Date: Thu, 16 Jan 1992 14:59-EST From: space-tech-request@cs.cmu.edu To: "~/st/lists/stdigest" Subject: Space-tech Digest #94 Sender: mnr@DAISY.LEARNING.CS.CMU.EDU Contents: mars mission question (3 msgs) lightweight lasers (2 msgs), aluminum engines (3 msgs) Needed: Bi- and Monopropellant Thruster info (1 msg) Low Cost Expendable Boosters (8 msgs) ------------------------------------------------------------ Date: Mon, 16 Dec 91 16:27 PST From: trost@reed.edu (Bill Trost) To: BEAUFAIT@CEBAFVAX.BITNET, space-tech@DAISY.LEARNING.CS.CMU.EDU Subject: Mars mission Beaufait@cebaf2.cebaf.gov(?) writes: I ran across this paper that relates to a propulsion design for the mars mission (dates 1983). Its from a DOE sorce so I thought the powers that be may not be aware of it. The paper is titled "A LASER-FUSION ROCKET FOR INTERPLANETARY PROPULSION". It deals with the development of an interplanetary vehicile mass 486 ton without reaction mass.... No launching that dirty u word or p word. Less political hassel more chance of getting things done. While I'm not familiar with this particular work, most of these schemes require the availability of LARGE lasers (I mean, way big -- you know how big...never mind). Anyhow, I don't think you can use big lasers effectively from the earth's surface (gotta get the fools who put that silly atmosphere there), and I'm not certain about the implications of lasers that size in orbit -- they could easily push themselves up/down/all around. The best bet would be to push such lasers (and their power sources, and support facilities) on the moon, and that'll probably take a while. Which means we should get on the stick! -- Bill Trost If good planets are so hard to find, Reed College Systems Manager why aren't we looking harder? ------------------------------ Date: Tue, 17 Dec 91 10:09 EDT From: BEAUFAIT%CEBAFVAX.BITNET@BITNET.CC.CMU.EDU Subject: MARRS REPLY To: space-tech@DAISY.LEARNING.CS.CMU.EDU X-Original-To: space-tech@DAISY.LEARNING.CS.CMU.EDU, BEAUFAIT Bill. I think you are looking at this the wrong way. This is strictly an orbit to orbit vehicle. The lasers are integral to the system not externaly based,and will be facing aft. Laser power has to be 2Mj in 10ns* from artical original discusion (100 lbs thrust 10ns?) this would be insidental compared to the overall thrust but still added in the same direction. (Miner improvement in efficiency.) Answer to other reply; The most prominantly displayed call no# is ucrl-88857 writen in pen below that is conf-8310171-1. Maby that helps. ------------------------------ Date: Tue, 17 Dec 91 11:42:30 -0500 From: Jonathan Leech To: BEAUFAIT%CEBAFVAX.BITNET@BITNET.CC.CMU.EDU, space-tech@DAISY.LEARNING.CS.CMU.EDU Subject: Re: MARRS REPLY BTW, there is a citation of this paper in the sci.space FAQ. I forget who wrote the citation - Ted Anderson, perhaps - but it suggests contacting the "Technical Information Dept." at Livermore to get a copy. Jon __@/ ------------------------------ Date: Wed, 18 Dec 91 18:03:26 CST From: "Walker on Earth" To: space-tech@cs.cmu.edu Subject: light lasers A quick question: how much would an orbiting giga(tera?)-watt laser mass? Presumably it would be a gas laser ala CO2 or something more exotic confined in a 'bag' and would be optically pumped by solar energy. It seems possible that such a laser could have a high radia- tive output/mass ratio. have there been any experments or analyses concerning this concept? I know, I know, the efficiency would probably be quite low, but still. . . Thanks in Advance, Dwight Thieme ------------------------------ Date: Thu, 19 Dec 1991 01:51 EST From: "GORDON D. PUSCH" Subject: AtLas: a *REALLY* light laser :-D To: space-tech@cs.cmu.edu X-VMS-To: SPACE-TECH X-VMS-Cc: PUSCHG Dwight Thieme asks: > A quick question: how much would an orbiting giga(tera?)-watt laser > mass? Presumably it would be a gas laser ala CO2 or something more > exotic confined in a 'bag' and would be optically pumped by solar > energy. It seems possible that such a laser could have a high radia- > tive output/mass ratio. have there been any experments or analyses > concerning this concept? I know, I know, the efficiency would probably > be quite low, but still. . . Bob Forward once spoke of a *REALLY* low-mass laser of just such a type, in which the active gas was *the Earth's atmosphere*. It seems that there are layers of the Earth's (and also Mars's and Venus's) atmosphere that lases *naturally* at low power all the time, pumped by the sun (super- radiant cavityless mode, of course ;-). By orbiting a pair of large ( ~10 meter diameter) mirrors, I believed Forward claimed one should be able to produce a multi-gigawatt pulsed beam ( ~10-100 pulses/sec, w/ pulses ~10-100 usec long). 'Course, it'd take some pretty fancy ``formation flying'' of your mirrors to maintain adaquate allignment, and possibly some adaptive optics. I imagine a set of lower-powered pulses interleaved between the main pulses might be used to help sense spacing and alignment. Forward refered to the research proposal as ``AtLas,'' from ATmospheric LASer. It's discussed briefly in his USAF Advanced Propulsion System Research report; I'll dig it out of my files, if anyone's interested ... Gordon D. Pusch ! BITnet: TASCC A&D, Stn. 49a ! AECL, Chalk River Laboratories ! Phone: (613) 584-3311, X-4107 (off.) Chalk River, Ont. K0J 1J0 ! (613) 584-2368 (hm.) Canada ! ------------------------------ Date: Tue, 17 Dec 91 08:18:36 EST From: John Roberts Disclaimer: Opinions expressed are those of the sender and do not reflect NIST policy or agreement. To: space-tech@cs.cmu.edu Subject: Aluminum engines >From: freed@nss.FIDONET.ORG (Bev Freed) >Newsgroups: sci.space >Subject: Toward 2001 - 02 Dec >Date: 3 Dec 91 02:05:21 GMT >... >+ Wickman Space and Propulsion >Sacramento CA >Lunar soil broken down into liquid oxygen and aluminum powder powers >a new engine being developed by Wickman engineers. A subscale engine >has been sucessfully tested for periods of up to 35 seconds at >varying levels of thrust. Aha! Just what some of us have been discussing! :-) That would certainly seem to be attractive if it can be made to work, given the abundance of oxygen and aluminum on the moon. It could potentially be used for launches, and in conjunction with linear launchers. I may try to find out more about this. (Anyone else is welcome to do the same.) John Roberts roberts@cmr.ncsl.nist.gov ------------------------------ Date: Tue, 17 Dec 91 15:15:50 pst From: gwh@lurnix.LURNIX.COM (George William Herbert) To: uunet!cmr.ncsl.nist.gov!roberts@uunet.UU.NET Subject: Re: Aluminum engines Cc: space-tech@cs.cmu.edu Interesting. I tried doing some preliminary work on a hybrid Al fuel rocket with LOX oxidizer, but never got anywhere of significance (didn't have time to build a model 8-). It would seem to me that lightly sintered Al powder is a good fuel for such designs. -george william herbert gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ Date: Tue, 17 Dec 91 18:52:04 EST From: John Roberts Disclaimer: Opinions expressed are those of the sender and do not reflect NIST policy or agreement. To: space-tech@cs.cmu.edu Subject: Re: Aluminum engines I wouldn't expect the specific impulse to be as high as that of a hydrogen-oxygen engine, but on the other hand it doesn't take as high a specific impulse to launch from the moon as it takes to launch from Earth. John Roberts roberts@cmr.ncsl.nist.gov ------------------------------ Date: Thu, 19 Dec 91 17:22:42 -0800 From: George William Herbert To: space-tech@DAISY.LEARNING.CS.CMU.EDU Subject: Needed: Bi- and Monopropellant Thruster info I need some pointers to actual design info on existing Hydrazine and Hydrazine/Nitrogen Tetroxide thrusters and attitude control rockets. The adddress of companies that make them would help a lot, as would any publication that has lists or catalogs etc. General information on how they work etc. is _not_ part of this request. I already know that 8-) I just need to pick some specific examples as nominal hardware on a design I'm working on. Rough numbers don't cut it at the level of detail I'm trying to achive. Anything from ten newtons to a thousand newtons is useful (2 to 200 lb thrust). -george william herbert gwh@ocf.berkeley.edu ------------------------------ Date: Fri, 20 Dec 91 14:02:07 EST From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: Low Cost Expendable Boosters I just read an interesting article: Edward L. Keith, "Low-Cost Space Transportation: The Search for the Lowest Cost." In "Spaceflight Mechanics 1991," Advances in Astronautical Science Vol 75, part II. (AAS 91-169.) The article describes an expendable rocket that is projected to achieve a launch cost of $23 per pound of payload to low earth orbit. Features include: -- Use of LOX-kerosene propellant (the cost of this propellant mix is currently $.07/lb.) The propellant cost per pound of payload is < $3. -- Payload is 15,000 pounds. -- Simplified pressure-fed rocket engines. This is estimated to cut the cost of engines in half. Main engines are fixed, with steering by vernier rockets. -- Steel is used as a structural material, not aluminum or composites. The steel for the tanks is textured by cold-rolling, a simple and low-cost technique. --A propellant/dry weight ratio of 11 is assumed. This is conservative compared to an Atlas's stainless steel balloon tank, but would have to sustain higher pressure as well. If the ratio is decreased to 6.87, the launch cost rises to $62/lb. They have not shown that a ratio of 11 could be achieved; this is a target, not a provably achievable value. -- Isp is 250 seconds at sea level and 300 s. at altitude. If a range of 235 to 282 is assumed instead, the launch cost increases to $32/lb. Achieving the high Isp with low pressure engines is an engineering challenge, but in a pressure fed rocket lower chamber pressure --> lower tank pressure --> lighter tanks. -- A parallel-burn arrangement of seven identical boosters. All rockets operate simultaneously but feed from only two tanks at a time through a manifold system, giving, in effect, a rocket with 2.3 stages. Each segment weighs 8,170 pounds, is 8 feet in diameter, 40 feet long, and carries 89,871 pounds of propellant. Each segment has 4 main engines, for a total of 28 engines. They are all lit at launch, and are simple, so they should be reliable. -- If he plugs in all the worst-case assumptions (Isp, fuel/dry mass ratio) he still gets a cost of < $500/lb, similar to the cost projected for the "big dumb booster" and similar to the cost of the Soviet SL-4 and SL-13. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Fri, 20 Dec 91 13:50:38 pst From: gwh@lurnix.LURNIX.COM (George William Herbert) To: uunet!cs.rochester.edu!dietz@uunet.UU.NET Subject: Re: Low Cost Expendable Boosters Cc: space-tech@cs.cmu.edu Damn, and my _very similar_ hybrid engine'd proposal isn't ready for publication yet 8-( foo. My design was a large booster, using a hybrid (LOX/solid PBD fuel) motor cluster (6 units for first stage, 1 for second stage, smaller unit for third stage) and high strength steel structural elements. I'm not so sure if pressure-fed liquid is a better idea than pressure fed hybrid. And if they're working from 600-1000 PSI chamber pressure, their structural mass is low. I couldn't do that well, even using 100ksi yield steel (which, by the way, is absolutely great material to design with 8-) -george william herbert gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ Date: Fri, 20 Dec 91 22:10:54 EST From: dietz@cs.rochester.edu To: gwh@lurnix.LURNIX.COM Subject: Re: Low Cost Expendable Boosters Cc: space-tech@cs.cmu.edu One reason they went with pressure-fed liquid was so they could use parallel staging: all engines feed off common cross-connected tanks, with tank/engine segments dropping away as fuel is consumed. Also, at the costs they were projected, the cost of fuel is important. The paper admitted the structural mass was challenging. It said at one point that variable nozzle ratio engines would be needed (presumably, so that the chamber pressure could be kept low without making the engine overexpanded at sea level or underexpanded in vacuum). Details were not provided. What sort of fuel+oxidizer/dry mass ratio can be achieved with steel pressure-fed hybrids? Paul ------------------------------ From: sequent!techbook.com!szabo@uunet.UU.NET (Nick Szabo) Subject: Re: Low Cost Expendable Boosters To: uunet!cs.rochester.edu!dietz@uunet.UU.NET Date: Fri, 20 Dec 91 20:08:46 PST Cc: space-tech@cs.cmu.edu X-Mailer: ELM [version 2.4dev PL32] > Edward L. Keith, "Low-Cost Space Transportation: The Search for the > Lowest Cost." In "Spaceflight Mechanics 1991," Advances in Astronautical > Science Vol 75, part II. (AAS 91-169.) > > The article describes an expendable rocket that is projected to > achieve a launch cost of $23 per pound of payload to low earth orbit. >... By what method was this figure derived? What is cost breakdown for engines, structure, tank, control systems, facilities, operations, assumptions about flight rate? What is the cost to get to first flight? > -- Payload is 15,000 pounds. >... Then the proposed cost per flight is $345,000, about 20 times lower than Pegasus (which only lifts 1,000 pounds), or 120 times lower than an Atlas with similar payload. What are the amounts and sources of proposed savings, for each category, between the Atlas and this rocket? Can the shroud be used for existing upper stages and satellites? Will be the payload environment (thermal, electromagnetic, contamination, loads from venting, aerodynamics, acceleration, vibration, acoustic, etc) be significantly different from current launchers? > -- Simplified pressure-fed rocket engines. This is estimated to > cut the cost of engines in half. >... Does this mean all 28 engines together cost half of, for example, the two first-stage Titan 2 engines, or is this per engine? > Main engines are fixed, with > steering by vernier rockets. >... How much do the vernier rockets and control systems cost? Are there any cost savings in the control system? > -- Steel is used as a structural material, not aluminum or composites. > The steel for the tanks is textured by cold-rolling, a simple > and low-cost technique. >... How much lower? Can cold-rolling produce a tank that is sufficiently balanced? Can we use a standard rolling mill, or must a mill be custom-built for these tanks? > ...in a pressure fed rocket lower > chamber pressure --> lower tank pressure --> lighter tanks. >... At what point do aerodynamic, acceleration, and/or acoustic stresses become a more limiting factor than tank pressure? > -- A parallel-burn arrangement of seven identical boosters. All > rockets operate simultaneously...[4 engines per booster] >... How is the liquid distributed proportionately to the engines? If one or more of the engine's performance suffers, can the rocket compensate with the verniers, or will it go off course? How much doees this fuel distribution system cost? > -- If he plugs in all the worst-case assumptions (Isp, fuel/dry mass > ratio) he still gets a cost of < $500/lb, similar to the cost projected > for the "big dumb booster" and similar to the cost of the Soviet SL-4 > and SL-13. >... We need to know the accounting method for this or the preceding dollar figures to be meaningful. Also, I am curious as to how the Soviet booster costs were determined. Some other questions: * What are the proposed launch site(s)? * Can existing pads be used? If so, how much is the cost of modification, if not, how much is the cost of building a new pad? * How are the rocket sections transported to the launch sight and how much does that cost? * What kind of facilities are needed to interface payload and launcher? * What equipment is needed to connect the launchers stages and erect the payload on the pad? Of the factors mentioned, engines are only proposed to provide a factor of 2 savings (but I do not know if that includes vernier rockets and control systems). The amount of savings from going to cold-rolled steel is not mentioned, but tank manufacturing is only a small fraction of launch costs, as is fuel. Several major cost drivers (control systems, facilties, operations, etc.) are not mentioned. Altogether, the described design changes seem to provide savings of much less than a factor of 120 over the current art; I am not convinced they provide even a factor of 2. However, to determine the actual savings we need a cost comparison to existing launchers by category to see what savings the design changes do and do not provide. Sorry if this sounds like space-investors, but you did mention $$$ :-) szabo@techbook.COM ...!{tektronix!nosun,uunet}techbook!szabo Public Access UNIX at (503) 644-8135 (1200/2400) Voice: +1 503 646-8257 Public Access User --- Not affiliated with TECHbooks ------------------------------ Date: Sun, 22 Dec 91 10:32:33 -0800 From: George William Herbert To: dietz@cs.rochester.edu, gwh@lurnix.LURNIX.COM Subject: Re: Low Cost Expendable Boosters Cc: space-tech@cs.cmu.edu Using the highest strength steel that can be commonly purchased (100 thousand PSI Yield strength) and a maximum 600 psi pressurization, the best I could do was a tankage fraction of 0.12-0.14 . You can find, if you look and try real hard, steels at yield strengths up to 140,000 ksi, but they tend to be very hard to manufacture with, and may not deal with cryo temperatures well (If you use LOX, as I do...). IMHO it's not practical to drop below 0.12 and never will be, using a 600 PSI rocket. That chamber pressure seems to be the best for a hybrid of this type when trading off Isp (higher as chamber pressure increases) and tankage mass (pressure vessels get linearly more heavy as the pressure rises, in this structural design regime). Above 600 PSI, with a hybrid, you lose by increasing structural mass faster than you gain by increasing Isp. * note: the above conclusion is supported by a good deal of estimation from (scanty 8-) published information. My presumption is that a Hybrid at 600 psi chamber pressure will average an Isp of about 280 over its flight, presuming appropriate expansion ratios on the stages, and perhaps variable expansion nozzles (which are also convenient for thrust vector control...). The 280 is actually slightly conservative, as that's about the actual sea level Isp of the LOX/PBD+metal combination. At 1000 PSI, you only get a slight increase, perhaps to Isp=300, and the tankage fraction rises to 0.20 -george william herbert gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ Date: Sun, 22 Dec 91 10:38:23 -0800 From: George William Herbert To: space-tech@cs.cmu.edu Subject: More info... BTW, switching from hybrid to liquid will raise the tankage fraction an appreciable amount. RP-1 is significantly less dense than a hybrid's solid propellant (about 0.73-0.75 vs 0.98+). -george william herbert gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ From: henry@zoo.toronto.edu Date: Sun, 22 Dec 91 19:10:43 EST To: space-tech@cs.cmu.edu Subject: Re: More info... >BTW, switching from hybrid to liquid will raise the tankage fraction an >appreciable amount. RP-1 is significantly less dense than a hybrid's >solid propellant (about 0.73-0.75 vs 0.98+). Although this may be counterbalanced if your chamber walls have to be heavier than tank walls, e.g. for thermal reasons. Henry Spencer at U of Toronto Zoology henry@zoo.toronto.edu utzoo!henry ------------------------------ Date: Mon, 23 Dec 91 11:27:56 pst From: gwh@lurnix.LURNIX.COM (George William Herbert) To: uunet!zoo.toronto.edu!henry@uunet.UU.NET Subject: Re: More info... Cc: space-tech@cs.cmu.edu For nearly the entire thrust chamber of a hybrid, there is little thermal stress on the walls (the fuel will reasonably effectively insulate the walls until it's nearly burned away). I am mass budgeting some fiberglass ablative insulation for the sections that aren't covered by fuel. 8-) As far as I can tell, at the 600 PSI pressurization regime, the dominating loads are all pressurization related. Those that aren't are nearly all in the direction of the rocket's axis, and thus in the direction where pressurization loads are lowest (axial loads in a cylindrical pressure vessel are 1/2 of the circumfrential loads 8-) IMHO, and I haven't been able to poke holes in my design yet (including Nick's transportation costs etc 8-), it's possible to achive a $300/lb expendable, potentially $150-200 eventually, using a hybrid. I could see doing it with liquid fuels too, though it might be better to use a hybrid. Less parts need to be machined for a hybrid's chamber, for instance... all my design uses (except for some thrust vector control hardware... at this point, a vectoring nozzle extension) is welded parts. In standard steel, with easily assemblable thicknesses. I've priced and gotten firm offers on the materials involved and moderately firm prices on the welding (they keep wishy-washing over how much inspection would cost, but I think that's because I didn't set a firm order amount on the amount of welding...). I think that I ought to post design details... people seem interested in the general line of discussion. Maybe one of you will be able to kill some of my reasoning, or help resolve some of the issues 8-) -george william herbert gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ End of Space-tech Digest #94 *******************