Re: Google Cindy Sheehan (was: Re: Peace Mom)


"Josh Hill" wrote:
> "w.d.greene" wrote:
>>"Josh Hill" wrote:
>>> "w.d.greene" wrote:
>>>>"Josh Hill" wrote:
>>>
>>>>> Estimated cost of building the DC-Y single stage to take off vehicle
>>>>> -- a reusable spaceship that would slash the cost of putting stuff in
>>>>> orbit -- $6 billion.
>>>>
>>>>Um, no.
>>>>
>>>>Single-stage to orbit (SSTO) does not work. It cannot work. Any
>>>>fledgling
>>>>aerospace engineer with just a passing knowledge of the rocket equation
>>>>understands this. Space Shuttle is 1.5 stages and it can only pull it
>>>>off
>>>>by being the highest performing system ever built.
>>>
>>> I'm not sure I understand why you say that. The Shuttle was
>>> state-of-the art in the 70's, but it's hardly that now -- both engine
>>> design and structural materials have improved, with further
>>> improvements on the horizon.
>>
>>This is a common misunderstanding that is unforutnately repeated over and
>>over in the mass media. With regards to the engines, the space shuttle
>>main
>>engines are the highest performance, high thrust engines ever built.
>>Could
>>they be higher in performance? Yes. Theoretically, you might be able to
>>squeeze out another one or two percent in specific impulse. But that's
>>it.
>>Why? It has nothing whatsoever to do with 1970's technology or 2010
>>technology or even 2000 BCE technology. It has to do with the fundamental
>>properties of hydrogen and oxygen and the rest of the periodic table.
>
> You're oversimplifying, I think. According to what I've been reading
> (which is a fair amount since I logged on last), the SSME's were
> originally supposed to have had an expanding nozzle, but that was
> scrapped as part of a cost cutting move. So their specific impulse
> drops from 454 to 368 secs at sea level. An expanding nozzle or better
> yet an aerospike engine would be more efficient.

The SSME has an expanding nozzle (currently 69:1 expansion ratio). All
rockets have expanding nozzles. Or, more accurately, most have
converging-diverging nozzles.

The SSMEs have a standard bell nozzle. This was one of the few things that
was mandated in the original RFP with regards to "how." The reason that an
aerospike was not purused was because it was relatively low in technology
readiness. In fact, even today, its technology readiness is still low with
regards to flight-weight systems (as demonstrated within the last five years
by the testing of the XRS-2200 engines).

The addition of a usable aerospike engine would increase mission-average
specific impulse because it would raise the values for non-vacuum
conditions.

> Also, the sea-level F/W ratio of the Block II SSME's is 51; the
> shortened-nozzle Block II+ was (is?) supposed to increase that to 58,
> and the Block III to 70 by incorporating a channel-wall nozzle, jet
> boost pumps, electrical valves, an extra-large throat combustion
> chamber, and a new controller. That's significantly closer to the
> 75-80 F/W ratio required by an SSTO.

The Block 2 SSME is the currently flying SSME.

The addition of a channel-wall nozzle increases the weight of the engine for
a given expansion ratio. It does not decrease it.

Also, the extra large throat had nothing to do with thrust-to-weight. It
has to do with reliabilty.

Jet boost pumps are at a very low technology readiness. Weight savings?
Not sufficiently demonstrated to say.

Electro-mechanical valves are a legitimate option, but one currently fraught
with issues with regards to retrofitting. The XRS-2200 had some issues with
these but eventually overcame most of them.

Also, I'm a bit confused by the shortened nozzle comment. A shortened
nozzle reduces performance with regards to specific impulse.

> And then there are full-flow engines,

Full-flow refers to the engine cycle. It has no effect on performance,
i.e., Isp.

> hydrostatic bearings,

Reliability and life improvement that really only impact reusable engines.
However, the silicon nitride ball bearings currently used are pretty damn
good.

> thrust
> vectoring using carbon vanes or engine throttling as in the X-33,

X-33 did not use (or intend to use) vectoring blades.

> improvements in the manufacturing process (""Production costs of the
> current engines are also high because of their complexity, including
> the large number of parts needed and the manufacturing technology that
> was available when the SSME was developed") which don't affect
> performance but do affect cost and lead time,

Yes, improvements are possible in the area of fabrication. However, this is
only a really big issue for an expendable engine. As you said, not a
performance issue.

> and tripropellant
> engines.

Nice thought, but never been done other than on paper.

Note at this point that what I said regarding SSME performance still holds.
Nothing that you've mentioned, other than the aerospike nozzle, has any
effect on specific impulse, i.e., engine performance. As I said, that is
because engine performance in terms of specific impulse is largely a matter
of chemistry.

> So it seems to me that as things now stand, there's lots of room for
> improvements in engines for a new design, and some room for
> improvements in engines for the current one.

There's always room for improvement in terms of cost and reliability, but
not in performance. There might be a little in terms of weight,
particularly if you're taking about expendable applications, but not loads
and loads. Certainly not as much as you've claimed without serious
sacrifices in safety. Note that the original SSMEs and those that flew for
years, the Phase II engines, had higher thrust-to-weight, but had only about
one-fourth the reliabilty (and therefore safety) of the current
configuration engines.

>>The solid boosters could be of higher performance, but again probably only
>>marginally without radical changes (like maybe going to liquid propellant
>>boosters).
>
> Sure, but those solid fuel boosters were a serious compromise. One
> article I read claims that replacing the solid fuel boosters with
> recoverable H2O2/Kerosene boosters would increase payload mass by
> almost a third, from 24,950 kg to 33,140 kg.

H2O2/Kerosene are not solid propellants.

Yes, as I said, going to liquid boosters would boost performance. Everyone
recognized this from the beginning. As you said, it was a compromise.

Also, the current RSRMs are recoverable.

>>Yes, there have been materials improvements. However, some of them have
>>already been incorporated into the orbiter, the external tank, the
>>engines,
>>and the solid motors. See, these things are regularly and routinely
>>updated
>>and upgraded.
>
>>The Shuttle is, in fact, in many ways 2000's technology with only a few
>>exceptions here and there.
>
> Seems to me there are more than a few exceptions. The ceramic tiles,
> the aluminum frame, the fuel cells, the hydrazine-powered APU's, for
> example.

The tiles are an area for further improvement, unquestionably. However,
this is a reuse issue and not so much a performance issue.

Also, there's nothing wrong the APUs other than the fact that nobody wants
to deal with hydrazine. It's a green issue, not so much a performance
issue.

> And then there were design decisions that looked good at the
> time, such as the side-mounted fuel tank, that were in retrospect
> mistakes, and in a sense, those represent 70's technology too, because
> we didn't know better and now we do. As well as political/cost
> compromises, e.g., the SRB's.

The side-mount was essentially a requirement imposed by the DoD sticking
their nose into the process and insisting on cross-range capability. That's
not a matter of 1970's technology so much as requirements creep which is,
unfortunately, a timeless issue on every large project.

> Which isn't to say that we could do all that much better with the
> original cost constraints . . .
>
>>> I found a few papers that rough out SSTO proposals, including this
>>> one:
>>>
>>> http://www.ssdl.gatech.edu/main/ssdl_paper_archive/iaf-st-87-07.pdf
>>>
>>> Short on details and I don't know enough to fill them in, but I don't
>>> see any assumptions there that seem outlandish, e.g., they posit a 10%
>>> mass reduction for the vehicle.
>>
>>In order for SSTO to work, the propellant mass fraction has to be on the
>>order of 91% to 93% of the net liftoff weight (post engine ignition and
>>hold-down). That doesn't leave much for the rest of the vehicle as in
>>tanks
>>and structure and engines and wings and landing gear and, oh yeah,
>>payload.
>
> The figures I've seen are 89-90% but either way, it's a challenge.

Monumental challenge.

> Judging by what I've been reading, current materials technology may
> not be quite up to the task. But then, how will we get there if we
> don't do the research, linear aerospike engines, collapsing fuel
> tanks, and all?

Research is fine. What was being touted (DC-X/X-33 followed by RLV) was a
full development program, which is an entirely different thing. It was a
great big and expensive boondoggle based on unproven technology that nobody
believed would work (nobody with any experience with actual hardware that
is -- academics, like Dr. Olds, can convince themselves of anything).

>>Yes, the assumptions are outlandish once they're stacked up one on top of
>>another. It's interesting to note that the gross liftoff weight is
>>roughly
>>the same as Shuttle and yet it delivers less half of the payload.
>
> Well, that's to be expected from an SSTO -- no one has ever suggested
> that staging wasn't efficient in that regard. But the extra fuel costs
> a lot less than prepping a Shuttle or throwing away the equivalent of
> a 747 with every launch.

Actually, it would be cheaper to throw away more per launch. That is the
one lesson that Shuttle has talk us. In future reusable systems, if they're
ever developed again, will have to truly reusable and not Shuttle-like. Of
course the problem with that notion is that getting to orbit will always be
difficult. Thinking of any space vehicle as "like an airliner" has implicit
dangers when the task is so difficult.


.

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