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

On Sun, 28 Aug 2005 21:16:09 -0500, "w.d.greene" <bil64@xxxxxxxxxxx>
wrote:

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

>>>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.

Apropos staged combustion aerospikes, I just came across this --

"If you try to be cute and clever and build a continuous ring-shaped
chamber, or a ring broken up into a small number of segments, then
yes, instability is a worry. Some of Rocketdyne's aerospike
experiments with a ring chamber had a fairly strong oscillation
running around the ring. (The linear configuration a la X-33,
interestingly enough, had no trouble.)

"It helps a lot to inject one propellant as a warm gas; expander and
staged-combustion cycles have an advantage here. Neither the RL10 nor
the SSME had stability problems, but both the J-2 and J-2S had some,
and the less said about the LOX/kerosene and N2O4/hydrazine engines,
the better. (This is one area, by the way, where hypergolics are
probably worse rather than better.)"

>> 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.

Sure.

>> 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.

I'm wondering, given that the sea-level and vacuum impulse figures are
lower than the SSME, whether they wouldn't have been better off with
the RS-2100, which was apparently supposed to have a sea-level Isp of
384, vacuum Isp of 450, and a thrust-to-weight ratio of 83 . . .

>>>>> 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.

Didn't mean to imply it was. I'm sort of assuming that the RS-2200 and
RS-2100 would be cranky at first, and would be upgraded with time as
the SSME was . . .

>>>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.

I wish! I was just reading about it a few days ago.

>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.

I was wondering about that yesterday. You could make the drop-off
engines recoverable -- end up with most of the benefits of 1.5 stage
to orbit but without the dangers of a side-by-side cryogenic fuel
tank. Or how about dropping recoverable engine/kerosene tank boosters?
All the LOX would be supplied from the orbiter, so there wouldn't be
an ice/foam problem. (For that matter, you could use SRB's, but it
seems they cost about as much to recover and refurbish as they do to
replace.) And weren't they talking about a liquid-fueled fly back to
base replacement for the SRB's?

>>>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.

Yes, I've been reading about the Sabre engine. Which. however uses a
turbine.

--
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
.

Relevant Pages

  • Re: Google Cindy Sheehan (was: Re: Peace Mom)
    ... >>>engines are the highest performance, ... It has nothing whatsoever to do with 1970's technology or 2010 ... An expanding nozzle or better ... SSME could be used with a aerospike since there is no such supply of turbine ...
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  • Re: Vision of the three Rs: Regular, Reliable and Reusable
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  • Re: Google Cindy Sheehan (was: Re: Peace Mom)
    ... SSME, I think. ... Engines are heavy. ... >It is expensive and time-consuming to recover and refurbish the boosters, ...
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  • Re: Its 1972 all over!
    ... The aerospike. ... It turns out we don't even need it, if we just stick the SSME into the nosecone of the rocket and return the nose cone to earth. ...
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