Aero wrote:I would be interested in an evaluation of the conceptual design presented in the paper as applied to spacecraft power.
The authors propose a 60 MW (thermal?) cylindrical device ~2 m long and ~0.5 m diameter. Estimating the density at 5 metric tons per cubic meter gives a power source of about 2 tons mass. Couple that to a VASIMR engine about 70% efficient with an exhaust velocity, Ve, of 50 km/sec. ending up with about 60 MW * 0.33 * 0.7 ~= 14MW exhaust energy. (Assuming thermal to electric conversion at 33%) Now using energy = 1/2 m Ve^2 and thrust = m Ve, I would solve for exhaust mass then thrust if I were confident of the correct units to use. Once I know thrust, space vehicle acceleration is easy for any speculated vehicle mass.
Yes, Field Reverse Configuration can be used as space thruster. And claimed 50 or 250km/s is very attractive for this purpose if you have the corresponding power source.
The problem is how to produce required power.
Aero wrote:I would be interested in an evaluation of the conceptual design presented in the paper as applied to spacecraft power...
You've omitted the mass of the power conversion gear and the cooling gear.
Steam turbines, condensers and radiators and all the associated plumbing, pumps etc.
DPFs have an advantage here in that a rogowski coil around the exhaust can recover enough energy to recharge the capacitors while still allowing usable thrust.
A Helion-based drive must be externally powered.
A Helion-based (or any other neutronic fusion) power supply must have considerable extra mass dedicated to power conversion and cooling.
Aero wrote:I would be interested in an evaluation of the conceptual design presented in the paper as applied to spacecraft power...
You've omitted the mass of the power conversion gear and the cooling gear.
Steam turbines, condensers and radiators and all the associated plumbing, pumps etc.
DPFs have an advantage here in that a rogowski coil around the exhaust can recover enough energy to recharge the capacitors while still allowing usable thrust.
A Helion-based drive must be externally powered.
A Helion-based (or any other neutronic fusion) power supply must have considerable extra mass dedicated to power conversion and cooling.
I am not sure but think that not whole Helion but only its part: namely Inductive Plasma Accelerator is more attractive as space thruster.
As Inductive Plasma Accelerator produces velocity 250km/s vs. plasma focus having as I know zero velocity.
May be I am mistaken.
Yes, that should be externally powered.
Joseph Chikva wrote:
As Inductive Plasma Accelerator produces velocity 250km/s vs. plasma focus having as I know zero velocity.
May be I am mistaken.
Yes, that should be externally powered.
Nope. A DPF pinch can produce a beam of ions along the long axis of the anode.
It's a product of the extreme magnetic fields induced during the pinch. The composition of the beam depends on the fuel, and with boron 11 the beam is (mostly) alpha particles.
Joseph Chikva wrote:
As Inductive Plasma Accelerator produces velocity 250km/s vs. plasma focus having as I know zero velocity.
May be I am mistaken.
Yes, that should be externally powered.
Nope. A DPF pinch can produce a beam of ions along the long axis of the anode.
It's a product of the extreme magnetic fields induced during the pinch. The composition of the beam depends on the fuel, and with boron 11 the beam is (mostly) alpha particles.
This is regarding my earlier post on space vehicle power using VASIMR engines.
Yes, there is much more mass on a space vehicle than simply the prime mover, but that can all be estimated.
I did think of a simple way to estimate thrust from 14 MW exhaust energy from VASIMR engines. From Wikipedia:
it was expected that the VX-200 engine would have a system efficiency of 60-65 % and thrust level of 5 N.
200 kW in the VX-200 engine gives 5 N thrust, so by ratio, 14 MW would give 350 N. That won't get off the ground but it is enough thrust to move a lot of mass once in space. Of course heat radiators can be massive and there would be a lot of waste heat from this configuration, something like 46 MW. Still, such configuration might find a use.
Aero wrote:200 kW in the VX-200 engine gives 5 N thrust, so by ratio, 14 MW would give 350 N.
I think that is not quite correct to use such ratio.
As thrust depends on exhaust velocity but spent energy on square.
But if you would have a nuke power source I can spend more energy for spending less mass of working body.
PS: May be your terminology differs from mine as I am translating from my language.
Aero wrote:200 kW in the VX-200 engine gives 5 N thrust, so by ratio, 14 MW would give 350 N.
I think that is not quite correct to use such ratio.
As thrust depends on exhaust velocity but spent energy on square.
But if you would have a nuke power source I can spend more energy for spending less mass of working body.
PS: May be your terminology differs from mine as I am translating from my language.
Yes, with more energy you might find a different design operating point. Using ratios is fair only when all else is held constant and in this case that is what I have done. It is a starting point, nothing more. I agree that it is probably not practical to use 70 VASIMR engines on a single space vehicle, but if that were the case, then ratios would give absolutely the correct answer.
Note that I was only looking for thrust, not any of the myriad other critical parameters that must be considered for any space vehicle application. But, 350 N thrust for single digit ton mass might encourage one to look more closely.
Couple that to a VASIMR engine about 70% efficient with an exhaust velocity, Ve, of 50 km/sec. ending up with about 60 MW * 0.33 * 0.7 ~= 14MW exhaust energy. (Assuming thermal to electric conversion at 33%) Now using energy = 1/2 m Ve^2 and thrust = m Ve, I would solve for exhaust mass then thrust if I were confident of the correct units to use. Once I know thrust, space vehicle acceleration is easy for any speculated vehicle mass.
Why not directly eject the plasma from the reactor and save yourself the detour to VASIMIR?
And if that is no good, then directly use the heat to heat hydrogen and make a NERVA type engine?
The biggest problem with electrical propulsion from nuclear reactions that produce heat, is that in space, cooling is a huge problem. You can only radiate heat of, unless you have some sort of cooling agent that is released into space.
A NERVA engine expells the cooling as reaction mass and so does not require extra radiative cooling panels (or whatever the cooling solution looks like in the end).
Couple that to a VASIMR engine about 70% efficient with an exhaust velocity, Ve, of 50 km/sec. ending up with about 60 MW * 0.33 * 0.7 ~= 14MW exhaust energy. (Assuming thermal to electric conversion at 33%) Now using energy = 1/2 m Ve^2 and thrust = m Ve, I would solve for exhaust mass then thrust if I were confident of the correct units to use. Once I know thrust, space vehicle acceleration is easy for any speculated vehicle mass.
Why not directly eject the plasma from the reactor and save yourself the detour to VASIMIR?
And if that is no good, then directly use the heat to heat hydrogen and make a NERVA type engine?
The biggest problem with electrical propulsion from nuclear reactions that produce heat, is that in space, cooling is a huge problem. You can only radiate heat of, unless you have some sort of cooling agent that is released into space.
A NERVA engine expells the cooling as reaction mass and so does not require extra radiative cooling panels (or whatever the cooling solution looks like in the end).
I think that today's level of technology does not allow making jet engines with high specific momentums and at the same time high specific thrust.
I only said that if we would have compact and light enough high power source in the future (heat or electric), only in that case we will able to consider possibility of building engines with high exhaust velocity.
Not today. Any design.
Stefank wrote:That is an interesting paper, looks like Tri-Alpha has built a duplicate of the Helion device, pretty far removed from their published patents..
Skipjack wrote:And if that is no good, then directly use the heat to heat hydrogen and make a NERVA type engine?
Because there's a maximum temperature you can plausibly heat hydrogen to without falling back on your "directly eject plasma from the reactor" idea. Maximum Isp is limited by this value, and for solid heat exchangers (tungsten) it's ~1000 s, which is far too small for the applications you'd use a VASIMR for. Even a nuclear lightbulb only doubles that at best, and it's not clear how you'd translate the nuclear lightbulb propellant heating concept to a fusion reactor...
If you want the Isp up, you want either CSR or DFP.
Because there's a maximum temperature you can plausibly heat hydrogen to without falling back on your "directly eject plasma from the reactor" idea. Maximum Isp is limited by this value, and for solid heat exchangers (tungsten) it's ~1000 s, which is far too small for the applications you'd use a VASIMR for. Even a nuclear lightbulb only doubles that at best, and it's not clear how you'd translate the nuclear lightbulb propellant heating concept to a fusion reactor...
If you want the Isp up, you want either CSR or DFP.
You do understand that in order to create electricity from the heat, you need to have huge cooling "sails" in space and that is going to severely limit your efficiency, especially, if you account for the additional losses in the heat- >electricity conversion process. So simply ejecting the plasma is the way to go. Slough has several really good designs for that too, btw.
I wasn't addressing that. I was explaining why your suggestion of a NERVA-type thermal engine was in a completely different class from the VASIMR or direct-exhaust drives, and thus not appropriate for the same mission types.
Direct-exhaust, as I understand it, tends to be very high-Isp (and thus low-thrust), such that its low end roughly corresponds with VASIMR's high end. For Polywell, anyway; other reactor types may have different characteristics, and of course if direct conversion isn't available there's a substantial advantage to direct-exhaust unless you really need a larger thrust-to-power ratio than it can give you. Depending on the DFP drive's minimum Isp, the 'dead zone' where CSR is worse than DFP or ARC might cover most or all of its range in the case where the fusion reactor needs a thermal plant.