ESA's Dual Stage 4 Grid Ion propulsion

[Astronuc]

Any info yet on why the significant delays on DS4G since passing of the leader?

That's a good question, and I will have to find out. I suspect, other priorities, which happens in this area - a lot. I'm guessing it has to do with limited budgets and the lack of a mission at the moment. This happened in the 1970s, in 1988 (after the loss of Space Shuttle Challenger in 1986), in 1990s, 2005-2007 (following loss of Space Shuttle Columbia). In between we have various financial/political crises, e.g., the near economic collapse in 2008-2009, Brexit, and now the war in Ukraine and economic malaise. As for exotic missions, such as sending manned spacecraft to Mars, and elsewhere, it's been somewhat sinusoidal (oscillatory and volatile) in terms of funding and resources. I expected as much 35 years ago.

I will contact some of the principals and find out where the technology and interest stands. Part of the problem is the lack of a demonstrable nuclear power plant, and part is probably sufficient chemical alternatives such as those used on Galileo and Cassini, and the fact that folks are willing to accept long missions - 6+ years long - using gravity assists from Venus, Earth, and in the case of Cassini, Jupiter enroute to Saturn. I remember when JIMO spun up, then was abruptly cancelled.

Cassini was launched October 15, 1997 and achieved orbital insertion around Saturn July 1, 2004. It was sent into Saturn around September 15, 2017. It ran out of fuel. Launch mass was 5,712 kg (12,593 lb) and the dry mass 2,523 kg (5,562 lb), so about 7000 kg of fuel. A manned mission would require an order of magnitude greater mass, and even then it might take months one-way, and then return, which has yet to be demonstrated. A fast one-way trip is complicated, and a 'fast' round-trip even more complicated.

Ref: https://en.wikipedia.org/wiki/Cassini–Huygens

There needs to be some kind of infrastructure in place before committing people to a deep space mission - even to Mars - or even to the moon.

 

[darkdave3000]

That's a good question, and I will have to find out. I suspect, other priorities, which happens in this area - a lot. I'm guessing it has to do with limited budgets and the lack of a mission at the moment. This happened in the 1970s, in 1988 (after the loss of Space Shuttle Challenger in 1986), in 1990s, 2005-2007 (following loss of Space Shuttle Columbia). In between we have various financial/political crises, e.g., the near economic collapse in 2008-2009, Brexit, and now the war in Ukraine and economic malaise. As for exotic missions, such as sending manned spacecraft to Mars, and elsewhere, it's been somewhat sinusoidal (oscillatory and volatile) in terms of funding and resources. I expected as much 35 years ago.

I will contact some of the principals and find out where the technology and interest stands. Part of the problem is the lack of a demonstrable nuclear power plant, and part is probably sufficient chemical alternatives such as those used on Galileo and Cassini, and the fact that folks are willing to accept long missions - 6+ years long - using gravity assists from Venus, Earth, and in the case of Cassini, Jupiter enroute to Saturn. I remember when JIMO spun up, then was abruptly cancelled.

Cassini was launched October 15, 1997 and achieved orbital insertion around Saturn July 1, 2004. It was sent into Saturn around September 15, 2017. It ran out of fuel. Launch mass was 5,712 kg (12,593 lb) and the dry mass 2,523 kg (5,562 lb), so about 7000 kg of fuel. A manned mission would require an order of magnitude greater mass, and even then it might take months one-way, and then return, which has yet to be demonstrated. A fast one-way trip is complicated, and a 'fast' round-trip even more complicated.

Ref: https://en.wikipedia.org/wiki/Cassini–Huygens

There needs to be some kind of infrastructure in place before committing people to a deep space mission - even to Mars - or even to the moon.

I just found this article:
https://futurism.com/nasas-new-ion-thruster-breaks-records-could-take-humans-to-mars

In comparison to the ESA/Australian Dual Stage 4 grid is this NASA hall thruster higher or lower in specific impulse? What about thrust?

Is it a similar design to the ESA model?

Is this Hall Thruster and the ESA thruster currently the leading most powerful and efficient ion drives invented so far?

Or is your Dawn Engine still the leading contender?

Was there any difference between Deep Space 1 and Dawn's ion drives?

Why has there not been a design or prototype for a hydrogen version of the ion drive? Given that Xenon is hard to find in space!

Lastly what is the weight of Ion drives in general ? Are they relatively light weight?

 

[darkdave3000]

I'd go with an NTR or NEP (nuclear electric propulsion). I would not do a hybrid NTR/NEP; that would introduce more complication in an already complicated design. Depending on the mission, one might to a two stage, an NTR stage and an NEP stage.

I don't see a simultaneous dual mode being useful. Pumping hydrogen into the core 'softens' the neutron spectrum, which is a disadvantage for a fast spectrum reactor (and increased transmutation of various structural materials and parasitic neutron absorption). A challenge with the NTR is how much enthalpy rise in the core, which depends on mass flow rate and ΔT, and how one preheats the cryogenic hydrogen to gaseous hydrogen.

I wouldn't use Xe, but it's favored because of it's low ionization energy compared to the other noble gases. However, using Xe means a lower Isp for a given energy/enthalpy. Hydrogen has a high ionization energy (~13.6 eV), and recombination losses are an issue for electrostatic or electromagnetic systems. For high Isp, one wants to use as low a molecular/atomic mass of the propellant as possible, which is why the space shuttle main engines run rich on hydrogen (lean on oxygen).

Since there are no firm designs on the horizon, since there are not firm missions to Mars or outer planets, there is an opportunity for further innovations in orbital transfer infrastructure.

Would you agree that a NTR would have more waste heat that could be captured and recycles for the production of electricity to power ion drives too? That it makes more sense to build a hybrid around an NTR instead of building a hybrid around a Nuclear Electric Reactor?

A Nuclear Electric reactor would not be able to achieve a high enough temperature to heat up hydrogen to give us the Isp advantage over chemical rockets I think! BUT PLEASE CONFIRM. As I understand it's designed to turn mechanical dynomos not run as hot as possible! So a warm reactor couldndt possibly heat up h2 to levels required to match or exceed Isp of chemical engines I think!

So what do you think? If we were to have an NTR system might as well throw in the ability to generate electricity from it while it's powered? Otherwise so much heat goes to waste! The NTR obviously has the heating capacity to do both heating h2 and generating electricity. But the Nuclear Electric reactor is only warm enough to generate electricity efficiently but not heat up h2 to do any reasonable thrusting with high isp numbers.

 

[Astronuc]

Would you agree that a NTR would have more waste heat that could be captured and recycles for the production of electricity to power ion drives too?

No, the thermal energy from the core/fuel is transferred to the propellant. One might bleed some hydrogen to cool the exterior structure and control systems of the reactor, or simply use that region to preheat hydrogen before it enters the core.

That it makes more sense to build a hybrid around an NTR instead of building a hybrid around a Nuclear Electric Reactor?

No, one would use NTR or NEP, but not both. NTR has relatively low I_sp, so uses high mass flow rate to achieve high thrust. The propellant flow rate must be such to cool the reactor fuel.

A Nuclear Electric reactor would not be able to achieve a high enough temperature to heat up hydrogen to give us the Isp advantage over chemical rockets I think!

NEP systems use EM propulsive devices to achieve high I_sp to use low mass flow rate, so a transfer vehicle carries less propellant. A nuclear reactor provides thermal energy, which is transferred to some thermodynamic cycles, e.g., Rankine (liquid to vapor), which is used to turn a turbine driving a generator. The waste heat from the condensed vapor is rejected to the environment, which is a challenge in space (vacuum) - so one must use a radiator, which could be massive.

The generator would need cooling, which could come from the propellant (LH₂). The reactor coolant, thermodynamic system and propellant feed system are complicated.

NTRs have been tested during the ROVER and NERVA programs.
https://en.wikipedia.org/wiki/Project_Rover
https://en.wikipedia.org/wiki/NERVA

 

[darkdave3000]

Could the waste heat from the nuclear electric be used as a photon rocket? If the radiators are shaped ? 200 megawatts of infrared radiation with very high specific impulse.

 

[Astronuc]

I just found this article:
https://futurism.com/nasas-new-ion-thruster-breaks-records-could-take-humans-to-mars

In comparison to the ESA/Australian Dual Stage 4 grid is this NASA hall thruster higher or lower in specific impulse? What about thrust?

The Plasmadynamics & Electric Propulsion Laboratory has webpage that answers most of one's questions.
https://pepl.engin.umich.edu/project/x3-nested-channel-hall-thruster/

“Nested-channel Hall thrusters have been identfied as a means to increase Hall thruster power levels above 100 kW while maintaining acceptable device size and mass. In a recent Broad Agency Announcement, NASA identified high-power electric propulsion (up to 300 kW) as enabling for a variety of mission structures, including human space exploration. Additionally, a 2010 NASA team found that high-power electric propulsion was key to allowing affordable travel to asteroids and near-Earth destinations by reducing launch mass up to 50%. NASA hopes to implement a system that has a broad power and specific impulse range for maximum flexibility within a mission. The multiple discharge channels of a nested-channel Hall thruster allows for throttling far beyond that of a single-channel Hall thruster. . . .

“The X3 is designed to operate efficiently on both krypton and xenon propellants from 200–800 V discharge voltage and at total discharge currents up to 250 A. The total power throttling range of the X3 is 2–200 kW. The thruster is approximately 80 cm in diameter and weighs 230 kg. Each of the three discharge channels features an inner and outer electromagnet for a total of six, each of which is controlled separately.”

Some performance characteristics

During high-power tests of the X3, Hall et al. “successfully measured the performance of the X3 for a range of conditions spanning total power levels from 5 to 102 kW. These conditions consisted of discharge voltages from 300 to 500 V and current densities that were 0.63, 1.00, and 1.26 of a reference value. The seven channel combinations of the thruster were throttled across this range of settings. For each test point, [Hall et al.] directly measured thrust using a high-power inverted-pendulum thrust stand, and from those thrust measurements and thrust telemetry, we calculated specific impulse and efficiency values.” The “results demonstrated that a three-channel 100-kW class NHT can offer comparable or even improved performance over high-power single-channel thrusters. The X3 demonstrated total efficiencies ranging from 0.54–0.67 and total specific impulses from 1800–2650 seconds [during this test], experiencing the peak efficiency at 500 V discharge voltage.

https://phys.org/news/2018-02-x3-ion-thruster-propel-mars.html
https://www.nasa.gov/content/characterization-of-a-100-kw-three-channel-nested-hall-thruster/

https://en.wikipedia.org/wiki/Hall-effect_thruster

A Hall-effect thruster is different than the DS4G and a magnetic plasma dynamic (MPD) thruster, which was considered during the mid-1980s. Each technology has it's challenges, which are common, e.g., material erosion/degradation. They all need a power source, which would be nuclear.

 

[darkdave3000]

The Plasmadynamics & Electric Propulsion Laboratory has webpage that answers most of one's questions.
https://pepl.engin.umich.edu/project/x3-nested-channel-hall-thruster/
Some performance characteristicshttps://phys.org/news/2018-02-x3-ion-thruster-propel-mars.html
https://www.nasa.gov/content/characterization-of-a-100-kw-three-channel-nested-hall-thruster/

https://en.wikipedia.org/wiki/Hall-effect_thruster

A Hall-effect thruster is different than the DS4G and a magnetic plasma dynamic (MPD) thruster, which was considered during the mid-1980s. Each technology has it's challenges, which are common, e.g., material erosion/degradation. They all need a power source, which would be nuclear.

Could the waste heat from the nuclear electric be used as a photon rocket? If the radiators are shaped ? 200 megawatts of infrared radiation with very high specific impulse.

 

[Astronuc]

Could the waste heat from the nuclear electric be used as a photon rocket?

One could, but that would not produce much thrust, and would be rather impractical.

I suggest one do a calculation for the thrust obtained from a radiator rejecting 200 MW infrared radiation.

If the radiators are shaped ?

Note that the radiator emits photons from all of the surface.

 

[darkdave3000]

One could, but that would not produce much thrust, and would be rather impractical.

I suggest one do a calculation for the thrust obtained from a radiator rejecting 200 MW infrared radiation.Note that the radiator emits photons from all of the surface.

Actually the numbers have been done
https://space.stackexchange.com/questions/3591/could-radiated-heat-propel-space-craft-in-outer-space

The person who corrected my math believes 4kN could be yielded from 200MW. To me 4000 Newtons from 200MW that is useful addition to what the ion drives can produce.

 

[Astronuc]

Actually the numbers have been done
https://space.stackexchange.com/questions/3591/could-radiated-heat-propel-space-craft-in-outer-space

The person who corrected my math believes 4kN could be yielded from 200MW. To me 4000 Newtons from 200MW that is useful addition to what the ion drives can produce.

There is not enough detail to determine whether the calculations are correct. One would need to provide the basis. I don't see any rigor in the calculation, and I'm not going to put effort into correcting others' work.

Also, note radiator would need to face normal and opposite to the intended trajectory, AND it radiates from both sides, to have the radiation is in the correct/desired, and the other half is in the undesired direction of travel. One also needs to look at W/m² and the amount of sunshine. Certainly intercepting radiant heat from the sun would assist in propulsion, much like a solar sale.

There are a lot of material performance issues, especially with high temperature devices. Erosion of components will degrade performance, and then there is always the possibility of malfunction or system fault(s), and how quickly they can be overcome and system restored to designed performance. A long downtime may require corrective performance above design. Failure means potential loss of personnel.

 

[darkdave3000]

There is not enough detail to determine whether the calculations are correct. One would need to provide the basis. I don't see any rigor in the calculation, and I'm not going to put effort into correcting others' work.

Also, note radiator would need to face normal and opposite to the intended trajectory, AND it radiates from both sides, to have the radiation is in the correct/desired, and the other half is in the undesired direction of travel. One also needs to look at W/m² and the amount of sunshine. Certainly intercepting radiant heat from the sun would assist in propulsion, much like a solar sale.

There are a lot of material performance issues, especially with high temperature devices. Erosion of components will degrade performance, and then there is always the possibility of malfunction or system fault(s), and how quickly they can be overcome and system restored to designed performance. A long downtime may require corrective performance above design. Failure means potential loss of personnel.

What if the stream of ions from the ion drive was in close contact with the radiator panels as part of the nozzle design? Could heat be transferred this way to the ions thus turning them into hot ions or plasma to increase Isp and energy efficiency this way?

Or have the fuel (Xenon or H2) act as a coolant for the reactor before they are used and ejected by the ion drive? Thus transfering all the waste heat into the propellant. Would that also boost Isp/efficiency?

 

[Astronuc]

What if the stream of ions from the ion drive was in close contact with the radiator panels as part of the nozzle design? Could heat be transferred this way to the ions thus turning them into hot ions or plasma to increase Isp and energy efficiency this way?

No, it would not improve performance. Instead, the propellant plume would heat the exposed face of the radiator making it much less efficient.

Once the ions have been accelerated, they are neutralized through recombination with electrons that were stripped in the ionization process. The idea is to prevent an accumulation of a negative charge on the spacecraft, which would draw the ions back toward the spacecraft, which would make propulsion less efficient.

 

[darkdave3000]

No, it would not improve performance. Instead, the propellant plume would heat the exposed face of the radiator making it much less efficient.

Once the ions have been accelerated, they are neutralized through recombination with electrons that were stripped in the ionization process. The idea is to prevent an accumulation of a negative charge on the spacecraft, which would draw the ions back toward the spacecraft, which would make propulsion less efficient.

In case you missed this edited part what about:

"Or have the fuel (Xenon or H2) act as a coolant for the reactor before they are used and ejected by the ion drive? Thus transfering all the waste heat into the propellant. Would that also boost Isp/efficiency?"

 

[Astronuc]

"Or have the fuel (Xenon or H2) act as a coolant for the reactor before they are used and ejected by the ion drive? Thus transfering all the waste heat into the propellant. Would that also boost Isp/efficiency?"

Xenon is not an ideal coolant. Xenon has low thermal conductivity. Hydrogen has very high thermal conductivity. One challenge with hydogen is the relatively high (13.6 eV) ionization energy, with Xe (12.1 eV), which still relatively high compared to other elements. Propellant storage is another complication.

Thus transfering all the waste heat into the propellant.

Would have to be done as preheating.

 

[darkdave3000]

Xenon is not an ideal coolant. Xenon has low thermal conductivity. Hydrogen has very high thermal conductivity. One challenge with hydogen is the relatively high (13.6 eV) ionization energy, with Xe (12.1 eV), which still relatively high compared to other elements. Propellant storage is another complication.Would have to be done as preheating.

So it could be worth doing to utilise the otherwise wasted heat from a nuclear electric reactor? Would the hydrogen being slightly warmer actually produce more practical efficiency/ISP if that other 400,% of thermal wattage for every electric watt generated is captured by hydrogen? Assuming we have a higher voltage ion drive that can use hydrogen.

 

[Astronuc]

So it could be worth doing to utilise the otherwise wasted heat from a nuclear electric reactor? Would the hydrogen being slightly warmer actually produce more practical efficiency/ISP if that other 400,% of thermal wattage for every electric watt generated is captured by hydrogen? Assuming we have a higher voltage ion drive that can use hydrogen.

The idea of preheating the propellant would be to reduce the required radiator surface (mass), which is basically 'deadweight' mass. However, the system becomes more complicated with respect to propellant consumption (mass flow rate) and heat transport. The thruster may also need cooling, depending on how one allows (design operating temperature). One has to balance performance against material degradation, which includes erosion, creep (slow distortion of geometry), and fatigue (initiation and propagation of internal flaws to point of catastrophic failure). The nuclear fuel and reactor have there own performance issues, as do all other components, and they are intimately couple such that failure of one can cascade to failures of others and the entire system.

 

[mfb]

Actually the numbers have been done
https://space.stackexchange.com/questions/3591/could-radiated-heat-propel-space-craft-in-outer-space

The person who corrected my math believes 4kN could be yielded from 200MW. To me 4000 Newtons from 200MW that is useful addition to what the ion drives can produce.

Where do you see that number in the linked thread? Even with a perfectly collimated emission, 200 MW only provide 200MW/c = 0.66 N of thrust.

The paper discussing 20 µN/W isn't accessible but it's probably in a setup with mirrors, greatly amplifying the thrust - if you have a laser and a stationary mirror to use nearby.

 

[darkdave3000]

Where do you see that number in the linked thread? Even with a perfectly collimated emission, 200 MW only provide 200MW/c = 0.66 N of thrust.

The paper discussing 20 µN/W isn't accessible but it's probably in a setup with mirrors, greatly amplifying the thrust - if you have a laser and a stationary mirror to use nearby.

I think you're right but I'm not sure, I'm anxiously waiting for Astronuc to respond to this.

 

[Astronuc]

Even with a perfectly collimated emission, 200 MW only provide 200MW/c = 0.66 N of thrust.

Yes, perfectly collimated, i.e., if all the photons traveled in the same direction, which for thrust is the opposite direction of travel. In reality a radiator surface emits in a 2pi solid angle, so the resuting thrust will be much less.

I had previously indicated that a radiator will emit radiation from all surfaces (or both surface of a plane).

I think you're right but I'm not sure, I'm anxiously waiting for Astronuc to respond to this.

Why? I agree with mfb, but don't take my word for it. One should do the calculation(s) oneself.

For photons, the magnitude of momentum is given by p = E/c, where E is the energy and c is the speed of light.

 

[darkdave3000]

The idea of preheating the propellant would be to reduce the required radiator surface (mass), which is basically 'deadweight' mass. However, the system becomes more complicated with respect to propellant consumption (mass flow rate) and heat transport. The thruster may also need cooling, depending on how one allows (design operating temperature). One has to balance performance against material degradation, which includes erosion, creep (slow distortion of geometry), and fatigue (initiation and propagation of internal flaws to point of catastrophic failure). The nuclear fuel and reactor have there own performance issues, as do all other components, and they are intimately couple such that failure of one can cascade to failures of others and the entire system.

Hi, so I read through all this, are you trying to say that absorbing the waste heat with the propellant before electrifying it does nothing at all? I am trying to discern the bottom line here. So imagine a 400MW system and 100MW is turned into electrical power, the remaining 300MW is absorbed by the propellant before being shot out the ion thruster. Any additional thrust? Lets pretend it's hydrogen not Xenon.

 

[darkdave3000]

Yes, perfectly collimated, i.e., if all the photons traveled in the same direction, which for thrust is the opposite direction of travel. In reality a radiator surface emits in a 2pi solid angle, so the resuting thrust will be much less.

I had previously indicated that a radiator will emit radiation from all surfaces (or both surface of a plane).Why? I agree with mfb, but don't take my word for it. One should do the calculation(s) oneself.

For photons, the magnitude of momentum is given by p = E/c, where E is the energy and c is the speed of light.

Hi, yes I concur after adjusting the formula in my spreadsheet for p = E/c.

 

[Astronuc]

Hi, so I read through all this, are you trying to say that absorbing the waste heat with the propellant before electrifying it does nothing at all?

No, I did not indicate that. Certainly pre-heating propellant with waste heat is beneficial, one has to consider the enthalphy (temperature) change of the propellant (it can take only so much) and the mass flow rate. One has to do the thermal and mass balances.

 

[darkdave3000]

No, I did not indicate that. Certainly pre-heating propellant with waste heat is beneficial, one has to consider the enthalphy (temperature) change of the propellant (it can take only so much) and the mass flow rate. One has to do the thermal and mass balances.

I suppose there isn't sufficient flow rate to absorb 300MW, I am thinking the entire tank of hydrogen will have to absorb the heat and build up pressure as the hydrogen is slowly drained for there to be some benefit. I will see if I can calculate how much of the outflowing hydrogen can absorb. Is it possible for you to confirm that a hydrogen version of any existing xenon ion drive will offer 11 times more Isp seconds?

 

[mfb]

So imagine a 400MW system and 100MW is turned into electrical power, the remaining 300MW is absorbed by the propellant before being shot out the ion thruster. Any additional thrust? Lets pretend it's hydrogen not Xenon.

That'll melt your engine. If you try to heat your tank with it then it will melt your tank. Either way, you won't increase thrust notably without reaching unacceptable temperatures.

Is it possible for you to confirm that a hydrogen version of any existing xenon ion drive will offer 11 times more Isp seconds?

See the earlier discussion, this is an absurd comparison.

 

[darkdave3000]

That'll melt your engine. If you try to heat your tank with it then it will melt your tank. Either way, you won't increase thrust notably without reaching unacceptable temperatures.See the earlier discussion, this is an absurd comparison.

https://space.stackexchange.com/que...en-or-helium-been-explored-for-ion-propulsion

You might want to read this page.

 

[Astronuc]

https://space.stackexchange.com/que...en-or-helium-been-explored-for-ion-propulsion

You might want to read this page.

No, we really don't want to wade through that mess. It would be better to find textbooks or journal articles (which are ostensibly peer-reviewed), rather than read stackexhange.

Is it possible for you to confirm that a hydrogen version of any existing xenon ion drive will offer 11 times more Isp seconds?

Ideally, yes, which we have discussed, but in reality, the physics is more complicated than one simple equation.
https://www.physicsforums.com/threads/esas-dual-stage-4-grid-ion-propulsion.1051952/post-6879665

Lighter atoms can be accelerated faster than heavier atoms, but then one has to consider recombination (neutralization), which precludes further acceleration. Propellant storage is another matter.

Dawn is discussed here - https://solarsystem.nasa.gov/missions/dawn/technology/ion-propulsion/

 

[darkdave3000]

No, we really don't want to wade through that mess. It would be better to find textbooks or journal articles (which are ostensibly peer-reviewed), rather than read stackexhange.Ideally, yes, which we have discussed, but in reality, the physics is more complicated than one simple equation.
https://www.physicsforums.com/threads/esas-dual-stage-4-grid-ion-propulsion.1051952/post-6879665

Lighter atoms can be accelerated faster than heavier atoms, but then one has to consider recombination (neutralization), which precludes further acceleration. Propellant storage is another matter.

Dawn is discussed here - https://solarsystem.nasa.gov/missions/dawn/technology/ion-propulsion/

Why do you keep referencing Dawn, isn't the latest ion drive with the highest performing figures the NEXT-C used in the DART mission? Isn't that a better reference?

 

[Astronuc]

Why do you keep referencing Dawn, isn't the latest ion drive with the highest performing figures the NEXT-C used in the DART mission? Isn't that a better reference?

You have previously asked about Dawn and Deep Space 1

Or is your Dawn Engine still the leading contender?

Was there any difference between Deep Space 1 and Dawn's ion drives?

It's appropriate to reference existing, deployed systems, since they actually worked in service. The one would ask, can we do better. Note that all long distance missions to date have deployed gravity assist maneuvers, and have taken many years to complete.

Looking at Deep Space 1, the spacecraft had a mass of 1,071 pounds (486 kilograms), or less than 0.5 tonne. So, very small. If one want to seen a manned crew on a long mission, one would be looking at many metric tons of space craft.

Note that a heliocentric orbit for DS1 was achieved with a third stage (chemical rocket).
https://solarsystem.nasa.gov/missions/deep-space-1/in-depth/

On Nov. 10 controllers commanded the ion thruster to fire for the first time but it operated for only 4.5 minutes before stopping.

On Nov. 24, 1998, controllers once again fired Deep Space 1’s ion propulsion system (fueled by xenon gas) when the spacecraft was about 3 million miles (4.8 million kilometers) from Earth. This time, the engine ran continuously for 14 days and demonstrated a specific impulse of 3,100 seconds, as much as 10 times higher than possible with conventional chemical propellants.

On July 29, 1999, was traveling at a velocity of about 10 miles per second (15.5 kilometers per second) while passing near-Earth asteroid 9660 Braille.

By the end of 1999, DS1’s ion engine had expended 48.5 pounds (22 kilograms) of xenon to impart a total change in velocity (delta-v) of 4,265 feet per second (1,300 m/s), or 1.3 km/s, which is not fast.On its way to Borrelly, it set the record for the longest operating time for a propulsion system in space. By Aug. 17, 2000, the engine had been operating for 162 days as part of an eight-month run.

The spacecraft’s ion engine was finally turned off Dec. 18, 2001, having operated for 16,265 hours and provided a total change in velocity (delta-v) of about 3 miles per second (4.3 kilometers per second), the largest delta-v achieved by a spacecraft with its own propulsion system.

NSTAR was apparently used on DS1 - https://en.wikipedia.org/wiki/NASA_Solar_Technology_Application_Readiness
https://www.researchgate.net/public...opulsion_system_on_the_Deep_Space_One_mission

and DAWN - https://en.wikipedia.org/wiki/Dawn_(spacecraft)#Propulsion_system

Where are we today - NEXT-C in the DART space craft?
https://dart.jhuapl.edu/Mission/Impactor-Spacecraft.php

The total mass of the DART spacecraft was approximately 1,345 pounds (610 kilograms) at launch and roughly 1280 pounds (580 kilograms) at impact. DART carries both hydrazine propellant (about 110 pounds, or 50 kilograms) for spacecraft maneuvers and attitude control, and xenon (about 130 pounds, or 60 kilograms) to operate the ion propulsion technology demonstration engine.

https://dart.jhuapl.edu/Mission/Impactor-Spacecraft.php

NEXT-C is a solar-powered electric propulsion system, using a gridded ion engine producing thrust by electrostatic acceleration of ions (electrically charged atoms) formed from the xenon propellant. NEXT–C offers improved performance (higher specific impulse and throughput), fuel efficiency, and operational flexibility compared to the ion propulsion systems flown on NASA's previous planetary mission of Dawn and Deep Space 1.

https://www1.grc.nasa.gov/space/sep/gridded-ion-thrusters-next-c/

NEXT-C specifics:

Performance​

The NEXT engine is a type of electric propulsion in which thruster systems use electricity to accelerate the xenon propellant to speeds of up to 90,000mph (145,000km/h or 40 km/s). NEXT can produce 6.9 kW thruster power and 236 mNthrust. It can be throttled down to 0.5kW power, and has a specific impulse of 4,190 seconds (compared to 3,120 for NSTAR). The NEXT thruster has demonstrated a total impulse of 17 MN·s; which is the highest total impulse ever demonstrated by an ion thruster.[2] A beam extraction area 1.6 times that of NSTAR allows higher thruster input power while maintaining low voltages and ion current densities, thus maintaining thruster longevity.

https://www1.grc.nasa.gov/wp-content/uploads/NEXT-C_FactSheet_11_1_21_rev4.pdf
System input power - 0.6 – 7.4 kW
Thrust - 25-235 mN
Isp - 4220 s max, 4190 s was mentioned in the webpage article, which is a 1.34 improvement over NSTAR (used on DS1 and DAWN) ~3100 s (90 mN thrust). NEXT-C thrust is about 2.6 times that of NSTAR.
https://en.wikipedia.org/wiki/NASA_Solar_Technology_Application_Readiness#ApplicationsDART apparently experienced problems with the NEXT-C propulsion.
https://en.wikipedia.org/wiki/Double_Asteroid_Redirection_Test#Ion_thruster

Early tests of the ion thruster revealed a reset mode that induced higher current (100 A) in the spacecraft structure than expected (25 A). It was decided not to use the ion thruster further as the mission could be accomplished without it, using conventional thrusters fueled by the 110 pounds of hydrazine onboard.

NSTAR performance: "The 30-cm ion thruster operates over a 0.5 kW to 2.3 kW input power range providing thrust from 19 mN to 92 mN. The specific impulse ranges from 1900 s at 0.5 kW to 3100 s at 2.3 kW." Ref: https://en.wikipedia.org/wiki/NASA_Solar_Technology_Application_Readiness#Performance

NEXT-C performance: "NEXT can consume 6.9 kW power to produce 237 mN thrust, with a specific impulse of 4,170 seconds", or can be throttled down to 0.5 kW power, when it has a specific impulse of 1320 seconds. Ref: https://en.wikipedia.org/wiki/NEXT_(ion_thruster)#Performance
https://ntrs.nasa.gov/citations/20110000521

Some values of performance characteristics seem to vary among different sources.

NEXT-C gave greater thrust and Isp by using greater power levels. One has to consider available kW, and the voltage and current. It would be useful to dig deeper into the NSTAR and NEXT designs.

6.9 kW (NEXT)/2.3 kW (NSTAR) ~ 3, which gave an improvement of thrust 237 mN (NEXT)/90 mN (NSTAR) ~ 2.6, with an Isp improvement of 1.34.

The missions of DS1, DAWN and DART were different, so in addition to differences in power supply, it is difficult to compare the relative performance of the thruster involved.

Still, there are a long way from kN thrust levels, or multi-MW power supplies.

 

[Astronuc]

Here are some equations and numbers to consider
http://electricrocket.org/IEPC/IEPC-2009-157.pdf Argon is better than Krypton is better than Xenon in terms of Isp, but at a cost of lower thrust density. One has to consider thrust, power requirements and mass of power system (in addition to payload and propellant mass).https://arc.aiaa.org/doi/abs/10.2514/6.2004-4106

Gridded (electrostatic) technologies - https://en.wikipedia.org/wiki/Ion_thruster#Electrostatic_thrusters

Electromagnetic thrusters - https://en.wikipedia.org/wiki/Ion_thruster#Electromagnetic_thrusters
https://en.wikipedia.org/wiki/Magnetoplasmadynamic_thruster

On the thruster, reducing/minimizing beam divergence is one of many challenges.

 

[darkdave3000]

Here are some equations and numbers to consider
http://electricrocket.org/IEPC/IEPC-2009-157.pdf Argon is better than Krypton is better than Xenon in terms of Isp, but at a cost of lower thrust density. One has to consider thrust, power requirements and mass of power system (in addition to payload and propellant mass).https://arc.aiaa.org/doi/abs/10.2514/6.2004-4106

Gridded (electrostatic) technologies - https://en.wikipedia.org/wiki/Ion_thruster#Electrostatic_thrusters

Electromagnetic thrusters - https://en.wikipedia.org/wiki/Ion_thruster#Electromagnetic_thrusters
https://en.wikipedia.org/wiki/Magnetoplasmadynamic_thruster

On the thruster, reducing/minimizing beam divergence is one of many challenges.

Is it possible to build an ion thruster that can accomodate multiple types of propellant? For example able to use Xenon but can also use Hydrogen so that when you run out of Xenon in space you can mine some water and extract the hydrogen and use it?