Rocket engine performance?

Does the amount of heat in the exhaust have any bearing on the isp and
amount of thrust the engine produces?

i.e. the heat produced by the SSME is only a few thousand degrees, and
the SSME is a low isp but high thrust engine, and the VASMIR is
expected to have a plasma exhaust temp of a million or so degrees, but
it is a high isp with a low thrust engine as the plasma emmited is a
tenious plasma, and is only a fraction as dense as te exhaust emmited
by the SSME.

I guess the ideal rocket engine for sublight travel around this/a
system with our existing tech knowledge is a high thrust VASIMR type
plasma engine.


Christopher
+++++++++++++++++++++++++
"There's a light at the end
of the tunnel" says the optimist.
"It's probably a train coming
stright at us" responds the pessimist.

(Christopher) writes:

Does the amount of heat in the exhaust have any bearing on the isp and
amount of thrust the engine produces?

Heat in a rocket exhaust is an indication that the engine is less than 100%
efficient (as all heat engines must be!).


i.e. the heat produced by the SSME is only a few thousand degrees,
and the SSME is a low isp but high thrust engine, and the VASMIR is
expected to have a plasma exhaust temp of a million or so degrees,
but it is a high isp with a low thrust engine as the plasma emmited
is a tenious plasma, and is only a fraction as dense as te exhaust
emmited by the SSME.

First of all, don't confuse COMBUSTION temperature with EXHAUST temperature.
As the exhaust expands out the nozzle, its temperature falls as the heat
energy of the exhaust gases is converted into bulk kinetic energy of the
exhaust gases. In principle, a VERY long nozzle could expand the exhaust
gas to arbitrarily low temperature and pressure in a vacuum, but it would
be very, VERY heavy, and eventually the exhaust molecules would get so far
apart that their mean free path becomes comparable to the nozzle diameter,
and they start traveling "ballistically" instead of "flowing like a fluid."
There is little point in extending the nozzle past this "ballistic flow"
point, and in fact, most nozzles are much shorter.

Second, as I've noted before, if the exhaust temperature is high, this is
simply an indication that the engine is INEFFICIENT. There is no DIRECT
relationship between thrust-to-mass, efficiency, and I_sp --- they are
three orthogonal figures of merit.


I guess the ideal rocket engine for sublight travel around this/a
system with our existing tech knowledge is a high thrust VASIMR type
plasma engine.

As I believe Ian Stirling has noted in another thread, the "ideal" engine
is one whose exhaust velocity is about 2/3 of the mission delta-vee.
Since V_ex = g * I_sp, this would imply that a truly "variable" I-sp engine
would seem to be desirable, but for the near term that engine is _NOT_ VASMIR,
since its MINIMUM exhaust velocity is _TOO LARGE_ for most of the near-term
missions being considered. Likewise, for far-term missions, its MAXIMUM
exhaust velocity is _TOO SMALL_ (only a minute fraction of lightspeed).

The moral to this story is, there is =NO= single technology that can be
all things for all applications --- and when one tries, one gets a bloated
monster like EMACS that =STILL= can't do everything... :-T


-- Gordon D. Pusch

perl -e '$_ = \n"; s/NO\.//; s/SPAM\.//; print;'

On 03 Aug 2003 17:55:44 -0500, (Gordon D.
Pusch) wrote:

(Christopher) writes:

Does the amount of heat in the exhaust have any bearing on the isp and
amount of thrust the engine produces?

Heat in a rocket exhaust is an indication that the engine is less than 100%
efficient (as all heat engines must be!).


i.e. the heat produced by the SSME is only a few thousand degrees,
and the SSME is a low isp but high thrust engine, and the VASMIR is
expected to have a plasma exhaust temp of a million or so degrees,
but it is a high isp with a low thrust engine as the plasma emmited
is a tenious plasma, and is only a fraction as dense as te exhaust
emmited by the SSME.

First of all, don't confuse COMBUSTION temperature with EXHAUST temperature.
As the exhaust expands out the nozzle, its temperature falls as the heat
energy of the exhaust gases is converted into bulk kinetic energy of the
exhaust gases. ...

snip

Right, thanks for the clarification, as always appreciated.




Christopher
+++++++++++++++++++++++++
"There's a light at the end
of the tunnel" says the optimist.
"It's probably a train coming
stright at us" responds the pessimist.

"Christopher" wrote:
Does the amount of heat in the exhaust have any bearing on the isp and
amount of thrust the engine produces?

i.e. the heat produced by the SSME is only a few thousand degrees, and
the SSME is a low isp but high thrust engine, and the VASMIR is
expected to have a plasma exhaust temp of a million or so degrees, but
it is a high isp with a low thrust engine as the plasma emmited is a
tenious plasma, and is only a fraction as dense as te exhaust emmited
by the SSME.

Very much so. Temperature is a measurement of the average
kinetic energy of the molecules in a gas (or a liquid or a
solid, but we're concerned with gasses right now). In fact,
they are related by the equation KE = 3/2 k T, where KE is
the average kinetic energy of a molecule in the gas, k is
Boltzmann's constant (1.38e-23 J/K), and T is the
temperature of the gas in Kelvin. Further, using the
equation KE = 1/2 m v^2, you can determine the average
velocity of a gas at a specific temperature. For example,
take CO2 at 1,000 K, the mass of the molecule is 14 amu
(2.32e-26 kg), the average kinetic energy per molecule is
2.07e-20 J, and so the average speed of the molecules is
around 1.3 km/s. Neat, huh? As you can see, doubling the
temperature raises the molecular speed by sqrt(2), and
halving the molecular mass has the same effect.


I guess the ideal rocket engine for sublight travel around this/a
system with our existing tech knowledge is a high thrust VASIMR type
plasma engine.

Not necessarily. VASIMR is a tricky beast, because it,
like all electric rockets, has separate propellant and
power systems. This is disadvantages compared to many
other high-thrust, high-ISP designs which have combined
power and propellant (i.e. the propellant generates its
own power to heat itself, as in chemical rockets) such
as NSWR, Orion, fusion rockets, etc. This separation
means that the rocket equation needs a little mending to
accurately describe the functioning of the propulsion
system. If it takes a certain amount of, for example,
fission fuel to create the energy needed to expell a
certain amount of propellant from the VASIMR rocket then
you can't just figure the mass of the propellant alone
into the equation. Fortunately, nuclear reactions are
very energetic so they hardly don't need accounting for
mass usage, but not entirely, and there are other
concerns as well, such as the mass overhead of the
entire reactor system needed to power the rocket.


Personally, I don't see VASIMR as being capable of
providing the performance necessary for interstellar
travel in reasonable (sub millenial) travel times and I
think especially fusion, NSWR, and Orion are the best
bets there.