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

On Sat, 27 Aug 2005 07:37:26 -0500, "w.d.greene" <bil64@xxxxxxxxxxx>
wrote:

>
>"Josh Hill" wrote:
>> "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.

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.

Didn't know that.

>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.

Isn't the SSME nozzle optimized as much as possible for vacuum
performance? In which case, the nozzle wouldn't grow much longer than
it is now.

>>>> 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.

Isn't differential throttling more weight efficient?

>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.

Do annular aerospike designs have similar accessibility problems?

>> 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.

But is that so? An S-IVB has a mass fraction of .905. Using this as a
baseline, and borrowing the higher figure from your other post for the
weight of the recovery system, we have about a 10% shortfall -- make
that 20% to allow for payload mass, the TPS, thrusters and avionics.
Now use modern materials. Say they save a conservative 5%, which is
lower than the figures I've seen. We're still 15% too heavy. But
Salked's calculations suggest a 20% reduction in dry mass, which
should be enough even if his figure is 5% too high. Assuming, of
course, that his figures aren't overoptimistic.

>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.

It seems that Rockwell's RLV was supposed to use supercooled
propellants: "This provides a 10 percent volumetric savings with the
LOX tanks and a 6?7 percent volumetric savings with the LH2 tank. This
technology would be demonstrated with the SSME in the X-33 program."

>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.

Clever.

Since most of the thrust is needed early on, I wonder if you couldn't
get sufficient performance by combining sea-level optimized
kerosene/LOX engines with hydrogen/LOX engines.

>>>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.

Sure. But the tradeoff seems to yield a net gain, although from what
I've read the magnitude of the gain is very dependent on assumptions.

>> 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.

Some of them, apparently.

>> 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.

Which, however, doesn't mean that they don't save weight when used in
the places they /do/ work.

I'm seeing some very polarized claims here -- at one extreme, there
are claims that lop 35% off dry weight and on the other, the
implication that they don't produce any benefits at all.

>> 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.

Oh, I don't know. The return, presumably -- I think he wanted to use
jet engines and wings to land, don't know what he was going to do for
TPS. But I don't see any reason why the expendable version wouldn't
have made orbit, and I doubt that the technical director of the
program that designed the S-IVB would have spouted total nonsense.

>> 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.

However, the sense I have from what I've read is that a lot of
nonsense has been spouted on /both/ sides.

On one hand, the manufacturers and SSTO advocates made some wildly
overoptimistic predictions.

On the other, I hear a lot of people saying that "this can't possibly
be done." They point to the rocket equation as if it presented some
kind of fundamental limit, rather than an engineering challenge. Sure,
it can't be done if expectations are unrealistic, whether on the part
of the advocates or NASA. But by the same token, it can't be done if
you rule out every technology, whether it's new structural materials
or improved engine design.

It seems to me that it's as much about having realistic expectations,
KISS, and a realistic, incremental R&D approach as it is about
anything else. You're going to fail if you try to build a DC-10 rather
than a DC-3, and, conversely, you're never going to accomplish
anything if you don't go for the DC-3.

>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.

Don't know if it's correct, but according to what I've read, the fuel
weight penalty for vertical landing is about the same as the wing and
control weight penalty for horizontal landing:

"However, the McDonnell Douglas X-33 concept and DC-X
flight demonstration vehicles are VTVL designs. These vehicles have
ballistic missile aerodynamic characteristics and no wing structures,
providing an advantage during ascent because there are no parasitic
drag losses due to wings. However, this type of vehicle design can
result in high reentry speeds and high aeroshell heating rates during
reentry. This may lead to the disadvantage of greater thermal
protection requirements on reentry and increased vehicle mass
for the vehicle thermal protection system (TPS). Landing is
accomplished by restarting and firing the main engines. This increases
total mission propellant requirements, but results in reduced
structural weight because wings and related structures are not needed.
An increase of approximately 1000 ft/sec in ideal velocity is needed
for vertical powered landing. Studies have indicated that there is no
overwhelming advantage or difference in overall vehicle weight (GLOW)
between vehicles using horizontal and vertical landing modes. However,
there are increased risks of mission failure with vertical landing
systems because of requirements for main engine restart, the high
thrust levels potentially needed, and precise thrust vector control
needed at landing and after reentry and exposure to the space
environment."

http://64.233.161.104/search?q=cache:aW7UZY9AQd0J:www.rand.org/publications/MR/MR890/MR890.chap3.pdf+%22needed+at+landing+and+after+reentry+and+exposure+to+the+space%22&hl=en&lr=lang_en

>>>> 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.

But they bit off way more than they could chew. Let's develop linear
aerospike engines -- 1 year delay. Let's develop multi-lobed composite
LH2 tanks -- forget it; from what I've read even a plain composite LH2
tank is pushing things. Let's use a lifting body -- oops, the thing
isn't stable at transonic velocities and we need multilobed tanks. And
this in addition to metallic TPS and the ultralight composite
structure. Oh, and while we're at it, we'll give it a shuttle-sized
payload bay and a 49,000 lb. to LEO payload capacity. And we're going
to go directly from a costly, oversized X vehicle to a production run
of 2 or 3 economical short-lived vehicles, rather than going for a
full-sized Y vehicle or taking an even more cautious approach. All in
no time and with a tiny budget, natch.

It was NASP Lite.

>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."

In what sense is 40,000 pounds reasonable? You aren't going to get
enough 40,000 pound never mind the Venture Star's 49,000 pound
missions to make an SSTO even remotely economical.

The only conceivable reasons to build an SSTO are to save money on
launches or develop the technology that will allow you to do so in the
future as new technology comes on line. Conventional expendable or
semi-reusable vehicles make more economic sense if you're lofting the
occasional > 20,000 pound space station module or satellite.

>> 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?

No, but I just read G. Harry Stine's book on the origins of the
program. Delta Clipper was hatched in a living room by a bunch of
enthusiastic engineers.

>>>>>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.

--
Josh

"You know I could run for governor but I'm basically
a media creation. I've never done anything. I've
worked for my dad. I worked in the oil business. But
that's not the kind of profile you have to have
to get elected to public office." - George W. Bush
.