# Gemini: We can rebuild it, we have the technology

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#### mrmorris

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<font color="yellow">"While it is a plus to keep cable runs to a minimum, putting the avionics in the front defeats the dynamics of a capsule design. The major mass should be as close to the back as possible to increase the inherent stability. "</font><br /><br />The calculations for the ACRV patent (5,064,151 ) I mentioned earlier actually indicate the opposite as I read it:<br /><br /><i>"The other aerodynamic consideration was the moment slope (Cm-alpha) which determines the aerodynamic stability of the vehicle. A negative Cm-alpha defines a statically stable vehicle in which restoring forces are generated if the angle-of-attack is disturbed away from zero (note that dynamic stability is not guaranteed). The more negative the Cm-alpha, the more statically stable the vehicle behaves. FIG. 5 shows the Cm-alphas of the three original candidates and the final return vehicle design. They are plotted as functions of the c.g. position from the nose of the vehicles (Xcg). In all cases, the closer the c.g. is to the nose, the more negative Cm-alpha becomes and the more stable the vehicle is. "</i><br /><br />The seems to indicate we want more weight forward. I haven't run the calculations myself (nor do I have the means to), nor do I know if they are correct. I only know that I don't know enough to make a statement either way.<br /><br />

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#### najab

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That actually makes sense (after giving it some thought). The motion you are trying to fight is a tumbling of the capsule. The center of lift of the vehicle is the heat shield at the bottom. If the centre of mass was at the bottom as well then it would be easy for the capsule to rotate around the centre of mass and start tumbling. With the centre of mass way up near the nose it is much harder for the vehicle to rotate around it.<p>It's not the easiest thing to put into words - let me try another analogy: you have a stick fixed normal to the shaft of a motor at one end and free to move at the other, like the hand of a clock. Mounted on the stick is a mass. It takes more effort for the motor to rotate the stick when the mass is near the free end, than when it is near the shaft.<p>In other words, as the centre of mass of our capsule gets closer and closer to the nose, the bigger the disturbance required to upset the attitude.</p></p>

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#### nacnud

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I didn’t quite follow that but is it similar to why it is harder to balance a pencil vertically on the end of your finger as compared to a broom stick?

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#### najab

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Okay, let me try again...I'll use the clock analogy again since it's the one that's clearest in my mind.<p>Take a clock off the wall and put flat it on the table. Take a mass and attach it to the second hand so that it can slide from end to end of the hand. The motor will have to work much harder to move the hand when the mass is near the outer end than when it is near the centre, correct?<p>Now to relate this to the capsule. The capsule is the second hand, and the centre of lift (the heatshield) is the pivot. If the mass is concentrated near the heat shield then it doesn't take much to rotate the capsule around the heat shield (bad!). If the mass is near the nose (the outer end of the second hand), then it takes a lot more to rotate the capsule around the heatshield.<p>Any clearer?</p></p></p>

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#### nacnud

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Yeah that’s better, <br /><br />You just mean that the capsule will be more stable if it has a larger angular moment. Same as the reasoning behind my pencil and broom stick example.<br /><br />Does this mean thought that the lifting capsule designs (Apollo, Soyuz etc) all have to be actively controlled during re-entry rather than the designs that have passive stability like vostok and SSO (sorta)<br />

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#### najab

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><i>You just mean that the capsule will be more stable if it has a larger angular moment.</i><p>(Thinks about it for a moment...) Yeah, actually, that <i>is</i> what I meant! <img src="/images/icons/laugh.gif" /><p>><i>Does this mean thought that the lifting capsule designs (Apollo, Soyuz etc) all have to be actively controlled during re-entry...</i><p>I don't think so. As long as the lift/drag ratio is low enough, then they should be stable - even Apollo only had a l/d ratio something like 0.3 so drag was by far the dominant force. And we know that Soyuz is still stable enough to make a purely ballistic entry, even if it is a bit rougher on the crew.</p></p></p>

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#### mrmorris

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I'm looking for some Solid Rocket Motors for my DO boosters. As always -- I'm interested in pre-existing ones if possible. Does anyone know of current <b>small</b> SRMs on the market -- and where I could find specs on them?<br /><br />So far -- the only one I've found that looks at all possible is the Orion 38 -- made by Thiokol. It is the motor for the Pegasus stage 3. However -- I haven't been able to locate specs on it, so I dunno if it truly makes sense.

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#### mrmorris

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The only SRMs I can find are too big. Pratt & Whitney's Orbus 1 looked interesting... until I scrolled down and saw how big it is.<br /><br />Of course in searching -- I realized that there's no reason the DO stage couldn't use hybrid instead of straight solids. The ability to start and stop the DO engines would make this very attractive. At that point -- Spacedev and eAc become viable sources for the engines.

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#### mrmorris

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<font color="yellow">"How about a ATK Star kick stage: "</font><br /><br />That one is still much too large. However -- it gave me a search start point that led me to Astronautix section on engines (never found that bit before). From there -- I was able to get data on several Thiorkol SRMs. I didn't find the original Gemini one they made, but the Star 12A is reasonably close. Looking to Gemini-X3 -- the Star 17 looks promising.<br /><br />Star 12A <br /><br />------------------------------------------------------------------------<br />Manufacturer Name: TE-M-236-3. <br />Gross Mass: 34 kg. <br />Empty Mass: 11 kg. <br />Thrust(vac): 739 kgf. <br />Isp: 270 sec. <br />Burn time: 8 sec. <br />Diameter: 0.31 m. <br />Status: In Production. <br />Super SARV Retro is a longer, higher impulse version of the STAR 12, used as a retrograde rocket for an unmanned satellite vehicle. For all future deliveries, changes will be made to this design. Due to the unavailability of outdated materials, the main grain propellant will be changed from TP-G-3085 to TP-H-3340. Because of the higher flame temperature of the TP-H-3340 propellant, the molybdenum throat insert will be changed to G-90 graphite material. Total impulse 6,232 kgf-s. Propellant mass fraction 0.67. <br /> <br /> <br />Star 17 <br />-----------------------------------------------------------------------<br />Manufacturer Name: TE-M-479. <br />Gross Mass: 79 kg. <br />Empty Mass: 9 kg. <br />Thrust(vac): 1,116 kgf. <br />Isp: 286 sec. <br />Burn time: 18 sec. <br />Diameter: 0.44 m. <br />Length: 0.98 m. <br />Status: In Production. <br />First Flight: 1963. <br />Last Flight: 1965. <br />Flown: 15. <br />Comments: Orbit insertion motor has been used as the apogee kick motor for the Radio Astronomy Explorer satellite, the SOLRAD satellite, and an S-3 satellite. Total impulse 20,177 kgf-sec. Motor propellant mass fraction 0.881. <br /> <br /><br />I want to work out how much SRM propellant mass the Gemini-X3 requires. Using the origina

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#### mrmorris

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<font color="yellow">"...there it would be simpler to use cryogenic tanks, like those used on Shuttle for the fuel cells. "</font><br /><br />At which point you run into the <b>other</b> problem of using LH2 as a propellant. While LOX/LH2 are very *mass* efficient -- they are very *volume* inefficient. This is why the Shuttle ET is so flipping huge.<br /><br />Working the same equations to get the volume of LH2 required to get 300 m/s of delta-v:<br /><br />Dv = Isp * g * ln (W0/Wf) <br /><br />300 m/s = 453s * 9.8 m/s2 * ln (Wo/Wf) <br />300 m/s = 4439 m/s * ln (Wo/Wf) <br />0.06757669955 = ln (Wo/Wf) <br />1.06991231825 = Wo/Wf <br />Wf/Wo = 93.4% (i.e. 6.6% propellant mass required) <br /><br />Figure 4200 kg dry weight for Gemini-X3 <br />6.6% propellant would be 277.2 kg <br /><br />From Astronautix, LH2 is 0.071 g/cm3<br /><br />Volume Needed for 300 m/s: (mass ratio of Oxidizer to propellant 4:1)<br />V = 69.25 kg / 0.071 g/cm3 <br />V = 69,250 g / 0.071 g/cm3V = cm3 <br />V = 975,352 cm3 LH2<br /><br />5120 g * 1.14 g/cc = 5836.8 cm3 <br /><br />V = 207.74 kg / 1.14 g/cm3 <br />V = 207,740 g / 1.14 g/cm3<br />V = 236,835 cm3 <br /><br />The volume needed for an equivalent delta-v from LOX/ethanol was 474,018 cm3 of ethanol and 5,837 cm3 of LOX. So where the ethanol will fit into four 22.5" spherical tanks, for LH propellant we would need about eight, and the LOX would take another two -- whereas before it fit with plenty of room to spare in two 18" tanks. Gemini-X3 simply doesn't have that kind of volume to spare. The reduced mass is less important than the increased space requirement.

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#### mikejz

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Just making a suggestion: You might want to look into the cost of the fuel tanks vs. the cost of a Solid Retro Rocket. It might be more cost effective to use Lox/ethanol in larger disposable tanks.

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#### mrmorris

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<font color="yellow">"You might want to look into the cost of the fuel tanks vs. the cost of a Solid Retro Rocket. "</font><br /><br />I've already considered using more fuel tanks and dispensing with the adapter module. If I eliminated that -- then there would be plenty of space to add fuel to make up for the lost dv of the SRMs *and* add additional space to the pressure vessel. It also eliminates one more separation event.<br /><br />However, there's other things at stake here. The SRMs also act as part of the escape mechanism. In de-orbit mode -- the engines fire sequentially to slow the craft for re-entry. However -- at launch they can be fired in conjunction to act in lieu of an escape tower to push the capsule free of a failed booster. Also -- as NajaB indicates -- having SRMs for the DO module adds more of a failsafe to the system. In the event of problems with the OMS/RCS -- having the SRMs increases the chances that the capsule can still return from orbit. The simplicity of solids (especially ones this small) makes them less failure-prone than liquids. <br /><br />It very possible that Gemini-X3 Mark II might make such a change. If multiple flights of the FalconV booster prove it to be extremely robust, and the LOX/Ethanol engines of GX3 do the same -- a change of this nature might be considered worth the reduced margin of safety.

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#### mrmorris

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I've been re-thinking the re-entry of Gemini-X3 recently. Originally, of course, I was planning on a horizontal landing via parasail -- ala X-38. However -- it's always been the biggest technical question mark of the whole business. A horizontal landing greatly adds to the complexity of the engineering and the avionics.<br /><br />With a modern INS/GPS, a computer with several times the processing power of anything that's ever been used for a capsule re-entry, and an RCS/OMS with delta-v to spare -- I think the landing footprint of the Gemini-X3 could be reduced to a few square miles.<br /><br />I mentioned in the M&L forum on the Shuttle redesign thread the Genesis sample return. Ignoring the failure of the parachute -- NASA was anle to bring in that capsule to a <b>very</b> accurate re-entry. A similar desert re-entry for the G-X3 could have two helicopters waiting for recovery. One will pick up the personnel -- the second would be a heavy-lift copter to bring the X3 back to the hanger. With this -- the capsule could use the three-chute system of the Apollo return module. Since the RCS/OMS is present in the RM landing (unlike hydrazine-based ones that are jettisoned in orbit), this could be used to cushion the landing further -- much like Soyuz.<br /><br />While not quite as 'elegant' as a horizontal landing -- it has fewer failure points and greater safety redundancy.

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#### mikejz

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The only problems with that that I could think of is possible damage to the heatshield (if it lands on a rock....) and that the RCS in all odds does not have the kind of thrust needed.

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#### scottb50

##### Guest
<<Working the same equations to get the volume of LH2 required to get 300 m/s of delta-v>><br /><br />I wouldn't want to put all the propellant and oxidizer inside the X-3. First of all the only way to do that would be to have a compromised heat-shield. I was thinking more along the lines of a separate stage that mounts to the nose docking adapter of the X-3. For launch it acts as an escape tower, once in orbit it is used to modify the orbit. Sort of a small ET with a couple of engines. At a destination, in orbit, the X-3 undocks and docks with the station. The upper stage could be docked to a holding adapter for storage. For return the X-3 undocks from the station, docks to an upper stage and is de-orbited. If propellant is available at the station the upper stage could return to the station, automatically and be used for the next mission after being refueled. The more upper stages in orbit the more mobility we have and the more places we can go.<br /><br />The whole idea is to make a cheap way to get into orbit and back. I realize that LH2 consumes a lot of volume, but the physical size of the upper stage is basically irrelevant.<br /><br />As for the RCS system it would take up more room than hypergolics, but I would rather not have Hydrazine or Nitrogen Tetroxide in very close proximity. A rough landing could be a big problem. That also seems to be the main problem with keeping Soyus at the ISS, the storage system is compromised after a certain time, a definite problem if we want to refly the X-3 on a regular basis. <div class="Discussion_UserSignature"> </div>

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#### najab

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><i>For launch it acts as an escape tower, once in orbit it is used to modify the orbit.</i><p>I have no problem with a LOX/LH2 engine being used for on-orbit operations other than the fact that it's likely to be too powerful (I don't recall any low-thrust LOX/LH2 engines out there, but lets assume we can manage to find one) but I have a <b>BIG</b> problem with using one as an escape rocket.<p>First off, they can't be ignited instantly - the propellant lines have be chilled down first or the thermal shock causes them to break. Secondly, I don't know that I want to be hauling a tank of hydrogen around with me when things are obviously going horribly wrong. Thirdly LOX/LH2 engines are favoured for their high ISP, but don't give you the kick-in-the-pants that a solid would give you. For my launch escape system I want something that will get me out of &%\$#@!ville in a hurry!</p></p>

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#### mrmorris

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<font color="yellow">"possible damage to the heatshield..."</font><br /><br />My bad -- I keep posting the concepts whirling in my head in dribs and drabs. I envision the heat shield as being much more modular than that of the orbiters. The TPS material (whether ceramic or metal based) will be secured to a rigid framework that can be removed from the main body of the G-X3 (I envision it being bolted to the rear bulkhead -- but the attachment method isn't a critical factor). This concept has several advantages, and no disadvantages that I can determine.<br /><br />1. The heat shield is likely to be one of the most maintenance-intensive portions of the spacecraft. With this design -- there can be numerous heat shields manufactured for each craft -- and can be swapped in a matter of hours.<br />2. Repairing damaged tiles on the heat shield will be much easier if it is not attached to the orbiter at the time. <br />3. For metallic TPS -- tiles can be secured to the frame using bolts from the frame side out. This will make replacing damaged sections much simpler.<br />4. Gemini-X3 isn't locked into a single TPS system. They can have multiple subcontractors create a shield using their material -- then each can be tested on separate flights. Likewise if a new (better/cheaper) technology comes out -- it can be implemented immediately.<br /><br />Because of this -- I'm less worried about damage to the TPS than would otherwise be the case.<br /><br /><font color="yellow">"...the RCS in all odds does not have the kind of thrust needed. "</font><br /><br />I don't expect the RCS to slow Gemini-X3 to a soft landing by itself. The parachutes are intended to perform the vast majority of the deceleration. I mention the RCS because it's there -- and may well have propellant left at this time. Most spacecraft jetisson their RCS system before re-entry (again -- because hydrazine is generally considered too dangerous to land with). As for thrust -- it won't have

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#### mrmorris

##### Guest
<font color="yellow">"At a destination, in orbit, the X-3 undocks and docks with the station. The upper stage could be docked to a holding adapter for storage. For return the X-3 undocks from the station, docks to an upper stage and is de-orbited. "</font><br /><br />That's a very intelligent 'ET' to be able to perform that particular shell game. I can't view that as 'keeping things simple'. I believe I'll stick to SRMs and LOX/Ethanol in my personal imaginary spaceship.<br /><br /><font color="yellow">"...I would rather not have Hydrazine or Nitrogen Tetroxide in very close proximity..."</font><br /><br />As I've said -- Gemini-X3 (the one in *my* head anyway) contains no hypergolics. I want LOX/Ethanol as the RCS/OMS. Ergo -- I don't know where this paragraph comes from.

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#### najab

##### Guest
Did you ever get a chance to look into hydrogen peroxide as a monopropellant?

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#### mrmorris

##### Guest
<font color="yellow">"...to look into hydrogen peroxide ..."</font><br /><br />Nope. I had a daughter born 12 days ago and my wife is already calling me 'Spaceman Spiff' for the amount of time I spend on SDC posting the things I already *have* researched or postulated. Haven't taken the time to go off on a tangent into a new propulsion tech. Is anyone working wit this that you know of? I'm really trying to avoid technologies that require starting from ground zero. Ideally I want to have a plan that for the most part can be contracted out with most of the research completed and only the development to be done. Right now -- the Gemini-X3 consists of:<br /><br /><b>Structure</b> -- subcontracted to small aerospace firm (i.e. Cessna)<br /><b>Pressure vessel/seats, interior</b> -- same as above<br /><b>TPS</b> -- Select TPS from several competing technologies developed in the past few years (TUFI/AETB-8 Tiles, QUIC-TUFI, BF Goodrich Aerospace -- Metallic TPS)<br /><b>RCS/OMS</b> -- Ethanol/LOX (Aerojet is working on one now for NASA)<br /><b>SRMs</b> -- Thiorkol STAR-17 or equivalent (obviously Thiorkol is the prime candidate, but Aerojet & others might also wish to compete)<br /><b>Avionics</b> -- Honyewell's Primus Elite, Rockwell's Pro-Line 21, Garmin's G1000. Whichever can do the job the best for the least.<br /><b>Parachutes</b> -- (I forget who manufactured the Apollo parachutes, but they'd get first nod)<br /><b>Docking assembly</b> -- Energiya I believe makes the APAS-89 docking assembly to order. Failing that -- a US firm could probably reverse-engineer it -- much as the Chinese did. I doubt they sell it for less than the Russians, though, so why bother?<br /><b>Propellent Tanks</b> -- 22.1-in. Atlas-Centaur tank (or equivalent)<br /><b>Batteries</b> -- Li-Ion -- several manufacturerers have space-rated systems -- shop by price and performance.<br /><br />Obviously there's a great deal of specific design work to be done designing the structure, integrat

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#### najab

##### Guest
Congratulations on your new daughter! <img src="/images/icons/laugh.gif" />

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#### mikejz

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I would stick with stright lithium bats as they have more power-to-mass as opposed to Li-Ion batteries and a really long shelf life. The only thing you give up is being able to recharge them, which I would not recommend for G-X3 as Li-ion batteries are not known to last that long (My Li-Ion Laptop that is 2 1/2 years old only lasts half of what it did when I purchased it)

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#### mrmorris

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<font color="yellow">"Li-ion batteries are not known to last that long "</font><br /><br />What kills Li-Ion batteries isn't age, but repeated rechargings. Unlike a laptop which is constantly charging and discharging, the G-X3 at best would only discharge and recharge a few times a year.<br /><br />I like the idea of a battery system that can be trickle-charged at the space station while it is docked. However -- once again -- this is a system that needn't be with the G-X3 "til death do they part". I'm not dead-set on any particular technology. I want the basic concept to be batteries, as opposed to fuel cells or solar -- but beyond that, I'm simply looking to optimize cost, mass, and volume. As with the engineer's mantra -- it's unlikely to find a system that is the best at all three, so it will come down to which factor(s) are the most critical in order to pick one technology over another.

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