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Microsoft salesmen say support for Windows 7 has ended, but that doesn't mean it doesn't work. It still works. And you can still download all the updates, they're just not writing any more. My laptop still runs Windows 7. That version of Windows allows *YOU* to control when/if updates happen, and you can select which updates to install.
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I suspect tahanson43206 has confused two mission plans that I proposed. An expedition with a Mars Direct size habitat. This would include a second habitat dedicated to in-space use only. The space habitat would be reusable, flying from Earth orbit to Mars orbit and back. The Mars Ascent Vehicle would have over-size propellant tanks, because it would act as the TEI stage. That allows for ISPP for return propellant. Because the in-space habitat goes with the surface habitat, that means if free return is necessary they have all the food with them.
The big ship was in response to Elon Musk's statement that Starship version 2 would be 8 times the volume. Starship itself is huge. I said if you're going to build a reusable ship that big, then design it to remain in space and use artificial gravity. The big wheel ship was designed to have the same volume as Starship version 2... but then the hub, so it's a little bigger. The big ship is not for exploration or expedition, it's definitely for settlement. Would use Elon's original Starship as a shuttle for passengers.
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Robert,
I liked your concept. In order to truly colonize space, we'll need ships of the size you proposed. I think I've been advocating for the same thing. The SpaceX Starship is a tinker toy compared to what would actually be required for colonization to occur.
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Nuclear thermal doesn't offer order of magnitude improvements, but "merely doubling" of the most powerful chemical system currently in use.
Reported Isp of nuclear thermal using Hydrogen as the ejected fluid has been reported as 900 sec, versus hydrolox at 450 sec.
For missions to intermediate distances beyond Mars orbit--say the asteroid belt and moons of Jupiter--this would make possible reasonable transit times for humans. For deeper space human manned missions such as Saturn and satellites, probably not good enough.
My choice for a deeper space objective has become Callisto, the most distant from Jupiter of the Galilean satellites. Has water, acceptable surface radiation environment, and even a thin atmosphere of CO2.
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Reported Isp of nuclear thermal using Hydrogen as the ejected fluid has been reported as 900 sec, versus hydrolox at 450 sec.
Nope. That's solid core, such as NERVA. Increase ISP by increasing temperature. Engineers had a problem, too hot melts the reactor core. So they designed one intended to melt. Then boil the uranium to gas. A couple different designs. Highest Isp is open cycle, theoretical only, no prototype. Estimated Isp varies, highest 9,000 seconds.
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What I saw is decades old now. Open-cycle gas core is up to 2000-2500 sec Isp at T/Wengine > 10 with regenerative cooling. With an installed radiator system and higher nuclear operating power, Isp is up to 6000 sec at T/Wengine < 1, probably less than 0.1. One of the advocates of this approach was somebody named Maxwell Hunter. I don't honestly remember what outfit he worked for.
That concept is open-cycle, meaning the nuclear fireball is in the same can as the working-fluid propellant, with a flow scheme such that hydrogen flow rate/uranium flow rate equals or exceeds 1000:1, which with burnup rates, is effectively "perfect" containment. This is very most definitely is NOT the nuclear light bulb concept, which separates the two flows with some sort of cooled clear window.
The operation of a nuclear reactor in gas phase was demonstrated. The 1000:1 flow scheme was demonstrated in plasma flow devices. The two were never demonstrated in the same device. THAT is were we were, and still are.
GW
Last edited by GW Johnson (2020-07-03 17:46:54)
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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Ok. 6,000 seconds is still an order of magnitude greater than chemical. BTW Oldfart1939, there are hydrolox engines optimized for use exclusively in space with Isp 460 sec.
NASA decided to replace the upper stage engine for SLS with 4 RL-10C engines instead of a single J-2X. RL-10 engines are a lot lower thrust, even 4 add up to a fraction of the thrust of a single J-2. But RL-10C has very high Isp, so the stage can be turned on and just left on gradually accumulating velocity until propellant runs out. Net result is greater throw mass to the Moon or Mars.
But Thrust to Engine Weight ratio less than one? Please tell me we can achieve the higher Isp with more thrust than that. Pretty please? Timberwind dramatically reduced engine mass dramatically vs NERVA. I realize there are physics limitations with heat transfer at high exhaust velocity.
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RobertDyck:
Those thrust results have nothing to do with high exhaust velocity heat transfer, and everything to do with the heat transfer of high core power output in the chamber. Above something like 2000-2500 sec Isp, the chamber temperatures were just too high to cool with regenerative cooling. The propellant throughput flow was insufficient to absorb the heat at those conditions, complicated by thermal radiation transparency effects. That required an active cooling system with a powered waste heat radiator, which is inherently very heavy.
The 6000 Isp gas core engine was the backup design for the solid core NERVA engine for the 1983 Mars mission, as understood in the 1960's. Third choice was chemical propulsion. Orbit-to-orbit, the low thrust/weight ratio of high-Isp gas core propulsion made no difference. This is the mission I was trying to prepare for, when I accepted that appointment to the Naval Academy in 1969. The idea was to serve combat tours in Vietnam, then go to test pilot school, then enter the astronaut corps. For a 1983 mission, I would have been 33 years old, which is just about right.
All that died with Nixon's executive order in 1972 killing all human flight outside Earth orbit. NERVA died by 1974, after its last ground test in 1973 or 1974, getting ready to fly as an alternate 3rd stage for Saturn-5. "Why build the rocket if we are not going to go" was NASA's management approach.
I washed out of the Naval Academy in Sept/Oct 1969 because I got too sick to stay, and returned to my previous college career at UT Austin, which had started back in 1968. Thus I went civilian, building the weapons for others to use in Vietnam and subsequent, instead of using them in combat myself.
If you stayed under 2000-2500 sec Isp, which is a lot less core power output, then regenerative cooling was feasible with LH2 with open-cycle gas core. That's enough thrust/weight to take off the launch pad, and it's enough Isp to fly single stage to the moon and back, with around 20% payload.
After thinking about it for a while, I seem to remember that Maxwell Hunter worked for United Aircraft in the 50's and 60's. He and they were competitors to General Atomics and some others, back in those days. Memory fades as you age. I can't help that.
Timberwind was an alternative fuel element and core design for solid core nuclear thermal. It was not reusable, not even re-startable. One-shot only! Never ever actually tested on the thrust stand. Paper design only! Open-cycle gas core is VERY, VERY different from Timberwind!
Just the memories and musings of someone just about to turn 70 years old next week. It takes time to remember these things. Once upon a time, they were instant recall. No more.
GW
Last edited by GW Johnson (2020-07-04 14:17:36)
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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Robert-
I was going strictly on the Nerva data, as that was a demonstrated and flyable design, versus Hydrolox. Hydrolox Isp is generally listed as 450 sec, but a vacuum optimized design--probably can get the additional 10 sec.
I seem to recall Elon saying something about a vacuum optimized Raptor getting an Isp maybe 380 with Methylox.
Last edited by Oldfart1939 (2020-07-04 15:13:26)
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The 6000 Isp gas core engine was the backup design for the solid core NERVA engine for the 1983 Mars mission, as understood in the 1960's.
1990 update to NERVA used new high temperature materials, increased Isp from 825 to 925. Paper only, no prototype. How could that improve gas core?
Just the memories and musings of someone just about to turn 70 years old next week.
I turned 58 this week. Am I really that far behind?
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RD-0126, hot tested 1998, Isp 476 sec.
RD-0146, proposed replacement for RL-10-A-1, Isp 463 sec. Main modification was turbopump.
Of course the RD series are Russian.
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One figure I saw for the space shuttle main engines was listed as 467 sec Isp. I think that would be vacuum Isp for what had to be a sea level nozzle expansion design. They had to have max possible thrust at ignition. Which means the backpressure term had to be zero at sea level ignition conditions.
The disparities among all these numbers reflect mainly the operating chamber pressure for the engine design. Things like c* and CF are just higher for the higher pressures. It's quite the significant effect. You usually need a slightly different r mixture ratio as you change chamber pressures, too. That also affects c*, but not so much CF.
GW
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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This has been an interesting topic. What I've been able to take away is that the Nerva type rocket engine has it's Isp strictly limited by thermal properties of the construction materials; 1960-197 metallurgy was not sufficiently advanced to allow higher temperatures, and that the observed Isp is not being limited by the reaction so much as by the ability of core construction being unable to support higher temperatures.
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Manufacturer's designation for Space Shuttle Main Engine (SSME) was RS-25. For SLS they're now using that designation. NASA requirement was for Isp in vacuum 455 seconds. Manufacturer didn't quite achieve that, the engine as built had Isp in vacuum 453 seconds, and at sea level 366 seconds. NASA said close enough.
Nozzle extension could improve Isp in vacuum, but the engine couldn't be used in atmosphere. Atmospheric pressure would cause exhaust flow to separate from the nozzle extension, which would cause excessive heating. The extension would melt off in atmosphere.
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So why not a moveable extension which is housed at launch and extended at high altitude?
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Bob Clark has been the lead exponent of unusual and extendible nozzle designs on these forums.
GW
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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Why compromise on performance or require further costly and time-consuming basic development work when the existing RS-25 engine's nozzle could be optimized for expansion ratios associated with in-vacuum-only operation or people could simply learn to accept that all engineering refinement has limitations that preclude substantially better performance without brand new engine development?
Isn't this just another way to ensure that the mission never happens by subjecting ourselves to a continual development cycle directed at chasing down those last few percentage points of performance / design optimization while not actually building flight ready hardware and preparing it for useful work on a real mission?
Engineers, as amazing as you people are, our operations people need to put their foot down and call it "good enough for government work".
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trouble is "good enough for government work" is a major part of the problem due to the extraction of taxpayer cash being the objective.
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There's nothing wrong with the RS-25 design. It is a very well-performing LOX-LH2 sea level engine-and-nozzle design. That design combines a very careful balancing act to drive the two propellant species turbopumps, and another very careful balancing act getting the regenerative cooling of the chamber and nozzle bell with the LH2.
That last is having the hardware run cool enough without boiling the LH2, all at an LH2 feed pressure you can tolerate, which also impacts the turbopump design balance. This stuff is all very intimately connected. So you don't go making significant changes lightly. There is no "lightly" to it.
It might be possible to add an uncooled bell extension of limited size to the RS-25 without too much design change impact. The operative word in that statement is "might". It cannot be so large as to incur flow separation when you fire the thing at sea level. But, if that can be done with an uncooled extension the way the Merlin bell extension was added, then you have a practical design modification to pursue.
You will pay a price (relative to the unmodified RS-25) in reduced sea level thrust and Isp, for a modest gain in vacuum thrust and Isp. However, if adding an uncooled bell extension is not feasible, then it is hardly worth trying to change the RS-25 at all, because the entire regenerative cooling balancing act has to be redone from scratch. Which really means an essentially new engine design. Sorry, but that's just an ugly little fact of life.
The folks who could best determine whether an uncooled bell extension is even feasible would be the manufacturer. Not any of us here in these forums. They have the database and the test experience. I just offer the concept as one they might possibly pursue.
The baseline RS-25 nozzle expands from chamber pressure to something right at one standard atmosphere expanded pressure, when operating at some design-point throttle setting that is probably not 100%. I don't know what any of these numbers are for that engine. But that design-point selection sets the bell expansion ratio and engine operating limitations.
What you would do with the modification is expand further to a pressure somewhat less that one atmosphere, but not so low as to incur flow separation. You probably would have to revisit the design throttle setting and the whole set of operating limitations. But if this works uncooled, then you don't have to redo the bell regenerative cooling balance, you just tack-on the radiation-cooled heat balance for the extension kit. If it proves to be feasible at all. It might not.
Now if and when you do this, the sea level thrust's massflow x exit velocity momentum term is a bit bigger, but the exit minus ambient pressure x exit area term is now significant and rather negative. You thus inherently lose sea level thrust (and Isp) doing this. Out in vacuum, the momentum term is the same, which is slightly larger than the baseline design by the effect of extra expansion ratio, and the exit pressure term is crudely comparable to what it was before, which is positive. Vacuum thrust with any design is always higher than sea level thrust, because of that exit pressure difference term. That's just physics.
So, this modification is only attractive if you can afford to lose significant sea level thrust capability, in order to gain better vacuum Isp capability. It's not a trivial change, so this is not a trivial tradeoff choice. Even if it is feasible.
GW
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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elderflower,
I sincerely doubt that the RS-25 was developed to fleece the tax payers.
Why is historical achievement always judged from the standpoint of modern capabilities?
Improvements are always possible, but at what cost?
How much more can the vacuum performance of the RS-25 engine actually be improved?
It's already 453s in a vacuum. A complete optimization for vacuum operation could conceivably obtain another 15 seconds, or less?
If we spend that money, then we have an unproven engine that can't be used at sea level. To what end?
Those are the questions I'd ask if I held the purse strings.
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I looked up the RS-25 engine. There will be no bell extension that takes it from perfect sea level expansion to an over-expanded but unseparated condition. It's already there.
The existing bell is already 69:1 area ratio, expanded to 4-5 psia at the exit plane. They over-curled the exit lip back toward axial, which reduces local Mach but raises expanded pressure, right around around the rim. It is way-overexpanded to around 2 psia in the center core.
Any bell extension could not operate at sea level, only out in vacuum. Is that what you want? You could not ground test that down here on Earth, except inside some sort of giant vacuum chamber.
The specs show for the RS-25D 452.3 s Isp in vacuum, at the rating point thrust level. That was the latest version of the Shuttle engine. The SLS engine is a one-shot/reduced cost variant. It wears out far quicker.
GW
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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The fleecing took place on the J2-x engine since they are not using it.
https://www.nasa.gov/exploration/system … index.html
https://en.wikipedia.org/wiki/J-2X
https://old.texasarchive.org/a_journey_ … et_engine/
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NASA Announces Nuclear Thermal Propulsion Reactor Concept Awards
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