NASA changing opinion on the Direct HLV launcher.

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edkyle99

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vulture4":2gb7bvst said:
Both SpaceX and Boeing have proposed evolutionary variants of the Delta and Falcon up to 100mt to LEO. I don't think it's coincidence that even though the designs are completely different both use all-liquid propulsion. So if NASA was really interested in actually launching a heavy payload they could just issue an RFP for an HLLV and leave the design up to the contractor. But that's not the goal here at all. NASA is being forced by Congress to create jobs at MSFC and in Utah. Whether the rocket ever carries a payload is unimportant.
HEFT reported that it would cost too much to develop the high thrust kerolox engine needed for the Shuttle-Derived alternative HLV. The only way to fund it would be for NASA to share costs with the Pentagon, but the Pentagon is not interested.

As for "creating jobs in Utah", etc., a kerosene rocket would create those jobs (probably even more of them) somewhere else and would cost more money. I just don't understand the oft-repeated ATK hatespeak.

According to HEFT, the SRBs would cost far less than the core and its RS-25 engines. For the 70 tonner, the core cost to first flight (not including the cost of the first flight!) would be $5.6 billion, twice as much as the SRB cost. A kerolox stage would cost more than either. A cryo upper stage would cost nearly as much as the SRBs. Those relatively affordable SRBs produce 6.4 million pounds of thrust. Where else could NASA get so much power for that amount of money?

- Ed Kyle
 
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exoscientist

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edkyle99":vhrhjaza said:
HEFT reported that it would cost too much to develop the high thrust kerolox engine needed for the Shuttle-Derived alternative HLV. The only way to fund it would be for NASA to share costs with the Pentagon, but the Pentagon is not interested...

- Ed Kyle
In which HEFT study is the comparison made to a kerolox fueled lower stage? The one I saw posted on NasaWatch only assessed SRB and SSME powered lower stages.

Bob Clark
 
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edkyle99

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exoscientist":1c7po5we said:
edkyle99":1c7po5we said:
HEFT reported that it would cost too much to develop the high thrust kerolox engine needed for the Shuttle-Derived alternative HLV. The only way to fund it would be for NASA to share costs with the Pentagon, but the Pentagon is not interested...

- Ed Kyle
In which HEFT study is the comparison made to a kerolox fueled lower stage? The one I saw posted on NasaWatch only assessed SRB and SSME powered lower stages.

Bob Clark
It is discussed in the report posted on nasawatch, on page 31, where the following note appears:
"An RP‐based HLV (100‐120 t) and a replacement for the
(Russian) RD‐180 is higher cost to NASA and therefore
requires supplemental funding from DoD to offset increased
costs."

The rocket concepts themselves were covered in NASA's "Heavy Lift Launch Vehicle Study" published in, I believe, May of this year. One was "Evolved Atlas", or whatever NASA called it, which used an ET diameter tank augmented by Atlas 5 CCBs. Another was Falcon XX from SpaceX.

- Ed Kyle
 
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EarthlingX

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When i see a heavy lifter, which doesn't use Apollo infrastructure, which costs 2 500 000 000$/year without doing anything, it might be worth discussion.

Actually, there are, using operational vehicles (Delta IV, Atlas V, Falcon 9) and their upgrades, very likely much cheaper than SDHLV variants.
 
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exoscientist

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edkyle99":2jtbtm3l said:
exoscientist":2jtbtm3l said:
In which HEFT study is the comparison made to a kerolox fueled lower stage? The one I saw posted on NasaWatch only assessed SRB and SSME powered lower stages.

Bob Clark
It is discussed in the report posted on nasawatch, on page 31, where the following note appears:
"An RP‐based HLV (100‐120 t) and a replacement for the
(Russian) RD‐180 is higher cost to NASA and therefore
requires supplemental funding from DoD to offset increased
costs."
The rocket concepts themselves were covered in NASA's "Heavy Lift Launch Vehicle Study" published in, I believe, May of this year. One was "Evolved Atlas", or whatever NASA called it, which used an ET diameter tank augmented by Atlas 5 CCBs. Another was Falcon XX from SpaceX.

- Ed Kyle
Thanks. That seems to be based on assuming you are also developing the new heavy thrust kerosene engine. As a stop gap we could just use the RD-180 engines. At $10 million these are half the cost of the RS-68 expendable hydrogen engines, and probably a third of the cost of the expendable versions of the SSME's.
Also, considering that SpaceX has demonstrated it is able to provide launchers at half the cost of the big aerospace companies, we could cut cost significantly by using them to build the booster structures while using the RD-180 engines.
There is also the fact that both the Obama administration and Congress want the new heavy thrust kerosene engine so would likely provide the funding for it anyway. And the development cost for it really would not be that great. This article said the development cost for two different heavy thrust engines would only be $1.3 billion:

TICKET TO RIDE.
"Potential replacements for the Space Shuttle are taking shape as NASA struggles to finalise the requirements for a second-generation reusable launch vehicle."
GRAHAM WARWICK / WASHINGTON DC
8-14 OCTOBER 2002 FLIGHT INTERNATIONAL
"Engine development"
"The success of our architecture depends on the success of NASA's engine development programme," says Young. The space agency is funding work on four main engine candidates, two hydrogen fuelled and two kerosene-fuelled. Pratt &
Whitney and Aerojet are developing the Cobra, a 600,0001b-thrust (2,670kN) hydrogen-fuelled, staged-combustion, first and second-stage engine, while Boeing's Rocketdyne division is working on the 650,0001b thrust-class RS-83. Rocketdyne
is also pursuing the RS-84, a kerosene fuelled, staged-combustion, first-stage engine generating 1,100,0001b thrust,
while TRW is developing the 1,000,0001b thrust-class TR107. The plan is to test two prototype engines at a cost of $1.3 billion. "NASA will go for prototype engines that bracket the requirements of the three contractors," says Ford. He suggests the emphasis has shifted towards the kerosene-fuelled engines. "NASA wants to address kerosene first to reduce risk," he says. The USA has little experience with kerosene-burning rocket motors, having focused for decades
on cryogenic engines."
http://www.flightglobal.com/pdfarchive/ ... 02996.html

So for one heavy thrust engine it might only be $650 million. And considering that much development already went into the RS-84, it would probably be even less than this.

Bob Clark
 
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exoscientist

Guest
Ed, perhaps you can answer a question for me. Suppose we had a variant of the Jupiter Direct SD-HLV: this version will be used only to launch propellant to orbital depots. It does not have a payload canister or an interstage atop the external tank. It is just SSME's or RS-68's below the ET and the two SRB's. Since the shuttle orbiter plus payload could total 100+ Mt that reached orbit, it would seem to me this could loft 100 Mt of propellant, i.e., LH2/LOX, carried within the ET.
IF we could get a reason for this large amount of propellant to be launched on a regular basis then we would save on the launch costs on this system by having a high launch rate. This would then also pay for the launches of the usual form of the Jupiter Direct launcher to launch cargo, satellites, spacecraft, to orbit.
This is what I'm thinking. I did some back of the envelope calculations. The Dragon capsule only weighs 3,100 kg empty. Call it 4,000 kg with 1 to 4 passengers/crew. Now take a look at the delta-V requirements for a lunar landing mission that departs from LEO and returns to LEO:

Delta-v budget.
2.3 Earth–Moon space
http://en.wikipedia.org/wiki/Delta-v_bu ... Moon_space

If you cut the return trip delta-V by using aerobraking to re-enter Earth orbit, then the total round trip delta-V will be somewhat less than the delta-V needed for a launch from Earth to LEO. Now hydrogen fueled stages have gotten better than .90 propellant mass fractions for a while now. Say the propellant to inert mass ratio is 10 to 1. If we have a hydrogen-fueled rocket stage with 50,000 kg propellant mass, inert mass of 5,000 kg and Dragon capsule as payload with crew and cargo of 4,000 kg, and use a vacuum Isp of 465 s, then we get a delta-V of 465*9.8ln(1 + 50,000/(5,000 + 4,000)) = 8,568 m/s, about what's needed for the round-trip from LEO to the Moon's surface and back to LEO, with aerobraking.
About the cost of launching the propellant, estimates of the yearly costs of the STS system are around $3 billion annually, regardless of the number of launches, which has made shuttle launches quite expensive with the low number of launches that now obtain. So if we get 15 launches per year this is $200 million per launch (actually it likely will be even lower if you don't have the shuttle orbiter in the mix.) And at 100,000 kg of propellant as payload this would be $2,000 per kg.
For the 50,000 kg of the propellant needed for each Moon mission, only actually $100 million would cover the cost of the propellant needed. The Dragon capsule would be reusable and LH2/LOX engines such as the RL-10 were shown to be reusable multiple times on the DC-X test vehicle. So the cost of each flight only from LEO to the Moon would likely be less than the $100 million cost of getting the propellant into orbit.
Then this is a total cost that would well be affordable to most nations to make manned missions to the Moon.

Bob Clark
 
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