Sideways Engineering the SpaceX Dragon

[gunsandrockets]

"...Given the very volume-restricted crew quarters, and a desire to minimize ECLSS requirements, the transit time from launch to docking will be minimized. Soyuz takes ~2-3 days from launch to docking, and ~4 hours from undocking to touchdown. While the undocking->recovery timeframe is acceptable, the launch timeframe is impractical without significant ECLSS upgrades. Significant reductions should be possible via improved positioning and orientation capabilities on the craft. The actual Hohmann transfer orbit from orbital insertion to space station rendezvous takes just under an hour. Most of the 2-3 days on the Soyuz flight is spent refining position and trajectories to maximize the burn accuracy (and thereby minimize propellant usage)."<br /><br /><br />I believe these assumptions about ISS rendezvous time are in error.<br /><br />Most launch windows which match the orbital plane of the ISS can require a spacecraft to loiter in a lower orbit for up to two days in order to catch up to the ISS position. It's only when a spacecraft has caught up that it boosts it's orbit up to the same altitude as the ISS.

 

[gunsandrockets]

"I just found this link on another blog. An interesting concept from LM for a capsule to launch tourists from an Atlas V. In particular, their LES is shade of what has been discussed for G-X3 and Dragon -- namely beneath the capsule and doubles as their OMS... I have to wonder exactly how that OMS/LES generates enough thrust for an abort. From the artists' concept, I have to call it unlikely."<br /><br />Supposedly a base mounted pusher-type launch escape system enclosed by the insterstage does not need the same power as a tractor-type launch escape tower. Perhaps the enclosed space multiplies the force by providing an effect like a gun-tube.

 

[mrmorris]

<font color="yellow">"Supposedly a base mounted pusher-type launch escape system enclosed by the insterstage does not need the same power as a tractor-type launch escape tower"</font><br /><br />In later posts, discussions about the LES brought this up. Because a tower LES must thrust off-axis, it requies more thrust than one below that provides thrust through the c.g. Also the mass of the tower and blast shield are eliminated.

 

[mrmorris]

<font color="yellow">"Most launch windows which match the orbital plane of the ISS can require a spacecraft to loiter in a lower orbit for up to two days in order to catch up to the ISS position. "</font><br /><br />Let's look at the orbits. From SpaceX, the Falcon 9 is supposed to place the Dragon into a 185km elliptical orbit. A *circular* 185km orbit has a period of 88.19 minutes, without more info, we'll use this for the period. The altitude of the ISS generally ranges from 350km to 400km. I'll use 370km, with a period of 91.95 minutes for this calculation. <br /><br />There's 3.76 minutes difference between the two orbits, so on a per-orbit basis, the Dragon will 'catch up' by that much. Worst-case, this means it would take about 24 orbits (~36 hours) to make a complete circuit. The ISS travels at a constant velocity, so simple math says that a quarter of the time, the time required to catch up would be nine hours or less, and an eighth of the time it would be 4.5 hours or less. <br /><br />Launch windows to the ISS occur twice daily. The shuttle can only make use of only one a day, however. The ascent trajectory must be headed from south to north for the orbiter to be able to discard its external tank into the ocean, and to reach its abort landing sites if needed. Neither restriction should be in effect for Dragon. This immediately doubles the number of potential launch windows for SpaceX. The current daylight restrictions on the Shuttle launch and rendezvous times cut <b>way</b> back on the remaining set of windows. The launch lighting restriction shouldn't affect Dragon. The rendezvous one may or may not, depending on NASA rules.<br /><br />Given 365 days in a year, with two launch windows daily, this means that we have 730 annual windows. 91 of these should require a 4.5 hours or less of orbital 'hang time' before the Hohmann transfer is initiated. If we assume that NASA requires the rendezvous to occur in daylight, then this drops to ab

 

[mrmorris]

<font color="yellow">"I wonder if some on-orbit testing will be specified for the Orion heat shield? "</font><br /><br />I've read that Lockheed plans to use the SLA (Super Lightweight Ablator) for Orion's heat shield. This ablator has been around for quite some time and has been used on numerous spacecraft (including the shuttle ET). It's well-tested and robust.<br /><br />Since SLA is manufactured *by* Lockheed, and they have no desire to see Dragon succeed, I looked for an alternative for Dragon's heat shield. Applied Research Associates makes an ablator extremely similar in mass and ablative characteristics to SLA. Theirs is called SRAM (dunno the meaning of the acronym). Like SLA, it is a cork/silicone-based ablator. They make at least three varieties (SRAM-14, SRAM-17, and SRAM-20), but I have not taken the time to determine the specific differences or try to pick a 'best' one for Dragon. Their masses are all reasonably similar, and I've simply assumed a thickness comparable to the ablator for the Apollo CM in making a total heat-shield mass estimate.

 

[mrmorris]

OK -- durned if I know what I did stupid... but I certainly did something. My calculation on the volume available in the forward equipment section was way off. Unfortunately, I didn't post the calculation, so I can only speculate. Doesn't matter anyway.<br /><br />I took the time to add the new equipment section ito the database today and reshuffled equipment and calculations around. In doing this, I found my numbers were off. On the *plus* side, while I was troubleshooting, I found a better formula for calculating the volume of a truncated cone (previously I was using an approximation I devised). Anyway -- there still appears to be room for the parachutes, but it's certainly a lot cozier than I thought initially.

 

[mrmorris]

I finally remembered the proper formula to calculate the propellant mass. It'd been posted at one point in my G-X3 thread. I dug it up and I've got a good bit of the work done to automate propellant mass/volume calculations in my application. Seeing as we're not sure what propellant mix Dragon will use, I decided to put the physical specs of several propellant/oxidizer combinations into a table. With that, I can calculate the mass and volumes for multiple combinations at once. Currently I have the following combos:<br /><br />N2O4/UDMH<br />LOX/Ethanol<br />90% H2O2<br />LOX/Kerosene<br /><br />Can anyone think of other combinations? They don't really have to be <b>likely</b> (although that would be preferable). I may add LOX/LH2 just for the heck of it -- even though it makes a ridiculous RCS option. I may also add 95% and 100% H2O2.

 

[bpfeifer]

Have you considered hypergolics, or has SpaceX made some announcement that makes them unlikely? I assume you're talking about just the Dragon, and not the Falcon boosters. <div class="Discussion_UserSignature"> Brian J. Pfeifer http://sabletower.wordpress.com<br /> The Dogsoldier Codex http://www.lulu.com/sabletower<br /> </div>

 

[mrmorris]

N2O4/UDMH is a pretty standard hypergolic combination.

 

[mrmorris]

OK -- the attached picture is the propellant calculation sections from my application. It calculates the mass fraction using the isp for the given propellant/oxidizer combination, a maximum delta-v parameter value (currently set to 300 m/s), and an original mass (Mo) parameter value that is set to the max mass-to-LEO for the Falcon 9 (currently set to 9300kg). The beauty of the app is that if I change any of those three parameters, all the calcs get redone automagically. <img src="/images/icons/smile.gif" /> Right now I'm largely using iSP values from Astronautix, so these are likely to be 'best case' rather than being based on actual hardware results. The equations I'm using to get mass fraction and propellant mass are:<br /><br />mass fraction = e ^ (dV / Ve)<br />propellant mass = Mo - (Mo/ mass fraction)<br /><br />I then use the propellant-oxidizer ratio to break out the mass of each, then calculate volumes based on their densities. <br /><br />Based on the volumes, the app then calculates the number of tanks that would be required to hold the propellant & oxidizer. Right now I'm just assuming a single tank style -- spherical with an interior radius of .5m, and assuming 20% ullage for pressurization. At some point I may try to locate two actual COTS tanks (one cryogenic and the other non-cryogenic) to refine the tank calculations further.<br /><br />Finally, I calculate the 'gross' volume, since spherical tanks won't pack neatly. Basically I simply assume each tank uses up a .5m cube of volume, or .125 m3.

 

[mrmorris]

<font color="yellow">"N2O4/UDMH is a pretty standard hypergolic combination. "</font><br /><br />Looking for likely on-orbit thrusters, it appears that N2O4/MMMH (Monomethylhydrazine) is more commonly used than N2O4/UDMH. I've added the specs for that to my application. I also have located another power user -- the thruster valves. For Aerojet RCS-sized thrusters, they use about 35-45 watts. That's presumably only drawing power when the thurster is being opened/closed, but I'll have to factor that into the power usage and peak wattage figures.

 

[mrmorris]

Periodically I browse around through the known locations of Dragon data to see if something catches my eye to produce an epiphany. I was on SpaceX's site reading their September Update when the pictures of the SpaceX Launch Escape Tower caught my eye. The first thing I thought was how short it appeared. Given all of the discussion about the Apollo LET with a height of 33 feet, it was immediately obvious that Dragons was not anywhere close to that. A quickie measurement using the dimensions I've engineered for Dragon puts the height at ~5.7 m tall vs ~10 meters for the Apollo LET. While I'll wholeheartedly admit that the graphic I have to work with is tiny, there's several things that jump out at me.<br /><br />-- It looks like this design does away with both the pitch control motor (50 lbs) and the tower jettison motor (525 lbs). Given the 'thrusters at top' design, both functions can be incorporated into them rather than requiring dedicated SRMs. One of the four can be given a slightly larger nozzle/throat to provide the pitch control vector. Tower jettison is probably accomplished by a similar 'uneven' thrust -- possibly only a single thruster firing (although I can't say how they'll accomplish that via a solid unless there's a separate propellant chamber for each thruster).<br /><br />-- With the thrusters at the top, the 500-pound 'tower' required for standoff is eliminated entirely.<br /><br />-- Likewise, even without the tower, the thrusters are farther away than the Apollo LES (~13ft vs ~10ft), meaning a lighter boost protective cover is feasible. If the protective needs diminish in a linear fashion with distance, the cover with the same engineering could be reduced by 30%. However, it's almost certain to atenuate by the <b>square</b> of the distance. This would mean a ~40% reduction. I believe the boost cover could simply be a second layer of AFRSI. The amount of AFRSI involved would only be about 20 pounds. Even with framework, etc, the Dragon b

 

[mrmorris]

OK -- with the propellant information available, I've been able to add a lot more calculations to my Dragon Designer application. The attached photo shows where the main calculation section stands at the moment. Almost every number is a calculated value (the ones with <b>bold</b> labels are parameters), so if I change my underlying parameters, the whole thing recalculates automagically. Right now it's got the calcs separated into four areas: <br /><br /><b>Surface Area/Volume</b>: Reasonably obvious<br /><b>Mass Limits</b>: An attempt to locate the upper mass limits based on known/published figures.<br /><b>Mass Calcs</b>: Sums up all of the mass elements I've stored in the database. This number will be low -- hopefully getting closer to the upper limit as I locate 'the missing mass'.<br /><b>Propellant Calcs</b>: Determines the GLOW mass for Dargon Cargo and Dragon Crew using the 'Mass Calcs' with the data from the selected Oxidizer/Propellant and the LES figures added in. Again -- this number will be low until I can locate all of the mass elements that I haven't accounted for (or have, but set too low).<br /><br />As always -- it's a work in progress, but the results are becoming more interesting.

 

[mrmorris]

I thought I'd supply an image of what my parameter page looks like. This is where I store many of the numbers used in the calculations. If I'm able to refine a number, I'll change it in here to adjust all of the results in the section above.

 

[mdodson]

"Right now I'm just assuming a single tank style -- spherical with an interior radius of .5m, and assuming 20% ullage for pressurization."<br /><br />A hazard of doing pressurization that way would be that the pressure supplied to the thrusters would taper off, meaning that the thrust now becomes variable over the mission, and ends up being small. That complicates your maneuvering calculations. A tank with a smaller percentage of propellant aggravates the problem of getting it to feed after being in free fall. There's no way to predict WHERE the blobs of propellant might be.

 

[mrmorris]

<font color="yellow">"A hazard of doing pressurization that way..."</font><br /><br />I calculated it that way because when I was reseaching propellant tanks for G-X3, I found a mission planning document (can't recall the mission) which calculated tank available volume using that method. If you have another method and can provide information/links -- I'm always interested in alternative options. <br /><br />Right now, my tank volume calculations are at the level of a first draft. Using the .5m spherical volume is a rough approximation even before we get into pressurization questions. At some point, as I indicated, I'll get data on actual COTS tanks and use that information to refine the tank data.

 

[mdodson]

"If you have another method..."<br /><br />Shuttle RCS and OMS have separate systems arranged as follows: A high pressure (4000 psi?) filament wound helium tank, at least one shutoff valve, a 250 psi regulator, and a 280 psi burst disk that pressurized one propellant tank.<br /><br />Inside each propellant tank was a series of screens to help exclude gas and keep a steady flow of liquid going out. Testing the integrity of each screen after firings was a laborious process. I don't want to guess what the screen system added to the tank weight.

 

[mdodson]

"If you have another method..."<br /><br />There was a post Nov. 17th on sci.space.news about a report stating that the Russians, on a vehicle called R36, would pressurize each tank of hypergolics by injecting a small amount of the opposite propellant. Simple, but sporty! (laughing)

 

[gunsandrockets]

"Can anyone think of other [propellant] combinations?"<br /><br />Hydrazine monopropellant<br /><br />Nitrous oxide monopropellant<br /><br />http://adsabs.harvard.edu/abs/2004gpsp.confE..20W<br /><br />N20 + hydrocarbon bipropellant<br /><br />GOX + propane bipropellant

 

[gunsandrockets]

"Orbital mechanics is not my specialty (although for that matter, neither is designing spacecraft). Do you see something wrong with the assumptions/math that I've used for this?"<br /><br />The math seems fine.<br /><br />Just for the heck of it, I roughly calculated the orbital position of the ISS for northeast launch windows for twenty days based on an orbital period of 92 minutes and the ISS starting above Florida. I came up with some non-intuitive results. During the first ten days there was only one launch window with the ISS less than 1/4 of it's orbit ahead of Florida. During the next ten days there were four windows! So even though the average was about 1/4 of all launch windows the close orbital opportunities could be clustered or scattered.<br /><br />As for assumptions, I don't think the southward launch direction is viable. A quick eyeballing of the region using google maps shows overflight of Carribean islands and a good chunk of South America. If something went wrong I don't think Cuba would respond too favorably to a Falcon IX dropping on their heads. So only one launch window per day is open to the ISS.<br /><br />One other obstacle to available launch windows is the Russian launch schedule of Soyuz and Progress spacecraft to the ISS. To that must be added the future ESA launch of the ATV spacecraft and JAXA launch of the HTV spacecraft to the ISS. I know NASA likes to keep the traffic clear of simultaneous launches to the ISS.<br />

 

[mrmorris]

OK -- additional combinations added and the new propellant list is attached.<br /><br />Some caveats: <br /><br />1. There's no such thing as a density for GOX, per se. Since it's a gas, the pressure can be anything from near-vacuum to something under 1.140 g/cm3 (i.e. LOX). I put it at 0.8 g/cm3.<br /><br />2. I couldn't find anything on NO2 and a hydrocarbon. Instead I added N2O4 and Kerosene. Possibly you meant a hybrid (i.e. N2O and rubber/etc.). Dunno.<br />

 

[mrmorris]

<font color="yellow">" If something went wrong I don't think Cuba would respond too favorably..."</font><br /><br />If my assumptions about the southward flightpath from KSC -- it looks like it would be just east of Cuba. It'd overfly the Dominican republic and a good bit of Brazil though. I don't know if this would rule out the southward flightpath or not (mind you Brazil might be perfectly happy to have some booster hardware land in their laps. <img src="/images/icons/smile.gif" /> <br /><br />However... what about Kwajalein? It's pretty well clear in whatever direction they feel like it. Even if the KSC-South path is ruled out, it's conceivable that Kwaj provides two more daily for a total of three potential windows per day.

 

[gunsandrockets]

"However... what about Kwajalein? It's pretty well clear in whatever direction they feel like it. Even if the KSC-South path is ruled out, it's conceivable that Kwaj provides two more daily for a total of three potential windows per day."<br /><br />Sure. Though I suspect NASA would insist on using their own launch facilities for ISS missions, and SpaceX would be happy to avoid the cost and headaches of transporting their hardware so far away from the continental United States.

 

[gunsandrockets]

"2. I couldn't find anything on NO2 and a hydrocarbon. Instead I added N2O4 and Kerosene. Possibly you meant a hybrid (i.e. N2O and rubber/etc.). Dunno."<br /><br />I meant a bipropellant rocket using nitrous oxide liquid oxidizer (N20, not NO2) plus any common liquid hydrocarbon fuel such as RP-1.<br /><br />N20 as a monopropellant is interesting stuff. Even the Chinese are looking into it...<br /><br />http://www.tsinghua.edu.cn/docsn/lxx/mainpage/a/Web/index_files/page0001.htm<br /><br />Lockheed has proposed it as a multiple use system. The N20 storage not only provides a non-toxic monopropellant for an RCS, it also provides a compact non-cryogenic storage source for Oxygen and Nitrogen.<br /><br />http://66.102.7.104/search?q=cache:BSnI_9mj4a0J:www.mesasphere.com/Mark.htm+lockheed+nitrous+oxide+monopropellant&hl=en&gl=us&ct=clnk&cd=2&ie=UTF-8<br /><br />Robert Zubrin has another pseudo-monopropellant idea using N20. He proposed a monopropellant mixture of N2O and hydrocarbon fuel to boost the ISP over that of a pure N2O monopropellant RCS. [update] Whoops! Found a new link which suggests Zubrin couldn't get it to work!<br /><br />http://sbir.gsfc.nasa.gov/SBIR/abstracts/00/sbir/phase1/SBIR-00-1-20.02-8728.html<br />

 

[mrmorris]

<font color="yellow">"(N20, not NO2)"</font><br /><br />Call it a persistant typo. My fingers keep wanting to type it the wrong way for some reason. <br /><br />*edit* Saw the Zubrin link *after* I replied. It is showing an isp of 300s, but obviously given the context, that's theoretical. Call it a practical iSP of 290s at best. I'll assume that the addition of the hydrocarbon doesn't change the density from straight N20 (not a *great* assumption, but all I've got to work with).