From: Henry Spencer <henry@zoo.toronto.edu>
Newsgroups: sci.space.science
Subject: Re: URGENT: CNN Live: Water found at Lunar Poles
Date: Mon, 16 Dec 1996 03:45:42 GMT

In article <58lcfn$3aq@nntp.interaccess.com> dietz@interaccess.com (Paul F. Dietz) writes:
>...Unfortunately, superhot steam tends to destroy the materials
>you want to make a reactor from (carbon, uranium carbide, and
>so on.)  You can make the reactor with uranium oxide as fuel
>and ceramic structure, but UO2 is comparatively volatile...

This has been looked at, a little bit.  You use ThO2 for the fuel-element
structure, with embedded UO2 for the fuel itself.  ThO2 is chosen for its
physical and thermal properties; its melting point is very high and it is
not volatile.
-- 
"We don't care.  We don't have to.  You'll buy     |       Henry Spencer
whatever we ship, so why bother?  We're Microsoft."|   henry@zoo.toronto.edu

From: Henry Spencer <henry@zoo.toronto.edu>
Newsgroups: sci.materials,sci.space.policy,sci.space.science,sci.space.tech
Subject: Re: Mars Mission - Solar Flare Question...
Date: Sat, 4 Apr 1998 06:17:12 GMT

In article <35232812.2000@Boeing.com>, Erik Zahn  <Erik.Zahn@boeing.com> wrote:
>} Small problem:  we don't have an operational nuclear-rocket system, and
>} the idea of developing one has huge political problems (not to mention
>} being rather expensive).
>
>	The data and test bed is old.  NASA had a project in the 1960's for
>this thing.  Just have to make it...

Not that easy.  Subcontractors gone, some materials hard to get, can't
duplicate the old hardware so exactly that we don't need to re-run the
test program.  No longer acceptable to test nuclear rockets in the open
air.  Scrubber systems which can handle the exhaust of a large rocket
engine are difficult and expensive to build.  Feasible, yes; easy, quick,
cheap, no.

>} >...H2 as fuel on the outbound leg, and CO2 on the return leg...
>}
>} That's going to be an interesting engine -- what sort of materials did you
>} envision using?
>
>	How about Magneseum?  For reference see the 30+ years of Magnox
>reactors use in Britain.  They use CO2 as the coolant...

Yes, but (a) they run at relatively low temperatures and (b) they don't
also run hot hydrogen through the reactor.

A rocket engine must run *much* hotter than a power plant.  Pure magnesium
melts at only about 650degC, which is useless in a rocket.  By comparison,
the NERVA engines ran their cores at about 2700degC, and would have gone
higher if materials problems had not been so severe.

To handle really hot CO2, you use oxide materials.  Such materials tend to
quietly but rapidly wash away in hot hydrogen.  Similarly, the materials
which will handle hot hydrogen don't last long in hot CO2.

Chemical rockets can avoid most of these problems by keeping hot gases
away from the solid materials.  Solid-core nuclear rockets *can't*,
because it's the solid materials that are heating the gases!  They must
have intimate contact between the solid core and the exhaust gas, at the
full working temperature.  The materials problems are really bad even if
you stick strictly to reducing gases or oxidizing gases; trying to run
both through the same reactor is a nightmare.

>...Besides, a little corrosion in
>flight is ok.  Won't matter much, not like interplanetary space isn't
>already loaded with radiation, and who is going to live there??? ...

The issue is not "a little corrosion", but mass loss that would be rapid
going on catastrophic.  As a very rough rule of thumb, many chemical
reaction rates double for each 10degC rise in temperature.  It doesn't
take much thought to realize what this does to corrosion rates when the
hardware is running at 2700degC.  (Don't take the numbers too literally;
the doubling rule is not accurate over such a huge temperature range.  But
the magnitude of the problem should be clear.)

Unless you do something very clever indeed, you just can't run a
metal/graphite nuclear engine on CO2, or an oxide engine on H2.
--
Being the last man on the Moon                  |     Henry Spencer
is a very dubious honor. -- Gene Cernan         | henry@zoo.toronto.edu

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.history
Subject: Re: THE Limiting Factor to Exploring Space
Date: Tue, 23 Feb 1999 06:37:49 -0600

Henry Spencer wrote:

> I don't think so.  Unfortunately, ammonia and water perform a *lot* worse
> in an NTR than hydrogen -- they cut the Isp by about a factor of three,
> which usually brings a solid-core NTR down into the chemical-rocket range.

Lumping water and ammonia together is a bit unfair
to ammonia.  Ammonia decomposes at high temperature
into nitrogen and hydrogen, giving a mixture with
an effective molecular weight of 8.25.  Water remains
mostly intact, with a MW of 18.  Also, water is
oxidizing, and will react with the engine components
(like carbon), which forces the use of oxidation-
resistant engine components, limiting the temperature
that can be achieved.

	Paul

Newsgroups: sci.space.history
From: henry@spsystems.net (Henry Spencer)
Subject: Re: THE Limiting Factor to Exploring Space
Date: Tue, 23 Feb 1999 14:44:44 GMT

In article <36D2A11D.89B0B8E0@interaccess.com>,
Paul F. Dietz <dietz@interaccess.com> wrote:
>> I don't think so.  Unfortunately, ammonia and water perform a *lot* worse
>> in an NTR than hydrogen...
>
>Lumping water and ammonia together is a bit unfair
>to ammonia.  Ammonia decomposes at high temperature
>into nitrogen and hydrogen, giving a mixture with
>an effective molecular weight of 8.25.  Water remains
>mostly intact, with a MW of 18.

True, although the ammonia decomposition reaction is discouraged by high
pressure, and is also endothermic (absorbs some heat).  The experience of
using ammonia as a fuel in chemical rockets has not been a happy one, with
many development difficulties, although it's hard to say whether that
would carry over to NTRs.

And even at 8.25, the advantage over chemical rockets is rather limited.

(For those who aren't up on this...  The key problem is that solid-core
nuclear thermal rockets must be run cool enough that the core doesn't
melt, while chemical rockets are under no such constraint; temperatures in
a high-pressure LOX/LH2 flame are very high.  *The* big advantage of
solid-core nuclear rockets is that their exhaust is all hydrogen, which
gives very high performance because of its very low molecular weight and
some other reasons, while the hotter chemical rockets have to accept an
exhaust with a lot of heavier and less helpful molecules.)

>Also, water is
>oxidizing, and will react with the engine components
>(like carbon), which forces the use of oxidation-
>resistant engine components, limiting the temperature
>that can be achieved.

As I recall, the problem is not so much that it's hard to build an
oxidation-resistant NTR, as that it's hard to build one that can run well
on *both* oxidizing and reducing mixtures.  That is, you can probably
build one which will run hot on water (or carbon dioxide, another
propellant of interest that turns oxidizing at high temperatures), but you
can't run the same engine hot on hydrogen or ammonia.  So there hasn't
been much interest in trying.
--
The good old days                   |  Henry Spencer   henry@spsystems.net
weren't.                            |      (aka henry@zoo.toronto.edu)

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.history
Subject: Re: THE Limiting Factor to Exploring Space
Date: Tue, 23 Feb 1999 19:24:47 -0600

> And even at 8.25, the advantage over chemical rockets is rather limited.

I should have written 8.5 (oops).

> As I recall, the problem is not so much that it's hard to build an
> oxidation-resistant NTR, as that it's hard to build one that can run well
> on *both* oxidizing and reducing mixtures.  That is, you can probably
> build one which will run hot on water (or carbon dioxide, another
> propellant of interest that turns oxidizing at high temperatures), but you
> can't run the same engine hot on hydrogen or ammonia.  So there hasn't
> been much interest in trying.

When I looked at this a while ago, I found that the
really refractory compounds tend to be carbides and
metals.  Oxides tended to be relatively more volatile.
I wonder what compound they proposed using for the
structure of the oxidation-resistant engine --
boron nitride (the boron is acceptable, since it
would be a fast reactor)?

Note followup.

	Paul

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.history
Subject: Re: THE Limiting Factor to Exploring Space
Date: Tue, 23 Feb 1999 20:28:18 -0600

> Nuclear rockets on the other hand are power limited. There is only so
> much heat that can be transferred by a solid (the fuel elements) at a
> temperature below its melting point.

Perhaps more precisely, nuclear rockets are entropy
limited.  Since there is an upper bound T on the temperature
set by the materials, for a given quantity Q of heat
transfered the entropy transfered is at least Q/T.
The entropy has to be carried away by the reaction
mass (neglecting thermal radiation), and the reaction
mass has a limited capacity to do this.

What this suggests is that a nuclear engine be carefully
designed to avoid unnecessary creation of additional
entropy.  For example, one might expand the
reaction mass through internal turbines quasi-isothermally
before injecting it into the main engine, and use the
work produced by the turbines to generate electricity
to drive an arc "afterburner" in the nozzle.

Note followup.

	Paul

From: henry@spsystems.net (Henry Spencer)
Newsgroups: sci.space.tech
Subject: Re: oxygen-resistant coatings for NTR's
Date: Thu, 30 Dec 1999 00:57:04 GMT

In article <84c3ml$3h7$1@nnrp1.deja.com>,  <cray74@hotmail.com> wrote:
>> Are there any good references to coatings (metallic
>> or ceramic) that would protect NTR elements from an
>> oxygen-rich propellant (such as water or even straight LOX)?
>
>My first response is to suggest switching to alternate
>reaction masses lacking oxygen. Hydrogen and ammonia are
>plentiful enough in space...

Not in near-Earth space, alas.  Hydrogen is scarce in the inner solar
system, except for Earth's oceans.

However, ammonia in particular is an excellant propellant for an NTR,
because it's reasonably dense and unlike water, it will dissociate to a
really useful extent (enough to change its propellant properties for the
better).  It's not clear what source would make water available but not
ammonia.

Anyway... if you're using water in particular, I've seen it suggested that
adding a *little bit* of hydrogen to the water will bias the dissociation
equilibrium very strongly away from oxidizing forms, possibly enough to
permit use with a conventional NTR.

If you're using LOX, then you'd better make the reactor hardware out of
oxides, or at least coat it with oxides.  Nothing else will really do.
Even fluorides will react with oxygen under heat and pressure.
--
The space program reminds me        |  Henry Spencer   henry@spsystems.net
of a government agency.  -Jim Baen  |      (aka henry@zoo.toronto.edu)

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.tech
Subject: Re: solar-nuclear electro-thermal engine
Date: Sat, 21 Oct 2000 04:58:20 -0500

"Geoffrey A. Landis" wrote:

> You're describing a nuclear electric rocket with a resistojet thruster.
> The point here is, since a nuclear reactor produces heat and then
> generates electricity by heating a working fluid, why not just eliminate
> the middle step, and just exhaust the working fluid?  What point is there
> in running the working fluid through a turbine to produce electricity,
> and then use the electricity to heat a fluid?

Well, the idea doesn't make a whole lot of sense
with a resistojet, but it might be worthwhile with
a higher temperature electric heater, as part
of a nuclear thermal rocket.

The key observation here is that, unlike a chemical
rocket, a nuclear thermal rocket is limited not by
energy, but by entropy.  It may make sense to examine
the engine for steps that unnecessarily produce entropy.
The step of injecting cold hydrogen into a hot reactor
is just such a step.  Instead, the engine could be designed
to extract useful work during this step.  This work,
converted to electrical energy, could be used to power
an electrical afterburner.

In more detail: the cycle could work by first (1)
compressing the incoming liquid hydrogen to high
temperature, (2) heating it,  (3) near-isothermally
expanding it (in turbines?) to the pressure for
injection into the thrust chamber.

This could yield higher Isp than a purely nuclear
thermal engine, yet it would not require the large
radiators of a nuclear electric rocket.

	Paul

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.tech
Subject: Re: solar-nuclear electro-thermal engine
Date: Sun, 22 Oct 2000 13:56:20 -0500

I wrote:

> compressing the incoming liquid hydrogen to high
> temperature,

High *pressure*.  Ack.

	Paul

Newsgroups: sci.space.tech
From: henry@spsystems.net (Henry Spencer)
Subject: Re: solar-nuclear electro-thermal engine
Date: Sun, 22 Oct 2000 22:31:43 GMT

In article <39F168BC.3111CEDC@interaccess.com>,
Paul F. Dietz <dietz@interaccess.com> wrote:
>...It may make sense to examine
>the engine for steps that unnecessarily produce entropy.
>The step of injecting cold hydrogen into a hot reactor
>is just such a step.

It would be, if this were done.  The hydrogen is invariably used to cool
the nozzle and chamber first, which means it's no longer cold.  Especially
since the only obvious way to run the LH2 pump is to use an expander cycle
(or a dump cycle), using the warm GH2 to drive the pump turbine.
--
Microsoft shouldn't be broken up.       |  Henry Spencer   henry@spsystems.net
It should be shut down.  -- Phil Agre   |      (aka henry@zoo.toronto.edu)

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.tech
Subject: Re: solar-nuclear electro-thermal engine
Date: Sun, 22 Oct 2000 19:26:39 -0500

Henry Spencer wrote:

> It would be, if this were done.  The hydrogen is invariably used to cool
> the nozzle and chamber first, which means it's no longer cold.  Especially
> since the only obvious way to run the LH2 pump is to use an expander cycle
> (or a dump cycle), using the warm GH2 to drive the pump turbine.

Yes, but the scheme I described (which pumps the LH2 to a higher
pressure, and taps off some of the work as electricity)
would still allow the hydrogen to be used as thrust chamber/
nozzle coolant.  There'd just be additional heating stages
for further expansion before the (relatively low pressure)
gas was injected into thrust chamber.

	Paul

From: James A Davis <jimdavis2@primary.net>
Newsgroups: sci.space.tech
Subject: Re: solar-nuclear electro-thermal engine
Date: Mon, 23 Oct 2000 18:31:15 -0500

"Paul F. Dietz" wrote:

> Yes, but the scheme I described (which pumps the LH2 to a higher
> pressure, and taps off some of the work as electricity)
> would still allow the hydrogen to be used as thrust chamber/
> nozzle coolant.  There'd just be additional heating stages
> for further expansion before the (relatively low pressure)
> gas was injected into thrust chamber.

What you seem to be advocating is some form of thermal regeneration to
modify the rocket's thermodynamic Brayton cycle so that it more closely
approaches a Carnot cycle. This technique is used commonly with
stationary power plants but has found almost no airborne applications
because the additional weights overwhelm any potential thermodynamic
gain.

For rockets the potential improvement is a lot less. Rockets can have
thermal efficiencies of ~75% (compared to about 25 - 35% for gas
turbines). If you can't justify regeneration techniques for aircraft gas
turbines there seems little chance of doing so for rockets which are so
much more sensitive to inert weights.

Jim Davis

From: "Paul F. Dietz" <dietz@interaccess.com>
Newsgroups: sci.space.tech
Subject: Re: solar-nuclear electro-thermal engine
Date: Tue, 24 Oct 2000 05:39:30 -0500

James A Davis wrote:

> What you seem to be advocating is some form of thermal regeneration to
> modify the rocket's thermodynamic Brayton cycle so that it more closely
> approaches a Carnot cycle. This technique is used commonly with
> stationary power plants but has found almost no airborne applications
> because the additional weights overwhelm any potential thermodynamic
> gain.
>
> For rockets the potential improvement is a lot less. Rockets can have
> thermal efficiencies of ~75% (compared to about 25 - 35% for gas
> turbines). If you can't justify regeneration techniques for aircraft gas
> turbines there seems little chance of doing so for rockets which are so
> much more sensitive to inert weights.


Jim,

  You misunderstand.  In a chemically fueled system, you'd
be right.  Those systems are limited by the energy available
in the propellant.  Regeneration doesn't add new energy to
the propellant, and rocket engines (especially space engines
with large expansion ratios) are already very efficient.

  In the nuclear design I was discussing, the purpose of
the design is not to convert a given quantity of fission
energy into work more efficiently, but rather to get more
fission energy into the propellant.  No need to recycle
heat regeneratively when the reactor can give you additional
thermal energy 'for free'.

  Sensitivity to inert weight:  for space engines, it
can be worthwhile to have a lower thrust/mass ratio if
it increases specific impulse.  In this sense the
airplane analogy is misleading.

	Paul

Newsgroups: sci.space.tech
From: henry@spsystems.net (Henry Spencer)
Subject: Re: Nuclear Propulsion Questions
Date: Wed, 14 Feb 2001 15:15:47 GMT

In article <3A892CBF.1523A933@knutsen.dk>,
Peter Knutsen  <peter@knutsen.dk> wrote:
>> nuclear reactor creating a core temperature of approximately 2000K, with
>> hydrogen gas as the propellant...
>
>Offhand, I'd say that it's better to use a denser reaction mass,
>otherwise tankage and structural mass would get too high. Like
>water. Or at least nitrogen if you want something inert.

Unfortunately, performance suffers severely.  For a thermal rocket where
the energy source and the propellant are separate, hydrogen is miles out
in front for performance.  Exhaust velocity is proportional to the square
root of temperature over molecular weight, so hydrogen's molecular weight
of 2 gives it a huge advantage over water (18), nitrogen (28), etc.  A
conventional solid-core nuclear rocket using dense propellants has trouble
achieving better performance than a LOX/LH2 chemical rocket.

If you must use a non-hydrogen propellant, ammonia (NH3) is about the
best.  It will mostly decompose into nitrogen and hydrogen at rocket
temperatures, giving an effective molecular weight of about 6.  The one
snag is that you get atomic hydrogen as an intermediate step in the
decomposition, which makes it somewhat corrosive.
--
When failure is not an option, success  |  Henry Spencer   henry@spsystems.net
can get expensive.   -- Peter Stibrany  |      (aka henry@zoo.toronto.edu)

Newsgroups: sci.space.tech
From: henry@spsystems.net (Henry Spencer)
Subject: Re: Nuclear Propulsion Questions
Date: Wed, 21 Feb 2001 14:32:25 GMT

In article <3a8b6b0d.10819957@news.earthlink.net>,
Trakar <shardrukar@yahoo.com> wrote:
>Yeah, ammonia is a pretty good choice, so is methane (still cryo, but
>much better densities, and the insulation isn't near as critical) I
>don't know that I'd completely rule out many non-cryo fuels, you
>suffer a lot of ISP loss and possible coking problems with some of the
>hydro-carbon fuels, but...

The nice thing about ammonia is that all of its decomposition products are
gases, which unfortunately isn't true of carbon-based fuels except at very
high temperatures.  Carbon particles in the exhaust not only give you
potential problems with deposits in the engine, but they also tend to
reduce Isp, because they hold thermal energy and don't transfer it well to
the gases on the time scales involved.  Maybe if you use methanol, where
there's an oxygen atom to mate up with the carbon...
--
When failure is not an option, success  |  Henry Spencer   henry@spsystems.net
can get expensive.   -- Peter Stibrany  |      (aka henry@zoo.toronto.edu)

Newsgroups: sci.space.tech
From: henry@spsystems.net (Henry Spencer)
Subject: Re: Nuclear Propulsion Questions
Date: Wed, 21 Feb 2001 15:13:28 GMT

In article <3A91938E.F6941BB2@knutsen.dk>,
Peter Knutsen  <peter@knutsen.dk> wrote:
>> ...Exhaust velocity is proportional to the square
>> root of temperature over molecular weight, so hydrogen's molecular weight
>> of 2 gives it a huge advantage over water (18), nitrogen (28), etc...
>
>OK, so exhaust velocity is x9 better for hydrogen than for water...

Only x3, actually, because of the square root.  Still a very large
difference if the delta-V requirement is large enough for the mass ratio
to be substantial -- a x3 hit in exhaust velocity *cubes* the mass ratio.
A mass ratio of 5 (reasonable) turns into 125 (pretty much out of the
question).

For chemical rockets, dense propellants are competitive with hydrogen in
many applications, with the smaller tanks and more powerful engines and
fewer hassles making up for the poorer exhaust velocity... but there the
exhaust-velocity difference is only about x1.3, and because we're dealing
with exponential relationships, that's a vastly better case than x3.

Even ammonia, with an effective molecular weight of about 6 after it
decomposes, is still x1.7, which is a lot worse than x1.3.

>but what about tank volume? The density of water is 1 ton/m3,
>what about liquid hydrogen? I'm sure it's much lower than
>0.11 ton/m3 and then water seems to still be the winner, since
>you save a lot of tank mass.

In space, tank mass doesn't need to be very large.  Hydrogen's very low
temperature is a bigger headache:  you probably need active refrigeration
unless you're planning to use it all within days of departure.  Yes,
hydrogen's density is very low -- there is more hydrogen per liter in
water than in LH2! -- but with a nuclear engine, the difference between
hydrogen and anything else is so huge that it can overwhelm a lot of such
overheads.

>> conventional solid-core nuclear rocket using dense propellants has trouble
>> achieving better performance than a LOX/LH2 chemical rocket.
>
>What's your definition of performance?

Specific impulse, aka exhaust velocity -- propulsion per kilogram of
propellant.  That is, overwhelmingly, what matters for in-space propulsion.
Thrust is very much a secondary consideration, unless you're talking about
differences of whole orders of magnitude.

>> If you must use a non-hydrogen propellant, ammonia (NH3) is about the
>> best.  It will mostly decompose into nitrogen and hydrogen at rocket
>> temperatures, giving an effective molecular weight of about 6...
>
>What's the density for liquid ammonia?

About 0.6 at its boiling point (which is a bit below room temperature,
but not so far down that it qualifies as cryogenic).  Denser if you
chill it further; like water, it has a wide liquid range.

>And what is the availability of ammonia in the solar system?

In the outer solar system, probably excellent, as a contaminant in ice and
a minor component of atmospheres.  In the inner solar system, not so good
(except possibly from extinct comets masquerading as near-Earth
asteroids), although it can be made from nitrogen and hydrogen.

>One advantage of water is that it can be had on Mars (which has,
>as far as I know, frozen water at the poles)...

Mars definitely has quite a bit of ice at its north pole in particular,
and may have widespread permafrost deposits at medium latitudes.  It also
has a trace of water vapor in its atmosphere, although extracting that is
costly in energy.  Its atmosphere is about 3% nitrogen.

>and possibly also on Luna.

There is definitely hydrogen at the lunar poles, and if it's not already
water, converting it to water should be trivial.  There may be ammonia
there as well, given that the polar hydrogen deposits are probably from
comet impacts.

>Of course hydrogen beats water in availability, as it
>can be had on Mars (and Luna?) via electrolysis of water, and
>on Titan, and by skimming the Jovian planets (or Venus?).

Titan has no significant hydrogen in its atmosphere, which is mostly
nitrogen with some methane, but both water and ammonia probably exist as
solids on the surface.  There's no hydrogen on Venus except for traces of
water vapor and sulfuric acid in the atmosphere.

>What's Ammonia's availability rating, in the grand scheme of things?

Middling.  Not as good as water, but better than a lot of other things.
--
When failure is not an option, success  |  Henry Spencer   henry@spsystems.net
can get expensive.   -- Peter Stibrany  |      (aka henry@zoo.toronto.edu)

From: henry@spsystems.net (Henry Spencer)
Newsgroups: sci.space.tech
Subject: Re: Exhaust speed
Date: Thu, 2 Sep 1999 00:05:51 GMT

In article <7q4oac$jo@peabody.colorado.edu>,
Frank Crary <fcrary@rintintin.colorado.edu> wrote:
>...Neon, hot
>or cold, isn't an oxidizer. I'm not sure about hot nitrogen, but
>it definitely wouldn't be as bad as hot oxygen. A nitrogen/argon
>mix would also be interesting, if you wanted to use the rocket
>on Mars (where a nitrogen/argon mix is easier to get than anything
>containing hydrogen.)

Another interesting possibility for Mars is carbon monoxide, which isn't
found on Mars but isn't hard to make from CO2, which is.  (CO2 has the
same problem as water -- when it gets hot it's an oxidizing agent -- but
the oxygen in CO is much more tightly bound.)

A further possibility is the idea of a chemical afterburner for a nuclear
rocket, injecting LOX into hydrogen exhaust from the nuclear core.  It
does sacrifice Isp, but again, for that price it does give you more thrust.
--
The good old days                   |  Henry Spencer   henry@spsystems.net
weren't.                            |      (aka henry@zoo.toronto.edu)

From: henry@spsystems.net (Henry Spencer)
Newsgroups: sci.space.tech
Subject: Re: Exhaust speed
Date: Thu, 2 Sep 1999 00:12:31 GMT

I wrote:
>All unattractive, in practice.  Staging nuclear systems means either
>throwing away "hot" reactors (where?), or bringing them back and reusing
>them (awkward)...

I should note that there have been a few hair-raising concepts which did
in fact propose to stage "hot" nuclear engines.

In particular, I recently ran into an old IAF proceedings -- from 1961, I
think -- which had a paper giving a status report on the then-new NERVA
project.  It talked briefly about the RIFT concept, which would have been
a NERVA flight test circa 1967.  At that time, at least, RIFT was going to
be suborbital, dumping a spent nuclear engine into the Atlantic!  They did
talk about safety issues of the test, including launch failures, and the
discussion of what happens if it all works basically just said "it'll go
into in deep water and it looks like there's no real chance of any
significant amount of fission products going anywhere from the spent
engine".  Eesh.

(The drawing of the test concept suggests a nuclear S-IV on top of an S-I
first stage.  Impact of the upper stage was roughly mid-Atlantic.)
--
The good old days                   |  Henry Spencer   henry@spsystems.net
weren't.                            |      (aka henry@zoo.toronto.edu)

From: James A Davis <jimdavis2@primary.net>
Subject: Re: nuclear hybrid rocket
Newsgroups: sci.space.tech
Date: 07 Oct 1999

Henry Spencer wrote:

> >I believe that the data for temperature, pressure and dissociation are
> >well established.
> 
> Dissociation is not the issue; recombination time is the problem.  

Actually, recombination is a minor matter for nuclear thermal rockets,
what matters most is the amount of dissociation in the reactor and this
is determined by chamber pressure for a given temperature:

Isp - H2 Tc = 3000 K

pc, atm		Equilibrium 	Frozen
200		961		956		
20		986		970
2		1059		994

Isp - NH3 Tc = 3000 K

pc, atm		Equilibrium 	Frozen
200		467		466		
20		478		472
2		509		481

Dissociation is very slight for hydrogen except at very
(unrealistically) low pressures. Dissociation for ammonia on the other
hand is total (for the NH3, not the H2 and N2 products) at any pressure.
The performance differences between equilibrium flow (maximum
recombination, infinite reaction rate) and frozen flow (no
recombination, zero reaction rate) are minor except again at unrealistic
(low) chamber pressures.

Jim Davis