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:
>>>>> "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.
>
> I'm talking about a nozzle with two bells, one of which slides over
> the other to provide attitude compensation:
>
> "The Space Shuttle Main Engine was originally designed to use an
> expanding nozzle. Since these engines are fired from takeoff right
> into space, any sort of altitude compensation could dramatically
> improve the overall performance of the engines. This feature was later
> abandoned in one of the many cost-cutting phases of the Shuttle
> design, and today the SSME suffers a 1/4 drop in performance at low
> altitude as a result."
>
> http://fixedreference.org/en/20040424/wikipedia/Expanding_nozzle

This 1/4 drop in performance is again making that same mistake that I
mentioned in my discussion regarding aerospike nozzles. It assumes that
vacuum conditions are the same as optimal expansion. This is not true.

> Apparently, the Russians used an expanding nozzle on the unfinished
> RD-701. McDonel Douglas also planned to use it on half of the engines
> in the DC-Y because of the risk and cost of developing a plug nozzle
> design.

A deployable nozzle extension was developed and used for one of the
incarnations of the RL10 (from Pratt & Whitney). The RL10 is about 1/30th
the thrust level of the SSME.

Such a system is typically very heavy. When it's put into the trade space
it is typically dropped off very quickly in all cases except for in-space
applications for this reason. And for this reason, and others, it would
never have been a serious player in the trades for Shuttle. The other
reason has to do with geometry. The current bells come within inches of
smacking each other when they gimbal. Add a nozzle extension and you lose
your control authority because the engines can no longer gimbal. To
overcome this, you'd have to expand the size of the boattail. This is a
weight a aerodynamics issue.

>>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.
>
> I know.
>
>>The addition of a channel-wall nozzle increases the weight of the engine
>>for
>>a given expansion ratio. It does not decrease it.
>
> And that I didn't know. But I've seen the Block III nozzle referred to
> as a lightweight nozzle:
>
> "Technology programs are also planned to improve the SSME engine. Two
> improved versions, Block II+ and Block III, are planned. Block II+
> incorporates a 57:1 shortened nozzle; the Block III engine features a
> lightweight nozzle, jet boost pumps, electrical valves, a new
> combustion chamber, and a new controller. The Block II+ engine could
> support the X-33, but Block III would be available only for the RLV.
> The goal for Block II+ is a sea-level F/W of 58; the target for Block
> III is a F/W value near 70."
>
> http://books.nap.edu/openbook/0309054370/html/63.html

A shortened nozzle does not represent a performance increase. It represents
a performance decrease, but one that you might consider for considerations
of weight and assuming that you were using the engine in a different
capacity such as a, say, a first stage of a two-stage vehicle.

All practical channel-wall nozzle designs that I've seen (or laid out
myself) are hundreds of pounds heavier than the current tube-wall nozzle.
Period. The nozzle is the only major component of the SSME that has not
been completely overhauled and redesigned over the last thirty years. It is
an astonishing marvel of craftsmanship that results in a lightweight,
high-performance product. It is, however, also the component that is the
biggest headache on the project simply because all that performance and
craftsmanship means that they are difficult to build, expensive, and no
nearly as durable as they need to be. Proposals for building channel-wall
nozzles are based not on performance increases, but on producibility and
durability arguments.

Also note that the new main combustion chamber proposals are in the similar
vein. While the MCC was upgraded in the 90's, there does exist some
potential for further improvements in producibility (and therefore cost
reduction).

BTW, I was just discussing jet pumps for a different application and our
resident turbomachinery expert was explaining that they're typically very
low in efficiency and very limited in applicable range. So, they're big,
heavy, and probably couldn't work in this environment.

> Also:
>
> "SSME Block III Study -- This study would investigate the
> incorporation of an Extra Large Throat Combustion Chamber and a more
> robust channel wall constructed nozzle, increasing performance margins
> and abort thrust capability and eliminating main combustion chamber
> and nozzle failure modes. However, due to technical issues, this study
> is being discontinued."
>
> http://www.google.com/search?sourceid=navclient&ie=UTF-8&rls=GGLG,GGLG:2005-33,GGLG:en&q=Extra+Large+Throat+Combustion+Chamber+and+a+more+robust+channel+wall
>
> I don't know how this was all supposed to fit together, or what the
> "technical issues" were.

Again, this is all reliability, producibility, durability, and reuse stuff.
All good, but not performance based. Increasing performance margins in this
context means that you've actually backed off on performance for the sake of
safety.


>>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.
>
> I gather they intend to increase the sea-level F/W ratio at the cost
> of some vacuum Isp. Haven't come across anything that explains why
> they want to do this, though.

It wouldn't be for Shuttle applications.


>>> 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.
>
> Didn't mean to imply it had -- it used engine throttling. The carbon
> vanes are from a shuttle modification proposal I came across.

It used differential engine throttling because the notion of gimballing a
linear aerospike, while it can be done (and was demonstrated in the 1970's),
is difficult.

Some interesting notes about the XRS-2200, the recently tested linear
aerospike, includes the fact that it was based on good old, reconstituted
J-2 hardware. Thus is was, at its heart, "1960's technology." Also, it was
artocious in terms of reusibility considerations in that the hardware was
inaccessible and the inspections required between firings were onerous.
Part of this was due to fitting old piece parts into a new application.
Part of it, though, is inherent in the design and something that would have
to be considered if ever it comes up as a viable option for a future
vehicle.

>>> 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.
>
> I'm thinking about cost here.
>
>>> and tripropellant
>>> engines.
>>
>>Nice thought, but never been done other than on paper.
>
> It seems that the Russians bench tested a prototype of a tripropellant
> engine:
>
> "Although invented in the US, the only tripropellant engines built
> were in Russia. Kosberg and Glushko developed a number of experimental
> engines in the early 1990s for a SSTO spaceplane called MAKS, but both
> the engines and MAKS were later cancelled due to a lack of money.
> Glushko's RD-701 was built and test fired, however, and although there
> were some problems, Energomash feels that the problems are entirely
> solvable and that the design does represent one way to reduce launch
> costs by about 10 times."
>
> http://en.wikipedia.org/wiki/Tripropellant_rocket
>
> Specs: Thrust (each chamber, vac): 1960 kN mode 1 and 785 kN mode 2
> (throttle 40-100%). Isp: mode 1: 415 sec and 330 sea level. mode 2:
> 460 sec.
>
> Apparently NASA planned to use the RD-701in the X-2000.
>
> From the same source:
>
> "[Salked] concluded that tripropellant engines would produce gains of
> over 100% in payload fraction, reductions of over 65% in propellant
> volume and >20% in dry weight. A second design series studied the
> replacement of the Shuttles SRBs with tripropellant based boosters, in
> which case the engine almost halved the overall weight of the
> designs.."
>
> Also,
>
> "Use of tripropellant propulsion would reduce the weight of the
> vehicle from (249.475t dry / 2494.75t gross) to approximately (170.55t
> / 2267.961t). The volume would be reduced as well, since the average
> density of the oxygen/hydrogen/kerosene propellant combination is
> 880kg/m3 vs. 352kg/m3 for oxygen+hydrogen."
>
> "Consequently, Salkeld's baseline vehicle from 1973 used a single type
> of dual fuel high pressure engine operating at the same chamber
> pressure as the SSME. The sustainer engines would have burned a
> mixture of kerosene, hydrogen & oxygen during liftoff before switching
> to LH2 & LOX as the vehicle gets lighter. Salkeld claimed the gravity
> losses are reduced by up to 150m/s as a result of more rapid T/W
> buildup during ascent to orbit, if tripropellant propulsion is used.
> The total weight savings vs. LOX/LH2 SSTOs could be as great as 25%
> and the reduction in dry mass & size also reduces the sensitivity to
> weight growth. "
>
> "He concluded that mixed mode propulsion would produce gains of over
> 100% in payload:dry mass ratio for VTVL SSTOs and reductions of over
> 65% in propellant volume and >20% in dry weight."
>
> http://www.abo.fi/~mlindroo/SpaceLVs/Slides/sld039.htm
>
> They clearly have a ways to go, though I gather you could get some of
> the benefits by combining conventional hydrogen and kerosene engines.

Still, in this country, only paper studies.

Tri-propellant engines could have their place if an when we could get
structural mass down to a point low enough to take advantage of their
benefits. We're not even close right now.

Note that the notion of using a lower performing, higher thrust propellant
at low altitude is the foundation of Shuttle propulsion system, Saturn V,
and, for that matter, Saturn I. Basically you give up Isp for the
volumetric efficiency of kerosene or methane or solid propellant. The day
that somebody invents 50 lb/ft3 hydrogen is the day that all of this gets
easier.

Note that one creative idea I saw once in a published paper was the notion
of floating methane crystals in suspension within hydrogen. This
effectively gives you a tri-propellant issues yet avoids the hardware tax of
requiring parallel engine cycles.

>>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.
>
> The aerospike nozzle and, it would seem, tripropellant engines, both
> -- not that the latter would be practical in the existing orbiter even
> if they were available.

The tri-propellant engine lowers average engine performance though it may
have vehicle/stage efficiencies. This, again, is a trade.

> However, I gather that a better F/W ratio
> would be important to SSTO designs, as would any applicable
> reliability/safety/maintenance improvements.

In order to even approach the numbers necessary for SSTO, safety margins get
tossed out the window.

>>> 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.
>
> Of course not. I was comparing the SRB's with one proposal for
> liquid-fueled replacements -- the authors chose peroxide because the
> engines would have about the same mass and size as the SRB's.
>
>>Yes, as I said, going to liquid boosters would boost performance.
>>Everyone
>>recognized this from the beginning. As you said, it was a compromise.
>
> Or IIRC Sen. Garn pushing the booster proposal NASA liked least
> because Thiokol was in his home state?
>
>>Also, the current RSRMs are recoverable.
>
> Of course.
>
>>>>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.
>
> I thought it was a cost and safety issue?
>
> "Hydrazine's handling requirements are a significant contributor to
> the costs and time for preparing the shuttle for flight. Specially
> trained personnel and special equipment are needed because of the
> volatility, toxicity, and the caustic nature of hydrazine. Replacement
> of hydrazine with a nontoxic fuel would significantly reduce shuttle
> operation costs while enabling increased flight rates."
>
> http://www.grc.nasa.gov/WWW/RT1997/6000/6920viterna.htm

Yeah, as I said, a green issue. Anything not green is a safety issue and
therefore a cost issue. They go together. However, the performance of the
APUs is fine.

>>> 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 just read, though, that there are already at least 16 launch
> vehicles/launch vehicle stages that have a mass fraction greater than
> .90, most or all designed before the availability of lightweight
> composites.

"Lightweight composites" were to be the saviour of everything. Basically
they suck for many applications and, in some, are actually heavier.

> Apparently Philip Bono of McDonnell Douglas calculated
> that the S-IVB could actually put a Gemini capsule into orbit and
> designed a version that allowed for a powered return.

Sorry, but that's absolutely ridiculous.

> So -- what could we do with composite materials and slightly improved
> engines? I'm having trouble seeing how, even with the addition of
> landing struts and a TPS, SSTO isn't doable. I don't mean to minimize
> the R&D effort and risks involved, but the Delta Clipper team claimed
> that even the 40,000 lb. capacity DC-Y design allowed for 15% weight
> growth, 20% with a reasonable reduction in payload. Not the bloated,
> impractically cutting-edge X-33 at all.

Josh, I really don't have the time or energy to dissuade you of this notion
that you've picked up from multiple sales pitches. Every single study that
I've ever seen on the subject starts to fall apart as soon as you look at
the details. Also, the Delta Clipper was probably the most stupid proposal
of the bunch. Think about the propellant necessary for your landing burn.
For the purposes of ascent phase, that mass gets counted as unusable
propellant.

>>> 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).
>
> It seems to me that the X-33 and DC-X/Y were very different programs.
> The X-33 made too many questionable assumptions -- weight growth of
> 5%, irregularly-shaped composite tanks in a weird self-supporting
> arrangement, linear aerospike engines, etc. -- and set it out to build
> a large-scale prototype right off the bat. A perfect example of
> bloatware.

DC-X was the same. They went off, built a prototype, and proved that it
could go up a few hundred feet and then come down again. Based on that,
they thought they could get to orbit.

Note that all of the "questionable assumptions" and unusual design solutions
within X-33 were attempts to overcome is obvious fudges builts into the
Delta Clipper paper design. Even with these, once you started thinking
about real hardware rather than just paper studies, all of your margins
quickly disappeared. The follow-on to X-33, RLV, went through design cycle
after design cycle, squeezing every drop of performance out of the
propulsion system and out of the structure, but in the end they could never
yield any positive payload without adding a second stage.

> The Delta Clipper program, OTOH, originally had a very different
> philosophy -- use as many proven off-the-shelf components as possible,
> test the concepts in the field with low-cost experimental vehicles,
> set reasonable goals, e.g., 20,000 lbs to LEO (though government bloat
> pushed that up to 40,000),

"Government bloat"? That should read, "reasonable and useful mission
requirements."

> work up to a full-scale prototype rather
> than making the "Y" vehicle the production vehicle as so often
> happens. It seems to me that it's the program we should have gone
> with, because as far as I can tell, it wasn't a jobs program, but a
> realistic, incremental attempt to build a reusable SSTO.

Nope, it was technically completely unsound. The assumptions were
outrageous and fudged here, there, and everywhere. The Delta Clipper
proposal was a disaster, IMHO, and properly dismissed.

I'm also a bit intrigued by the "as far as I can tell" portion of your
statement. Were you involved with it somewhere along the say ten or twelve
years ago?

>>>>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.
>
> But then, the same thing could have been said of airliners before the
> DC-3 . . . I don't think any of the proposals I've seen would actually
> turn out to be a DC-3 or a 707, but I like to think that they're steps
> along the way, or at least the best ones are, like the original DC-Y.

I wish that you were right. The Delta Clipper / DC-Y / DC-X was an absurd
idea pushed by a company desperate to stay alive propped up by congressmen
with no technical knowledge whatsoever.


.

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