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Rather than focusing on Isp; the use of Id is a far better number to look at for mass control of a spacecraft. The Id for a combination of LF2 with Hydrazine is nearly 3x that of LF2/LH2: 432 kg-sec/l. Another potential exotic oxidizer is FLOX, a 50-50 mixture of LOX and LF2; when used in combination with Aerozine 50, it yields Id of 403, and with UH 25 (25% UDMH/N2H4), it goes up to 411. These are all significantly better than RP-1 using FLOX, at 386 kg-sec/l. When we subsequently look at using Dinitrogen Tetroxide, the use of Aerozine 50 jumps off the page w/r to safety in handling over that of N2H4, but with a significant performance edge at 326 kg-sec/l.
Note: all values are at sea level, so vacuum performance should be significantly higher.
These exotic propellant schemes are pretty impressive on paper, but they aren't going to be allowed at Cape Canaveral if they produce oceans of hydrogen fluoride as exhaust products. For upper stages, on the Moon in vacuo--go for it.
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Oldfart1939,
I agree that Id is important for mass control of tankage and support structures. Solids have the best Id, but the Isp of solids is best described as poor when compared to most common liquid bi-propellants. I was pointing out that both Isp and Id should be very high using LF2/HAN bi-propellant or LOX/HAN bi-propellant. I never suggested LF2/HAN for propulsion to Earth orbit. I was stating that the Mars ascent stage (the mission architecture component that propels the astronauts back into Mars orbit from the surface of Mars), specifically, would benefit from the use of powerful oxidizers. The oxidizer could be manufactured on Mars (LOX) or it could provide superlative Isp and Id (LF2).
I was curious about the relative difficulty of storing LF2 in comparison to LOX. If the ascent stage uses HAN-only for landing, then the internal components of the AJ-10 derivative engine are not exposed to hot flowing Fluorine for more than the few minutes of ascent. Various rocket engines, to include the F-1's and RS-25's amongst others, use super alloys in their turbo machinery to resist hot high pressure flowing oxygen. These alloys are also in use where hot flowing fluorine gas is present, such as in the manufacture of fluorinated polymers and Uranium processing. It's expensive, but cost should not be a major issue for relatively small engines.
I have a paper that indicated through various tests that common nickel bearing alloys are suitable for LF2 tankage. The alloys in question were not significantly degraded over the course of those tests and the tensile strength of the materials after the test was nearly identical to a control article that held LN2. This was a result of the fluoride films that formed on the surfaces of the metals and little to no further reaction was noted after initial film formation. If the tankage was pre-treated with fluorine gas, then there'd be no reaction once the LF2 was loaded. Tests with LF2 filled tanks ruptured with blasting caps failed to ignite various nickel alloys. The LF2 tests were conducted at standard atmospheric pressure and the LF2 was stored at -320F. That'd require substantial cryocooler power to maintain that temperature. In any event, I only wanted to get an opinion from a chemical engineer on the feasibility of long term storage of pressurized LF2.
There are some practical reasons I can think of as to why F-based oxidizers is not as good solution a solution as LOX made on Mars.
* The deep cryogen storage problem begins when the LF2 is loaded and doesn't end until the astronauts leave Mars 3 years later.
* The ascent stage is still so heavy that it requires a different HIAD design.
* Engine turbo machinery would be trashed after firing, so no practical test fire or restart capability.
* A launch vehicle failure on the pad would contaminate the launch pad area.
* Construction of expensive new test facilities is required.
* It's probably a safe bet that NASA has little experience with Fluorine handling.
* Astronauts and equipment in the launch area on Mars require protection from the combustion products, as neither are F-compatible.
LOX is looking better and better all the time.
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Back in the 1950's and especially the 1960's, Rocketdyne did a lot of fluorine work at Santa Susannah, CA. They eventually gave up on it as a bad deal. Lots of people got hurt doing this. The poisonous dangers and almost insuperable corrosion problems were just not worth the theoretical advantages over oxygen.
That's not to say it could not be reinvestigated. But when guys that talented back then gave up on it, there's considerable practical resistance to trying again, even though the ones with the experience are all retired or dead now. Chemistry hasn't changed, although some of the handling technologies have.
That history is why F2 and OF2 compounds are listed among the oxidizer-fuel combinations in my ancient Pratt handbook. The best "theoretical" is actually ozone (O3) and hydrogen (if you think H2O2 is unstable ....). They were still being investigated back then (late 1960's).
GW
Last edited by GW Johnson (2017-03-26 09:26:14)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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rbd512-
I arrived at a similar conclusion in my post #66, although not necessarily stated as such. From the energetics versus vehicle mass, it seems that if we are selecting a propellant couple for a Stage 1 liftoff, a combination of Hydrazine and LOX has the overall "best performance," but stability consideration accounted for, either Aerozine 50 or MMH with LOX are the winners. Whether the performance advantage is significantly better than RP-1 is an economic decision based on fuels cost. In the case of Falcon 9, when there has to be every bit of fuel reserve utilized for stage recovery, the difference in fuel costing $600,000 (a real WAG for Aerozine or MMH) versus $200,000 for RP-1) which allow for more reliable vehicle recovery is a decision beyond my pay grade to make.
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I did some checking on prices for some of these other fuels, such as Aerozine-50, and the stuff isn't cheap. Of course the cost isn't based on the chemicals but the handling and transportation under EPA rules. Roughly $135 per kg is pretty stiff. Even worse for NTO, though. It's certainly NOT the production costs, but the regulatory burden is certainly oppressive. Fueling a Delta II second stage would cost roughly $800,000, at this rate. I'm sure the Russians are having none of these environmental impediments to interfere with their launches. Looking at the Aerojet-Rocketdyne engine, it's rated for up to 7 restarts, and is non-regeneratively cooled, using instead an ablative lining for the combustion chamber. This puts it in the "use and throw away" category of equipment.
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Oldfart1939,
Even with the eminently more affordable mission hardware architecture I've proposed, there's nothing about this mission that will be inexpensive. It just won't be so expensive and complicated that NASA can't possibly afford to do it within our lifetimes, given existing and projected budgets.
The storable fuel is a non-negotiable line item for initial exploration missions. There is no such technology, nor experience with said technology, for launching a spacecraft loaded with cryogenic propellants, flying them tens of millions of miles to another planet, and then having them sit for a couple years on that planet with zero boil-off of propellants prior to use. Simply making and storing LOX in the oxidizer tank of the ascent stage is pushing our present technological capabilities to the limit.
Mission #1:
* Demonstrate production and storage of LOX from Martian CO2 using MOXIE and solid state cryocoolers.
* Experiment with collection and purification of the briny Martian ground water.
* Demonstrate the ability to grow a few common fruits, vegetables, cotton, and hemp on Mars.
Mission #2:
* Demonstrate ice mining on a scale suitable for use in a propellant plant and in agriculture.
* Demonstrate water electrolysis and storage of LOX and LH2.
Mission #3:
* Demonstrate LOX/LCH4 production in a propellant plant and fueling of a LOX/LCH4 powered ascent vehicle. Demonstrate production of Martian concrete.
Mission #4:
* Demonstrate a nuclear fission surface power system.
* Demonstrate production of Martian steel.
* Demonstrate construction techniques using steel-reinforced Martian concrete.
Mission #5:
* Demonstrate mining equipment for production of metals required for aerospace alloys.
* Demonstrate fabrication of simple tools and spare parts using 3D printing.
* Demonstrate fabrication of microchips.
That's as ambitious an ISRU demonstration plan as I'm willing to entertain. I think it's extraordinarily ambitious at that. Each successive mission is intended as a logical progression of technology demonstrations to confirm that the general capabilities required for establishment of a permanent human presence on Mars are executed by NASA so that corporations can then refine those technologies to make them cost effective and suitable for use in Martian colonies.
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Not going to disagree with you, but the demonstration of a Mars Nuclear Fission powerplant needs to be in mission 1 architecture, or the rest of the technology is a "no-go." Power is the all-important component of Martian colonization.
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Design the mission to use a common nuclear power source for all other demostrators to attach to making them with additional rovers to allow for cabling to be moved such that they can recieve the power. Have them all work to produce man's first insitu built base. These dedicated mission will be easily still useable for mans first missions with that many cycles for all of the missions....
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The fission surface power system demonstrator should be a 1MW unit, which means some development of SAFE-400 technology is required to achieve that output level. A 400kWe unit requires no significant modifications to the existing SAFE-400 core technology, but a 1MW unit would be far more useful for a starter outpost. There is obviously a requirement for a flight test, perhaps on the surface of the moon, to prove that the reactor functions as expected. I'd rather have a finalized design tested than a design that would require another flight test. SAFE-400 technology has never been tested in space, even though development of the ground demonstrator core unit was completed. It still requires completion of development of the radiator system, electrical power generator, and control unit. IIRC, the reactor was controlled manually. Apart from the heat exchange fluid that drives the electrical generator, the only moving parts in the core are the control drums.
NASA's tiny nuclear power and propulsion budget is better spent on development of fission power systems, rather than on re-creating the fuel elements from the NERVA program. We're never going to fly nuclear powered rockets, so there's no point to that development effort. If EMDrive merely functions as well as experiments thus far have demonstrated, then there's a requirement for a power source that provides reliable power anywhere in our solar system, irrespective of vehicle orientation or distance from the Sun.
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I agree that having a 1 MWe unit "would be nice," but we should fly something immediately available and not wait around for the "perfect unit," that simply delays implementation of any realistic program. My first mission would be a lot like the Zubrin Mars direct model. It would include a working COTS Nuclear reactor, of any power output beyond the minimum of the SAFE-400; a reasonable scale MOXIE unit, as well as a Sabatier reactor. Both reactors should include a scaled down cryo-liquifaction system and associated storage tanks. This demonstrates the feasibility of fuel/oxidizer ISPP. This is a lot more than some of the other landers have accomplished, since it will not get bogged down in a sand drift.
Last edited by Oldfart1939 (2017-03-28 15:15:39)
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Oldfart1939,
We do have 100kWe SAFE-400 core development completed. Even so, development of a 400kWe core should be completed. It's the exact same technology set as the 100kWe core, not the variant technology set required by the 1MWe core, and it's nearly identical in volume to the 100kWe core. The fuel element count is higher in the 400kWe cores because the fuel elements have smaller diameters. 1MWe represents a more significant development program and requires serious funding, whereas 400kWe does not.
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https://en.wikipedia.org/wiki/Safe_Affo … ion_Engine
Safe Affordable Fission Engine (SAFE) are NASA's small experimental nuclear fission reactors for electricity production in space. Most known is the SAFE-400 reactor producing 400 kW thermal power, giving 100 kW of electricity using a Brayton cycle gas turbine. The fuel is uranium nitride in a core of 381 pins clad with rhenium. Three fuel pins surround a molybdenum-sodium heatpipe that transports the heat to a heatpipe-gas heat exchanger. This is called a Heatpipe Power System. The reactor is about 50 centimetres (20 in) tall, 30 centimetres (12 in) across and weighs about 512 kilograms (1,129 lb). It was developed at the Los Alamos National Laboratory and the Marshall Space Flight Center under the lead of Dave Poston. A smaller reactor called SAFE-30 was made first.
The working fluid used in the reactor is a Helium Xenon gas mixture.
The project is funded with discretionary money in the lab's budget and done mostly outside the researchers' normal work.
Kilopower: Small Fission Power Systems for Mars and Beyond
http:\\spirit.as.utexas.edu\~fiso\telec … 2-1-17.pdf
Fission Surface Power
• 10 to 100 kWe
• 900 K Liquid metal (NaK) cooled reactor with UO2 fuel and stainless steel structure
• Stirling or Brayton power conversion
• 400 K composite radiators with H2O heat pipes
Nuclear Space Application Program 241pg
https://stuff.mit.edu/afs/athena/course … Report.pdf
Nuclear Systems Design project, this group designed a 100 kWe Martian/Lunar surface reactor system to work for 5 EFPY in support of extraterrestrial human exploration efforts.
The radiator is a series of potassium heat pipes with carbon-carbon fins attached. For each core heat pipe there is one radiator heat pipe. The series of heat pipe/fin combinations form a conical shell around the reactor. There is only a 10 degree temperature drop between the heat exchanger and radiator surface, making the radiating temperature 940 K.
At a distance of 11 meters from the reactor, on the shielded side, the radiation dose falls to an acceptable 2 mrem/hr; on the unshielded side, an exclusion zone extends to 14 m from the core. When combined together, the four systems comprise the MSR. The system is roughly conical, 4.8 m in diameter and 3 m tall. The total mass of the
reactor is 6.5 MT.
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quote from Kilopower: Small Fission Power Systems for Mars and Beyond, Comparison of solar and nclear
ISRU Demo Lander Study
Solar version:
• 4X 5.6m Ultraflex arrays
• Daytime ISRU only (1098 days to produce 4400 kg LOX*)
• Requires 4X 7.5m arrays and 1100 kg Li batteries for day/night ISRU ops excluding dust storm (527 days to produce 4400 kg LOX*)
Nuclear version:
• 1X 10 kW Kilopower reactor operated at 70% power
• Continuous ISRU operations (407 days to produce 4400 kg LOX*)
• Co-located reactor results in elevated radiation levels for ISRU equipment on lander deck
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Missions should use more than one power source and more than one technology to produce power where possible. This would avoid some possible failures.
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Power is a resource of which there can never be too much available. Should be a nuclear plant on virtually every flight, given how little a SAFE-400 weighs. Use some Solar panels for the backup to the nukes.
Last edited by Oldfart1939 (2017-03-29 10:10:15)
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What I found online about SAFE-400 says the core assembly is very light at 512 kg, but the cooling is not. The "400" refers to KW thermal, it only produces 100 KW electric. The other 300 KW must be carried away for dissipation to the environment. That is done with 127 heatpipe assemblies, each one of them 72 kg. So, the overall mass of the 100 KW electric power supply is 9656 kg.
It was not clear from what I found online whether this was for use in vacuo. But the articles did talk about it being used with an electric thruster "somewhere". It is harder to dissipate heat in vacuo, so perhaps on Mars you might not need quite so many heat pipe assemblies, and maybe they don't need to be fully 72 kg each. But I don't know, the articles didn't say. But you must pull 300 KW waste heat out of that core and get rid of it.
Text from article appended below.
GW
The SAFE-400 space fission reactor (Safe Affordable Fission Engine) is a 400 kWt HPS producing 100 kWe to power a space vehicle using two Brayton power systems – gas turbines driven directly by the hot gas from the reactor. Heat exchanger outlet temperature is 880°C. The reactor has 127 identical heatpipe modules made of molybdenum, or niobium with 1% zirconium. Each has three fuel pins 1 cm diameter, nesting together into a compact hexagonal core 25 cm across. The fuel pins are 70 cm long (fuelled length 56 cm), the total heatpipe length is 145 cm, extending 75 cm above the core, where they are coupled with the heat exchangers. The core with reflector has a 51 cm diameter. The mass of the core is about 512 kg and each heat exchanger is 72 kg.
SAFE has also been tested with an electric ion drive.
A smaller version of this kind of reactor is the HOMER-15 – the Heatpipe-Operated Mars Exploration Reactor. It is a15 kW thermal unit similar to the larger SAFE model, and stands 2.4 metres tall including its heat exchanger and 3 kWe Stirling engine (see above). It operates at only 600°C and is therefore able to use stainless steel for fuel pins and heatpipes, which are 1.6 cm diameter. It has 19 sodium heatpipe modules with 102 fuel pins bonded to them, 4 or 6 per pipe, and holding a total of 72 kg of fuel. The heatpipes are 106 cm long and fuel height 36 cm. The core is hexagonal (18 cm across) with six BeO pins in the corners. Total mass of reactor system is 214 kg, and diameter is 41 cm.
Last edited by GW Johnson (2017-03-29 13:46:42)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Seems to me that throwing away 300 KW as waste heat is squandering a lot of energy. Couldn't this heat be utilized in some useful manner? Maybe not initially, but later after we get men there who need heat?
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GW,
Is that 9656kg the weight of a properly shielded 400kWt output reactor?
Core is 58cm tall (includes reflector), so let's say the heat pipes have an outer diameter of 1.59cm and are 116cm in length. Instead of a pipe with a .089cm (edit: I missed the zero behind the decimal point) wall thickness, let's go with a solid bar of Mo:
V = pi*r^2*h
V = 3.14 * (.795*.795) * 116
V = 230.2
Mo bulk density is 10.28g/cm^3
10.28 * 230.2 = 2,366.5g = 2.3665kg
2.3665 * 127 = 300.5kg <- this is what 127 1.59cm x 116cm solid Mo rods should weigh
My math could be wrong, but either those heat pipes are a lot longer than I thought they were or they're not made out of Molybdenum.
IIRC, the 72kg refers to the weight of each Brayton cycle gas turbine that the He-Xe coolant spins. SAFE-300 (400kWt, 100kWe) was supposed to use 4 "heat exchangers" to produce 100kWe. I've no clue why they confused everyone by calling gas turbines heat exchangers. Heat exchanger, to me, implies heat pipe or part of a radiator system. Obviously they were referring to the electric generators.
If the core weighs 512kg and four gas turbines weigh 288kg, then the complete unit, less shielding, weighs less than 900kg. To properly shield the entire reactor at that output level would require multiple metric tons of W (gamma) and LiH (neutron). That could easily weigh 10t, depending on how close you want the astronauts to be able to get while it's in operation. If you bury it in a hole, you could get away with a 1t worth of shielding. The best new radiator materials are substantially lighter than the stuff ISS uses.
Last edited by kbd512 (2017-03-29 19:59:14)
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https://ntrs.nasa.gov/archive/nasa/casi … 005293.pdf
Closed Brayton Cycle Power System with a High Temperature Pellet Bed Reactor Heat Source for NEP Applications
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Mass of the SAFE-400 is very important.
Nuclear Reactors and Radioisotopes for Space
The SAFE-400 space fission reactor (Safe Affordable Fission Engine) is a 400 kWt HPS producing 100 kWe to power a space vehicle using two Brayton power systems – gas turbines driven directly by the hot gas from the reactor. Heat exchanger outlet temperature is 880°C. The reactor has 127 identical heatpipe modules made of molybdenum, or niobium with 1% zirconium. Each has three fuel pins 1 cm diameter, nesting together into a compact hexagonal core 25 cm across. The fuel pins are 70 cm long (fuelled length 56 cm), the total heatpipe length is 145 cm, extending 75 cm above the core, where they are coupled with the heat exchangers. The core with reflector has a 51 cm diameter. The mass of the core is about 512 kg and each heat exchanger is 72 kg.
SAFE has also been tested with an electric ion drive.
This says it has "two Brayton power systems", and "each heat exchanger is 72 kg". What is a "heat exchanger"? It also says "127 identical heatpipe modules". Does that 72 kg refer to a total power system, or a each heatpipe? I don't see how a single heatpipe could mass 72 kg. If it does, then that totals the extreme mass that GW Johnson calculated.
Another point is the radiation. "Nuclear Space Application Program" linked by SpaceNut says...
The specific mass of this radiator is only 1.6 kg/m² ... total radiator area of 150 m².
That means the radiator mass = 1.6 * 150 = 240 kg
Total system mass = 512 kg for reactor core + 240 kg for radiator + heat exchanger. So is the heat exchanger 72 kg * 2 or 72 kg * 127? If the latter, we have a mass problem.
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Rob,
I believe SAFE-300 (SAFE-400 series reactor that generates 400kWt at 100% of its rated thermal output and converts that at 25% efficiency to 100kWe) using 4 gas turbines connected to electric generators that each produce 25kWe. I can't recall where I saw that written, but I'm almost positive. In any event, 72kg for a gas turbine connected to a 25kWe electric generator is not an unreasonable mass and I can point to a commercial 25kWe DC electric generator that weighs 54.5kg.
This 25kWe DC electric generator weighs 120lbs:
The Franklin-Thomas Co., Inc. - 25 kW StarPower Generators
As I've shown above, 127 solid Molybdenum rods 1.59cm in diameter and 116cm in length would weigh 300kg. Obviously we're not using solid rods in a heat pipe system. Mo is awfully heavy, but the relatively thin Mo pipes contain liquid Na or a mix of He and Xe gases to drive the turbines, so the pipes would have to be awfully long to weigh 72kg each. On the other hand, 72kg is an entirely believable figure for a gas turbine powering a 25kWe electric generator. If you're right about the radiator mass, then SAFE-300 would weigh 1,040kg with no shielding.
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Well a part of the problem is info from various sources who don't really know what they are writing about. Typical "news reporter" problem. I know little about nuclear power, but I do have thermodynamics under my belt from mechanical engineering.
I took another look around trying to find substantive information about the SAFE 400 thing. My initial impression was wrong, interpreting the confusing source I quoted above. What follows is more sure. But less confident of numbers.
The heatpipes and reactor core are integral units. The liquid metal heatpipe has reactor fuel pins embedded inside. Clusters of these things make up the reactor's core. Maybe that weighs 512 kg, and maybe not. It depends upon what you include in your definition of "core".
These heat pipes heat up one side of the gas loop in heat exchangers that will power a turbine generator to make electricity. As near as I can tell there is one heat exchanger/one gas loop/one turbo generator as a "power conversion system". Most of what I read seemed to suggest there are two of these things on a SAFE 400. Nothing I read had a hard figure for how much these actually weigh, and again, it depends on your definition of "power converter" as what gets included on that category.
The overall thermal efficiency of the thing is 25% at max output. The core generates 400 KW heat, the heat pipes deliver it to the heat exchangers, which provide what I think I read to be 800 C gas as input to the turbo-generator. 100 KW electricity is produced. That means you either find uses for 300 KW of low-grade low temperature heat energy, or you dispose of it. In space that requires a radiator. NOBODY had figures for the weight of the radiator.
On Mars, you have two additional sinks for waste heat: the atmosphere and the solid surface. NOBODY indicated either of those was being used.
As for radiation shielding, yes, it is needed, and NOBODY had any figures for how much it weighs. If you power a manned vehicle in space with this thing, you will need lots of shielding, and some significant waste heat radiator wings. Unmanned, you can use less shielding, but not none, because most electrical/electronic gear and lots of materials are vulnerable to active fission core radiation.
I think the "512 kg core weight" that is so often quoted is a marketing ploy; the actual system installed weight will be a lot higher, especially if men are involved. But I cannot dope out how heavy.
GW
Last edited by GW Johnson (2017-03-30 09:26:26)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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GW,
MIT did a study on a Mars Surface Fission Power System. I think we have a link to it here in the forum. We already know what types of radiator materials would be used from NASA's documentation. SpaceNut determined what the radiator surface area would have to be to dissipate 300kWt.
We can use what we know about halving thicknesses of shielding materials and Mars regolith bulk density to gauge how much shielding material is required for an exclusion zone of a specific radius from the center of the core. We know exactly how thick the reflector is and what it's made of, we know what the required halving thicknesses are for W (High-Z) and LiH (Low-Z) materials are, we know how what the gamma and neutron fluxes will be for a given level of power output, etc. It's not hard to ballpark it and we've gone through this exercise before.
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Descriptions of SAFE-400 linked in this thread appear to state total system mass is: 512kg reactor core + 2*72kg gas turbine heat exchangers + 240kg radiator = 896kg. However, the fact they use the words "heat exchangers" would imply that does not include the alternator. One of the links SpaceNut provided (https://stuff.mit.edu/afs/athena/course … Report.pdf) pages 155-157 provides the only description I've ever read about the radiator. It also includes a schematic for the power conversion system. It shows a turbine driving an alternator and a compressor, as well as power conditioning and controls. Wording implies the electrical stuff is not including in mass, so add something more for alternator, power conditioning and controls.
Still, this is significantly less mass than SP-100. I expect that considering SAFE-400 produces 400 kWt while SP-100 must produce 2,000 kWt.
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The "heat exchangers" can be thought of in the form of a power transformer where the primary coil is the inlet from the reactor and the output from the secondary is the feed into the turbine. It is done to remove the excess heat which could be felt by the reactor temperatures from reaching the turbine. The outlet side of the turbine runs to the radiator to make that loop more efficient for the heat exchager side of the circuit. The heat exchange is a coupling of the reactor to the power generator system and is done via different media types. Some times the core of the tank which holds the tubing from the reactor side coils and the turbine side of the output to the power system will have a third type of media which allows for the core to act as a heat loss area as well so as to reduce the size of the radiator.
The transfer ratios of the reactor to power generation is simular to the transformer ratio as well in how it will work to control the heat content on the primary versus the output of it on the secondary side.
This is a power in to power out where the transfer of heat is the variable which we are trying to control which is a volume function of tubing size to coil length inside the holding tank for the exchanger.
The radiator measurement as are in the nuclear rover topic and can vary depending on the sink to remove heat as GW indicated with soil, dry ice ect.....
Now to gain more energy out of the left over 300 kw one can change the loop of power to contain another heat exchange for another different power generation unit and you can also insulate the holding tanks for the exchange to control the heat loss as well as covering the conneting tubing throughout the reactor to power systems unit or units.
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