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


"Josh Hill" wrote:
> "w.d.greene" wrote:
>
>>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.
>
> Any idea how far?

At this point, no. It was left as a lesson learned and a problem to be
resolved should the program ever be re-started. It may be that there is
more than one way to solve this issue and maybe, just maybe, the impact to
the engine performance could be minimized.

>>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.
>
> AFAIK that hasn't been proposed -- the RS-2200 was supposed to use a
> gas-generator cycle and the altitude-compensated SSME proposals I've
> seen mentioned use a passive plug nozzle or an expanding bell.

I heard suggestions of it, the staged-combustion aerospike, but I haven't
heard any discussion regarding how to overcome the cooling issue.

> I've seen it claimed that the aerospike's high expansion ratios make
> up for the lower performance of the gas-generator cycle. Also,
> aerospike engines are small and light and have an extra margin of
> safety and reliability.
>
> I wonder if you could make an aerospike with a bleed expander cycle?

Sure, it would be possible, but you'd still need to have something to cool
and pressurize the base. Any propellant that doesn't flow through the
throat represents a performance loss. That in a nutshell is the difference
in performance between gas-generator cycle and staged-combustion cycle, for
example.

>>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.
>
> The RS-2200 was apparently supposed to have a sea-level Isp of 347 sec
> and a vacuum ISP of 455 sec. Which I confess I don't understand, since
> the SSME is 363/453 -- where's the aerospike's altitude compensation?

I don't know. It's been a few years since I looked into this. Calculating
the effective expansion ratio for an aerospike as a function of altitude was
one issue for which I never came to straightforward solution short of CFD.

>>>> 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.
>
> I guess it depends on how you define 2000 technology, on whether you
> refer to what we have sitting on the shelf, e.g., SSME Block II, or
> what you would expect to develop if you were developing a new engine.
> The original SSME took a lot of R&D and really pushed the envelope
> back in the 70's: turbopumps were blowing up left and right. If we
> were developing an innovative vehicle today, I assume we'd also push
> the envelope as in the RS-2100 or RS-2200.

Yes, but as I mentioned elsewhere, the Block 2 SSME is not the same beast as
the original engine. Both high-pressure turbopumps have been replaced. The
MCC was replaced. The powerhead was replaced. The low-pressure pumps have
not been replaced wholesale, but they have been incrementally improved (and
they're not a amjor source of risk anyway). Only the nozzle stands out as a
near-original part that is still, truly 70's technology.

>>>> 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.
>
> I've read that.
>
>>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.
>
> The authors of this paper suggested H2O2/kerosene for a retrofit
> because the boosters would be about the size and mass of the SRB's.
> But I think you're right -- this was just another pie-in-the-sky
> proposal.

No, it's not pie-in-the-sky. It's a legitimate proposal that's been put
forward on several occasions. It would be a substantial and expensive
development program, but it is certainly feasible. As with everything like
this, though, it comes down to a cost-benefit analysis. Space hardware,
particularly human-rated space hardware, is extremely expensive.

>>> 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).
>
> Interesting. Serious proposals or pie-in-the-sky papers? I'm just
> beginning to appreciate how many of those there are.
>
>>>> 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.
>
> Weren't the early Atlas vehicles 1.5 stages? I think they dropped two
> of their engines in mid-flight -- it seems that when they were
> designed, no one knew if you could start an engine in a vacuum, so
> they decided to take the safe approach and start all the engines on
> the pad.

Yep, you're right. My mistake. Good memory. The original Atlas vehicle
had a booster-sustainer arrangement of engines. It also had balloon tanks
that would collapse unless pressurized. Performance was pushed to the
limits to make it work.

Interestingly, back in the early 1990's when we were looking at NLS
(National Launch System), one of the ideas that was thrown about was
something similar to this where the engines were put into pods and staged by
dropping these pods. Engines are heavy. If you don't need them, drop them.
On the other hand, engines are very expensive so by dropping them you're
throwing away tens of millions of dollars each. Again, it's another of
those trades that you have to do to demonstrate a feasible solution,
feasible both technically and from a business point of view.

>>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.
>
> I suspect that SSTO will initially take Ferarri technology, but
> produce Volkswagen results -- but that later versions would run on VW
> technology.

Here, I'll give you some hope: Combine-cycle engines. If you could build
an engine that functioned like a ram/scram jet in the atmosphere and as a
rocket once you're out of the atmosphere, then you can start talking about
specific impulse values over 1000 seconds. You can also avoid carrying all
of you're oxidizer with you from the ground. Thus, this starts to look
plausible as SSTO from a rocket equation standpoint. Of course the hitch is
that combine-cycle engines are still in their infancy with some significant
technical issues to overcome.


.

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