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What is the actual delta-V split and payload for the Delta IV Heavy? Using astronautix.com numbers, the total delta-V exceeds 10.5 km/s.Also, what would be the payload for a LV consisting of a single CBC and a J-2/J-2X upper stage? I felt that we needed a Q&A thread for the Delta IV.
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What are the injectors upgrades done for Ares V looking like from Delta IV:s point of view?
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tnphysics - 13/8/2007 7:08 PM
Also, what would be the payload for a LV consisting of a single CBC and a J-2/J-2X upper stage?
Since no one is working such a vehicle, there is not a good answer
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meiza - 13/8/2007 7:34 PM
What are the injectors upgrades done for Ares V looking like from Delta IV:s point of view?
The RS-68 upgrades were done for D-IV
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and lemme guess that the performance numbers are ITAR? :)
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I'm American, PM me if it's an ITAR problem. Both EELV upper stages are too small. The Delta IV's US should be the size of the S-IVB.
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it is also propriety
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All I need to account for is a 1 km/s overperformance when delivering a 25.8 mt payload.
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tnphysics - 13/8/2007 8:37 PM
. Both EELV upper stages are too small. The Delta IV's US should be the size of the S-IVB.
Not for GTO and GSO missions, that's what they were designed for
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What's the difference between the RL 10 B 2 engines D IV is using and the ones used on Atlas?
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pippin - 14/8/2007 7:50 AM
What's the difference between the RL 10 B 2 engines D IV is using and the ones used on Atlas?
Extendable nozzle, thrust level, etc. Many differences
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tnphysics - 13/8/2007 4:08 PM
What is the actual delta-V split and payload for the Delta IV Heavy? Using astronautix.com numbers, the total delta-V exceeds 10.5 km/s.Also, what would be the payload for a LV consisting of a single CBC and a J-2/J-2X upper stage?I felt that we needed a Q&A thread for the Delta IV.
Since you're asking a hypothetical question (current D-IV uses RL10B-2 and not J-2/J-2X), you can get a comparable number by taking the existing payload, delta-vee, and B-2's engine Isp, then swap out with the J-2/ J-2X Isp to see what additonal payload gain (if any) can be achieved. You'll need to run this throughtout the trajectory since the CBC will take a hit on delta-vee because of additional payload weight as well. I would also allow some gravity delta-vee losses on the vehicle.
You'll then need to substract the additional J-2/J-2X engine weight from that additional payload weight gain.
Now the fun part is, should you find yourself run out of 2nd stage propellant too early, it's time to re-size the 2nd stage tanks to allow to carry more propellant. But that would be difficult since you only allow a single CBC and soon we'll run into the thrust limitation of the RS-68! :laugh:
But don't be discouraged. Atlas V can't do any better as their initial liftoff thrust-to-weight is already dangerously low. :bleh:
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Jim - 13/8/2007 4:43 PM
meiza - 13/8/2007 7:34 PM
What are the injectors upgrades done for Ares V looking like from Delta IV:s point of view?
The RS-68 upgrades were done for D-IV
You mean for the Air Force and NASA....
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pad rat - 14/8/2007 6:43 AM
"I'm American, PM me if it's an ITAR problem."
How does one prove that assertion on the internet?
Same as "I'm an Iranian student doing a project on ground-to-air missile...."
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tnphysics - 14/8/2007 5:50 AM
All I need to account for is a 1 km/s overperformance when delivering a 25.8 mt payload.
I came up with only about half that much "overperformance" for such a payload, assuming full propellant loading and an extra tonne of "payload" used for adapter hardware, etc. - but still hundreds of meters/second delta-v margin on paper for a big LEO payload. Calculating Delta IV Heavy is tricky due to all of the throttling up and down during the CBC burns.
I can think of several explanations for the extra delta-v. One could be that published payload capabilities reflect structural limitations rather than propulsion limits. There's a lot of discussion in the Planners Guides about payload adapter/payload center of gravity/vibration limits and the like.
Another possibility is that Delta IV dry masses are higher than generally believed. I think that this is probably true, but I doubt that it accounts for all of the difference. The strap-on CBCs are heavier than the core CBC, for example, when the nose cones and interstages are included.
A third explanation is related to the low upper stage thrust - it only produces a 0.19 to 0.36 thrust to weight ratio for this payload mass if full propellant loading is assumed, which means quite a bit of gravity loss to counter during a LEO ascent. The Delta IV Planners Guide showed a total upper stage burn time of only 912 seconds for a LEO mission, which implies that the upper stage would only carry about 22 tonnes of propellant, 5 tonnes less than its maximum load. That knocks off about 230 m/s delta-v right there.
A final explanation may be that the Delta IV people are sandbagging a bit, downplaying the real capability that might be available.
- Ed Kyle
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edkyle99 - 14/8/2007 3:17 PM
tnphysics - 14/8/2007 5:50 AM
All I need to account for is a 1 km/s overperformance when delivering a 25.8 mt payload.
I came up with only about half that much "overperformance" for such a payload, assuming full propellant loading and an extra tonne of "payload" used for adapter hardware, etc. - but still hundreds of meters/second delta-v margin on paper for a big LEO payload. Calculating Delta IV Heavy is tricky due to all of the throttling up and down during the CBC burns.
........
A final explanation may be that the Delta IV people are sandbagging a bit, downplaying the real capability that might be available.
- Ed Kyle
The better to make it look like you can't loft Orion with it??
You could probably find out how much margin there was in the US by looking at the demonstration flight where the main's shut down early.. see how deltaV they manage to make up with the US. Wouldn't tell you anything about margin for the core though.
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TrueBlueWitt - 14/8/2007 2:31 PM
edkyle99 - 14/8/2007 3:17 PM
tnphysics - 14/8/2007 5:50 AM
All I need to account for is a 1 km/s overperformance when delivering a 25.8 mt payload.
I came up with only about half that much "overperformance" for such a payload, assuming full propellant loading and an extra tonne of "payload" used for adapter hardware, etc. - but still hundreds of meters/second delta-v margin on paper for a big LEO payload. Calculating Delta IV Heavy is tricky due to all of the throttling up and down during the CBC burns.
........
A final explanation may be that the Delta IV people are sandbagging a bit, downplaying the real capability that might be available.
- Ed Kyle
The better to make it look like you can't loft Orion with it??
Nah. The numbers were in the original Delta IV Planners Guide published in 2000 or so, years before CEV/Orion.
You could probably find out how much margin there was in the US by looking at the demonstration flight where the main's shut down early.. see how deltaV they manage to make up with the US. Wouldn't tell you anything about margin for the core though.
Yes. And terrific - there goes my evening!
- Ed Kyle
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edkyle99 - 14/8/2007 12:17 PM
tnphysics - 14/8/2007 5:50 AM
All I need to account for is a 1 km/s overperformance when delivering a 25.8 mt payload.
I came up with only about half that much "overperformance" for such a payload, assuming full propellant loading and an extra tonne of "payload" used for adapter hardware, etc. - but still hundreds of meters/second delta-v margin on paper for a big LEO payload. Calculating Delta IV Heavy is tricky due to all of the throttling up and down during the CBC burns.
I can think of several explanations for the extra delta-v. One could be that published payload capabilities reflect structural limitations rather than propulsion limits. There's a lot of discussion in the Planners Guides about payload adapter/payload center of gravity/vibration limits and the like.
- Ed Kyle
What delta-vee were you guys assuming? What orbit are you trying to get to?
Have you consider possible additional payload FAIRING weight due to a much larger payload? What size payload (L X D X H) would constitute a 25.8 mt payload?
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Propforce - 14/8/2007 2:48 PM
edkyle99 - 14/8/2007 12:17 PM
tnphysics - 14/8/2007 5:50 AM
All I need to account for is a 1 km/s overperformance when delivering a 25.8 mt payload.
I came up with only about half that much "overperformance" for such a payload, assuming full propellant loading and an extra tonne of "payload" used for adapter hardware, etc. - but still hundreds of meters/second delta-v margin on paper for a big LEO payload. Calculating Delta IV Heavy is tricky due to all of the throttling up and down during the CBC burns.
I can think of several explanations for the extra delta-v. One could be that published payload capabilities reflect structural limitations rather than propulsion limits. There's a lot of discussion in the Planners Guides about payload adapter/payload center of gravity/vibration limits and the like.
- Ed Kyle
What delta-vee were you guys assuming? What orbit are you trying to get to?
Have you consider possible additional payload FAIRING weight due to a much larger payload? What size payload (L X D X H) would constitute a 25.8 mt payload?
I'm assuming a 5.1 x nearly 23 meter fairing that weighs close to 5 tonnes. I'm also assuming payload adapter hardware that adds an extra three-quarters of a tonne. I'm guessing that something in the neighborhood of 9,500 meters/second ideal delta-v would be budgeted for a LEO ascent, which provides a bit of padding for unknowns like gravity losses and covers a variety of inclination assumptions. I'm remembering my propellant residuals (something like 1%). I'm guessing a bit when it comes to the CBC throttling and burns, and trying to be conservative with the specific impulse numbers during the first four minutes of flight. As I mentioned before, I've also gone ahead and offloaded five tonnes of upper stage propellant.
With all of that, I'm still coming up with no less than 9,800 meters/second ideal delta-v, roughly.
- Ed Kyle
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The Delta IV upper stage has a very low T/W (0.2) for large LEO missions. Gravity losses will be much higher than "normal".
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What payload do you calculate?
Could anyone calulate the numbers with POST, etc?
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edkyle99 - 14/8/2007 1:17 PM
I'm assuming a 5.1 x nearly 23 meter fairing that weighs close to 5 tonnes. I'm also assuming payload adapter hardware that adds an extra three-quarters of a tonne. I'm guessing that something in the neighborhood of 9,500 meters/second ideal delta-v would be budgeted for a LEO ascent, which provides a bit of padding for unknowns like gravity losses and covers a variety of inclination assumptions. I'm remembering my propellant residuals (something like 1%). I'm guessing a bit when it comes to the CBC throttling and burns, and trying to be conservative with the specific impulse numbers during the first four minutes of flight. As I mentioned before, I've also gone ahead and offloaded five tonnes of upper stage propellant.
With all of that, I'm still coming up with no less than 9,800 meters/second ideal delta-v, roughly.
- Ed Kyle
At risk of pointing out the obvious, we're talking about existing DIV-H capability, right?
If that's the case, although I don't have the LEO delta-vee number but I do have the GTO detla-vee number (for GTO: 167 x 35,788km (90 x 19,323 nmi), 27 deg). I get the delta-vee of 11.8 km/sec on the DIV-H based on a payload weight of 13,130 kg (13.13 tons?). I assume a metallic fairing weight a little over 4 tons, plus payload adaptor weight.
This works out roughly about 9,800 m/sec for 28.5 mt payload but I did not adjust the fairing weight.
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tnphysics - 14/8/2007 4:15 PM
What payload do you calculate?
Could anyone calulate the numbers with POST, etc?
POST doesn't help you if you don't have all the right models (aero, etc.).
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Propforce - 14/8/2007 6:26 PM
edkyle99 - 14/8/2007 1:17 PM
I'm assuming a 5.1 x nearly 23 meter fairing that weighs close to 5 tonnes. I'm also assuming payload adapter hardware that adds an extra three-quarters of a tonne. I'm guessing that something in the neighborhood of 9,500 meters/second ideal delta-v would be budgeted for a LEO ascent, which provides a bit of padding for unknowns like gravity losses and covers a variety of inclination assumptions. I'm remembering my propellant residuals (something like 1%). I'm guessing a bit when it comes to the CBC throttling and burns, and trying to be conservative with the specific impulse numbers during the first four minutes of flight. As I mentioned before, I've also gone ahead and offloaded five tonnes of upper stage propellant.
With all of that, I'm still coming up with no less than 9,800 meters/second ideal delta-v, roughly.
- Ed Kyle
At risk of pointing out the obvious, we're talking about existing DIV-H capability, right?
If that's the case, although I don't have the LEO delta-vee number but I do have the GTO detla-vee number (for GTO: 167 x 35,788km (90 x 19,323 nmi), 27 deg). I get the delta-vee of 11.8 km/sec on the DIV-H based on a payload weight of 13,130 kg (13.13 tons?). I assume a metallic fairing weight a little over 4 tons, plus payload adaptor weight.
This works out roughly about 9,800 m/sec for 28.5 mt payload but I did not adjust the fairing weight.
Yes. Existing Delta IV-H, or at least what we *think* the existing Delta IV-H is.
Playing around with what was reported about the first Delta IV-H, it appears that the 6 tonne payload ended up 476 meters/sec short of the planned near geostationary orbit. My model shows that the 8 second short CBC strap-on and 9 second short core CBC burns would have caused a shortfall of nearly 600 meters per second. The timing is important, but only rough numbers were provided in the Air Force press release about the failure. As a first guess, then, I'm going to assume that the upper stage must have had a bit more than 120 meters per second of planned delta-v margin that it ended up burning off in this instance, making up part of the loss.
That planned orbit would have required roughly 3,940 meters per second more delta-v than a circular 185 km orbit. My model gives about 13,720 meters per second ideal delta-v for a Delta IV-H with a 6 tonne payload. Subtracting 120 meters per second and 3,940 meters per second from that gives 9,660 meters per second to get to a 185 km LEO. There's a lot of wiggle room in there, of course, that makes it easy to see how a 9,700 to 9,800 meters/second ideal delta-v to LEO range might be expected.
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
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edkyle99 - 14/8/2007 10:00 PM
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
It's not the pure hydrogen nature of the rocket that affects this, it's almost all down to the upper stage design which one can only assume was influenced by the choice of reference missions, although some Delta folks may be able to come in and confirm or deny. GTO (and GSO) missions are much higher energy than LEO missions and so involve a lot more flight at low vertical accelerations and hence gravity loss.
The Atlas V mission planner's guide clearly shows this in the section on intermediate orbits from CCAFS. Looking at a 63.4 degree MEO and the 531/532 pair, the second RL10 buys you 1400 kg additional payload to a 1000km orbit but only 500 kg more to a 9000km orbit. Lockheed and now ULA only offer single-engine Centaurs to GTO.
Boeing designed the Delta III second stage around a single RL10-B2 for GTO missions with maybe the odd MEO flight here and there. 700 second burn time to GTO isn't so bad, especially with that big nozzle giving you 462 (!) seconds of specific impulse. Add 23,000 pounds of propellant and switch to an LEO mission, and all of a sudden you have a 900-second second-stage burn before you get to SECO-I, 1200 seconds after liftoff. Adding a second engine would really help things, but there's you can't fit two 2.2 meter diameter nozzles in a 5-meter stage with plausible clearances. So higher gravity losses it is.
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What if you used two short nozzle (Centaur like) RL-10:s on the upper stage, what would the numbers look like then? And which would be easier, making a short nozzle version of the current engine or use the Centaur engine?
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The Delta-IV doesn't have the attachments for a 2nd engine
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Could a Centaur be mated with Delta IV?
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not without a rather complex redesign
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Thorny - 15/8/2007 10:23 AM
Could a Centaur be mated with Delta IV?
Sure, with lotsa time and lotsa, lotsa money.
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You'd have to redesign a large part of the upper stage. Which is why Boeing was looking at using larger engines such as the MB-45 or RL-60. Again, no money, either on their part or on the part of their Japenese partners.
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What about adding a second engine and having both engines feed into one nozzle?
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tnphysics - 15/8/2007 8:21 PM
What about adding a second engine and having both engines feed into one nozzle?
?????
Why?
why start a new engine development, especially a concept that hasn't been done?
It would be better to have engine that is twice as large.
I have notice a trend in your suggestions, tnphysics. They don't take into account the time or money needed to solve engineering issues. There always is an "easy" answer for problems, the hard part is doing it for less cost or time
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Quick question on the RS-68 upgrades. When done what will the Delta IV be renamed to? Delta IVb? Delta IV+? Delta V?
Also, how many 'origonal' RS-68's is ULA on the hook for? I assume they will not start building the upgraded RS-68's until they burn through the current inventory. At the current flight rate are we talking years? decades? It would make sense to me to switch it out on vehicles where a new RS-68 will save having to bump up to the next more expensive Delta IV model.
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It won't be renamed. there are that many RS-68"s made
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Jim - 16/8/2007 3:09 PM
It won't be renamed. there are that many RS-68"s made
Jim, Please clarify, where you trying to say "there are not that many" or there are billions and billions of RS-68 already made.
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I meant "aren't". Just had carpal tunnel surgery, and trying to minimize words
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Jim - 16/8/2007 3:38 PM
I meant "aren't". Just had carpal tunnel surgery, and trying to minimize words
Sorry to hear that, hope things heal up.
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If you are using windows then I think there are passable speech recognition tools built in, you feel like a dork talking to your computer at first but once trained it's faster than typing.
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nacnud, LTNS.
Glad to see you surfaced on *this* site :cool:
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edkyle99 - 14/8/2007 10:00 PM
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
Ed,
I can only refer you to the published Delta IV payload planner's guide for published payload weight as function of altitude and inclination angle (fig 2-35 for DIV-H). LEO means lots of different orbits and not all assumptions are consistent (for example, 100 nmi/ 28 deg differs from 220 nmi/ 51.6 deg, etc.). Then there are in-house conservatisms & performance margins that are book-kept in details (example, decremented engine performance, etc.), then there are payload numbers for commercial vs. "assurred access", etc.
A few years ago, there was a recognized need among the DLS executives that there's a payload gap between DIV-M (5,4) and DIV-H. That market is being dominated by Atlas V and the Ariane. It was recognized that the RL10B-2 can use a higher thrust to fill the need, but the new engine & stage upgrade effort was squashed by the economic reality.
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What would be the payload of the Delta IV(all varients) with regen engines? The Delta IV Heavy with cross-feed? Numbers to LEO, GTO, and GEO.
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There is a chart on the Boeing website with it
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Since this is a Delta IV information thread, here are two really interesting documents regarding the development and design philosophy behind the RS-68:
http://www.engineeringatboeing.com/dataresources/PropulsionForThe21stCentury-CostDrivenRS-68.ppt
http://www.engineeringatboeing.com/dataresources/PropulsionForThe21stCentury-RS-68.doc
Also IMO this should be pinned like the Atlas V and Centaur Q&As are. Great info here, keep it coming. :)
Nick
P.S. Jim, hope your carpal tunnel clears up.
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edkyle99 - 14/8/2007 10:00 PM
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
One more question I forgot to ask. Were you assuming the 2nd stage tanks are fully loaded?
Tanks are not fully loaded for DIV-H LEO missions.
Propforce - 16/8/2007 4:04 PM
Ed,
I can only refer you to the published Delta IV payload planner's guide for published payload weight as function of altitude and inclination angle (fig 2-35 for DIV-H). LEO means lots of different orbits and not all assumptions are consistent (for example, 100 nmi/ 28 deg differs from 220 nmi/ 51.6 deg, etc.). Then there are in-house conservatisms & performance margins that are book-kept in details (example, decremented engine performance, etc.), then there are payload numbers for commercial vs. "assurred access", etc.
This is another reason why I'd refer you to the payload planners' guide.
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Propforce - 17/8/2007 1:54 PM
edkyle99 - 14/8/2007 10:00 PM
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
One more question I forgot to ask. Were you assuming the 2nd stage tanks are fully loaded?
Tanks are not fully loaded for DIV-H LEO missions.
Huh. Structural reasons, trajectory optimization, or both?
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Propforce - 17/8/2007 3:54 PM
edkyle99 - 14/8/2007 10:00 PM
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
One more question I forgot to ask. Were you assuming the 2nd stage tanks are fully loaded?
Tanks are not fully loaded for DIV-H LEO missions.
My spreadsheet assumption was that the upper stage carried about 5 tonnes less propellant for a LEO mission. I based that offloaded amount on the burn times given in the user's guide flight profiles. The 9,660-ish m/s number mentioned above, however, was merely a backtrack from the Delta IV-Heavy orbital data back down to a 185 km parking orbit.
Propforce - 16/8/2007 4:04 PM
Ed,
I can only refer you to the published Delta IV payload planner's guide for published payload weight as function of altitude and inclination angle (fig 2-35 for DIV-H). LEO means lots of different orbits and not all assumptions are consistent (for example, 100 nmi/ 28 deg differs from 220 nmi/ 51.6 deg, etc.). Then there are in-house conservatisms & performance margins that are book-kept in details (example, decremented engine performance, etc.), then there are payload numbers for commercial vs. "assurred access", etc.
This is another reason why I'd refer you to the payload planners' guide.
Yep. I'm sure that if I spent a few more hours contrasting every publicly available bit of information I would be able to narrow down to a better understanding of the Delta IV gravity losses. I suspect that I'm already in the ballpark, but I've been meaning to dig at this a bit more, perhaps this weekend.
- Ed Kyle
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yinzer - 17/8/2007 2:11 PM
Propforce - 17/8/2007 1:54 PM
edkyle99 - 14/8/2007 10:00 PM
9,700 to 9,800 meters per second means 1,900 to 2,000 meters per second worth of drag, gravity, and steering losses. That is high compared to Shuttle's 1,700 meters per second, Ariane's 1,300 meters per second, or the old Saturn IB's 1,400 meters per second, or Atlas V's 1450-ish meters per second. If true, then that apparently is the price paid to fly a pure hydrogen machine with its bigger tanks and lower-than desirable upper stage thrust.
- Ed Kyle
One more question I forgot to ask. Were you assuming the 2nd stage tanks are fully loaded?
Tanks are not fully loaded for DIV-H LEO missions.
Huh. Structural reasons, trajectory optimization, or both?
No need to carry extra. Has enough propellant to get there, plus reserves.
Doesn't Centaur offload for LEO missions?
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edkyle99 - 17/8/2007 2:33 PM
My spreadsheet assumption was that the upper stage carried about 5 tonnes less propellant for a LEO mission. I based that offloaded amount on the burn times given in the user's guide flight profiles. The 9,660-ish m/s number mentioned above, however, was merely a backtrack from the Delta IV-Heavy orbital data back down to a 185 km parking orbit.
Some missions offload more than 5 tons. Also what payload weight?, ETR vs. WTR? and which LEO (51.6 vs. 28 deg inclination) orbit?
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Thank you for bringing that to my attention. My calculations assumed no offload, no fairing, no gravity losses, no nose cone on the strap-ons, and a core that burned only 60% of its propellant while the strap-ons are firing.
When does the fairing separate? What is the mass of the fairing and of the strap-on CBC's nose cones?
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I've looked again at the Delta IV Heavy mission example provided in the 2002 Planners Guide. This was a 20 tonne payload launched to a 780 km x 96 deg circular orbit. My simplified rocket equation spreadsheet shows the Delta IV Heavy producing roughly 10,080 meters per second ideal delta-v for such a launch. This ascent profile includes a brief, second upper stage burn at first apogee to circularize the orbit.
My best guess is that this example has about 1980 m/s of losses, including drag and gravity. An equivalent Cape Canaveral launch would gain an extra 430 m/s from the earth's rotation and so would only have roughly 1550 m/s losses, which is about 100 m/s more than STS.
- Ed Kyle
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Thanks Ed,
I'll take a look and see if I can provide helpful comments.
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Propforce - 14/8/2007 3:04 PM
tnphysics - 13/8/2007 4:08 PM
What is the actual delta-V split and payload for the Delta IV Heavy? Using astronautix.com numbers, the total delta-V exceeds 10.5 km/s.Also, what would be the payload for a LV consisting of a single CBC and a J-2/J-2X upper stage?I felt that we needed a Q&A thread for the Delta IV.
Since you're asking a hypothetical question (current D-IV uses RL10B-2 and not J-2/J-2X), you can get a comparable number by taking the existing payload, delta-v, and B-2's engine Isp, then swap out with the J-2/ J-2X Isp to see what additional payload gain (if any) can be achieved. You'll need to run this throughout the trajectory since the CBC will take a hit on delta-vee because of additional payload weight as well. I would also allow some gravity delta-v losses on the vehicle.
You'll then need to subtract the additional J-2/J-2X engine weight from that additional payload weight gain.
Now the fun part is, should you find yourself run out of 2nd stage propellant too early, it's time to re-size the 2nd stage tanks to allow to carry more propellant. But that would be difficult since you only allow a single CBC and soon we'll run into the thrust limitation of the RS-68! :laugh:
But don't be discouraged. Atlas V can't do any better as their initial liftoff thrust-to-weight is already dangerously low. :bleh:
I was going to make stage 2 bigger and shorten stage 1.
Or you could add small SRBs and lengthen stage 2.
Or you could use 3 CBCs.
Or you could use an RS-800.
They have considered a WBC on Atlas V. Try it on Delta IV.
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Why do the Delta IV upgrade options give such a dramatic boost in payload?
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Why did the Delta IV go with a cryogenic first stage?
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tnphysics - 28/8/2007 8:33 PM
Why do the Delta IV upgrade options give such a dramatic boost in payload?
The solids on the Heavy give a big boost because they reduce gravity losses in flight. Basically as the solids burn, they help to propel the vehicle higher so that for a given burn time, the vehicle covers more altitude. Thus, less of the main engines' thrust is wasted fighting gravity on the way up and can be used to increase delta-v, which is what really helps get heavy payloads into orbit. A new upper stage would have the same effect.
It probably also removes any concerns about thrust-to-weight ratio. Six GEMs would give one hell of a kick!
As for why they went with a cryogenic first stage, I don't know. Increased ISp maybe? The RS-68, even with its simplified design, has a better ISp than the much more complex RD-180 (408 vs 338 sec, respectively). But then you do give up thrust, so I don't really know.
My understanding of this is probably wrong, so the more experienced/qualified people here can probably answer your question more completely and more correctly than I ever could. ;)
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What kind of payload could the Delta IV Heavy put in orbit, if
a) the upper stage tanks were fully loaded?
b) the upper stage was deleted?
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Why can't the Delta IV put as much in LEO as the Atlas V?
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tnphysics - 3/9/2007 10:06 AM
Why can't the Delta IV put as much in LEO as the Atlas V?
Because it can't.
Obvious answers, tanks were sized wrong and dry weights are higher than predicted
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Nick L. - 3/9/2007 1:23 AM
As for why they went with a cryogenic first stage, I don't know. Increased ISp maybe? The RS-68, even with its simplified design, has a better ISp than the much more complex RD-180 (408 vs 338 sec, respectively). But then you do give up thrust, so I don't really know.
IIRC, When D4 was initially designed, McDonnell Douglas had conducted studies showing that using the same propellant combos for both stages would reduce overall costs.
In hindsight, most people will look at D4 and say that LOX-LH2 was a pretty bad choice for the lower stage. A better choice might have been a two-stage rocket with LOX-Kerosene in both stages. That seems to be SpaceX's reasoning.
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How much larger should the first stage tanks have been?
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How large would the upper stage of an LEO-optimized Delta IV be?
Assume that the CBC is unchanged (or offloaded, if necessary, although adding SRBs would be preferable)
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CFE - 3/9/2007 10:26 AM
Nick L. - 3/9/2007 1:23 AM
As for why they went with a cryogenic first stage, I don't know. Increased ISp maybe? The RS-68, even with its simplified design, has a better ISp than the much more complex RD-180 (408 vs 338 sec, respectively). But then you do give up thrust, so I don't really know.
IIRC, When D4 was initially designed, McDonnell Douglas had conducted studies showing that using the same propellant combos for both stages would reduce overall costs.
In hindsight, most people will look at D4 and say that LOX-LH2 was a pretty bad choice for the lower stage. A better choice might have been a two-stage rocket with LOX-Kerosene in both stages. That seems to be SpaceX's reasoning.
1) Choosing a LO2/LH2 cryo first stage has NOTHING to do with sizing engine THRUST. What I mean is, one could have sized the RS-68 to 800K lbf if McDD chose to. The fact was that RS-68 thrust level was picked before all the weight growth were frozen.
2) One could argue all day till the cows come home between a LOX/LH2 vs. a LOX/HC first stage. A LOX/HC first stage would've required a much higher thrust engine because of lower Isp of LOX/HC engine and heavier HC propellant weight ( a combination of low Isp engine requires MORE propellant + higher density of HC fuel).
3) DIV was designed for high energy, high performance missions to GTO. It is NOT optimized for LEO missions. The market study said that the money is on the GTO comm sat missions.
4) Until I see an Atlas V Heavy and how much they'd charge for a heavy launch. DIV Heavy is still the only game in town for U.S. government missions.
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How much would developing the RS-800 cost? Probably no more than $470 million, or 2 Heavy launches.
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tnphysics - 11/9/2007 11:19 PM
How much would developing the RS-800 cost? Probably no more than $470 million, or 2 Heavy launches.
more and who is going to pay. But anyways, you just only got an engine so what are you going to do with it. integrating it into a vehicle costs big money
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The RS-68 cost only $470 million to develop. That is where I got my cost number.
The AUS could have helped tremendously because a Medium+ (5,4) with an AUS could have replaced a Heavy without one.
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tnphysics - 16/9/2007 8:27 AM
The RS-68 cost only $470 million to develop. That is where I got my cost number.
The AUS could have helped tremendously because a Medium+ (5,4) with an AUS could have replaced a Heavy without one.
No, the AUS was not big enough to make a 5,4 equivalent to a heavy
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tnphysics - 16/9/2007 8:27 AM
The RS-68 cost only $470 million to develop. That is where I got my cost number.
The AUS could have helped tremendously because a Medium+ (5,4) with an AUS could have replaced a Heavy without one.
How are you coming up with these assertions? IIRC there have been no published data regarding the planned performance of the AUS. A Medium+(5,4) with the AUS could fill the gap between the standard M+(5,4) and the Heavy, but it could never replace it. There are also probably cheaper ways of boosting the performance other than developing an all new upper stage.
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Why is the Delta IV Medium+ (5,2) unable to lift as much to LEO as the Delta IV Medium?
And why is there no Delta IV Medium+ (4,4)?
Or even (5,6) or (5,9)?
And why does the Delta IV cost more than the Delta II?
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tnphysics - 20/9/2007 7:56 PM
Why is the Delta IV Medium+ (5,2) unable to lift as much to LEO as the Delta IV Medium?
And why is there no Delta IV Medium+ (4,4)?
Or even (5,6) or (5,9)?
And why does the Delta IV cost more than the Delta II?
1) Because of the extra weight of the larger second stage and fairing, it creates more gravity loss on LEO missions. On GEO missions the bigger upper stage makes up for the added weight.
2) It wasn't worth the cost of certificating it considering the lack of payloads in the weight range.
3) It would require extra cost to add more GEM attach points.
4) Because it's bigger and has more payload capability.
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tnphysics - 20/9/2007 6:56 PM
Why is the Delta IV Medium+ (5,2) unable to lift as much to LEO as the Delta IV Medium?
And why is there no Delta IV Medium+ (4,4)?
Or even (5,6) or (5,9)?
And why does the Delta IV cost more than the Delta II?
1. It has the same thrust but is lifting a wider, heavier payload fairing. Atmospheric drag loss and mass properties probably account for most of the 1,600 kg discrepancy to GTO.
2. Probably because most of the heavy payloads that need 4 GEMS are also wide enough to need the 5-meter fairing.
3. That's the purpose of the Delta IV-H
4. The smallest Delta IV lifts twice the payload of the Delta II. It's just a more capable vehicle and that translates into cost. The Delta IV is also newer and has very low flight rate, which does nothing good for costs.
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Why is it more expensive to build from a supply side point of view?
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tnphysics - 20/9/2007 8:17 PM
Why is it more expensive to build from a supply side point of view?
supply side ??????
It is bigger, more material therefore more costs
It also has to buy down development costs
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tnphysics - 20/9/2007 7:56 Pm
Or even (5,6) or (5,9)?
The pad can't support the installation of anymore than 4
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Jim - 20/9/2007 8:29 PM
tnphysics - 20/9/2007 8:17 PM
Why is it more expensive to build from a supply side point of view?
supply side ??????
It is bigger, more material therefore more costs
It also has to buy down development costs
What about all those SRBs on the Delta II?
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So what about them? They work.
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Why they don't drive up the costs.
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Drive up what costs? Delta has been using 9 SRM for more than 35 years
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Because Delta II was designed with the GEMs. It needs them to be able to fly at all. Delta II is also only a derivative of the early Delta (1000/2000/3000 series), so it was cheaper to develop, so development costs are already paid for, which makes the cost lower. Delta IV is a clean-sheet design and the development costs haven't been paid off yet which drives up the per launch cost.
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As for the question about more solids on a Medium+… Delta has looked at configurations other than the 4 currently offered. That includes the neccessary pad mods. Its less developed than AUS or RS-68A, but Delta has invested some time and effort to develop conceptual designs that have up to 8 solids on a “single stick”… But these have not been put on the manifest because of a lack of need. Right now the customer needs are at the high end of the vehicle capability, to upgrade the Heavy as indicated on the COTS RFI response. Quite simply there isn’t a costumer need for a better Medium. But there’s enough work to feel confident that everything from a 4/4 to a 5/8 could be quickly brought on-line if a costumer indicates a need.
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At least the Delta-4M+(4,4) could be implemented rather quickly, as no pad modification would be necessary.
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TrueGrit - 21/9/2007 3:50 AM
Quite simply there isn’t a costumer need for a better Medium. But there’s enough work to feel confident that everything from a 4/4 to a 5/8 could be quickly brought on-line if a costumer indicates a need.
The real issue is that the need can be satisified by an Atlas 4XX or 5XX more cheaply and thus that is why we won't see any 4/4 or 5/8
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What would be the 5/8's payload?
How much would it cost to stretch the CBC's tanks and increase the payload of such a configuration?
TrueGrit, are you stating that ULA could provide a 5/8 to a customer for less than they would charge for a Heavy?
This includes pad mods, to whatever extent ULA would decide to charge its first 5/8 customer for them as opposed to amortizing them over many 5/8 missions.
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Could the CBC be modified for parachute recovery and reuse?
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tnphysics - 21/9/2007 7:41 AM
What would be the 5/8's payload?
How much would it cost to stretch the CBC's tanks and increase the payload of such a configuration?
TrueGrit, are you stating that ULA could provide a 5/8 to a customer for less than they would charge for a Heavy?
This includes pad mods, to whatever extent ULA would decide to charge its first 5/8 customer for them as opposed to amortizing them over many 5/8 missions.
You figure it out. You would get just as good of answer
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tnphysics - 21/9/2007 7:53 AM
Could the CBC be modified for parachute recovery and reuse?
It just takes time, money and some performance. which all would be a waste
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Shouldn't a Delta IV Heavy with a multi-engine upper stage be able to put about 35 metric tons in LEO?
This assumes that the additional engines are staged after they are no longer needed.
And couldn't the three common booster cores be able to put about 19 metric tons in LEO?
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tnphysics - 21/9/2007 8:58 PM
Shouldn't a Delta IV Heavy with a multi-engine upper stage be able to put about 35 metric tons in LEO?
Define "multi-engine upper stage", please.
This assumes that the additional engines are staged after they are no longer needed.
Say what??
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Multi-engine upper stage means more than one upper stage engine.
I was going to jettison each engine once its thrust was less important than its mass.
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tnphysics - 21/9/2007 10:28 PM
Multi-engine upper stage means more than one upper stage engine.
I was going to jettison each engine once its thrust was less important than its mass.
Yeah, we know. The question is, what kind of multi-engine stage? 2 engines? 4? How big? How much fuel? How long of a burn time? How much would it cost?
1.5 staging (like you are proposing) is very complex, and to be able to do such a thing successfully would probably not be cost effective considering the lack of payloads in the weight range.
Developing a new upper stage would be too expensive and would not be worth it; there are not enough payloads out there to absorb the cost of development. This is why the Advanced Upper Stage was canceled in the first place. There are cheaper ways to improve the Delta IV's performance. RS-68 upgrades and GEMs are some of these.
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tnphysics - 21/9/2007 9:28 PM
Multi-engine upper stage means more than one upper stage engine.
I was going to jettison each engine once its thrust was less important than its mass.
Both D-IV and Centaur have pretty efficient upper stage engines now (and with Centaur you can choose either one or two engines based on payload mass and type of orbit injection requirements).
You do realize how much extra hardware would be required to do an Atlas-1-type jettison of upper stage engines, right?
(And that they -upper stage engine(s)- have to be balanced both thrusting and stopping and jettisoning so the CoP and CoG and overall thrust vector are balanced and designed/programmed to place the payload in its contracted orbit.)
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You need more than 2 to fight the gravity losses.
4 RL-10 engines is more like it.
This will improve payload even if the switch must be made to less efficient RL-10-A-4 engines (instead of RL-10-B-2).
I did not realize how much extra hardware was required for jettisoning upper stage engines.
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tnphysics - 21/9/2007 10:28 PM
Multi-engine upper stage means more than one upper stage engine.
I was going to jettison each engine once its thrust was less important than its mass.
why??
And drop an engine on some people
This is the real world Theorical solutions are not always practical nor cost effective. Delta IV and Atlas V were designed to meet specific requirements, which they do. Any other "requirements" would have to paid for by the org that needs the change.
Right now, the only upgrade to the EELV's, is the Delta IV Heavy enhancements. Any others don't exist until a paying user is found
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Why would you need to strech the CBC to incorporate more solids? Any studies certianlly didn't involved development a new GEM (big $$), but utilize the existing 40, 46, or 60 designs (first two from Delta II and III). Any additional GEMs would be placed radially around the CBC. But I think it's safe to say that the M+4/4 and M+5/8 along with WBC and Atlas Heavy are all a question marks at the least with ULA around. That doesn't mean the advanced product development teams are sitting on their hands. But the focus has changed from "beating the other guy" to "working with the other guy".
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So Delta spends countless hours keeping the CBC internal enviornment "dry" and you want it to be dropped in the ocean to soak in sea water for hours. No one has conviced me you can keep the propellant systems from getting contaminated in that process. And once contaiminated how do you clean a 15000+ cuft tank with isogrid cutouts without disassembling the entire structure. The disassembly, cleaning, and reassembly process would be more expensive than simply making another tank. And that doesn't begin to talk about the 1000+ lbm weight hit for parachutes, costs to incorporate it, or costs to support the recovery boat/team.
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Isn't the tank cleaned after welding too? Are the segments cleaned (washed) after machining before welding? It seems to be a small part of the manufacturing process.
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All cleaning is done while the tank is still in subassembly... There is no forward or aft covers, or small/electrical equipement installed. As such you have great access through the top and bottom and don't have to worry about small/electircal parts. With the vehicle all assembled removing the covers means removing the feedlines, electrical equipment, and small lines/valves. Nothing short of a compelte disassembly of the tank. And this doesn't being to address the composite structure... All of which degrades when exposed to a high moistore enviornment. My guess would be that the composite structure structure would be a loss. So now your talking about disassembling the stage into subsections, diassembling the tank for cleaning, and completely replacing the Centerbody/Instage... Building the tank and engine section structure is actually quite cheep. Particuarly now that the tooling is built and the processes arer developed.
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Okay, you've presuaded me. It's not cost effective.
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TrueGrit - 22/9/2007 10:34 PM
All cleaning is done while the tank is still in subassembly... There is no forward or aft covers, or small/electrical equipement installed. As such you have great access through the top and bottom and don't have to worry about small/electircal parts. With the vehicle all assembled removing the covers means removing the feedlines, electrical equipment, and small lines/valves. Nothing short of a compelte disassembly of the tank. And this doesn't being to address the composite structure... All of which degrades when exposed to a high moistore enviornment. My guess would be that the composite structure structure would be a loss. So now your talking about disassembling the stage into subsections, diassembling the tank for cleaning, and completely replacing the Centerbody/Instage... Building the tank and engine section structure is actually quite cheep. Particuarly now that the tooling is built and the processes arer developed.
Thanks, this is good info! I for one didn't have this good insight into the whole process. That makes the arguments powerful and justified - reuse would be costly.
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What about scrapping the tankage and focusing on reusing the engines?
Or making a flyback CBC?
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tnphysics - 22/9/2007 10:38 PM
What about scrapping the tankage and focusing on reusing the engines?
Or making a flyback CBC?
Not worth the effort
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Would using RS-800 engines require pad mods?
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If the need or financial justification arose, could the RS-68 be modified for use as an upperstage engine? Stock RS-68 ISP in a vacum (410) can't compete w/ RL-10 (as high as 462?) but I wonder if perhaps the Ares upgrades to the RS-68 would open up this possibility. Or would it be better to complete the AUS?
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Firehawk153 - 25/9/2007 4:38 PM
If the need or financial justification arose, could the RS-68 be modified for use as an upperstage engine? Stock RS-68 ISP in a vacum (410) can't compete w/ RL-10 (as high as 462?) but I wonder if perhaps the Ares upgrades to the RS-68 would open up this possibility. Or would it be better to complete the AUS?
RS-68 as an upperstage engine is a nogo. See other posts
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Firehawk153 - 25/9/2007 3:38 PM
If the need or financial justification arose, could the RS-68 be modified for use as an upperstage engine? Stock RS-68 ISP in a vacum (410) can't compete w/ RL-10 (as high as 462?) but I wonder if perhaps the Ares upgrades to the RS-68 would open up this possibility. Or would it be better to complete the AUS?
As mentioned by others here in different threads, the RS-68 design just takes too much ground support equipment to prep for ignition and through the ignition cycle. To try to add all it needs into an upper stage would add so much mass it would overwhelm any advantages the engine could offer.
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Jetteson the excess startup hardware after RS-68 ignition.
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Ahhhh, okay, it makes sense then why it wouldn't work. I forget that these vehicles require extensive ground support.
tnphysics - 25/9/2007 5:48 PM
Jetteson the excess startup hardware after RS-68 ignition.
Wouldn't this incur a mass penalty to the overall payload capacity to haul the excess equipment thru 1st stage and 2nd stage ignition?
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You're kidding, right?
I'm not sure it couldn't be done, but do you realize the amount of extra engineering, testing, and costs involved to try to accomplish that?
Just some advice, but you really need to keep in mind that everything requires money - $$$$$. Designs cost $$$$$. Engineering costs $$$$$. Development and testing costs $$$$$. Construction costs $$$$$. Money ($$$$$) rules over everything, *everything* regardless of how neat, nifty, cool, or even practical a new idea for a booster, propulsion system, mission vehicle, or mission plan design might be.
The primary emphasis is to design and build the hardware needed to accomplish the ordered mission for as little $$$$$ as possible. The $$$$$$ is the main priority for the designs and hardware, not the hardware or mission itself.
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tnphysics - 25/9/2007 6:48 PM
Jetteson the excess startup hardware after RS-68 ignition.
The LV might not even be able to get off the ground! The hardware is MUCH bigger than you think. I would bet that carrying all of the starting equipment would reduce payload to nil, if you could even get it to fit into the LV interstage. From an old post of Jim's, I believe the helium tank for starting the engine is the size of a tractor-trailer. Engineering the ability to carry all of this equipment into the rocket, and then jettisoning all of it, would cost very much - almost certainly too much to be worth it.
The RS-68 is also not capable of air-starting, due to the differences in air pressure and density, and re-engineering it to be able to do that basically kills the cost benefits of the cheaper RS-68 design.
All in all, too much money for too little reward.
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What about using Atlas V CCBs as LRBs on a Heavy?
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First fund the heavy testing and development.
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"What about using Atlas V CCBs as LRBs on a Heavy?"
Delta Heavy with the RS-68A exceeds Atlas Heavy capability... Atlas and Delta attachment points points don't match up... Altas and Delta avionics systems are not compatible... LC37/SLC6 don't have the proper GSE (umbicals to electrical) to support Atlas... LC37/SLC6 don't have any RP1 loading capability... LC37/SLC6 LOx storage capability sized to Delta and Atlas requires more... And the problems go on and on
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RS-68 as an upperstage engine? First question is what upper-stage would require ~800klb of thrust? Second what upperstage would want the huge mass penalty (RS-68 is quite heavy)? RS-68 would drain the huge Ares upperstage dry in less than 3 minutes... And pull almost 2 g's at the time.
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I am amused with all these seemly naive questions. But I think the problem is that the folks who are not in this industry doesn't know how a launch vehicle and engine is designed. As such, in their minds, they're trying to fit an existing engine into a "new" vehicle for whatever application they have in mind.
In reality, those of us who are in this business will design or modify an existing engine/ vehicle to meet whatever the new mission(s) calls for, assuming there's a "need" for this new mission and whoever the customer is will pay for it. Sometime we change the design so much just about the only "heritage" part is the nameplate and everything else is different. We, D-IV, do it. The Atlas folks do it. The SpaceX folks do it. Everyone who's in the LV business does it.
We will start with these "new" missions, develop representative flight profiles, call them "design reference missioins" (DRM) and start to evaluate existing vehicle/ engine to see if we can modify or need to build new ones, and how long and how much they will cost. All these study results are fed back to management/ customers in order to evaluate if the mission is worth the money that will be required.
Like Dr. Antonio Elias of OSC said, any fools can design an expensive new launch vehicle, but it takes experiences and brains to design a "cost effective" vehicle. The ingenius part is to design a vehicle that is inexpensive while meeting launch accuracy and relaibility. In the case of "mature" infrastructure such as the Atlas and the Delta, any decision on design changes on the vehicle must take into account its impact on existing launch pad, ground system, factory, suppy chain, and logistic. Sometime it's harder to make changes to these infrastructure and it's easier if one is starting from scratch such as the SpaceX and OSC.
In the case of Delta IV upgrades, yes we have studied putting a brand new engine (MB-60) on a new "stretch" versioin of upper stage (AUS), or putting 2~3 smaller new engines (MB-35), or existing engines (RL10B-2), more solids on the CBC, more liquid strap-on (upto 6) around the core CBC, build a larger diameter of CBC (over 8 meters) and put multiple RS-68s on each CBC, up-rating RS-68 to a much higher thrust than the existing upgrade, etc. All these were done for various *missions* with various customers, each comes with impact on cost, schedule and varying degree of risks. But don't ask me about these results, they are proprietary first of all, then the issue of ITAR, FOUO, etc... yikes!
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Where can I find information about the Delta IV Heavy upgrades beyond what is on astonautix.com etc.?
I didn't know about the widend CBC.
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tnphysics - 26/9/2007 4:54 PM
Where can I find information about the Delta IV Heavy upgrades beyond what is on astonautix.com etc.?
I didn't know about the widend CBC.
Why do you need to know?
:cool:
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Just curious.
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tnphysics - 26/9/2007 8:13 PM
Just curious.
pre ULA they where clearly posted on the boeing site. I am sure one of the forums digital pack rats has a copy somewhere. If not google is your friend... I am sure they are findable... I know the graphic has been posted several times on this forum.
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Yeah, it was named "delta evolution chart" or something like that.
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Okay. Why didn't the Delta IV developers go with a wider twin engine upper stage?
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DeltaIV upperstage is an evolution of that designed for Delta III... The single RL10 was the best choice to meet the reference missions (GSO commercial) the Delta III was designed against. Another engine would increase the cost of the rocket and hurt the reliability, for little gain. And there is no in production option other than RL10... Is that simple enough. Low Earth might be the target for all you space hobbiest, but those of us in the buissness know the money is to be made in GSO.
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tnphysics - 27/9/2007 3:30 PM
Okay. Why didn't the Delta IV developers go with a wider twin engine upper stage?
They could. But not enough sales to justify the "upgrade".
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I know... but the Delta IV and Delta III upper stages are too small even for GTO missions. Just look at the Atlas V and the WBC.
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tnphysics - 27/9/2007 7:06 PM
I know... but the Delta IV and Delta III upper stages are too small even for GTO missions. Just look at the Atlas V and the WBC.
They are correctly sized for GTO
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Then why is Atlas going to the WBC?
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tnphysics - 27/9/2007 7:54 PM
Then why is Atlas going to the WBC?
Atlas is not going to the WBC. That is just a study. WBC is a way of increasing performance without upsetting too many ground interfaces.
This is system engineering. Not everything is optimized for flight performance, cost is always the driver
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Too small for GSO? Guess the most capable rocket in the US inventory isn't enough :sigh:
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Jim - 27/9/2007 10:11 PM
tnphysics - 27/9/2007 7:54 PM
Then why is Atlas going to the WBC?
Atlas is not going to the WBC. That is just a study. WBC is a way of increasing performance without upsetting too many ground interfaces.
This is system engineering. Not everything is optimized for flight performance, cost is always the driver
I was mixed up. Sorry.
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Why won't the AUS be built?
It will if the WBC flies on Atlas and SpaceX fails.
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tnphysics - 30/9/2007 9:27 PM
Why won't the AUS be built?
It will if the WBC flies on Atlas and SpaceX fails.
There are requirements for AUS or WBC
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tnphysics - 30/9/2007 9:27 PM
Why won't the AUS be built?
It will if the WBC flies on Atlas and SpaceX fails.
Again, not enough payloads in the weight range to spread the costs. It will cost a lot to develop the AUS. You need to finish developing the MB-60 engine, new tankage, possibly new interstages, fairings, guidance system changes, a lot of work needs to be done. Delta IV Heavy already outperforms every other rocket out there right now (Ariane 5 ECA is the next most powerful), and they can't even find enough payloads for it! There are just not enough payloads that heavy. And Boeing/ULA can't compete with Ariane on costs so no commercial dice either.
If WBC flies (which is HIGHLY unlikely, see above), there will be no need to develop the AUS; they will have already saturated the market, there are that few payloads out there that take advantage of the extra capability. SpaceX doesn't have an LV, currently available or planned, that can even match Delta IV Heavy (F9-Heavy has less performance to LEO and GTO than D-IVH).
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Isn't ULA now talking about a combined AUS and WBC for both Delta IV and Atlas V called ACES? At least they have it in there power point responding to NASA's ISS Cargo RFI ( http://www.ulalaunch.com/docs/publications/ULA/ISS_Cargo_RFI_Final_09062007.pdf ).
Interesting how closely matched the Delta IV Heavy and Atlas V Heavy performance is based on ULA's numbers in the just mentioned .pdf. Only a 200 kg difference.
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Jim - 30/9/2007 9:33 PM
tnphysics - 30/9/2007 9:27 PM
Why won't the AUS be built?
It will if the WBC flies on Atlas and SpaceX fails.
There are requirements for AUS or WBC
I meant no requirements
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I probably know the answer to this but I'm going to ask it anyway to be sure.
D-IVH roll control... is it through gas generator exhaust like all the other Delta IVs or is through outboard CBC engine gimballing?
How would each compare wrt to roll authority?
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hyper_snyper - 1/10/2007 10:31 PM
I probably know the answer to this but I'm going to ask it anyway to be sure.
D-IVH roll control... is it through gas generator exhaust like all the other Delta IVs or is through outboard CBC engine gimballing?
How would each compare wrt to roll authority?
The Delta IV PPG says that Heavy roll control is by differential gimbaling of the outboard engines. I suspect that the central CBC still has the turbopump exhaust vectoring, for roll control after the strap-on CBCs are jettisoned.
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I hope that the AUS is eventually built.
It would allow a 3-launch lunar architecture.
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It won't be
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How much would it increase the LEO payload?
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Figure it yourself from the ULA product page...
http://www.ulalaunch.com/images/product_sheet/Delta_Product_Sheet_FINAL.pdf
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Why dosen't the Delta IV US use the RL-10C, which has more thrust?
This would help the heavy dramatically.
It does have less Isp.
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tnphysics - 17/10/2007 6:49 PM
Why dosen't the Delta IV US use the RL-10C, which has more thrust?
This would help the heavy dramatically.
It does have less Isp.
A mentioned in the Atlas Q & A thread, there is no such engine.
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Could the MB-XX, which was tested, be incorporated into the Delta IV US?
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Maybe - there would probably be a lot of integration work required (in addition to actually qualifying the engine). But, as with the Atlas thread and the RL10C, why? It would have to support some HUGE payloads, and those do not exist. There are a lot of "what if" scenarios out there, but unless any of them come up with money, no one will pursue them
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tnphysics - 17/10/2007 7:15 PM
Could the MB-XX, which was tested, be incorporated into the Delta IV US?
MB-XX is still in development and a long way from flight status, it seems.
If development is finished, and someone wants to pay for it, yes. That was Boeing's long term thinking before ULA and the reason for the RL60 development program for Centaur, which has definitely been shelved after testing a prototype and starting to assemble a production prototype for testing.
Right now, the most advanced new small cryogenic engine that's anywhere close to flight status is Vinci:
http://cs.space.eads.net/sp/LauncherPropulsion/LaunchVehiclePropulsion.html#Vinci
For a detailed description of Vinci and lots of internal details of an expander cycle
cryogenic engine, see this largish PDF file:
http://cs.space.eads.net/sp/PDF/vinci.pdf
Any changes in design are going to be driven by need for additional performance, and the necessary funding being available to make those changes. It'd be nice if more advanced engines were available off the shelf but even then the upper stages might need significant changes just to swap in a single engine, let alone add multiple engines.
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MB-XX family was supposed to be a “drop-in” replacement for the RL-10, and along with the AUS represented the next step in Delta evolution. It was all part of Boeing’s strategic planning… Where the launch vehicle and engine would be all under one house, just like the CBC. They were to team with Mitsubishi who would share the risks and costs to develop the MB-XX. And then strengthen the existing similarities with the H-II and Delta upper stages. This interestingly turns out to be way the 787 is being done, with major suppliers who share the risks and costs in exchange for more power and profit potential. In the end the MB-XX would filter down to replacing the RL-10 on all Delta upper stages… Of course all work in the direction and development money dried up with the commercial satellite market collapse.
MB-XX testing reached demonstrator level, having tested “battleship” subassemblies (pumps, injector, and chamber). But I don’t believe ever tested an integrated engine, and is therefore years and hundreds of millions dollars away from production. Of course any discussions of the MB-XX are moot now that Pratt owns Rocketdyne… No corporate will to replace the RL-10. Although MB-XX isn’t completely dead as the lessons learned are living on in the J-2X and RS-68 update programs. And who knows… If the RL10 is upgraded as part of the lunar lander program may end up with significant MB-XX like design changes.
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Damon Hill - 18/10/2007 4:06 AM
For a detailed description of Vinci and lots of internal details of an expander cycle
cryogenic engine, see this largish PDF file:
http://cs.space.eads.net/sp/PDF/vinci.pdf
Any changes in design are going to be driven by need for additional performance, and the necessary funding being available to make those changes. It'd be nice if more advanced engines were available off the shelf but even then the upper stages might need significant changes just to swap in a single engine, let alone add multiple engines.
Quite high pressure, 60 bars on an expander. RL-10A-4 is under 40 bars, if astronautix is to be believed... Guess the Germans and French want to one-up. :) Also the stowable ceramic nozzle... I hope all goes well with this ambitious design, I wouldn't be that surprised if there was a failure during the early flights.
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Pratt & Whitney Rocketdyne's web site no longer catalogs RL60, and of course MB-XX is not there either. Instead, the company "offers" RD-0146, which is roughly speaking a Russian RL10 equivalent.
Why PWR would offer to "shorthaul" itself this way is a bit mystifying. The only explanation that makes sense to me is that PWR would rather not have to build RL10 at all.
- Ed Kyle
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Before launch, LH2 boiloff on the shuttle is piped away from the pad to be burned safely.
How does Delta IV differ in terms of what do to with LH2 boiloff? For that matter how does any vehicle with any sort of hydrogen stage deal with boiloff before launch?
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It's a range requirement to capture and burn-off, or dilute to 1/4-th flamability limit all hydrogen (gaseous or liquid). GO2 is allowed to free vent, and liquid oxygen is dumped into a dedicated pond (depression). Interestingly the RL-10 powered DC-X free vented hydrogen... But that was White Sands, and each range site has there own rules abou these things.
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Besides the DC-X how many LH fueled vehicles have flown out of White Sands?
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kevin-rf - 19/10/2007 2:25 AM
Besides the DC-X how many LH fueled vehicles have flown out of White Sands?
no other vehicles (counting the DC-XA as DC-X)
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Wouldn't the Delta IV Heavy's payload to the ISS be about 32 metric tons with both an RS-68 upgrade and propellant crossfeed?
The RS-68 upgrade alone gives about 27 metric tons to the ISS.
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tnphysics - 22/10/2007 6:55 AM
Wouldn't the Delta IV Heavy's payload to the ISS be about 32 metric tons with both an RS-68 upgrade and propellant crossfeed?
The RS-68 upgrade alone gives about 27 metric tons to the ISS.
Dan Collins, Boeing's VP of Delta Programs back when Delta still was a Boeing product, presented a paper on Delta IV growth options during the 2004 Space Congress. Here is a link.
http://www.spacecongress.org/2004/Panel-4/2Collins.pdf
Lockheed Martin and ATK also provided presentations. http://www.spacecongress.org/sessions.htm
Nothing there about propellant cross feed, which doesn't sound to me like something that would provide a large benefit at any rate.
According to the paper, Delta IV can get to 30 tonnes with an RS-68 upgrade combined with a new upper stage. Atlas V would also need a reconfigured upper stage to make 30 tonnes.
- Ed Kyle
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There is a delta growth path presentation from Boeing somewhere that has the multiple options with both benefit and risk listed. Cross feed is one of them... I think it was from 2004 or 2005 and related to the VSE somehow. It had options to the future for 8 meter cores I think...
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http://www.ninfinger.org/~sven/models/vault2004/delta_iv_heavy_lift.jpg
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meiza - 22/10/2007 10:58 AM
There is a delta growth path presentation from Boeing somewhere that has the multiple options with both benefit and risk listed. Cross feed is one of them... I think it was from 2004 or 2005 and related to the VSE somehow. It had options to the future for 8 meter cores I think...
I stand corrected. There was one mention of propellant crossfeed for one of the projected growth options in that presentation, but it was combined with propellant densification, which probably accounted for a majority of the payload growth.
- Ed Kyle
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edkyle99 - 22/10/2007 12:56 PM
meiza - 22/10/2007 10:58 AM
There is a delta growth path presentation from Boeing somewhere that has the multiple options with both benefit and risk listed. Cross feed is one of them... I think it was from 2004 or 2005 and related to the VSE somehow. It had options to the future for 8 meter cores I think...
I stand corrected. There was one mention of propellant crossfeed for one of the projected growth options in that presentation, but it was combined with propellant densification, which probably accounted for a majority of the payload growth.
- Ed Kyle
What is propellant densification?
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hyper_snyper - 22/10/2007 1:09 PM
edkyle99 - 22/10/2007 12:56 PM
meiza - 22/10/2007 10:58 AM
There is a delta growth path presentation from Boeing somewhere that has the multiple options with both benefit and risk listed. Cross feed is one of them... I think it was from 2004 or 2005 and related to the VSE somehow. It had options to the future for 8 meter cores I think...
I stand corrected. There was one mention of propellant crossfeed for one of the projected growth options in that presentation, but it was combined with propellant densification, which probably accounted for a majority of the payload growth.
- Ed Kyle
What is propellant densification?
Chilling the LH to a lower slush like state. That way you can fit more in the same volume. I would have thought you would take an ISP hit doing that with the RS-68. Does it not have an issue with the LH entering at to low a temp to begin with? I wonder what that does to the LOM numbers on the LH turbine. Solids in the flow and all.
It was also a trick proposed for NASP.
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Densification has been an idea floating around for some time now, but is at a high technical risk level... There are multiple modifications required to the ground storage/transfer system and vehicle tank system. And would require a new engine development and certification program to test the engine at the ultra low propellant temps (meaning test stand mods too). The general feeling is that it isn't worth the cost...
Crossfeed is something both Atlas and Delta have studied in depth... It is a very large performance improvement. The idea being that you don't deplete the Core Booster propellant when the Strapons are burning. This maximizes the stagging effect of the three-body booster system. By not using any Core propellant until the Strapons drop away you make sure the impulse provided by the proepllant isn't wasting lifting Strapon dry weight. It is actually somewhat more prefered than adding solids by some in the community. Difficulties are in getting the propellant transfer switching on the Core and umbilical seperation system to work under very high lfowrates ~1600lbm of LOx. There's some expereience doing this at engine test stands and thru the Shuttle program to anchor designs to.
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Why is the LH2 tank below the LOX tank, as opposed to the more structurally efficient method of putting the LOX tank on the bottom?
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tnphysics - 9/11/2007 9:59 PM
Why is the LH2 tank below the LOX tank, as opposed to the more structurally efficient method of putting the LOX tank on the bottom?
controllability
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Here (http://www.lpi.usra.edu/opag/announcements.html) you can find some very detailed Outer Planets Mission Flagship Study Reports (Cassini class missions to Jupiter/Saturn and their moons). They require a Delta IVH or an Atlas 551. The cost given for the launch service are:
- Delta IVH: $486 million
- Atlas 551: $190 million
The Atlas number is exactly in the New Horizons / Juno range. I am surprised by the Delta IVH, 2.5 times more than Atlas for less than 50% more performance. This is almost the Titan 4B Centaur price class Cassini paid. Any comments.
Analyst
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-- >> from the live thread
Jim - 10/11/2007 6:33 PM
This isn't much different than a Titan IV-B Centaur
Jim please confirm - is it closer to a Proton?
Edited by Avron 10/11/2007 6:45 PM
looking for idea in pct..
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Lift performance or first stage thrust?
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Jim - 10/11/2007 7:06 PM
Lift performance or first stage thrust?
1st stage
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1.95 Mlb vs 2.0/2.1 Mlb
DIV Proton
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How is that possible, when DIV lifts so much more? Much better mass fraction?
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thrust isn' everything
ISP, launch site, mass fraction
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GncDude - 10/11/2007 6:29 PM
How is that possible, when DIV lifts so much more? Much better mass fraction?
Proton actually has a propellant better mass fraction (much smaller dry mass), but Delta IV uses higher energy propellant. It gets better "gas mileage", in effect, than Proton, so that it can weigh slightly less at lift off. But that propellant, liquid hydrogen fuel and liquid oxygen oxidizer, is less dense than Proton's hypergolic propellants, which means that Delta IV must be physically larger than Proton (and thus weigh much more dry - and cost much more as a result). Delta IV stands about 15 meters taller and measures 15.3 meters across its widest point compared to 7.4 meters for Proton.
- Ed Kyle
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I thought 1st stage thrust was closely related to lift off weight, just a fraction above (with some exceptions).
Edit: Oh I see. It's the liftoff weight what depends on ISP, not just mass fraction. Thanks. Darn, gotta got back to those equations.
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GncDude - 10/11/2007 12:49 AM
I thought 1st stage thrust was closely related to lift off weight, just a fraction above (with some exceptions).
If you want a really good side-by-side of this not being true - watch the MRO launch and the New Horizons launch side by side. MRO crawls away, New Horizons set's off like a scalded cat.
Doug
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MKremer - 25/9/2007 5:20 PM
As mentioned by others here in different threads, the RS-68 design just takes too much ground support equipment to prep for ignition and through the ignition cycle. To try to add all it needs into an upper stage would add so much mass it would overwhelm any advantages the engine could offer.
I thought that RS-68 is "simpler" than SSME and is cheaper to operate. From the above I learn that it is not completely simple. Can someone give a bit more details where RS-68 stands in comparison to SSME and to, say, RD-180? I mean, how much work/money is needed at the pad to support it?
Another question: it looks counter-intuitive to me that rocket engines, which are designed to withstand huge temperatures and pressures when they operate, are not able to withstand ordinary sea water. It was said several times that first stage engine recovery is not cost efficient because of required cleanup / replacement of parts damaged by sea water. Any more details on what happens to engine which was immersed in sea water for an hour?
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Analyst - 10/11/2007 2:18 PM
Here (http://www.lpi.usra.edu/opag/announcements.html) you can find some very detailed Outer Planets Mission Flagship Study Reports (Cassini class missions to Jupiter/Saturn and their moons). They require a Delta IVH or an Atlas 551. The cost given for the launch service are:
- Delta IVH: $486 million
- Atlas 551: $190 million
The Atlas number is exactly in the New Horizons / Juno range. I am surprised by the Delta IVH, 2.5 times more than Atlas for less than 50% more performance. This is almost the Titan 4B Centaur price class Cassini paid. Any comments.
Analyst
Wow. It takes one's breath away, doesn't it? I knew it had to be a lot more money than everyone thought, but nearly half a billion dollars?
Scaled with payload capability from the 551 cost, Delta 4H should cost $307 million. $486 million is just outrageous, scandalous even, especially considering that Boeing once listed these at $150 million (!) and that EELV Heavy was sold to the taxpayers as a cost savings compared to Titan 4.
First the EELV scandal. Then the FIA disaster. Now this (although I know that Delta 4 now falls within the United Launch Alliance realm, Boeing developed it). Is Boeing working for, or against, the U.S. government?
- Ed Kyle
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gospacex - 11/11/2007 11:28 AM
1. I thought that RS-68 is "simpler" than SSME and is cheaper to operate. From the above I learn that it is not completely simple.
Another question: it looks counter-intuitive to me that rocket engines, which are designed to withstand huge temperatures and pressures when they operate, are not able to withstand ordinary sea water. It was said several times that first stage engine recovery is not cost efficient because of required cleanup / replacement of parts damaged by sea water. Any more details on what happens to engine which was immersed in sea 2. water for an hour?
1. It is simpler and cheaper. It was not designed for airstart and by putting start systems on the ground it made it simpler and less complex. The bulk of the GSE required is a large He tank
2. Plain and simple. Corrosion
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Jim - 11/11/2007 12:10 PM
gospacex - 11/11/2007 11:28 AM
Any more details on what happens to engine which was immersed in sea water for an hour?
Plain and simple. Corrosion
Lots of corrosion after only an hour or two? Can you quantify it?
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gospacex - 11/11/2007 2:06 PM
Jim - 11/11/2007 12:10 PM
gospacex - 11/11/2007 11:28 AM
Any more details on what happens to engine which was immersed in sea water for an hour?
Plain and simple. Corrosion
Lots of corrosion after only an hour or two? Can you quantify it?
The engine isn't going to retrieved immediately. The SRB's spend at least a day in the water. Nevertheless, even if it was pulled out after 1 -2 hours, it is going to be disassembled for days or weeks.
It is not just the salt water on the engines, it effects the other components also (avoinics, hydraulics, structure, etc) . It is water impact loads on the structure, components, etc
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edkyle99 - 11/11/2007 9:06 AM
Analyst - 10/11/2007 2:18 PM
Here (http://www.lpi.usra.edu/opag/announcements.html) you can find some very detailed Outer Planets Mission Flagship Study Reports (Cassini class missions to Jupiter/Saturn and their moons). They require a Delta IVH or an Atlas 551. The cost given for the launch service are:
- Delta IVH: $486 million
- Atlas 551: $190 million
The Atlas number is exactly in the New Horizons / Juno range. I am surprised by the Delta IVH, 2.5 times more than Atlas for less than 50% more performance. This is almost the Titan 4B Centaur price class Cassini paid. Any comments.
Analyst
Wow. It takes one's breath away, doesn't it? I knew it had to be a lot more money than everyone thought, but nearly half a billion dollars?
Scaled with payload capability from the 551 cost, Delta 4H should cost $307 million. $486 million is just outrageous, scandalous even, especially considering that Boeing once listed these at $150 million (!) and that EELV Heavy was sold to the taxpayers as a cost savings compared to Titan 4.
First the EELV scandal. Then the FIA disaster. Now this (although I know that Delta 4 now falls within the United Launch Alliance realm, Boeing developed it). Is Boeing working for, or against, the U.S. government?
If the Delta 4H costs almost as much as a Titan 4 Centaur when flying at a much lower flight rate, it is a cost savings of sorts. And you of all people should know that linearly scaling cost with payload weight is completely meaningless. Based strictly on payload weight, the DSP could have flown on an Atlas 521.
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Not true. The adapter weighed over 2000 ibs.
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gospacex - 11/11/2007 10:28 AM
MKremer - 25/9/2007 5:20 PM
As mentioned by others here in different threads, the RS-68 design just takes too much ground support equipment to prep for ignition and through the ignition cycle. To try to add all it needs into an upper stage would add so much mass it would overwhelm any advantages the engine could offer.
I thought that RS-68 is "simpler" than SSME and is cheaper to operate.
Another question: it looks counter-intuitive to me that rocket engines, which are designed to withstand huge temperatures and pressures when they operate, are not able to withstand ordinary sea water. It was said several times that first stage engine recovery is not cost efficient because of required cleanup / replacement of parts damaged by sea water. Any more details on what happens to engine which was immersed in sea water for an hour?
There's a slight confusion here. RS-68 as currently used requires virtually no maintenance after acceptance. Heck, the Heavy Demo engine covers blew into the river during one of the 2004 hurricanes, and those 3 worked just fine. An air-start RS-68, actually its stage, would not be simple because of the spin-start equipment and new ignition system that would have to be carried along.
As others have said, corrosion is the big problem. Nothing highly electronegative touches any engine in normal use after assembly, AFAIK. Specifically, it's stress corrosion cracking that worries everyone, which is a 3-legged stool. Elevated electrons, electromotive force and an electron transfer medium are required. Well, anything pre-stressed has elevated electrons, so those are all over engines and rockets. As for the other two legs, there are two words: SEA WATER.
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People are overdramatizing it. This is what happens when you design a vehicle for customers that subsequently disappear. Your costs go up because there are just not enough payloads out there. If they could get customers like Arianespace can and get four or more flights a year then I'll bet the costs go down a LOT.
Also, it costs more because a) there's more raw material cost and b) there are three of them strapped together, aggravating a). If you built an A-V heavy and flew it at D-IVH flight rates it would likely cost about the same and have roughly the same performance.
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Jim - 11/11/2007 1:21 PM
gospacex - 11/11/2007 2:06 PM
Lots of corrosion after only an hour or two? Can you quantify it?
The engine isn't going to retrieved immediately. The SRB's spend at least a day in the water. Nevertheless, even if it was pulled out after 1 -2 hours, it is going to be disassembled for days or weeks.
I still don't understand. RS-68 is much smaller than SRB, should be easier to retrieve. So, you separate RS-68 from the (remains of) CBC tank, lift it on recovery ship's deck, wash it out with clean water and then with suitable non-water solvent to remove corrosive agents. All while transporting it to land. It should be possible to do it in a few hours, not days.
It is not just the salt water on the engines, it effects the other components also (avoinics, hydraulics, structure, etc) . It is water impact loads on the structure, components, etc
Avionics are tiny and I am not concerned with reusing those, just replace those. I am thinking about big expensive hardware parts. Engine, hydraulics and structure are designed to withstand at least static 5-6 g, and severe vibrations and aerodynamic loads in flight. Why they are so fragile when it comes to water impact? Can some modest parachute make impact less severe?
(I am not saying that you are wrong, just trying to get more info to understand this stuff. This is Q&A thread after all)
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yinzer - 11/11/2007 1:22 PM
edkyle99 - 11/11/2007 9:06 AM
Analyst - 10/11/2007 2:18 PM
Here (http://www.lpi.usra.edu/opag/announcements.html) you can find some very detailed Outer Planets Mission Flagship Study Reports (Cassini class missions to Jupiter/Saturn and their moons). They require a Delta IVH or an Atlas 551. The cost given for the launch service are:
- Delta IVH: $486 million
- Atlas 551: $190 million
The Atlas number is exactly in the New Horizons / Juno range. I am surprised by the Delta IVH, 2.5 times more than Atlas for less than 50% more performance. This is almost the Titan 4B Centaur price class Cassini paid. Any comments.
Analyst
Wow. It takes one's breath away, doesn't it? I knew it had to be a lot more money than everyone thought, but nearly half a billion dollars?
Scaled with payload capability from the 551 cost, Delta 4H should cost $307 million. $486 million is just outrageous, scandalous even, especially considering that Boeing once listed these at $150 million (!) and that EELV Heavy was sold to the taxpayers as a cost savings compared to Titan 4.
First the EELV scandal. Then the FIA disaster. Now this (although I know that Delta 4 now falls within the United Launch Alliance realm, Boeing developed it). Is Boeing working for, or against, the U.S. government?
If the Delta 4H costs almost as much as a Titan 4 Centaur when flying at a much lower flight rate, it is a cost savings of sorts. And you of all people should know that linearly scaling cost with payload weight is completely meaningless. Based strictly on payload weight, the DSP could have flown on an Atlas 521.
Nonlinear scaling should actually work the other way. Delta 4H, since it carries more payload, should cost LESS than Atlas 551 on a dollar per kg payload basis.
- Ed Kyle
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gospacex - 11/11/2007 6:25 PM
1. I still don't understand. RS-68 is much smaller than SRB, should be easier to retrieve. So, you separate RS-68 from the (remains of) CBC tank, lift it on recovery ship's deck, wash it out with clean water and then with suitable non-water solvent to remove corrosive agents. All while transporting it to land. It should be possible to do it in a few hours, not days.
2. Avionics are tiny and I am not concerned with reusing those, just replace those. I am thinking about big expensive hardware parts. Engine, hydraulics and structure are designed to withstand at least static 5-6 g, and severe vibrations and aerodynamic loads in flight. Why they are so fragile when it comes to water impact? Can some modest parachute make impact less severe?
1. The CBC has to be recovered to get the RS-68. If the CBC parachutes with the nozzle down like the SRB, splash down loads on the nozzle are going to break things. If the CBC is going to land nose down, then there has to be some engineering to make the parachute riser lines avoid the engine.
Removing an engine is not an easy thing and can't be done in the water. Both the engine and CBC have to be supported to allow the connections to be undone, in a unloaded condition. "suitable non-water solvent" on a ship? This would have to be done on land to keep the EPA happy. Days not hours.
Also it wasn't designed for reused. It has smaller margins.
2. Static loads are not the same as shock loads or splash down loads. Water impact is the same as the ground. One shuttle flight's SRB parachutes did not deploy. The casings shattered on impact. The bottom segments of the SRB's have reinforcing bands to prevent them from crushing when the splashdown "hole"* in the water collapses
*cannonball effect
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Jim - 11/11/2007 5:56 PM
1. The CBC has to be recovered to get the RS-68. If the CBC parachutes with the nozzle down like the SRB, splash down loads on the nozzle are going to break things. If the CBC is going to land nose down, then there has to be some engineering to make the parachute riser lines avoid the engine.
Removing an engine is not an easy thing and can't be done in the water. Both the engine and CBC have to be supported to allow the connections to be undone, in a unloaded condition.
Why CBC's bottom with attached RS-68 cannot be severed by small explosive charges? CBC has those charges anyway as part of range safety system, why not place them near the bottom of CBC? Detonate those, and slow down the aft part with parachutes. RS-68 weight is ~6.5 tonnes. There you are - no need to detach RS-68 from CBC, just fish it out from the ocean, Still won't work?
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gospacex - 11/11/2007 8:12 PM
Why CBC's bottom with attached RS-68 cannot be severed by small explosive charges? CBC has those charges anyway as part of range safety system, why not place them near the bottom of CBC? Detonate those, and slow down the aft part with parachutes. RS-68 weight is ~6.5 tonnes. There you are - no need to detach RS-68 from CBC, just fish it out from the ocean, Still won't work?
Nope. Range safety charges are longitudinal to split the tanks to allow mixing of the propellants, not to sever them. The linear shaped charges are placed in the system tunnels running the length of the tanks.
Range safety systems can't be used for other purposes.
Splitting/severing the tanks would also destroy the buoyancy "system" and the engine sinks
Changes for re usability makes the system more complicated (which is less reliable) and more expensive
Re usability doesn't pay off until higher flight rates, >20 or so flights per year
40-60 for an RLV
40
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edkyle99 - 11/11/2007 11:06 AM
First the EELV scandal. Then the FIA disaster. Now this (although I know that Delta 4 now falls within the United Launch Alliance realm, Boeing developed it). Is Boeing working for, or against, the U.S. government?
And you forgot the KC-767 Lease scandal and the CH-47 CSAR helicopter fiasco. It does seem on the surface that Boeing is working diligently against the US taxpayer.
But wasn't Boeing's pricing that you cite actually based on them winning the EELV contract in its entirety? In other words, they didn't so the prices are no longer valid. Still, nearly half a million dollars for a D-IVH is ridiculous.
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yinzer - 11/11/2007 8:22 PM
If the Delta 4H costs almost as much as a Titan 4 Centaur when flying at a much lower flight rate, it is a cost savings of sorts.
Together with the other Delta IV variants, the CBC flies as often or even more than Titan 4 did. This is the whole concept of the CBC: Add two boosters to the core and you get a Heavy. So to classify Delta IVH as an independent vehicle (with no connection to the other variants) is wrong. Because of this the flight rate is higher. And because of this § 486 million is extreme, imho.
Analyst
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Analyst - 12/11/2007 3:16 AM
So to classify Delta IVH as an independent vehicle (with no connection to the other variants) is wrong.
It is an as an independent vehicle, unfortunately. The all three CBC's of a D-IV heavy are unique and can't be used in any other configuration. In total, there are 5 or 6 unique, CBC's with the Medium, and Medium Plus 4/5 as the others. This all due to the vehicle unable to meet performance goals.
Unlike, Atlas V which only has one core and could do its heavy with the same 3 cores. Which would make only 2 variations and I believe the heavy config CCB can be converted back to STD
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What's the deal with this $486M figure? I could not find the details at that link provided earlier but it's got to be way off the mark. It was probably based on what Boeing had to pay before the Buy 3 contracts were put in place. The COTS RFI ULA put out shows the DIVH costs less per Kg than the Atlas 551 18K/Kg vs 20K/Kg. Using that chart the DSP launch would have been 108M. Now I know there's other costs on top of that but 300M+? Please provide a more direct link or something that backs up that number.
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bigdog - 12/11/2007 8:00 AM
What's the deal with this $486M figure? I could not find the details at that link provided earlier but it's got to be way off the mark. It was probably based on what Boeing had to pay before the Buy 3 contracts were put in place. The COTS RFI ULA put out shows the DIVH costs less per Kg than the Atlas 551 18K/Kg vs 20K/Kg. Using that chart the DSP launch would have been 108M. Now I know there's other costs on top of that but 300M+? Please provide a more direct link or something that backs up that number.
It is here
http://www.lpi.usra.edu/opag/announcements.html
and more specifically here, in the cost section of the Jupiter System Observer Mission Study report.
http://www.lpi.usra.edu/opag/jso_final_report.pdf
A shame too. They really want to use Delta 4H, but apparently can't afford it.
- Ed Kyle
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bigdog - 12/11/2007 3:00 PM
What's the deal with this $486M figure? I could not find the details at that link provided earlier but it's got to be way off the mark. It was probably based on what Boeing had to pay before the Buy 3 contracts were put in place. The COTS RFI ULA put out shows the DIVH costs less per Kg than the Atlas 551 18K/Kg vs 20K/Kg. Using that chart the DSP launch would have been 108M. Now I know there's other costs on top of that but 300M+? Please provide a more direct link or something that backs up that number.
Open this (http://www.lpi.usra.edu/opag/europa_explorer_public_report_1.pdf) document, and go to pdf page 116 for Delta IV and 122 for Atlas 551 and you will find the numbers (Atlas is actual §191 million). Or take the Enceladus Study linked above via the OPAG site. The numbers in the document you used are for LEO missions, Delta IVH takes ~ 25.000kg to LEO, which gives ~$450 million taking the 18k/kg number. Quite close to $486 million. And quite expensive.
Analyst
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With these prices, I think we need to consider outsourcing the launch of all but the most sensitive spacecraft. Looking at the projected launch schedule the 'Heavy' is only scheduled to fly a few NRO launches .. perhaps at a rate of one per year. At that usage, costs will only continue to escalate.
If I were managing that 2017 Jupiter bound spacecraft, I'd have to size it to ride out of here on Ariane or engage in creative back room politicking to get the government to subsidize a domestic launcher.
I have the utmost respect for the intellect required to finally get the 'Heavy' working and am not trying affront those who worked diligently. BUT the entire project, from a business viewpoint, will be a failure if no one can afford to use it.
--- CHAS
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One can not quote the prices in NASA announcements of opportunity as launch vehicle costs. There are other costs in the numbers that are in addition to the vehicle price. Among the costs are payload processing facility, telemetery, nuclear approval, mission unique mods, NASA overhead, etc. Those can amount to 20% of the vehicle cost. NASA missions never use "stock" vehicles. Commercial prices would be cheaper
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Who was the last commercial customer to book a Delta IV Heavy?
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HIPAR - 12/11/2007 11:25 AM
If I were managing that 2017 Jupiter bound spacecraft, I'd have to size it to ride out of here on Ariane or engage in creative back room politicking to get the government to subsidize a domestic launcher.
The taxpayers would already be BUYING the launch. Exactly how would SUBSIDIZING the launch be different? Neither NASA or DoD can buy a foreign ride, though they can be gotten for free like JWST. Don't ask me what the out-clause is for Soyuz.
But in general, the Heavy launch rate doesn't really determine the price. Rather the overall launch rate of Delta IV. The components are basically the same, as is the engineering staff. The standing army costs have to be spread over the number of vehicles in a year, not to mention the fact that the parent companies are still trying to recover their investments, especially Boeing since it developed a new engine.
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Avionics are tiny and I am not concerned with reusing those, just replace those. I am thinking about big expensive hardware parts.
Um, a flight computer costs about as much as an upper stage engine. Many avionics boxes are not tiny, though they may be tinny.
Why CBC's bottom with attached RS-68 cannot be severed by small explosive charges?
That adds a few more failure modes and a heavy structural joint. Granted, MA-5 had that, but that was accounted for in the performance and reliability allocations.
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HIPAR - 12/11/2007 11:25 AM
With these prices, I think we need to consider outsourcing the launch of all but the most sensitive spacecraft.
That would end up making the "most sensitive" launches outrageously expensive. Your cost savings would vanish. The EELV concept was the right way to go. The mistake was the DoD's hare-brained decision to fund two of them.
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Jim - 12/11/2007 7:33 AM
...snip... there are 5 or 6 unique, CBC's with the Medium, and Medium Plus 4/5 as the others. This all due to the vehicle unable to meet performance goals.
...more snip
Can you be more specific about DIV's underperformance? I have read that: 1) tanks sized smaller than optimal, 2) dry weight of CBC turned out to be heavier than anticipated. Am I missing anything?
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Thorny - 12/11/2007 1:34 PM
HIPAR - 12/11/2007 11:25 AM
With these prices, I think we need to consider outsourcing the launch of all but the most sensitive spacecraft.
That would end up making the "most sensitive" launches outrageously expensive. Your cost savings would vanish. The EELV concept was the right way to go. The mistake was the DoD's hare-brained decision to fund two of them.
The mistake was that they started with funding for only half the NRE of a single contractor(ie. $1B). After the contractors were committed, they compounded this mistake by then splitting it between two contractors in an attempt to force recurring price competition.
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Thorny - 12/11/2007 12:34 PM
That would end up making the "most sensitive" launches outrageously expensive. Your cost savings would vanish. The EELV concept was the right way to go. The mistake was the DoD's hare-brained decision to fund two of them.
I like your analysis.
I worked with the DoD for 35 years; a major portion of the cold war. The one thing about outrageously expensive national security funding that I noticed was they always paid the price. Somewhere they always found money that 'wasn't there'. I never asked how because I knew they wouldn't tell so I just worked the programs. Only really special sensitive programs enjoy that big ticket luxury. NASA science missions do not.
What ULA needs is some good old Yankee entrepreneurial spirit. Work the books .. let's somehow entice the likes of Echostar, DirecTV or whoever to buy in for Heavy launches. Even if they go for a per launch net loss, there is some income to defray taxpayers costs. You also keep that expensive rocket factory in Alabama working.
--- CHAS
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Jim - 12/11/2007 6:36 PM
One can not quote the prices in NASA announcements of opportunity as launch vehicle costs. There are other costs in the numbers that are in addition to the vehicle price. Among the costs are payload processing facility, telemetery, nuclear approval, mission unique mods, NASA overhead, etc. Those can amount to 20% of the vehicle cost. NASA missions never use "stock" vehicles. Commercial prices would be cheaper
All valid. But all this is true for Atlas 551 too. Substract 20% each and you end up with the same difference.
Never has a Delta IVH be ordered commercially, Boeing stopped offering all Delta IV sometime arround 2003.
Is there anyone out who believes Ares I will fly for ~$150 million variable, Ares V for~$500 million variable? I don't want a distraction from Delta IV, but the Delta IV (and Atlas V) numbers put these Ares dream numbers into perspective: Pretty impossible.
Analyst
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EE Scott - 12/11/2007 2:40 PM
Jim - 12/11/2007 7:33 AM
...snip... there are 5 or 6 unique, CBC's with the Medium, and Medium Plus 4/5 as the others. This all due to the vehicle unable to meet performance goals.
...more snip
Can you be more specific about DIV's underperformance? I have read that: 1) tanks sized smaller than optimal, 2) dry weight of CBC turned out to be heavier than anticipated. Am I missing anything?
The Medium requirement was 10K lb to GTO
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I have a question on last Saturday's D4 Heavy launch. Looking at the pics taken by Ben Cooper, the lower parts seem nice and orange, and not the black charred appearance of the debut launch. It seems the ignition burn off mushroomed up along the boosters just like last time, but they didn't seem as affected as before. Did something change... wind conditions?
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Jim - 12/11/2007 4:53 PM
EE Scott - 12/11/2007 2:40 PM
Jim - 12/11/2007 7:33 AM
...snip... there are 5 or 6 unique, CBC's with the Medium, and Medium Plus 4/5 as the others. This all due to the vehicle unable to meet performance goals.
...more snip
Can you be more specific about DIV's underperformance? I have read that: 1) tanks sized smaller than optimal, 2) dry weight of CBC turned out to be heavier than anticipated. Am I missing anything?
The Medium requirement was 10K lb to GTO
Understood. DIV was designed for specific requirements, back in the '90s. Outside of those requirements lies non-optimal land, possibly even kludgville. I suppose that Atlas' higher thrust kero first stage gives it inherently more versitility in this regard.
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Analyst - 12/11/2007 3:16 PM
...snip...
Is there anyone out who believes Ares I will fly for ~$150 million variable, Ares V for~$500 million variable? I don't want a distraction from Delta IV, but the Delta IV (and Atlas V) numbers put these Ares dream numbers into perspective: Pretty impossible.
Analyst
Dead on. Have preliminary estimates of LV costs ever been accurate? Ares I/V, if they ever get built will dwarf the estimates above, just as the Shuttle did before it.
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RedSky - 12/11/2007 5:02 PM
I have a question on last Saturday's D4 Heavy launch. Looking at the pics taken by Ben Cooper, the lower parts seem nice and orange, and not the black charred appearance of the debut launch. It seems the ignition burn off mushroomed up along the boosters just like last time, but they didn't seem as affected as before. Did something change... wind conditions?
Yep, wind and weather have a lot to do with whether or not you get lots of charring. You can get significant charring even with the smaller Medium fireballs (see the first flight Flight Readiness Firing, GOES-N, and NROL-22 for examples). Just depends on the weather.
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EE Scott - 12/11/2007 4:20 PM
Analyst - 12/11/2007 3:16 PM
...snip...
Is there anyone out who believes Ares I will fly for ~$150 million variable, Ares V for~$500 million variable? I don't want a distraction from Delta IV, but the Delta IV (and Atlas V) numbers put these Ares dream numbers into perspective: Pretty impossible.
Analyst
Dead on. Have preliminary estimates of LV costs ever been accurate? Ares I/V, if they ever get built will dwarf the estimates above, just as the Shuttle did before it.
The "V" in Ares V could end up standing for $5 billion!
- Ed Kyle
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EE Scott - 12/11/2007 1:40 PM
Can you be more specific about DIV's underperformance? I have read that: 1) tanks sized smaller than optimal, 2) dry weight of CBC turned out to be heavier than anticipated. Am I missing anything?
The early Isp predictions on the RS-68 were about 10sec higher than what was achieved. Turbopump efficiency estimates were high, IIRC. I'm sure one can find early marketing numbers for RS-68 that quoted 420sec.
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I've heard that RS-68 produced less thrust than anticipated but I guess that's just another way of saying the same... (Ie same mass flow but less ISP.)
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McDew - 12/11/2007 1:40 PM Thorny - 12/11/2007 1:34 PM HIPAR - 12/11/2007 11:25 AM With these prices, I think we need to consider outsourcing the launch of all but the most sensitive spacecraft.
That would end up making the "most sensitive" launches outrageously expensive. Your cost savings would vanish. The EELV concept was the right way to go. The mistake was the DoD's hare-brained decision to fund two of them.
The mistake was that they started with funding for only half the NRE of a single contractor(ie. $1B). After the contractors were committed, they compounded this mistake by then splitting it between two contractors in an attempt to force recurring price competition.
Correct. Look at it this way - we believe in 1) one expensive sole-source that we watch like a hawk (this because we're realistic about launch rates and the business), or 2) artificially expand the number of launch vehicle vendors, knowing after the early cut we'll only have one (this because we aren't sure of which one of the vendors is to be THE vendor), or 3) we expand the market with multiple vendors by committing to a greater launch rate that teases for supporting a competitive launch environment.
What happened was a "head fake". 3) then 2) then 1). Except we have two vehicles in one "company".
The only way access to space will get cheaper is increased launch rate. It's a chicken-and-egg puzzle, because you need as a premise a cheaper LV that exploits the launch rate to permanently expand launch rates. Every attempt to force (shuttle, eelv) this has failed in the past. The argument for the prior Jarvis vehicle to Boeing was to capitalize on the opportunity of having the whole engine development costs underwritten by the prior lunar program.
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nobodyofconsequence - 13/11/2007 11:24 AM
McDew - 12/11/2007 1:40 PM Thorny - 12/11/2007 1:34 PM HIPAR - 12/11/2007 11:25 AM With these prices, I think we need to consider outsourcing the launch of all but the most sensitive spacecraft.
That would end up making the "most sensitive" launches outrageously expensive. Your cost savings would vanish. The EELV concept was the right way to go. The mistake was the DoD's hare-brained decision to fund two of them.
The mistake was that they started with funding for only half the NRE of a single contractor(ie. $1B). After the contractors were committed, they compounded this mistake by then splitting it between two contractors in an attempt to force recurring price competition.
Correct. Look at it this way - we believe in 1) one expensive sole-source that we watch like a hawk (this because we're realistic about launch rates and the business), or 2) artificially expand the number of launch vehicle vendors, knowing after the early cut we'll only have one (this because we aren't sure of which one of the vendors is to be THE vendor), or 3) we expand the market with multiple vendors by committing to a greater launch rate that teases for supporting a competitive launch environment.
What happened was a "head fake". 3) then 2) then 1). Except we have two vehicles in one "company".
The only way access to space will get cheaper is increased launch rate. It's a chicken-and-egg puzzle, because you need as a premise a cheaper LV that exploits the launch rate to permanently expand launch rates. Every attempt to force (shuttle, eelv) this has failed in the past. The argument for the prior Jarvis vehicle to Boeing was to capitalize on the opportunity of having the whole engine development costs underwritten by the prior lunar program.
You guys are all getting your panties in a bunch. These launch cost numbers are nothing compare to the value of spacecrafts they put up in orbit, reliably, accurately, and ITAR controlled.
You have to pay to play. The initial DoD EELV procurement took advantage of both companies' rosey outlook in the commercial market and "low-balled" the entry into this program. DoD saved billions of RDT&E $$$ as a result. Now the market has changed. Simple fact is, if you still want to launch classified payloads, you have to use the EELV. They existed now only because the DoD wants it, so pay for them. Both companies made it clear, it's you baby now and if you don't want it then shut it down. We've washed our hands.
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Propforce - 13/11/2007 1:41 PM You guys are all getting your panties in a bunch. These launch cost numbers are nothing compare to the value of spacecrafts they put up in orbit, reliably, accurately, and ITAR controlled.
You have to pay to play. The initial DoD EELV procurement took advantage of both companies' rosey outlook in the commercial market and "low-balled" the entry into this program. DoD saved billions of RDT&E $$$ as a result. Now the market has changed. Simple fact is, if you still want to launch classified payloads, you have to use the EELV. They existed now only because the DoD wants it, so pay for them. Both companies made it clear, it's you baby now and if you don't want it then shut it down. We've washed our hands.
... which is why any enhancement of EELV will be limited because there will be no ROI and limited ability to pay for a high-end LV beyond DIVH. They'll want to mine out Atlas V medium and Delta IVH for all they can get.
All decisions have repercussions. Including the quality and reliability if underfunded. E.g. don't gloat.
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NASA could have quadrupled the US launch market, but it didn't. It's not NASA's core mission, but it's not NASA's core mission to build custom boosters when alternatives are available on the market either.
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Antares - 13/11/2007 12:04 PM
EE Scott - 12/11/2007 1:40 PM
Can you be more specific about DIV's underperformance? I have read that: 1) tanks sized smaller than optimal, 2) dry weight of CBC turned out to be heavier than anticipated. Am I missing anything?
The early Isp predictions on the RS-68 were about 10sec higher than what was achieved. Turbopump efficiency estimates were high, IIRC. I'm sure one can find early marketing numbers for RS-68 that quoted 420sec.
Interesting. Thanks.
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Propforce - 13/11/2007 1:41 PM
You guys are all getting your panties in a bunch. These launch cost numbers are nothing compare to the value of spacecrafts they put up in orbit, reliably, accurately, and ITAR controlled.
You have to pay to play. The initial DoD EELV procurement took advantage of both companies' rosey outlook in the commercial market and "low-balled" the entry into this program. DoD saved billions of RDT&E $$$ as a result. Now the market has changed. Simple fact is, if you still want to launch classified payloads, you have to use the EELV. They existed now only because the DoD wants it, so pay for them. Both companies made it clear, it's you baby now and if you don't want it then shut it down. We've washed our hands.
The reason for complaint is that if the costs haven't fallen from pre-EELV days, then why were these things even developed? I'm starting to think that it might have been better for the U.S. Government to have stuck with Titan IV, Atlas IIAS, and Delta II.
- Ed Kyle
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Well, RS-68 will now be used for Ares V as well. :)
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What is the core mass at strap-on jettison?
And why do some spacecraft require the old Titan IV fairing?
And why did this Delta IV Heavy seem to accelerate so much more quickly than the last?
I know that it did not, but that's how it appeared.
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tnphysics - 13/11/2007 7:46 PM
What is the core mass at strap-on jettison?
And why do some spacecraft require the old Titan IV fairing?
And why did this Delta IV Heavy seem to accelerate so much more quickly than the last?
I know that it did not, but that's how it appeared.
a) Probably the CBC dry mass, about 27.7 tons, which can be found on Ed's site here, plus about 90 second's worth of propellant (which could be calculated I think):
http://www.geocities.com/launchreport/delta4.html
b) I don't know. Some have suggested something to do with the PAF that DSP requires, could be EMI, RF, or other things too.
c) Could be because of the night launch. Also DSP is lighter than the Demosat (about 4400 kg including the PAF versus 6500 kg for Demosat) so it would move slightly faster off the line, so to speak.
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edkyle99 - 13/11/2007 5:59 PM The reason for complaint is that if the costs haven't fallen from pre-EELV days, then why were these things even developed? I'm starting to think that it might have been better for the U.S. Government to have stuck with Titan IV, Atlas IIAS, and Delta II. - Ed Kyle
Exactly what I was trying to imply. Full circle to nowhere.
The whole point was to "evolve" the technology and the market through competition.
Not sad to see Titan IV go - that mother scared the sh*t out of me! But really, we haven't changed things much - if we didn't have EELV's, we'd have still done the launches we've done, and the net cost difference, even considering NPV/FV backwards/forwards, wouldn't pay diddly squat. CBC's only payoff when we're at 6x our current rate - fat chance.
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Just for note the RAND Nation Security Space Report clearly states the EELV program has meet it's goal of a 25-50 percent reduction in launch costs. Believe what you will, but I believe this more than a bunch of internet gossip.
And those that think the Delta and Atlas vehicles weren't developed in a competitive enviornment clearly weren't paying attention. That's what makes ULA such a culture shock... working "arm-in-arm" with your #1 competitor.
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Thrust and Isp are no the same thing... Thrust is a function of the the chamber pressure setting and chamber/nozzle geometry. Isp is the function of the flowrate required to acheive said chamber pressure. They are both required to determine the performance of an engine. You then need mixture ratio to determine how best size the launch vehicle tanks to the optimal system desing point.
RS-68 actually exceeds the thrust originally proposed, running now at 102% or 663klb sea level thrust. It also has a lower throttled thrust, 57% vs. 60%, for optimal Heavy vehicle performance. Of course these changes and some vehicle structure changes were made to mitigate the engine not acheive the desired Isp. Anyone with a historical perspective will recognize that meeting Isp is the biggest risk to any new engine development, and RS-68 lived up to past expectations. And finally the fundamental design causes of the lower Isp have been investigated since and have resulted in the RS-68A/B developments.
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Thanks for the clarification.
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I just checked out the new payload planners guide and went over to the Delta Product Sheet.
WOULD SOMEONE AT ULA FIX THE INSULTING GRAPHIC ON THE FRONT NOW THAT YOU'RE ONE BIG HAPPY FAMILY? It gets the Razzie award for most egregious use of PhotoShop in the launch industry. It shows a Delta IV lifting off from Complex 41, not Complex 37. The view would be from the top of the VIF. Plus, the building in the background IS the VIF, not the LC-37 MST. I can't really tell if those are the LC-37 lightning towers, but I'll give the PR department the benefit of the doubt that they didn't use photoshopped versions of the Atlas V MLP Mast.
Sheesh. How insulting to the other half of your company, not to mention your customers.
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Antares - 15/11/2007 2:59 PM
I just checked out the new payload planners guide and went over to the Delta Product Sheet.
WOULD SOMEONE AT ULA FIX THE INSULTING GRAPHIC ON THE FRONT NOW THAT YOU'RE ONE BIG HAPPY FAMILY? It gets the Razzie award for most egregious use of PhotoShop in the launch industry. It shows a Delta IV lifting off from Complex 41, not Complex 37. The view would be from the top of the VIF. Plus, the building in the background IS the VIF, not the LC-37 MST. I can't really tell if those are the LC-37 lightning towers, but I'll give the PR department the benefit of the doubt that they didn't use photoshopped versions of the Atlas V MLP Mast.
Sheesh. How insulting to the other half of your company, not to mention your customers.
Maybe that's part of the ULA merger. Perhaps they plan to consolidate launch sites soon? ;)
I never noticed that until you pointed it out. It seemed a little weird though...
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pad rat - 15/11/2007 3:17 PM
That is the MST, not the VIF (the cruciform panels are an obvious giveaway), and the lightning towers are the SLC-37 towers.
OK. Well, it's still the back of the MST, which in some ways is even weirder.
Edit: I don't live in Mims, I never see that side.
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Is it planned to add SRBs to the DIVH (either a mostly-current version or the upgraded version that will fly in 2011)?
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tnphysics - 26/11/2007 4:21 PM
Is it planned to add SRBs to the DIVH (either a mostly-current version or the upgraded version that will fly in 2011)?
nope
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tnphysics - 26/11/2007 10:21 PM
Is it planned to add SRBs to the DIVH (either a mostly-current version or the upgraded version that will fly in 2011)?
A version with six solid strap ons ist under study, but there are no plans.
This Version would look a little bit strange, as each two solids would be mounted on each of the CBCs - all six on the same side as the other is occupied on the pad by the pad structures. But even this modification would require pad and and non-trivial-vehicle modifications.
See page 265 on http://www.ulalaunch.com/docs/product_sheet/Delta_IV_Payload_Planners_Guide_2007.pdf
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Skyrocket - 26/11/2007 4:55 PM
tnphysics - 26/11/2007 10:21 PM
Is it planned to add SRBs to the DIVH (either a mostly-current version or the upgraded version that will fly in 2011)?
A version with six solid strap ons ist under study, but there are no plans.
This Version would look a little bit strange, as each two solids would be mounted on each of the CBCs - all six on the same side as the other is occupied on the pad by the pad structures. But even this modification would require pad and and non-trivial-vehicle modifications.
See page 265 on http://www.ulalaunch.com/docs/product_sheet/Delta_IV_Payload_Planners_Guide_2007.pdf
I would have thought that solids would be the perfect solution if one needs to bump up DIVH payload capacity (for non manrated duty). Obviously someone has done the analysis of the costs/other trade-offs of developing higher performance versions of the RS-68 vs solids/pad alterations, since uprating the RS-68 is going on right now I read. Does there exist any data on costs of these two options?
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2 pads would have to be modified. Also uprating the RS-68 is applicable to all the other versions
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Jim - 27/11/2007 8:57 PM
2 pads would have to be modified. Also uprating the RS-68 is applicable to all the other versions
I will go digging around to see if I can find a figure for development costs of the RS-68B or whatever they are calling it. I forgot about the Vandenberg pad - that obviously makes a difference. Plus additional processing costs would be significant. And now that you make the point about the applicability of the uprated engine to other versions, I could see how perhaps some D4 medium launches might be able to delete a solid compared to the baseline engine, if they were borderline.
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D-IV adds SRM's in pairs
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What are the changed parts that make up the RS-68 upgrade? (injectors, nozzle, etc.) I searched but couldn't find anything. Also, is the upgraded RS-68 a "bolt-on" upgrade for Delta IV, i.e. there aren't significant changes to the vehicle itself?
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Nick L. - 9/12/2007 12:45 AM
What are the changed parts that make up the RS-68 upgrade? (injectors, nozzle, etc.) I searched but couldn't find anything. Also, is the upgraded RS-68 a "bolt-on" upgrade for Delta IV, i.e. there aren't significant changes to the vehicle itself?
It wouldn't be on the web. It is ITAR info
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What do some of the calls in the final countdown mean? Some of them are self explanatory like "hydraulic press at 4000" and "NMEQ report swing arm system ready" but others are more cryptic.
T-120 seconds, "DPA script running"
T-40 seconds, "GE main panel off"
Roughly T-25 seconds, "flight lock in (?)"
Roughly T-10 seconds, "**** deactivated" (can't tell what's been deactivated though because the countdown goes over it)
I've always been curious as to what these phrases mean. If they are proprietary then sorry, but just wondering.
Nick
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Jim - 9/12/2007 6:36 AM
Nick L. - 9/12/2007 12:45 AM
What are the changed parts that make up the RS-68 upgrade? (injectors, nozzle, etc.) I searched but couldn't find anything. Also, is the upgraded RS-68 a "bolt-on" upgrade for Delta IV, i.e. there aren't significant changes to the vehicle itself?
It wouldn't be on the web. It is ITAR info
There are AIAA papers on this.
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Antares - 20/12/2007 10:35 PM
Jim - 9/12/2007 6:36 AM
Nick L. - 9/12/2007 12:45 AM
What are the changed parts that make up the RS-68 upgrade? (injectors, nozzle, etc.) I searched but couldn't find anything. Also, is the upgraded RS-68 a "bolt-on" upgrade for Delta IV, i.e. there aren't significant changes to the vehicle itself?
It wouldn't be on the web. It is ITAR info
There are AIAA papers on this.
Ha ha ha... :laugh: :laugh: :laugh: classic !! just classic !! :laugh: :laugh: :laugh:
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Antares - 21/12/2007 12:35 AM
Jim - 9/12/2007 6:36 AM
Nick L. - 9/12/2007 12:45 AM
What are the changed parts that make up the RS-68 upgrade? (injectors, nozzle, etc.) I searched but couldn't find anything. Also, is the upgraded RS-68 a "bolt-on" upgrade for Delta IV, i.e. there aren't significant changes to the vehicle itself?
It wouldn't be on the web. It is ITAR info
There are AIAA papers on this.
I couldn't find any at AIAA.org. Do you have any paper numbers?
This site,
http://www.spaceref.com/news/viewsr.html?pid=23819
mentions that work is underway on RS-68A and RS-68B designs, and that subscale injector testing of a variety of alternative injector designs, performed at MSFC, "benefits the RS-68A and RS-68B engines, as well as the J-2X engine".
My guess is slightly higher thrust and improved ISP provided by slightly revved up turbopumps and/or modified injectors.
- Ed Kyle
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120 seconds is when the flight HTPA (Hydraulic Turbine/Pump Assembly) is spun up. Next has to with locking-out specific commands just prior to launch to assure that the system configuration can not change due to an inadvertant command. Flight lock-in is a command used to set the flight computer in ready-to-fly mode. The last one is simply deactivation of parameter alarms as all control is handed-off to the launch computer in the last few seconds.
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edkyle99 - 28/12/2007 9:44 AM
Antares - 21/12/2007 12:35 AM
Jim - 9/12/2007 6:36 AM
Nick L. - 9/12/2007 12:45 AM
What are the changed parts that make up the RS-68 upgrade? (injectors, nozzle, etc.) I searched but couldn't find anything. Also, is the upgraded RS-68 a "bolt-on" upgrade for Delta IV, i.e. there aren't significant changes to the vehicle itself?
It wouldn't be on the web. It is ITAR info
There are AIAA papers on this.
I couldn't find any at AIAA.org. Do you have any paper numbers?
This site,
http://www.spaceref.com/news/viewsr.html?pid=23819
mentions that work is underway on RS-68A and RS-68B designs, and that subscale injector testing of a variety of alternative injector designs, performed at MSFC, "benefits the RS-68A and RS-68B engines, as well as the J-2X engine".
My guess is slightly higher thrust and improved ISP provided by slightly revved up turbopumps and/or modified injectors.
- Ed Kyle
The paper number is AIAA-2007-5833
I'll just make it easy for you to download :laugh:
From the paper
Planned modifications to the current RS-68 are:
1. Increased power level of roughly 6 percent.
2. Main injector changes to improve Isp roughly 2 percent.
3. New bearing material to decrease stress corrosion susceptibility.
4. Redesigned turbopump pump inlets to incorporate tip vortex suppression.
5. Redesigned fuel turbopump second stage blisk to increase robustness.
6. Redesigned gas generator igniter that eliminates potential foreign object debris.
7. Higher reliability oxidizer turbopump bearing chill sensor.
8. Higher reliability hot gas sensor.
9. Redesigned oxidizer turbopump to reduce pre-start and operational helium usage.
10. Modified engine start sequence/configuration to reduce free hydrogen on the pad during engine start.
11. Redesigned ablative nozzle to accommodate the longer-duration Ares V mission profile.
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OK... can someone tell me how to up load a 8 megabytes PDF document ? This thing won't let me upload anything more than 100K :frown:
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Hmmm. I'm not sure which takes precedence, that the paper was produced by taxpayer dollars or AIAA (copy)rights. I've had problems uploading too. I can't remember what I did to fix it. Maybe PM Chris.
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Interesting, how were these improvements discovered? Were they in the pipeline during the original RS-68 development but time ran out since EELV had a deadline?
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The RS-68 was biased towards cost rather than performance; the current round of enhancements mostly involve improving the existing design without major changes like a fully regenerative nozzle. Others, like reduction of the hydrogen flare at ignition, are driven more by possible safety issues, or just reducing the appearance of problems. I'd like to see a COBRA-style high-performance engine (sort of a simplified SSME with higher thrust), but the money isn't there for radical changes in design.
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Interesting how much they are trying to reduce helium use ... Any chance that by the time they are done it will use less helium than LH ;)
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Helium requirements are almost entirely driven by KSC infrastructure limits. Future helium costs were secondary when the requirements were written, though the white papers have helped make the point.
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Wow, so much great info! This thread is this forum at its best. That's why I love this place. :)
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Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
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Clean burning.
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Actually the decision was driven by what was available in the US industry. Atlas wanted LOX/RP, but Rocketdyne could not develop a cost effective current technology high thrust/performance engine which led them to go the RD-180 route. Boeing's strategy was to stay with a US built engine which drove them to a RS-68 shuttle derived engine with the hope of maintaining shuttle/NASA cost synergy which is paying off with planned upgrades.
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McDew - 1/1/2008 9:42 PM
Actually the decision was driven by what was available in the US industry. Atlas wanted LOX/RP, but Rocketdyne could not develop a cost effective current technology high thrust/performance engine which led them to go the RD-180 route. Boeing's strategy was to stay with a US built engine which drove them to a RS-68 shuttle derived engine with the hope of maintaining shuttle/NASA cost synergy which is paying off with planned upgrades.
Not quite. RS-68 was designed specifically for the Delta-IV. Atlas went to the RD-180 when Rocketdyne couldn't handle two engine developments.
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Jim - 1/1/2008 6:47 PM
Atlas went to the RD-180 when Rocketdyne couldn't handle two engine developments.
Rocketdyne could not compete with the price of RD-180 offerred by the Russians.
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sticksux - 1/1/2008 4:02 PM
Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
Why is it backward to have a H2 first stage?
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Boeing did not go the Rocketdyne LOX/RP route for the same reasons as Atlas. The US was 20 years behind in the LOX/RP engine technology and the cost to catch up would kill any proposal.
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Couldn't you just restart RS-84 development?
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Propforce - 2/1/2008 2:59 AM
sticksux - 1/1/2008 4:02 PM
Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
Why is it backward to have a H2 first stage?
Because LH stage costs more than equivalent (in lifting capability) kerolox one. Ask yourself - why SpaceX didn't go LH route?
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sticksux - 2/1/2008 10:54 AM
Propforce - 2/1/2008 2:59 AM
sticksux - 1/1/2008 4:02 PM
Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
Why is it backward to have a H2 first stage?
Because LH stage costs more than equivalent (in lifting capability) kerolox one. Ask yourself - why SpaceX didn't go LH route?
An LH2 gas-generator engine (with an ablative nozzle!) is much much simpler and cheaper than an oxygen rich kerolox staged combustion one, if produced in the US. And still has better ISP. I bet that was one of the reasons... There was LH2 experience from the SSME and STME that were clearly assets and competetive advantages.
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The Delta IV decision on LH2/LOx first stage is an example of a decision based on all four aspects of real-world engineering performance-cost-schedule-manufacturing (both recurring and nonrecurring). Anyone who considers it a "step-backward" isn't taking into account all considerations... Delta (under MacDac) traded LH2 and RP1 during the proposal phase and found significant benefits to go with LH2, a big one being the high maturity of the Rocketdyne engine proposal. While I don't want to go into details the Delta and Atlas trades are good examples for those of you guys in school planning on joining us in the real-world... Real trade studies often involve all the aspects of the business.
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TrueGrit - 2/1/2008 5:22 PM
The Delta IV decision on LH2/LOx first stage is an example of a decision based on all four aspects of real-world engineering performance-cost-schedule-manufacturing (both recurring and nonrecurring). Anyone who considers it a "step-backward" isn't taking into account all considerations... Delta (under MacDac) traded LH2 and RP1 during the proposal phase and found significant benefits to go with LH2, a big one being the high maturity of the Rocketdyne engine proposal. While I don't want to go into details the Delta and Atlas trades are good examples for those of you guys in school planning on joining us in the real-world... Real trade studies often involve all the aspects of the business.
Well, LH2 seems to lose on all the other sectors than ISP (which is an important thing ofc).
Ground equipment: pumps, vacuum jacketed piping, hydrogen leaks easily, issue of fuel cost and storability, density, temperature...
Rocket (other than the engine): Low density makes a bulky stage which is expensive to make and transport and is draggy in flight, low temp means you need more insulation, purges, bellows in the lines etc...
But there was experience in all of this.
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TrueGrit - 2/1/2008 11:22 AM
The Delta IV decision on LH2/LOx first stage is an example of a decision based on all four aspects of real-world engineering performance-cost-schedule-manufacturing (both recurring and nonrecurring). Anyone who considers it a "step-backward" isn't taking into account all considerations... Delta (under MacDac) traded LH2 and RP1 during the proposal phase and found significant benefits to go with LH2, a big one being the high maturity of the Rocketdyne engine proposal. While I don't want to go into details the Delta and Atlas trades are good examples for those of you guys in school planning on joining us in the real-world... Real trade studies often involve all the aspects of the business.
The results of this very interesting "trade study" are still pending, but seem, based on recent and projected EELV flight rates, to be leaning right now toward the Atlas staged combustion RP/LOX choice. The results are not final. One monster Atlas pad explosion, Sea Launch style for example, could change everything.
- Ed Kyle
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sticksux - 2/1/2008 2:54 AM
Propforce - 2/1/2008 2:59 AM
sticksux - 1/1/2008 4:02 PM
Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
Why is it backward to have a H2 first stage?
Because LH stage costs more than equivalent (in lifting capability) kerolox one. Ask yourself - why SpaceX didn't go LH route?
... and you think SpaceX is more successful than Delta IV? :laugh:
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TrueGrit - 2/1/2008 9:22 AM
Delta (under MacDac) traded LH2 and RP1 during the proposal phase and found significant benefits to go with LH2, a big one being the high maturity of the Rocketdyne engine proposal.
Very good points, both you and Meiza made.
While I don't want to go into details the Delta and Atlas trades are good examples for those of you guys in school planning on joining us in the real-world... Real trade studies often involve all the aspects of the business.
Well... easy there now. Remember those young kids are the FUTURE of this country !!!
... God help us !!! :laugh:
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meiza - 2/1/2008 10:09 AM
TrueGrit - 2/1/2008 5:22 PM
The Delta IV decision on LH2/LOx first stage is an example of a decision based on all four aspects of real-world engineering performance-cost-schedule-manufacturing (both recurring and nonrecurring). Anyone who considers it a "step-backward" isn't taking into account all considerations... Delta (under MacDac) traded LH2 and RP1 during the proposal phase and found significant benefits to go with LH2, a big one being the high maturity of the Rocketdyne engine proposal. While I don't want to go into details the Delta and Atlas trades are good examples for those of you guys in school planning on joining us in the real-world... Real trade studies often involve all the aspects of the business.
Well, LH2 seems to lose on all the other sectors than ISP (which is an important thing ofc).
Ground equipment: pumps, vacuum jacketed piping, hydrogen leaks easily, issue of fuel cost and storability, density, temperature...
Rocket (other than the engine): Low density makes a bulky stage which is expensive to make and transport and is draggy in flight, low temp means you need more insulation, purges, bellows in the lines etc...
But there was experience in all of this.
Not necessary. LH2 first stage also has additional benefits. First despite a "bulkier" first stage, the GTOW of vehicle is actually less than a RP first stage, so the booster engine require less thrust than a comparable RP engine. You can even afford to loss LH2 propellant with less insulations and still outperform a RP first stage. Common ground propellant equipment. This is a BIG deal and costly.
For those of you comparing Delta IV with Atlas V as a "definitive study" on LH2 vs RP first stage really doesn't give enough credits to a LH2 first stage. Delta IV made a few "mistakes" during its development while Atlas has the advantage of buying an essentially "off-the-shelf" cheap import with plenty of power to spare.
D-IV mistakes include 1) weight growth on the first stage design, and 2) a weak RL10B-2 as 2nd stage engine. The 1st stage weigth growth greatly reduced the take-off thrust-weight of the vehicle. Unfortunately; the thrust requirements of RS-68 was already negotiated with Rocketdyne at the time. It would take an act-of-God, and all his money, to re-negotiate this requirement. As result, the "common" booster core (CBC) were forced to be re-designed to become "individual" booster core in order to take weights out.
D-IV was designed for payloads to go to geosynchronous orbit (GSO), rather than at GTO and wave goodbye to payloads. In fact, it's probably the only vehicle that can take payload to GEO than separate. This way, payloads do not need to impart its propellant energy to circularize orbit, thus prolong payload life in GEO. As result, the RL10B-2 was selected for its high Isp. Unfortunately; its relative low thrust means a poor performance for LEO payloads. Notice Atlas V has the same problem, but they use two RL10A-4 engines (dual engine centaur, DEC) to compensate for this shortfall. So Atlas V has a higher LEO payload capability than D-IV. D-IV attempted to solve this with its advanced upper stage (AUS) upgrade, with a 60K lbf upper stage engine. This would greatly enable D-IV to compete with Atlas and the Arianne in the commercial payload sector. But that funding got squashed along with the dot.com constellation business plan.
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Propforce - 2/1/2008 8:15 PM
sticksux - 2/1/2008 2:54 AM
Propforce - 2/1/2008 2:59 AM
sticksux - 1/1/2008 4:02 PM
Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
Why is it backward to have a H2 first stage?
Because LH stage costs more than equivalent (in lifting capability) kerolox one. Ask yourself - why SpaceX didn't go LH route?
... and you think SpaceX is more successful than Delta IV? :laugh:
No, I think SpaceX is much more cost-averse than Boeing -> it is more likely to pick more economical solution, by necessity.
If you want to compare apples to apples, compare Atlas V to Delta IV. Atlas is almost twice as cheap (!).
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sticksux - 2/1/2008 2:52 PM
Propforce - 2/1/2008 8:15 PM
sticksux - 2/1/2008 2:54 AM
Propforce - 2/1/2008 2:59 AM
sticksux - 1/1/2008 4:02 PM
Why it was decided to use LH+LOX first stage for Delta IV? It seems to be a step backwards from Delta II and III.
Why is it backward to have a H2 first stage?
Because LH stage costs more than equivalent (in lifting capability) kerolox one. Ask yourself - why SpaceX didn't go LH route?
... and you think SpaceX is more successful than Delta IV? :laugh:
No, I think SpaceX is much more cost-averse than Boeing -> it is more likely to pick more economical solution, by necessity.
SpaceX launch failure rate is 100%, success rate is 0%.
If you want to compare apples to apples, compare Atlas V to Delta IV. Atlas is almost twice as cheap (!).
How did you arrive at that number? Cost per lbm of payload?
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I hear you Propforce but I do hope Space X is successful long term as we can always use another launch option especially one that is funded by the private sector. What I worry about with Space X is if they don't get real lucky they will end up spending a lot more time and money than they plan to and being a private venture the money could dry up very quickly. I hope Mr. Musk hangs in there because even though launching rockets looks easy the reality is our rocket technology today is layered with a lot of unknowns and probabilities. But I understand what he is trying to do....better to be early to the party than late.
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Just to correct something... The complexity of LH2 systems, particularly on the ground, is overrated. Yes you need vacuum jacketing, but vacuum jacketing is recommended for LOx systems in order to eliminate heat leak. And in return for vacuum jacketing you eliminate the transfer pumps. That's because unlike heavy LOx or RP lighter LH2 can be transfer via pressure alone... Vaporizor, storage tank, lines, and flow control valve that's all. And yes you need higher spec valves, but most LOx designs work fine at LH2 temps with little modification. And in return there is no need to be concerned with storage vessel heaters in order assure your propellant doesn't get too cold during a freeze before a launch. In all there's a favorable trade by eliminating the RP storage system completely, and increasing the size of the upperstage system.
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Propforce - 2/1/2008 11:57 PM
sticksux - 2/1/2008 2:52 PM
Propforce - 2/1/2008 8:15 PM
sticksux - 2/1/2008 2:54 AM
Propforce - 2/1/2008 2:59 AM
Why is it backward to have a H2 first stage?
Because LH stage costs more than equivalent (in lifting capability) kerolox one. Ask yourself - why SpaceX didn't go LH route?
... and you think SpaceX is more successful than Delta IV? :laugh:
No, I think SpaceX is much more cost-averse than Boeing -> it is more likely to pick more economical solution, by necessity.
SpaceX launch failure rate is 100%, success rate is 0%.
You miss the point. I am saying than when company has no monetary backing from DOD, it *has to* pick the most economical solutions, not the high-tech fancy ones.
If you want to compare apples to apples, compare Atlas V to Delta IV. Atlas is almost twice as cheap (!).
How did you arrive at that number? Cost per lbm of payload?
Yes. It was mentioned on this forum. LV and Services Cost ($FY06M):
A 551 $180 - ~21 ton to LEO, 8.2 ton to GTO
D IVH $475 - ~25 ton to LEO, 10.8 ton to GTO
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TrueGrit - 3/1/2008 5:14 AM
Just to correct something... The complexity of LH2 systems, particularly on the ground, is overrated. Yes you need vacuum jacketing, but vacuum jacketing is recommended for LOx systems in order to eliminate heat leak. And in return for vacuum jacketing you eliminate the transfer pumps. That's because unlike heavy LOx or RP lighter LH2 can be transfer via pressure alone...
Why are you comparing LOX with LH? You should compare _kerosene_ with LH, since LOX is used in both cases anyway. No difference in LOX piping. The difference is in fuel piping.
...is no need to be concerned with storage vessel heaters in order assure your propellant doesn't get too cold during a freeze before a launch.
Sure, you don't need heaters for LH. You need BIG FRIDGE instead, though, and need to run it all the time (say, even during launch delay), which has to be much more expensive. And what is done during month-long launch delay (like current Shuttle one)? I suppose they run cryo cooler all this time?
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Propforce - 2/1/2008 8:43 PM
LH2 first stage also has additional benefits. First despite a "bulkier" first stage, the GTOW of vehicle is actually less than a RP first stage, so the booster engine require less thrust than a comparable RP engine.
This is true. You need bigger kerolox engine for the same LEO capability. For example, RD-180 has approx. ~1.25 higher thrust than RS-68, and therefore can be considered "equivalent".
We should always think in terms of $$$, not newtons, if we want to have affordable access to space. I am trying to estimate cost of US kerolox engine. Russians build and sell RD-170 for $7.5 million apiece to Ukrainians. They also build "half of RD-170" engine - RD-180 - for $10 million and sell it to ULA. So, it is already sold at approx x3 profit! RS-68 costs $14 million. I'd hazard to guess than American engine in RD-180 ballpark should cost no more than $25 million apiece.
You can even afford to loss LH2 propellant with less insulations and still outperform a RP first stage.
I very much doubt you can have less insulation on your LH tank that on kerosene tank, as kerosene tank does not need any insulation!
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sticksux - 3/1/2008 4:43 AM
Sure, you don't need heaters for LH. You need BIG FRIDGE instead, though, and need to run it all the time (say, even during launch delay), which has to be much more expensive. And what is done during month-long launch delay (like current Shuttle one)? I suppose they run cryo cooler all this time?
LH2 ground systems do not have refrigerant systems. The insulation on the tanks keep the LH2 cold. Only a little venting is required.
Cryotanks at the launch sites are not allowed to go dry. They are replenished even during long stand downs (years)
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sticksux - 3/1/2008 6:11 AM
I very much doubt you can have less insulation on your LH tank that on kerosene tank, as kerosene tank does not need any insulation!
Incorrect. Some vehicles require the RP-1 to be at certain temps
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Jim - 3/1/2008 11:41 AM
sticksux - 3/1/2008 4:43 AM
Sure, you don't need heaters for LH. You need BIG FRIDGE instead, though, and need to run it all the time (say, even during launch delay), which has to be much more expensive. And what is done during month-long launch delay (like current Shuttle one)? I suppose they run cryo cooler all this time?
LH2 ground systems do not have refrigerant systems. The insulation on the tanks keep the LH2 cold. Only a little venting is required.
Cryotanks at the launch sites are not allowed to go dry. They are replenished even during long stand downs (years)
Didn't know that. So, when/where LH is actually produced (where it becomes L - liquid?).
How much LH boils off in a month?
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Jim - 3/1/2008 11:43 AM
sticksux - 3/1/2008 6:11 AM
I very much doubt you can have less insulation on your LH tank that on kerosene tank, as kerosene tank does not need any insulation!
Incorrect. Some vehicles require the RP-1 to be at certain temps
My main point still stands - LH tank needs more insulation, not less.
Heating/cooling kerosene is for squeezing last 0.3% of engine performance, right? Is it really worth the trouble - I mean - does it make $/kg to orbit lower?
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sticksux - 3/1/2008 7:35 AM
Didn't know that. So, when/where LH is actually produced (where it becomes L - liquid?).
How much LH boils off in a month?
Louisana
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What I'm saying is that the cost for ground system cryo lines and valves is minor... And technically LH2 is not that different than LOx. The trade is in eliminating the RP system completely vs. enlarging the upperstage LH2 system that was already going to be there. Buying and maintaining large, simple, common LH2 storage/transfer system vs. buying and maintaining a smaller LH2 storage/transfer system aswell as a more complex large RP system.
The trade isn't a clear cut as some of you like to make it out... It's is actually a very close trade. And that's why you don't see a clear cut winner... Shuttle and Delta IV are LH2, while Atlas V and Russian systems are RP. That big differnce maker for Delta IV was the high maturity of the Rocketdyne engine, which lowered development costs tremendously compared with starting with something fairly imature.
Anyone that's been around large scale LOx or LN2 stoarge will be familar with the liquid gas trucks that come and go during all times of the night and day... Those same trucks also carry LH2 and LHe. LH2 and LHe, just like LOx and LN2, are produced in a large plant that takes in gas, seperates it into it's respective components, and sells off the result in either gas or liquid form. Look up Praxair and Air Liquide for more information...
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TrueGrit - 3/1/2008 4:24 PM
maintaining large, simple, common LH2 storage/transfer system vs. buying and maintaining a smaller LH2 storage/transfer system aswell as a more complex large RP system.
I don't understand. RP system will be more complex than similarly-sized LH system? Why?
And that's why you don't see a clear cut winner... Shuttle and Delta IV are LH2, while Atlas V and Russian systems are RP.
I do see clear cut winners here, as soon as I recall $/kg for these vehicles.
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sticksux - 3/1/2008 3:39 AM
A 551 $180 - ~21 ton to LEO, 8.2 ton to GTO
D IVH $475 - ~25 ton to LEO, 10.8 ton to GTO
I'm not sure I believe that Atlas price/cost. In fact, I'm almost certain that I don't believe it. It probably costs almost that much per launch just for the facility maintenance contract. Then again, I don't know if I believe the Delta number either.
- Ed Kyle
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$190M (not $180M) was the advertised (by NASA) value of the Juno contract.
http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=10071&posts=15&start=1
Believe but you will, but if you think ULA would sell a vehicle at a loss, you're mistaken.
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sticksux - 3/1/2008 1:39 AM
You miss the point. I am saying than when company has no monetary backing from DOD, it *has to* pick the most economical solutions, not the high-tech fancy ones.
Delta IV was built with 80% internally funded money.
Cost is not a clear driver to discriminate between LH2 and RP first stage.
ps- you don't think SpaceX is building Falcon 9 with government funded money?
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sticksux - 3/1/2008 3:11 AM
We should always think in terms of $$$, not newtons, if we want to have affordable access to space. I am trying to estimate cost of US kerolox engine. Russians build and sell RD-170 for $7.5 million apiece to Ukrainians. They also build "half of RD-170" engine - RD-180 - for $10 million and sell it to ULA. So, it is already sold at approx x3 profit! RS-68 costs $14 million. I'd hazard to guess than American engine in RD-180 ballpark should cost no more than $25 million apiece.
You know, according to your logic we should just out-sourced to the Russians or the Chinese to launch our national security payloads. It would be a lot cheaper this way.
For commercial payloads, both LM and Boeing steer their commercial customers to their Russian launchers anyway.
I very much doubt you can have less insulation on your LH tank that on kerosene tank, as kerosene tank does not need any insulation!
Insulations are cheap. Not a discriminator. Bigger tanks are minor cost increase. Also not a discriminator.
Pay someone to develop you a brand new U.S. made staged combustion cycle engine, either RP or LH2. EXPENSIVE !!!
RS-68 costs $14 million
May I ask where does this number come from ?
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Ok lets review again...
LH2 storage/transfer = tank, lines, vaporizor, and valves
RP storage/transfer = tank, lines, pump, and valves
LH2 system only has valves that move vs. RP which will also have a speciality built remote operated pump with associated control unit.
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sticksux - 3/1/2008 9:05 AM
TrueGrit - 3/1/2008 4:24 PM
maintaining large, simple, common LH2 storage/transfer system vs. buying and maintaining a smaller LH2 storage/transfer system aswell as a more complex large RP system.
I don't understand. RP system will be more complex than similarly-sized LH system? Why?
He's saying it's cheaper to build one common propellant system than building two on the ground.
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Don't know where you got your performance numbers... But just looking at the ULA COTS RFI they don't come close:
DIV M40 - 8400kg to ISS - 4300kg to GTO - 1100 kg to GEO
AV 401 - 8400kg to ISS - 4900kg to GTO - N/A to GEO
DIV M54 - 12800kg to ISS - 7000kg to GTO - 2700kg to GEO
AV 551 - 16100kg to ISS - 8700kg to GTO - 3900kg to GEO
DIV Hvy - 21500kg to ISS - 12900kg to GTO - 6100kg to GEO
And DIV Heavy is undergoing an engine upgrade that will extend the capability to 27,000 kg to ISS, and once available the upgraded engine can be applied to the Medium CBCs.
As for costs... How many times do we have to say it. You can't simply pull some number off a press release and assume it's a comparible cost. Each contract is filled with lots of mission specific work in the form of multiple trajectory studies, payload integration, oversight support, and vehilce mission mods that all are signifcant cost factors.
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WHAP - 3/1/2008 3:23 PM
$190M (not $180M) was the advertised (by NASA) value of the Juno contract.
http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=10071&posts=15&start=1
Believe but you will, but if you think ULA would sell a vehicle at a loss, you're mistaken.
The best explanation of true EELV costs was given a couple of years ago by Jim McAleese in the following Space News article.
http://www.space.com/spacenews/archive05/McAleese_112805.html
In it, Jim wrote that "ultimately each EELV launch has an average unit-procurement cost of $226 million over the 137 total planned EELV launches. The average unit cost is $232 million per launch, once the $834 million of U.S. Air Force research development, testing and evaluation expenses also are allocated over the 137 total planned launches through 2020".
He also noted that "total EELV sustainment payments from 2004-2020 average $818 million per year on a straight-line basis". The $818 million annual subsidies are "infrastructure sustainment payments". The launches costs extra. Note that there were only five EELV launches this year - that's $164 million per launch even before the launch vehicle costs are added!
Bottom line - there is no way that an Atlas 551 only costs the taxpayers $190 million. It might only cost NASA $190 million, but not the taxpayers.
- Ed Kyle
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edkyle99 - 3/1/2008 6:41 PM
The best explanation of true EELV costs was given a couple of years ago by Jim McAleese in the following Space News article.
http://www.space.com/spacenews/archive05/McAleese_112805.html
In it, Jim wrote that "ultimately each EELV launch has an average unit-procurement cost of $226 million over the 137 total planned EELV launches. The average unit cost is $232 million per launch, once the $834 million of U.S. Air Force research development, testing and evaluation expenses also are allocated over the 137 total planned launches through 2020".
He also noted that "total EELV sustainment payments from 2004-2020 average $818 million per year on a straight-line basis". The $818 million annual subsidies are "infrastructure sustainment payments". The launches costs extra. Note that there were only five EELV launches this year - that's $164 million per launch even before the launch vehicle costs are added!
Bottom line - there is no way that an Atlas 551 only costs the taxpayers $190 million. It might only cost NASA $190 million, but not the taxpayers.
- Ed Kyle
Define infrastructure. My understanding is that the sustainment payment includes money to companies like Aerospace, which have huge contracts with the government. Say what you will, but those costs are not part of the true launch vehicle costs.
I'm struggling with the numbers: If the government gave $1 billion (total) to LM and BA initially, then $834 million for R&D, plus $818 million per year for 17 years, that totals $15.74 billion. Over 137 vehicles, I get $115 million per vehicle. McAleese comes up with a number double that. Where's my math error?
Of course, none of this is relevant to the original post, which was comparing the price of an Atlas V to a Delta IV. All of the supposed costs in the article apply to both vehicles. You may not believe the price numbers, but both vehicles benefit equally from government support.
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TrueGrit - 3/1/2008 3:10 PM
Ok lets review again...
LH2 storage/transfer = tank, lines, vaporizor, and valves
RP storage/transfer = tank, lines, pump, and valves
LH2 system only has valves that move vs. RP which will also have a speciality built remote operated pump with associated control unit.
Atlas V has pumps for the RP system, but they're nowhere near as complicated as you make them sound. The RP system is pretty straightforward, and doesn't require a lot of the maintenance required of the cryo systems.
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The RD-180 is an extremely complex engine. Its performance at sea level approaches the theoretical limit. Payload performance requires that it do so.
Because of the higher Isp of LH2/LOX, the RS-68 can be much simpler and thus cheaper.
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WHAP - 3/1/2008 10:00 PM
Define infrastructure. My understanding is that the sustainment payment includes money to companies like Aerospace, which have huge contracts with the government. Say what you will, but those costs are not part of the true launch vehicle costs.
The article spelled it out quite clearly.
"Specifically, the EELV launch capability contracts include virtually all the costs of launch vehicle development, production and launch, other than materials and direct labor for EELV launch vehicle assembly.
Even day-to-day program management, and actual launch costs themselves, are excluded from the EELV launch service contracts. Instead, all of the recurring overhead costs, plus all infrastructure costs have been consolidated into the sole-source EELV launch capability contracts. The Air Force is funding virtually all of the annual working capital, long-term capital expenditures and recurring overhead for all of United Launch Alliance's EELV launches, including both commercial customers and NASA launches as well."
I don't know what you mean by "true launch vehicle costs". What matters in the long run is total program costs. Dividing that by the number of launches gives an average launch vehicle mission cost, which to me is the closest thing one can get to a per-launch cost estimate.
I'm struggling with the numbers: If the government gave $1 billion (total) to LM and BA initially, then $834 million for R&D, plus $818 million per year for 17 years, that totals $15.74 billion. Over 137 vehicles, I get $115 million per vehicle. McAleese comes up with a number double that. Where's my math error?
The total program cost is projected to be $31.8 billion for 137 launches, an average of $232.12 million per launch. Nearly $14 billion of the total is for "launch infrastructure sustainment". That averages out to about $823 million per year, or $102.2 million per launch if all 137 launches occur during the 17 year period - a rate of 8 per year. But the EELV program hasn't done better than 5 per year yet.
Of course, none of this is relevant to the original post, which was comparing the price of an Atlas V to a Delta IV. All of the supposed costs in the article apply to both vehicles. You may not believe the price numbers, but both vehicles benefit equally from government support.
Prior to the formation of ULA, Boeing was getting a bigger slice of the "sustainment" money than Lockheed Martin. Higher fixed costs, presumably. Delta flies from massive Slick 6 while Atlas will go from much more basic SLC 3E, for example. The Heavy needs more infrastructure.
- Ed Kyle
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WHAP - 3/1/2008 11:03 PM
TrueGrit - 3/1/2008 3:10 PM
Ok lets review again...
LH2 storage/transfer = tank, lines, vaporizor, and valves
RP storage/transfer = tank, lines, pump, and valves
LH2 system only has valves that move vs. RP which will also have a speciality built remote operated pump with associated control unit.
Atlas V has pumps for the RP system, but they're nowhere near as complicated as you make them sound. The RP system is pretty straightforward, and doesn't require a lot of the maintenance required of the cryo systems.
I don't think he's saying that an RP1 system is more complex/more maintenance intensive than a LH2 system. He's saying that with an LH2 first stage you don't need a separate RP1 system at all, you can just have a single LH2 system for the entire vehicle. The combined complexity of having an LH2 system for the second stage and a separate RP1 system for first stage is greater than just having one big LH2 system for the entire thing.
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With apologies to my off-line colleagues:
When was the last time you saw an RP booster lifting off 2/3rds charred?
Anything you don't have to purge with an increasingly scarce commodity and worry about vaporizing into a combustible mixture with the atmosphere sounds like a better plan to me. For me:
Booster: RP > LH on grounds of usability
Exoatmospheric: LH > all on grounds of performance
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Antares - 3/1/2008 10:35 PM
With apologies to my off-line colleagues:
When was the last time you saw an RP booster lifting off 2/3rds charred?
Smoke & Fire baby !!! :laugh:
Isn't it why we're all into rockets?
Diversity is the mother of all inventions !!! How boring would it be if we only have same type of rockets?
You have to admit, the poor slobs who have to clean & work on the pad post launch would prefer a H2 first stage.
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TrueGrit - 3/1/2008 10:28 PM
Don't know where you got your performance numbers... But just looking at the ULA COTS RFI they don't come close:
DIV M40 - 8400kg to ISS - 4300kg to GTO - 1100 kg to GEO
AV 401 - 8400kg to ISS - 4900kg to GTO - N/A to GEO
DIV M54 - 12800kg to ISS - 7000kg to GTO - 2700kg to GEO
AV 551 - 16100kg to ISS - 8700kg to GTO - 3900kg to GEO
DIV Hvy - 21500kg to ISS - 12900kg to GTO - 6100kg to GEO
As for costs... How many times do we have to say it. You can't simply pull some number off a press release and assume it's a comparible cost.
Ok, I am taking above performance figures. If you think my cost figures are wrong, what data do you have on flyaway cost of Atlas and Delta?
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Propforce - 3/1/2008 10:07 PM
You know, according to your logic we should just out-sourced to the Russians or the Chinese to launch our national security payloads. It would be a lot cheaper this way.
I am saying that we should make our LVs more competitive, or at least try to. Otherwise we will never be a spacefaring species.
I very much doubt you can have less insulation on your LH tank that on kerosene tank, as kerosene tank does not need any insulation!
Insulations are cheap. Not a discriminator. Bigger tanks are minor cost increase. Also not a discriminator.
RP: smaller tank with no insulation
LH: bigger tank with insulation.
For me it's obvious that second one costs and weighs more, needs bigger machines to make, bigger barges to transport, insulation requires protection in transportation (danger! woodpeckers!) and on pad (hailstorm?).
Right now, NASA has Shuttle delayed because of cryo-induced problems in LH tank. How much does it cost to have Shuttle delayed by a month?
Pay someone to develop you a brand new U.S. made staged combustion cycle engine, either RP or LH2. EXPENSIVE!!!
Yes. Something like $1 bn, but it needs to be done only once.
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sticksux - 4/1/2008 5:52 AM
RP: smaller tank with no insulation
LH: bigger tank with insulation.
For me it's obvious that second one costs and weighs more, needs bigger machines to make, bigger barges to transport, insulation requires protection in transportation (danger! woodpeckers!) and on pad (hailstorm?).
Assuming the larger tank tank costs more than using an RP engine instead of an LH engine.
Since the Atlas engine is made in russia and the Delta engine is made in the US we can not do a fair comparison on price.
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kevin-rf - 4/1/2008 5:19 AM
Since the Atlas engine is made in russia and the Delta engine is made in the US we can not do a fair comparison on price.
Yes we can. This is the Free Trade era. Japan makes cars. China, computers. Korea, flat screens. Russia, kerosene rocket engines.
- Ed Kyle
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There is a very minimal amount of extra cost do to a larger tank... Material costs aren't a big factor and the majority of costs are in manufacturing. The processes and time required to manufacture a 5m tank is equivelent to that of a 3m tank.
As for the launch fireball... That is more of an unintented consequence of how PWR and Delta does the RS-68 start sequence. Call if a lesson learned... RS-68 is completely new and you wouldn't expect the first Block to be perfect. PWR with Delta and NASA assistance have developed start sequence and design changes to mitigate the fireball. In the end the 5-engined Ares V will have less of a fireball the the 3-engined Shuttle.
Once again, and please listen so I don't have to repeat myself another time... I'm not saying and LH2 stage is clearly more supperior than and RP stage... What I'm saying is that the trade is a lot closer than advocates of either would make it seem. A very big facotr in that decision is the available technology. And dispite your feelings on the issue US LH2 engine technology is head-and-shoulders above that of RP. And the EELV program specificaly said they prefered a US solution... It's not a minor thing considering it is a military system that the DoD doesn't want a critical spy/cummications satelitte held hostage to some Russia political desicion. Hard enough dealing with the wavering of US political desicion makers...
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edkyle99 - 4/1/2008 9:01 AM
kevin-rf - 4/1/2008 5:19 AM
Since the Atlas engine is made in russia and the Delta engine is made in the US we can not do a fair comparison on price.
Yes we can. This is the Free Trade era. Japan makes cars. China, computers. Korea, flat screens. Russia, kerosene rocket engines.
- Ed Kyle
Agreed. The Russians have become excellent capitalists. :)
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Given the current state of the dollar, and the money Russia is raking in from oil and gas, if one were to request bids right now for say a long-term 30 engine contract, I have to wonder if the RS-68 would end up being cheaper per unit.
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sticksux - 4/1/2008 4:33 AM
Ok, I am taking above performance figures. If you think my cost figures are wrong, what data do you have on flyaway cost of Atlas and Delta?
Some years back, for data for a summer research project, I was able to call LMA and get a catalog price for Athenas. It's a ROM, but at least it's a primary source. I think the same person is still working for ULA and listed on web sites.
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Tru Grit said something interesting over in one of the Ares-1 threads : http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=11584&start=61&posts=72
And finally... By the time the administrators successor gets around to begging from congressional money for the heavy booster ULA will have moved on to RL-60 and ACES. They'll also conveniently forget this speech when the same successor comes asking for support because NASA is stuck in LEO again.
Is ACES really moving foward at ULA and more interesting will it also replace Centuar on the ULA Atlas V?
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I have some questions about the RS-68 start sequence. I was watching a movie of the last Delta IV flight. Since it was at night, you could see some interesting thing happening.
1) The engine pump starts at T-15 seconds. This seems like along time before start up, my guess is that there is a chill sequence before start.
2) The engine nozzles appear to be below the three holes in the flame chute. Does anyone have close up photos on how the rocket mounts on the pad? It would be interesting to see detail on this
3) When the turbo pump starts it is only pumping hydrogen. The oxygen is shut off, I would be curious to see how they handle the O2 during that part of the startup. Most high-pressure pumps are very unhappy running with the flow shut off.
4) At T-5 is when you get the fireball; I assume that is when the igniters are turned on.
5) You notice also that the tapered aft end of the first stage allows the hydrogen to vent upward quickly
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1. It starts later than T-15,
2.Each CBC sits on 4 posts of the Launch Mate Unit which in turn sits on the launch table
http://www.ulalaunch.com/index_gallery.html
3. It is pumping both LO2 and LH2 at the sametime
4. No, that is caused by the burn off igniters (ROFIs) which are like the shuttle's and are lit at L-15
5. Inconsequential. That is just the aeroshell of the engine section. The H2 is going to rise no matter what.
Will have more info on Monday.
Some of the real details will touch upon propriety and ITAR issue.
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The pumps start about T-8.5 seconds (right about when the TCSR takes control of the vehicle), this can be seen as the smaller flames coming from the heat exchanger and roll control nozzles on either side of the aft cone. This is the gas generator exhaust and indicates that the gas generator has begun operating and producing gas to run the turbines.
The fuel rich exhaust from the turbopumps/gas generator is ignited by the radial outward-firing ignitors (ROFIs), which as Jim stated are lit off at T-15.
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Thanks for the great pictures!
Upon further review I realized that the RS-68 uses two turbo-pumps, this would give a lot more flexibility in the start sequence. I guess I was thinking of the Spacex Merlin with its single turbo-pump when I asked the question.
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Big Al - 2/2/2008 2:55 PM
Thanks for the great pictures!
Upon further review I realized that the RS-68 uses two turbo-pumps, this would give a lot more flexibility in the start sequence. I guess I was thinking of the Spacex Merlin with its single turbo-pump when I asked the question.
It uses one (corection) gas generator. Even, vehicles with two, like the shuttle, would start them at the same time. An engine can not run unbalanced
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I bet my timing was off too. I have been watching the video over at Spaceflight Now.com and as soon as The Fat Man Doing The Countdown (he looks just like me!) says “15” you see the red glow of hydrogen deep within the flame chute
Another comment on the taper of the aft end of the rocket is that the heat blankets on both the Delta and Atlas have a very tight fit. I’m sure this is good for aerodynamics, but I would think it would trap all the heat generated buy the upper part of the engine, but it does work and is widely used.
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Big Al - 2/2/2008 3:37 PM
1. I bet my timing was off too. I have been watching the video over at Spaceflight Now.com and as soon as The Fat Man Doing The Countdown (he looks just like me!) says “15” you see the red glow of hydrogen deep within the flame chute
2. Another comment on the taper of the aft end of the rocket is that the heat blankets on both the Delta and Atlas have a very tight fit. I’m sure this is good for aerodynamics, but I would think it would trap all the heat generated buy the upper part of the engine, but it does work and is widely used.
1. That is from the ROFI's
2. If Spacex had such a structure,they are purged with conditioned air prelaunch, they wouldn't have had the corrosion on the first launch. Most launch vehicles enclose the upper engine sections, not just the EELV's. Thor, Atlas I & II, Jupiter, Redstone, etc
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Are you sure about the single turbo pump?
Wikipedia in their write up states that it has two pumps. Also over at Spaceandtech.com they have a very nice CAD rendering of the RS-68 with a good view of the top of the engine. It shows two pumps. Each pump exhausts on the opposite side of the engine. One of them, which I think is the hydrogen pump because it looks like it has bigger scroll housing, has a high pressure nozzle on it which I assume is used for roll control. The other pump looks like it uses a lower pressure exhaust system.
The over all design looks like a classical rocket engine…except it burns hydrogen….and its on the first stage of a rocket…..and it has an ablative lower nozzle…
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I corrected my post to single gas generator
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Over on the Ares I TO thread it was mentioned that the RS-68 uses quite a lot of He and that was a potential problem due to the limited supply of He. Can someone explain exactly what the He is used for and why can't another inert gas be used, say N?
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Well N freezes at LH temps while He stays a gas. So purging the system with cold He gas gets rid of all the air with might freeze just where you don't want it, but I'm no expert, best check with them.
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nacnud - 23/4/2008 4:26 PM
Well N freezes at LH temps while He stays a gas. So purging the system with cold He gas gets rid of all the air with might freeze just where you don't want it, but I'm no expert, best check with them.
You very correct on that description. Frozen air in the RL-10 was the cause of 2 failures for the Atlas II program back in the late 80's.
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Rocket Rancher - 23/4/2008 4:40 PM
nacnud - 23/4/2008 4:26 PM
Well N freezes at LH temps while He stays a gas. So purging the system with cold He gas gets rid of all the air with might freeze just where you don't want it, but I'm no expert, best check with them.
You very correct on that description. Frozen air in the RL-10 was the cause of 2 failures for the Atlas II program back in the late 80's.
early 90's
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This may not be as sophisticated as the other questions, but what is a good book as an overview of the Delta IV system and its development? And if any other books of other launchers if know.
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Jim - 23/4/2008 10:42 PM
Rocket Rancher - 23/4/2008 4:40 PM
nacnud - 23/4/2008 4:26 PM
Well N freezes at LH temps while He stays a gas. So purging the system with cold He gas gets rid of all the air with might freeze just where you don't want it, but I'm no expert, best check with them.
You very correct on that description. Frozen air in the RL-10 was the cause of 2 failures for the Atlas II program back in the late 80's.
early 90's
Atlas II never failed. It has been Atlas I.
Analyst
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Ronsmytheiii - 23/4/2008 5:43 PM
This may not be as sophisticated as the other questions, but what is a good book as an overview of the Delta IV system and its development? And if any other books of other launchers if know.
Rocketdyne had an internal and PR magazine called Threshold. There was one called "First Flight Edition" for RS-68 that gave a very good overview of the development of the engine. I can't find it on Google. Maybe a request to Canoga to get a soft copy of it.
Doubtful there's a book on this, exactly. I suggest looking at AIAA papers over the 1997-2004 and more recent time period.
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Antares - 24/4/2008 5:40 PM
Doubtful there's a book on this, exactly. I suggest looking at AIAA papers over the 1997-2004 and more recent time period.
These papers are extremly expensive. Sadly its very often this way in academics and engineering. $30 for a pdf with 15 pages - while I am interested in this stuff, I have to pay my bills too.
Analyst
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Libraries, anyone? Many companies, agencies and universities have subscriptions to pdf-dot-aiaa-dot-org.
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There is an interesting Word document and PowerPoint about the RS-68 and its design philosophy and development here:
http://www.engineeringatboeing.com/data.htm
Look for "Propulsion For the 21st Century"
HTH,
Nick
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Space News carried an op-ed critical of RS-68, but a simpler engine always looks like a better deal.
The processes and time required to manufacture a 5m tank is equivelent to that of a 3m tank. In the end the 5-engined Ares V will have less of a fireball the the 3-engined Shuttle.
Interesting points
Ironically--it was the time period right after Communism's collapse that was dangerous for the investor. Putin put a lid on things, and more investments started coming in. Things are a little less friendly now--but oil and gas are still good investments.
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As for the launch fireball... That is more of an unintented consequence of how PWR and Delta does the RS-68 start sequence. Call if a lesson learned... RS-68 is completely new and you wouldn't expect the first Block to be perfect. PWR with Delta and NASA assistance have developed start sequence and design changes to mitigate the fireball. In the end the 5-engined Ares V will have less of a fireball the the 3-engined Shuttle.
I think this worth some clarification.
The main reason for the "fireball" on the RS-68 and the Delta IV has much to do with how the exhaust duct was designed as well as with the RS-68 itself.
The Delta IV chosed an "enclosed" exhaust duct design over the water suppression approach, such as used on the Shuttle launch pad, for the simple reason of economic. They did not want to spend the money to put in a water-suppression system during the early development days when money was tight (keep in mind that the majority of development cost was paid for by Boeing and not by the government). The enclosed duct satisfied the main concern at the time, which was Ignition Over Pressure (IOP) that could adversly impact payloads acoustic qualification requirements.
The realization of have a pre-mixed detonable hydrogen-air inside the duct was an after-thought. Internal trades were conducted to evaluate the alternatives, including what was proposed for the Shuttle pad at Vandenberg of using a jet engine to pump out the exhausts. Again, the key figure of merit (FOM) was cost. The radially outward flame ignitor (ROFI), a.k.a. the sparklers, used for the Shuttle launch pad, were slected out of pure economics and not that it was the most effective solution. In fact, I doubt that they really make that much difference at all.
The physics problem was equivalent to pumping gas into your oven for ~30 seconds before you light a match to it. You get a big BANG and that pressure wave doesn't just go out the back end, it goes toward the vehicle as well. They had to structurally beef up the engine aft-section as well as adding extra thermal insulation as a result. Since hydrogen gas is lighter than air, some do floated up via the gap between the launch table and the vehicle. So as flame propagates forward, you see a nice flash and burning of these hydrogen next to the vehicle, resulting in the charring of vehicle paints on lift-off.
Another unwelcome result is losing some initial vehicle thrust due to drag. When the 3 engines going on full blast, e.g., 104% power level, it created a 'jet pump' effect inside the duct which created a suction at the based of vehicle. This further reduces vehicle's T/W ratio at take-off.
But this will NOT happen on Ares 1 or V, unless NASA fails to learn this lesson from Delta and use the same enclosed exhaust duct approach. If the vehicle is on an 'open-air' plateform, such as the current Shuttle launch pad, the hydrogen pumped out during engine start will quickly dissipate and not accumulate at the based of vehicle to cause the similar problem.
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While the flame-duct does have some influence on the "fireball" that is release when the Delta IV is started it is not the primary reason. The primary reason is the way the RS68 is started... I've seen simulations of the predicted Ares V start that shows a very similar situation as Delta IV, even with an open duct and water suppression. That's the reason NASA has identified "start sequence" changes for the RS68B program. Heck even the CBC during its stage test had similar problems with "fireballs" despite being multiple stories in the sky on an open test stand. The primary reason for the water on the Shuttle pad is for sound/acoustic suppression due to the two massive SRBs, not fireball control.
The reason for the "fireball" is that RS68, like most LO2/LH2 engines, has a "fuel lead" during the start sequence. This means the first thing the RS68 does when it starts is to dump a heck of a lot of hydrogen...
As the risk of getting off topic... It is this initial dump of hydrogen why the Delta IV, like Shuttle and SSME, has "sparklers". The "sparklers" only try to assure the hydrogen released doesn't poll anywhere. The "sparklers" do this job, and without them there's a risk of the hydrogen pooling and exploding when the engine is ignited. On the test stand they have the same concerns, but utilize propane torches (like the flame stack).
Getting back to the start fireball... This initial release of hydrogen is lit by the "sparklers" which adds to its buoyancy... It's is this "free hydrogen" that floats up and becomes the "fireball" seen prior to each launch. What happens is that once the engine is started the duct aspiration you mentioned start to suck down the air around the vehicle, including any remaining fire around the bottom of the vehicle. So the problem is that the amount of hydrogen released combined with relatively long time it takes to achieve aspiration results in too much fire released.
The NASA RS68B plan to take some conservative margins out of the start sequence to both reduce the "free hydrogen" and quicken the time to reach aspiration. As I said these design changes are known and will reduce the amount of "free hydrogen" of 5 RS68s to that of 3 SSMEs.
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Okay here is a tinker toy question on something that has been bugging me since the DOD is ending the Delta-II 30 day callup for launching GPS birds.
Different Delta-IV cores are used for different Delta-IV configurations. In the case of the GPS, where the Delta-IV M is not maxed out can the heavier M+(4,2) and M+(5,4) core (ment for a different launch that is being bumped) be used in zero solid Delta-IV M configuration to launch a GPS bird quickly?
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Okay here is a tinker toy question on something that has been bugging me since the DOD is ending the Delta-II 30 day callup for launching GPS birds.
Different Delta-IV cores are used for different Delta-IV configurations. In the case of the GPS, where the Delta-IV M is not maxed out can the heavier M+(4,2) and M+(5,4) core (ment for a different launch that is being bumped) be used in zero solid Delta-IV M configuration to launch a GPS bird quickly?
Not the M+(5,4) core, it has a 5m interstage. Interstage is part of the core.
Physical at M+(4,2) core could be used but all the analyses would have to be done before hand
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The acoustic environments that solids put out could easily prevent a payload from moving off of an all-liquid stack. G's would be higher too since there's about 2/3rds more thrust while the solids are on. That could even present a problem for max-Q violations on the rocket, not just the payload.
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The acoustic environments that solids put out could easily prevent a payload from moving off of an all-liquid stack. G's would be higher too since there's about 2/3rds more thrust while the solids are on. That could even present a problem for max-Q violations on the rocket, not just the payload.
I think he is asking if it is possible to fly a Delta IV M+ core (which is structurally different from the Medium core) without solids if the need arises.
I wonder now, with the RS-68 performance boost in the cards (for Ares V), is it possible that now they can do away with having separate, lighter parts specifically for the Medium and just use the M+(4,2) core, reducing cost?
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Very good, sir. That precisely is or has been investigated. The performance boost is not for Ares V. Ares V is using the performance boost from the Delta IV Heavy Upgrade.
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GPS IIF's are currently planned to be launched on a M+ with 2 GEMs, not a Medium. And would be by default capable of living with the environments from the 2 GEMs. As for the launch vehicle... The GPS IIF SV is similar in mass to the GOES SV and as such would result in similar loads.
As for flying a M+ CBC without GEMs... The M+ structure is heavier and stronger than the Medium and therefore would be capable of flying without GEMs if the situation called for it. The question is a matter of performance... The M+ is multiple thousands of lbs heavier than a Medium.
As for use of the upgraded RS68A being developed for the USAF... There are specific Heavy EELV performance needs driving its development. That engine once developed is getting a close look at for "single stick" Delta IVs. Part of that trade involves more commonality across the range. But is being traded against other considerations, including performance...
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What would the performance of a D-IV Medium with an RS-68A look like? Assuming everything stays the same as it is now except for the engine.
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As it has been said "rockets are not legos"... Nothing is completely unchanged with the RS-68A incorporated. Everything in the rocket will get a second look and some minor changes are needed at a minimum.
The Delta IV perfromance with the upgraded RS-68A is detailed in back of the mission planners guide, which is available on the ulalaunch.com website. Assuming the RS-68A full thrust option is pursued performance to a 19,323 nmi x 100 nmi minimum at 27.0 deg orbit is:
Medium: Current 9.4 klb - Upgrade 10.5 klb
M+4,2: Current 13.1 klb - Upgrade 14.0 klb
M+5,2: Current 10.3 klb - Upgrade 11.3 klb
M+5,4: Current 14.1 klb - Upgrade 15.1 klb
Heavy: Current 28.4 klb - Upgrade 32.1 klb
In the same section there is detailed proposals for additional M+ configurations. The most ready to go is the M+4,4. But also M+5,6 and M+5,8 configurations. And has a more upto date version of the potential future Heavy growth chart.
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Cool, thanks. Didn't think to look in the PPG.
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'Nuther question:
On today's launch before the scrub one of the cameras on the east LPT was showing "LOW PRESSURE". Also saw this too on a couple other videos. What pressure is this referring to?
Nick
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'Nuther question:
On today's launch before the scrub one of the cameras on the east LPT was showing "LOW PRESSURE". Also saw this too on a couple other videos. What pressure is this referring to?
Nick
I suspect the camera's have enviromental enclosures with a positive pressure seal and that is what the message is referring to, i could be wrong on that since i know next to nothing about the GSE in question but i can almost guarantee you its not related to the LV.
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As I understand it currently you have to throttle down anyway so you don't exceed constraints on acceleration and/or dynamic pressure. If
The Heavy doesn't have this issue.
So why does the center core throttle down? Just to allow the outer cores to be jettisoned early? And does that outweigh gravity losses from not running at full thrust?
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Does anyone know what the max q of the heavy is?
Danny Deger
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As I understand it currently you have to throttle down anyway so you don't exceed constraints on acceleration and/or dynamic pressure. If
The Heavy doesn't have this issue.
So why does the center core throttle down? Just to allow the outer cores to be jettisoned early? And does that outweigh gravity losses from not running at full thrust?
The strap on's aren't jettisoned early, they are jettisoned when they burnout.
The core throttles down to reduce propellant consumption, so that it can burn after the strap on's are jettisoned and function as another stage. It is similar to the booster package on the heritage one and half stage Atlas.
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snip
So why does the center core throttle down?
So the outers will use their fuel before the core does.
Just to allow the outer cores to be jettisoned early?
If you realize the outers are empty when jettisoned, the answer it yes.
And does that outweigh gravity losses from not running at full thrust?
It must or they wouldn't do it.
Danny Deger
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If I understand correctly, the DoD had a concern about the nozzle extension of the RL10B-2 being provided by a French company. To mitigate the perceived risks associated with that, DoD funded the purchase of a "large number" of the nozzles and is warehousing them.
I have not found a source indicating the exact (or even approximate) number of nozzles purchased. Does anyone know? Also, given the large inventory is it fair to assume the French manufacturer's production line is "cold", i.e. it could take them awhile to resume production?
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Does anyone know what the max q of the heavy is?
Danny Deger
In the last AIAA JPC, Jim Sponick (VP of Delta) said the max q for DIV-H is down to the low 300s.
He is citing this as 'advantage' because of the benign environment.
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It was recently mentioned that Delta needs very strict alignment of its cores and solids. What causes the need for extreme precision? What bad things would happen if they were only roughly aligned? I'm trying to understand if the same thing would apply to Microcosm/Beal/Truax style, very strong and stiff pressure-fed rockets.
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It was recently mentioned that Delta needs very strict alignment of its cores and solids. What causes the need for extreme precision? What bad things would happen if they were only roughly aligned? I'm trying to understand if the same thing would apply to Microcosm/Beal/Truax style, very strong and stiff pressure-fed rockets.
Off centered thrust from a solid would be inefficient and require the core to counter, causing further inefficiencies.
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So how accurate do the alignments have to be? Fractional degrees?
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Re RL10B-2 nozzles: the supplier is SNECMA, but I'm not familiar with any advance purchase of nozzles. There are lots of whole RL10s, not just nozzles, sitting around due to low flight rate.
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Very good, sir. That precisely is or has been investigated. The performance boost is not for Ares V. Ares V is using the performance boost from the Delta IV Heavy Upgrade.
Well... this will all be OBE soon ;D
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Ronsmytheiii - 23/4/2008 5:43 PM
This may not be as sophisticated as the other questions, but what is a good book as an overview of the Delta IV system and its development? And if any other books of other launchers if know.
Rocketdyne had an internal and PR magazine called Threshold. There was one called "First Flight Edition" for RS-68 that gave a very good overview of the development of the engine. I can't find it on Google. Maybe a request to Canoga to get a soft copy of it.
Doubtful there's a book on this, exactly. I suggest looking at AIAA papers over the 1997-2004 and more recent time period.
Here's the new link to re-named PWR Engineering where you'll find many excellent technical articles on engine designs
http://www.pwrengineering.com/data.htm
Here's the link to article on RS-68, published in 2002
http://www.pwrengineering.com/dataresources/EvolvedExpendableLaunchVehicleSystem-RS-68MainEngineDevelopment.doc
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Drool,... Thanks Propforce
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Looking at the mission booklet for the upcoming WGS launch, it says that only two of the four GEM-60s have TVC, why not have it on all four?
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Looking at the mission booklet for the upcoming WGS launch, it says that only two of the four GEM-60s have TVC, why not have it on all four?
It is simply not necessary.
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In one of the EELV alternative threads, I recall PadRat saying the Delta IV requires its pad to be 'pretty much the most level surface on earth'......
I'm curious why that tolerance is so tight?
Sure, you want a level pad- but the millisecond the bird lifts, guidance is gimballing the engines, yes? So is the critical levelling requirement related to GSE/propellant connections.... cryo settling.... or what?
Advance thanks for your answers!
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In one of the EELV alternative threads, I recall PadRat saying the Delta IV requires its pad to be 'pretty much the most level surface on earth'......
I'm curious why that tolerance is so tight?
Sure, you want a level pad- but the millisecond the bird lifts, guidance is gimballing the engines, yes? So is the critical levelling requirement related to GSE/propellant connections.... cryo settling.... or what?
Advance thanks for your answers!
Not the pad, but the floor in the HIF for stage alignment
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I wonder if any new info on base heating can be gleaned from this latest config with the RS-68 sandwiched between 4 solids?
I've read in other threads that clustering is going to be a major problem? (even without shuttle SRB's)
Beautiful looking launch by the way!
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You mean base heating on a potential NASA HLV? It's too geometry-dependent to translate to a different configuration. I'm sure they instrumented the heck out of the Delta IV boat tail on this launch.
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I wonder if any new info on base heating can be gleaned from this latest config with the RS-68 sandwiched between 4 solids?
I've read in other threads that clustering is going to be a major problem? (even without shuttle SRB's)
Beautiful looking launch by the way!
You're probably thinking using the RS-68s for Ares V where six RS-68 is packaged in between two 5.5 segment of SRBs.
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Pretty sure that this is a Delta II second stage, can anyone confirm?
not any type of Delta hardware
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High Bypass Turbofan by the looks of things.
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High Bypass Turbofan by the looks of things.
Bingo
With a composite fan housing.
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Question:
Can the Delta IV heavy take a payload directly to EML 1? Or any langrange point for that matter? Has this been done before?
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GEO is further delta-v wise, so it should be able to take substantial payloads to L1. As I understand it it would need a long duration mission kit to stay alive for long enough to do the insertion burn, unless the payload did that itself, which would not be a strange thing. If you used a quasi-ballistic trajectory that should require only stationkeeping capability.
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Hypothetical question if I may, suppose DIVH is selected to launch a ISS & BEO Orion. Is it likely the crossfeed option (from booster to core) would happen?
What kind of extra performance could be expected? (assuming 68A) Thanks!
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Too complicated to answer. Really requires a trajectory optimization program.
I don't think it would happen anyway. NASA is going to have to take the Hippocratic oath when it gets serious about humans on EELV.
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Hypothetical question if I may, suppose DIVH is selected to launch a ISS & BEO Orion. Is it likely the crossfeed option (from booster to core) would happen?
What kind of extra performance could be expected? (assuming 68A) Thanks!
Crossfeed isn't needed, but RS-68A with human rating upgrades is needed. The 2009 Aerospace study linked below examined this question. It found that the existing rocket (once upgraded with RS-68A) could do either job, but Delta IV would need some "human rating" upgrades to meet crew safety requirements.
http://www.nasa.gov/pdf/377875main_081109%20Human%20Rated%20Delta%20IV.pdf
The rocket essentially exists. Now, if only there were a spacecraft...
- Ed Kyle
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In addition, for an ISS Orion, a second stage isn't even needed. (bottom of page 14, section 2.2.5) So close yet so far away?
Since J-2X will likely be a reality, who knows?
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When the last shuttle flight has flown and NASA begins a new wish list for heavy payloads to ISS, I wonder if DIVH will see some action?
Any thoughts on this?
Maybe using that shuttle payload carrier mentioned in one of the ULA papers?
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When [...] NASA begins a new wish list [...] I wonder if DIVH will see some action? [...] Maybe using that shuttle payload carrier mentioned in one of the ULA papers?
Apparently NASA would like to develop a new spacecraft, the "Flagship Service Vehicle":
http://nspires.nasaprs.com/external/viewrepositorydocument?cmdocumentid=230989&solicitationId={980D21C5-AF8F-7252-C1BA-507EA54906BB}&viewSolicitationDocument=1
Possibly some flagship missions would end up requiring DIVH lift, though that doesn't seem to be how they are currently envisioned.
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Thanks! I hadn't seen that one.
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When the last shuttle flight has flown and NASA begins a new wish list for heavy payloads to ISS, I wonder if DIVH will see some action?
Any thoughts on this?
Maybe using that shuttle payload carrier mentioned in one of the ULA papers?
There is no need for heavy payloads to ISS. The truss is done. NASA doesn't have plans for anything big. Any "expansion" will be small or inflatable modules.
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There are "supposed" to be 3 Delta-IV Heavies over the next twelve months. Looking through the currently posted launch manifest (on NSF) I don't see any additional Heavies. After the next three fly will it really be years before we see another heavy? Are there any announced Heavy contracts that have not been assigned a launch date?
I also noticed next up will sport the new RS-68a's. Is it really needed for the next flight, or is it being tested for a future flight that needs it?
Edit: Okay, feeling a bit sheepish, missed: 2015 - NRO L-37 - Delta IVH - TBD.
Meaning it will be fourish years before the next Heavy after this batch flys?
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it will be fourish years before the next Heavy after this batch flys?
This speaks to the role EELV-Heavy plays in the overall DoD spacelift acquisition strategy. They only had a few missions in mind that could not be lifted by "standard" EELV-Medium vehicles. Rather than having a separate launch system specifically for these payloads, the EELV-Heavy approach using three "standard" EELV cores was (quite rightly) seen as cost effective, especially because there are so few payloads in this class.
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Will there be 2 versions of RS-68 going forward? (RS-68 & RS-68A)?
Or, are the 68A's the only kind being produced & when the old 68's are used up that's it?
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Jim correct me if I am wrong, but didn't Boeing pre-ULA buy a large number of RS-68's that need to get used up before additional RS-68A production is needed? The only reason we have RS-68A's today is it was cheaper to do a batch of them than the other options available for increasing needed Delta IV Heavy performance.
Has ULA release the new RS-68A performance number for non heavy Delta IV's?
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I would also appreciate correction if wrong, but isn't the expectation that once RS-68A flies successfully, all subsequent Delta flights will use RS-68A? (Perhaps there is a possibility that RS-68 inventory could be retrofitted?)
Further, RS-68A will allow a reduction in the number of different CBC variants, and some Delta vehicles will thus provide the same capability but in a different way, i.e. with a more efficient engine but also with a more massive CBC.
Fascinatingly, the February 9, 2010 press release from PWR that describes RS-68A status is no longer listed as being available from their archive. PRNewsWire has it here:
http://www.prnewswire.com/news-releases/pratt--whitney-rocketdyne-begins-certification-testing-of-the-rs-68a-rocket-engine-84050757.html
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Perhaps there is a possibility that RS-68 inventory could be retrofitted?
You wouldn't throw away the old MCC and turbopumps. Too expensive. Useful pitches I just found that discuss the engine changes and potential booster changes.
(NASA Ares V pitch from 2007) http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20070032980_2007031295.pdf
(ULA pitch from 2009 JPC) http://www.ulalaunch.com/site/docs/publications/BoosterInnovationJPC2009.pdf
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The ULA doc has the RS-68A integrated into the Delta IV M+(5,4) CCB in the 2014 time frame. So the answer is (if we can believe ULA and power point) is the RS-68A will replace the RS-68 and go to a single CCB design starting in 2014.
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So, how will the HEFT study affect the Delta IV launch rate and price, especially for the Delta IV Heavy?
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Currently the price is fixed ...
http://www.losangeles.af.mil/library/factsheets/factsheet.asp?id=5324
Acquisition History
The initial phase of the EELV program, Low Cost Concept Validation (LCCV), was successfully completed in November 1996. LCCV emphasized competition in preliminary designs and risk reduction demonstrations. Four $30-million contracts were awarded during this phase to Alliant Techsystems, The Boeing Company, Lockheed Martin Corporation and McDonnell Douglas Aerospace. (Note: Boeing acquired McDonnell Douglas at about the time this competition ended.)
During the second phase, pre-engineering and manufacturing development, two $60-million, 17-month contracts were awarded to The Boeing Company and Lockheed Martin Corporation to continue refining their system concepts and complete a detailed system design.
EELV phase three began in October 1998 with the award of two development agreements and two initial launch services contracts (known as Buy 1) totaling more than $3 billion. The development agreements will run through fiscal year 2007 and the initial launch services contracts through fiscal year 2012. Additional launch service awards were made in Buy 2.
EELV Buy 3 is under negotiation. Buy 3 transitions the EELV contract from a commercial services contract to a hybrid of government contracts. The Buy 3 contracts consist of an EELV Launch Services (ELS) contract and an EELV Launch Capability (ELC) contract. The ELS is a firm fixed-price contract that buys the launch service to include mission hardware. The ELC is a combined firm fixed price/cost plus award fee contract that maintains workforce and facility readiness for mission success and provides the ability to deal with contingencies.
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What does that mean? Does that really mean 10 Delta IV mediums a year will have the same price as 1? Don't you think ULA will charge more for just the one, since their fixed costs are very high (compared to marginal costs) if you only launch one a year?
Anyone?
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The fixed (facility) costs are taken care of in the ELC contract.
edit: for clarity/specificity.
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The fixed costs are taken care of in the ELC contract.
So... Only the marginal costs are paid for, then?
Is that the only way that NASA can acquire launch services for the Delta IV Heavy?
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The fixed costs are taken care of in the ELC contract.
Is that strictly true? I thought that the DoD covered the fixed costs of the facilities, but that other customers picked up a share of the development costs.
That would also be a fixed cost, yes? If the company paid $1.5 billion to develop the launcher, and amortized them over, say, 20 years, then that would be $75 million a year, regardless of the number of flights. If you had two launches a year, that would be over $30 million a launch, if ten, $7.5
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Yes, development costs* and conversely any production efficiencies due to scale are shared amongst all customers. That is reflected in the ELS price which is roughly the price NASA would be paying too.
http://science.ksc.nasa.gov/shuttle/nexgen/EELV_main.htm
*not including those costs already agreed and paid for by specific customer (e.g. Government) in advance
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Yes, development costs and conversely any production efficiencies due to scale are shared amongst all customers. That is reflected in the ELS price which is roughly the price NASA would be paying too.
http://science.ksc.nasa.gov/shuttle/nexgen/EELV_main.htm
"Eventually it became clear that the costs perspective provided early in the EELV program hinged on volume, whereby commercial customers would be so abundant as to cumulatively contribute, in the commercial prices charged, to defraying an assortment of costs."
If it hinges on volume, and the volume goes way up because of more launches per year, then price can come down, right?
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Yes but costs may also go up as the ELC contract is for a certain DoD volume only and if more personnel/facilities are needed than provided by that capability that will be included in the price charged to other customers.
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Yes but costs may also go up as the ELC contract is for a certain DoD volume only and if more personnel/facilities are needed than provided by that capability that will be included in the price charged to other customers.
Right, but that is included in the incremental costs.
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The fixed costs are taken care of in the ELC contract.
So... Only the marginal costs are paid for, then?
Is that the only way that NASA can acquire launch services for the Delta IV Heavy?
No, NASA has a separate contract.
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Yes, development costs and conversely any production efficiencies due to scale are shared amongst all customers. That is reflected in the ELS price which is roughly the price NASA would be paying too.
http://science.ksc.nasa.gov/shuttle/nexgen/EELV_main.htm
This guy doesn't know what he is talking about
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Yes but costs may also go up as the ELC contract is for a certain DoD volume only and if more personnel/facilities are needed than provided by that capability that will be included in the price charged to other customers.
Right, but that is included in the incremental costs.
The costs we are talking about is about $400 million for the fixed ones and $300 million per manned Orion launch. Even at this rate that is 4 launches and a total of 1.6 billion. This would be sufficient to put up 100 tons in LEO. Which could be more than enough to do BOE exploration.
Heft's 100 ton brute has a unit cost of 1.8 billion dollars just for each launch and unkown fix costs perhaps around 2 billion a year.
You would have to launch the smaller delta IV five time for a total of 125 tons before the 1 launch of the 100 ton rocket is cheaper.
In addition small upgrades to the Delta IV to bring it up to say 30-50 could be a lot cheaper than building SLS.
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If you're going to make an assertion like "more Delta IV Heavies launched per year won't lower costs significantly," then you need evidence. Especially when one is launched only about every two years! Come on...
You're putting words in my mouth. Strange as it may seem on this board lately, I'm not trying to push an ideological stance here; I'm simply attempting to add a caution to what I saw as a misapprehension in someone's post.
What I was trying to convey was simply that looking at the very low Delta IV Heavy flight rate in isolation and trying to calculate savings as though it were the sole tenant in its fixed-/amortized-cost structure results in exaggeration of the potential per-flight savings that would be realized by increasing the flight rate.
Do you contest this?
I suppose it is actually a quantitative result supported by a quantitative analysis, if you allow the application of the term 'quantitative' to the > and < concepts...
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If you're going to make an assertion like "more Delta IV Heavies launched per year won't lower costs significantly," then you need evidence. Especially when one is launched only about every two years! Come on...
You're putting words in my mouth. Strange as it may seem on this board lately, I'm not trying to push an ideological stance here; I'm simply attempting to add a caution to what I saw as a misapprehension in someone's post.
What I was trying to convey was simply that looking at the very low Delta IV Heavy flight rate in isolation and trying to calculate savings as though it were the sole tenant in its fixed-/amortized-cost structure results in exaggeration of the potential per-flight savings that would be realized by increasing the flight rate.
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I understand, thanks for the clarification.
Of course, as I noted, Delta IV mediums don't have a heck of a very high flight rate, either... 1.25 flights per year.
What is the factory capacity for CBCs per year right now? I'm sure this sort of depends on what part of the pipeline, as I'm sure different parts of the pipeline have different capacities.
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Dunno. Jim?
It looks to me like DIRECT's graphs have fixed costs uptick at 40 cores per year, which is 13 flights of a heavy EELV, plus one medium. They don't specify which EELV; I believe I found out at one point and forgot...
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Dunno. Jim?
It looks to me like DIRECT's graphs have fixed costs uptick at 40 cores per year, which is 13 flights of a heavy EELV, plus one medium. They don't specify which EELV; I believe I found out at one point and forgot...
Decatur was originally built mid-90s for (IIRC) 40 DIV cores/yr. (There are document links around here somewhere.) Maybe they've scaled that infrastructure back some to lower their fixed costs, separately from the reduced staffing/shifts personnel costs, but I haven't seen any newer production limit figures reported. It's also complicated by the co-location of AV core production. Did they move tooling down from Colorado? Is Centaur still manufactured in San Diego?
-Alex
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Okay, let us be clear. ELC pays for the fixed cost (or ~95% IIRC) of both facilities and human capital for something like 5 launches of each EELV per year. An educated guess says this includes one heavy. ELS then pays for the marginal cost of each DoD launch.
Any non-DoD launch (NASA, commercial, whatever) pays both that marginal price, (though the price might not equal the ELS price since different customers negotiate different specific terms for special services, insight, etc) plus they pay part of the ELC, which is either refunded to the DoD or they just don't pay part of the next ELC installment.
HEFT is irrelevant until everything shakes out in Congress and there are RFPs or JOFOCs on the street.
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If NASA/ESMD chooses Delta-IV Heavy (DIVH) for a test flight of Orion in 2013, when would ULA need to get the order? The booster CBCs aren't exactly stock inventory, so what lead time is required for DIVH flights in general? Then also, could the Orion/DIVH configuration be purchased under the terms of the existing NASA/ULA contract for science missions? Or would ESMD need to compete the launch, in which case ULA could conceivably offer only AVH?
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If NASA/ESMD chooses Delta-IV Heavy (DIVH) for a test flight of Orion in 2013, when would ULA need to get the order? The booster CBCs aren't exactly stock inventory, so what lead time is required for DIVH flights in general? Then also, could the Orion/DIVH configuration be purchased under the terms of the existing NASA/ULA contract for science missions? Or would ESMD need to compete the launch, in which case ULA could conceivably offer only AVH?
Good questions.
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One answer is that, as discussed on a thread in the Live News section, Delta IV Heavy is not on the existing NASA/ULA contract. If NASA contracts directly with ULA, it would need another, most likely standalone contract. That takes a lot of calendar time to navigate the procurement hoops. There are regulations that allow not having to compete something if there's only one thing that can do it. One could also look at the GOES launches, where the launch was part of a delivery-on-orbit satellite contract; so the launch was contracted for as a commercial launch by the satellite contractor. But that contract was always intended to go that way. I doubt any existing contracts have that provision.
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Wouldn't they also have to sign a contract to make pad mods they need to make to support manned spaceflight?
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Wouldn't they also have to sign a contract to make pad mods they need to make to support manned spaceflight?
Not at first, and maybe not ever. The Orion team is proposing using DIVH for an uncrewed test flight, for which they might have a flight test article ready in 2013. That test flight gives them operational experience with Orion in orbit. Once SLS is available, the plan would be to use it to launch the crewed Orion flights.
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Wouldn't they also have to sign a contract to make pad mods they need to make to support manned spaceflight?
Not at first, and maybe not ever. The Orion team is proposing using DIVH for an uncrewed test flight, for which they might have a flight test article ready in 2013. That test flight gives them operational experience with Orion in orbit. Once SLS is available, the plan would be to use it to launch the crewed Orion flights.
That is one plan, but the HEFT report suggested that "commercial crew" was still a possibility for these missions, keeping SLS unmanned to launch an uncrewed "Orion" like spacecraft. It isn't clear to me that any final decisions have been made.
- Ed Kyle
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One answer is that, as discussed on a thread in the Live News section, Delta IV Heavy is not on the existing NASA/ULA contract. If NASA contracts directly with ULA, it would need another, most likely standalone contract. That takes a lot of calendar time to navigate the procurement hoops. There are regulations that allow not having to compete something if there's only one thing that can do it. One could also look at the GOES launches, where the launch was part of a delivery-on-orbit satellite contract; so the launch was contracted for as a commercial launch by the satellite contractor. But that contract was always intended to go that way. I doubt any existing contracts have that provision.
The DoD once "gifted" a Titan IV to NASA, Perhaps they would do the same with the Delta heavy and get a reimbursement from NASA.
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The DoD once "gifted" a Titan IV to NASA, Perhaps they would do the same with the Delta heavy and get a reimbursement from NASA.
NASA didn't have a contract with LM/Titan. Now, NASA has a contract with ULA that the DOD is envious of. The DOD actually wanted to go thru NASA for vehicles. Delta IV can be added to the contract easily.
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Now, NASA has a contract with ULA that the DOD is envious of. The DOD actually wanted to go thru NASA for vehicles.
Care to elaborate? Does this mean NASA negotiated a shrewder contract than the DOD did? Or DOD likes NASA's bureaucracy better than its own? The grass is always greener I guess ;)
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Now, NASA has a contract with ULA that the DOD is envious of. The DOD actually wanted to go thru NASA for vehicles.
Care to elaborate? Does this mean NASA negotiated a shrewder contract than the DOD did? Or DOD likes NASA's bureaucracy better than its own? The grass is always greener I guess ;)
Different sections of FAR
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When is the first launch planned for the RS-68A?
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The next Delta-IV Heavy I believe... It was listed in one of the write-ups on the upcoming launch.
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Do you have a link for that article?
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Hypothetically, when RS-25e is developed for SLS, could it also be used on Delta-IV for commonality with SLS, if it was cost-effective to do so? What kind of mods to Delta-IV would be needed, and what would it do to or for its performance?
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Hypothetically, when RS-25e is developed for SLS, could it also be used on Delta-IV for commonality with SLS, if it was cost-effective to do so? What kind of mods to Delta-IV would be needed, and what would it do to or for its performance?
Insufficient thrust for fully loaded Delta-IV cores..
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I don't like playing with Frankenrockets, but a structural feasibility study of a dual-SSME CBC would not be a horrible idea. It could solve some performance problems. If SLS can get the flight rate up, it might even be more economical than RS-68.
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Would it have enough thrust for the M+ versions?
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I've noticed that the first Heavy from Vandenberg has been resting on SLC-6 for a whole year before launch. Is this normal procedure for a heavy or was just because they were testing the new pad?
I'm curious because I've read that SpaceX moved Falcon 1 testing from SLC-3 to Omelek because range safety prevented them from doing launches while there was another LV in a near pad, and the one on pad could be there for months. Was it the last Titan or a Delta 4?
Is it normal to have a Delta for months on a pad? I ask because I think I've read that ULA wanted the RS-68A to standarize the CBC (which isn't exactly "common" nowadays) so it could offer a "quick" turnaround and flexibility in swapping LV and payloads. If a LV has to be on the pad for years, it would require not less than three years of lead time. No wonder that the way to have backups is to have them on orbit.
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1. I've noticed that the first Heavy from Vandenberg has been resting on SLC-6 for a whole year before launch. Is this normal procedure for a heavy or was just because they were testing the new pad?
2. I'm curious because I've read that SpaceX moved Falcon 1 testing from SLC-3 to Omelek because range safety prevented them from doing launches while there was another LV in a near pad, and the one on pad could be there for months. Was it the last Titan or a Delta 4?
3. Is it normal to have a Delta for months on a pad? I ask because I think I've read that ULA wanted the RS-68A to standarize the CBC (which isn't exactly "common" nowadays) so it could offer a "quick" turnaround and flexibility in swapping LV and payloads. If a LV has to be on the pad for years, it would require not less than three years of lead time. No wonder that the way to have backups is to have them on orbit.
1. new pad
2. Titan IV
3. 3 months
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The first flight of the RS-68A is decided? It was supposed to be this year, so it should be on the manifest, but I couldn't find it anywhere. May be some pointers to ULA's manifest?
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The first flight of the RS-68A is decided? It was supposed to be this year, so it should be on the manifest, but I couldn't find it anywhere. May be some pointers to ULA's manifest?
2012
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And the last plain-vanilla RS-68 flight?
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Would it have enough thrust for the M+ versions?
Yes, and also with a Heavy if the two boosters are RS-68. (the payload jumps to almost 45mT)
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How does 2012 work if it due to fly first on the Heavy? Did NRO L-15 slip to 2012 already? The next listed heavy is in 2015.
Will it not be on a Heavy?
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I would think one would definitely want to flight-certify RS-68A on a Medium before risking a higher value payload on a Heavy.
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Am I just remembering from a few years back it was being developed first for the Heavy and the plan was to fly out the inventory of RS-68's on the mediums before switching them.
Is that still the plan, or did ULA change things up?
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Looking at photos of yesterdays DIV-H launch... immediately after liftoff it can be seen that the foam on one side of all three first stage cores is burned black, and near the top of one booster it is still actively on fire. Is this not a problem?
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Looking at photos of yesterdays DIV-H launch... immediately after liftoff it can be seen that the foam on one side of all three first stage cores is burned black, and near the top of one booster it is still actively on fire. Is this not a problem?
Why it should be a problem?
EDIT: I watched videos and I admit it looked *far* worse than I imagined! :o This indeed looks... eh.... off-nominal ;)
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Looking at photos of yesterdays DIV-H launch... immediately after liftoff it can be seen that the foam on one side of all three first stage cores is burned black, and near the top of one booster it is still actively on fire. Is this not a problem?
It's not desirable, but it's not a "problem" per se.
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How does one keep the soot from the charing from penetrating the payload shroud?
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Positive pressure, like for any clean room environment? Who would want that dirty outside air getting inside, regardless of whether it contains a bit of soot?
Face reality: the main reason to avoid having your rocket set itself on fire is simply because outside observers don't like to see that happen.
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How does one keep the soot from the charing from penetrating the payload shroud?
The air is leaving the fairing as it climbs
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Positive pressure, like for any clean room
There isn't any added
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They actually had a significant redesign of the vents after Heavy Demo to prevent ingestion of free H2 and combustion products.
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Face reality: the main reason to avoid having your rocket set itself on fire is simply because outside observers don't like to see that happen.
Well, this "outside observer" thought it was a reasonable question.
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They actually had a significant redesign of the vents after Heavy Demo to prevent ingestion of free H2 and combustion products.
Didnt the pad have this problem during the shuttle era, preventing its use?
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They actually had a significant redesign of the vents after Heavy Demo to prevent ingestion of free H2 and combustion products.
Didnt the pad have this problem during the shuttle era, preventing its use?
He was referring to vents on the vehicle
The "problem" was fixed
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
No, the first D-IV flew as is after an FRF.
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
No, the first D-IV flew as is after an FRF.
FRF supposedly means Flight Readiness Firing.
Very well. Thanks.
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
No, the first D-IV flew as is after an FRF.
"What's a 'furf'?"
LOL, guess the movie kids! ;)
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They actually had a significant redesign of the vents after Heavy Demo to prevent ingestion of free H2 and combustion products.
Didnt the pad have this problem during the shuttle era, preventing its use?
He was referring to vents on the vehicle
The "problem" was fixed
I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time. Even before the Challenger, the work began to slow down and many were laid off. The Challenger accident made it officially shut down the launch pad.
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
No, the first D-IV flew as is after an FRF.
"What's a 'furf'?"
LOL, guess the movie kids! ;)
FRF = Flight Readiness Firing
But I recalled that was only a 1 second hot fire on the pad with the vehicle hold down bolts engaged.
The first DIV-H flight was when most were surprised at the amount of gH2 flames engulfing the vehicle. Someone told me that they had to call Range Safety to assure them this is 'normal' for this vehicle and DO NOT PUSH THE DESTRUCT BUTTON !!! ;D
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
No, the first D-IV flew as is after an FRF.
"What's a 'furf'?"
LOL, guess the movie kids! ;)
FRF = Flight Readiness Firing
Yeah, I know....
But take a shot at what movie the above quoute came from.....just some fun and games instead of the "seriousness" around here all the time.
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In case, would they do anything about the charred insulation after a pad abort?
Does the black surface (reduced albedo) impact thermal conditioning of cryo tanks in a subsequent launch attempt?
No, the first D-IV flew as is after an FRF.
"What's a 'furf'?"
LOL, guess the movie kids! ;)
FRF = Flight Readiness Firing
Yeah, I know....
But take a shot at what movie the above quoute came from.....just some fun and games instead of the "seriousness" around here all the time.
Oh sorry, didn't realize that was a joke. I guess we're just a bunch of serious nerds here ! :)
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But I recalled that was only a 1 second hot fire on the pad with the vehicle hold down bolts engaged.
By the time the engine reached full thrust, there was no free H2 being vented anymore. It could have been 10 seconds, wouldn't have made much difference. The FRF day also looks like it was somewhat windy and the vehicle wasn't really charred - at least the side I see in a LQ video.
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Strictly speaking LOX hydrogen engines run rich of stoichiometric, so there's plenty of free hydrogen in the plume at any time.
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Yes, but strictly speaking it's being thrown out the flame trench and actually sucking in the air above the engine in the process.
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I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time.
Incorrect. The jet engines were to be used to heat the ET to prevent ice formation in foggy/rainy conditions.
The problem you're talking about is called hydrogen entrapment. Hydrogen would be trapped in the closed exhaust duct leading to significant build-up prior to engine start.
The fix was steam injection which would would dilute the hydrogen preventing an explosion.
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I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time.
Not true.
A. the duct of concern was dedicated for the SSME's, It can be seen in some of the photos and is filled with concrete. The Shuttle SRB's had dedicated ducts were were modified for Delta IV
2. I see DaveS had answered.
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Some discussion in this thread about whether certain NRO payloads can fit on a M+(5,x) vehicle.
http://forum.nasaspaceflight.com/index.php?topic=11127.45
The 5m fairing on the M+ models is shorter than the one on the Heavy. Can the long Heavy fairing (the composite one, as opposed to the aluminum trisector fairing) be used on the M+?
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Some discussion in this thread about whether certain NRO payloads can fit on a M+(5,x) vehicle.
http://forum.nasaspaceflight.com/index.php?topic=11127.45
The 5m fairing on the M+ models is shorter than the one on the Heavy. Can the long Heavy fairing (the composite one, as opposed to the aluminum trisector fairing) be used on the M+?
aluminum trisector is only for the heavy and certain missions.
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Strictly speaking LOX hydrogen engines run rich of stoichiometric, so there's plenty of free hydrogen in the plume at any time.
Yes, but not all LOX/H2 engines exhibit this amount of free H2 at startup to generate such a fireball as the RS-68 does. For example, the SSME and Vulcain engines have remarkable clean startups in comparison. Either that, or the excess H2 is vented away much more efficiently.
I know you are not arguing this personally, but it seems like several people here have been suggesting that this kind of start is a "feature" of all LOX/H2 engines. But this is clearly not the case.
Clearly the RS-68 was designed to do what it does, and it works, but it still raises eyebrows.
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Strictly speaking LOX hydrogen engines run rich of stoichiometric, so there's plenty of free hydrogen in the plume at any time.
Yes, but not all LOX/H2 engines exhibit this amount of free H2 at startup to generate such a fireball as the RS-68 does. For example, the SSME and Vulcain engines have remarkable clean startups in comparison. Either that, or the excess H2 is vented away much more efficiently.
I know you are not arguing this personally, but it seems like several people here have been suggesting that this kind of start is a "feature" of all LOX/H2 engines. But this is clearly not the case.
As stated earlier, the RS-68 has only 25% more H2 at start. The water system at LC-39 help suck it away
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A tighter start sequence also has a greater possibility of going "hardware" rich. Pick your poison.
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Some discussion in this thread about whether certain NRO payloads can fit on a M+(5,x) vehicle.
http://forum.nasaspaceflight.com/index.php?topic=11127.45
The 5m fairing on the M+ models is shorter than the one on the Heavy. Can the long Heavy fairing (the composite one, as opposed to the aluminum trisector fairing) be used on the M+?
aluminum trisector is only for the heavy and certain missions.
Understood, but is it possible to use the Heavy fairing on the M+, for payloads that need more volume but not more payload capacity?
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I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time.
Incorrect. The jet engines were to be used to heat the ET to prevent ice formation in foggy/rainy conditions.
The problem you're talking about is called hydrogen entrapment. Hydrogen would be trapped in the closed exhaust duct leading to significant build-up prior to engine start.
The fix was steam injection which would would dilute the hydrogen preventing an explosion.
Thanks. But I believe the correct word is called Hydrogen Entrainment. The RS-68 has a much higher gaseous hydrogen "lead" during engine start-up than SSME. The fear was that these GH2 will be "pre-mixed" with air in the exhaust duct so, when engine ignites will set up a perfect bomb scenario, creating a detonation wave blow back to the vehicle. The Delta IV fix was to install ROFIs (sparklers) in the duct to burn away as much pre-mixed hydrogen/air as possible.
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Strictly speaking LOX hydrogen engines run rich of stoichiometric, so there's plenty of free hydrogen in the plume at any time.
Yes, but not all LOX/H2 engines exhibit this amount of free H2 at startup to generate such a fireball as the RS-68 does. For example, the SSME and Vulcain engines have remarkable clean startups in comparison. Either that, or the excess H2 is vented away much more efficiently.
I know you are not arguing this personally, but it seems like several people here have been suggesting that this kind of start is a "feature" of all LOX/H2 engines. But this is clearly not the case.
As stated earlier, the RS-68 has only 25% more H2 at start. The water system at LC-39 help suck it away
Huh? Explain how does water "suck" away gaseous H2
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Clearly the RS-68 was designed to do what it does, and it works, but it still raises eyebrows.
It was not designed to do this on purpose. It was a by-product as the design of vehicle and ground system came together. But it was an acceptable outcome as it posed no threat to the vehicle and missions.
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Some discussion in this thread about whether certain NRO payloads can fit on a M+(5,x) vehicle.
http://forum.nasaspaceflight.com/index.php?topic=11127.45
The 5m fairing on the M+ models is shorter than the one on the Heavy. Can the long Heavy fairing (the composite one, as opposed to the aluminum trisector fairing) be used on the M+?
aluminum trisector is only for the heavy and certain missions.
Understood, but is it possible to use the Heavy fairing on the M+, for payloads that need more volume but not more payload capacity?
The 5m composite fairing is the same on the heavy and Medium + (5,X)
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Huh? Explain how does water "suck" away gaseous H2
Hydraulic ventilation
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I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time.
Incorrect. The jet engines were to be used to heat the ET to prevent ice formation in foggy/rainy conditions.
The problem you're talking about is called hydrogen entrapment. Hydrogen would be trapped in the closed exhaust duct leading to significant build-up prior to engine start.
The fix was steam injection which would would dilute the hydrogen preventing an explosion.
Thanks. But I believe the correct word is called Hydrogen Entrainment. The RS-68 has a much higher gaseous hydrogen "lead" during engine start-up than SSME. The fear was that these GH2 will be "pre-mixed" with air in the exhaust duct so, when engine ignites will set up a perfect bomb scenario, creating a detonation wave blow back to the vehicle. The Delta IV fix was to install ROFIs (sparklers) in the duct to burn away as much pre-mixed hydrogen/air as possible.
The term used wrt to SLC-6 was entrapment
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I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time.
Incorrect. The jet engines were to be used to heat the ET to prevent ice formation in foggy/rainy conditions.
The problem you're talking about is called hydrogen entrapment. Hydrogen would be trapped in the closed exhaust duct leading to significant build-up prior to engine start.
The fix was steam injection which would would dilute the hydrogen preventing an explosion.
Thanks. But I believe the correct word is called Hydrogen Entrainment. The RS-68 has a much higher gaseous hydrogen "lead" during engine start-up than SSME. The fear was that these GH2 will be "pre-mixed" with air in the exhaust duct so, when engine ignites will set up a perfect bomb scenario, creating a detonation wave blow back to the vehicle. The Delta IV fix was to install ROFIs (sparklers) in the duct to burn away as much pre-mixed hydrogen/air as possible.
The term used wrt to SLC-6 was entrapment
So the physics is different at SLC-6 than at LC-17? ;D
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Some discussion in this thread about whether certain NRO payloads can fit on a M+(5,x) vehicle.
http://forum.nasaspaceflight.com/index.php?topic=11127.45
The 5m fairing on the M+ models is shorter than the one on the Heavy. Can the long Heavy fairing (the composite one, as opposed to the aluminum trisector fairing) be used on the M+?
aluminum trisector is only for the heavy and certain missions.
Understood, but is it possible to use the Heavy fairing on the M+, for payloads that need more volume but not more payload capacity?
The 5m composite fairing is the same on the heavy and Medium + (5,X)
Really? The Payload Planners' Guide shows 14.3 m length for the (5,x) fairing and 19.1 m for the Heavy fairing.
I know the long fairing is not currently offered on 5-meter versions. I guess what I'm trying to get at is, is using the longer fairing on a M+ something that they could do relatively easily if the customer requested it?
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I recalled that there was a 'strong' concern regarding to the exhaust duct at SLC 6 to safely funnel the rocket plume exhausts away from the Shuttle. At the time, turbojet engines were used to provide suction power to draw the exhaust plumes away, but there were enough concerns voiced at the time.
Incorrect. The jet engines were to be used to heat the ET to prevent ice formation in foggy/rainy conditions.
The problem you're talking about is called hydrogen entrapment. Hydrogen would be trapped in the closed exhaust duct leading to significant build-up prior to engine start.
The fix was steam injection which would would dilute the hydrogen preventing an explosion.
Thanks. But I believe the correct word is called Hydrogen Entrainment. The RS-68 has a much higher gaseous hydrogen "lead" during engine start-up than SSME. The fear was that these GH2 will be "pre-mixed" with air in the exhaust duct so, when engine ignites will set up a perfect bomb scenario, creating a detonation wave blow back to the vehicle. The Delta IV fix was to install ROFIs (sparklers) in the duct to burn away as much pre-mixed hydrogen/air as possible.
The term used wrt to SLC-6 was entrapment
So the physics is different at SLC-6 than at LC-17? ;D
Same Physics, different physical situation.
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Some discussion in this thread about whether certain NRO payloads can fit on a M+(5,x) vehicle.
http://forum.nasaspaceflight.com/index.php?topic=11127.45
The 5m fairing on the M+ models is shorter than the one on the Heavy. Can the long Heavy fairing (the composite one, as opposed to the aluminum trisector fairing) be used on the M+?
aluminum trisector is only for the heavy and certain missions.
Understood, but is it possible to use the Heavy fairing on the M+, for payloads that need more volume but not more payload capacity?
The 5m composite fairing is the same on the heavy and Medium + (5,X)
Really? The Payload Planners' Guide shows 14.3 m length for the (5,x) fairing and 19.1 m for the Heavy fairing.
I know the long fairing is not currently offered on 5-meter versions. I guess what I'm trying to get at is, is using the longer fairing on a M+ something that they could do relatively easily if the customer requested it?
I was getting at that the aluminum one is not used on any type of non Heavy vehicles. As for different lengths of composite 5m fairings, ULA should be able to meet any request.
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Another question about the fireball:
if the same amount of H2 that is producing the fireball during a delta IV heavy launch were to somehow thoroughly mix with air prior to ignition, you would have a much more violent explosion.
Is the delta IV designed to handle that as well? And if not, how is it ensured that the H2 does not mix with air? CFD simulations of the volume around the vehicle in all possible weather conditions?
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Another question about the fireball:
if the same amount of H2 that is producing the fireball during a delta IV heavy launch were to somehow thoroughly mix with air prior to ignition, you would have a much more violent explosion.
Is the delta IV designed to handle that as well? And if not, how is it ensured that the H2 does not mix with air? CFD simulations of the volume around the vehicle in all possible weather conditions?
I guess sparklers' purpose is exactly to prevent that: they guarantee that LH ignites before there is a big amount of it released and mixed with the air.
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What is the purpose of the Al fairing vs composite fairings? RFI?
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Would it have enough thrust for the M+ versions?
Yes, and also with a Heavy if the two boosters are RS-68. (the payload jumps to almost 45mT)
Neato! Sounds very cost-effective. $30 million per launch, nearly double the payload. I am surprised they have not considered this.
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I don't like playing with Frankenrockets, but a structural feasibility study of a dual-SSME CBC would not be a horrible idea. It could solve some performance problems. If SLS can get the flight rate up, it might even be more economical than RS-68.
Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
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Seeing how much the vehicle was cryo-smoking (yeah, my word) on this launch compared to other Heavys I'd almost venture to say that all of the foam was burned off in some spots.
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I don't like playing with Frankenrockets, but a structural feasibility study of a dual-SSME CBC would not be a horrible idea. It could solve some performance problems. If SLS can get the flight rate up, it might even be more economical than RS-68.
Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
From what I understand it is similar to this scenario (not saying these are the SSME costs, just that this is how the cost is determined):
To run the production line costs $60 million a year.
Per-engine cost is $20 million.
The more engines you produce, the cheaper each one is. On a production run of 1 engine per year, as is usual, the engine is $80 million each. For a 3 engine vehicle, like the J-130 or Delta V Heavy, however, in a two-launch per year situation we'd be producing 6 engines per year. This means the cost per-engine, is $30 million, only a bit more than the RS-68A at $24 million. If we have the 5-engine Ares V Classic, now the cost per-engine is $26 million. Launch that three times in a year instead of two, the cost per-engine is now the same as the RS-68A, $24 million.
The RS-68's are produced in bulk, placing a large order for multiple years all at once, so reducing it's cost to produce is not going to happen. The SSME, however, has always been ordered as-needed, so the sticker cost is higher. If we produced them in the same quantities as the RS-68, the cost would be quite comparable, but the SSME's are a more capable engine over a longer period of flight, which makes them ideal for parallel staged vehicles. I even did a calculation once, if you took a Delta IV Heavy, and replaced the RS-68 in the center with an SSME, it's payload would go up by almost 10mT.
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Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
The price hike was due to the fixed costs of PWR that had been borne by the SSME contract being spread over the other engine programs after the FY11 budget seemed to indicate the coming demise of the SSME contract. I'm extrapolating a bit that the unit cost of SSMEs would not be any worse than it had been.
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Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
The price hike was due to the fixed costs of PWR that had been borne by the SSME contract being spread over the other engine programs after the FY11 budget seemed to indicate the coming demise of the SSME contract. I'm extrapolating a bit that the unit cost of SSMEs would not be any worse than it had been.
One could seriously consider the RD-0120 for the expendables such as the replacement for RS-68 and for the SLS considering the price advantage. The RD-0120 performance is very comparable to the SSME but at a significant lower price.
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Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
The price hike was due to the fixed costs of PWR that had been borne by the SSME contract being spread over the other engine programs after the FY11 budget seemed to indicate the coming demise of the SSME contract. I'm extrapolating a bit that the unit cost of SSMEs would not be any worse than it had been.
One could seriously consider the RD-0120 for the expendables such as the replacement for RS-68 and for the SLS considering the price advantage. The RD-0120 performance is very comparable to the SSME but at a significant lower price.
Not really, not in the order quantities being talked about. Plus, the RD-0120 is not produced anymore. NASA has, however, studied it extensively with an eye to adapting it to the SSME, This combined design is referred to as the RS-25e. Best of both worlds, higher performance, lower cost. SSME has better thrust, equal isp. With the proposed changes, restarting the RD-0120 production would cost more, take longer, and give you an engine with less performance.
Something I noticed as well. The SSME is not *that* much less thrust than an RS-68, only 25% less. The RD-0120, however, is 45% less than the RS-68. With the SSME's better isp, it still wins. If we do follow NASA's plan to reduce the cost to produce by adapting technologies from other engines, there is no compelling argument against the SSME in my opinion.
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400k / 663k is 40% less not 25%
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400k / 663k is 40% less not 25%
Why are you running the SSME under maximum thrust? And compare against back thrust as well. But this time run it at the SSME's max.
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Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
The price hike was due to the fixed costs of PWR that had been borne by the SSME contract being spread over the other engine programs after the FY11 budget seemed to indicate the coming demise of the SSME contract. I'm extrapolating a bit that the unit cost of SSMEs would not be any worse than it had been.
One could seriously consider the RD-0120 for the expendables such as the replacement for RS-68 and for the SLS considering the price advantage. The RD-0120 performance is very comparable to the SSME but at a significant lower price.
Not really, not in the order quantities being talked about. Plus, the RD-0120 is not produced anymore. NASA has, however, studied it extensively with an eye to adapting it to the SSME, This combined design is referred to as the RS-25e. Best of both worlds, higher performance, lower cost. SSME has better thrust, equal isp. With the proposed changes, restarting the RD-0120 production would cost more, take longer, and give you an engine with less performance.
Something I noticed as well. The SSME is not *that* much less thrust than an RS-68, only 25% less. The RD-0120, however, is 45% less than the RS-68. With the SSME's better isp, it still wins. If we do follow NASA's plan to reduce the cost to produce by adapting technologies from other engines, there is no compelling argument against the SSME in my opinion.
RS-0120 has a better Isp than the SSME.
SSME sea level thrust@ 109% power is at 419K lbf, so that's only 63% of RS-68A, the RD-0120 sea level thrust is at 341K, at 51% of RS-68A.
But that makes the RD-0120 a better fit as a replacement of RS-68s.
Two RD-01020 is just about identical as one RS-68A thrust, therefore minimum structural modification on the CBC other than the obvious (2 engine feed & gimbals). At a significantly higher Isp, itwill make the Delta-IV a higher performance vehicle with minimum mods.
SSME, on the other hand, will require way too much throttling than currently capable.
As far as on currently in production, well; 1) the market place is a tremendous motivator, and 2) SSME is currently not in production either. Just try to find the toolings for them.
As for the "NASA studies", there's a huge difference between "study" and "done it".
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Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
The price hike was due to the fixed costs of PWR that had been borne by the SSME contract being spread over the other engine programs after the FY11 budget seemed to indicate the coming demise of the SSME contract. I'm extrapolating a bit that the unit cost of SSMEs would not be any worse than it had been.
One could seriously consider the RD-0120 for the expendables such as the replacement for RS-68 and for the SLS considering the price advantage. The RD-0120 performance is very comparable to the SSME but at a significant lower price.
Not really, not in the order quantities being talked about. Plus, the RD-0120 is not produced anymore. NASA has, however, studied it extensively with an eye to adapting it to the SSME, This combined design is referred to as the RS-25e. Best of both worlds, higher performance, lower cost. SSME has better thrust, equal isp. With the proposed changes, restarting the RD-0120 production would cost more, take longer, and give you an engine with less performance.
Something I noticed as well. The SSME is not *that* much less thrust than an RS-68, only 25% less. The RD-0120, however, is 45% less than the RS-68. With the SSME's better isp, it still wins. If we do follow NASA's plan to reduce the cost to produce by adapting technologies from other engines, there is no compelling argument against the SSME in my opinion.
RS-0120 has a better Isp than the SSME.
SSME sea level thrust@ 109% power is at 419K lbf, so that's only 63% of RS-68A, the RD-0120 sea level thrust is at 341K, at 51% of RS-68A.
Where are you getting the stats for the RS-68A, as it's not yet published? I've been comparing against the RS-68, which makes your comparison far weaker.
But that makes the RD-0120 a better fit as a replacement of RS-68s.
Two RD-01020 is just about identical as one RS-68A thrust, therefore minimum structural modification on the CBC other than the obvious (2 engine feed & gimbals). At a significantly higher Isp, itwill make the Delta-IV a higher performance vehicle with minimum mods.
You'd still have to restart production, and you would have the situation of the RD-180 off of Atlas, a foreign made engine.
SSME, on the other hand, will require way too much throttling than currently capable.
As far as on currently in production, well; 1) the market place is a tremendous motivator, and 2) SSME is currently not in production either. Just try to find the toolings for them.
As for the "NASA studies", there's a huge difference between "study" and "done it".
SSME has a wider throttling range than both RS-68 and RD-0120, so not certain what you are referring to. There has even been discussion of further pushing it's thrust.
Also, SSME can burn for a longer range than RS-68, compare the vac thrust and you find the RS-68 not as strong.
Now, you're still paying for startup and R&D costs in either case (RD-0120 and RS-25e) but one would give us more thrust, and have a domestic production history to work from.
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Considering the recent price hike on engines, any idea how much the SSME/ RS-25e will cost these days?
Am I the only one that think it is a shame to throw away all these SSMEs on every flight? How is that affordable?
The price hike was due to the fixed costs of PWR that had been borne by the SSME contract being spread over the other engine programs after the FY11 budget seemed to indicate the coming demise of the SSME contract. I'm extrapolating a bit that the unit cost of SSMEs would not be any worse than it had been.
One could seriously consider the RD-0120 for the expendables such as the replacement for RS-68 and for the SLS considering the price advantage. The RD-0120 performance is very comparable to the SSME but at a significant lower price.
Not really, not in the order quantities being talked about. Plus, the RD-0120 is not produced anymore. NASA has, however, studied it extensively with an eye to adapting it to the SSME, This combined design is referred to as the RS-25e. Best of both worlds, higher performance, lower cost. SSME has better thrust, equal isp. With the proposed changes, restarting the RD-0120 production would cost more, take longer, and give you an engine with less performance.
Something I noticed as well. The SSME is not *that* much less thrust than an RS-68, only 25% less. The RD-0120, however, is 45% less than the RS-68. With the SSME's better isp, it still wins. If we do follow NASA's plan to reduce the cost to produce by adapting technologies from other engines, there is no compelling argument against the SSME in my opinion.
RS-0120 has a better Isp than the SSME.
SSME sea level thrust@ 109% power is at 419K lbf, so that's only 63% of RS-68A, the RD-0120 sea level thrust is at 341K, at 51% of RS-68A.
Where are you getting the stats for the RS-68A, as it's not yet published? I've been comparing against the RS-68, which makes your comparison far weaker.
If that is the case, then your numbers are full of mistakes.
RS-68 sea level thrust is 650K lbf.
But that makes the RD-0120 a better fit as a replacement of RS-68s.
Two RD-01020 is just about identical as one RS-68A thrust, therefore minimum structural modification on the CBC other than the obvious (2 engine feed & gimbals). At a significantly higher Isp, itwill make the Delta-IV a higher performance vehicle with minimum mods.
You'd still have to restart production, and you would have the situation of the RD-180 off of Atlas, a foreign made engine.
and that has been a problem for National Security payload launches how?
SSME, on the other hand, will require way too much throttling than currently capable.
As far as on currently in production, well; 1) the market place is a tremendous motivator, and 2) SSME is currently not in production either. Just try to find the toolings for them.
As for the "NASA studies", there's a huge difference between "study" and "done it".
SSME has a wider throttling range than both RS-68 and RD-0120, so not certain what you are referring to. There has even been discussion of further pushing it's thrust.
Also, SSME can burn for a longer range than RS-68, compare the vac thrust and you find the RS-68 not as strong.
Now, you're still paying for startup and R&D costs in either case (RD-0120 and RS-25e) but one would give us more thrust, and have a domestic production history to work from.
Americanization of Russian engines is no longer a myth. We are not comparing SSME with RS-68, we are comparing the SSME with RD-0120 as option to replace the RS-68s, also as core engines for the SLS LO2/LH2 option.
In this case, more thrust is not necessary better.
For throw-away engines, cost is a big issue.
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Pratt & Whitney Rocketdyne Successfully Completes Hot-Fire Test Series on Second RS-68A Certification Engine (http://www.pw.utc.com/Media+Center/Press+Releases/Pratt+%26+Whitney+Rocketdyne+Successfully+Completes+Hot-Fire+Test+Series+on+Second+RS-68A+Certification+Engine)
November 22, 2010
"Like the RS-68, RS-68A engines will be used to boost the Delta IV rocket, but each RS-68A engine will provide 705,000 pounds of lift-off thrust, or 42,000 more pounds of thrust than a basic RS-68 engine."
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Where are you getting the stats for the RS-68A, as it's not yet published? I've been comparing against the RS-68, which makes your comparison far weaker.
If that is the case, then your numbers are full of mistakes.
RS-68 sea level thrust is 650K lbf.
Right, and the SSME's SL thrust @ 109% is 420k lbf.
420/650 == 65%
Now, at vac, we're dealing with RS-68 at 758k and the SSME at 515k, so, 515/758 == 68%. Yes?
And we're looking at 410 isp vs 453 isp.
But that makes the RD-0120 a better fit as a replacement of RS-68s.
Two RD-01020 is just about identical as one RS-68A thrust, therefore minimum structural modification on the CBC other than the obvious (2 engine feed & gimbals). At a significantly higher Isp, itwill make the Delta-IV a higher performance vehicle with minimum mods.
You'd still have to restart production, and you would have the situation of the RD-180 off of Atlas, a foreign made engine.
and that has been a problem for National Security payload launches how?
It's a political non-starter. Atlas got away with it only due to the agreement that Delta would use a domestic engine.
SSME, on the other hand, will require way too much throttling than currently capable.
As far as on currently in production, well; 1) the market place is a tremendous motivator, and 2) SSME is currently not in production either. Just try to find the toolings for them.
As for the "NASA studies", there's a huge difference between "study" and "done it".
SSME has a wider throttling range than both RS-68 and RD-0120, so not certain what you are referring to. There has even been discussion of further pushing it's thrust.
Also, SSME can burn for a longer range than RS-68, compare the vac thrust and you find the RS-68 not as strong.
Now, you're still paying for startup and R&D costs in either case (RD-0120 and RS-25e) but one would give us more thrust, and have a domestic production history to work from.
Americanization of Russian engines is no longer a myth. We are not comparing SSME with RS-68, we are comparing the SSME with RD-0120 as option to replace the RS-68s, also as core engines for the SLS LO2/LH2 option.
In this case, more thrust is not necessary better.
For throw-away engines, cost is a big issue.
And we're back to the original issue, SSME at $28 mil vs RD-0120 at $25 mil, with the added bonus of the SSME production reducing the cost of RL-10 production as well. Whole system cost still gives SSME the edge by my math.
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Source for a $28M SSME? Non-credible in my estimation.
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Source for a $28M SSME? Non-credible in my estimation.
If you wish to call NASA a liar, feel free. That is the cost estimate put forward last year in their SD-HLV study papers. They've had cost estimates out since 2003, and PWR has reams and reams of documentation for it if you want to ask them nicely for it. (Which is what I did)
The RS-25e is a well documented upgrade path to the existing RS-25d SSME engine. It incorporates numerous upgrades, and system elimination methods, to bring the cost of the engines down. If you're calling it non-credible, you should call the cost estimates of the RS-68A non-credible, for the same people did both.
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Source for a $28M SSME? Non-credible in my estimation.
If you wish to call NASA a liar, feel free. That is the cost estimate put forward last year in their SD-HLV study papers. They've had cost estimates out since 2003, and PWR has reams and reams of documentation for it if you want to ask them nicely for it. (Which is what I did)
The RS-25e is a well documented upgrade path to the existing RS-25d SSME engine. It incorporates numerous upgrades, and system elimination methods, to bring the cost of the engines down. If you're calling it non-credible, you should call the cost estimates of the RS-68A non-credible, for the same people did both.
Antares is saying that anything done before the SSME shut down is no longer applicable and yes, it applies to the RS-68. It is much more expensive now.
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Source for a $28M SSME? Non-credible in my estimation.
If you wish to call NASA a liar, feel free. That is the cost estimate put forward last year in their SD-HLV study papers. They've had cost estimates out since 2003, and PWR has reams and reams of documentation for it if you want to ask them nicely for it. (Which is what I did)
The RS-25e is a well documented upgrade path to the existing RS-25d SSME engine. It incorporates numerous upgrades, and system elimination methods, to bring the cost of the engines down. If you're calling it non-credible, you should call the cost estimates of the RS-68A non-credible, for the same people did both.
Antares is saying that anything done before the SSME shut down is no longer applicable and yes, it applies to the RS-68. It is much more expensive now.
Right now, it's all a bit of an unknown. I used the best data available for the discussion, which was current as of 9 months ago. Whenever they post new information, I'll go by that.
But right now, almost all rocket costs have been going up, no matter the company or even country.
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Source for a $28M SSME? Non-credible in my estimation.
If you wish to call NASA a liar, feel free. That is the cost estimate put forward last year in their SD-HLV study papers. They've had cost estimates out since 2003, and PWR has reams and reams of documentation for it if you want to ask them nicely for it. (Which is what I did)
The RS-25e is a well documented upgrade path to the existing RS-25d SSME engine. It incorporates numerous upgrades, and system elimination methods, to bring the cost of the engines down. If you're calling it non-credible, you should call the cost estimates of the RS-68A non-credible, for the same people did both.
Antares is saying that anything done before the SSME shut down is no longer applicable and yes, it applies to the RS-68. It is much more expensive now.
Right now, it's all a bit of an unknown. I used the best data available for the discussion, which was current as of 9 months ago. Whenever they post new information, I'll go by that.
But right now, almost all rocket costs have been going up, no matter the company or even country.
If the tooling and know how still exists then restarting is not going to be too big of a deal.
I think Russia had restarted production on engines such as the RD-170 after a more longer hiatus.
Restarting the SSME now would be cake compared to the idea of restarting F1 production back in the 80s or designing the J-2X.
It probably would even be less work then finishing the RL-60 a much smaller engine for use on upper stages.
Now the Delta really could use the AUS with the RL-60 or even ACES.
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I think Russia had restarted production on engines such as the RD-170 after a more longer hiatus.
This is not correct. The RD-170 family has been flying more or less continuously since it became operational.
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Nothing before February 2010 on PWR engines is valid now. When the SSME contract was given an end date, the question of who bears the fixed costs of PWR became unknown. I wouldn't believe anything south of $50M, and I start getting confident around $80M given the costs of other PWR engines. 'Course, it's all dependent on how many engines are built each year.
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I know that ULA has been working on a "fuel densification" technologies. I understand that this means to keep the cryogenic hydrogen at higher temperature/pressure. Or does it means to still use the same temperature at a higher pressure?
On a related note, what temperature could you keep the hydrogen if you used 400Bar pressure vessels? All public data that I've found on hydrogen is on the 1~12Bar range. In fact, I'd love to know what pressure would keep the hydrogen liquid at 20C.
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I know that ULA has been working on a "fuel densification" technologies. I understand that this means to keep the cryogenic hydrogen at higher temperature/pressure. Or does it means to still use the same temperature at a higher pressure?
Actually it is to make it colder.
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What are the chances of Delta 4 medium picking up additional missions over the next several years? If it isn't on NLS2, and it's more expensive than Atlas V, it seems like the only reason a non-heavy Delta 4 would be flying is DoD/NRO working around current Atlas flight rate constraints.
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What are the chances of Delta 4 medium picking up additional missions over the next several years? If it isn't on NLS2, and it's more expensive than Atlas V, it seems like the only reason a non-heavy Delta 4 would be flying is DoD/NRO working around current Atlas flight rate constraints.
Delta IV will be flying approx 4 missions per year.
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What are the chances of Delta 4 medium picking up additional missions over the next several years? If it isn't on NLS2, and it's more expensive than Atlas V, it seems like the only reason a non-heavy Delta 4 would be flying is DoD/NRO working around current Atlas flight rate constraints.
Delta IV will be flying approx 4 missions per year.
As it stands now, leaving D4H out, what profiles are better covered by D4 and which by A5? Being from the same company I'd think they would be sort of LV agnostic. But as an economist, I would think that until they truly merge at every level, there would be an amount of launches for each vehicle that would require further investment, and thus be cheaper to cover with the other LV. So, the question is:
Are some profiles better for one and some for the other or are they using the A5 at maximum capacity and filling the rest with Delta 4?
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NRO for DIV
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NRO for DIV
Future policy or past? NROL-41 was on AV-501 last year, plus three others on 401/411.
-Alex
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As it stands now, leaving D4H out, what profiles are better covered by D4 and which by A5? Being from the same company I'd think they would be sort of LV agnostic. But as an economist, I would think that until they truly merge at every level, there would be an amount of launches for each vehicle that would require further investment, and thus be cheaper to cover with the other LV. So, the question is:
Are some profiles better for one and some for the other or are they using the A5 at maximum capacity and filling the rest with Delta 4?
Delta and Atlas are still owned by Beoing and lockhead Martin. ULA just builds and launches them. Getting rid of Delta is probably not easy as Both Lockhead Martin and Boeing are on ULA board.
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NRO for DIV
I have to understand that for the general orbits and payloads that the NROL needs, Delta 4 is better optimized? I still don't parse Jim that easily.
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No, just less Russians
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No, just less Russians
See? That was short and clear. Thank you.
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What is the current status of Delta IV (Heavy) upgrades other than RS-68A? Is there a real need for the upgrades? Are they planned to actually happen, or just powerpoint right now?
Also, what's the status of RS-68A? What performance will this enable to LEO for Delta IV Heavy? To C3=0?
Thanks!
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Apparently they have passed the first acceptance test of the first engine for the Delta 4 Heavy mission.
You can read the Press Release (http://www.pw.utc.com/Media+Center/Press+Releases/Pratt+%26+Whitney+Rocketdyne+Demonstrates+First+RS-68A+Production+Engine+is+Ready+for+Flight)
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Quick Thor, aka delta question. Something Jim recently mentioned had me thinking. Are there any good pics. of the flamey side of the old Thor IRBM. I am now a little curious how the Thor/Delta/Delta II Verniers evolved with time. I had never given it much thought.
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They never changed
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Now that is an answer I did not expect ...
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This is a question about Delta Mariner, the ship that transports Delta (and now Atlas) cores from the assembly plant to the launch site. What facilities are needed at the receiving end of those shipments? Could cores be delivered to the same location where Shuttle ET cores were delivered? Could cores be delivered to Michoud and then be transported on from there?
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This is a question about Delta Mariner, the ship that transports Delta (and now Atlas) cores from the assembly plant to the launch site. What facilities are needed at the receiving end of those shipments? Could cores be delivered to the same location where Shuttle ET cores were delivered? Could cores be delivered to Michoud and then be transported on from there?
The Delta Mariner uses the dock at VAFB that was intended for shuttle ET's.
The Delta Mariner is a RO RO (like the ET Barge) and only a Delta IV EPT (transporter) is needed to remove the core for Delta. The Atlas core had its tractor with it.
Why would cores be delivered to Michoud?
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The Delta Mariner uses the dock at VAFB that was intended for shuttle ET's.
The Delta Mariner is a RO RO (like the ET Barge) and only a Delta IV EPT (transporter) is needed to remove the core for Delta. The Atlas core had its tractor with it.
Why would cores be delivered to Michoud?
This goes slightly off topic but it really is about the newly expanded Delta Mariner capabilities.
For a design like AJAX there are (at least) two different scenarios for how Atlas CCBs would get to the VAB by water transport. One is that Delta Mariner would take them directly to the turn basin and offloads them there. It can get there and do that? If for any reason it couldn't, the cores could presumably be cross-loaded onto Pegasus at Michoud.
The AJAX theory is that the CCBs could be used completely unmodified ("white tail"). But if it turns out they can't, the motivations for the USAF white tail requirement might be at least partially addressed by modifying the cores at some location outside Decatur. So the other imaginable scenario for Michoud delivery is where, for whatever reason (e.g. political), work to modify the CCBs for AJAX is given to Michoud rather than KSC.
More generally, Vitter has done a pretty good job keeping Michoud relevant in the changing world of NASA exploration systems.
http://www.youtube.com/watch?v=0NyQWp92htk
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Hypothetical question:
What is the current max number of DIVH flights that could be flown per year if only heavies were flown?
EDIT: from Florida
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Not sure if this is the proper thread and forgive me if it's been answered already (I've done some digging here and didn't find an answer), but I'm curious - what route does the Delta Mariner take from Decatur, AL. to the Cape to deliver rocket stages, etc. ?
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Another question. With the RS-68A, the Medium fleet is standardizing its body to the M+(5,4), right? Are they also moving to AlLi?
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Not sure if this is the proper thread and forgive me if it's been answered already (I've done some digging here and didn't find an answer), but I'm curious - what route does the Delta Mariner take from Decatur, AL. to the Cape to deliver rocket stages, etc. ?
Two ways to go
1. Tennessee river - Ohio river - Mississippi
2. Tennessee river - Tennessee–Tombigbee waterway
2. is much shorter, but depending on river conditions, the Delta Mariner can be too big to fit.
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Not sure if this is the proper thread and forgive me if it's been answered already (I've done some digging here and didn't find an answer), but I'm curious - what route does the Delta Mariner take from Decatur, AL. to the Cape to deliver rocket stages, etc. ?
Two ways to go
1. Tennessee river - Ohio river - Mississippi
2. Tennessee river - Tennessee–Tombigbee waterway
2. is much shorter, but depending on river conditions, the Delta Mariner can be too big to fit.
Ok, thanks. That helps. I was trying to follow the route from Decatur to Kentucky Lake where it struck the bridge. Curious though that it looks like there's a couple of dams along the route.
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Curious though that it looks like there's a couple of dams along the route.
With locks, I assume?
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Curious though that it looks like there's a couple of dams along the route.
With locks, I assume?
Yep, thanks. I took a closer look at the Kentucky Dam thru google earth and I see it.
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1) Would a notional Delta 4 heavy with a SSME or RS_25e core engine instead of the actual RS-68 increase or reduce payload? Assume the two outboard boosters would still be RS-68 powered. Seems the trade would be higher ISP and slightly lower weight (lighter engine but added apu's) vs the gravity loss of the longer burn time. Any other considerations?
2) What would the payload effect be of adding a second RS-68 to the outboard boosters of a Delta 4 heavy? Burn duration would obviously be reduced by half (to about 210 seconds, remakably similar to the STS SRB's). From what I can tell the current 2 booster configuration is underpowered, but would this be ovekill or just overly expensive?
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Not enough real estate for a 2nd -68, but given the PWR cost increases with the demise of the NASA SSME contract, I opined that ULA/PWR should study a dual SSME CBC. I like your idea of an SSME center CBC too. Thinking about it, though, the decrease in liftoff T/W might be a killer. You'd have to offload a lot of propellant.
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1) Would a notional Delta 4 heavy with a SSME or RS_25e core engine instead of the actual RS-68 increase or reduce payload? Assume the two outboard boosters would still be RS-68 powered. Seems the trade would be higher ISP and slightly lower weight (lighter engine but added apu's) vs the gravity loss of the longer burn time. Any other considerations?
I modeled this, in a crude way, and found that payload to LEO would likely decrease by about 1 tonne, give or take (about a 3.5-4% performance reduction) and by about 0.5 tonnes to GTO (about 5% reduction). About 88 tonnes of propellant would have to be offloaded from the boosters (combined total) to maintain the 1.2 liftoff T/W ratio. That's about 21.5% of the propellant capacity for each booster. The core's better specific impulse would make up much, but not all, of the difference.
- Ed Kyle
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1) Would a notional Delta 4 heavy with a SSME or RS_25e core engine instead of the actual RS-68 increase or reduce payload? Assume the two outboard boosters would still be RS-68 powered. Seems the trade would be higher ISP and slightly lower weight (lighter engine but added apu's) vs the gravity loss of the longer burn time. Any other considerations?
I modeled this, in a crude way, and found that payload to LEO would likely decrease by about 1 tonne, give or take (about a 3.5-4% performance reduction) and by about 0.5 tonnes to GTO (about 5% reduction). About 88 tonnes of propellant would have to be offloaded from the boosters (combined total) to maintain the 1.2 liftoff T/W ratio. That's about 21.5% of the propellant capacity for each booster. The core's better specific impulse would make up much, but not all, of the difference.
- Ed Kyle
A Delta IV Medium+ (5,4) with a Shuttle SSME for the core in place of the RS-68 would still maintain a much higher than 1.2 liftoff T/W so wouldn't need prop-unloading (which, no surprise, reduces performance).
Adding an SSME to the core would probably make the most sense (as far as maximizing payload) in the case of the variants of Delta IV Heavy proposed with extra GEMs. It would provide a lot of the benefit of cross-feeding since the SSME has a lot less thrust compared to RS-68. Adding two GEM60s to a Delta IV Heavy (with SSME core) should be plenty so no propellant unloading would be required.
It's all academic, anyway, since I supremely doubt the SSME in any variant (expendable or no) will get anywhere near as cheap as RS-68 while maintaining the same performance as the SSME currently has. Plus you have all the re-design necessary to make this worthwhile. It may well be that finishing Atlas V Heavy (and possibly even finishing wide body Centaur) would be cheaper than switching Delta IV to use partly SSMEs.
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It's all academic, anyway,
Agreed!
It may well be that finishing Atlas V Heavy (and possibly even finishing wide body Centaur) would be cheaper than switching Delta IV to use partly SSMEs.
Another cheaper alternative might be to steadily improve RS-68 performance over time.
- Ed Kyle
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Adding an RS-25 doesn't help the rest of the fleet. But let's consider the other alternatives:
AlLi: lowers the structural weight. Can be done for both the core and the US. You could get a benefit to the whole fleet. Relatively simple change, but I ignore if it would need new tooling.
Fuel Densification: Improves your mass fraction. Gives benefit to the whole fleet. Could improve a lot the results of a regen RS-68. Would need new GSE.
Regen Nozzle RS-68: Would improve the T/W of the engine, allow for more thermal environment. Might even allow for thrust of isp improvements. Again, improves the whole fleet. Probably expensive.
Better US: This is sort of needed for many reasons, would lower the costs for ULA, improve performance and give better capabilities. Improves the whole fleet. Expensive.
Solids for the Heavy: improves the payload for very little cost. Only applies to Heavy. Cheap, might require minimum pad changes.
Cross feeding: improves the payload, probably more to the higher orbits than to LEO. Applies only to Heavy. Might require minimum pad changes.
With all this things applied, you could probably go close to 55ton to 60 ton to LEO. But the really interesting part is the GSO performance. It could be as high as 16tonnes. Please note, that's GSO, not GTO.
GTO would be close to 30tonnes. That's 166% of Falcon Heavy's performance. You've gotta love H2 vehicles for high energy orbits.
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Fuel densification changes loads. It would move the loaded mixture ratio toward rich, which would lower thrust. You'd have to densify LOX too, which would be an even bigger hit on loads. Loads increases increase weight due to stronger structure needed. Plus, the hit on T/W.
The rest are DOA too for the listed reasons.
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Fuel densification changes loads. It would move the loaded mixture ratio toward rich, which would lower thrust. You'd have to densify LOX too, which would be an even bigger hit on loads. Loads increases increase weight due to stronger structure needed. Plus, the hit on T/W.
The rest are DOA too for the listed reasons.
Sad about the Better Upper Stage, though. That would be quite handy even for Atlas V launches (and could also benefit missions launched on SLS, since iCPS would likely be based off of the Delta IV upper stage).
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Fuel densification changes loads. It would move the loaded mixture ratio toward rich, which would lower thrust. You'd have to densify LOX too, which would be an even bigger hit on loads. Loads increases increase weight due to stronger structure needed. Plus, the hit on T/W.
The rest are DOA too for the listed reasons.
Couldn't you just shorten the tanks (I'm assuming both H2 and LOX densification). With the just understanding that you'd have to re qualify, certify and change lot's of GSE/Pad attachments.
In any case I was pointing out that those are just as possible as putting dual SSME on the Delta IV core. And those were proposed by Boeing.
In any case I don't see a point of talking about "improvements" for a rocket that's barely flying due to a crime (i.e. if Boeing hadn't stolen the bidding numbers of LM the Heavy would have been Atlas V and it would all have been a very different story).
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... if Boeing hadn't stolen the bidding numbers of LM the Heavy would have been Atlas V and it would all have been a very different story. ..
I'm not sure that's the way things happened. According to the following RAND report, for example,
http://www.rand.org/pubs/monographs/2006/RAND_MG503.pdf
Lockheed Martin did not defer Atlas V Heavy development until after the 1998 EELV competition. It actually "gave back" two launches to Boeing at that time, "when Lockheed Martin decided not to develop an Atlas V heavy-lift vehicle or to build a West Coast launch facility" for Atlas V.
Regardless, Boeing was soon caught (it turned itself in, didn't it?), and severely penalized to the tune of $1 billion in immediate launch re-allocations back to Atlas V, a lawsuit by Lockheed Martin, and much lost good will at the Pentagon. Lockheed Martin was awarded money to build that West Coast pad, etc., but the company never saw a need to develop its own Heavy, because 551/552 could compete for much of that business.
- Ed Kyle
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... if Boeing hadn't stolen the bidding numbers of LM the Heavy would have been Atlas V and it would all have been a very different story. ..
I'm not sure that's the way things happened. According to the following RAND report, for example,
http://www.rand.org/pubs/monographs/2006/RAND_MG503.pdf
Lockheed Martin did not defer Atlas V Heavy development until after the 1998 EELV competition. It actually "gave back" two launches to Boeing at that time, "when Lockheed Martin decided not to develop an Atlas V heavy-lift vehicle or to build a West Coast launch facility" for Atlas V.
I had understood that since Boeing had better pricing information they underbid Atlas enough to make the AH a non viable development. The fact that they got so close in performance with the 551 also helped, of course.
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... if Boeing hadn't stolen the bidding numbers of LM the Heavy would have been Atlas V and it would all have been a very different story. ..
I'm not sure that's the way things happened. According to the following RAND report, for example,
http://www.rand.org/pubs/monographs/2006/RAND_MG503.pdf
Lockheed Martin did not defer Atlas V Heavy development until after the 1998 EELV competition. It actually "gave back" two launches to Boeing at that time, "when Lockheed Martin decided not to develop an Atlas V heavy-lift vehicle or to build a West Coast launch facility" for Atlas V.
Is there a contradiction here? Boeing used their stolen data to create their bid such that they won most of the West Coast and Heavy flights. The few that Lockheed did win weren't enough to justify actually building the West Coast pad or the Atlas V Heavy, so they decided not to.
By the time Boeing got caught, they'd made enough progress that giving those launches back to Lockheed was going to add a lot of cost and slip a lot of the schedule, so they stayed with Boeing. Of course, by then Boeing was losing money hand over fist on the Delta IV and was staring down the face of a huge lawsuit from Lockheed that they were certain to lose, so they threatened to pull out of the market unless they could form ULA.
And, here we are.
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1) Would a notional Delta 4 heavy with a SSME or RS_25e core engine instead of the actual RS-68 increase or reduce payload? Assume the two outboard boosters would still be RS-68 powered. Seems the trade would be higher ISP and slightly lower weight (lighter engine but added apu's) vs the gravity loss of the longer burn time. Any other considerations?
I modeled this, in a crude way, and found that payload to LEO would likely decrease by about 1 tonne, give or take (about a 3.5-4% performance reduction) and by about 0.5 tonnes to GTO (about 5% reduction). About 88 tonnes of propellant would have to be offloaded from the boosters (combined total) to maintain the 1.2 liftoff T/W ratio. That's about 21.5% of the propellant capacity for each booster. The core's better specific impulse would make up much, but not all, of the difference.
- Ed Kyle
Thanks Ed,
To a not-so-knowledgable person, this seems to indicate that the delta 4 heavy is inefficient, in that an imaginary smaller booster with a core powered by an rs-25e would have a similar payload.
Would it be possible to model a two barrel core powered by two rs-25e and tne boosters powered by 2 rs-68's each?
My thought here is that if SLS ia cancelled, we might want a somewhat larger booster that can loft Orion and a significant payload in one launch. And if it isn't already obvious, I wish to preserve a role for the descendants of the SSME.
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You do realize that the RS-68 was designed from the ground up with lessons learned in developing the SSME, and is a rightful descendant. One of the main drivers in designing the RS-68 was reducing the manufacturing costs (compared to the SSME). The lower ISP and ablative nozzle are a direct result of that.
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Thanks Ed,
To a not-so-knowledgable person, this seems to indicate that the delta 4 heavy is inefficient, in that an imaginary smaller booster with a core powered by an rs-25e would have a similar payload.
The idea with Delta IV was to design for minimum cost, using lower-cost gas generator RS-68 engines rather than higher-cost staged combustion engines like RS-25. In addition, RS-68 was designed to be an expendable booster engine while RS-25 was designed to work largely as a reusable, high-altitude sustainer engine for Shuttle.
I don't know what these engines really cost now, but a few years ago we talked about projected RS-68 costs at $20 million versus $70 million for SSME and perhaps $25 million for RS-25. Whatever the truth, and I think it is safe to say that it is not among these numbers, it is clear that RS-25 will cost more than RS-68 despite providing only about 60% as much thrust.
Would it be possible to model a two barrel core powered by two rs-25e and tne boosters powered by 2 rs-68's each?
My thought here is that if SLS ia cancelled, we might want a somewhat larger booster that can loft Orion and a significant payload in one launch. And if it isn't already obvious, I wish to preserve a role for the descendants of the SSME.
You really would only need one more RS-25 on the core, and no additional RS-68 engines, to get full tanks and maximum results. I'll model that in a bit.
But, of course, there are ways to significantly improve Delta IV performance without tossing out its basic propulsion system.
- Ed Kyle
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You do realize that the RS-68 was designed from the ground up with lessons learned in developing the SSME, and is a rightful descendant. One of the main drivers in designing the RS-68 was reducing the manufacturing costs (compared to the SSME). The lower ISP and ablative nozzle are a direct result of that.
Precisely.
I find the idea that RS-25 will be roughly the cost of RS-68 laughable.
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The old cost numbers don't compare to today's because the SSME contract was supporting a lot of Rocketdyne fixed costs at its facilities, support functions and management. Now the other PWR engines carry those cost burdens.
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The old cost numbers don't compare to today's because the SSME contract was supporting a lot of Rocketdyne fixed costs at its facilities, support functions and management. Now the other PWR engines carry those cost burdens.
Well in the case where $30 million is tacked on to any engine to support fixed costs, then yeah, it masks the price differences.
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Could the Delta IV be used as a launcher from VAFB to resupply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
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It would have to fly a dogleg to get posigrade. It would be a huge performance hit.
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Could the Delta IV be used as a launcher from VAFB to resuply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
The orbit ATV used has no bearing on whether spacecraft can launched to the ISS from VAFB
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It would have to fly a dogleg to get posigrade. It would be a huge performance hit.
A Delta II took Jason to 66 degrees after deploying TIMED at 74.
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Gary Hudson, mentioned in one of the threads that Titan II from Vandenberg to ISS would take a 20% performance hit when he looked at it. Not the Delta IV, and performance will vary.
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Could the Delta IV be used as a launcher from VAFB to resuply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
The orbit ATV used has no bearing on whether spacecraft can launched to the ISS from VAFB
Let’s ask this a different way, discount weather etc.
We have two choices using the same launcher the Delta IV.
1) Launch from Alaska
2) Launch from VAFB
Given the same launcher, which site would give max payload (for ISS) per same fuel load etc.
Why?
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Could the Delta IV be used as a launcher from VAFB to resuply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
The orbit ATV used has no bearing on whether spacecraft can launched to the ISS from VAFB
Let’s ask this a different way, discount weather etc.
We have two choices using the same launcher the Delta IV.
1) Launch from Alaska
2) Launch from VAFB
Given the same launcher, which site would give max payload (for ISS) per same fuel load etc.
Why?
Question is moot. Can't launch Delta IV from Alaska, because the liquid hydrogen would freeze.
-Alex
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Question is moot. Can't launch Delta IV from Alaska, because the liquid hydrogen would freeze.
-Alex
Say what? Oh, I see it's April first...
Another thing is the min. possible launch inclination for Kodiak (ignoring any overflight issues) is 57 degrees, ISS is at 51.6 degrees.
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Could the Delta IV be used as a launcher from VAFB to resuply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
The orbit ATV used has no bearing on whether spacecraft can launched to the ISS from VAFB
Let’s ask this a different way, discount weather etc.
We have two choices using the same launcher the Delta IV.
1) Launch from Alaska
2) Launch from VAFB
Given the same launcher, which site would give max payload (for ISS) per same fuel load etc.
Why?
CCAFS (or KSC), presumably.
cheers, Martin
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Could the Delta IV be used as a launcher from VAFB to resuply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
The orbit ATV used has no bearing on whether spacecraft can launched to the ISS from VAFB
Let’s ask this a different way, discount weather etc.
We have two choices using the same launcher the Delta IV.
1) Launch from Alaska
2) Launch from VAFB
Given the same launcher, which site would give max payload (for ISS) per same fuel load etc.
Why?
Depends on where in Alaska your launch site is and what you are launching on top of it. If you are launching from somewhere in the aluetians where you can head east without overflying land and can carry an upper stage wiith enough fuel to do the required plane change in orbit, then you could launch a lot more from there than from VAFB.
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Could the Delta IV be used as a launcher from VAFB to resuply the ISS?
Was thinking of the Orbit the ATV used, with that in mind would it be possible?
The orbit ATV used has no bearing on whether spacecraft can launched to the ISS from VAFB
Let’s ask this a different way, discount weather etc.
We have two choices using the same launcher the Delta IV.
1) Launch from Alaska
2) Launch from VAFB
Given the same launcher, which site would give max payload (for ISS) per same fuel load etc.
Why?
Depends on where in Alaska your launch site is and what you are launching on top of it. If you are launching from somewhere in the aluetians where you can head east without overflying land and can carry an upper stage wiith enough fuel to do the required plane change in orbit, then you could launch a lot more from there than from VAFB.
Only thought atm of posting of Kodiak. The question however was a general location question, so thx for you info.
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How does the delta iv get erected and put in the launch pad as all of the photos I see is when it is less than 50° from the ground?
EDIT- after rereading it some times I mean how does the erector lift the rocket above the posts that stick up. From the photos I have found the posts must go down?
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How does the delta iv get erected and put in the launch pad as all of the photos I see is when it is less than 50° from the ground?
EDIT- after rereading it some times I mean how does the erector lift the rocket above the posts that stick up. From the photos I have found the posts must go down?
What are the "posts"?
The rocket is already part way up the erector. The pivot point is several feet below the engine nozzles.
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How does the delta iv get erected and put in the launch pad as all of the photos I see is when it is less than 50° from the ground?
EDIT- after rereading it some times I mean how does the erector lift the rocket above the posts that stick up. From the photos I have found the posts must go down?
What are the "posts"?
The rocket is already part way up the erector. The pivot point is several feet below the engine nozzles.
http://www.patrick.af.mil/shared/media/photodb/photos/111011-F-vs513-0025.jpg
Computer is not working with me so I will post a link
The posts are the grey things in the bottom right corner.
Also how do they drop the rocket down as in the bottom left of this next photo there seems to be some support does that get taken away or cut off?
http://www.kcet.org/updaily/socal_focus/delta-iv-launch.jpg
Sorry for so many edits I am still waking up ( I am in Australia)
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From today's launch (attempt?), what's a SMTU battery?
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What is the Delta IV heavy LEO payload with the new upgraded engines? What would it's payload be with a 5m upgraded Centaur second stage if it is ever developed?
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I think that the payload to LEO with the four-engine Common Centaur and 6 GEMs is in the 40-50 metric tonne range. I don't know if the new RL-10 gives any significant change in overall performance as it has lower Isp than RL-10B-2 but better thrust.
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I think that the payload to LEO with the four-engine Common Centaur and 6 GEMs is in the 40-50 metric tonne range. I don't know if the new RL-10 gives any significant change in overall performance as it has lower Isp than RL-10B-2 but better thrust.
I think the OP was referring to the RS-68A, not a new RL-10 (but I could be wrong). Where are people getting performance for this RL-10? There was discussion about a new RL-10 over on the PWR sale thread, too. Seems odd that anyone would build a new engine with lower Isp but higher thrust, since Isp is more important for the upperstage (although thrust is probably more important for LEO).
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I'm not sure but, IIRC, the variant is called RL-10C and it's supposed to be 30klbf and 450s vacuum impulse. PWR have been advertising it for a while but only recently has DoD shown any interest - the New Generation Engine (NGE) project. RL-10C isn't certain to win but you've got to think it has an advantage.
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Greater LEO payload assumes that it is not limited by the structural limits of the vehicle.
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I'm not sure but, IIRC, the variant is called RL-10C and it's supposed to be 30klbf and 450s vacuum impulse. PWR have been advertising it for a while but only recently has DoD shown any interest - the New Generation Engine (NGE) project. RL-10C isn't certain to win but you've got to think it has an advantage.
RL-10C isn't NGE, and it's not 30klbf.
Suspect that's the target thrust for NGE, which is at proposal stage at the moment I think.
cheers, Martin
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I was refering to the RS-68A. Currently the Delta IV heavy in standard configuration is 29 tons to LEO. Just wondering how much increase with the new engines. Then with the new engines and a new upper stage. Without Gems. Then maybe with Gems. I was thinking it has greater capability for LEO and L1 or L2 station building capability with the 5m + width of payload. If things don't change soon, I don't see SLS being built anytime soon. Just wanted to know if we could use the Delta IV heavy for some space exploration with equipment and spacecraft launched to dock with using commercial. then go beyond LEO without SLS.
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Just wanted to know if we could use the Delta IV heavy for some space exploration with equipment and spacecraft launched to dock with using commercial. then go beyond LEO without SLS.
Sure you can. Centaur would be a good EDS for small payloads, and that's all you really need if you use Lagrange point staging. It's a bit small, too small for Orion, but not unworkably small. The Centaur-derived ACES would be excellent, but plain old Centaur would do.
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I was refering to the RS-68A. Currently the Delta IV heavy in standard configuration is 29 tons to LEO. Just wondering how much increase with the new engines. Then with the new engines and a new upper stage. Without Gems. Then maybe with Gems. I was thinking it has greater capability for LEO and L1 or L2 station building capability with the 5m + width of payload. If things don't change soon, I don't see SLS being built anytime soon. Just wanted to know if we could use the Delta IV heavy for some space exploration with equipment and spacecraft launched to dock with using commercial. then go beyond LEO without SLS.
RS-68 is around 23 tonnes to LEO, RS-68A was expected to be around 29tonnes to LEO. It's anyone's guess at this time.
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RS-68A was expected to be around 29tonnes to LEO.
If it's that much, you could even use a 5m Delta upper stage. You couldn't get it to orbit fully fueled, but you could get it to orbit with significantly more propellant than a Centaur.
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I was thinking if they built a 5m upper stage on the Centaur design for efficiency using a common bulkhead design instead of separate tanks. Would that improve payload? Then when you add solids or have crossfeed to central core I was wondering if this could be done incrementally to get to 50 tons?
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I was thinking if they built a 5m upper stage on the Centaur design for efficiency using a common bulkhead design instead of separate tanks. Would that improve payload? Then when you add solids or have crossfeed to central core I was wondering if this could be done incrementally to get to 50 tons?
Congrats, you just describes the next-gen upper stage being developed by ULA, ACES:
http://www.ulalaunch.com/site/docs/publications/UpperStageEvolutionJPC2009.pdf
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How soon will the new upper stage be developed? Seems like if this takes place, SLS will not be needed. Money spent there could be used for in space equipment to get the show on the road.
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How soon will the new upper stage be developed? Seems like if this takes place, SLS will not be needed. Money spent there could be used for in space equipment to get the show on the road.
ACES was a proposal, several years old now. It isn't being developed to my knowledge. SLS may end up with an in-space stage that looks a bit like a version of ACES, but it hasn't been put up for bids (won't for awhile now) and once competed there is no guarantee that ULA's (Lockheed Martin's) balloon tank concept would be the winner. Boeing has proposed a rigid tank alternative.
- Ed Kyle
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ULA is working on the Integrate Fluid Vehicle concept. Is cooperating on with XCOR with a piston pumped H2/LOX rocket engine, and has contracted an auto manufacturer for a four piston GH2/GOX engine for electricity generation. So I don't know if exactly ACES, but a pretty close thing is being worked on. Since there's no extra money, it's being done slow. But they are working. I remember Downix stating that ULA would work on the Delta IV Fleet Standardization Program (CDR on October), then on ACES, and then on Atals V Phase 2. Since they are investing out of their own pocket, they are doing it slow. So ACES might be a 2020 IOC, and who know about AVP2.
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ULA is working on the Integrate Fluid Vehicle concept. Is cooperating on with XCOR with a piston pumped H2/LOX rocket engine, and has contracted an auto manufacturer for a four piston GH2/GOX engine for electricity generation. So I don't know if exactly ACES, but a pretty close thing is being worked on. Since there's no extra money, it's being done slow. But they are working. I remember Downix stating that ULA would work on the Delta IV Fleet Standardization Program (CDR on October), then on ACES, and then on Atals V Phase 2. Since they are investing out of their own pocket, they are doing it slow. So ACES might be a 2020 IOC, and who know about AVP2.
That is what they have detailed in the upgrade path document I linked to above anyways. Ed is right that Boeing did propose a rigid tank alternative, mind you, but until the fleet standardization is completed, I would not expect either path to be explored.
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This may be a crazy question but with an increase of 50,000 lbs of thrust with the new RS-68A and the desire to use the smaller second stage from the Atlas saving about 10,000 lbs how hard would it be to add another 30,000-70,000 lbs to the first stage?
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This thread is a much better read than the Atlas Q&A....more space launch drama...ehhehe.
Not found my question asked so here goes.
Did the Centaur G Prime Influence the Second stage design of the Delta IV?
The General looks are of a "clone" with different materials.
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This thread is a much better read than the Atlas Q&A....more space launch drama...ehhehe.
Not found my question asked so here goes.
Did the Centaur G Prime Influence the Second stage design of the Delta IV?
The General looks are of a "clone" with different materials.
No. The HDCSS was an upgrade of the DCSS of Delta III. The Centaur G Prime had a common bulkhead and HDCSS has separate tanks connected by a truss. Plus they were designed by different contractors.
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With the Fleet Consolidation Program, will the DIVUS get Centaur avionics? Will the Delta IV Heavy get the new unified avionics and the Atlas V fairing?
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This thread is a much better read than the Atlas Q&A....more space launch drama...ehhehe.
Not found my question asked so here goes.
Did the Centaur G Prime Influence the Second stage design of the Delta IV?
The General looks are of a "clone" with different materials.
No. The HDCSS was an upgrade of the DCSS of Delta III. The Centaur G Prime had a common bulkhead and HDCSS has separate tanks connected by a truss. Plus they were designed by different contractors.
Maybe different contractors, but not necessarily different people. However, there are a number of factors that result in the designs looking similar, even with different structural baselines.
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Maybe different contractors, but not necessarily different people. However, there are a number of factors that result in the designs looking similar, even with different structural baselines.
Different people too. The fat Centaur goes back 20-30 years before Delta iV
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From a question on the L-15 thread, RS-68 does not test fire with its ablative nozzle - so no need to replace between tests.
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On this web page:
http://www.boeing.com/defense-space/space/bss/factsheets/702/702b.html
where Boeing lists the launch systems compatible with the 702MP, why do they mention, "Atlas, Ariane, Proton and Sea Launch," but not Delta IV?
Is there really some requirement imposed by the 702MP which Delta IV cannot meet? That seems very odd, as it was essentially a requirement that the two EELV systems have very similar capabilities....
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Well, before ULA Boeing did withdraw the Delta IV from the commercial market, and more likely than not post ULA any EELV launch would fly on an Atlas. Just an educated guess.
To be fair Falcon 9 and Falcon Heavy are also not yet on the list. Though I suspect a Falcon 9 does not currently have the needed performance.
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Well, before ULA Boeing did withdraw the Delta IV from the commercial market
Ah, good point! Thanks for that reminder. The short version is probably, "It's a marketing thing." ;D
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On this web page:
http://www.boeing.com/defense-space/space/bss/factsheets/702/702b.html
where Boeing lists the launch systems compatible with the 702MP, why do they mention, "Atlas, Ariane, Proton and Sea Launch," but not Delta IV?
Is there really some requirement imposed by the 702MP which Delta IV cannot meet? That seems very odd, as it was essentially a requirement that the two EELV systems have very similar capabilities....
The only Delta able to lift a 6,215 kg Intelsat 27 to the planned Zenit 3SL insertion orbit would be Delta 4 Heavy, which may cost a half-billion dollars. So there you go.
- Ed Kyle
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The only Delta able to lift a 6,215 kg Intelsat 27 to the planned Zenit 3SL insertion orbit would be Delta 4 Heavy, which may cost a half-billion dollars. So there you go.
- Ed Kyle
What about the Delta M(5,4)?
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The only Delta able to lift a 6,215 kg Intelsat 27 to the planned Zenit 3SL insertion orbit would be Delta 4 Heavy, which may cost a half-billion dollars. So there you go.
- Ed Kyle
What about the Delta M(5,4)?
Not to the planned 0 deg. inclination GTO...
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The only Delta able to lift a 6,215 kg Intelsat 27 to the planned Zenit 3SL insertion orbit would be Delta 4 Heavy, which may cost a half-billion dollars. So there you go.
- Ed Kyle
What about the Delta M(5,4)?
Not to the planned 0 deg. inclination GTO...
Neither can Atlas, Ariane, or Proton
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The only Delta able to lift a 6,215 kg Intelsat 27 to the planned Zenit 3SL insertion orbit would be Delta 4 Heavy, which may cost a half-billion dollars. So there you go.
- Ed Kyle
What about the Delta M(5,4)?
Not to the planned 0 deg. inclination GTO...
Neither can Atlas, Ariane, or Proton
How about comparing with the same delta-v to GTO?
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The only Delta able to lift a 6,215 kg Intelsat 27 to the planned Zenit 3SL insertion orbit would be Delta 4 Heavy, which may cost a half-billion dollars. So there you go.
- Ed Kyle
What about the Delta M(5,4)?
Not to the planned 0 deg. inclination GTO...
Right. Delta 4M+5,4 can only lift 5.85 tonnes to an equivalent energy 0 deg inclination standard GTO orbit (1,500 m/s delta-v from GEO) - at least the one with a basic RS-68. Both of the two most powerful Atlas 5 variants could lift an Intelsat 27 though.
Of course if you could launch a "5,4" from the equator, it could do the job.
- Ed Kyle
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I was thinking of something else. What's with the rarity of RocketCams on the Delta IV's first/second stages? If the EFT-1 launches on time, I've yet to see some RocketCams.
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I was thinking of something else. What's with the rarity of RocketCams on the Delta IV's first/second stages?
Nobody wants to pay for them
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
The deployable extension is connected in one piece, but it consists of two parts (called the "B-cone" and "C-cone"). A third nozzle extension (called the "A-cone") is fixed to the main nozzle. The fixed versus extended division point is seen as the broad black band in the attached image.
- Ed Kyle
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
The deployable extension is connected in one piece, but it consists of two parts (called the "B-cone" and "C-cone"). A third nozzle extension (called the "A-cone") is fixed to the main nozzle.
- Ed Kyle
Thanks Ed, I thought it might have been something like that as a two-piece deployment mechanism would have introduced as additional failure modes.
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
The deployable extension is connected in one piece, but it consists of two parts (called the "B-cone" and "C-cone"). A third nozzle extension (called the "A-cone") is fixed to the main nozzle.
- Ed Kyle
Thanks Ed, I thought it might have been something like that as a two-piece deployment mechanism would have introduced as additional failure modes.
RL-10C, created from rebuilt excess RL-10B-2 engines, will, as I understand it, dispense with the B and C-cone extensions, eliminating the entire extension sequence.
- Ed Kyle
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
The deployable extension is connected in one piece, but it consists of two parts (called the "B-cone" and "C-cone"). A third nozzle extension (called the "A-cone") is fixed to the main nozzle.
- Ed Kyle
Thanks Ed, I thought it might have been something like that as a two-piece deployment mechanism would have introduced as additional failure modes.
RL-10C, created from rebuilt excess RL-10B-2 engines, will, as I understand it, dispense with the B and C-cone extensions, eliminating the entire extension sequence.
- Ed Kyle
Interesting, so they're going with a fixed nozzle extension? I wonder how all of it is going to fit into the interstage adapter, or is there enough space for the entire nozzle?
Edit:
You don't happen to know the base diameter of each section do you?
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Interesting, so they're going with a fixed nozzle extension? I wonder how all of it is going to fit into the interstage adapter, or is there enough space for the entire nozzle?
Edit:
You don't happen to know the base diameter of each section do you?
RL10C will fly on Atlas V/Centaur, which has enough interstage length. I'm not sure if it will fly on Delta IV.
Here's the information that I have on hand.
The A-Cone fixed section is 22.1 in. (570 mm) long. The combined B/C-Cone translating section is 77.2 in. ( 2082 mm) long. The three cones increase the expansion ratio from 77:1 at the regenerative primary nozzle to 285:1. The primary renerative nozzle is 44 in. (1113 mm) in diameter. At the connection point between the A and B Cones (connection between fixed and translating nozzles) the diameter is 57 in. (1445 mm). At the exit plane the maximum diameter is 84 in. (2136 mm).
Snecma Moteurs of Le Haillan, France built the nozzle extensions.
Here's an image with some more dimensions.
- Ed Kyle
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
The deployable extension is connected in one piece, but it consists of two parts (called the "B-cone" and "C-cone"). A third nozzle extension (called the "A-cone") is fixed to the main nozzle.
- Ed Kyle
Thanks Ed, I thought it might have been something like that as a two-piece deployment mechanism would have introduced as additional failure modes.
RL-10C, created from rebuilt excess RL-10B-2 engines, will, as I understand it, dispense with the B and C-cone extensions, eliminating the entire extension sequence.
- Ed Kyle
That is only for Atlas. Delta will still use the extensions.
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The nozzle extension used on the RL10B-2 engine on the DCSS, is it a two-piece extension or just one? From various photos I have found of it in the stowed config, it looks like it's a two-piece design rather than a single-piece design.
The deployable extension is connected in one piece, but it consists of two parts (called the "B-cone" and "C-cone"). A third nozzle extension (called the "A-cone") is fixed to the main nozzle.
- Ed Kyle
Thanks Ed, I thought it might have been something like that as a two-piece deployment mechanism would have introduced as additional failure modes.
RL-10C, created from rebuilt excess RL-10B-2 engines, will, as I understand it, dispense with the B and C-cone extensions, eliminating the entire extension sequence.
- Ed Kyle
Interesting, so they're going with a fixed nozzle extension? I wonder how all of it is going to fit into the interstage adapter, or is there enough space for the entire nozzle?
Atlas already uses a fixed nozzle extension.
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Does anyone have a photo of the RL-10B-2 with the thermal shield installed? Does it cover the entire exterior pipes/boxes or just part of them?
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Details of the RL10 plans from the new Delta IV users guide (page 244)...
8.2.1 RL10C-2 2nd Stage Engine Upgrade
To improve commonality between the Atlas and Delta launch vehicles, ULA and Pratt & Whitney Rocketdyne (PWR) are currently developing the RL10C-1 engine for the Centaur upper stage of the Atlas launch vehicle. This engine uses similar chamber and nozzle configuration as the RL10B-2 engine currently used on Delta. Use of this common engine allows for future upgrades to the RL10B-2 engine, to be called the RL10C-2 (Figure 8-2).
The RL10C-2 engine will incorporate all improvements from the RL10C-1, including an upgraded redundant ignition system to improve reliability, changes to the engine plumbing to improve starting operations, a propellant valve design update, and a number of improvements previously qualified under the Assured Access to Space program including a revised gear train and seal improvements.
The RL10C-2 development will be managed through
the RL10 Sustainment and Modernization Program.
This program is intended to incorporate improved manufacturing methods for turbomachinery, propellant valves, and injector hardware, revised large plumbing to reduce weight, and more robust solenoid valves. Additionally, the RL10C-2 is intended to be qualified to operate with active Mixture Ratio control, a capability available on Atlas/Centaur missions dating back to 1965. This feature, enabled on Delta IV by the addition of Common Avionics (Section 8.3.2), could result in a performance improvement of up to 200 lb for certain Delta missions. The RL10C-2 will continue to use the 3-segment extendible nozzle currently used on the RL10B-2. The C-2 will look virtually the same as an RL10B-2, with slight changes to the Ignition and Engine Instrumentation Boxes and realignment of some of the large plumbing.
Changes incorporated as part of the Sustainment and Modernization effort will be qualified for both the RL10C-1 for Atlas and the RL10C-2 for Delta at the same time, using the same common core engine. The end result will be an engine that can be built and acceptance tested using a common bill of material and test program, and then configured as necessary with bolt-on hardware to support either Atlas V or Delta IV vehicles.
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Anyone have a general idea of what the weight of the LH2 tankage of the Delta IV is?
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Could DIV be made to use the Atlas V solid boosters?
If so how many per core could be used.?
With maximum amount of Atlas V solid boosters on a DIV core what LEO payload mass could this get?
Note for questions:
This is all for tech and not political.
OK if it needs a new pad for this idea.
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My guess is that you could.. while doing Delta V. Those SRB are not thrust vectored, like the GEM-60. I don't know if the RS-68A has enough control authority to handle the Atlas V's SRB. Then they have more thrust, so you'd probably need reinforcements in the tanks. And let's remember that the big work was actually having a common core. Thus, if you reinforce the structure for bigger SRB, you add weight and take out performance of the Medium.
I'm pretty sure that they could also try to make a composite H2 tank out of Boeing's 5.5m demonstrator (would have to adapt the tooling to 5.2m, of course). And that will increase performance a lot. But then you are so far from a Delta IV that's probably simply too expensive. ULA has always talked about an Atlas V Phase II.
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Anyone have a general idea of what the weight of the LH2 tankage of the Delta IV is?
Rough guess?
((stage burnout mass) - (mass of engines)) * (29 / (12 + 29))
((27 t) - (7 t)) * 0.71 = 14.2 t
The LOX tankage would be (20 - 14.2 = 5.8 t.
You probably wonder where 29 and 12 come from. STS ET.
http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/et.html
The ratio isn't going to be exactly correct, since RS-68 and RS-25 have different O/F ratios. But hey, you asked for a general idea...! Maybe the data is available from Boeing somewhere?
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How come ULA does not have as many Delta IV launches planned as they have of Atlas V? Delta has only three for 2015 and in that same year, Atlas has nine including MUOS 3 seventy-two hours ago. I understand that the workhorse Delta II has only one this year because it is on the verge of retirement.
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ULA is, AIUI, planning on retiring the Delta IV as well, because it is far more expensive to produce, prepare and launch than the Atlas V, and gives little in extra performance. Even the D-IVH can't outperform a fully loaded-out Atlas V 500 series to the extent needed to maintain it as a viable commercial launch vehicle.
At least, that's the rationale I've heard.
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How come ULA does not have as many Delta IV launches planned as they have of Atlas V? Delta has only three for 2015 and in that same year, Atlas has nine including MUOS 3 seventy-two hours ago. I understand that the workhorse Delta II has only one this year because it is on the verge of retirement.
A. it is more expensive
b. it has a lower launch rate capability
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Even the D-IVH can't outperform a fully loaded-out Atlas V 500 series to the extent needed to maintain it as a viable commercial launch vehicle.
Delta 4 Heavy can lift about 1.7 times as much to GTO as Atlas 5-551. 11.4 versus 6.7 tonnes to GEO-1,500 m/s - more than any other rocket in the world. That seems substantial to me.
- Ed Kyle
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Holy smokes! I didn't know the Delta IV was near retirement! :(
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Holy smokes! I didn't know the Delta IV was near retirement! :(
It probably has another decade or more of life. ULA plans to develop a new first stage powered by BE-4 engines that will replace the existing Atlas 5 first stage for EELV Medium work, but the company cannot use this new rocket to replace Delta 4 Heavy until it also develops a new upper stage to replace Centaur. No substantial work is underway on the new upper stage to my knowledge, which means Heavy is going to be needed for years.
- Ed Kyle
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Sorry for the confusion, didn't someone at ULA or Boeing recently say that Delta IV production would be stepped up to protect against Atlas V engine availability issues? Is that still that case? Thanks.
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Holy smokes! I didn't know the Delta IV was near retirement! :(
...
No substantial work is underway on the new upper stage to my knowledge, which means Heavy is going to be needed for years.
- Ed Kyle
For which payloads? NROL-37 is the only one baselined for DIVH, and there are rumours, that it has been cancelled.
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Holy smokes! I didn't know the Delta IV was near retirement! :(
...
No substantial work is underway on the new upper stage to my knowledge, which means Heavy is going to be needed for years.
- Ed Kyle
For which payloads? NROL-37 is the only one baselined for DIVH, and there are rumours, that it has been cancelled.
The block buy was said to include four Delta 4 Heavies. That's for 2017 to maybe 2019-20.
- Ed Kyle
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If it's chosen when would the Delta IVH for the Solar Probe Plus mission be ordered?
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Hi, i have some questions about the Delta 4:
Why is it being retired instead of being evolved into a less costly rocket? Basically ULA has this great vehicle and are going to just throw it away! F9 had to evolve to meat the commercial market, why not D4?
Gotta be cheaper than developing Vulcan from scratch^.^
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In reality, Vulcan is that evolution.
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Hi, i have some questions about the Delta 4:
Why is it being retired instead of being evolved into a less costly rocket? Basically ULA has this great vehicle and are going to just throw it away! F9 had to evolve to meat the commercial market, why not D4?
Gotta be cheaper than developing Vulcan from scratch^.^
The company that builds Delta 4 already has a lower-cost alternative for all but the Heavy missions (Atlas 5). Last year it considered increasing the Delta 4 production rate to offset the RD-180 issue, but soon switched to the plan to develop Vulcan instead. Presumably Vulcan will cost less than Delta 4 over its life, including development costs.
Atlas 5/Vulcan already has a better upper stage (Centaur). Delta 4 would need a vastly improved first stage (cost and performance) to survive. That means a new main engine, which means just as much development as Vulcan.
- Ed Kyle
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Hi, i have some questions about the Delta 4:
Why is it being retired instead of being evolved into a less costly rocket? Basically ULA has this great vehicle and are going to just throw it away! F9 had to evolve to meat the commercial market, why not D4?
Gotta be cheaper than developing Vulcan from scratch^.^
Delta IV is not that great. You can't get Atlas V 431/531/541/551 performance with a single stick. The integration process is longer and more expensive. Also, it has less launch rate per pad. It needs two different upper stages and there's a reason that the Common Upper Stage is based on Centaur. Its avionics were replaced with Atlas'. It had a lot of different fairings. You need at least four different cores models to offer the full performance design. It's not human rated. And its engine is ablatively cooled and thus has no reuse potential. Also its manufacturer has a lot of overhead because it was designed to mount on the SSME infrastructure. And there are no equivalent engines in the world to replace it with like they are able to with the RD-180.
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Delta IV is not that great. You can't get Atlas V 431/531/541/551 performance with a single stick. The integration process is longer and more expensive. Also, it has less launch rate per pad. It needs two different upper stages and there's a reason that the Common Upper Stage is based on Centaur. Its avionics were replaced with Atlas'. It had a lot of different fairings. You need at least four different cores models to offer the full performance design. It's not human rated. And its engine is ablatively cooled and thus has no reuse potential. Also its manufacturer has a lot of overhead because it was designed to mount on the SSME infrastructure. And there are no equivalent engines in the world to replace it with like they are able to with the RD-180.
Keep in mind the Areojet SRBs on the Atlas have about seven and a half tons more propellant then the GEM 60s on the Delta IV M+ and may be responsible for a good deal of the performance difference.
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In this image of the WGS-9 launch you can see that the Delta IV boosters have different aft skirts.
Why is that?
https://twitter.com/johnkrausphotos/status/843299763747479553 (https://twitter.com/johnkrausphotos/status/843299763747479553)
(https://pbs.twimg.com/media/C7QAp72UwAATtqn.jpg)
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In this image of the WGS-9 launch you can see that the Delta IV boosters have different aft skirts.
Why is that?
https://twitter.com/johnkrausphotos/status/843299763747479553 (https://twitter.com/johnkrausphotos/status/843299763747479553)
One pair has a fixed nozzle, the other pair has a gimballed nozzle. I don't know which is which.
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Ah, thanks!
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In this image of the WGS-9 launch you can see that the Delta IV boosters have different aft skirts.
Why is that?
https://twitter.com/johnkrausphotos/status/843299763747479553 (https://twitter.com/johnkrausphotos/status/843299763747479553)
One pair has a fixed nozzle, the other pair has a gimballed nozzle. I don't know which is which.
The one will the longer skirt has TVC
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Thanks again.
Once you know it it's quite obvious how different the exhaust directions are for the different boosters.
Do the ones with TVC usually fire at less of an angle at liftoff? To decrease cosine losses or to limit the effect on the pad?
Is the impact of plume impingement lower at liftoff than at altitude so that they can simply do it?
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Has anyone ever published good, worked-out performance figures for a version of the single-stick with 8x GEM-60 solids, aluminum-lithium structures and the 5-meter D-IVH upper stage? And with a regenerative RS-68 prospect and all the above gear? :) I'm willing to bet performance closing in on the Delta IV-Heavy, even with the standard RS-68A engine, and cheaper than the Heavy to boot.
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Has anyone ever published good, worked-out performance figures for a version of the single-stick with 8x GEM-60 solids, aluminum-lithium structures and the 5-meter D-IVH upper stage? And with a regenerative RS-68 prospect and all the above gear? :) I'm willing to bet performance closing in on the Delta IV-Heavy, even with the standard RS-68A engine, and cheaper than the Heavy to boot.
Sorry I don't know the answer to your question but am responding anyway! ;)
I think maybe no serious consideration has been given to 8x solids on DIV because one or more of:
- mounting 8x solids would require somewhat major modifications to the pads;
- supporting the loads would require a redesign of the core, interstage, upper stage, etc.
- DIV-H exists;
- Atlas V exists;
- Vulcan is in development.
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Delta with 8x solids should outperform Atlas V-551. I was recently told by someone who knows these things that 8x GEM-60 solids would be the most that could be accommodated with Delta on the corestage without major redesign, and even then the corestage engine would likely have to run a different throttling profile. The 8x solid concept was looked at early in the Delta's career but was abandoned early on. Anyway, if they ever were to continue with Delta; 4x better solids such as the GEM-63XL would suffice over 8x GEM-60s. But with Vulcan on the horizon, the mighty RS-68 will be retired and the Delta IV tooling heritage will live on with Vulcan.
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But with Vulcan on the horizon .... Delta IV tooling heritage will live on with Vulcan.
Maybe. Maybe not. Some hints have suggested that Vulcan diameter could grow fatter than Delta 4, to as much as 5.4 meters, the same as the payload fairing. Meanwhile, a 2 x AR-1 version, if that should happen, would likely be 3.81 meters (150 inches) in diameter, the same as the Atlas 5 CCB.
- Ed Kyle
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How many pyro bolts does the Delta IV Common Booster Core use for separation from the Delta CSS?
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I noticed during the recent NROL launch that the Heavy seems to ascend with the boosters "top and bottom" (I assume to reduce aerodynamic forces) and then rolls "wings level" for booster jettison. But then it appears to roll 45 degrees after jettison and through second stage ignition. Why is that?
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How many pyro bolts does the Delta IV Common Booster Core use for separation from the Delta CSS?
none, they don't use pyro bolts there
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Well, then: what do they use?