ESA's Dual Stage 4 Grid Ion propulsion

In summary: It is not clear what you are asking. The reactor and power conversion system would be a nuclear engineering topic. The wattage per ton would be determined by the power conversion system.
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  • #5
darkdave3000 said:
Does it work?
The paper is a for a proposed concept. As far as I can tell it hasn't been built, so as of yet it doesn't work. But i think it probably would.

But that's not really what you want to know. You want to know why we aren't powering spaceships with this technology - what's wrong with it The main answer is given in the wiki article: it has a high power requirement.
 
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  • #6
The ESA paper indicates that the DS4G was in a laboratory (R&D) prototype stage, which was used to demonstrate the capability. The demonstration was back in 2005-2006. The maximum thrust was 5.4 mN. See below the comment on scalability.

The DS4G was discussed on PF about 17 years ago.
https://www.physicsforums.com/threa...ough-new-spacecraft-ion-engine-tested.108685/ - some links got broken since.

ESA summary of their activities (from 2006 and appears in the ESA pdf).
https://www.esa.int/gsp/ACT/projects/ds4g_overview/

Defense Systems Information Analysis Center (DSIAC) overview of various propulsion concepts
https://dsiac.org/articles/space-travel-aided-by-plasma-thrusters-past-present-and-future/

It does seem that not much work has been done since the prototype was tested (in 2006).

The ESA paper in the OP mentions mission application, which is conceptual.
An optimum design solution was found for an 8000 kg spacecraft wet mass, 400 kg payload, with 65 kW reactor and
DS4G-type thruster operating at 10,200 seconds specific impulse and thrust of 0.9 N. A single 25 cm diameter DS4G-type thruster with a beam potential of 13000 Volts would be sufficient for these requirements. The benefits compared to a 1-ton, 1 kW-class RTG-powered 3-grid ion thruster spacecraft include reduced trip time (23 years compared to 30 years), increased payload, . . . .
A thrust of 0.9 N would require 167 of the prototype DS4Gs at 0.0054 N thrust each (5.4 mN). I don't believe they have demonstrated a scaled up version.Back ~35+ years ago, I was part of team that looked into similar concepts. We however were considering kN levels of thrust. The power plant was a compact fast reactor of about 300 MWt (operating at ~1500 K, or as high as we could get it subject to thermomechanical limits). Besides the mass of the core, control system and containment, one has to consider masses of the power conversion system, shielding, propellant and storage, and masses of the thrusters. The largest component was the radiator - to reject heat. I found a way to reduce radiator.

This is an Aerospace Engineering/Science topic. The nuclear reactor and NPP (design and operation) would be a nuclear engineering topic. The power conversion and propulsion system (design and operation) is mostly electrical engineering and applied physics.
 
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  • #7
300MWt does that mean 300 Megawatts per metric ton? And does that include the cooling mechanism?
 
  • #8
darkdave3000 said:
300MWt does that mean 300 Megawatts per metric ton?
The 300 MWt was the total thermal output. I was mainly concerned with the reactor and plant design. We used the knowledge of the previous decades from the 1940s through the early 1980s, but we employed some exotic technology. It was a conceptual preliminary design. NASA however was less ambitious in favor of smaller systems, e.g., SP-100, and the Russian TOPAZ nuclear system.

darkdave3000 said:
And does that include the cooling mechanism?
That did include the radiator, which in space is how one removes heat from a thermal system. There is no atmosphere to conduct heat away. The radiator mass was a significant fraction of the total mass.
 
  • #9
Astronuc said:
The 300 MWt was the total thermal output. I was mainly concerned with the reactor and plant design. We used the knowledge of the previous decades from the 1940s through the early 1980s, but we employed some exotic technology. It was a conceptual preliminary design. NASA however was less ambitious in favor of smaller systems, e.g., SP-100, and the Russian TOPAZ nuclear system.That did include the radiator, which in space is how one removes heat from a thermal system. There is no atmosphere to conduct heat away. The radiator mass was a significant fraction of the total mass.

What electrical power wattage output per ton of reactor mass can we hope to produce in the near future?
 
  • #10
darkdave3000 said:
300MWt does that mean 300 Megawatts per metric ton?
The 't' here is not a separate unit. It doesn't mean tonne or anything else. It's merely a conventional way of denoting power (I.e. W) output that is thermal. You may also encounter 'We' for electric output. See Wikipedia article on Watt (unit).
It would probably be clearer if written as a subscript, though.
 
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  • #11
Bandersnatch said:
The 't' here is not a separate unit. It doesn't mean tonne or anything else. It's merely a conventional way of denoting power (I.e. W) output that is thermal. You may also encounter 'We' for electric output. See Wikipedia article on Watt (unit).
It would probably be clearer if written as a subscript, though.
Is this thermal power output useful at all? Can it be used to heat hydrogen in the style of Nuclear Thermal Reactors to shoot it out at 900 seconds of Isp? Also what amount of We can we hope to produce in the near future per metric ton of reactor mass in space?
 
  • #12
darkdave3000 said:
What electrical power wattage output per ton of reactor mass can we hope to produce in the near future?
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.
 
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  • #13
darkdave3000 said:
Is this thermal power output useful at all?
The nuclear reactor provides thermal energy, which is transferred to a working fluid, which could be hydrogen propellant (as in direct thrust, nuclear thermal rocket (NTR)), or a fluid, e.g., potassium, which is vaporized and feed to a turbine system (as in a potassium Rankine cycle).

With a low Isp, high thrust (from high mass flow), the thrust would be applied for a few days, then turned off, for a long coasting period, then near the destination, the spacecraft would turn and the thrust applied to slow the craft to whatever orbit velocity is required. The return trip (assuming a round trip) would repeat the cycle (thrust - coast - thrust).

Nuclear electric systems are designed to operate continuously at higher Isp, but producing lower thrust.
 
  • #14
Astronuc said:
The nuclear reactor provides thermal energy, which is transferred to a working fluid, which could be hydrogen propellant (as in direct thrust, nuclear thermal rocket (NTR)), or a fluid, e.g., potassium, which is vaporized and feed to a turbine system (as in a potassium Rankine cycle).

With a low Isp, high thrust (from high mass flow), the thrust would be applied for a few days, then turned off, for a long coasting period, then near the destination, the spacecraft would turn and the thrust applied to slow the craft to whatever orbit velocity is required. The return trip (assuming a round trip) would repeat the cycle (thrust - coast - thrust).

Nuclear electric systems are designed to operate continuously at higher Isp, but producing lower thrust.
But in this case the reactor is producing waste heat, not heat designed to heat hydrogen to a specific temperature. Will the temperature of such a reactor optimized for electricity generation be hot enough to give hydrogen passing through it the temperature it needs to give specific impulse of 800s or more? NTR are supposed to give at least that much Isp. If im not mistaken I think NTR designs heat h2 to 1500K?
 
  • #15
darkdave3000 said:
But in this case the reactor is producing waste heat, not heat designed to heat hydrogen to a specific temperature.
The reactor produces heat. In the case of an NTR, the heat goes into the propellant, e.g., H2. The temperature achieved by the reactor is constrained by the ability of the reactor to perform for a time at temperature, while maintaining its geometry and fuel integrity. There are very few alloys that can withstand 1500 K for a long time, e.g., 100 hrs, 1000 hrs, or 10,00 hrs. The longer the time, the lower the temperature because of creep, fatigue (from cavitation and vibration) and/or erosion, as well as constraints of tensile strength at temperature. Hydrogen embrittlement would be another factor, in conjunction with fatigue and tensile strength. A colleague showed me a video of NERVA/ROVER tests in which the core/fuel vibration was so bad that the fuel disintegrated and flew out of the core.

In the case of a nuclear electric system, the thermal to mechanical system must reject heat. For a thermal-to-mechanical efficiency of 33% (0.33), two-thirds of the energy must be rejected. When comparing an electric system to a thermal NTR, one has to consider the trade-offs.

For a nuclear electric system, we propsed a liquid lithium cooled reactor, which transferred heat to a potassium Rankine cycle. This required a heat exchanger to keep the lithium and potassium separated. Note that the large components also provide shielding to the other parts of the spacecraft. For the reactor core, the containment vessel provides much of the shielding, in addition to 'blankets' in the core. Blankets are comprised of special assemblies, usually natural or depleted U in a form compatible with the enriched fuel. For neutron shielding, one would make use of Y-hydride, which is relatively stable.
 
  • #16
Astronuc said:
The reactor produces heat. In the case of an NTR, the heat goes into the propellant, e.g., H2. The temperature achieved by the reactor is constrained by the ability of the reactor to perform for a time at temperature, while maintaining its geometry and fuel integrity. There are very few alloys that can withstand 1500 K for a long time, e.g., 100 hrs, 1000 hrs, or 10,00 hrs. The longer the time, the lower the temperature because of creep, fatigue (from cavitation and vibration) and/or erosion, as well as constraints of tensile strength at temperature. Hydrogen embrittlement would be another factor, in conjunction with fatigue and tensile strength. A colleague showed me a video of NERVA/ROVER tests in which the core/fuel vibration was so bad that the fuel disintegrated and flew out of the core.

In the case of a nuclear electric system, the thermal to mechanical system must reject heat. For a thermal-to-mechanical efficiency of 33% (0.33), two-thirds of the energy must be rejected. When comparing an electric system to a thermal NTR, one has to consider the trade-offs.

For a nuclear electric system, we propsed a liquid lithium cooled reactor, which transferred heat to a potassium Rankine cycle. This required a heat exchanger to keep the lithium and potassium separated. Note that the large components also provide shielding to the other parts of the spacecraft. For the reactor core, the containment vessel provides much of the shielding, in addition to 'blankets' in the core. Blankets are comprised of special assemblies, usually natural or depleted U in a form compatible with the enriched fuel. For neutron shielding, one would make use of Y-hydride, which is relatively stable.
So if it were up to you would you build a ship that is nuclear electric or hybrid nuclear electric and thermal to shoot out both xenon and hydrogen? Would you have both modes cycle between cooling the reactor in hydrogen mode and heating it up in xenon electric mode or operate in dual mode simultaneously?
 
  • #17
darkdave3000 said:
So if it were up to you would you build a ship that is nuclear electric or hybrid nuclear electric and thermal to shoot out both xenon and hydrogen?
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.
darkdave3000 said:
Would you have both modes cycle between cooling the reactor in hydrogen mode and heating it up in xenon electric mode or operate in dual mode simultaneously?
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.
 
  • #18
Astronuc said:
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 go with dual stage 4 grid for NEP?Also what is the ISP if hydrogen was substituted for xenon?
 
  • #19
darkdave3000 said:
Would you go with dual stage 4 grid for NEP?
I would certainly consider it. However, I'd have to study the concept. I'd like to know what has happened in terms of development since 2006/2007, and why the lack of reporting in the literature (budget limitations?), or as others mentioned, are the power requirements too great. Are the alternatives more suited to the missions (e.g., Galileo, Juno, Cassini, . . . ), which relied strongly on gravity-assist.

https://en.wikipedia.org/wiki/Exploration_of_Jupiter

darkdave3000 said:
Also what is the ISP if hydrogen was substituted for xenon?
On a simple basis of equivalent energy and using Isp = ve * g, where g is acceleration equivalent to Earth's gravity at sea level (9.80665 m/s2)

E(Xe) = E(H) and using E = 1/2 mv2, then mXevXe2 = mHvH2, or

vH = vXe sqrt(131), or vH = 11.4 * vXe, or

Isp(H) = 11.4 * Isp(Xe) with the caveat that the coefficient of 11.4 could be lower if there are energy losses in the hydrogen from recombination, or incomplete dissociation of molecular H2. The 11.4 is an ideal value, assuming that the average hydrogen atom has the same energy as the average Xe atom.

Edit/correction: I was writing Xe but thinking Cs (133). Xe has several stable isotopes, and one would want to used the lightest Xe-129, -130, -131, -132, so I took 131, and sqrt(131) = 11.45.
 
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  • #20
Astronuc said:
I would certainly consider it. However, I'd have to study the concept. I'd like to know what has happened in terms of development since 2006/2007, and why the lack of reporting in the literature (budget limitations?), or as others mentioned, are the power requirements too great. Are the alternatives more suited to the missions (e.g., Galileo, Juno, Cassini, . . . ), which relied strongly on gravity-assist.

https://en.wikipedia.org/wiki/Exploration_of_JupiterOn a simple basis of equivalent energy and using Isp = ve * g, where g is acceleration equivalent to Earth's gravity at sea level (9.80665 m/s2)

E(Xe) = E(H) and using E = 1/2 mv2, then mXevXe2 = mHvH2, or

vH = vXe sqrt(131), or vH = 11.4 * vXe, or

Isp(H) = 11.4 * Isp(Xe) with the caveat that the coefficient of 11.4 could be lower if there are energy losses in the hydrogen from recombination, or incomplete dissociation of molecular H2. The 11.4 is an ideal value, assuming that the average hydrogen atom has the same energy as the average Xe atom.

Edit/correction: I was writing Xe but thinking Cs (133). Xe has several stable isotopes, and one would want to used the lightest Xe-129, -130, -131, -132, so I took 131, and sqrt(131) = 11.45.

Im having trouble understanding your answer 11.45 is that seconds? I was expecting the answer in seconds.

And yes I also want to know why there is a lack of update regarding dual stage, I am wondering if this is government surpressing technology from the public for their own use or if it simply doesn't work. I am gravitating toward the first.
 
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  • #21
darkdave3000 said:
Im having trouble understanding your answer 11.45 is that seconds?
One asked:
darkdave3000 said:
what is the ISP if hydrogen was substituted for xenon?
I gave the answer in terms of an equivalence between the Isp of hydrogen and that of Xe.

Assuming that the atoms have the same kinetic energy, using hydrogen atoms would give an Isp that is 11.45 times that of Xe atoms (assuming atomic mass of Xe ~131 amu). The Isp is determined by the exhaust velocity of the atoms.

https://www.qrg.northwestern.edu/pr...ion/3-how-you-calculate-specific-impulse.html

https://www.grc.nasa.gov/www/k-12/airplane/specimp.html

The specific impulse is determined by the exhaust velocity of the propellant stream, by definition. The thrust (force) is a product of the mass flow rate and exhaust velocity. The challenge is to put as much energy into the propellant, or change in momentum of the propellant, within the nozzle/thruster.

darkdave3000 said:
I am wondering if this is government surpressing technology from the public for their own use
Sounds like hypothesis of a government conspiracy theory. We don't do that at PF, so don't go there.

VASIMR suffers from the same issue - power requirement - and putting a lot of energy into a small amount of mass very quickly. One has to work with the constraints that Nature imposes on us. Solid materials of melting points, as well as strength that diminishes with temperature, which constrains the temperature and pressure at which a thruster performs. Then one has to become creative and devise a way, e.g., electrostatic, electromagnetic or magneto-plasma-dynamic, to heat the propellant to whatever temperature can be achieved, then directing the propellant in the opposite direction of the desired direction. Also, in the case of a plasma, one must contend with the matter of neutralizing positive ions in the plasma with the electrons that were stripped away in order to accelerate the positive ions. In any solid system, one must address the cooling of the thrusters to minimize erosion of the structural material.

The DS4G concept was based on neutral beam injectors, which are used to push deuterium/tritium fuel into a plasma. It's not a trivial matter.
 
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  • #22
Astronuc said:
One asked:

I gave the answer in terms of an equivalence between the Isp of hydrogen and that of Xe.

Assuming that the atoms have the same kinetic energy, using hydrogen atoms would give an Isp that is 11.45 times that of Xe atoms (assuming atomic mass of Xe ~131 amu). The Isp is determined by the exhaust velocity of the atoms.

https://www.qrg.northwestern.edu/pr...ion/3-how-you-calculate-specific-impulse.html

https://www.grc.nasa.gov/www/k-12/airplane/specimp.html

The specific impulse is determined by the exhaust velocity of the propellant stream, by definition. The thrust (force) is a product of the mass flow rate and exhaust velocity. The challenge is to put as much energy into the propellant, or change in momentum of the propellant, within the nozzle/thruster.Sounds like hypothesis of a government conspiracy theory. We don't do that at PF, so don't go there.

VASIMR suffers from the same issue - power requirement - and putting a lot of energy into a small amount of mass very quickly. One has to work with the constraints that Nature imposes on us. Solid materials of melting points, as well as strength that diminishes with temperature, which constrains the temperature and pressure at which a thruster performs. Then one has to become creative and devise a way, e.g., electrostatic, electromagnetic or magneto-plasma-dynamic, to heat the propellant to whatever temperature can be achieved, then directing the propellant in the opposite direction of the desired direction. Also, in the case of a plasma, one must contend with the matter of neutralizing positive ions in the plasma with the electrons that were stripped away in order to accelerate the positive ions. In any solid system, one must address the cooling of the thrusters to minimize erosion of the structural material.

The DS4G concept was based on neutral beam injectors, which are used to push deuterium/tritium fuel into a plasma. It's not a trivial matter.
Thanks your answer is very useful, so that means that if Isp with Xenon is 19,000s as quoted in wikipeidia along with 0.2m diameter nozzel,250kW and 2.5Newtons over that area. That means the Isp with hydrogen in the same system will be almost 60,000 seconds. And what's great about this is that hydrogen you can replenish in space. Xenon not so easily! Mars in 1 week and Titan in 3 months.

Any theories about why there is remarkably little recent info/updates about dual stage?
 
  • #23
darkdave3000 said:
Any theories about why there is remarkably little recent info/updates about dual stage?
I'd have to find out who is now involved in the development of the concept.

At the time (2006) that results were published, it was stated that it would take 10 years to develop the concept. Unfortunately, in 2007, the chief proponent/developer of the DS4G, Dr. David Fearn, died, apparently unexpectedly. https://www.telegraph.co.uk/news/obituaries/1563505/David-Fearn.html

The status of DS4G is mentioned here: https://beyondnerva.wordpress.com/electric-propulsion/gridded-ion-thrusters/
There remain significant engineering challenges, but nothing that’s incredibly different from any other high powered ion drive. Indeed, many of the complications concerning ion optics, and electrostatic field impingement in the plasma chamber, are largely eliminated by the 4-grid design. Unfortunately, there are no missions that currently have funding that require this type of thruster, so it remains on the books as “viable, but in need of some final development for application” when there’s a high-powered mission to the outer solar system.
The University of Southampton has named a laboratory in his honor: David Fearn Electric Propulsion Lab
https://www.southampton.ac.uk/engineering/research/facilities/analytical_facilties.page

I do not know who took up the work after Fearn's death. I will have to make some inquiries.
 
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  • #24
Astronuc said:
I'd have to find out who is now involved in the development of the concept.

At the time (2006) that results were published, it was stated that it would take 10 years to develop the concept. Unfortunately, in 2007, the chief proponent/developer of the DS4G, Dr. David Fearn, died, apparently unexpectedly. https://www.telegraph.co.uk/news/obituaries/1563505/David-Fearn.html

The status of DS4G is mentioned here: https://beyondnerva.wordpress.com/electric-propulsion/gridded-ion-thrusters/
The University of Southampton has named a laboratory in his honor: David Fearn Electric Propulsion Lab
https://www.southampton.ac.uk/engineering/research/facilities/analytical_facilties.page

I do not know who took up the work after Fearn's death. I will have to make some inquiries.
Many thanks, I hope that he died from natural causes painlessly. Though unexpected deaths tend to be engineered. But we won't go there.
 
  • #25
darkdave3000 said:
Though unexpected deaths tend to be engineered.
Most of the time it is an undiagnosed condition, e.g., sudden heart attack (myocardial infarction, atrial or ventricular fibrillation, . . . ). I understand from the obituary that Fearn had a heart condition, and previously, a partially successful coronary by-pass surgery.
darkdave3000 said:
But we won't go there.
Please don't.

Please adhere to PF guidelines and stick to the scientific and technical discussion of DS4G and related topics.

As I recall, a benchmark of sorts is a specific power of 1 kW/kg or a specific mass of 1 kg/kW. Ideally, the specific power is >1 kW/kg or specific mass <1 kg/KW.

As I recall, the NTR Isp capability (with hydrogen propellant) was < 800 s, but in theory could have been slight greater, e.g., 850 sec. It is a challenge to design a custom nuclear power system for a long-range mission, especially if one requires a round trip. There are a lot of aspects, e.g., how and where to store the used stages, or used propellant tanks.

For trips to Mars, there was some notion to send a slow support structure to get to Martian orbit, and a separate personnel rapid transfer vehicle, which could rendezvous with the support system in transit (perhaps during deceleration) or on orbit. The crew transfer system would the refuel in Martian orbit and return to earth. It's a rather complicated concept.

To be viable, the DS4G or other devices would have to be scaled up to by about 3 orders of magnitude, from mN to N range, or preferably even more given that a large mass would require 100s N or kN range of thrust for an appreciable acceleration. Then there is the reliability and fault tolerance of the system to be demonstrated. For example, if one loses 10% thrust, or if the entire system trips off-line, how quickly can one recover, or conversely, how long can one coast before getting outside of the permissible or safe envelope. It would reqire an unmanned test before committing a crew to a round trip mission.
 
  • #26
darkdave3000 said:
Thanks your answer is very useful, so that means that if Isp with Xenon is 19,000s as quoted in wikipeidia along with 0.2m diameter nozzel,250kW and 2.5Newtons over that area. That means the Isp with hydrogen in the same system will be almost 60,000 seconds.
The same energy per hydrogen atom means 130 times the energy per mass, or 11.5 times the power for the same thrust. If you want to keep the power then your thrust decreases by the same factor 11.5. Changing the propellant is not a miracle solution. It's very similar to running with xenon at a different acceleration voltage. Hydrogen also comes with other disadvantages.

Existing ion thrusters are already pretty efficient, somewhere around 2/3 conversion of electrical power into the minimum power required for a given thrust and I_sp combination. Even a completely lossless system could only get ~50% more thrust with the same power and I_sp. If you see people developing completely new concepts then it's typically done in order to get larger thrusters instead of using hundreds of existing ones.
 
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  • #27
Astronuc said:
Most of the time it is an undiagnosed condition, e.g., sudden heart attack (myocardial infarction, atrial or ventricular fibrillation, . . . ). I understand from the obituary that Fearn had a heart condition, and previously, a partially successful coronary by-pass surgery.

Please don't.

Please adhere to PF guidelines and stick to the scientific and technical discussion of DS4G and related topics.

As I recall, a benchmark of sorts is a specific power of 1 kW/kg or a specific mass of 1 kg/kW. Ideally, the specific power is >1 kW/kg or specific mass <1 kg/KW.

As I recall, the NTR Isp capability (with hydrogen propellant) was < 800 s, but in theory could have been slight greater, e.g., 850 sec. It is a challenge to design a custom nuclear power system for a long-range mission, especially if one requires a round trip. There are a lot of aspects, e.g., how and where to store the used stages, or used propellant tanks.

For trips to Mars, there was some notion to send a slow support structure to get to Martian orbit, and a separate personnel rapid transfer vehicle, which could rendezvous with the support system in transit (perhaps during deceleration) or on orbit. The crew transfer system would the refuel in Martian orbit and return to earth. It's a rather complicated concept.

To be viable, the DS4G or other devices would have to be scaled up to by about 3 orders of magnitude, from mN to N range, or preferably even more given that a large mass would require 100s N or kN range of thrust for an appreciable acceleration. Then there is the reliability and fault tolerance of the system to be demonstrated. For example, if one loses 10% thrust, or if the entire system trips off-line, how quickly can one recover, or conversely, how long can one coast before getting outside of the permissible or safe envelope. It would reqire an unmanned test before committing a crew to a round trip mission.
Why do you say mN to N? A single unit was quoted in Wikipedia to generate 2.5N per 0.2m diameter for an input of 240kN. That is not mN. That is 2 whole newtons and a bit for the area and power input mentioned.
 
  • #28
mfb said:
The same energy per hydrogen atom means 130 times the energy per mass, or 11.5 times the power for the same thrust. If you want to keep the power then your thrust decreases by the same factor 11.5. Changing the propellant is not a miracle solution. It's very similar to running with xenon at a different acceleration voltage. Hydrogen also comes with other disadvantages.

Existing ion thrusters are already pretty efficient, somewhere around 2/3 conversion of electrical power into the minimum power required for a given thrust and I_sp combination. Even a completely lossless system could only get ~50% more thrust with the same power and I_sp. If you see people developing completely new concepts then it's typically done in order to get larger thrusters instead of using hundreds of existing ones.
That is fine, so 11.5 time less expenditure of propellant for 11.5 times more power requirements. In some situations that might be desirable. Especially if you ever run out of Xenon in space. One could always replace hydrogen from ice or from Jupiter if youre desperate enough to take the risk. But where can one find Xenon in space? Very difficult.

Im only talking about ion propulsion here regarding hydrogen btw. Im assuming the waste heat is not hot enough to get us the 800s of isp for hydrogen using thermal.

Regarding the death, it still seems a long time ago for nothing to have progressed since his passing.
 
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  • #29
darkdave3000 said:
Why do you say mN to N? A single unit was quoted in Wikipedia to generate 2.5N per 0.2m diameter for an input of 240kN. That is not mN. That is 2 whole newtons and a bit for the area and power input mentioned.
I'm going by the article (pdf file) by Cristina Bramanti et al, "The Innovative Dual-Stage 4-Grid Ion Thruster Concept – Theory and Experimental Results," DS4G-IAC-06-C4.4, International Astronautical Federation, 2006. ACT-RPR-PRO-IAC2006-DS4G-C4.4.7.pdf

Tables 1 and 2 provide the performance and operating characteristics of the prototype DS4G device for the first and second testing campaigns in Nov 2005 and May 2006, respectively. In the first campaign, the best value of thrust is 2.85 mN; maximum values of beam power and RF power are given as 260 W and 490 W, respectively, and beam diameter is 2.3 cm. Table 2 provides a best value for thrust as 5.4 mN, maximum values for beam power and RF power as 398 W and 316 W, respectively, and a beam diameter of 2 cm.

The tables also shows improvements in efficiency in the second campaign compared to the first.

The Wikipedia statement ("A 4-grid ion thruster with only 0.2 m diameter is projected to absorb 250 kW power. With that energy input rate, the thruster could produce a thrust of 2.5 N.") is a projection of what someone would expect for the capability of a scaled up version. It is an unsubstantiated claim; no basis is provided in support of such a claim.

In the second table of the ESA article, the total energy (sum of beam and RF power) is 614 W, or 0.614 kW. The thrust of 0.0054 N (5.4 mN) divided by 0.614 kW gives 0.00879 mN/kW (or 8.79 N/MW). Assuming that ratio can be applied to 250 kW (i.e., the same efficiencies can be realized in the larger version), then the device could produce about 2.2 N of thrust. This requires scaling the beam diameter from 2 cm to 200 cm, which implies a commensurate increase in the grids.

I am not aware that a DS4G device with a 0.2 m (200 cm) diameter has been built and tested. it's not clear what that would look like.

I don't take published articles at face value as they are summaries. For application or utilization of data, I'd want to see the design details and test reports.

Note that in the Wikipedia article, the section "Experiments proposed and tests done," is empty. The article certainly does not provide a basis for scalability from the prototype to the large scale.

Consider thrust requirements for Galileo and Cassini spacecraft.
https://en.wikipedia.org/wiki/Galileo_(spacecraft)#Propulsion
The propulsion subsystem consisted of a 400 N (90 lbf) main engine and twelve 10 N (2.2 lbf) thrusters, together with propellant, storage and pressurizing tanks and associated plumbing. The fuel for the system was 925 kg (2,039 lb) of monomethylhydrazine and nitrogen tetroxide.
To save on fuel and propulsion demands, the Galileo craft was sent from Earth to Venus and back passed Earth using gravity assist.
Launched during the STS 34 flight of the Atlantis orbiter, the two spacecraft were kicked out of Earth orbit by an inertial upper stage (IUS) rocket, sending them careening through the inner solar system. The trajectory which the spacecraft followed was called a VEEGA (Venus-Earth-Earth Gravity Assist), traveling first in toward the Sun for a gravity assist from Venus before encountering the Earth two times (spaced two years apart). These encounters with Venus and the Earth allowed Galileo to gain enough velocity to get it out to Jupiter.
https://nssdc.gsfc.nasa.gov/planetary/galileo.html#overview
https://www.jpl.nasa.gov/news/galileo-heads-towards-second-gravity-assist
https://spaceflightnow.com/galileo/030920overview.html

https://articles.adsabs.harvard.edu//full/1997ESASP.398...53K/0000056.000.html

Cassini had a slightly more powerful propulsion system - "Main (445 Newton) engine for propulsive maneuvers" and 16 monopropellant hydrazine thrusters of which eight were prime and eight were backups. https://solarsystem.nasa.gov/missions/cassini/engine/

Cassini was launched from Earth on 15 October 1997, "followed by gravity assist flybys of Venus (26 April 1998 and 21 June 1999), Earth (18 August 1999), and Jupiter (30 December 2000). Saturn arrival was on 1 July 2004."
https://solarsystem.nasa.gov/resources/11776/cassini-trajectory/

Getting a crewed spacecraft of many tons in short time, e.g., several weeks to a month, would require a substantial power plant and propulsion system. What would scaling from 2 N to 2 kN look like? From 250 kW to 250 MW?

The ESA paper has one comment, "Taken to the extreme, a human Mars mission using EP would require about 8-15 MW power and deliver a thrust of 90-180 N." That's more modest than 250 MW. On the other, the thust is fairly low 0.09-0.18 kN, and no basis is given for those thrust requirements. Taking the maximum values of 15 MWe and 180 N, that gives 12 N/MW, which is greater than the 8.79 N/MW achieved by the prototype DS4G. One then has to devise a nuclear power plant to provide up to 15 MWe, which might be a 50 MWt plant if an efficiency is 0.3.

The Kiwi NTR was tested at full power, 1 GW, briefly with a thrust of about 74,000 pounds, or 334 kN.

One can find some references on rocket vehicles here:
https://www.sciencedirect.com/topics/physics-and-astronomy/rocket-vehicles
 
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  • #30
Astronuc said:
I'm going by the article (pdf file) by Cristina Bramanti et al, "The Innovative Dual-Stage 4-Grid Ion Thruster Concept – Theory and Experimental Results," DS4G-IAC-06-C4.4, International Astronautical Federation, 2006. ACT-RPR-PRO-IAC2006-DS4G-C4.4.7.pdf

Tables 1 and 2 provide the performance and operating characteristics of the prototype DS4G device for the first and second testing campaigns in Nov 2005 and May 2006, respectively. In the first campaign, the best value of thrust is 2.85 mN; maximum values of beam power and RF power are given as 260 W and 490 W, respectively, and beam diameter is 2.3 cm. Table 2 provides a best value for thrust as 5.4 mN, maximum values for beam power and RF power as 398 W and 316 W, respectively, and a beam diameter of 2 cm.

The tables also shows improvements in efficiency in the second campaign compared to the first.

The Wikipedia statement ("A 4-grid ion thruster with only 0.2 m diameter is projected to absorb 250 kW power. With that energy input rate, the thruster could produce a thrust of 2.5 N.") is a projection of what someone would expect for the capability of a scaled up version. It is an unsubstantiated claim; no basis is provided in support of such a claim.

In the second table of the ESA article, the total energy (sum of beam and RF power) is 614 W, or 0.614 kW. The thrust of 0.0054 N (5.4 mN) divided by 0.614 kW gives 0.00879 mN/kW (or 8.79 N/MW). Assuming that ratio can be applied to 250 kW (i.e., the same efficiencies can be realized in the larger version), then the device could produce about 2.2 N of thrust. This requires scaling the beam diameter from 2 cm to 200 cm, which implies a commensurate increase in the grids.

I am not aware that a DS4G device with a 0.2 m (200 cm) diameter has been built and tested. it's not clear what that would look like.

I don't take published articles at face value as they are summaries. For application or utilization of data, I'd want to see the design details and test reports.

Note that in the Wikipedia article, the section "Experiments proposed and tests done," is empty. The article certainly does not provide a basis for scalability from the prototype to the large scale.

Consider thrust requirements for Galileo and Cassini spacecraft.
https://en.wikipedia.org/wiki/Galileo_(spacecraft)#Propulsion

To save on fuel and propulsion demands, the Galileo craft was sent from Earth to Venus and back passed Earth using gravity assist.

https://nssdc.gsfc.nasa.gov/planetary/galileo.html#overview
https://www.jpl.nasa.gov/news/galileo-heads-towards-second-gravity-assist
https://spaceflightnow.com/galileo/030920overview.html

https://articles.adsabs.harvard.edu//full/1997ESASP.398...53K/0000056.000.html

Cassini had a slightly more powerful propulsion system - "Main (445 Newton) engine for propulsive maneuvers" and 16 monopropellant hydrazine thrusters of which eight were prime and eight were backups. https://solarsystem.nasa.gov/missions/cassini/engine/

Cassini was launched from Earth on 15 October 1997, "followed by gravity assist flybys of Venus (26 April 1998 and 21 June 1999), Earth (18 August 1999), and Jupiter (30 December 2000). Saturn arrival was on 1 July 2004."
https://solarsystem.nasa.gov/resources/11776/cassini-trajectory/

Getting a crewed spacecraft of many tons in short time, e.g., several weeks to a month, would require a substantial power plant and propulsion system. What would scaling from 2 N to 2 kN look like? From 250 kW to 250 MW?

The ESA paper has one comment, "Taken to the extreme, a human Mars mission using EP would require about 8-15 MW power and deliver a thrust of 90-180 N." That's more modest than 250 MW. On the other, the thust is fairly low 0.09-0.18 kN, and no basis is given for those thrust requirements. Taking the maximum values of 15 MWe and 180 N, that gives 12 N/MW, which is greater than the 8.79 N/MW achieved by the prototype DS4G. One then has to devise a nuclear power plant to provide up to 15 MWe, which might be a 50 MWt plant if an efficiency is 0.3.

The Kiwi NTR was tested at full power, 1 GW, briefly with a thrust of about 74,000 pounds, or 334 kN.

One can find some references on rocket vehicles here:
https://www.sciencedirect.com/topics/physics-and-astronomy/rocket-vehicles
Any info yet on why the significant delays on DS4G since passing of the leader?
 
  • #31
darkdave3000 said:
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.
 
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  • #32
Astronuc said:
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?
 
  • #33
Astronuc said:
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.
 
  • #34
darkdave3000 said:
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.
darkdave3000 said:
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 Isp, so uses high mass flow rate to achieve high thrust. The propellant flow rate must be such to cool the reactor fuel.

darkdave3000 said:
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 Isp 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
 
  • #35
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.
 

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