Subject: Space-tech Digest #112 Contents: Re: P2 Launcher (8 msgs) Delta Clipper refs. (1 msg) ------------------------------------------------------------ Date: Tue, 31 Mar 92 11:25:19 -0800 From: George William Herbert To: space-tech@cs.cmu.edu Subject: Hydrogen Peroxide Pricing I just got a quote of $0.48/lb for bulk quantities of H2O2 70%, in tankers delivered to Florida from New Jersey. The people I talked to got nervous at the proposal of shipping 80% and 90%, though these were sales people. If we had to distill 95%-100% from the 70%, it looks like total price including the processing would run around $0.75/lb. I would guess that truly large lots (500 to 1000+ tons) would bring the price down another 10%. Still fairly uncheap. That's about the price of the vehicle structure I was looking at for my hybrid, and it's probably not cost effecient to do it that way. I'm going to go grind some numbers and check, though. -george william herbert gwh@lurnix.com gwh@ocf.berkeley.edu ------------------------------ From: henry@zoo.toronto.edu Date: Tue, 31 Mar 92 18:32:31 EST To: space-tech@cs.cmu.edu Subject: Re: P2 Technical Issues >Trying to employ the cryogenic strength of the steel would require that the >whole tank be insulated so that no part of the wall warms up... If memory serves, LOX tanks tend to be self-insulating because they form a layer of frost on the outer surface, at the cost of some initial boiloff. Of course, the frost probably won't stick around after liftoff. The LOX tank on an Atlas is some flavor of stainless steel, with LOX on one side and the outside on the other. However, it's thin enough that the whole thing can probably be assumed to be at LOX temperatures, which presumably wouldn't be the case with more normal wall thickness. Max Hunter's SSX proposal used propane plus LOX, with the propane chilled to LOX temperatures to increase its density (and incidentally avoid an insulated intertank area). Henry Spencer at U of Toronto Zoology henry@zoo.toronto.edu utzoo!henry ------------------------------ Date: Thu, 2 Apr 92 10:01:26 -0500 From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: Re: P2 Launcher Some more comments on he P2 launcher... I have some problems with the choice of the shuttle SRM nozzles. (1) The nozzles have a lousy expansion ratio. At 900 psia chamber pressure, one could boost the sea level Isp considerably by going to optimal expansion ratio. The 20-1 ratio for the second stage is also too low. At higher expansion ratio, denser hydrocarbon fuels like RP-1 may be more advantageous, as their performance relative to more hydrogen-rich fuels improves as the nozzle exit temperature declines. (2) The chamber pressure may be too high. High chamber pressure --> higher tank pressure --> heavier/more expensive tanks & more pressurization gas. Lower chamber pressure does reduce mass flow, but this could be compensated for by using a bigger nozzle. (3) The nozzle does not use a regeneratively cooled throat or upper nozzle section. Doing so could reduce nozzle weight. By using the heat to vaporize pressurant gases, complexity coul be reduced as well (with possible startup problems, as Bruce noted). (4) The nozzle is mechanically complex. It would be simpler, I think, to use several smaller engines, steered by throttling, or to include a number of vernier rockets for steering. (5) The nozzle is designed for use with solid rocket exhaust, which is hot, fiercely radiative, and contains erosive particulates. Peroxide/propane exhaust will contain no aluminum oxide, and, because of the high hydrogen/carbon ratio, will not contain much soot. The thermal load on the nozzle should be less. This suggests to me that the SRM nozzles are overdesigned for this task. (6) The use of SRM nozzles forces the vehicle to be too big. A payload of around 10 tons is more in line with the potential market. Scaling the launcher down but holding chamber pressure fixed also improves the thrust/weight ratio (square/cube law). This would push the design towards higher expansion ratios or lower chamber pressures as vehicle size declines. I consider (6) to be the most serious criticism. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Thu, 2 Apr 92 22:05 PST To: space-tech@cs.cmu.edu Subject: P2 Nozzle From: Bruce_Dunn@mindlink.bc.ca (Bruce Dunn) Paul Deitz writes: > I have some problems with the choice of the shuttle SRM nozzles. > > (1) The nozzles have a lousy expansion ratio. At 900 psia chamber > pressure, one could boost the sea level Isp considerably > by going to optimal expansion ratio. The 20-1 ratio for the > second stage is also too low. At higher expansion ratio, denser > hydrocarbon fuels like RP-1 may be more advantageous, as their > performance relative to more hydrogen-rich fuels improves as the > nozzle exit temperature declines. The approximately 7.5 expansion ratio quoted for the P2 booster is that of the standard SRM nozzle. It is not clear to me if this expansion ratio is optimum for the SRM, or is a compromise forced upon the SRM designers who got caught between a minimum size for the throat (to get sufficient thrust) and a maximum exit cone diameter (forced by the slimness of the SRBs). From some published literature, for RP-1 and LOX the theoretically correct expansion ratio for 6 MPa to atmospheric looks like about 9.3. In practice, it is desirable to have a slightly overexpanded engine at sea level (as long as there is no flow separation) with the consequent loss of Isp made up for by a higher Isp later in the flight. The F1 RP-1/LOX engine has a theoretical optimum expansion ratio which looks like it should be about 11, but the engine itself is built with a ratio of 16 to 1. Paul is no doubt correct in that the expansion ratio of 7.5 is less than optimum, but how much bigger it could be is not clear to me without some calculations based on propane/peroxide exhaust products, and some knowledge of how much overexpansion could be tolerated. I am leery of relying too much on with analogies with RP-1/LOX engines, as the products of combustion of hydrocarbons with peroxide are different than those of hydrocarbons with LOX (see below). In the absence of this information, performance has been conservatively calculated using the standard SRB expansion ratio. This doesn't mean that the first stage nozzle couldn't be extended if a trade study showed that the cost of making modifications to the standard SRB nozzle were worth it in terms of improved performance. For the second stage, it is clearly desirable to have the expansion ratio as high as possible. This has been done firstly by restricting the throat. There is a limit to how much the throat can be restricted however, as restriction lowers thrust. The lower the thrust, the higher the gravity losses. For preliminary performance estimates, the throat was sized to give the second stage a thrust/mass ratio of 1.2 at ignition. The exit cone of the nozzle was set at 5 meters, which makes it only a meter less than the diameter of the vehicle. A further extension might be possible, but this seemed a reasonable limit to me which would allow the second stage tank to be directly supported by the launch pad during vehicle mating and stacking operations (with the nozzle projecting through a hole in the platform). > (2) The chamber pressure may be too high. High chamber pressure --> > higher tank pressure --> heavier/more expensive tanks & more > pressurization gas. Lower chamber pressure does reduce mass > flow, but this could be compensated for by using a bigger nozzle. Particularly with pressure fed rockets, there is a complex relationship between mass, thrust and Isp. The desire to lower chamber pressure and tank mass directly conflicts with the wish to have a high expansion ratio and a high Isp. When designing a completely new vehicle, the chamber pressure would be a free parameter, with the engine being optimized for each pressure. Gary Hudson's design for the Liberty pressure fed LOX/RP-1 booster used a first stage engine with a chamber pressure of 250 psi. This low pressure kept the tanks light (fabricated from aluminum, flight pressure 360 psi, hydrotested at 375 psi, calculated yield point at 420 psi). However, he paid for this lightweight tank by having a first stage engine with a low expansion ratio and an indifferent Isp. Different design choices could have led to a similar performing vehicle with higher pressure tanks, a better expansion ratio engine, and a higher Isp. The P2 is designed to use an existing nozzle - this limits the ability to trade off pressure (and tank mass) vs. engine efficiency. If a designer is required to use a certain size engine throat because the engine is off-the-shelf, then if the chamber pressure is cut by a factor of 2, to a first approximation the thrust will be cut by a factor of 2. Since the rocket must still take off the ground, this immediately cuts the total vehicle mass by a factor of 2 with a corresponding effect on the payload. To make matters worse, when the chamber pressure is lowered, the expansion ratio must be lowered also (for a first stage engine), killing Isp. As a final insult, the engine mass is unlikely to drop by a factor of two when the pressure is lowered by 2, and so the stage propellant fraction suffers as well. > (3) The nozzle does not use a regeneratively cooled throat or upper > nozzle section. Doing so could reduce nozzle weight. By using > the heat to vaporize pressurant gases, complexity could be reduced > as well (with possible startup problems, as Bruce noted). This agains revolves around using an existing nozzle to reduce development costs. It would be perfectly feasible to design a regeneratively cooled throat and upper nozzle, and would probably lead to some weight savings. However, the delivered payload is not particularly sensitive to the mass of the lower stage boosters. The current throat and nozzle have a mass of 10 tons. Part of this mass probably reflects the high degree of mechanical strength needed for the nozzle to survive as it hits the water during recovery (and could probably be pared down if it was decided to expend the first stage boosters). Assuming that a completely new nozzle assembly could be produced that had a mass of 8 tons rather than 10 tons, the benefits for this would be only about a 1.1 % increase in payload. In a trade study, the issue becomes not whether the new nozzle would increase the payload, but whether the money spend on the development of a new nozzle would be better off spent on improving something else in the vehicle. If the vehicle has excess payload capacity, then the increase in ultimate performance may lead to no savings at all for lower mass payloads (see below for comments on payload capacity). For the sake of low development costs and getting a working vehicle quickly, I think that it is simplest to use something very close to a stock SRB nozzle. I am not convinced that pressurization by robbing heat from the combustion chamber, throat and upper nozzle is any simpler or lighter than using a gas generator. I think that it is simplest to generate the heat by burning a small proportion of the pressurizing hydrogen with peroxide (catalytic peroxide decomposition followed by hydrogen injection into the hot oxygen should give autoignition) and quenching the exhaust products with liquid hydrogen (or if helium or nitrogen is used, to burn propane with peroxide and quench the exhaust with the liquid). The gas generator approach would certainly be easier to develop, as the gas generator could be tested in isolation without running the engine. This was one of the major problems in SSME development - the various components could not easily be tested in isolation. > (4) The nozzle is mechanically complex. It would be simpler, I think, > to use several smaller engines, steered by throttling, or to include > a number of vernier rockets for steering. The nozzle is not particularly complex. It has two actuators and a flexible joint based on steel and rubber laminations. Building it uses different technologies than building a regeneratively cooled nozzle, but not necessarily a more difficult technologies. If I remember correctly, there have been some concerns about insulation burn-through and some modifications have been made in later production versions (does anyone have the details?). However, generally the nozzle has proved itself and has the outstanding virtue that it has been repeatedly tested on real functioning rockets and not just in some computer simulation. Using several smaller engines and steering by throttling probably would raise complexity, as this would require a large multiplication in the number of valves and propellant feed lines. This in turn leads to multiplying the number of failure points. In such a system, failure of any one engine leads to total vehicle failure at many stages of the flight. Remember the Ariane failure that was due to reduced thrust from only one of its four main stage engines. > (5) The nozzle is designed for use with solid rocket exhaust, which > is hot, fiercely radiative, and contains erosive particulates. > Peroxide/propane exhaust will contain no aluminum oxide, and, > because of the high hydrogen/carbon ratio, will not contain much > soot. The thermal load on the nozzle should be less. This suggests > to me that the SRM nozzles are overdesigned for this task. I fully agree with this. Not only is peroxide/propane a very gentle combination relative to the corrosive standstorm that comes out of a solid motor, it is also considerably gentler than RP-1/LOX. Hydrocarbon/peroxide gives optimum Isp when near stochiometric ratios are used. In spite of not being fuel-rich, the combustion temperature is only about 3000 K, vs. about 3700 for RP-1/LOX. This is because each oxygen atom used for combustion is accompanied by a water molecule which lowers the exhaust molecular weight without adding much energy. This gives reasonable Isp even with low temperature. Once again, the use of a standard SRB nozzle in preliminary performance estimates was designed to be conservative. In practice, quite a bit of the ablative material could probably removed from the nozzle without affecting safety. > (6) The use of SRM nozzles forces the vehicle to be too big. A payload > of around 10 tons is more in line with the potential market. Too big? Not necessarily: 1) The size of the average "commercial" payload has been creeping up through the years. Commercial launch vehicles have been forced into a long series of upgrades to keep up with the increasing payload demands. Tanks are lengthened, engines are upgraded, and solid boosters are strapped on. All this has cost money. 10 tons may be about right now, but 10 to 20 years from now it may seem entirely inadequate. Don't laugh at the time scale - what is the age of the original Atlas, Delta, and Titan designs (and what were their original payloads?) Saying that a 10 ton payload is adequate reminds me somewhat of Microsoft who in their wisdom decided that MS-DOS machines were unlikely to ever have more than 640K of RAM (Paul, this is not a flame - your comments have been very helpful in getting me to think about certain issues). If the vehicle is cheap enough, it won't matter that it could have been made cheaper still by making it smaller. In fact having excess payload capacity may be an important factor in keeping a launcher inexpensive. With a weight margin to play with, components won't need to be as highly stressed, and less expensive materials and designs can be used. For example the question of whether to use an inexpensive medium strength steel or an expensive high strength steel for the tanks is an open question for the P2 (or at least it is after a bit of gentle prodding by George Herbert). With a payload margin to play with, the cheaper but heavier material is an option. 2) The fact that current payloads are about 10 tons to LEO may have more to do with current launch vehicles than with the most desirable mass for the payloads. I would be suprised if direct broadcast TV satellites weren't easier to build if designers were not so concerned with mass and size limitations. Given a chance to cheaply launch 10 or 15 ton satellites into geosynchronous orbit, designers might well welcome the option (is there anyone with comsat design experience who could comment on this?). 3) If we ever get a space station, spare capacity on a P2 flight could be used for consumables for the station. It would only require giving the third stage on-orbit maneuvering capability. Any unused payload capability could be devoted to the delivery of fuel, water, air etc. A lot of the cost of running the space station, if it ever gets built, will be in getting mundane consumables to it. 4) If a payload of 10 tons or so is all that is really required, the P2 vehicle doesn't have to be downsized, it merely has to be taken apart and reassembled. Take one of the standard booster stages as described in P2 technical description (about 600 tons fueled), remove the parachutes, and add a pump fed second stage (a close relative of the 200 ton third stage used in the full P2 vehicle). Launch as a two stage vehicle with both stages expended. Using a UDMH/N2O4 upper stage and a Viking IV engine gives the following: Mini P2 Capability Second stage size (metric tons gross) LEO Payload (metric tons) 38 (stock Ariane IV second stage) 7.8 50 (stretched Ariane IV second stage) 9.0 75 (same propulsion, new wider tanks) 11.0 100 (ditto) 12.4 The calculated payloads assume a stock P2 stage running with a chamber pressure of 6 MPa. At this pressure, the thrust to mass ratio is high and takeoff is rather brisk (about 1.6 to 1.8 G initial). Throttling will be required later in the first stage burn to keep the G loading reasonable, and throttling may be required earlier in the flight to lower aerodynamic loads while in the atmosphere. If such extensive throttling is needed, it may prove advantageous to lower the tank pressure and tank mass of the first stage (bearing in mind that this will tend to kill the allowable expansion ratio and Isp). If one looks solely at liftoff mass, a Mini P2 looks like a gross waste of resources relative to a similar performance (700 tons vs. about 450). However it looks positively simple when counting parts and failure points (2 engines, two tank sets and one stage separation vs. 10 engines, 7 tank sets and 6 separations). Pad operations are likely to be a lot simpler for the P2 and I think that the vehicle itself is likely to be less expensive. There is no reason that a Mini P2 couldn't be a strong commercial success, and have a good reliability record to boot. Disassembling a full P2 to produce a Mini P2 is in effect of the opposite of the current proposals to create a 50 ton capable heavy launchers by clustering a number of medium launchers. > Scaling the launcher down but holding chamber > pressure fixed also improves the thrust/weight ratio (square/cube law). This > would push the design towards higher expansion ratios or lower chamber > pressures as vehicle size declines An interesting point. Higher expansion ratios for lower stages would be ruled out by the problems of overexpansion, but would be the obvious way to go for upper stages. Lower chamber pressures make sense for first stages. As noted above, Gary Hudson settled on a chamber pressure of 250 psi (about 1.7 MPa) for a vehicle with a gross liftoff mass of about 25 tons. The P2 seems to a reasonable design at 6 MPa, but would clearly be even more efficient at high pressures. Provided that the combustion chamber and nozzle can take the heat and pressure, raising the tank pressure by say 50% would not only give 50% more thrust and allow a 50% larger vehicle, but would give a even better mass ratio (the non-tank masses such as the engine don't change when scaling the tanks and propellants up) and allow for a better expansion ratio for first stage nozzles. -- Bruce Dunn Vancouver, Canada Bruce_Dunn@mindlink.bc.ca ------------------------------ Reply-To: davidsen@crdos1.crd.ge.com Date: Fri, 3 Apr 92 12:37:45 EST From: davidsen@crdos1.crd.ge.com To: space-tech@cs.cmu.edu Subject: Re: P2 Launcher Sender: mnr@DAISY.LEARNING.CS.CMU.EDU > (2) The chamber pressure may be too high. High chamber pressure --> > higher tank pressure --> heavier/more expensive tanks & more > pressurization gas. Lower chamber pressure does reduce mass > flow, but this could be compensated for by using a bigger nozzle. I'm not sure I totally understand this point. The tank pressure needs to be adequate to deliver fuel to the injection pumps and should be related to input area and volume. The chamber pressure determines what you need in the way of fuel delivery pressure at the outlet of the pump. Am I missing something? I freely admit to not being an expert here, but I have played with pumping fuels into chambers (GE makes jet engines and gas turbines) so I know that nozzle and tank pressures are relatively independent. > > (3) The nozzle does not use a regeneratively cooled throat or upper > nozzle section. Doing so could reduce nozzle weight. By using > the heat to vaporize pressurant gases, complexity coul be reduced > as well (with possible startup problems, as Bruce noted). What implications does this have in terms of reliability and reusability? Seems like a throwaway "chunk" would be cheaper in terms of reliability and $/lb payload, but that's just a gut feeling, because I don't have cost and reliability figures for cooled nozzles. > (4) The nozzle is mechanically complex. It would be simpler, I think, > to use several smaller engines, steered by throttling, or to include > a number of vernier rockets for steering. Reliability? Is one steerable nozzle cheaper overall (and possibly more reliable) than one simple nozzle plus N gymbols and/or throttle valves? Gut feeling is that one would be more reliable and that overall weight would be about a wash. > (5) The nozzle is designed for use with solid rocket exhaust, which > is hot, fiercely radiative, and contains erosive particulates. > Peroxide/propane exhaust will contain no aluminum oxide, and, > because of the high hydrogen/carbon ratio, will not contain much > soot. The thermal load on the nozzle should be less. This suggests > to me that the SRM nozzles are overdesigned for this task. On the other hand it's designed, tested, and in production. > (6) The use of SRM nozzles forces the vehicle to be too big. A payload > of around 10 tons is more in line with the potential market. Scaling > the launcher down but holding chamber pressure fixed also improves > the thrust/weight ratio (square/cube law). This would push the > design towards higher expansion ratios or lower chamber pressures as > vehicle size declines. As long as there is enough demand for launches and the ability to handle multiple payloads, I think the big capacity is a plus, since it allows competition in the small job category as well as offering a large job capacity which would allow ideas not currently feasible. One of the things I like about the design is the reuse factor. By having two units 1st stage and only one 2nd, it allows units which have serious problems to be turned into parts instead of refurbed at high cost. It could allow building of units from "spare parts," if the ecconomics of selecting and assembling the parts were correct. ------------------------------ Date: Fri, 3 Apr 92 14:29:53 -0500 From: dietz@cs.rochester.edu To: davidsen@crdos1.crd.ge.com Subject: Re: P2 Launcher Cc: space-tech@cs.cmu.edu > (2) The chamber pressure may be too high. High chamber pressure --> > higher tank pressure --> heavier/more expensive tanks & more > pressurization gas. Lower chamber pressure does reduce mass > flow, but this could be compensated for by using a bigger nozzle. I'm not sure I totally understand this point. The tank pressure needs to be adequate to deliver fuel to the injection pumps and should be related to input area and volume. The chamber pressure determines what you need in the way of fuel delivery pressure at the outlet of the pump. The P2 concept's first and second stages are *pressure fed*. There are no pumps. Rather, the propellant tanks are at higher pressure than the combustion chamber (7.5 vs 6 MPa). This is simple but requires heavy tanks. > As long as there is enough demand for launches and the ability to > handle multiple payloads, I think the big capacity is a plus, I think the point is that the demand is not currently there. Multiple payloads increase integration difficulties, reduces flexibility, and raises problems with insurance. I do like the Mini P2, though. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Fri, 3 Apr 92 14:37:40 -0500 From: dietz@cs.rochester.edu To: space-tech@cs.cmu.edu Subject: P2 Launcher One other potential problem with the SRM nozzle... The P2's exhaust gas will be considerably enriched in water relative to SRM (or LOX/RP-1) exhaust. Moderately hot steam (> 1000 K) gasifies carbon by the water gas reaction: H2O + C --> H2 + CO The high water content of the gases will tend to push the equilibrium further to the left. High temperatures will also tend to push the reaction to the left; high pressure, to the right. In the F-1 engine carbon deposition helped insulate the nozzle, but I could imagine the reaction going the other way here. It would be undesirable for the graphite nozzle components (especially the nozzle insert) to be chemically eroded. Is this a showstopper for use of the SRM nozzle? Someone should sit down with the JANAF tables and figure the thermochemical equilibrium to see if it could be a problem. Kinetics might still make it ok, granted. Paul F. Dietz dietz@cs.rochester.edu ------------------------------ Date: Mon, 6 Apr 92 23:20:40 CDT From: "lucius chiaraviglio" To: space-tech@cs.cmu.edu Subject: P2 questions: why use separate pressurization gas and/or H[2]O[2]? If you need to pressurize a cryogenic liquid tank, why not just boil a little of the cryogenic liquid stored within? That way you don't have to bother with separate storage (and possible reactivity problems) of a pressurization gas. Also, the idea of using hydrogen peroxide for an oxidizer seems kind of counterproductive -- it is slightly over twice as heavy as oxygen for a given amount of reducing equivalent. From what I have heard of it, it is not the nicest stuff to have to handle either. | Lucius Chiaraviglio | Internet: chi9@midway.uchicago.edu ------------------------------ Sender: reed@intelhf.hf.intel.com (Reed College UUCP) To: space-tech@cs.cmu.edu Subject: Delta Clipper Date: Mon, 06 Apr 92 01:33:40 -0700 From: Doug Reeder First, two more article listings: "Launching the Delta Clipper" _Space_ November-December 1991 pp 17-19 "Delta Clipper Partners Set Goal for Single-Stage-to-Orbit Vehicle" _AW&ST_ February 3 1992 pp 55-56 The following information is from one or more of Single Stage To Orbit Program Phase I Concept Definition Perferred Concept Selection Exeutive Overview 11 December 1990, McDonnel Douglas Space Systems Company Single Stage to Orbit Program Delta Clipper Tomorrow's Transportation Today's Technology January 1991, McDonnell Douglas Space Systems Company Single-Stage-Rocket: Status and Opportunities Lt Col Pat Ladner, Maj. Jess Sponable SDIO/TNE While the information in these papers is public, I do not believe these particular papers are available. Sorry. DC-X Structure -------------- material & structures aluminum & steel tanks aluminum intertank & thrust structure graphite epoxy aeroshell reaction control system GO2/HG2 thrusters propulsion system four RL-10A-5 engines, throttleable from 30% to 100% (4,384-14764 lb.) (in another reference 2,600-13,400 lb.) at the lower thrust rating, ath especific impulse would be 380.5 sec (in another reference 221 sec) and the chamber pressur 142 psia. the values for 100% thrust are 373 sec specific impulse and 475 psia. chamber pressure. gross liftoff weight: 35,980 lb. empty weight: 15,940 lb. payload: 500 lb. vehicle heigt: 39 ft DC-Y structure -------------- materials propellant tanks: aluminum/lithium alloys and/or graphite epoxy main tank insulation: "blanket insulation" return fuel tank insulation: "high efficiency mulitilayer insulation" aeroshell: graphite epoxy and aluminum-stiffend honeycomb landing gear: titanium metal matrix flaps: aluminum siffened graphite epoxy honeycomb section with carbon-carbon material overlay to protect the surfaces wheels (for ground towing): graphite epoxy propulsion system eight engines of about 200,000 lb. thrust four engines will have extendable nozzles, for efficient high altitude operation gross liftoff weight: 1,020,530 lb. empty weight: 79,810 lb. payload to polar orbit: 10,000 lb. payload to Space Station: 15,000 lb. return payload: 15,000 lb. cargo bay size: 15 ft dia x 20 ft long vehicle height: 127.14 ft. mission duration nominal capability: 2 crew: 4 days 0 crew: 7 days ground crew size estimate: 200-220 people Heat Protection --------------- "The reentry temperatures on the vehicle are well withing the capabilites of current carbon carbon technology. Most of the outer surface is also within the capability of advanced flexible reusable surface insulation (AFRSI). AFRSI is lighter, but the robustness of carbon carbon is more proven." Abort Safety ------------ single engine failure: return to launch site from up to 200 sec into flight (M=8.5) abort to orbit from 100 sec onward (M=1.5) (main engine cutoff 385 sec) Rentry ------ retro fire (main engines) with base pointed forward reorient to nose forward fly constant angle of attack (~15 degrees) 20,000 ft altitude initiate main engine idle 10,000 ft altitude rotate to base down 5000 ft inctrease power decelerate to landing "During most of the return flight, the Delta Clipper remains in a nose-up attitude at a contant angle of attack. It attains crossrange maneuver by rolling around its flight vector. The reentry capabilites, particulary the crossrange, will greatly increase the space ship's operation utility for takeoff and landing from the same base." DC-X Flight Test Series ----------------------- series 1: launch ascend hover at 500 ft descend deploy landing gear land Turnaround Time Reduction ------------------------- from a chart in the Ladner&Spoonable paper: STS/ELV Experience: 680 hours (85 days) vehicle accessibility: -14 days aircraft-lik certification process: -13 days standardized payload/launch vehicle interfaces: -12 days autonomous vehicle: -8 days sealing technology: -8 days subsystem operational rqmts/design margins: -7 days automated paperless documentation: -7 days engine opeation rqmts: -6 electromechanical actuators: -2 fiber optics: -1 speciialized handling rqmt: -1 reduced tem/hemid sensitivities: -1 net turnaround time: 6 days (numbers may not sum due to roundoff errors) Doug Reeder USENET: ...!tektronix!reed!reeder Internet: reeder@reed.EDU BITNET: reeder@reed.BITNET I'm looking for a grad school or a job as a research assistant where I can work on tethers for space propulsion or robotics, in particular, walking machines. ------------------------------ End of Space-tech Digest #112 *******************