ESA's Dual Stage 4 Grid Ion propulsion

[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?

 

[Astronuc]

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

Certainly. If one reads various studies, one finds that thrusters have been tested with different propellants.

However, it is unlikely that a propulsion system would use alternative propellants during a mission. If hydrogen offered superior performance, then hydrogen would be used throughout the mission.

when you run out of Xenon in space you can mine some water and extract the hydrogen and use it?

Current practice is to carry propellant with the craft until mission is completed. Mining water to extract hydrogen would require an entirely different infrastructure. I would imagine that a plant would be established whereby a transport ship docks with a refueling station near or en route to a destination. It would make more sense to extract ammonia or methane from a moon, e.g., Titan (Saturn), that is rich in hydrogen.

https://www.nasa.gov/mission_pages/cassini/media/methane20060302.html

However, let's not get to far off topic, which is the DS4G, which is a type of electrostatic thruster.
https://beyondnerva.com/electric-propulsion/gridded-ion-thrusters/

In theory, the DS4G is more efficient in terms of I_sp, e.g., 10000-15000 (using Xe), compared to about 4000-8000 for more conventional grid thrusters. The greater I_sp, the lower the thrust for a given power level. The 'best' thrust reported for the DS4G was 5.4 mN vs 237 mN for NEXT-C. However, NEXT-C used approximately 6.9 - 7 kW, vs about 0.61 kW for DS4G. One really needs to compare technologies on the same basis, e.g., same kW level, and mass flow rate.

Erosion of the grid is a long term problem.
https://beyondnerva.com/electric-propulsion/gridded-ion-thrusters/

 

[darkdave3000]

Certainly. If one reads various studies, one finds that thrusters have been tested with different propellants.

However, it is unlikely that a propulsion system would use alternative propellants during a mission. If hydrogen offered superior performance, then hydrogen would be used throughout the mission.Current practice is to carry propellant with the craft until mission is completed. Mining water to extract hydrogen would require an entirely different infrastructure. I would imagine that a plant would be established whereby a transport ship docks with a refueling station near or en route to a destination. It would make more sense to extract ammonia or methane from a moon, e.g., Titan (Saturn), that is rich in hydrogen.

https://www.nasa.gov/mission_pages/cassini/media/methane20060302.html

However, let's not get to far off topic, which is the DS4G, which is a type of electrostatic thruster.
https://beyondnerva.com/electric-propulsion/gridded-ion-thrusters/

In theory, the DS4G is more efficient in terms of I_sp, e.g., 10000-15000 (using Xe), compared to about 4000-8000 for more conventional grid thrusters. The greater I_sp, the lower the thrust for a given power level. The 'best' thrust reported for the DS4G was 5.4 mN vs 237 mN for NEXT-C. However, NEXT-C used approximately 6.9 - 7 kW, vs about 0.61 kW for DS4G. One really needs to compare technologies on the same basis, e.g., same kW level, and mass flow rate.

Erosion of the grid is a long term problem.
https://beyondnerva.com/electric-propulsion/gridded-ion-thrusters/

I thought DS4G solves the problem of erosion????

Also thought you might want to see this:

The left rectangular is using your numbers, looks like according to you the Dual Stage 4 Grid has a better power efficiency. But on my right with the numbers I extracted from the wikipedia the numbers are shifted in favor of Next-C. I guess maybe I got theoretical numbers.

 

[Astronuc]

The left rectangular is using your numbers, looks like according to you the Dual Stage 4 Grid has a better power efficiency.

They are not 'my' numbers. I did not do the tests, nor measure performance. I simply reported what is cited in the literature. I've not made any endorsement, expressed or implied, regarding the validity of the data. I do point out that one must compare different technologies on an equal basis.

I thought DS4G solves the problem of erosion????

It's a 'gridded' technology. Metal grids are subject to erosion from the impingement of atoms/ions, which pass through the grids. The greater the speed (kinetic energy) of the ions/atoms, the greater the erosion potential for a given fluid density, as well as temperature of the grid.

As I've indicated, further development of the DS4G is required.

 

[darkdave3000]

They are not 'my' numbers. I did not do the tests, nor measure performance. I simply reported what is cited in the literature. I've not made any endorsement, expressed or implied, regarding the validity of the data. I do point out that one must compare different technologies on an equal basis.It's a 'gridded' technology. Metal grids are subject to erosion from the impingement of atoms/ions, which pass through the grids. The greater the speed (kinetic energy) of the ions/atoms, the greater the erosion potential for a given fluid density, as well as temperature of the grid.

As I've indicated, further development of the DS4G is required.

Can you reply my conversation I started with you?
https://www.physicsforums.com/conversations/the-space-plane-corporation.240119/#convMessage-362895

 

[Peter034]

The answer is not simple, because is depends on the thermodynamic efficiency of the entire system. We were targeting 100 MWe from 300 MWt, or about 0.33 efficiency, which could be greater or lesser depending on the thermodynamic cycle for thermal to mechanical conversion. I don't have my notes at hand, but it was something like 100 MWe and perhaps 10 tonnes (metric tons, or 1000 kg) for the reactor, or about 10 MWe/tonne, or 10 kWe/kg. However, one has to consider the rest of the mass of all the equipment, which would reduce about an order or magnitude, or about 1 kWe/kg. As mentioned earlier, the radiator was the single largest mass. The mass depends on how much heat must be rejected (area), the temperature at which the radiator operates, and the alloys used to construct the radiator.

Well that's real interesting. I can engineer how to fix my garage but when I tried to imagine a hundreds of megawatts radiator fit for an advanced propulsion system in near space then it started out looking like about 88 watts per kilogram with the latest thermal salt which would not have been known to science 35 years ago. If maybe you call that kind of stuff a coolant then I am curious what kind of it you would have been using for temperatures in the vicinity of 1500 K?