant flow is channeled through the ion plenum, the DCA, and the central common NCA. The Ion Anode Power Supply, the Ion Beam/Hall-Effect Anode Dual Power Supply, and the Ion Accelerator Electrode Power Supply are all energized. S1 switch is closed, and S2 switch is open. The ion component of the dual thruster is then operated as a conventional ion thruster. Operating range would be typically 2000-4000 seconds specific impulse as seen at 1120.
In yet another example, the dual thruster can be operated in a ‘burst mode’ 1130 where the propellant flow is channeled through all 4 locations, all power supplies are energized, and both S1 and S2 switch are closed. The Hall-effect and ion components of the dual thruster are then both operated simultaneously. Operation would be typically in the 1800-2200 seconds specific impulse range, in a zone where both components are capable of functioning with some overlap in capability 1130. This mode is theoretically possible but may not be operationally advantageous. However, two potential reasons for operating in this burst mode include: (a) Providing a seamless-transition in specific impulse throttling between Hall-effect mode 1110 and ion mode 1120 operation; and (b) assuming there is sufficient capability in the power electronics and propellant management system, operating at a total input power and generating a total thrust level for the dual thruster exceeding that which could be achieved by either the Hall-effect or ion components alone.
As mentioned above in discussing the annular discharge chamber ion thruster, when the dual thruster is operated in ion mode 1120 higher efficiencies at fixed input power, and higher input power at fixed specific impulse are possible, as compared to conventional ion thrusters. For example, the flat-geometry of the ion optics electrodes afforded by the annular design will improve the efficiency of the thruster as compared to a conventional ion thruster of equivalent beam area. This is because the thrust-losses associated with beam divergence due to the domed shape of conventional ion thrusters are eliminated in the annular design.
Moreover, both the specific impulse and the thrust are proportional to the thrust-loss correction factor due to off-axis beam vectoring (Ft); hence the overall thruster efficiency, which includes the specific impulse and thrust terms, is proportional to Ft2. The correction Ft includes both the beam divergence due to the electrode dome shape Ft-d and beam divergence due to beamlet expansion Ft-b.
From equation 9 of Soulas, G. C., “Design and Performance of 40 cm Ion Optics,” the NEXT ion thruster optics (dome height of approximately 2.35 cm and chord of 18 cm) Ft-d is estimated to be approximately 0.983. Hence, the reduction in thrust and specific impulse, and overall thruster efficiency of the NEXT ion thruster due to off-axis beam vectoring caused by the domed ion optics is expected to be 0.017, or −1.7% {100*(Ft-b−1.00)} in thrust and specific impulse, and 0.034, or −3.4% {100*(Ft-b2−1.00)} in efficiency at all throttle conditions. Therefore, all else being equal, for an equivalent beam area dual thruster with flat ion optics (Ft-d equal to 1.00), a 3.4% increase in efficiency as compared to the NEXT ion thruster would be expected across the entire specific impulse range as seen below:
One of the design issues with the NEXT thruster, and other conventional partial-conic or cylindrical discharge chamber ion thrusters (such as the NSTAR ion thruster, or the NASA GRC 50 cm laboratory model ion thruster), is that they cannot take full advantage of the ion current extraction capability of the ion optics technology—and hence operate at-or-near their maximum theoretical input power capability.
The NEXT neutralizer cathode, discharge cathode, magnets, ion optics, and high voltage propellant isolators all have adequate thermal and/or operational margins for operation at extremely-high input power levels; well in excess of the 7 kW maximum input power of the NEXT throttle table. Additionally, conventional NEXT ion optics are capable of operating at beam currents well-in-excess-of the maximum 3.52 A of the NEXT throttle table; >>7.0 A beam currents at full total voltage (2010 V). Application of advanced high-perveance design ion optics to the NEXT thruster would allow for operation at beam currents >>7.0 A at low total and beam voltages, thus enabling truly-high Thrust-to-Power operation.
Unfortunately, although operation at these high beam currents is consistent with the ion extraction system electrostatic functionality, they require a maximum sustainable discharge current which exceeds that which can be supported by the available anode surface area of the NEXT thruster discharge chamber. The maximum sustainable discharge current for the NEXT thruster is estimated to be in the range of about 32-35 Amperes, yielding a maximum beam current of about 7.0 A.
As documented in Table 3, for a given beam area, implementation of the annular discharge chamber either in a dual thruster arrangement or as a stand-alone annular ion thruster increases the effective and available anode surface area, as compared to conventional ion thrusters. The larger anode surface area increases the permissible sustainable discharge current and therefore beam current and thruster input power for a given beam voltage (specific impulse). The increase in sustainable discharge current (and approximate equivalent increase in beam current and input power) will be either directly proportional to the increase in anode surface area—or—will be equal to the discharge current necessary to support the maximum perveance-limited beam current that the ion optics are capable of extracting, whichever is less.
For an ion thruster with beam area equivalent to that of the NEXT ion thruster, an increase in anode surface area of 2.4× is possible. This permits an increase in sustainable discharge current and increase in beam current and thruster input power for a given beam voltage (specific impulse).
The increase in input power for a dual thruster versus a NEXT ion thruster of equivalent beam area is documented in Table 5 for a range of specific impulse. As indicated in Table 5 the input power to the NEXT ion thruster is limited by the anode-area down to about 3340 seconds specific impulse. Below this level, both the anode-area and the ion optics current-extraction-capability limit the thruster input power.
The annular ion and dual thruster is also anode-area limited at high specific impulse (4430 seconds); the ion optics are capable of supporting >20 A beam current, which would require a discharge current of greater than 84 Amperes. Although at this specific impulse the engine is anode-area limited, the maximum input power and thrust for this engine are 2.4× higher than that feasible with the equivalent-beam area NEXT thruster, due to the larger anode area of the dual thruster. At 4140 seconds specific impulse and below the dual thruster maximum input power is limited by the ion optics current-extraction-capability. That is, the engine is much more capable of taking full-advantage of the ion optics electrostatics than the NEXT thruster. At about 3340 seconds specific impulse and below, the maximum input power and thrust are equivalent to the NEXT thruster.
The data in Table 5 for the dual thruster are conservative estimates. This is because: a) the 1.7% increases in both specific impulse and thrust expected with the dual thruster due to the flat-geometry electrodes were not included; and b) the design of the dual thruster assumed the same ion optics electrostatic design as that of the NEXT thruster. The dual thruster could in fact incorporate an advanced-perveance ion optics design in an annular configuration—while maintaining the same beam area as the NEXT thruster optics. This dual thruster design would yield an increase in input power and thrust as compared to that documented in Table 5. It should be noted that there would be no advantage to applying advanced-perveance ion optics configuration to the NEXT thruster since its maximum beam current is already limited or inhibited by the anode-area.
With reference now to
With reference now to
It can now be appreciated that many variations of the dual thruster described here are possible. For example: using a second, or interior thruster other than a Hall-effect component, combining various sizes of Hall-effect components and ion components, to yield difference total specific impulse throttling ranges, and different total ranges in input power and thrust capability—depending upon the application need. One case example is provided here to define overall performance of a dual thruster. This example is for illustration, with the performance characteristics quoted specific to the selected configuration. As such the performance numbers documented do not imply any particular limits in capability to the dual thruster concept in general.
Table 6 lists the projected performance capabilities of one embodiment of a dual thruster, with a Hall-effect component having physical characteristics similar to that of the NASA GRC 300M Hall thruster, and an Ion component with beam area equivalent to the NASA NEXT Ion thruster. The performance of the Hall-effect component with specific impulse was modeled assuming a similar efficiency-specific impulse characteristic as the BPT-4000 Hall-Effect thruster, with a nominal input power of 20 kW at 2000 seconds specific impulse which is the design basis for the 300M.
The ion component of the dual thruster was modeled in a fashion consistent with prior ion thruster performance modeling conducted by this author assuming a discharge chamber, propellant efficiency of 0.92. The performance gains due to the elimination of the thrust-losses associated with domed ion optics were included in these calculations. Additionally, an advanced-perveance ion optics electrode design was assumed using equation 15 of Patterson, M. J., “NEXT Study of Thruster Extended-Performance II (NEXT STEP II).
Note in Table 6, ‘Hall-Effect Mode’ refers to the Hall-effect component operating solely (from 1220-1840 seconds specific impulse), ‘Ion Mode’ refers to the ion component operating solely (from 2150-4500 seconds specific impulse), and ‘Burst Mode’ refers to both components operating simultaneously. In this example case, burst mode is operated in the specific impulse ‘overlap-zone’ of 1840-2150 seconds specific impulse. The individual input power and thrust contributions of the two components in burst mode are documented in Table 6, with the first number associated with the Hall-effect component and the second number associated with the ion component.
As noted in Table 6, this dual thruster provides a continuous-throttling capability from 1220-4500 seconds specific impulse (3.7:1 range), with an input power range of 4.44-32.5 kW (7.3:1), thrust of 351-1146 mN (3.3:1), and efficiency of 0.47-0.78. A peak in input power and thrust occur over the specific impulse overlap-zone of the burst mode, with slight reduction in efficiency in this region.
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As seen in
The dual thruster could be applied to any application for which the high thrust-to-power characteristics of a Hall-effect thruster and high specific impulse capability of an ion thruster would be advantageous. For example: those Earth-orbital applications requiring both primary and auxiliary electric propulsion functions; Earth-orbital applications requiring both rapid orbit changes and fuel-efficient less-time-critical orbital changes; and Planetary mission applications requiring both fuel-efficient high-delta-V transfers and on-orbit high thrust-to-power operations.
In one embodiment of a dual thruster, the Hall-effect or other suitable second thruster component and the ion component may have approximately the same input power capabilities at their nominal design specific impulses (about 2000 seconds for Hall-effect and 3600 seconds for ion). This geometry is referred to here as a ‘Matched-Dual-Mode.’ This configuration is that provided in the example case earlier, and would be appropriate if both components were required to perform a primary-propulsive application.
In an alternative embodiment, referred to here as a ‘Mixed-Dual-Mode,’ involves combining a Hall-effect or other suitable second thruster component and an ion component with markedly-different input power capabilities at their respective nominal design specific impulses. This may be most-appropriate if one component were required to provide primary propulsion and the other component provide auxiliary propulsion; for example, combining a high-power Hall-effect component core (such as the NASA GRC 300M) for rapid (Earth) orbital transfers with a low-power ion component (for example, beam area equivalent to the NASA NSTAR thruster) for station-keeping. These configuration options are captured in Table 7.
While the systems, methods, and so on have been illustrated by describing examples, and while the examples have been described in considerable detail, it is not the intention of the applicants to restrict or in any way limit the scope of the appended claims to such detail. It is, of course, not possible to describe every conceivable combination of components or methodologies for purposes of describing the systems, methods, and so on provided herein. Additional advantages and modifications will readily appear to those skilled in the art. Therefore, the invention, in its broader aspects, is not limited to the specific details, the representative apparatus, and illustrative examples shown and described. Accordingly, departures may be made from such details without departing from the spirit or scope of the applicants' general inventive concept. Thus, this application is intended to embrace alterations, modifications, and variations that fall within the scope of the appended claims. Furthermore, the preceding description is not meant to limit the scope of the invention. Rather, the scope of the invention is to be determined by the appended claims and their equivalents.
As used herein, “connection” or “connected” means both directly, that is, without other intervening elements or components, and indirectly, that is, with another component or components arranged between the items identified or described as being connected. To the extent that the term “includes” or “including” is employed in the detailed description or the claims, it is intended to be inclusive in a manner similar to the term “comprising” as that term is interpreted when employed as a transitional word in a claim. Furthermore, to the extent that the term “or” is employed in the claims (e.g., A or B) it is intended to mean “A or B or both”. When the applicants intend to indicate “only A or B but not both” then the term “only A or B but not both” will be employed. Similarly, when the applicants intend to indicate “one and only one” of A, B, or C, the applicants will employ the phrase “one and only one”. Thus, use of the term “or” herein is the inclusive, and not the exclusive use. See, Bryan A. Gamer, A Dictionary of Modern Legal Usage 624 (2d. Ed. 1995).
This application claims priority to U.S. Provisional Application Ser. No. 61/295,326 filed Jan. 15, 2010.
The invention described herein was made by employees of the United States Government and may be manufactured and used by or for the Government for Government purposes without the payment of any royalties thereon or therefore.
| Number | Name | Date | Kind |
|---|---|---|---|
| 3156090 | Kaufman | Nov 1964 | A |
| 3613370 | Paine | Oct 1971 | A |
| 3956666 | Reader et al. | May 1976 | A |
| 3969646 | Reader et al. | Jul 1976 | A |
| 4011719 | Banks | Mar 1977 | A |
| 4825646 | Challoner et al. | May 1989 | A |
| 5448883 | Meserole et al. | Sep 1995 | A |
| 5924277 | Beattie et al. | Jul 1999 | A |
| 7180243 | Secheresse et al. | Feb 2007 | B2 |
| 7560870 | Tamida et al. | Jul 2009 | B2 |
| 20060010851 | Cagan et al. | Jan 2006 | A1 |
| Number | Date | Country | |
|---|---|---|---|
| 61295326 | Jan 2010 | US |
