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#40 by Propforce on 16 Aug, 2007 23:04
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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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#41 by tnphysics on 16 Aug, 2007 23:20
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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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#42 by Jim on 16 Aug, 2007 23:38
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There is a chart on the Boeing website with it
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#43 by Nick L. on 17 Aug, 2007 02:37
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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. -
#44 by Propforce on 17 Aug, 2007 20:54
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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.QuotePropforce - 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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#45 by yinzer on 17 Aug, 2007 21:11
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Propforce - 17/8/2007 1:54 PM
Quoteedkyle99 - 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? -
#46 by edkyle99 on 17 Aug, 2007 21:33
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Propforce - 17/8/2007 3:54 PM
Quoteedkyle99 - 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.QuoteQuotePropforce - 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 -
#47 by Propforce on 18 Aug, 2007 04:16
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yinzer - 17/8/2007 2:11 PM
QuotePropforce - 17/8/2007 1:54 PM
Quoteedkyle99 - 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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#48 by Propforce on 18 Aug, 2007 04:20
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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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#49 by tnphysics on 18 Aug, 2007 14:36
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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? -
#50 by edkyle99 on 22 Aug, 2007 01:28
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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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#51 by Propforce on 22 Aug, 2007 01:43
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Thanks Ed,
I'll take a look and see if I can provide helpful comments.
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#52 by tnphysics on 23 Aug, 2007 11:27
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Propforce - 14/8/2007 3:04 PM
Quotetnphysics - 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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#53 by tnphysics on 29 Aug, 2007 00:33
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Why do the Delta IV upgrade options give such a dramatic boost in payload?
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#54 by tnphysics on 03 Sep, 2007 03:37
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Why did the Delta IV go with a cryogenic first stage?
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#55 by Nick L. on 03 Sep, 2007 07:23
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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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#56 by tnphysics on 03 Sep, 2007 14:00
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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? -
#57 by tnphysics on 03 Sep, 2007 14:06
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Why can't the Delta IV put as much in LEO as the Atlas V?
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#58 by Jim on 03 Sep, 2007 15:15
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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 -
#59 by CFE on 03 Sep, 2007 17:26
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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.
