Subject: Space-tech Digest #153 Contents: Business Week article on DC-X (1 msg) ISECCo update (1 msg) Comments on P2 (6 msgs) ------------------------------------------------------------ Date: Wed, 16 Jun 93 13:40 PDT From: jean@opus.dgi.com (J. Kim) To: space-tech@cs.cmu.edu Subject: Business Week mag. For those interested in a brief write up of DC-X and Gary Hudson, check out Business Week/June 21, 1993. -jean ------------------------------ Date: 17 Jun 1993 22:14:02 -0800 From: FSRRC@acad3.alaska.edu To: space-tech@cs.cmu.edu Subject: ISECCo Update: President goes wandering 6/17/93 ISECCo President Ray Collins (yours truly) is planning a tour of the States (and parts of Canada). The primary purpose of this trip is to meet and talk to space enthusiasts, with an emphasis on CELSS (Closed Ecological Life Support Systems). For those of you who aren't familiar with us and our project, ISECCo (the International Space Exploration and Colonization Co.) has undertaken the construction of a CELSS, which we call Nauvik (an Eskimo term meaning nurturing place) to help develop the technology needed for building CELSS in space. Though we expect down-to-eath applications (eg underwater habitats), our emphasis will be for use in space. Our biosphere design is nearing completion and we want to be sure we have all the major elements planned for before we fix the structure design. As our design currently stands we have a 40' (12m) diameter dome which will be buried completely underground. It will have an airlock attached to the outside and on the other end there will be an emergency escape hatch. Inside there will be 2.5 floors. The bottom floor will be devoted to crop growth, aquaculture (fish) and any critters we choose to raise for meat, like chickens and rabbits. In the center a column of large, doughnut-shaped tables will be stacked vertically 4' apart all the way to the peak of the dome. These tables will be used exclusively for short crops (eg carrots). The second floor will be primarily a living area. This will not be very large for the inward curve of the dome is reaching in toward the central column of tables. The top floor will be smaller yet (not to mention quite low). It will be used to house our equipment such as dehumidifiers, water tanks and air conditioners. I am eager to meet with anyone who is interested in space exploration. My itinerary has not yet been set, (it will be pretty much determined as I go along) b ut I am going to pretty well cover the eastern states. I will also be venturing to such places as Houston, Tucson and Denver, and may (time permitting) visit the West coast. But even if I don't get to visit with you in person I would very much like to hear from you, so even if you don't live near my travel route do send a note! Since I shall be spending ISECCo money on this trip (though I would like to point out that it originated from my wallet!) I shall be traveling by bus and if you want me to visit a place to stay (floor space is fine!) would be appreciated. Unfortunately I can't give exact times because my schedule will vary (depending on meetings in each stop). I will be posting updates regularly though, so you can track my progress and if it looks like I will pass near you send me your phone # and I'll give you a call! Currently my plan is to depart Fairbanks June 25th, spend the 26th in Anchorage and then head for Canada. I expect to arrive in Edmonton around the first of July (plus or minus a couple of days). From there I hesitate to predict arrival times, so I'll wait on that until I'm under way. (I'll be checking in to this computer network about once or twice a week.) Those of you who want more general--or specific--information about us feel free to write. My e_mail address is FSRRC@ACAD3.ALASKA.EDU or (on Compuserve) 74010,3722. A regular postal address is below. Please include a postal mailing address for all initial correspondence--we occasionally have letters whose computer return address fails for one reason or another. --Ray :: President, ISECCo :::The International Space Exploration and Colonization Company::: :::P.O. Box 60885::Fairbanks::Alaska::99706::: 907/457-2674 Researching and Developing space oriented technology for the betterment of mankind. * * * * * * * * * * * * {end} ------------------------------ To: Bruce_Dunn@mindlink.bc.ca, space-tech@cs.cmu.edu Subject: Re: P2 2/5 Design Description Date: Wed, 16 Jun 93 02:09:23 -0400 From: Randy Appleton Here's a question or two on the P2 booster: - In the expanded P2, the 'core' stage has an extended tank, and fires for much longer than normal. This either means that it needs a different engine than the normal first stage, or that the normal engine is somewhat over-designed. Would it be possible to delay the ignition of the core stage until the boosters are expended? - There seems to be no accounting for fuel that sits in the bottom of the tanks unburnt. I know I cannot run my car all the way to empty but must have some small resurve. Should your booster? -Randy ------------------------------ Date: Fri, 18 Jun 93 14:36:58 -0400 From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: comments on P2 Here are some comments and questions about the latest P2 design... One thing I don't remember us talking about it passivation of the peroxide tank. Iron and other transition metals are catalysts for peroxide decomposition. I don't think it's going to be obviously ok to just pour peroxide into a field-welded tank. At the least, the tank should be coated with some passivating layer. Thoughts? One possibility would be plating some inert metal (is cadmium inert with peroxide?); this could be done by filling the tank with plating solution and inserting an electrode through an opening. A more general problem I have with the design is that design decisions (for example, use of the particular high strength steel chosen) are being justified by the fact that they increase payload. This is dangerous: you want to optimize on cost/pound. It very well may be that stronger, more expensive steels are worthwhile, but I would like to see an explicit argument why a larger, dumber first stage would not be cheaper. I like the move to lower pressure. As an aside: if you went to even lower pressure, you could perhaps use self-pressurizing propane. Alternately, RP-1 + a low molecular weight gas (methane?) could be self-pressurizing with relatively low mass overhead. Now that you are using a custom nozzle, how about going to a plug nozzle using the bottom of the oxidizer tank to either support or actually be the plug (suitably coated with ablator)? This at the very least makes the "space frame" support for the nozzle much more compact, and gives some altitude compensation. This idea is reminiscent of how Amroc's industrial launch vehicle is/was going to have nozzles clustered around the LOX tank. Paul F. Dietz dietz@cs.rochester.edu "Never have so many done so little for so much." ------------------------------ Date: Fri, 18 Jun 93 15:33:20 -0400 From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: more P2 comments Bruce says: > chamber pressure was varied, keeping a constant 1 MPa of pressure > available for injector pressure drop. It is not clear to me why the injector pressure drop should remain constant. A lower chamber pressure engine will require a larger throat and chamber to achieve the same thrust, so it can have a larger injector. A larger injector can have more or larger orifices, which should reduce the pressure drop for a given propellant flow rate. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Fri, 18 Jun 93 23:06:23 -0400 From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: another idea for P2 Some additional ideas for P2... During the last go-round on this, I was taken by the idea of molten zinc as a propellant. With peroxide, it has a density of about 3 g/cc or better. Even with rather inefficient utilization in a rocket, it would have a good density-Isp product. The idea here is to add a third tank containing molten zinc. The engine would be designed to run on zinc/peroxide at liftoff, switching to RP1/peroxide later (or, perhaps, running on all three propellants at once with a gradually shifting ratios). Molten zinc would be at about 400 C. At this temperature, the zinc tank could be pressurized to 4 MPa with steam (the steam being produced by pouring water onto the zinc). After the zinc is exhausted (when the vehicle is going at roughly the exhaust velocity of the Zn/peroxide combination), the valve would be shut off, and the Zn tank slightly overpressurized. The steam would then be used to heat helium in the other two tanks. I think there is enough energy in the steam to be interesting. If zinc has too high a vapor pressure at this temperature, or is too reactive with steam, this idea has problems. It also would not apply to the second stage. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Mon, 21 Jun 1993 08:53:15 -0400 (EDT) From: Ted_Anderson@transarc.com To: Bruce Dunn Subject: Re: P2 1/5 Introduction Cc: space-tech@cs.cmu.edu This is a very nice design. I especially appreciate the comparisons with DC-1, Ariane, etc. Your comment that there are no moving parts except for values really caught my attention! I have been aware for a long time of the argument that the basic fleet of expendables used in the US is ICBM derived technology from the 50s. I guess I had assumed that rockets like Ariane sort of suggest that even a full redesign for commercial purposes couldn't do much better. Otherwise, Ariane would be undercutting Delta more aggressively (why settle for 50% of market share?). As a conseqence I've sort of disregarded that sort of argument, in spite of earlier BDB suggestions. This design proposal suggests that efforts like Ariane and even the Japanese H-2 are geared at something else besides commercial space launcher development! I must be crazy to think that goverment pork is something restricted to the US. The scales drop from my eyes... Keep up the good work. Thanks, Ted Anderson ------------------------------ Date: Mon, 21 Jun 93 08:22 PDT To: space-tech@cs.cmu.edu Subject: P2: Replies to Paul Dietz From: Bruce_Dunn@mindlink.bc.ca (Bruce Dunn) Ever inventive, Paul Dietz offers a slew of suggestions about the P2 vehicle. I will try my best to respond: > One thing I don't remember us talking about it passivation of the > peroxide tank. Iron and other transition metals are catalysts for > peroxide decomposition. I don't think it's going to be obviously > ok to just pour peroxide into a field-welded tank. At the least, > the tank should be coated with some passivating layer. Thoughts? > One possibility would be plating some inert metal (is cadmium inert > with peroxide?); this could be done by filling the tank with plating > solution and inserting an electrode through an opening. I agree that the basic tank, right from the welding shop is likely to catalyze peroxide decomposition. I don't know what common plating metals would be best for passivation of the surface. In my lab, we have 50% peroxide in polyethylene plastic bottles, with stability times of years. A plastic lining to the tank, or the moral equivalent is probably what is needed. This could consist of painting the inside of the tank with an epoxy resin or other paint, or perhaps simply coating the steel with a thin layer of paraffin wax. I regard this as a non-show-stopping detail to be worked out later. > A more general problem I have with the design is that design decisions > (for example, use of the particular high strength steel chosen) are > being justified by the fact that they increase payload. This is > dangerous: you want to optimize on cost/pound. It very well may be > that stronger, more expensive steels are worthwhile, but I would like > to see an explicit argument why a larger, dumber first stage would not > be cheaper. I agree that minimizing cost per unit payload is the primary design criterion. For some design choices, I am not sure what would result in the lowest cost. However, the use of lower strength steels looks highly unattractive, at least for this two stage vehicle (see below for comments on stage number). The D6aC steel proposed is not an exotic high alloy steel, as was the 18Ni250 originally used in the P2 design. Even if the steel is several times the cost of lower strength steels it would pay to use it, as steel costs are likely to be only a smallish fraction of the total vehicle costs. If the D6aC steel comes in at $10 per kg (several times what 100,000 psi steel costs according to George), the 38 metric tons of steel in the P2 vehicle will cost $380,000. The roughly 800 tons of hydrogen peroxide, at $2 per kg, has a cost of $1,600,000. Using steel at 100,000 psi strength, as proposed by George, results in a vehicle over double the size for a given payload. Even if the lower strength steel were given to the designer for free, the extra propellant alone for the larger vehicle would cost many times the savings realized. > I like the move to lower pressure. As an aside: if you went to even > lower pressure, you could perhaps use self-pressurizing propane. > Alternately, RP-1 + a low molecular weight gas (methane?) could be > self-pressurizing with relatively low mass overhead. I don't think that propane has enough vapor pressure. On a warm Florida afternoon, it will only self pressurize to about 1 MPa. This is just enough for the injector pressure drop (see below), leaving nothing much for chamber pressure. Hydrocarbon mixtures with an appreciable percentage of lower molecular weight gas might be ok, as would ethane (a liquid with a vapor pressure of 4 MPa at room temperature). However, all these suffer greatly from having a high molecular weight pressurizing gas, leaving a lot of mass in the tank at burnout, unless the gas could be burned with oxidizer. Because of their content of low density hydrocarbons as well, tanks will be bigger and heavier. Finally, having the fuel self pressurize doesn't solve the oxidizer pressurization problem. > Now that you are using a custom nozzle, how about going to a plug > nozzle using the bottom of the oxidizer tank to either support or > actually be the plug (suitably coated with ablator)? This at the very > least makes the "space frame" support for the nozzle much more > compact, and gives some altitude compensation. This idea is > reminiscent of how Amroc's industrial launch vehicle is/was going to > have nozzles clustered around the LOX tank. In some ways, using an aerospike engine is attractive, particularly for the lower stage where there is a loss of Isp due to the compromise nature of the engine expansion ratio. There are some potentially neat packaging benefits as well - the liquid helium tank and associated equipment could fit inside the aerospike nozzle, shortening the interstage area. I haven't however proposed an aerospike engine for a number of reasons: 1) It presents a perceived developmental risk, which is the last thing that I want in the vehicle. This is aside from the problem of whether there is indeed a developmental risk. 2) I don't have the software to be able to properly predict the Isp of the engine. This isn't so much a real impediment to using an aerospike engine as an impediment to my proposing to use one. 3) The improvement in Isp is mainly for the lower stage, and this does not have a very strong impact on the overall payload for the vehicle. Regarding my comment that I assumed a constant 1 MPa pressure drop for the injectors, no matter what the chamber pressure, Paul writes: > It is not clear to me why the injector pressure drop should remain > constant. A lower chamber pressure engine will require a larger > throat and chamber to achieve the same thrust, so it can have a larger > injector. A larger injector can have more or larger orifices, which > should reduce the pressure drop for a given propellant flow rate Unfortunately, this is like arguing that because a fire hose has a larger diameter than a garden hose, it needs only a fraction of the water pressure! The function of the injector is to spray or otherwise introduce the propellant into the chamber in finely divided droplets to promote rapid reaction. The pressure differential needed to get adequate spray velocity and droplet formation is independent of the chamber pressure itself. The figure of 1 MPa is taken from some literature on conventional LOX/hydrocarbon engines, where pressure is "cheap" because pumps are used. There may well be injectors which do a reasonable job with a lower pressure drop. If so, they would be usable at all chamber pressures. > During the last go-round on this, I was taken by the idea of molten > zinc as a propellant. With peroxide, it has a density of about 3 g/cc > or better. Even with rather inefficient utilization in a rocket, it > would have a good density-Isp product. > > The idea here is to add a third tank containing molten zinc. The > engine would be designed to run on zinc/peroxide at liftoff, switching > to RP1/peroxide later (or, perhaps, running on all three propellants > at once with a gradually shifting ratios). > > Molten zinc would be at about 400 C. At this temperature, the zinc > tank could be pressurized to 4 MPa with steam (the steam being > produced by pouring water onto the zinc). After the zinc is exhausted > (when the vehicle is going at roughly the exhaust velocity of the > Zn/peroxide combination), the valve would be shut off, and the Zn tank > slightly overpressurized. The steam would then be used to heat helium > in the other two tanks. I think there is enough energy in the > steam to be interesting. > > If zinc has too high a vapor pressure at this temperature, or is > too reactive with steam, this idea has problems. It also would not > apply to the second stage. I am not too keen on complicated cycles such as this, even if the zinc-peroxide engine were a developed technology with a proven and reasonable Isp. The issue is not can it be made to work, but is there any advantage to this over simply making the peroxide/RP-1 stage bigger. > Why two stages? A three stage launcher would have 50% more engines, > but it could also be built with a crude first stage, as it would be > thrown away sooner. It could be that the savings in cheaper materials > and larger margins would make up for the need for another engine + > control system, although perhaps even that could be avoided by doing > some kind of parallel staging. The selection of the number of stages certainly deserves thought. As mentioned in the postings, I think that a single stage is too difficult a proposition for an expendable vehicle. With the mass fractions and Isp that I am working with, a single stage can't get to orbit by itself, let alone carry any payload. Mitchell Burnside Clapp thinks that a pressure fed peroxide /JP-5 vehicle can get to orbit with a single stage. This however assumes abandoning low cost construction in order to lower vehicle dry mass. Two stages appear to be the minimum to get to LEO. If I design a three stage vehicle for the same payload, I can get to orbit with a total vehicle mass of 550 metric tons rather than 1000. However, in order to save the 450 tons, which is mostly cheap steel, peroxide and fuel, I have had to increase the number of engines, helium supply systems and tank sets by 50%, and double the number of interstage structures, stage separation maneuvers and in-flight engine ignitions. Granted, the engines are somewhat smaller than they would be otherwise, but this does not necessarily make them that much less expensive. Alternately, I could stick with the 1000 metric ton vehicle, but make it three stages and a lower level of technology. The counter argument to this is that I think that the vehicle is already likely to be very cheap to build, and that there are unlikely to be enough cost savings from the "lower level of technology" to pay for the complications and increased part count of a three stage vehicle. This type of argument will only be settled by detailed cost calculations which include not only the production costs of the vehicle, but also development and operations costs. I think that I am on one side of the argument (using D6aC steel but only two stages), while George Herbert is on the other side (conventional steel, but three (or more?) stages). Finally I should also point out that for the purposes of geosynchronous transfer orbits, the P2 is already a three stage vehicle, with a solid upper stage. Everything else being equal the two stage vehicle will be more reliable. The three stage vehicle has at least 50% more critical subsystems than a two stage system. When I think about the Ariane 44L vehicle, I shudder. It has no fewer than 10 engines (four liquid fueled boosters, four bottom stage engines, a second stage engine, and a third stage engine). There are 6 separations which must take place flawlessly (the four boosters from the first stage, the second from the first, and the third from the second). The failure of any one engine or staging is likely fatal. Even the first stage with 4 engines does not have complete engine out capability. An Ariane was lost when 4 of the four quit partway through atmospheric flight and the remaining engines could not vector their thrust enough to keep the vehicle from being pushed at an angle through the atmosphere (it broke up structurally because of aerodynamic forces). > A comment on gas pressurization... > > You have the tanks pressurized with helium at full pressure until > 2/3 of the propellant is gone, then blowing down adiabatically. > > You could get extra performance (or, tolerate a heavier pressurization > system) if you designed the pressurization system to be dropped at > this point, when it is no longer needed (this would be at about 1/2 > the delta-V for the stage, I think). You needn't even include solid > rockets to do the separation; you could exploit residual gas pressure. > That is also true for separating parallel boosters in your heavy lift > variant. You would need good check valves to prevent escape of gas, > as well as a means of breaking the connection (explosive?), but > you need check valves anyway to prevent mixing of propellants. I have to listen to Paul about pressurization, because he set me straight on the importance of not assuming that gas expansion was iso-thermal. The advantages of getting rid of the helium pressurization system are very small. The total mass involved is only about 3 tons for the lower stage, compared with a total burnout mass of approximately 70 tons. I didn't want to rewrite my spreadsheets to accommodate dropping dry mass partway through a flight, so I ran a quick calculation of what the payload would be if the helium system weight nothing at all, right from liftoff. You pick up only about 1.2 tons of extra payload. Perhaps half of this might be achievable by discarding the helium pressurization system at mid burn. The gain is thus modest, even before subtracting the mass of the equipment and modifications necessary to make the system drop correctly. An even bigger problem is where to put the large spherical helium tank so that it can be dropped easily. It is currently in the interstage area at the top of each stage, and can't be dropped from there without some sort of Rube Goldberg system of doors, ejector rails and explosive cartridges. It would be possible to make the helium tank cylindrical and build it and its associated equipment into a streamlined pod strapped to the side of the booster (similar to the N2O4 supply system for the Titan solid rocket boosters). The extra mass incurred by having a non-spherical tank would be paid for by the shortened interstage. One would then have to figure in the extra drag to see if the system is workable, even with the advantage of droppability. This configuration would not be so easy to do with the upper stage in its current semi-conical shape. > You mentioned that use of steel with 100,000 psi yield strength > resulted in lower performance. Did this calculation assume both > stages used the lower strength steel? It would be interesting > to see the results for the lower stage only (even if just the > oxidizer tank is made of the heavier material, so that the large > spherical tank remains common). I had assumed that the lower strength steel would be used throughout the vehicle, in order to maintain commonalty. Here are some results of various assumptions, in terms of percent original payload to orbit: 100,000 psi steel used in fraction of original payload lower stage oxidizer tank 87% both lower stage tanks 85% upper stage only 57% both stages 41% Clearly, the use of high strength steel is more critical for the upper stage than for the lower. -- Bruce Dunn Vancouver, Canada Bruce_Dunn@mindlink.bc.ca ------------------------------ End of Space-tech Digest #153 *******************