Author Topic: Maximum temperature Solar Thermal Rocket  (Read 1412 times)

Offline matterbeam

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Maximum temperature Solar Thermal Rocket
« on: 10/26/2017 10:43 AM »
Hello.
I have been looking at ways to allow a Solar Thermal Rocket to achieve the maximum temperature possible using concentrated sunlight (5800K) and unlock its full 12km/s exhaust velocity potential.

Here is the blog post: http://toughsf.blogspot.com/2017/10/liquid-rhenium-solar-thermal-rocket.html

Here is the description of the design I am attaching:
The diagram is for illustrative purposes only - a functional schematic would be more detailed. Here is an explanation for each component:

Solar collector: A very large, very lightweight reflective film based on solar sails that can collect sunlight and focus it through a series of lens onto the heat exchanger fluid's inner surface.

Rotating drum: The drum's inner surface contains a liquid heat exchanger. The outer surface is actively cooled. The drum is dotted with tiny channels that allow the propellant to enter the liquid from the bottom and bubble through to the top. It is made of Tantalum-Hafnium Carbide.

Fluid surface: The fluid here is liquid rhenium. Its surface is heated to 5800K by concentrated sunlight. The lower layers nearer the drum holding the fluid is cooler. The centripetal forces hold the fluid in place

Pressure chamber: The rotating gas mix gets separated here. Dense rhenium vapours fall back down, hot hydrogen escapes.

Bubble-through heating: The rotation induces artificial gravity, allowing the hydrogen to heat up and rise through the denser rhenium. As it rises, it reaches hotter layers of the fluid heat exchanger. At the surface, it has reached 5800K. Small bubbles in direct contact with the rhenium allows for optimal thermal conductivity. More detail below.

Active cooling loop: liquid hydrogen from the propellant tanks makes a first pass through the drum walls, lowering the temperature below the melting point of THC. It emerges as hot, high pressure gaseous hydrogen.

High pressure loop: The heated hydrogen is forced through the channels in the drum. It emerges into the fluid heat exchanger as a series of tiny bubbles.

And for the performance:
We will calculate the performance of two versions of the RD-FHE STR. The first version uses modern materials and technologies, such as a 7g/m^2 Mylar sheet to collect sunlight and a 167kW/kg engine power density. The second version is more advanced, using 0.1g/m^2 sunlight collectors and a 1MW/kg power density.

Modern RD-FHE
5 ton collection area => 714285m^2
927MW of sunlight focused onto the drum.
5.56 ton propulsion system
Exhaust velocity: 12km/s
Thrust: 123.4kN (80% efficiency)
Thrust-to-weight ratio: 1.19
Overall power density: 87kW/kg

Advanced RD-FHE
5 ton collection area =>50000000m^2
64.9GW of sunlight received
64.9 ton propulsion system
Exhaust velocity: 12km/s
Thrust: 10.8MN
Thrust-to-weight ratio: 15.75
Overall power density: 928kW/kg

The principal argument against solar thermal rockets, that their TWR is too low and their acceleration would take too long to justify the increase in Isp, can be beaten by using very high temperatures and very low mass sunlight collectors.

For example, a 50 ton propulsion system based on the modern RD-FHE STR design, would be able to push 100 ton payloads to Mars (6km/s mission deltaV) using only 97 tons of propellant. It would leave Earth orbit at a decent 0.24g of acceleration, averaging 0.32g. The departure burn would take only 20 minutes. Using the advanced version of the RD-FHE solar thermal rocket would allow for a positively impressive acceleration of 3.1g.
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Offline Athrithalix

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Re: Maximum temperature Solar Thermal Rocket
« Reply #1 on: 10/27/2017 09:13 AM »
In your blog post, you talk about the conversion of 'sunlight to heat, and heat to kinetic energy' this is an odd mixture of measures that doesn't really make sense. A more accurate way to talk about this would be transferring the kinetic energy of photons to the exhaust mass.
Your blog also appears to trail off mid sentence when discussing seeding the hydrogen propellant.

It's not really clear from your post what sort of questions you're asking, so I've just said what jumps into mind.

Offline matterbeam

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Re: Maximum temperature Solar Thermal Rocket
« Reply #2 on: 10/27/2017 11:45 AM »
Thanks. I closed that sentence on the hydrogen seeding. I was intending to calculate the molar mass gain from seeding, but I did not manage to find any information on how much seed material was required. My only reference was a Gas-Core Nuclear Reactor that seeded some percent of the hydrogen propellant's density in tungsten, which worked out to the molar mass dropping from 1g/mol to 1.04g/mol...

Sunlight - W, heat -W, kinetic energy -W*s. It is not a proper equivalence of measures, but I was trying to shorthand this sentence: converting the collected sunlight into heat energy in the heat exchanger, which is then transferred to the hydrogen propellant, which through its acceleration and ejection through the nozzle, allows the spacecraft to gain kinetic energy.

The post isn't really asking any questions. I was putting some light on a propulsion system that is often overlooked when compared to nuclear, solar-electric and chemical-fuel options. By showing that it could achieve better exhaust velocity than solid-core nuclear thermal rockets, better thrust than solar-electric propulsion and with a simpler ISRU cycle than chemical fuels, I might even the scales.
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Offline Asteroza

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Re: Maximum temperature Solar Thermal Rocket
« Reply #3 on: 10/29/2017 10:39 PM »
You are going to need to do bit more drawing showing the mirror arrangement and the drum containment, perhaps importing some of the design of that laser OTV in your blog post. Holding down the leakage from the top of the drum is not a trivial issue. Then you have the interface issues between the drum and the nozzle proper.

There may be some usable mechanical tricks to deal with the drum, such as using low temperature gaseous H2 from other cooling operations to use as a gas bearing and spinup mechanism for the drum (rotational and thrust bearing).

I wonder if a turboinductor approach (from the NTER designs) might have more applicability, as upping the heat there to plasma conditions allows a magnetic nozzle.

Offline john smith 19

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Re: Maximum temperature Solar Thermal Rocket
« Reply #4 on: 10/30/2017 08:00 AM »
Hello.

Exhaust velocity: 12km/s

Thrust-to-weight ratio: 15.75
Overall power density: 928kW/kg
Exhaust velocity of 12Km/sec gives Isp of 1223secs. Much better than solid core NTR and (by Ion thruster standards) also high acceleration.

But
The mechanical and material issues of this design are very serious. Physically the structure (loosely) recalls the water boiler on the Shuttle, which provided cooling above 250Kft by spraying water on pipes inside a cylinder, creating a layer of boiling water (boiling around 40-65c due to near vacuum IIRC)

OTOH you're looking at a "transpiration" system (from the walls) through a layer of molten Rhenium, all of which is spinning. The walls are (AIUI) Thorium Carbide, a brittle ceramic which will be slightly radioactive and one of the candidates for use in UHT leading edges (whose major issue has been how to make larger pieces without internal flaws so they don't fail in use). 

AFAIK there is no knowledge base for doing any of this.  :(

[EDIT. There is another "sort of" relevant system. One of the National Labs (Oak Ridge?) worked on a Uranium enrichment scheme using a centrifuge like the kind in hospitals, with 2 (or any even number for balance) crucibles spun at high speed and the lighter enriched U235 extracted from the top surface. IIRC they dissolved the Uranium in Bismuth to lower the operating temperature so the crucible could retain more of its "hot strength." AFAIK it was abandoned because engineering a system to handle a dense, molten, chemically aggressive radioactive alloy moving at high speed, with acceptable safety levels in the event of failure, proved to be unexpectedly challenging. ]

The nearest I can think of are vacuum furnace crucibles made of reinforced carbon carbon.

So drill holes in the walls of one, put it inside another that's a close fit,with some separators between, seal the top edge and you've got a start.

IRL what I've just described is close to impossible.  :(

But what really kills you is that T/W ratio. It might be good by NTR standards, but it's beaten by every other high thrust chemical propellant. a 3:1 TWR means 1/3 of your whole vehicle is engine and fuel.  :(
[EDIT. Oops. You have a T/W ratio of the advanced system > 15:1 IE 6.6%. That's still poor for an engine. SSME was 59:1 in vacuum, and that was considered poor by engine standards (but very impressive by LH2 engine standards). ]
« Last Edit: 10/30/2017 09:44 AM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline matterbeam

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Re: Maximum temperature Solar Thermal Rocket
« Reply #5 on: 10/30/2017 02:15 PM »
You are going to need to do bit more drawing showing the mirror arrangement and the drum containment, perhaps importing some of the design of that laser OTV in your blog post. Holding down the leakage from the top of the drum is not a trivial issue. Then you have the interface issues between the drum and the nozzle proper.

There may be some usable mechanical tricks to deal with the drum, such as using low temperature gaseous H2 from other cooling operations to use as a gas bearing and spinup mechanism for the drum (rotational and thrust bearing).

I wonder if a turboinductor approach (from the NTER designs) might have more applicability, as upping the heat there to plasma conditions allows a magnetic nozzle.

I can only go into as much detail as I am able to find information about or as I know of. In this case, I do not know of any propulsion system that uses a rotating fluid bed as a heat exchanger, or of any mechanism designed to hold a fluid at a temperature higher than the melting temperature of the container.

One of the solutions for the various high-temperature-moving-component issues I mentioned in the comments of the blog post was to use a very thick drum. The hot inner surface is far removed from the outer surface. You'd have a thick cross-section to pump a lot of liquid hydrogen through. The outer walls of the drum might even be lowered to room temperature if it weren't very thermodynamically inefficient to do so. Either way, you'd have a relatively cool surface to mount wheels, bearings, insulation, suspension ect. onto.

The bottom of the drum can be shaped to form the top part of the nozzle if it ends up being a non-trivial challenge to mate a rotating pressure chamber to the top of a nozzle. The expansion skirt can extend from this top section once the propellant gasses have expanded and cooled to the point where lubrifiants between the top part, the sealant and the bottom part can operate without boiling off. Or maybe just make the drum and nozzle one continuous piece and just have a small segment of the drum's inner surface rotate like an inverted impeller and hope that spinning up the deepest part of the fluid is enough to spin the spin up the surface of the fluid.

Plasma temperatures for hydrogen are at least 10000K, which is nearly twice the maximum temperature achievable with concentrated sunlight.
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Offline matterbeam

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Re: Maximum temperature Solar Thermal Rocket
« Reply #6 on: 10/30/2017 02:23 PM »
Exhaust velocity of 12Km/sec gives Isp of 1223secs. Much better than solid core NTR and (by Ion thruster standards) also high acceleration.
But
The mechanical and material issues of this design are very serious. Physically the structure (loosely) recalls the water boiler on the Shuttle, which provided cooling above 250Kft by spraying water on pipes inside a cylinder, creating a layer of boiling water (boiling around 40-65c due to near vacuum IIRC)
OTOH you're looking at a "transpiration" system (from the walls) through a layer of molten Rhenium, all of which is spinning. The walls are (AIUI) Thorium Carbide, a brittle ceramic which will be slightly radioactive and one of the candidates for use in UHT leading edges (whose major issue has been how to make larger pieces without internal flaws so they don't fail in use). 
AFAIK there is no knowledge base for doing any of this.  :(
[EDIT. There is another "sort of" relevant system. One of the National Labs (Oak Ridge?) worked on a Uranium enrichment scheme using a centrifuge like the kind in hospitals, with 2 (or any even number for balance) crucibles spun at high speed and the lighter enriched U235 extracted from the top surface. IIRC they dissolved the Uranium in Bismuth to lower the operating temperature so the crucible could retain more of its "hot strength." AFAIK it was abandoned because engineering a system to handle a dense, molten, chemically aggressive radioactive alloy moving at high speed, with acceptable safety levels in the event of failure, proved to be unexpectedly challenging. ]
The nearest I can think of are vacuum furnace crucibles made of reinforced carbon carbon.
So drill holes in the walls of one, put it inside another that's a close fit,with some separators between, seal the top edge and you've got a start.

IRL what I've just described is close to impossible.  :(
But what really kills you is that T/W ratio. It might be good by NTR standards, but it's beaten by every other high thrust chemical propellant. a 3:1 TWR means 1/3 of your whole vehicle is engine and fuel.  :(
[EDIT. Oops. You have a T/W ratio of the advanced system > 15:1 IE 6.6%. That's still poor for an engine. SSME was 59:1 in vacuum, and that was considered poor by engine standards (but very impressive by LH2 engine standards). ]

I agree with the engineering difficulties. It might be simpler to just spin up the fluid inside the drum wall instead of spinning up the entire drum.

The walls are THC, which I described as standing for Tantalum Hafnium Carbide, with the alternative being HNC, or Hafnium Nitrogen Carbide.

If using a high pressure 'transpiration' mechanism to inject the hydrogen and have it bubble through the fluid is difficult to make work, one of the alternatives I suggested at the end of the blog post was to pump hydrogen from the top of the chamber and have to absorb heat by simply being in contact with the heat exchanger's surface. An elongated pressure chamber would be needed to increase the hydrogen/rhenium contact area and ensure that most of the heat is absorbed by the propellant. The downside would be that the hydrogen would reach a velocity of a dozen km/s at the bottom of the tube and would start dragging along the rhenium like wind over a sea creating waves - a lot of fluid could be lost because of this.
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Offline john smith 19

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Re: Maximum temperature Solar Thermal Rocket
« Reply #7 on: 10/30/2017 05:58 PM »
I agree with the engineering difficulties. It might be simpler to just spin up the fluid inside the drum wall instead of spinning up the entire drum.
If the fluid is H2, either liquid or gas, you'd be right. If it's a molten metal it's more doubtful.
Quote from: matterbeam
The walls are THC, which I described as standing for Tantalum Hafnium Carbide, with the alternative being HNC, or Hafnium Nitrogen Carbide.
I'm used to chemical names, so I read that as ThC. What you've described is TaHfC. It's like Nichrome being an actual alloy of NiCr or Fecralloy being FeCrAl and (usually) Yttrium.
Quote from: matterbeam
If using a high pressure 'transpiration' mechanism to inject the hydrogen and have it bubble through the fluid is difficult to make work, one of the alternatives I suggested at the end of the blog post was to pump hydrogen from the top of the chamber and have to absorb heat by simply being in contact with the heat exchanger's surface. An elongated pressure chamber would be needed to increase the hydrogen/rhenium contact area and ensure that most of the heat is absorbed by the propellant. The downside would be that the hydrogen would reach a velocity of a dozen km/s at the bottom of the tube and would start dragging along the rhenium like wind over a sea creating waves - a lot of fluid could be lost because of this.
For spinning heat exchangers the "Dumbo" and "spinning pebble bed reactor" NTR reactor concepts come to mind, specifically the idea of heat transfer through pebbles immobolized by centripetal force.

I'm not going to read  your blog. The thrust level and acceleration is attractive (as is the lack of a nuclear reactor) but the T/W ratio is abysmal and the TRL is basically 0. If you can't improve those, and/or show where at least some of this tech has been tried out, you're going nowhere.

Several people here have pointed you in several useful directions you should consider investigating further (I'd also suggest "Non imaging optics" for high ratio light concentration in the > 1000:1 range).
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline matterbeam

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Re: Maximum temperature Solar Thermal Rocket
« Reply #8 on: 10/31/2017 11:17 PM »
If the fluid is H2, either liquid or gas, you'd be right. If it's a molten metal it's more doubtful.

For spinning heat exchangers the "Dumbo" and "spinning pebble bed reactor" NTR reactor concepts come to mind, specifically the idea of heat transfer through pebbles immobolized by centripetal force.

I'm not going to read  your blog. The thrust level and acceleration is attractive (as is the lack of a nuclear reactor) but the T/W ratio is abysmal and the TRL is basically 0. If you can't improve those, and/or show where at least some of this tech has been tried out, you're going nowhere.

Several people here have pointed you in several useful directions you should consider investigating further (I'd also suggest "Non imaging optics" for high ratio light concentration in the > 1000:1 range).

Why is molten metal hard to keep flowing?

An pebble-bed reactor is on the blog post, which is why I specified 'fluid' heat exchanger.

T/W ratio is not relevant in the way you think it is. An spacecraft using this propulsion system necessarily only travels in vacuum, from orbit to orbit. It needs enough acceleration to perform departure and insertion burns within reasonable timescales if humans are part of the payload, or within half an orbit if it wants to benefit from the Oberth effect.

For Hohmann transfers, such as Earth to Mars, the departure burn is on the order of 3.5km/s. If you want to spend less than 1 day leaving Earth, you need to have a minimum acceleration of 4.1 milligees. If you want to spend only an hour leaving Earth, then the minimum acceleration rises to 99 milligees. The modern STR described in the blog's 'worked examples' section has a TWR of 1.19. If the propulsion system represents 100% of the spacecraft's dry mass, it would have an acceleration of 1.19G. If it represented 25% of the dry mass, the acceleration is  0.3G. We can afford to have the propulsion system represent only 8.3% of the spacecraft's dry mass before the departure burn has to be split into several orbits.

Unlike a liftoff from Earth's surface, there is no 'minimum' TWR of 1G.

The TRL is zero only on NASA's scale, but it is not a useful argument as just about every single thread on 'Advanced Concepts' won't break out of TRL 0-1.
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Offline john smith 19

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Re: Maximum temperature Solar Thermal Rocket
« Reply #9 on: 11/01/2017 08:04 PM »
Why is molten metal hard to keep flowing?
Depends. You have a very high Mp, dense metal. You have problems with either inertia or heat loss. Thin layers have to be kept hot continually. Thicker layers needs a lot of energy to spin them up. BTW this will also have substantial angular momentum (like a Control Moment Gyro) so changing the thrust vector will be another issue.
Quote from: matterbeam
An pebble-bed reactor is on the blog post, which is why I specified 'fluid' heat exchanger.
Which is mentioned on several threads on this site. It did not work out well. As a BOTE I'd expect you'd need to exceed the force with which the molten metal layer is being pressed to the wall, so lower speed is better.
Quote from: matterbeam
T/W ratio is not relevant in the way you think it is. An spacecraft using this propulsion system necessarily only travels in vacuum, from orbit to orbit. It needs enough acceleration to perform departure and insertion burns within reasonable timescales if humans are part of the payload, or within half an orbit if it wants to benefit from the Oberth effect.

For Hohmann transfers, such as Earth to Mars, the departure burn is on the order of 3.5km/s. If you want to spend less than 1 day leaving Earth, you need to have a minimum acceleration of 4.1 milligees. If you want to spend only an hour leaving Earth, then the minimum acceleration rises to 99 milligees. The modern STR described in the blog's 'worked examples' section has a TWR of 1.19. If the propulsion system represents 100% of the spacecraft's dry mass, it would have an acceleration of 1.19G. If it represented 25% of the dry mass, the acceleration is  0.3G. We can afford to have the propulsion system represent only 8.3% of the spacecraft's dry mass before the departure burn has to be split into several orbits.

Unlike a liftoff from Earth's surface, there is no 'minimum' TWR of 1G.
TWR does matter because because your designs TWR is not impressive and if the tugs acceleration drops far enough (because of the payload it's pushing) it's no longer in chemical or NTR thrust territory, it's in ion drive, and they can do better already. [EDIT. It's high thrust certainly helps quite a bit though. The question then becomes will it run out of fuel before it completes the burn
BTW it's little "g" (acceleration due to gravity. G is the Universal Gravitational Constant).
Quote from: matterbeam
The TRL is zero only on NASA's scale, but it is not a useful argument as just about every single thread on 'Advanced Concepts' won't break out of TRL 0-1.
It's been a while since Reaction Engines SABRE/Skylon was transferred over to "Commercial."

That said there are few high thrust short duration burn systems. An interesting question would be if it can push a payload to very close to escape while still being in a very elliptical orbit, so returning to LEO when done.

[EDIT your reaction chamber also has a surface temperature above the melting point of Tungsten.

There isn't any realistic way to attach a nozzle to this chamber. ]
« Last Edit: 11/01/2017 10:07 PM by john smith 19 »
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Offline matterbeam

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Re: Maximum temperature Solar Thermal Rocket
« Reply #10 on: 11/02/2017 11:47 AM »
Depends. You have a very high Mp, dense metal. You have problems with either inertia or heat loss. Thin layers have to be kept hot continually. Thicker layers needs a lot of energy to spin them up. BTW this will also have substantial angular momentum (like a Control Moment Gyro) so changing the thrust vector will be another issue.

The need for the top layer of the rhenium to be much hotter (+1000K)  than the bottom layer favours a very thick fluid heat exchanger depth.
The good acceleration means that the spacecraft only needs to spin up the heat exchanger fluid for relatively short periods of time and point in a single direction during burns.

Quote
Which is mentioned on several threads on this site. It did not work out well. As a BOTE I'd expect you'd need to exceed the force with which the molten metal layer is being pressed to the wall, so lower speed is better.

I don't understand what you mean. What is BOTE? What needs to exceed the centrifugal force? If you mean pushing the propellant through the fluid, then it is no problem at all. Hydrogen is massively less dense than rhenium, so it will have a significant buoyancy in the fluid heat exchanger. Once near the surface of the fluid, the temperature is too great for it to be dissolved  and the pressure too low to keep the hydrogen from expanding out and down to the nozzle.

Quote
TWR does matter because because your designs TWR is not impressive and if the tugs acceleration drops far enough (because of the payload it's pushing) it's no longer in chemical or NTR thrust territory, it's in ion drive, and they can do better already. [EDIT. It's high thrust certainly helps quite a bit though. The question then becomes will it run out of fuel before it completes the burn
BTW it's little "g" (acceleration due to gravity. G is the Universal Gravitational Constant).

I think you overestimate the importance of g-level TWR. If the propulsion system represents 8.3% of the entire spaceship's initial mass, you can complete a burn headed to Mars in under an hour!  If the spaceship is 100 tons, a trip to mars might require 40 tons of propellant, 8.3 tons of STR propulsion, and the rest of the 51.7 tons are just payload and other modules.

Electric rockets so far have a power density of under 1kW/kg. Paired with an exhaust velocity of 2000s or more, they produce a thrust of only 0.1N/kg. That's a TWR of 0.01. A 100 ton spaceship headed for Mars would need to dedicate only 26.3 tons to propellant, but if it dedicated 8.3 tons to propulsion, it would have an acceleration of only 0.85 milligees. The Mars departure would take a full 4.88 days of burning in the sunlight and in the right direction, so in total it would take more than two weeks as the orbit gets elongated and the period increases after each split burn...

All this to save on some propellant?

Quote
That said there are few high thrust short duration burn systems. An interesting question would be if it can push a payload to very close to escape while still being in a very elliptical orbit, so returning to LEO when done.

That's just a question of keeping enough deltaV in reserve after the payload is released. Any rocket can do that, there is nothing notable about the STR in that regard.

Quote
[EDIT your reaction chamber also has a surface temperature above the melting point of Tungsten.

There isn't any realistic way to attach a nozzle to this chamber. ]

The reaction chamber has a 5800K surface on the fluid heat exchanger, a 4000K drum inner surface, an arbitrarily cool drum outer surface and arbitrarily cool nozzle. Tungsten is not relevant as was not considered for any of the components due to its melting point.

If an impeller can spin the rhenium over the drum's inner basin without having to spin the whole drum, then the connection between reaction chamber bottom and nozzle opening is a non-issue.
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Offline john smith 19

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Re: Maximum temperature Solar Thermal Rocket
« Reply #11 on: 11/02/2017 06:34 PM »
I don't understand what you mean. What is BOTE? What needs to exceed the centrifugal force? If you mean pushing the propellant through the fluid, then it is no problem at all. Hydrogen is massively less dense than rhenium, so it will have a significant buoyancy in the fluid heat exchanger. Once near the surface of the fluid, the temperature is too great for it to be dissolved  and the pressure too low to keep the hydrogen from expanding out and down to the nozzle.
BOTE is Back Of The Envelope AKA a "Fermi Approximation." In this case for a gas bubble to come through that metal layer it has to create a bubble from that porus wall whos internal pressure exceeds the force excerted on it.. That means it has to exceed the pressure at the bottom of the layer of molten metal. Hence the lowest rotation rate possible is better.
Quote from: matterbeam
I think you overestimate the importance of g-level TWR. If the propulsion system represents 8.3% of the entire spaceship's initial mass, you can complete a burn headed to Mars in under an hour!  If the spaceship is 100 tons, a trip to mars might require 40 tons of propellant, 8.3 tons of STR propulsion, and the rest of the 51.7 tons are just payload and other modules.

Electric rockets so far have a power density of under 1kW/kg. Paired with an exhaust velocity of 2000s or more, they produce a thrust of only 0.1N/kg. That's a TWR of 0.01. A 100 ton spaceship headed for Mars would need to dedicate only 26.3 tons to propellant, but if it dedicated 8.3 tons to propulsion, it would have an acceleration of only 0.85 milligees. The Mars departure would take a full 4.88 days of burning in the sunlight and in the right direction, so in total it would take more than two weeks as the orbit gets elongated and the period increases after each split burn...

All this to save on some propellant?

Quote
That said there are few high thrust short duration burn systems. An interesting question would be if it can push a payload to very close to escape while still being in a very elliptical orbit, so returning to LEO when done.

That's just a question of keeping enough deltaV in reserve after the payload is released. Any rocket can do that, there is nothing notable about the STR in that regard.

Quote
[EDIT your reaction chamber also has a surface temperature above the melting point of Tungsten.

There isn't any realistic way to attach a nozzle to this chamber. ]

The reaction chamber has a 5800K surface on the fluid heat exchanger, a 4000K drum inner surface, an arbitrarily cool drum outer surface and arbitrarily cool nozzle. Tungsten is not relevant as was not considered for any of the components due to its melting point.

If an impeller can spin the rhenium over the drum's inner basin without having to spin the whole drum, then the connection between reaction chamber bottom and nozzle opening is a non-issue.
That's an impeller that won't melt while spinning liquid Rhenium.

I'll leave other to comment on the viability of that concept.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline stefan r

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Re: Maximum temperature Solar Thermal Rocket
« Reply #12 on: 11/09/2017 01:53 AM »
In school (both high school and  general chem college) they taught us about 3 states of matter.  Solid, liquid gas.  Then they add plasma and critical fluids.  I would do the same if I had to teach a chemistry class.  Good profs/textbooks mention that other things like foams, aerosols, emulsions, gels and aerogels exist.  They usually leave out "froth" and "spray". 

Experiments with rhenium are expensive.  Carbon dioxide and water are much cheaper (and can be fun); gas from bubbles should vent out of a beer bottle.  The only water loss should be H2O molecules that evaporate and dissolve in the gas phase.  However, if you shake the beer before opening the top, the CO2 moves into the gas phase faster.  The fluid that comes out of the opening is not liquid or gas, it is both.  It is technically not "foam" because foam is not in equilibrium.  Froth looks a lot like foam and if you can freeze or stabilize a froth then it becomes a foam.  If you set up the beer experiment right you can also demonstrate "spray".  Sprays are a lot like aerosols but, like froth, sprays are not in equilibrium. 

Even if you avoid making a froth you have another problem.  Most of the energy going into the rocket is adsorbed by the surface of the rhenium puddle.  Heat transfer depends on the thermal conductivity of liquid rhenium.  Metals have high thermal conductivity so that might be fine.  Your blog looks like > 105 solar which might be a bit much.   When a bubble approaches the surface the heat transfer depends on the thermal conductivity of hydrogen gas.  Gasses have very low thermal conductivity.  A thin film of rhenium will be exposed to full energy flux.  The film's temperature will rise rapidly.  You have the puddle temperature within 100 degrees of rhenium's boiling temperature.

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