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

On Mon, 29 Aug 2005 21:45:50 -0500, "w.d.greene" <bil64@xxxxxxxxxxx>
wrote:

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

>> 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.
>
>It is optimized for altitude performance, but I do not remember the
>altitude. It is over-expanded at sea level, which is why we need a diffuser
>on the test stand for lower power level operation in testing.
>
>>>>>> 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?
>
>Yes and no. Sure, with differential throttling you can get rid of the
>mechanisms for gimballing the engines. However, you pay the price with the
>XRS-2200 by requiring twenty separate combustion chambers. Of course, you
>could differentially throttle the different engines of a cluster, but you'd
>have to evaluate the control authority that such a scheme gave you. In
>theory it ought to be possible.

>>>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 all depends upon how its put together. One of the nice things in terms
>of vehicle design is that the guts of the powerpack for these kinds of
>engines can be sunk within the nozzle. This reduces the engine length and
>helps you some with the vehicle profile. But nothing is free.
>
>>>> 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.
>
>Remember with a tri-propellant engine your average Isp is lower so your
>propellant mass fraction has to increase. The tri-propellant engine can
>help you because you can build a more compact vehicle but you pay for that
>with lower performance. Again, a trade.

I don't think I fully understand the tradeoffs. On one hand, kerosene
engines have a higher thrust/weight ratio. Also, kerosene means you
have lower drag losses, lower gravity losses, and less massive fuel
tanks, even with the extra bulkhead. And all of these should ripple
through to the TPS and recovery systems.

Against that, you have a lower average specific impulse, some extra
plumbing, and, in a tripropellant system, the mass of the extra
turbines and perhaps some practical performance compromises compared
to a bipropellant engine.

That being said, why switch propellants in an SSTO? The only reasons I
can think of are the advantage of high thrust at takeoff, which should
lower your engine mass somewhat even with the extra turbines, and the
ability to tune the plume to altitude to some unknown extent by
changing the propellant mix on the way up. Am I missing something?

>> 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."
>
>That was also proposed for the Lockheed Martin RLV. In fact, I worked on
>that subject matter specifically, propellant densification, for two and a
>half years. That's where I got my first patent. It was to be a flight
>experiment on the X-33 when it flew. Yes, it helps and it was considered.

Didn't know that. Any significant roadblocks?

>>>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.
>
>You've just described Shuttle but substitute solid propellants instead of
>Lox/kerosene. That's basically the idea we use.

Yes, I was thinking the same thing.

>>>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.
>
>Exactly.

>>>> 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.
>
>Depends. I've seen claims that composite propellant tanks are the solution
>to our problems. But when designed, the composites tanks were actually
>heavier. Plus, they had significant reuse issues. (And the ones designed
>and tested were self-destructive ... as I predicted at the time they would
>be.)

Heh.

>> 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.
>
>Exactly.
>
>>>> 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.
>
>The rocket equation does indeed represent a fundamental limit. Whatever
>else one wants to do, you cannot violate physics.

Sure. But that particular limit doesn't rule out SSTO the way the laws
of thermodynamics, say, rules out a perpetual motion machine. It just
makes it impossible to do with the materials and technologies we have
on the shelf, and perhaps with those that we can reasonably hope to
develop in the next few years. Which may be quibbling, but I think
it's important to distinguish between something that can't be done and
something that we can't do, lest we fall into the trap of those who
trotted out formulas that proved that you couldn't achieve heavier
than air flight, when what they really meant was that you couldn't
achieve heavier than air flight with boilers and coal. (At least, I
you couldn't then -- there are always carbon composites . . . )

>> 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.
>
>As I said before, I'm all in favor of research. But initiated a development
>program without technology either in hand or clearly in the near term is a
>recipe for disaster, IMHO.

I think you're right, and I think that for some reason we've made that
mistake repeatedly in recent years. Or maybe it's most accurate to say
that you /can/ push things some -- I mean, we went from ICBM's to
Apollo 11 in only a few years -- but you have to be realistic about
how far you can push technology at each stage and allow for some
shortfalls and errors.

>>>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
>
>Lot's of supposition there. As I've said in another post, any reusability
>apparatus makes SSTO even more untenable than an expendable SSTO concept.

Why do you say that an expendable SSTO is untenable? Seems to me we
could do it now, given that there have long been expendable stages
with a good enough mass ratio and engines with sufficient specific
impulse, though from an economic perspective I assume it would be
senseless -- very high performance stuff to send up a small payload.

>> 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.
>
>Muti-lobe tanks were demonstrated. The X-33 lox tank, aluminum, worked just
>fine in testing. A subscale composite hydrogen tank also worked in testing
>as well. The choice of the final design for the hydrogen tanks, on the
>other hand, was a stroke of pure idiocy perpetrated by egotistic aircraft
>designers.
>
>> 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.
>
>And even, with all the paper efficiencies incurred by the risky
>technologies, they still couldn't achieve a positive payload.
>
>>>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 business case was based upon the mid-90's market which fueled lots of
>plans for 20k to 40k payloads. When the satellite market fell through the
>floor, the business case was hopelessly broken. I'd want something more
>than a paltry 20k just for purposes of effectively replacing Shuttle.

Sure, if you're talking about the Shuttle replacement. But even if you
could get an SSTO Shuttle replacement to work, I think it would be a
serious mistake. Much better to go with something reasonably
incremental when you're building a workhorse, and leave the cutting
edge stuff to x projects, prototypes, and then smaller stuff.

>> 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.
>
>Yes. And, as I said, the market fell to pieces. Even if every technology
>adventure was successful, it probably still wouldn't have made sense in the
>end. This is why I'm a very strong advocate for NASA building stuff for
>NASA and not to attempt to solve everyone else's issues (purely my opinion,
>not in any way official policy).

I guess I see NASA's role as comprising scientific and technical
research and exploration (and in that last context I think manned
exploration has an importance beyond by science done or economics). So
I can see NASA developing a cutting-edge vehicle or technologies that
can't be developed by private enterprise because the risks are too
great or the ROI is beyond the horizon of venture capitalists.

In that context, I look at some of the "business arguments" that are
supposed to justify SSTO, private space vehicles, and the like with a
jaundiced eye because they don't seem to me realistic -- but at the
same time, I believe that some of these technologies /will/ pay off
over the long term. And I guess you could say that that's true of
basic research in general.

>>>> 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.
>
>And how much beer?

Let's hope it was just that.

--
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
.

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