Date: Sat, 18 Jan 1992 13:33-EST From: space-tech-request@cs.cmu.edu To: "~/st/lists/stdigest" Subject: Space-tech Digest #95 Sender: mnr@DAISY.LEARNING.CS.CMU.EDU Contents: Big Dumb Hybrid writeup, liquid rocket design (16 msgs) ------------------------------------------------------------ Date: Tue, 24 Dec 91 12:14:45 -0800 From: George William Herbert To: space-tech@cs.cmu.edu Subject: Bid Dumb Hybrid writeup Feel free to criticize any and all aspects of this paper. -george Bid Dumb Hybrid Booster George William Herbert What is the Big Dumb Hybrid Booster design? It simply is a complete re-examination of all aspects of booster vehicle design and manufacture in an attempt to find a more effecient design. Most aerospace vehicles are manufactured of high-strength, low weight alloys and composites, in an attempt to absolutely minimize mass and therefore maximize payload. The side effect of this is that they cost extremely large amounts to manufacture and design. The BDHB is a complete reversal of classical space vehicle design philosophy. Rather than minimize mass, we make it cheaper and bigger. Rather than use high-strength Aluminum and composites, use simple high- strength steel. Fuel is cheap; if the booster dry mass gets significantly cheaper as it gets bigger, you win overall. Simmilarly, a choice to use a three-stage design minimizes the performance pressure on each stage, allowing economical (and feasible) low-cost higher-weight design. A two stage design of this nature is probably not reasonable. The BDHB is a three-stage pressure-fed hybrid (solid fuel, liquid oxidizer) design, with all stages of the design optimized for minimum cost. A hybrid type engine was selected for its intrinsic ease of manufacture; the only components needing precise engineering and manufacture are an oxidizer flow valve and thrust vector control mechanisms. The rest of the vehicle is a simple pressure vessel, cylindrical in shape, manufactured from rolled and welded A514 steel. The main nozzle is manufactured from welded cone sections, as are the tank ends and the area between the oxidizer tank and the thrust chamber. The steel involved is widely available and easily weldable in thicknesses that this design calls for. The performance of the fuel/oxidizer combination of any rocket is critical. Solid rockets tend to provide low specific impulse though with relatively cheap (and heavy) casings. Liquid fuels are better, but there are limits to how inexpensively a standard rocket engine can be manufactured. Some of the more exotic and high-performance fuel/oxidizer combinations have other disadvantages, like toxicity (Flourine oxidizer) or very low density and handling problems (Hydrogen fuel). Hybrids tend to avoid most of these problems. Using a Liquid Oxygen oxidizer and a metallic particle laden rubber fuel, a specific impulse of at least 280 (sea level, 600 psi chamber) is a reasonable assumption. This is high enough to allow a three-stage design with a mass ratio of very close to 3:1 at each stage (and payload ratio of 1/6 per stage) to achive orbital velocities. Should this prove unreasonable, even an Isp of 260 will work, albeit with about twice the initial (launch) mass for the same payload. But the figure of 280 should be reliable. In addition, the hybrid avoids the handling problems associated with solid boosters (the fuel, while flammable, is not explosive nor significantly hard to handle). DESIGN ASSUMPTIONS I have made a number of judgement calls (mostly based on insufficient data) in some areas in preparing this design. One is that a metalized PBD/LOX hybrid will have a fuel burn or regression rate of about 7.5mm/sec at a chamber pressure of 600 psi. This is supported by smaller scale data on hybrid performance, but I have been unable to get confirming data from literature or Amroc, who actually built some large (though not _this_ large) hybrids. A regression rate of as low as 5mm/sec can be accomidated without seriously affecting the design (about a 25% initial mass increase). That regression rate consequently limits the overall diameter. A maximum allowable diameter of 4 meters is assumed. THE DESIGN General Configuration: The example BDHB design is a three-stage launch vehicle. The first stage is a pair of 4m diameter and 100m long cylinders, each massing very close to 900 tons including fuel. Over half this mass (585 tons each) is Liquid Oxygen oxidizer. About 180 tons is metalized Polybutadiene fuel. The rest, 135 tons worth, is the "dry" mass of the booster, approxomately 98 tons structural mass, and the rest auxiliary equipment (thrust vector paddles, ablative insulation in the thrust chamber and nozzle, insulation in the LOX tank, the LOX valve, and a LOX tank pyro pressurization system and pressure release valves). All the auxiliary items have generous design margins (50% above authors estimate, overall), and the structural mass is well developed. The second and third stages, along with payload, are of similar design, and are nestled along the lower half of one side of the two first-stage engines: TOP VIEW SIDE VIEW /--\/--\ /\ \__/\__/ <= - - - - - first stage motors - - - - - => || /--\ || \__/ <= second and third stages stacked with payload /|| \- => ||| ||| ^^^ Structural Design: The vehicle (all stages) is manufactured from (mostly 10mm thickness) A514 steel plate, rolled into cylindrical shapes and welded. Some specific high-stress areas have reinforcement or increased thickness, but the main body of both the fuel and LOX sections are simply a welded cylinder. Unpressurized, it is sufficiently strong and rigid to handle its own shipping stresses. In operation, at operating pressures of 720 PSI (LOX tank) dropping to 600 PSI in the fuel chamber, it is stressed in tension by the internal pressure loads. All other calculated flight loads were significantly lower than the pressure loads, and all flight loads are in the axial direction. The mechanics of stress in cylindrical pressure vessels lead to a stress axially that is half that of the circumfrential stress; thus, all the flight loads are in the direction that is already least stressed. At all stages of design, a nominal safety factor of 1.3 will be maintained. It is intended that high-stress areas be designed to a higher factor, to be conservative. As stated, the normal thickness of plate used is 10mm, with some areas stronger. Welding of this thickness A514 steel is well within current welding technology's limits, and can be performed by a large number of shops (or in-house). Once welded (and inspected), it can reasonably be assumed to retain full strength in the joints. A radiographic and other inspections of all welds will be specified, to ensure against uncorrected weld flaws. Again, these inspections are well within current limits, and will be relatively inexpensive. At this point, a 15% margin has been included for structural weight increases. Equipment Design: The critical outfit items are the oxygen valve and the thrust vector paddle system. The insulations (etc) are straightforward. The valve is currently being investegated, but a design allowing over five tons has been allowed for. This would be an order-of-magnitude more weight than known valves of the same size and operating condition, which will allow inexpensive construction. A rellatively novel plunger-type valve is also being considered The thrust vector paddles are nozzle extensions, manufactured of the same steel as the rest of the vehicle. Actuators are baselined as aircraft-type electrical actuators (flap, etc). Similar (perhaps identical) actuators should be used in the valve. Design selection has not been made, though a variety of designs that are suitable exist. Further detail analysis is required for a choice. Guidance and control are baselined as PC based (hardened) designs, similar to that used on the Pegasus vehicle. Inertial and GPS recivers will be fitted. COSTING Strucutral: The cost of the structure of the vehicle can be well estimated; a total of about 240 tons of steel will be required, and the amount of welding needed (and inspections etc) can be accurately estimated. Steel cost, including forming and shipping to any arbitrary location within the US, will be less than $350,000 [hard bids were obtained for this]. Welding and inspection performed anywhere but California can reasonably be estimated at $700,000 [bids were solicited on this estimate, and that's a reasonable extrapolation from the (wishy-washy) responses, with some margin...] total: abt. $1.05 million Fuel: Fuel and oxidizer are cheap: LOX runs about $25/ton in bulk, and metalized PBD should be less than $.25/lb in bulk (formed in place and inspected). Total: about $0.275 million Equipment: The valves are baselined at $100,000 apiece; this is lower than current valving, but they are simpler and heavier, so ought to be about 50% cheaper. Thrust control vanes and actuators are a bit harder to estimate. An estimate of about $100,000 each, including actuators, is reasonable with some margin. Insulation is low cost; about $75,000. Other items should run under $100,000. total: abt. $1.65 million subtotal: abt. $3 million dollars. Transportation of the vehicle via sea to the pad should run under $100/ton, with only the "dry" weight and the PBD weight included; or about eight hundred tons for $80,000 The pad itself needs little in the way of equipment (just LOX supply and electrical connections, and erection gear). It should cost under $5 million, amortized over 10 launches (ignoring interset...) about $500,000. Assembling, erecting, and preflight checks on the vehicle are baselined as requiring about 40 persons dedicated. At an average payroll of $40,000 and two launches per year (conservative), $0.8 million per launch. Total so far: $4.4 million Assuming normal engineering economy, a gross margin of roughly 40% would be required to recover investment and such; that would yield a total launch cost in the $7.35 million range. Estimating $8 million, with a margin to go to $10 million, is reasonable. At those costs, the vehicle will provide launch services at a cost of about $365-$450/lb to LEO. This is notably about 1/10 that of current vehicles, and equivalent to several current reusable system estimates. Design and development of the vehicle, due to its inherent simplicity, are estimated at about $14 million, but including a safety margin, about $18 million. It should be possible to accomplish development in on the order of 24 months. ------------------------------ Date: Tue, 24 Dec 91 17:25:40 EST From: dietz@cs.rochester.edu To: gwh@ocf.Berkeley.EDU Subject: Re: Bid Dumb Hybrid writeup Cc: space-tech@cs.cmu.edu George W. Herbert wrote: > Feel free to criticize any and all aspects of this paper. Just a few... I am concerned about the thrust-vector control mechanism. There will be large forces on the paddles that will have to be countered by the actuators. It would be preferable, I think, to have a system in which no powerful actuators are necessary. Vernier rockets, for example, or thrust vectoring by fluid injection. Another problem I have is that the burn rate of the hybrid may not be easily controllable. With liquid fuel engines, one can carefully meter both fuel and oxidizer, and leave less in the tank at the end. The mixture ratio is also closer to optimal. A regeneratively cooled nozzle should also stay closer to the optimal shape than an ablative nozzle. In a liquid fueled rocket, one also has more leeway in the shape of the pressure vessel for the fuel. Perhaps it can be shorter and more squat (closer to a sphere) than a hybrid engine, although I imagine aerodynamics limit what one can do here. I think LOX costs more than $25/ton in the quantities you are discussing; the figure I've seen is about $80/ton. Note also that the cost of kerosene, per pound, is about 1/3 that you quote for PBD. Fuel cost is a sufficiently small part of the launch cost that that doesn't make much difference. My biggest concern would be the development risk. Why develop a new hybrid engine? Can't you go to the patent office and pull out designs for low-tech pressure-fed kerosene-LOX engines? The patents should have expired by now. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Fri, 27 Dec 91 12:11:12 pst From: gwh@lurnix.LURNIX.COM (George William Herbert) To: space-tech@cs.cmu.edu Subject: Re: Re: Big Dumb Hybrid writeup Paul Deitz writes: >Just a few... > >I am concerned about the thrust-vector control mechanism. There >will be large forces on the paddles that will have to be countered >by the actuators. It would be preferable, I think, to have a system >in which no powerful actuators are necessary. Vernier rockets, >for example, or thrust vectoring by fluid injection. Doing vernier rockets with a hybrid is a bit hard... and adding a liquid fueled (even pressure fed) vernier system would add systems, which can't be done cheaply, which will increase cost... this is a cost-constrained design. Fluid injection was considered, but it looks like one of the big costs for the design is the valving. Adding more variable-flow valves may not be a good idea. A tradeoff between the paddle and the injection methods is worthwhile, but probably can't be done at this level of development. I need more info on valve costs. >Another problem I have is that the burn rate of the hybrid may not >be easily controllable. With liquid fuel engines, one can carefully >meter both fuel and oxidizer, and leave less in the tank at the end. >The mixture ratio is also closer to optimal. A regeneratively >cooled nozzle should also stay closer to the optimal shape >than an ablative nozzle. The burn rate criticism is valid, and cannot be completely ignored. Hybrids are in a lot of ways not the best rocket engine system. Having it slightly throttlable is a plus, though. Regenerative cooling of the nozzle will require extensive and expensive manufacturing of the nozzle. I think it's a poor choice. Remeber, this design's cardinal rule is bigger and cheaper, not better first. Most of these pluses to performance increase cost more than they will performance, thus are losing propositions economy-wise. >In a liquid fueled rocket, one also has more leeway in the shape >of the pressure vessel for the fuel. Perhaps it can be shorter >and more squat (closer to a sphere) than a hybrid engine, although >I imagine aerodynamics limit what one can do here. It's possible to design the LOX tank for the first stage as one tank, shorter and squatter, rather than two as I did. The disadvantage is incrased wall thickness (harder to weld). Getting it closer to spherical is probably not worth it... it might save a few percent, but forming doubly-curved (spherical) steel sections isn't cheap. Cylindrical and conic sections decrease manufacturing expense immensely. >I think LOX costs more than $25/ton in the quantities you are discussing; >the figure I've seen is about $80/ton. Note also that the cost of >kerosene, per pound, is about 1/3 that you quote for PBD. Fuel cost >is a sufficiently small part of the launch cost that that doesn't >make much difference. LOX for industrial processes was (supposedly) available in the $25/ton range. More purified might cost more, i guess... but there's no reason for this rocket to have to use 99.999 % LOX 8-) it's not that picky. You're right that fuel cost is low enough that it doesn't make a difference here, and IMHO (from having played with numbers a bit) it's not feasible to get down into that region. >My biggest concern would be the development risk. Why develop a new >hybrid engine? Can't you go to the patent office and pull out >designs for low-tech pressure-fed kerosene-LOX engines? The patents >should have expired by now. Cost. Mostly manufacturing cost. Liquid fuel rockets require very carefully machined and manufactured elements. They take time, effort, etc., and I doubt that you can build them for as inexpensively as you can a hybrid. Development cost is the one potential killer, but if you develop it the quick and dirty way (i.e. build one and see what goes wrong...), then it can be done cheaply. I personally think that there's nothing wrong with the Cessna wing design process in an expendable booster (build one with about half the safety margin you need, with somea allowances, bend the test example till it fails, then reinforce the failure points until it's as strong as you need...). > Paul F. Dietz > dietz@cs.rochester.edu George William Herbert gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ Date: Sat, 28 Dec 91 01:40:09 EST From: dietz@cs.rochester.edu To: gwh@lurnix.LURNIX.COM Subject: Re: Re: Big Dumb Hybrid writeup Cc: space-tech@cs.cmu.edu > >My biggest concern would be the development risk. Why develop a new > >hybrid engine? Can't you go to the patent office and pull out > >designs for low-tech pressure-fed kerosene-LOX engines? The patents > >should have expired by now. > Cost. Mostly manufacturing cost. Liquid fuel rockets require very > carefully machined and manufactured elements. There are two places where cost matters: (1) Getting the project off the ground, you want development cost -- which will dominate manufacturing cost -- to be as low as possible. (2) Once production has ramped up, you want manufacturing cost to be low. Testing your hybrid rocket will require the expenditure of a fully fueled tank for each test. Testing a liquid rocket engine does not use up the fuel tank. So hybrids could very well be more expensive to develop than liquid fueled rockets, unless you can reuse the tank. But refilling it with solid fuel is going to be timeconsuming. Now, what happens at high production volume? There, you want to make many small units, so that manufacturing economies of scale can come into play. This is what the Russians do on their boosters -- they gang lots of engines of moderate size. I believe they make the engines for the SL-4 in lots of size ~1000. Lots of small engines should also reduce development cost, since testing a small engine should be easier than testing a large one. This applies to hybrids as well as liquid engines, of course. Look at the AMROC "ILV" design -- they wanted to gang many smallish hybrid engines running off a common LOX tank. You should as well. They would steer by varying the LOX flow to the different engines, but then they did have problems with the LOX valve. I am not convinced that manufacturing a pressure-fed liquid engine should be all that hard, especially if the design is conservative. You should be able to make cooling channels in the chamber and nozzle by electroplating of (say) nickel onto a positive form that is then melted or dissolved away. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ From: henry@zoo.toronto.edu Date: Tue, 31 Dec 91 18:40:54 EST To: space-tech@cs.cmu.edu Subject: Re: Re: Big Dumb Hybrid writeup > Fluid injection was considered, but it looks like one of the big >costs for the design is the valving. Adding more variable-flow valves >may not be a good idea... You could end up needing them for hydraulic actuators for the paddles. The fewer different fluid systems, the better. >Regenerative cooling of the nozzle will require extensive and expensive >manufacturing of the nozzle. I think it's a poor choice... Huzel and Huang had an interesting observation that is relevant to this: there appears to be no existing rocket engine that uses pure regenerative cooling. *Everyone* gets some of their wall/nozzle cooling by other means, such as arranging injectors to get a film of unburned fuel along the wall. >>... Can't you go to the patent office and pull out >>designs for low-tech pressure-fed kerosene-LOX engines? The patents >>should have expired by now. > >Cost. Mostly manufacturing cost. Liquid fuel rockets require very >carefully machined and manufactured elements... Not quite; *traditional* liquid-fuel rockets require careful machining etc. There is no reason why a pressure-fed liquid should need better manufacturing than your hybrid; indeed the manufacturing may be easier because the chamber can be much smaller. And there is much more knowledge about how to and how not to build liquids, e.g. Huzel&Huang for starters. The Long March series of boosters is built by a company that also builds refrigerators. Don't confuse Western military-missile construction methods (used for all current Western launchers, including the ones that are nominally civilian like Ariane) with natural laws. Henry Spencer at U of Toronto Zoology henry@zoo.toronto.edu utzoo!henry ------------------------------ Date: Wed, 1 Jan 92 13:56:06 EST From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: Liquid rocket design This discussion of simple, low-tech engines made think about how simple a liquid fueled engine could be. The advantage of the hybrid rocket George proposed is that no active cooling of the combustion chamber is required (well, the LOX no doubt cools the LOX injector, but that seems straightforward). The walls of the combustion chamber, nozzle and nozzle throat are either the fuel (which ablates faster than it conducts heat) or some refractory material that ablates slowly (graphite or phenolic composite, perhaps). But: one can do the same thing in a liquid fueled rocket. Line the combustion chamber with a refractory material. Because the mass ablated from the walls need not be large, the chamber can be much smaller than a hybrid rocket's combustion chamber. The wall material can be chosen to erode more slowly than metalized PHB. The injectors and surrounding area have to be cooled, but that is true in hybrids anyway. One would have to be careful that the changing combustion chamber geometry does not cause the engine to enter a regime of unstable burning. Presumably this could be at least partially solved by designing the wall material to erode unevenly, so that there are an odd number of cavities around the chamber. Anyway, the same problem has to be solved in hybrid rocket and solid rockets. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Reply-To: davidsen@crdos1.crd.ge.com Date: Thu, 2 Jan 92 08:56:03 EST X-Mailer: Mail User's Shell (6.5 4/17/89) From: davidsen@crdos1.crd.ge.com To: space-tech@cs.cmu.edu Subject: Re: Low Cost Expendable Boosters > * 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...). If variable expansion nozzles must be used (and I think they must), then could these be used to provide the steering by differential thrust? If so, that would save the weight, cost, and reliability penalties for having a set of steering nozzles as mentioned earlier. This sounds so obvious I'm sure someone has thought of it, if not used it. I have none of the refs I need to even do this on the back of an envelope. ------------------------------ From: henry@zoo.toronto.edu Date: Fri, 3 Jan 92 14:43:58 EST To: space-tech@cs.cmu.edu Subject: Re: Liquid rocket design >But: one can do the same thing in a liquid fueled rocket. Line >the combustion chamber with a refractory material. Because the >mass ablated from the walls need not be large... Indeed, ablation cooling for liquid motors is quite a respectable technique that has seen considerable use. It's most common in small motors, such as the Apollo Command Module attitude thrusters, but it has been used in some larger ones, notably both descent and ascent engines for the Lunar Module. Some of the Big Dumb Booster work that was done in the 70s looked at much larger ablation-cooled motors, and concluded that there should be no problem. >One would have to be careful that the changing combustion chamber >geometry does not cause the engine to enter a regime of unstable >burning... The classic approach to this is to use an ablative material in which the shape is defined by a refractory matrix material while the cooling is done by a less-refractory filler. Midway through the burn, then, you have a skeletal matrix filled with (relatively) cool gas forming the actual chamber wall, with the ablation front considerably deeper down supplying the gas. It's common to use a non-ablative refractory insert to define nozzle throat geometry, however, since this is rather critical. The geometry of the rest of the engine actually isn't a big problem apart from the stability issues: in the chamber itself the gas velocities are low and the exact geometry is not crucial, while the heating problems are less severe in the nozzle and holding a desired shape is less difficult. Even regeneratively-cooled engines have to worry a little bit about effects due to geometry changes. Hydrocarbon engines, in particular, routinely deposit carbon on the chamber walls during operation. I don't know if this changes geometry enough for stability worries, but I know the cooling design has to consider it. Henry Spencer at U of Toronto Zoology henry@zoo.toronto.edu utzoo!henry ------------------------------ Date: Fri, 3 Jan 92 23:33:59 EST From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: Re: Liquid rocket design Henry wrote: > [ some stuff about ablation cooled liquid motors ] Indeed, I should have looked in Sutton before I posted that. The ablation rates given for pyrolytic graphite are quite modest -- .001 in/s -- and a few times that for lower grade (if cheaper) graphite. He doesn't give ablation rates for phenolic/graphite composites. What are the numbers there for "typical" chamber conditions? Pyrolytic graphite is described as being formed by pyrolysis of methane. This seems similar to the processes involved in making diamond films. Has anyone looked at these films for use in rocket engines, I wonder? One thing that came up off line: just how hard is it to pressurize a LOX tank? LOX is fairly cold, so more gas has to be present to achieve a given pressure. The alternatives to LOX are expensive, nasty, and/or of lower performance, but perhaps an ambient temperature oxidizer like nitrogen tetroxide could have a more compact pressurization system (and a system pressurized by something cheaper than helium). Interestingly, the pressure George chose for his tank, 720 psi, is only a shade below the critical pressure for oxygen (734.1 psia). The critical density of oxygen is poor (~.4 g/cc), so you clearly want to start with colder stuff. Still, supercritical oxygen would get around some problems of sloshing or starting in zero g. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Sat, 4 Jan 92 23:55:36 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: Diamond films >Date: Fri, 3 Jan 92 23:33:59 EST >From: dietz@cs.rochester.edu >To: space-tech@cs.cmu.edu >Subject: Re: Liquid rocket design >Pyrolytic graphite is described as being formed by pyrolysis of >methane. This seems similar to the processes involved in making >diamond films. Has anyone looked at these films for use in rocket >engines, I wonder? I don't recall the exact method of making diamond films, but as I recall it starts with an organic gas, and temperatures/pressures aren't nearly as high as one might intuitively think. About five or six doors down from my office is the diamond film deposition lab - the only one used by the US government that I'm aware of. I don't know the people involved (different organizational unit), but if anyone has some questions about the process or the characteristics of the material (not potential applications), perhaps I could pass them on. John Roberts roberts@cmr.ncsl.nist.gov ------------------------------ From: henry@zoo.toronto.edu Date: Mon, 6 Jan 92 23:22:11 EST To: space-tech@cs.cmu.edu Subject: Re: Liquid rocket design >One thing that came up off line: just how hard is it to pressurize a >LOX tank? ... Not very. Just run some of the LOX through a heat exchanger and you have pressurization gas without needing a separate gas system. (Admittedly, you do need to get the heat from somewhere; maybe cool your nozzle throat a bit with it?) This does have the disadvantage of pressurizing with a relatively high-molecular-weight gas; on the other hand, it's stored much more compactly than something like helium would be. See Huzel & Huang's chapter on pressurization systems. Henry Spencer at U of Toronto Zoology henry@zoo.toronto.edu utzoo!henry ------------------------------ Date: Mon, 6 Jan 1992 22:46:21 -0600 From: Fraering Philip G To: henry@zoo.toronto.edu, space-tech@CS.CMU.EDU Subject: Re: Liquid rocket design Could someone explain what "Huzel and Huang" is, and where I could get a free copy? Phil ------------------------------ From: henry@zoo.toronto.edu Date: Tue, 7 Jan 92 00:49:22 EST To: Fraering Philip G Cc: space-tech@CS.CMU.EDU Subject: Re: Liquid rocket design >Could someone explain what "Huzel and Huang" is, and where I could >get a free copy? Dieter K. Huzel and David H. Huang, Design of Liquid Propellant Rocket Engines, NASA SP-125. This is the bible on liquid-rocket design; Sutton has references to it everywhere, and rightly so, because it spends chapters on things Sutton covers in a page or two. Pressurized-gas feed systems, for example, are pages 151-176. Not for the faint of heart or the innumerate. The bad news is that there ain't no such thing as a free Huzel&Huang. This thing is an inch thick; don't expect anyone to xerox it for you for fun. It's long out of print but can be bought (in xeroxed form) from NTIS. NTIS quality varies, but the reproduction on this one is actually pretty decent. I dimly recall them charging $70 or so. I paid it cheerfully. Henry Spencer at U of Toronto Zoology henry@zoo.toronto.edu utzoo!henry ------------------------------ Date: Tue, 7 Jan 92 11:00 CST From: jws3@engr.uark.edu (JW Smith) To: space-tech@daisy.learning.cs.cmu.edu Subject: Huzel and Huang For those who haven't seen it, the references are to a book called "Design of Liquid Propellant Rocket Engines" by Huzel and Huang. I have a copy which I received as part of my NASA training, and it's a darn good reference as well as a primer for those who want to know the state of the art. I do have a question though. My copy is 2nd edition (1971), which just about qualifies as state of the art in the US, admittedly, but I'd like to know if there's a newer edition out. Please email me if you've seen one. For the guy who wanted to know where to get this book, my copy says "For sale by the National Technical Information Service, Springfield, VA 22151." You might write to that address and see if anyone's home.... As for getting a free copy, well, you could apply for a job at NASA when the hiring freeze goes off. :) | James W. Smith, University of Arkansas | jws3@engr.uark.edu | | I'm so depressed. If I didn't have so much to do, I'd be a nihilist. | | Neither NASA nor the U of Ark. is responsible for what I say. Mea culpa. | ------------------------------ Date: Tue, 7 Jan 92 13:09:11 pst From: gwh@lurnix.LURNIX.COM (George William Herbert) To: space-tech@DAISY.LEARNING.CS.CMU.EDU Subject: LOX pressurization After some quick analysis last night, I've concluded that pressurizing the LOX tank by boiling LOX in a vessel surrounding the nozzle (therefore also cooling the nozzle...) is the most effecient low-cost pressurization system I can locate. It also will reduce the number of hard-to-handle items (it's structural, not systemic in nature, and _simple_. Plus, we don't need ablative lining anymore... I figure it will weigh a bit less 8-) -george (comments? 8-) gwh@ocf.berkeley.edu gwh@lurnix.lurnix.com ------------------------------ End of Space-tech Digest #95 *******************