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

Just a few more thoughts, FYI.

"w.d.greene" wrote:
> "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.

Another note about aerospikes.

The only place that an aerospike has actually been tested on a large engine
is with a modifed J-2 powerpack. This is significant because the J-2 is a
gas generator cycle. In this case, the turbine discharge gas is used as the
supply for the base pressure and coolant flow. On the XRS-2200 testing,
achieving sufficient cooling was an ongoing and unresolved issue. The
solution, if taken to a flight system, would have been a significant hit in
performance down from typical J-2 levels.

I've never seen it fully explained how a staged-combustion cycle like the
SSME could be used with a aerospike since there is no such supply of turbine
discharge gas to use (it get funnelled back into the main chamber(s)). If
you do divert coolant flow to this area, then you immediately lose
performance and it starts to look a lot like a gas-generator cycle.

The vacuum specific impulse for a typical Lox/H2 gas generator is about 420
seconds. Even with an aerospike, the sea level impulse does not come close
to this (a common error in academic calculations) because the plume expand
optimally based upon altitude, not always to vacuum conditions.

The vacuum specific impulse for the SSME is currently about 452 seconds.

There is no design that would achieve a 454 second Isp from sea level up to
vacuum.


>> 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.

It's important to note here that in the U.S., we've never built and
demonstrated a large, flight-weight channel-wall nozzle. We've built some
prototypes but not yet the real thing.

> 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.

The only major engine development program other than SSME in the last thirty
years in this country was the RS-68 for the Delta IV vehicle. The RS-68 is
a workhorse engine with performance significantly lower than the SSME. This
was by design. The purpose of this engine was to demonstrate cheaper,
faster production. It's amazing what they can do in terms of turning these
things out. But again, these are fabrication and assembly things and not
performance issues.

>> 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.

And, my original point that the SSME is 2000 technology is still true simply
because there hasn't been a great deal of development work in this specific
area of large rockets.

>> 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.

And you couldn't get me anywhere nearly a H2O2/Kerosene system. This is
another one that looks good on paper but is extremely difficult to truly
implement. H2O2, hydrogen peroxide, taken to concentration levels necessary
to make it volume efficient and performance efficient as a propellant is
extremely unstable. The test guys love telling stories about this stuff
virtually spontaneously exploding.

The better solution is good old Lox/Kerosene, the propellant combination
used for the Saturn V first stage. This has always been a top-notch combo
for first stage applications. Lox/CH4 is another viable choice. I'd take
either one of these before H2O2.

> 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.

I don't know what papers you've read, but be aware that one of the most
common mistakes that I've seen in dozens of purported SSTO solutions is a
failure to take into account propellant residuals. There are simply
thousands of pounds of unusable propellants at the end of the mission. Some
of it is gaseous in the form of pressurants. Much of it is liquid. You
simply cannot allow the liquids to run dry or you will have a very bad day
(catastrophic engine failure within fractions of a second).

>> 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.

That's my point with regards to Shuttle. Nobody else has built and operated
a 1.5 stage vehicle. Therefore, nobody else has built and operated a
vehicle system with the level of outright performance of the Shuttle. It's
a Ferrari. The question is whether you want a Ferrari or whether you want a
BMW. Note that I didn't say pickup truck because that analogy is old and
totally misused. Any vehicle that gets you to orbit is a high performance
vehicle. It's a question of levels of high performance and the associated
costs versus benefits.

> > 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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