NASASpaceFlight.com Forum
Commercial and US Government Launch Vehicles => ULA - Delta, Atlas, Vulcan => Topic started by: sdsds on 05/02/2010 08:57 pm
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For a mission payload launched to LEO on Atlas V, could a separately launched Centaur (maybe dual-engine) rendezvous with it? Suppose the Centaur were launched as a passive payload on a Delta-IV Heavy. Could it reach rendezvous still almost fully loaded with propellant? After rendezvous, could the Centaur then act as an Earth departure stage for the separately launched mission payload?
Delta IV-Heavy payload to LEO is roughly 20 mT, and that's roughly the wet mass of Centaur.... Is this concept a "no brainer," or is it somehow "brain dead"?
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It's been proposed before, by serious people, and it is one of my favourite near term ways to go beyond earth orbit.
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IIRC, ULA have already done a study on replacing the Delta-IV upper stages with the Single-Engine Centaur. Their studies have shown a favorable increase in performance over the DIVUS family.
However, I believe that you are talking about launching the Centaur as a payload, specifically as an EDS for a BEO mission. For this to work, you would need to solve the following issues:
1) Power supply - How long can Centaur's batteries last? That is the amount of time you have to launch the payload;
a) Can you fit solar arrays to the Centaur?
2) Rendezvous - How does the Centaur rendezvous with the mission module? (Accepting that Orion would need one for all by the shortest and least-ambitious missions);
3) TOI payloads - Centaur is fairly small as EDS options go. It is fine for launching a robotic interplanetary probe, but a crewed space probe would be as much as an order of magnitude greater. Would Centaur be able to push even just an Orion through TOI to any useful destination?
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1) Power supply - How long can Centaur's batteries last? That is the amount of time you have to launch the payload;
a) Can you fit solar arrays to the Centaur?
2) Rendezvous - How does the Centaur rendezvous with the mission module? (Accepting that Orion would need one for all by the shortest and least-ambitious missions);
ULA (or LM before them) proposed an extended duration mission kit for that. You certainly couldn't use the Centaur as is.
3) TOI payloads - Centaur is fairly small as EDS options go. It is fine for launching a robotic interplanetary probe, but a crewed space probe would be as much as an order of magnitude greater. Would Centaur be able to push even just an Orion through TOI to any useful destination?
If it's unmanned, without return propellant, yes. Otherwise Orion is too heavy. If you want to do TLI with a Centaur you need a smaller capsule or a bigger Centaur. A Delta IV upper stage would be big enough, and the new Delta IV Heavy with RS-68A or else Atlas V Heavy might barely be able to lift that to orbit. Otherwise you'd need cryogenic propellant transfer.
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could you not design a framework which would enable 3 Centaur to hook onto a single unit, as the EDS? Launch each of them seperately to rendevous up in orbit.
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could you not design a framework which would enable 3 Centaur to hook onto a single unit, as the EDS? Launch each of them seperately to rendevous up in orbit.
I don't think that anyone has ever done a lateral rendezvous without the assistance of a crewed spacecraft equipped with an RMS. I'm sure it could be done, but whether the R&D would be cheaper than just building a larger (4- or 6-engine) Centaur-derived EDS is questionable.
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could you not design a framework which would enable 3 Centaur to hook onto a single unit, as the EDS? Launch each of them seperately to rendevous up in orbit.
I don't think that anyone has ever done a lateral rendezvous without the assistance of a crewed spacecraft equipped with an RMS. I'm sure it could be done, but whether the R&D would be cheaper than just building a larger (4- or 6-engine) Centaur-derived EDS is questionable.
The issue with larger units is what will carry it up?
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could you not design a framework which would enable 3 Centaur to hook onto a single unit, as the EDS? Launch each of them seperately to rendevous up in orbit.
ESA was considering something similar, which is not surprising since its cryogenic upper stage is so small. The Mars DRM did something similar, which is not surprising given the large mass it had to put through TMI. But apart from complexity the drawback would be that you would still have to mitigate substantial boil-off issues.
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could you not design a framework which would enable 3 Centaur to hook onto a single unit, as the EDS? Launch each of them seperately to rendevous up in orbit.
I don't think that anyone has ever done a lateral rendezvous without the assistance of a crewed spacecraft equipped with an RMS. I'm sure it could be done, but whether the R&D would be cheaper than just building a larger (4- or 6-engine) Centaur-derived EDS is questionable.
The issue with larger units is what will carry it up?
A 5.4m-diameter stage could fit on the current Atlas-V and Delta-IV cores without too much trouble.
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could you not design a framework which would enable 3 Centaur to hook onto a single unit, as the EDS? Launch each of them seperately to rendevous up in orbit.
ESA was considering something similar, which is not surprising since its cryogenic upper stage is so small. The Mars DRM did something similar, which is not surprising given the large mass it had to put through TMI. But apart from complexity the drawback would be that you would still have to mitigate substantial boil-off issues.
An option for boil-off is to house a shade around the mounting point of what you are mounting them to.
What if you did not do this in a single go, and you did staged EDS, you use one to push into a large eccentric orbit, then dump the first EDS. It then meets up with the second EDS, which has burned up more of its fuel to reach the item, and it pushes even more, to get things even more eccentric, where it then meets up with the last EDS, already en route. A way to get a *bit* more out of the system.
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I believe that you are talking about launching the Centaur as a payload, specifically as an EDS for a BEO mission.
Yes, and this specific configuration only because of a coincidental similarity: the payload to LEO of Delta IV-H is approximately the wet mass of a Centaur.
1) Power supply - How long can Centaur's batteries last?
There was a recent Centaur mission that did some R&D activity after the primary payload was safely separated. Wasn't that extended mission duration 72 hours or more?
That is the amount of time you have to launch the payload;
Unless the payload launches first and can meet the necessary loiter requirements. (Understood this is the reverse of the CxP architecture.) For a payload using hypergolic propellants with a design flight duration measured in weeks, the loiter requirement might be met within existing margins.
2) Rendezvous - How does the Centaur rendezvous with the mission module? (Accepting that Orion would need one for all by the shortest and least-ambitious missions);
Perhaps the mission module could be already at the destination, delivered by a prior launch campaign?
3) TOI payloads - Centaur is fairly small as EDS options go. It is fine for launching a robotic interplanetary probe, but a crewed space probe would be as much as an order of magnitude greater. Would Centaur be able to push even just an Orion through TOI to any useful destination?
If it's unmanned, without return propellant, yes. Otherwise Orion is too heavy
So the concept of operations would be the departure stage sending the payload (possibly even including crew) into a nominal "free return" lunar flyby orbit. The payload would have enough propellant to:
a) in the nominal case, adjust its trajectory and perform an "insertion" burn (or burns) to reach its pre-positioned return propellant. (Or a pre-positioned return propulsion module.)
b) in the case of a minor departure stage under-burn, supplement the burn to get back on the flyby return trajectory.
c) in the case of a major departure stage under-burn that only reaches a highly elliptical Earth orbit, perform appropriate re-entry burns.
I hope lunar orbit, EML1 and EML2 would be reasonable crewed mission destinations for this architecture. Is there a hope that dual-launch could be competitive with single-launch architectures for long duration robotic missions? Does it provide enough more mass to Earth-escape (compared to a single Atlas V 551 or Delta IV-H) to interest the science mission community? For example would it, in combination with Mars-orbit rendezvous, enable robotic Mars sample return?
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An option for boil-off is to house a shade around the mounting point of what you are mounting them to.
That is a good idea, and ULA has been doing some work on that.
What if you did not do this in a single go, and you did staged EDS, you use one to push into a large eccentric orbit, then dump the first EDS. It then meets up with the second EDS, which has burned up more of its fuel to reach the item, and it pushes even more, to get things even more eccentric, where it then meets up with the last EDS, already en route. A way to get a *bit* more out of the system.
That could work, but rendez-vous would be more complicated and time-critical. You would also have more van Allen transits, but presumably only the last one would be manned, so that wouldn't be a show stopper.
But rendez-vous in a nearly circular high energy orbit sounds easier. Most likely that orbit would be L1/L2, but MEO or GEO could be used too if the lifeboat scenario is a concern. A bit less efficient, but that might be worth it.
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So the concept of operations would be the departure stage sending the payload (possibly even including crew) into a nominal "free return" lunar flyby orbit. The payload would have enough propellant to:
a) in the nominal case, adjust its trajectory and perform an "insertion" burn (or burns) to reach its pre-positioned return propellant. (Or a pre-positioned return propulsion module.)
It depends a lot on the size of the stage, the size of the capsule, the choice of staging point and how long the trip is allowed to take. I did a lot of sums on DCSS (Delta IV 5m upper stage, which has a bewildering array of names) and Orion, back when I thought an SDLV was inevitable. With that you could get an Orion to LLO, but not with enough propellant to get back, which seemed too risky. With L1/L2 as a staging point you could get there with enough propellant to both the insertion burn and the TEI. Not with a whole lot of margin, but it looked doable. A slightly bigger stage would be useful.
A Centaur could only get you to L1/L2 on an almost ballistic trajectory, so you wouldn't need an insertion burn. This means the journey takes >=100 days, so this is not suitable for crew. Not being able to bring your return propellant with you also rules it out. But with a smaller capsule a Centaur might work.
I hope lunar orbit, EML1 and EML2 would be reasonable crewed mission destinations for this architecture.
A Lagrange point is all you need. From there everything but LEO can be reached at reasonable cost, even with storable propellant. And since storable propellant transfer is a mature technology, you could easily use refueling at L1/L2. In order to stimulate development of RLVs (and cryogenic depots, maybe even EELV Phase 1) you would want to buy that propellant in LEO, at L1/L2 and also buy transportation services between the two. All that would happen slowly, in the background, without it appearing on the critical path.
Is there a hope that dual-launch could be competitive with single-launch architectures for long duration robotic missions? Does it provide enough more mass to Earth-escape (compared to a single Atlas V 551 or Delta IV-H) to interest the science mission community? For example would it, in combination with Mars-orbit rendezvous, enable robotic Mars sample return?
So far no one has done it, which is not a good sign. In part that may be because it would be too expensive to develop on the budget of a single science mission. Another reason would be that science missions do not now max out the payload size of existing launchers, so there's little point in developing technology that won't be needed for a long time from a science budget.
Of course, once NASA developed the technology for it (or had it developed by industry), it could make it available to science missions as well. An alternative would be refueling at a Lagrange point. Again, NASA would have to pay for the initial development, but once it was available, it could be used by science missions too. In light of current missions not maxing out current launch vehicles, I suspect this would only make sense once things like RLVs had made launching propellant a lot cheaper than it is today.
MSL may be a perfect example. I think the proposal to split it up stems from a lack of funding for developing the spacecraft itself, not because it doesn't fit on a launch vehicle.
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This is an interesting idea. Sounds a bit like accelerated ACES or mini ACES. Perhaps, in order to make centaur capable of being the EDS all one needs to do is: Add solar wings, add in the ACES low boil off system [could this simply consist of uprated insulation and cold, boil off gas tubes? (only in this case so as to make centaur more capable for the least amount of money)], finally, add the ACES propellant depot sunshade. In short: mini ACES. This could also fly on jupiter130 as a smaller, cheaper eds
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It makes sense to put some numbers on the proposed Earth departure burn performed by the Centaur. For the CxP lunar sortie mission, the Orion TLI mass was 20.2 t. Spacelaunchreport shows the Centaur total mass as 22.83 t. That's approximately 43 t between the two for m0. With a Centaur dry mass of 2 t that's approximately 22 t for m1.
Using a Centaur Isp of 450 s:
450 * 9.8 * ln(43 / 22) = 2,955 m/s of delta-v.
That's nearly the CxP sortie TLI of 3,175 m/s. The Orion for that mission was to be loaded with propellant for a 1,560 m/s TEI burn from LLO. It seems like off-loading some of that propellant demonstrates the architecture could "close".
For the return I'm suggesting use of a full pre-positioned propulsion module (rather than propellant) because rendezvous and docking are fundamental capabilities for Orion. With Centaur providing the Earth departure burn, all Orion needs to provide is the destination "insertion" burn, and then use its rendezvous and docking capability to dock with a propulsion module that will provide the boost to get home.
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This is an interesting idea. Sounds a bit like accelerated ACES or mini ACES.
Yes, very much the same idea. The term "accelerated" here would mean, "using existing stages (or very minor modifications thereof) for the major propulsive maneuvers." Doing this would allow NASA to focus for the next few years on constructing payloads, principally a "deep space" Orion.
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This is an interesting idea. Sounds a bit like accelerated ACES or mini ACES.
Yes, very much the same idea. The term "accelerated" here would mean, "using existing stages (or very minor modifications thereof) for the major propulsive maneuvers." Doing this would allow NASA to focus for the next few years on constructing payloads, principally a "deep space" Orion.
I'm a fan of this idea. Existing stages means lower costs (unless you have a LOT of money to devote to exploration, in which case an exploration-focused stage can make sense, maybe) and lower time for development.
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On the performance calculations: I'm getting about 17.5mT through TLI. You also have to take the mass of a docking mechanism and an extended duration mission kit into account. That might be enough to do L1/L2 insertion as well, even without the capacity of a Delta upper stage. I'd be uneasy with not having return propellant. A smaller capsule sounds like a better idea to me.
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Ok question 2: How soon can the "mini ACES" be ready? The sooner the better. Also I would suggest using an avionics pacakge that is easily compatible with both EELV and SDHLV. That way this could fly on whatever rocket is availble.
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On the performance calculations: I'm getting about 17.5mT through TLI. You also have to take the mass of a docking mechanism and an extended duration mission kit into account. That might be enough to do L1/L2 insertion as well, even without the capacity of a Delta upper stage. I'd be uneasy with not having return propellant. A smaller capsule sounds like a better idea to me.
I do not like the idea of a smaller capsule............what about making centaur longer, in addition to adding solar wings, sunshades, low boiloff insulation/cold gas tubing ect?? I'm thinking mini ACES 71. Add second engine or switch to RL-10-B2 for more thrust.
Jim, would really like your input on the "mini ACES/ EDS centaur" idea.
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Doing this would allow NASA to focus for the next few years on constructing payloads, principally a "deep space" Orion.
Better yet, a deep-space Altair precursor which could also serve as a makeshift hypergolics depot. To go to the moon, we need a lander. We already have launch vehicles. An uprated commercial capsule could then serve as the return vehicle. The Altair work could even start now, instead of the CRV Orion, using the same workforce, but doing it a few years from now using CRV Orion as a basis for the lander would also be a good idea.
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On the performance calculations: I'm getting about 17.5mT through TLI. You also have to take the mass of a docking mechanism and an extended duration mission kit into account. That might be enough to do L1/L2 insertion as well, even without the capacity of a Delta upper stage. I'd be uneasy with not having return propellant. A smaller capsule sounds like a better idea to me.
It sounds like you will need the return propellant waiting at EML.
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Yeah, and I'm a bit uneasy with that. That's why I'd rather use an uprated smaller commercial capsule than Orion. Once at L1/L2 you could dock with a mission module, say an Altair precursor or a Bigelow style hab, or both. Orion could still serve as a CRV if necessary.
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On the performance calculations: I'm getting about 17.5mT through TLI. You also have to take the mass of a docking mechanism and an extended duration mission kit into account. That might be enough to do L1/L2 insertion as well, even without the capacity of a Delta upper stage. I'd be uneasy with not having return propellant. A smaller capsule sounds like a better idea to me.
It sounds like you will need the return propellant waiting at EML.
Centaur (modified i.e. mini aces based) Depot :D
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Yeah, and I'm a bit uneasy with that. That's why I'd rather use an uprated smaller commercial capsule than Orion. Once at L1/L2 you could dock with a mission module, say an Altair precursor or a Bigelow style hab, or both. Orion could still serve as a CRV if necessary.
I was under the impression that bigelow habs are better for stations or for surface bases, not deep space mission modules...........
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The Bigelow modules started their life as NASA's Mars TransHab.
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Suppose the Centaur were launched as a passive payload on a Delta-IV Heavy. [...] Is this concept a "no brainer," or is it somehow "brain dead"?
Oh. I don't suppose the propellant tanks at SLC-47 already have the capacity to fill a Centaur after having filled the three CBCs and the 5 m Delta upper stage?
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If not, then just offload the CBCs slightly.
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It looks like there is a really good idea here for a future upper stage/EDS. Great job SDSDS and all. :D
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You can launch A Delta IV Heavy without an upper stage too... :) Of course modified
http://gravityloss.wordpress.com/2009/08/13/delta-iv-manrating/
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An obvious aspect of this approach is that it requires some sort of interstage between the Delta vehicle and the Centaur. Estimating a mass for that should be easy, i.e. it should be about the mass of the interstage used on Atlas between the booster and the Centaur.
Looking at spacelaunchreport.com, though, makes this a bit confusing. The 400-series interstage is reported as 0.8 t, but the 500 series interstage is reported as 1.57 t. What do those interstages look like? Is one more appropriate than the other for supporting a Centaur?
On a related question, for an EELV Heavy vehicle there was at one point a bolt pattern specified for the "Standard Interface Plane" at the top of the launch vehicle that was a circle 173 inches in diameter. Did this match the interface to a pre-existing Titan payload adapter? Has this interface actually been implemented for Delta IV-Heavy launches?
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1. Looking at spacelaunchreport.com, though, makes this a bit confusing. The 400-series interstage is reported as 0.8 t, but the 500 series interstage is reported as 1.57 t. What do those interstages look like? Is one more appropriate than the other for supporting a Centaur?
2. On a related question, for an EELV Heavy vehicle there was at one point a bolt pattern specified for the "Standard Interface Plane" at the top of the launch vehicle that was a circle 173 inches in diameter. Did this match the interface to a pre-existing Titan payload adapter? Has this interface actually been implemented for Delta IV-Heavy launches?
1. The 400 series interstage is visible. It the truncated cone between the Centaur and Atlas. . The 500 series interstage is inside the fairing and it also supports the fairing. There are pics in the planner's guide.
2. The 173" bolt circle was from the outer Centaur G' forward adapter bolt circle on the Titan IV. Delta IV can provide a payload attach fitting that has the same dimensions.
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Why is it a Centaur in the first place? Why not a D-IV upperstage?
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So it will fit on a Delta Heavy. Would a Delta upper stage fit on a Delta Heavy with RS-68A? Or else on an Atlas Heavy? You could always offload some propellant of course.
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there are two D-IV upperstages and adapters can be made
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Doh, of course! It even has better Isp than Centaur. On the other hand, I thought Centaur was considered to be the best upper stage in the world. Is it?
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Why is it a Centaur in the first place? Why not a D-IV upperstage?
My thinking was motivated by two factors:
1) Duration: Centaur has a proven record of (I think) 9 hours between launch and final burn, and ULA has talked about mission kits that would extend that further. Has there been a D-IV upper stage conducting a mission of that duration? The goal is to provide a reasonably large window of time for the rendezvous and docking conducted by the departure stage's payload.
2) Workforce: The Delta upper stage workforce will be busy with activities related to the second stage. They'll be tired out once that stage has completed its mission. Some portion of the (still rested) Centaur workforce would sit in the Delta "payload" chairs; some portion would sit in their regular seats at the Atlas control center, ready to take over once Delta had done its job.
Also, from a political perspective this approach might help to "spread the wealth" and thus might garner wider support.
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1) Duration: Centaur has a proven record of (I think) 9 hours between launch and final burn, and ULA has talked about mission kits that would extend that further. Has there been a D-IV upper stage conducting a mission of that duration? The goal is to provide a reasonably large window of time for the rendezvous and docking conducted by the departure stage's payload.
2) Workforce: The Delta upper stage workforce will be busy with activities related to the second stage. They'll be tired out once that stage has completed its mission. Some portion of the (still rested) Centaur workforce would sit in the Delta "payload" chairs; some portion would sit in their regular seats at the Atlas control center, ready to take over once Delta had done its job.
1. The D-IV GSO mission is over 6 hours long. Also D-IV has kits.
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a. There is no savings there. For a mission to deliver anything to LEO, including a fully fueled upperstage, the duration is only a few minutes.
b. Anyways, the Centaur people will be on consoles during the countdown and launch, so there are no savings.
c. This is all mute. The Centaur is autonomous and there is no need for a "control" team. All ELV "flight" team only monitor telemetry and there is no uplink. The flight team is the same as the launch team.
ELV upperstages ha
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Doh, of course! It even has better Isp than Centaur. On the other hand, I thought Centaur was considered to be the best upper stage in the world. Is it?
Depends on your payload and your mission needs. DIVHUS has more fuel, more thrust (than DES), and higher ISP, but higher dry mass (presumably coming from both the big nozzle on RL-10-B-2 and the less efficient structural design). IIRC, running the numbers through the rocket equation for typical high-energy, lightweight space probe missions, it was about a wash. (And AV5x1 is much cheaper than DIVH). For heavy payloads, presumably DIVHUS makes a big difference.
But DIVH was never about price/performance, it was about substituting for Titan IV, some of whose payloads couldn't perform their own GSO burns. So, DIVHUS has mission packs (as Jim alludes to) to shepherd the national security payload (all?) the way.
I assume that all the same abilities were offered for Atlas V/Centaur, and indeed LockMart put up plenty of documents on upgrades they've studied (and don't seem hard to put into production), but like AVH I assume they're on hold pending a customer.
It's been said around here that Centaur's avionics are more advanced than Delta's. Is it fair to say that "ACES" as presently contemplated would look more like Centaur than DCSS?
-Alex
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The currently proposed budget seems to keep Orion alive as a CRV. As far as I know, the only vehicle that would be capable of lifting it would be a human-rated Delta IV Heavy. Since the existing Delta pad at LC-37 is not suitable for crewed launches, a second pad would have to be constructed for that purpose (which also means that NASA launches wouldn't interfere with DOD launches, which is a good thing).
If all this turns out to be the case, we'll have two Delta IV Heavy pads and an HR'd Delta IV Heavy capable of placing an Orion spacecraft into LEO.
Can we get from there to manned BEO missions without any new launch vehicle development? I believe the answer is yes.
One scenario would be to use the existing Delta IV Heavy to put its own mostly-fueled upper stage in orbit, launched from LC-37A. A block II Orion would be launched on a second Delta IV Heavy from LC-37B, less than 24 hours later.
The DIVUS would have a Soft Capture Module bolted to the front of it (similar to the one that was recently mounted on Hubble). The Orion would dock with the DIVUS, which would take it through TLI on a "free return" lunar trajectory or to a rendezvous at EML.
Time for some numbers...
According to http://www.astronautix.com/lvs/deltaiv.htm, the Delta IV Heavy can place 25,800 kg in LEO. With no payload, that means it can place its own upper stage in orbit with that amount of fuel remaining.
According to that same page, The upper stage has a dry mass of 3,490 kg.
According to http://en.wikipedia.org/wiki/Delta-v_budget, delta V from LEO to C3 is 3220 m/s. Let's assume the DIV upper stage has a vacuum Isp of 450 sec.
Solving the rocket equation for the dry mass we find that a DIV upper stage with 25,800 kg of fuel can put 23,990 kg through TLI. Subtracting the stage's dry weight of 3,490 leaves us with 20,500 kg of payload. That payload would be a block II Orion, placed in orbit by the second Delta IV Heavy.
According to http://astronautix.com/craft/orion.htm, Orion has a mass of 21,500 kg plus the LAS. However, a large part of that is fuel for the main propulsion system. Since our mission scenarios don't require that the Orion SM do TEI, we can reduce that mass by several thousand kg which would bring it within our budget of 20,500 kg with suitable margins.
So manned lunar flybys and round-trip missions to EML1 and EML2 are feasible with no new launch vehicle development. The only cost would be the development of the block II Orion, which would basically mean heavier heat shields and some radiation shielding.
Later on you could use the Delta IV Heavy to position a Sundancer-derived outpost at EML1 and/or EML2. Lunar landings would be harder, and would require some ACES-like technology (long-loiter cryogenics, possibly cryogenic fuel depots as well as a lander).
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One scenario would be to use the existing Delta IV Heavy to put its own mostly-fueled upper stage in orbit, launched from LC-37A. A block II Orion would be launched on a second Delta IV Heavy from LC-37B, less than 24 hours later.
The CONOPS for dual pads at SLC-37 does not include dual LCC's. There never has been a requirement for Delta IV's to be launched within days of each other. So this is not feasible. I believe Boeing/ULA have not even thought of this.
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The CONOPS for dual pads at SLC-37 does not include dual LCC's. There never has been a requirement for Delta IV's to be launched within days of each other. So this is not feasible. I believe Boeing/ULA have not even thought of this.
This would not be a problem if the ISS were used as a staging point and if the crew were launched first.
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Just fly the DIV Heavy with no payload and only burn the upper stage for a very short while or even not at all.
Requires significant analysis though and some mods. Low throttling and longer burn times? Higher q? I understand the computer on DIV is old fashioned and not very flexible so don't know how much it can be changed.
The DIV upper stage has two separate tanks and probably is very different from a thermal protection point of view than a Centaur. IIRC ULA still has some papers on it as a depot basis. Dallas Bienhoff?
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RE OP: Good idea. Great discussion.
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The CONOPS for dual pads at SLC-37 does not include dual LCC's. There never has been a requirement for Delta IV's to be launched within days of each other. So this is not feasible. I believe Boeing/ULA have not even thought of this.
This would not be a problem if the ISS were used as a staging point and if the crew were launched first.
Not only that, but then you also wouldn't even need another launch pad or a crew tower or a LAS for the Orion, which could be "just another" Delta IV Heavy payload. Crew could arrive at ISS whichever way they feel like: commercial crew, Shuttle (if that continues somehow), Soyuz, or Shenzou (if that ever happens).
Of course, it would change slightly the payload capabilities, but not by a heck of a lot (besides, we were going to offload Orion to begin with).
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The currently proposed budget seems to keep Orion alive as a CRV. As far as I know, the only vehicle that would be capable of lifting it would be a human-rated Delta IV Heavy.
If as a CRV it would be launched without a crew, why would the launcher need to be human-rated?
According to http://www.astronautix.com/lvs/deltaiv.htm, the Delta IV Heavy can place 25,800 kg in LEO. With no payload, that means it can place its own upper stage in orbit with that amount of fuel remaining.
Are you concerned about the ascent profile for this? With the first stage carrying roughly half the mass of a nominal mission, its acceleration at the end of its burn is roughly double. Can the upper stage tolerate that? Also, what about max-Q? And where does the expend core CBC fall?
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Not only that, but then you also wouldn't even need another launch pad or a crew tower or a LAS for the Orion, which could be "just another" Delta IV Heavy payload. Crew could arrive at ISS whichever way they feel like: commercial crew, Shuttle (if that continues somehow), Soyuz, or Shenzou (if that ever happens).
Yes, I believe that was Steidle's argument. It also conveniently mitigates boil off issues.
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Not only that, but then you also wouldn't even need another launch pad or a crew tower or a LAS for the Orion, which could be "just another" Delta IV Heavy payload. Crew could arrive at ISS whichever way they feel like: commercial crew, Shuttle (if that continues somehow), Soyuz, or Shenzou (if that ever happens).
Yes, I believe that was Steidle's argument. It also conveniently mitigates boil off issues.
Do you have a link to Steidle's argument? I'd be interested in hearing something along these lines, but by someone far, far (far) more experienced in the field than I. ;)
Also, then you would have two "Orions" docked to the ISS, right? If something bad happened to the "exploration" Orion while still in LEO (say, a botched docking with the Centaur that damaged the craft, causing the crew to don pressure suits before the cabin depressurized), couldn't the "lifeboat" Orion rescue the crew from the exploration Orion, thereby decreasing the LOC risk of another rendezvous and docking?
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the Delta IV Heavy can place 25,800 kg in LEO. With no payload, that means it can place its own upper stage in orbit with that amount of fuel remaining.
No it doesn't.
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Could this scenario (which was mentioned before) actually work, which I mentioned in the latest live hearing thread:
A trip to EML-1/2 (if not LLO) in a Block II Orion (lunar reentry, full prop load, weeks-long life support, but LAS and human-launchable not required):
*unmanned Orion launched on Delta IV Heavy
*docks at ISS
*a modified Delta IV upper stage is launched via Delta IV Heavy as a payload so it is co-orbiting with ISS
*crew enters Orion from ISS and the Orion disembarks to dock with the D4US
*D4US burns and places Orion on the way to EML1 or 2
*Orion handles the rest (EML1/2 insertion, departure, reentry)
Is this possible? How long to do this, assuming we magically had a block-II Orion?
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Whay not just carry the extra fuel as a payload in tanks and refuel the second stage from it and discard the tanks instead of carrying a whole extra RL-10 engine?
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Whay not just carry the extra fuel as a payload in tanks and refuel the second stage from it and discard the tanks instead of carrying a whole extra RL-10 engine?
In-space propellant transfer on that scale has never been demonstrated, so that would make for a great technology development mission! The idea here though is to use known, proven technologies like staging and payload deployment, and to use (as much as possible) known, proven systems to implement those technologies.
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The idea here though is to use known, proven technologies like staging and payload deployment, and to use (as much as possible) known, proven systems to implement those technologies.
In regards to which: what are the tested limits for "promptness" of rendezvous and docking? If a Delta Cryogenic Upper Stage has an on-orbit shelf-life of 6 hours, during which time it is a passive rendezvous target, does that impose unreasonable requirements on the active rendezvous vehicle or the tracking systems?
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The idea here though is to use known, proven technologies like staging and payload deployment, and to use (as much as possible) known, proven systems to implement those technologies.
In regards to which: what are the tested limits for "promptness" of rendezvous and docking? If a Delta Cryogenic Upper Stage has an on-orbit shelf-life of 6 hours, during which time it is a passive rendezvous target, does that impose unreasonable requirements on the active rendezvous vehicle or the tracking systems?
6 hours? That's clearly far too little for anything. Is it batteries? Boiloff? Both have some solutions, the former problem being much cheaper to tackle...
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The idea here though is to use known, proven technologies like staging and payload deployment, and to use (as much as possible) known, proven systems to implement those technologies.
In regards to which: what are the tested limits for "promptness" of rendezvous and docking? If a Delta Cryogenic Upper Stage has an on-orbit shelf-life of 6 hours, during which time it is a passive rendezvous target, does that impose unreasonable requirements on the active rendezvous vehicle or the tracking systems?
6 hours? That's clearly far too little for anything. Is it batteries? Boiloff? Both have some solutions, the former problem being much cheaper to tackle...
For boiloff, an external fairing could envelope the Delta IV upper-stage-as-payload during launch, permitting it to be covered in MLI to reduce boiloff, if need be. Longer lived batteries (and/or power rationing techniques) could be used. Could certainly be done if it had to be done, with a few years to do it.
Also, it took less than ten minutes for the Apollo 11 Command/Service Module to separate from the stack (by at least 50 feet?) then turn around and dock with the Lunar Module. I'm guessing that once you get close, an Orion docking to an upper stage would be a lot more like that than something docking to the ISS.
EDIT: And how long did Gemini 6A take to rendezvous (i.e. get within 120 feet)? I think it was just barely under 6 hours from launch. So, it should be possible, if you are very clever with timing phasing and launch windows. But this is the 21st century, so it shouldn't be an issue. ;)
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Is this possible? How long to do this, assuming we magically had a block-II Orion?
Well, I've been arguing for about a year that we should do it this way (or with small variations on this theme), so I'd say yes. :) This assumes we can launch a fully fueled 5m Delta upper stage on Delta or Atlas Heavy. It doesn't take gravity losses and boil-off into account and it doesn't have a whole lot of margin, unless you allow for refueling the SM at L1/L2. There are papers on the ULA website that describe the required modifications to the upper stage to turn it into an EDS capable of EOR. Many of these would have to be developed for an ACES EDS too, so it would even be an incremental step in that direction.
I'm guessing a lunar capable capsule would be the long pole. A couple of years if we had the magical Block II Orion? Note that it doesn't have to be a Block II Orion, it could also be a Block II commercial capsule. In that case we could make sure it would fit on an existing upper stage + mission kit.
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Is this possible? How long to do this, assuming we magically had a block-II Orion?
Well, I've been arguing for about a year that we should do it this way (or with small variations on this theme), so I'd say yes. :) This assumes we can launch a fully fueled 5m Delta upper stage on Delta or Atlas Heavy. It doesn't take gravity losses and boil-off into account and it doesn't have a whole lot of margin, unless you allow for refueling the SM at L1/L2. There are papers on the ULA website that describe the required modifications to the upper stage to turn it into an EDS capable of EOR. ...
Linky-link? ;)
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There are three releases of basically the same paper that describe using Centaur as an EDS:
Atlas Centaur Extensibility to Long-Duration In-Space Applications (http://www.ulalaunch.com/docs/publications/AtlasCentaurExtensibilitytoLongDurationInSpaceApplications20056738.pdf)
Centaur Extensibility For Long Duration (http://www.ulalaunch.com/docs/publications/CentaurExtensibilityForLongDuration20067270.pdf)
Centaur Upperstage Applicability for Several-Day Mission Durations with Minor Insulation Modifications (http://www.ulalaunch.com/docs/publications/CentaurUpperstageApplicabilityforSeveralDayMissionDurationswithMinorInsulationModificationsAIAA20075845.pdf)
And one that briefly describes using the Delta upper stage as an EDS:
DELTA IV LAUNCH VEHICLE GROWTH OPTIONS TO SUPPORT NASA’S SPACE EXPLORATION VISION (http://www.ulalaunch.com/docs/publications/DeltaIVLaunchVehicle%20GrowthOptionstoSupportNASA%27sSpaceExplorationVision.pdf)
The above discussion centers on boil-off, power and navigation. There is another one that details the changes to turn it into an EDS capable of EOR, which adds docking capability as well. It doesn't seem to be listed in the current list of publications anymore, but I've posted a link before so I'll see if I can find it.
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I think I've found it:
The Advanced Cryogenic Evolved Stage (ACES)-A Low-Cost, Low-Risk Approach to Space Exploration Launch (http://www.ulalaunch.com/docs/publications/TheAdvancedCryogenicEvolvedStageACES2006LeBar7454.pdf)
It's on the website after all, but I overlooked it because it describes ACES, not Centaur or the existing Delta upper stage. The same changes apply to those too.
The modifications are mentioned on page 5. On the preceding page there is a list of EDS options. The second one is the one we are talking about. The third one might be more useful if you were to develop a whole new stage, though it would be less efficient if it had to haul more metal through L1/L2.
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Thanks!
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I have done some past brainstorm and simulation musings related with using and/or adapting something like the current Delta IV Heavy Upper Stage to achieve some kind of earlier crewed beyond LEO exploration goals (as well related with Delta IV Heavy being used for Orion CLV duties).
I ask the reader for patience when reading what will write next (thanks in advance!).
I. Slightly Heavier Lift plus more or less existing stages (single launch BEO crewed applications)
The mentioned above brainstorms were sometimes mostly focused on integration of such stage + Orion on a single launch vehicle (based on past DIRECT J12X / J13X vehicle design iterations but you could consider something else with equivalent conceptual capability, the important is the Earth Departure stack related stuff), please see the following links for (outdated) imagery...
http://www.flickr.com/photos/simcosmos/4014041934/
http://www.flickr.com/photos/simcosmos/3179561421/
... and the following related NSF posts for extra context:
http://forum.nasaspaceflight.com/index.php?topic=18139.msg477552#msg477552
http://forum.nasaspaceflight.com/index.php?topic=12379.msg323124#msg323124
I'm only sharing these links here because they contain information that might be useful for this thread (at least given my perception of recent posts).
The advantages of using some kind of slightly heavier lift (than current EELV Heavy but not going as far as super heavy lift) single launch approach are:
a) Near delivery or even full delivery into LEO of the departure stack composed by such stage + payload (in this case, some kind of crewed spacecraft): this means that the mission stage would have almost all its main prop. load at the start of TLI... The mission capability for such stage would depend of payload mass and boil-off mitigation assumptions...
b) Another advantage of a generic slightly heavier lift capability is that something like the Delta IV Heavy Upper Stage would then perhaps not require as much modifications, at least from the perspective of boil-off mitigation: this means that the current design could almost be used - again, from this perspective - under current operational assumptions for the main burn (TLI), which would be done soon after LEO insertion (if using a parking orbit before TLI; the mission design could also consider other slightly more 'aggressive' TLI assumptions).
II. The relevance of all the above for this thread?
The moment in which additional assumptions start to be included for further changes on a chosen existing hardware / stage (such as the case of Delta IV Heavy US on recent posts), it is the moment where such changes need to be a little more clearly accounted on things like dry mass estimations, performance calculations, simulation work (and in the implementation impact on time line / cost, etc)
This also means a number of extra assumptions that need to be analyzed. If the objective is to use a dual launch of some kind of currently existing or even slightly upgraded EELV to achieve some kind of crewed beyond LEO target goal then such changes need to be accounted in the implementation / development assumptions for the mission design trade space, else the analysis - even if just a very first order analysis - might lack in detail and not be suitable for comparison with other cases.
This to say that, as most people here are aware (even if just by looking at pdf links that were provided above), the idea of using existing or slightly upgraded EELV for beyond LEO crewed exploration is not new.
Focusing in DeltaIV related hardware, there are some famous images from the pre-Constellation times where Earth Departure / Transfer Stacks for lunar missions (just an example) would be made by joining a number of modified Delta IV Heavy Upper Stages: the 'problem' is that such stages dry masses would be higher than the current operational design if needing to assume some kind of rendezvous due to lower launch lift capability... and this beyond assuming other needed adaptations related with integration for crewed duties...
... Only as a very rough example, have somewhere in my download archives, a pdf that showed a first order estimation of ~7t or so (if remembering well) for such kind of modified Delta IV Heavy US (in order to account for extra mass related with boil-off mitigation for extended orbital / mission capability and for rendezvous + docking with other stages / crewed spacecraft). Of course that such data could be first order or even outdated or I might not remembering well (might try to (re)find the pdf, if someone wishes)...
This then brings the question about the assumption of using 'something else' a bit more optimized for the mission designs in question... Such 'something else' could be one or several of the following examples... new upper stage design (instead of modifying existing stage) and/or heavier lift capability and/or depots, etc... or, also as alternative, might lead to the reduction of mission design goals / capabilities (example: assumption of much lighter crewed spacecraft launching first... something like a modified Soyuz or Shenzhou equivalent mass class spacecraft) and rendezvous within hours with upper stage being used for departure (if such stage is a cryogenic stage with very little modifications regarding currently operational designs).
To conclude, the topic of this thread is very interesting indeed but it might become even more interesting if thinking about introducing a little more data on the starting assumptions in order to better study mission capabilities vs implementation constraints (technology assumptions vs development times vs cost vs impact on mass / propulsion assumptions vs impact on mission design ground rules).
António
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I don't think anybody's mentioned it yet, but the feasibility study for sending a manned Orion to a NEO uses as its lower bookend a mission consisting of a Centaur launched on a Delta IV-H which docks with an Orion:
http://ti.arc.nasa.gov/projects/neo_study/pdf/NEO_feasibility.pdf
Of course, if Orion isn't being man-rated for launch, you'll probably want to launch a separate crew taxi or get the crew from the ISS.
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I don't think anybody's mentioned it yet, but the feasibility study for sending a manned Orion to a NEO uses as its lower bookend a mission consisting of a Centaur launched on a Delta IV-H which docks with an Orion:
http://ti.arc.nasa.gov/projects/neo_study/pdf/NEO_feasibility.pdf
Of course, if Orion isn't being man-rated for launch, you'll probably want to launch a separate crew taxi or get the crew from the ISS.
This was interesting.
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I don't think anybody's mentioned it yet, but the feasibility study for sending a manned Orion to a NEO uses as its lower bookend a mission consisting of a Centaur launched on a Delta IV-H which docks with an Orion:
http://ti.arc.nasa.gov/projects/neo_study/pdf/NEO_feasibility.pdf
Of course, if Orion isn't being man-rated for launch, you'll probably want to launch a separate crew taxi or get the crew from the ISS.
I was (still am) trying to prepare a post where would clearly share starting assumptions about Earth Departure Stage (or stack of stages) for such kind of 'lower bookend' crewed missions, I mean, missions which would assume a minimum of two launches (up to 4 EELV 'class'): at least one vehicle to launch an adaptation of an existing EELV stage (either AtlasV Centaur or Delta IV Heavy US) plus at least another launch to deliver some kind of crewed spacecraft which would rendezvous with such departure stage(s)...
However, for the moment, I would ask the readers to please look with a little more attention to the assumptions available on documents such as the one that was shared above versus look at the date they were produced vs think a little more about the starting assumptions...
This all to say that in the special case of these 'lower bookend' missions it is enough to just be a little more conservative in some assumptions to quickly make a mission concept unfeasible (which would lead to the consideration of extra assumptions and their impact on time line + development + cost).
To give an example of a constrained mission design (really, really, lower bookend), for the case of using something like an adapted AtlasV Centaur US + some kind of Orion while, at the same time, assuming slightly more conservative numbers for those components: I'm not being able to achieve much better than ~3950m/s up to 4150m/s dV budgets (preliminary math for total US + Orion dV). The Orion spacecraft assumed on some older documents also seems to have a dV of 1700m/s (older iteration) but have seen some official public NASA documents where the maximum dV available for more recent Orion iterations was closer to ~1500m/s or so... Not sure if the work about 'lower bookend' missions was updated to reflect details like these or not...
... In any case, that is why wrote my previous post: unless the starting assumptions for the mission hardware and mission design constraints are more clearly shared, it becomes a bit difficult to really brainstorm - and compare - the impact of such starting assumptions on conceptual mission profiles...
...Without such 'work' done - or without other people's work somehow 'verified' and/or updated against last known information - it is very easy to write, in these forums, about assuming a 'Delta IV Heavy Upper Stage waiting for a crewed spacecraft outside of a given ISS envelope' or writing about 'using Centaur + Orion' for these 'lower bookend' BEO missions... (slightly more challenging and time expensive is to present some math / extra assumptions to backup the thoughts and for others to review / discuss a bit further).
So, on my humble opinion, it would be a lot more interesting to fully share / verify / adapt all starting assumptions so that everyone could have a common starting ground - even if just first order - to clearly follow calculations and mission design constraints (at least believe that was the spirit of this thread, please let me know if not).
If possible, will then try to share (although not sure when will be able to do it) the range of starting assumptions that, at least (and for what is worth!) I'm using here for a modified AtlasV US, DeltaIV Heavy US and for crewed spacecraft options versus eventual 'low bookend' conceptual mission assumptions (while trying to keep things constrained to a maximum of four EELV, minimum of two).
Thanks,
António
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'Lower Bookend' Beyond Earth Orbit Crewed Missions
Definition: mission designs using existing, near-existing and/or currently ongoing testing / in-development hardware to achieve,with Orion (or other crewed spacecraft), a given beyond-LEO mission goal by using a minimum of two launches (EELV or vehicles with similar payloads range) at the earliest and most cost-effective possible opportunity.
As noted on past posts, will try to share a few starting assumptions related with this thread’s topic so that:
- other forum participants are able to review the data
- some mission design and implementation constraints are shared
- comparisons with other past or future mission assumptions, documents, etc becomes slightly easier (at least a common ground / numerical starting point is provided for a specific scenario)
For this particular post will only focus on the following case: one launcher delivering an adapted Centaur SE which would act as Departure Stage for the rest of the mission components (crewed spacecraft and perhaps mission specific hardware would be delivered by the second launcher).
Note: on later occasion might focus the attention on Delta IV Heavy Upper Stage being used as EDS but, for the moment, starting then by a lower bookend case…
Whenever possible will try to use existing or near-existing hardware and reference to source data. More aggressive assumptions are certainly possible (subjected to the impact on time line + cost, etc…). As a last introductory note, of course that readers are free to review this text but, if doing so, please try to do it with some ‘substance’.
I. Atlas V Centaur US (single engine, adapted for beyond LEO missions)
3500 kg : inert mass
20830 kg : prop. mass (at lift-off)
99200 N : thrust
450.5s : ISP
Notes:
- Accordingly with ‘AtlasV Users Guide 2010 (rev11, March)’ inert mass for 5X1 Centaur SE is 2247 kg and, for AtlasV Heavy, such stage mass is given as 2316 kg; on the ‘Centaur Application to Robotic and Crewed Lunar Lander Evolution’ paper, the Centaur SE is 2500 kg: I’m using this latest value together with the addition of 800 kg for an ‘Extended Duration Mission Kit’ (also from public LM/ULA materials) in order to obtain what believe to be a better representation of a Centaur SE stage adapted for operation within 7 days of being launched (equal to say, a better representation for the kind of mission profiles being referenced in this thread…)
- An extra mass of ~200 kg is assumed on the stage’s ‘inert mass’ to account for a passive LIDS (similar to the hardware integrated on HST on STS-125): I’m not sure about the mass of such passive version (feel free to point for a source regarding that value)
- Will also share results calculated by using 450.5s ISP / 0.25% prop. left on departure stage versus 445s ISP / 1% prop. left on the stage, to compare with a more conservative dV budget calculation (and to calculate an average dV budget).
- Boil-off will be assumed at 0.2% / day of the initial LEO delivered prop. load (this assuming that the stage is fully delivered into LEO by the launch vehicle)
Sources:
http://www.ulalaunch.com/site/pages/Education_PublishedPapers.shtml
http://www.ulalaunch.com/site/pages/Products_AtlasV.shtml
http://www.ulalaunch.com/site/docs/product_cards/guides/AtlasVUsersGuide2010.pdf
II. Crewed Spacecraft & Mission Hardware
II.a) Orion
Will assume that some kind of BEO Orion implementation will happen in the future. Because I’m not sure about what is the most current state of Orion mass breakout will use past information / extrapolate eventual Orion mass breakouts for the intended mission profile.
Note: another option would be to look at this as a ‘black box’ (placeholder number) for a given total mass which could represent a generic crewed spacecraft alone (not necessarily Orion) or a crewed spacecraft + mission related module & payloads to, at least, calculate the dV provided by the modified Centaur SE.
Orion CM (lunar / BEO variant) has been noted, on earlier iterations, at ~8t… up to ~8.7t up to perhaps 9.3t, at lift-off, for a crew of four astronauts living inside it for periods up to ~18 days (?) (when accounting for Constellation’s Lunar Mission profile and for maximum loiter times on LEO, LOI, TEI phases + when taking in consideration the length of the journey to/from Moon).
Orion SM: have seen estimations ranging from ~3.6t on very early iterations up to ~4.3t up to ~>4.5t (?)… Main prop. load of at least 8t… Will assume main engine at 326s ISP and from 0.25% up to 1% prop. left on the service Module (for FPR / Residuals, etc.)
As mentioned above, not very easy to have a clear picture about the current Orion state… Some of its mass growth and configuration changes are/were very closely related with the specific AresI integration, other part of the spacecraft mass growth is ‘natural’ of any development process…
What will present next are then some ‘options’ for several Orion mass breakouts / rough first order performance calculations (kept SM mass the same and varied SM prop for two CM cases)… This might hopefully be representative and conservative enough for a first order estimation (while assuming full LEO delivery of the vehicle)...
22.3t : 9.3t (CM) + 4.5t (SM) + 8.5t (SMprop) ~1515 m/s
21.8t : 9.3t (CM) + 4.5t (SM) + 8.0t (SMprop) ~1443 m/s
20.2t : 9.3t (CM) + 4.5t (SM) + 6.4t (SMprop) ~1203 m/s
21.7t : 8.7t (CM) + 4.5t (SM) + 8.5t (SMprop) ~1568 m/s
21.2t : 8.7t (CM) + 4.5t (SM) + 8.0t (SMprop) ~1495 m/s
20.2t : 8.7t (CM) + 4.5t (SM) + 7.0t (SMprop) ~1343 m/s
Note:
- If the total mission scenario would take ~>18 days then the crew would probably have to be reduced to three or even to two astronauts, depending of extra mission requirements (such as desired contingency margins, mission hardware, etc)…
II.b) Dragon
Not sure about what would be the mass breakout for a lunar Dragon variant: perhaps some of the eventual mass growth could be achieved with crew number considerations and by carrying less cargo on the trunk. For the moment will share a first order estimation of a Dragon mass breakout for LEO, only to provide a starting point for eventual later brainstorms (feel free to correct data):
3250 kg : capsule mass
2500 kg : 7 (?) max. crew or less crew (3?, 4?) + some combination of internal / external cargo, etc
1290 kg : OMS / RCS prop (270s ISP?)
681 kg : Trunk
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7721 kg : Dragon mass (crewed flights)
Falcon9 assumes a payload of ~10450 kg into 200km, 28.8 inc. There seems to exist enough performance margin to either include a standard LAS design (released during ascent) or some kind of dual-duty abort system on the capsule itself (simulations needed!).
Assuming 270s ISP and ~1% left on the integrated SM tanks, the above numbers would result on a dV budget of ~479 m/s.
Only for reference, assuming a cargo only variant (2500 kg internal cargo + equipment, etc and ~2729 kg of external cargo carried on the trunk, for a total cargo of ~5.2t and a total spacecraft mass of ~10450 kg, the dV budget (again keeping 1% prop) could be ~345 m/s.
This all to write that in theory, a lunar Dragon with a crew of ~2 to ~3? astronauts could perhaps have enough resources to return from EML2 (using propulsive swing-by at the Moon): extra study needed about the mass / performance impact of the changes that would be required to transform the conceptual LEO Dragon into an even more conceptual Lunar Dragon.
II.c) Mission Hardware
Crew size could also be strongly dependent on the amount of extra mission hardware that would need to be carried in this kind of ‘lower bookend’ missions…
Back to Orion mass breakout examples, the reader can, for example, look at the heavier crew command module and think about reduced crew size (2 astronauts?) and some of that 9.3t mass as being ‘mission hardware’ (either transported inside the CM – such as EVA suits and other tools - or attached to the SM exterior…).
Another option would be to assume an extra mission module… However, such extra module (plus its internal contents / external payloads) could well reach a mass from ~3t up to ~>4t (depending of mission needs)…
… Always depending of extra assumptions – and focusing for the moment only on Orion - I’m not sure for which mission design such extra module would be feasibly integrated IF assuming only two launches of existing EELV / EELV Heavy (or equivalent payload vehicles) and without assuming something else beyond the AtlasV Centaur SE modifications mentioned above… Yes, extra lift capability under the form of EELV upgrades could also be assumed but then all that starts adding to the timeline (I’m already assuming several adaptations to Centaur SE upper stage, its eventual integration on a launcher that might not be an AtlasV, development of Orion for BEO, etc)…
III. Mission Design Constraints (Launch)
The mission would start with the launch of the crewed spacecraft or even the departure stage… A good number of extra simulation and related considerations would be needed to decide which payload would be launched first and to also better study injection targets for each payload vs a better definition of rendezvous procedures vs Earth Departure setup requirements…
If adding a mission module of ~3t up to ~>4t, such module could perhaps be launched (depending of mass and launch vehicle choice) together with a non crewed Orion vehicle: in such case, the delivery requirement could be something like ~>25.25t total spacecraft mass (not counting with adapters) into ~250 km altitude…
Will however focus on the ‘no-mission module case’ for this specific scenario where the departure stage would be a modified AtlasV SE US: as roughly represented above, the Orion could then have a lift-off mass of ~22t for an average dV of ~1500m/s or so...
The AtlasV Centaur SE based departure stage could have a total mass of ~24.33t…
Both these payload targets seem to be feasible (need to double-check / simulation work needed) for Delta IV Heavy (which in this time frame would use RS-68A): it would perhaps require dual launch within up to ~5 days (extra study needed, have somewhere a pdf about AresI+V regarding this issue of launch order vs lunar mission). If using dual Delta IV Heavy, one of them would need Human Rated Kit (as well related facilities adaptations).
Another Heavy vehicle option (perhaps AtlasV Heavy… or Falcon9 Heavy!, or AresI!... although AresI injection would be sub-orbital and require a few extra considerations, etc) or some other EELV upgraded iteration could also be assumed (which, depending of specific assumption, would also add to development and/or timeline and/or cost)…
Please remember that I’m still only talking here about an entry level for lower bookend type of missions. This all needs to be taken in consideration…
Yet another ‘nearer-time’ alternative could be to assume Delta IV Heavy launching the AtlasV upper stage (which poses its own integration challenges), something like an AtlasV 552 launching ~21t Orion with no crew (the dual Centaur would have to be fielded…) and ‘something else’ (not many options left, really...) launching the crew… although this would become a 3 launch scenario with things starting to complicate… near-ISS space could also be used as a stationary / assembly point for the crew to wait for the Earth Departure Stack docking but that would introduce yet another layer of complications for this type of ‘lower bookend’ missions... better to avoid that for this specific case.
Given all the constraints, not sure when even this kind of lower bookend mission would be possible… maybe I’m being pessimist…
Summing up, would perhaps then baseline two launches of the heaviest EELV (or similar class) possible vehicles available at the date: one perhaps launching EDS alone, the other – human rated - launching the crewed spacecraft. Considerations about which launch would be first or about eventual mission module would need a much better definition of mission requirements, mission hardware and specific simulation work.
III. Mission Design Constraints (dV budgets)
Regarding the dV budget, under what have assumed:
3500 kg : AtlasV Centaur SE US + Kit + Passive LIDS
20538 kg : Centaur main prop. load (after 7 days @ ~0.2% boil-off / day)
20900 kg : Orion: 9.3t (CM+mission hardware?) + 4.5t (SM) + 7.1t (SMprop, after rendezvous, initial load would be ~>8t, simulation needed)
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44938 kg : total mass at start of departure burn
Depending of how math is made:
dV1 (Centaur): ~2688 m/s (450.5s ISP / ~0.25% prop. left)
dV1 (Centaur): ~2661 m/s (450.5s ISP / ~1.00% prop. left)
dV1 (Centaur): ~2656 m/s (445s ISP / ~0.25% prop. left)
dV1 (Centaur): ~2629 m/s (445 ISP / ~1.00% prop. left)
dv2 (Orion): ~1311 m/s to 1322 m/s (depending of FPR, etc)
This all gives about ~3975 m/s total dV budget (in average) for the above Centaur SE + Orion assumptions (two launches).
III.a) Performance Comments
- Such kind of average dV estimation and ~18 day journey duration could be compatible with a 3 astronauts crew delivery into EML2 and return from there:
~1 x 3148.5 m/s : TLI (shared by Centaur SE + Orion)
+2 x 184.1 m/s : propulsive swing-by at the Moon (In/Out)
+2 x 147.5 m/s : EML2 (In/Out)
(with ~163 m/s available for MCC, station keeping)…
Source: http://forum.nasaspaceflight.com/index.php?topic=1337.msg18213#msg18213
(RE: An Alternative Lunar Architecture)
It would be better if the crew could have ‘something’ waiting for them at EML2, else they would arrive there only to return on the next hours (and if that would be the mission objective perhaps a simpler lunar fly-by profile would be better)
- Such kind of ~3975 m/s dV and more than 18 day mission having EML2 (or EML1, also using swing-by, see pdf below) could perhaps also be achieved by a reduced crew of just two astronauts while, at the same time, perhaps also carrying a little of extra mission specific payload (IF the goal of the mission would be to service something at EML2, for example): not sure however how much payload would be feasible (depending of assumptions for the crewed spacecraft, etc).
On my humble opinion and under my clumsy first order considerations, little else could perhaps be assumed (do not see great opportunities for reaching a BEO object) unless going for more capable launch vehicles and/or more capable propulsion stage assumptions and/or optimized spacecraft assumptions and/or depots and/or using specific date geometries together with extra trajectory tweaks for intended mission objectives…
… Each or several groups of those extra assumptions would need to be taken in account in order to better compare implementation constraints, but all that would add to the time line of the ‘simplest’ case that have tried to very preliminarily study and share here (extra study, assumptions refinement, simulation work and more specific design assumptions would be needed).
So, summing up, have tried to share some considerations about:
- AtlasV Centaur SE (with kit + passive LIDS) fully delivered into LEO
- Orion (also fully delivered into LEO) launched within ~5 days or so and rendezvous, within ~2 days, with departure stage
- departure dV provided by upper stage burn + Orion undocking and making the final commitment for such burn
- Orion making rest of middle course correction / main burns
- mission duration of ~18 days for crew of 3?
- mission duration of ~>18 days for crew of 2? (+ X? payload?)
III.b) Related readings and final comments:
http://servicingstudy.gsfc.nasa.gov/presentations_final/day2/Harley_Thronson/Thronson_servicing_workshop.pdf
(http://servicingstudy.gsfc.nasa.gov/workshop_1_presentations.htm)
DOES THE NASA CONSTELLATION ARCHITECTURE OFFER OPPORTUNITIES TO ACHIEVE SPACE SCIENCE GOALS IN SPACE?
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080032732_2008031420.pdf
(IAC-08-A5.3.6)
A few differences on the IAC-08-A5.3.6 pdf (versus what I have written above) are:
- No boil-off assumed? (absence of long duration kit on Centaur?)
- Quick (next orbit?) back-to-back launch of dual AresI from both KSC pads 39 (instead of assuming other launch vehicles) + rendezvous within hours (instead of allowing for up to ~5 days difference between launches, from the moment the Centaur would be launched?)
Comment: If using aggressive mission components delivery, the crewed spacecraft (perhaps launched first into a lower orbit) could rendezvous within hours with the departure stage but that does not leave much margin for unexpected events…
… On another hand if using slightly less aggressive rendezvous assumptions, the departure stage would need addition of extra boil-off reduction kit (probably similar to what I have assumed) and rendezvous time line of mission components could be a little more relaxed although the kit assumption would have impact on how early such boil-off reduction kit would be available for operational use… as far as I’m aware, an early sun shield demo is expected for ~2011 (that would be only one component of the kit).
- Inert mass for Centaur SE of ~1.8t (I have used ~2.5t for a modified Centaur SE + 800 kg Kit)
- Lower Orion mass breakouts (CM + SM) than what I have assumed: it would be really interesting to have an update regarding Orion status, despite all the indefinition we are experimenting nowadays…
- ~>2.7t airlock module (the differences on Centaur assumptions + CM + SM almost cover the minimum module mass range): on the IAC-08-A5.3.6 paper the AresI also seems to be defined as a ~>26t payload vehicle
- Different dV calculation methods and trajectory assumptions: haven’t fully verified those. On my calculations tried to keep some FPR / Residuals on departure stage, SM, etc, not sure if that was the case on the pdf. This post is already long enough (!!!): feel free to define ground rules, compare and refine estimations and methods.
Real Final Comments:
Quick note about assuming the crewed spacecraft with a total mass of ~15t instead: this could allow the Centaur to make full TLI. Such crewed spacecraft could look like an Orion CM with a smaller SM or like a Shenzhou with larger SM or with an extra mini-stage (or, more relevant, Lunar Dragon with some kind of SM connected to the trunk: mission capability would depend of assumptions).
To end, could also write a similar preliminary analysis / references about a possible brainstorm 'next level' (using / adapting DeltaIV Heavy Upper Stage, based on some pdf / past notes / simulation work / math) but that text could probably reach a point where it would be 'simply' better to start assuming or a new stage development or heavier lift capability than the one expected for Delta IV Heavy (or similar EELV-class) or both those changes and/or even several other extra assumptions...
So, will save forum space and, at least for now, post just this ‘lower bookend’ case where the Departure Stage would be an adapted AtlasV Centaur SE... At the same time provided a few related links / references and some numbers about what could be the mass breakout of such adapted stage, Orion, Dragon, etc which might provide a good starting point if someone else wishes to further refine calculations and mission design / implementation constraints for this kind of lower bookend cases.
Thanks,
António
PS: this was a loooong post: my apologies in advance for any less clear text / math / errors and/or typos (wrote this in pieces and then did a bit of intensive final copy+paste)
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Another reason to select Centaur rather than DCUS is because over the program's history Centaur vehicles have been exposed to (and survived) a wide variety of launch environments. DCUS has flown about a dozen times total, on Delta III and Delta IV vehicles only.
And speaking of launch environments, what are the chances a Centaur (or DCUS) could be flown inverted, i.e. nose-to-nose with the second stage of the launch vehicle? This might substantially reduce the mass of the interstage hardware.
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Another reason to select Centaur rather than DCUS is because over the program's history Centaur vehicles have been exposed to (and survived) a wide variety of launch environments. DCUS has flown about a dozen times total, on Delta III and Delta IV vehicles only.
And speaking of launch environments, what are the chances a Centaur (or DCUS) could be flown inverted, i.e. nose-to-nose with the second stage of the launch vehicle? This might substantially reduce the mass of the interstage hardware.
AtlasV Centaur vs DeltaIV Heavy US vs longer space endurance kits
On my opinion, something like the assumption of modifying an existing Centaur (vs DeltaIV Heavy US) could make sense (for the specific case of lower bookend missions while using existing or near-existing launchers) given the following reasons:
- on-going work about the topic (longer endurance kits) seems to be already happening and using Centaur as a test platform (at least in a more active way than on Delta IV Heavy US case)
- the total mass of such modified Centaur could apparently still be within the maximum payload range for a good and stable enough LEO insertion by an Heavy EELV (or equivalent ‘class’ vehicle)
If wishing to take the brainstorm to a ‘next level’ and extrapolate about modifications done to the Delta IV Heavy Upper Stage… Well, the current stage is ~3490 kg for a propellant load of ~27.2t, thrust of ~110 kN, vac. ISP of ~462s (459s if wishing to be conservative on calculations)…
… The above means that the total mass of the stage, *without modifications* is ~3.49t + 27.2t = 30.69t… This could be OK if assuming a heavier launch capability to single-launch the departure stack (crewed spacecraft + eventual mission related hardware + departure stage) and with such stage making departure burn soon after reaching orbit (in a similar way to what have mentioned above on DIRECT related brainstorms): such single-launch heavier lift capability could at least workaround the issue of not using the stage much beyond its current endurance limits (and/or minimize the modifications needed)…
However, if assuming instead something like current EELV / Delta IV Heavy (or even something like AtlasV Heavy, Falcon Heavy, etc)… Well, a modified Delta IV Heavy Upper Stage *with extra kits for longer space endurance* and also with extra mass related with mission hardware integrations (docking / integration with mission payload or extra stage, etc) could eventually weight more than 31t, depending of estimations for the modifications vs conceptual mission profiles… The mass of the stage could jump from those ~3.49t up to something closer to ~5t or even greater (have seen an older pdf, where a probably conservative brainstorm (?) assumed ~7t, including stuff such as densified prop. for a load of ~28.2t, boil-off reduction kit, docking hardware capability with other Delta IV Heavy US… if remembering well, the two-stage stack would allow for a TLI payload of ~35t or so).
Summing up, if not assuming heavier launch vehicles (between ~50t up to ~80t or so, being that such heavier vehicles could also make good use of longer endurance stages althought not requiring it for a first phase) and/or several other things (such as depots, etc), we could end-up by talking about a longer endurance modified Delta IV Heavy Upper Stage resulting in a total stage mass ranging from ~32t up to ~36t… Depending of extra assumptions, this could require a launch vehicle capable of at least ~35t up to 40t or so delivery into at least 240 km up to 300 km 29 inc Earth orbit… this, if the objective would still be to deliver a fully loaded departure stage – in this case based on Delta IV Heavy design - to rendezvous with mission payload (and/or with other stage to form a departure stack)…
So, from this perspective, not sure if considerations about flying the departure stage upside down in order to eventually save on interstage mass would be relevant because:
a) longer endurance kit additions to current Centaur + interstage, etc seems to still be in the range of something like Delta IV Heavy (or AtlasV Heavy or Falcon9 Heavy, if any of those would be fielded one day…); extra refinement / study / simulation work needed...
b) if talking instead about longer endurance kits for Delta IV Heavy US we could probably end-up by talking on requirements for larger delivery capabilities (= extra modifications) of the launch vehicles I mentioned above… If such would be the case, the vehicles would need to be modified (or any of their evolution phases implemented and/or something like at least an ‘entry-level’ SDLV or clean sheet design, etc) in order to meet lift requirements… Or else… some other changes regarding technology and/or mission design, etc would have to be assumed… (articulated with better assumptions for BEO crewed spacecraft configuration...)
… That is why wrote somewhere on the previous posts that this could quickly lead to considerations about new upper stage designs, stages which would be built and optimised - from day one - with larger endurance applications in mind… (versus talking about adapting current upper stages).
Extra Comment about AtlasV Centaur vs Lower Bookend Missions
Having written the above, that is why tried to focus the previous looooong post only on lower bookend missions using Centaur as a tesbed for slightly longer endurance applications (than what happens nowadays): anything else would probably have impacts on the implementation timeline as well need the assumption of extra development / monetary / technology stuff… In which case it would probably be better to leave behind the ‘lower bookend’ mission concepts / assumptions and go instead for something slightly more optimised and adapted to mission needs (in terms of hardware design assumptions) and less constrained in terms of mission capabilities…
Would also like to remember the readers that on my previous post, haven’t really taken in account things such as gravity losses: with a T/W of ~0.23 for something like the modified Centaur SE + Orion described on such post, that could mean gravity losses in the order of ~75 m/s or so (this looking at a graphic for a 400 km departure, from ESAS report)… This could be in part mitigated by using a much lower departure altitude and that is why also mentioned on such post that extra assumptions refinements (mass breakouts, for example) and a good amount of extra simulation work would be needed about LEO injection targets for the modified Centaur and for the crewed spacecraft vs order of launch vs rendezvous procedures vs departure burn setup, etc, etc
António
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And speaking of launch environments, what are the chances a Centaur (or DCUS) could be flown inverted, i.e. nose-to-nose with the second stage of the launch vehicle? This might substantially reduce the mass of the interstage hardware.
They are designed to take loads from a certain direction
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This topic is being brought up in another thread, so I'm bumping this thread.
Another thought: Dragon on Delta IV Heavy could go straight to EML1/2. I think Dragon would have plenty of delta-v capability (~500-700m/s) to handle EML1/2 insertion and departure (total of 350 m/s?).
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Hello all,
I'm trying to (slooowly) return to some addon development work for Orbiter Space Simulator and, when looking at 'simcosmos' archives, I have decided to unfreeze / revamp an old 2006 Centaur 3D model in order to transform it into an Earth Departure Stage for an Orion spacecraft (http://www.orbiter-forum.com/showthread.php?p=496674&postcount=25).
This is part of an 'what if' conceptual side-exercise about Constellation Program and is also a kind of a follow-up to the brainstorms I have written in this thread (please see previous posts).
It may also be more or less relevant for current days (Orion MPCV), because I'm assuming an heavier CxP Orion (~24.5t CSM at liftoff). This kind of Lower Bookend Mission does not give space for wild mission planning (at least not with an heavy assumption for the spacecraft mass)... As noted, with these heavy constraints, it will be all about a lunar flyby (unless doing extra tweaks on several assumptions)... As also referenced, a crewed spacecraft with a total mass between 6t up to 8t or so would obviously be much better (Orion's CM alone is already above those values).
To conclude, my apologies for bumping the old topic but, independently of something like this (Centaur modification for EDS duty on a crewed mission) ever having a chance or not of becoming reality (in particular with a lighter crewed spacecraft), at very least I think that this is a right place to complete my previous participation on this thread with (the attached below) 'work-in-progress' contribute to the 'eye-candy' department.
Cheers,
António Maia
PS: Anyone knows if there is still any intention to demonstrate a Centaur Sun shield in an actual flight? Thanks!