Why not just bring back the F-1 and J-2?

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vulture3

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The Titan was visually impressive but probably the most complicated, expensive, and unreliable of all the major ELVs. The last time solid fuel was economical in transportation was when ships and trains stopped burning coal. I believe the fundamental error with Shuttle was the attempt to predict operational costs without actual hands-on experience with reusable prototypes. I don't think anyone understood the total cost, particularly the tedious transportation and stacking and complete disassembly for reuse, the overhead for the facilities needed, the fact that all operations are hazardous, etc.
 
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oldAtlas_Eguy

Guest
NASA should have known better because AF Titan IV launch costs were astronomical and it wasn’t because of the UDMH and Nitrous Tetroxide core booster either. A Titan IV normal pad time was 6 months to 1 year. These very long pad times to get the solids erected and attached then once it was all built up finally add the payload on top resulted in a large amount of costs. NASA thought they could do it cheaper. HA!
 
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edkyle99

Guest
dwight_looi":2op1ulzg said:
In retrospect, it appears that the last 40 years has seen NASA, the USAF and private US companies subcontracting for them running around in circles trying to find more expensive and unsafe ways of accessing space. The Shuttle is the most expensive method of getting a pound of anything into orbit and it has the worst safety record of any man rated launch vehicle. The Atlas V and Delta IV offered nothing that a good old Saturn IB didn't. The ARES I again was another 15~20 tonner with nothing new to offer other than stability problems and resonance issues with the extended length SRB. Even the latest kid on the block -- the Falcon 9 -- is basically rehashing the medium lift option albeit using nine very small engines.
So many things wrong here. Where to begin?

STS and R-7/Soyuz have equivalent safety records. Soyuz would have had a worse crew-loss record were it not for one successful abort system performance.

Saturn IB was a two-stage human LEO launcher. It could not go to GTO or GEO or deep space or sun synchronous orbit, all of which the EELVs can do. Both EELVs can outperform Saturn IB to LEO from the Cape - and both can do it for less money (inflation adjusted).

Ares I did not have "stability problems". Thrust oscillation was solved. The problem with Ares I was cost, or insufficient funding depending on one's perspective.

Falcon 9 is an evolutionary advance in hydrocarbon launch technology, leveraging engine performance and materials technology, etc.. No other two-stage kerosene rocket can go to GTO.

It seems that NASA and the US government is more interested in preserving aerospace jobs in existing and unnecessary programs than finding the best and most efficient means of flying orbital missions. I mean, let's turn the clock back to 1970. Let's say the EELVs didn't exist, the Shuttle didn't exist, SpaceX didn't exist and ARES wasn't even thought of...

Can you imagine what we could have done and how many more missions we could have flown if we simply rolled out a simple 2-stage vehicle using a single F-1 engine on the first stage and a single J-2 engine on the upper stage? That's a 20-ton to LEO Proton class vehicle. It'll have all manrated engines with extensive flight histories. We could have launched the Hubble with it. We could have launched Chandra with it. We certainly could have pieced together the ISS with it. There wouldn't have been a need for the Delta IV, the Atlas V or the Falcon 9 for the most parts. There wouldn't have been a need for the RS-68 engine or the license built Russian RD-180. Most of the shuttle payloads would have flown on a 20 ton vehicle. If you really want, you can probably strap three side by side for 50~60 tons of LEO capacity.
This is still a rocket that, like Saturn IB, cannot do beyond LEO missions. It would need a third stage (probably a Centaur).

F-1 was considered for Shuttle, and bypassed in favor of SRB. The reason was that SRB cost less.

Your design, BTW, produces fairly high g-forces at the end of the first stage burn. A more realistic design would probably have a heavier upper stage, perhaps 100 tonnes.

The biggest problem with this idea is that it would have required the Air Force and NASA to share a human-rated rocket, which Shuttle showed does not work.

- Ed Kyle
 
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oldAtlas_Eguy

Guest
edkyle99":tefyvldb said:
F-1 was considered for Shuttle, and bypassed in favor of SRB. The reason was that SRB cost less.
- Ed Kyle
Actually the Boeing Flyback F-1 was passed up not because of its operational costs but its development costs. It would have been immensely cheaper to operate than the SRB but its development costs would not fit in the development cost budget. NASA originally wanted Shuttle to be a fully reusable 2 stage system. As the development costs estimates came in and Congress kept reining in what NASA could spend yearly for its development NASA had to back off from its goals to the eventual halfway compromise vehicle the Shuttle is today, resulting in much higher operational costs.

On another point a F-1 1st-stage vehicle to be price competitive to existing LV’s it would have to be reusable at least 3 times or more to spread out the cost of the engine over multiple flights. Even though the F-1 wasn’t designed as a reusable engine it demonstrated a capability to be reused as much as 5 times.
 
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neutrino78x

Guest
Falcon 9 Heavy will be quite capable of launching Orion plus its service module (total 21,250 kg) into low earth orbit.

Beyond that, you either give it another rocket in space, or you use a more powerful rocket on Earth.

--Brian
 
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stevekk

Guest
neutrino78x":27jv1o4v said:
Falcon 9 Heavy will be quite capable of launching Orion plus its service module (total 21,250 kg) into low earth orbit.

Beyond that, you either give it another rocket in space, or you use a more powerful rocket on Earth.

--Brian
What is the most optimistic date for the first F9H launch, 2015 ??

If the constellation program was able to meet any of its original goals / dates, it would have launched the Orion capsule into LEO using the Ares 1 before then. The Ares V was necessary to go to the moon, but can the F9H get that job done ??
 
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neutrino78x

Guest
stevekk":2vgehei9 said:
If the constellation program was able to meet any of its original goals / dates, it would have launched the Orion capsule into LEO using the Ares 1 before then. The Ares V was necessary to go to the moon, but can the F9H get that job done ??
Falcon 9 heavy will be able to put Orion + service module into LEO. Presumably, you would use the Service Module on the Orion to send it to the Moon from there. The concept of the orbital fuel depot plays a role here, if you can refuel the service module to go to Mars. I would think you would lift up another, more powerful service module for that, though.

A Falcon 9, non-heavy, can lift the Orion itself to LEO, if you lift the service module with another rocket.

--Brian
 
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edkyle99

Guest
oldAtlas_Eguy":p6v00zk1 said:
edkyle99":p6v00zk1 said:
F-1 was considered for Shuttle, and bypassed in favor of SRB. The reason was that SRB cost less.
- Ed Kyle
Actually the Boeing Flyback F-1 was passed up not because of its operational costs but its development costs. It would have been immensely cheaper to operate than the SRB but its development costs would not fit in the development cost budget. NASA originally wanted Shuttle to be a fully reusable 2 stage system. As the development costs estimates came in and Congress kept reining in what NASA could spend yearly for its development NASA had to back off from its goals to the eventual halfway compromise vehicle the Shuttle is today, resulting in much higher operational costs.
Operational costs. Development costs. It's all money. The most important comparison is total program costs. Since Shuttle will only end up flying 134-135 times, I suspect that the development cost difference would have trumped the operating cost difference.

But let's turn this upside down. In 1972 the U.S. chose to go with Shuttle as, essentially, a partial replacement for NASA's Saturn/Apollo. We've contemplated one alternative - building a new F-1/J-2 rocket instead. A second option would have been to keep Saturn IB and Apollo rather than Shuttle. Here's a third option: Rather than develop an all new rocket, or use an existing costly rocket, why not ask NASA to use the existing Titan IIIC system? This would have caused the same NASA/USAF sharing issues, but if money-saving was the goal, why not use a proven rocket that already existed?

- Ed Kyle
 
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oldAtlas_Eguy

Guest
edkyle99":bksso2v6 said:
oldAtlas_Eguy":bksso2v6 said:
edkyle99":bksso2v6 said:
F-1 was considered for Shuttle, and bypassed in favor of SRB. The reason was that SRB cost less.
- Ed Kyle
Actually the Boeing Flyback F-1 was passed up not because of its operational costs but its development costs. It would have been immensely cheaper to operate than the SRB but its development costs would not fit in the development cost budget. NASA originally wanted Shuttle to be a fully reusable 2 stage system. As the development costs estimates came in and Congress kept reining in what NASA could spend yearly for its development NASA had to back off from its goals to the eventual halfway compromise vehicle the Shuttle is today, resulting in much higher operational costs.
Operational costs. Development costs. It's all money. The most important comparison is total program costs. Since Shuttle will only end up flying 134-135 times, I suspect that the development cost difference would have trumped the operating cost difference.

But let's turn this upside down. In 1972 the U.S. chose to go with Shuttle as, essentially, a partial replacement for NASA's Saturn/Apollo. We've contemplated one alternative - building a new F-1/J-2 rocket instead. A second option would have been to keep Saturn IB and Apollo rather than Shuttle. Here's a third option: Rather than develop an all new rocket, or use an existing costly rocket, why not ask NASA to use the existing Titan IIIC system? This would have caused the same NASA/USAF sharing issues, but if money-saving was the goal, why not use a proven rocket that already existed?

- Ed Kyle
Ever hear of the Atlas The pratically same vehicle except for engine upgrades and tank streches flew from 1957 through the mid 1990's
 
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oldAtlas_Eguy

Guest
CAllenDoudna":f65f8cws said:
Actually, there was a proposal for an F-1 flyback that would have continued existing technology:

http://www.nasm.si.edu/collections/artifact.cfm?id=A19740733000

But it was scrapped because projected cost of the modifications was slightly beyond budget goals so it was scrapped and we went with the Shuttle because lower cost estimates were submitted for it :lol:
Yes the estimated budget overrun was $300 mil per year for 5 years. A total of $1.5 bil which would have saved over 100 missions a total $10 bil.
 
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dwight_looi

Guest
edkyle99":1yu8xjhc said:
STS and R-7/Soyuz have equivalent safety records. Soyuz would have had a worse crew-loss record were it not for one successful abort system performance.
The Soyuz had no in flight failures since 1975. It had no loss of life since 1971. The Soyuz over 1000 flights to the STS's 100+ flights. The Soyuz has much lower pad times and schedule reliability.

edkyle99":1yu8xjhc said:
Saturn IB was a two-stage human LEO launcher. It could not go to GTO or GEO or deep space or sun synchronous orbit, all of which the EELVs can do. Both EELVs can outperform Saturn IB to LEO from the Cape - and both can do it for less money (inflation adjusted).
The Saturn 1B was not a particularly simple or elegant design. It was eight bundled up H-1 boosters as its first stage. But to say that the EELVs outperform it is a misnomer. The EELVs cannot lift 20 tons without strap-ons, the 1B could. And to say that it cannot go to GTO or GEO is also misleading. You can always go to GTO or GEO if you reduce the payload; EELV payloads halves if you want to go to GTO. The similar limitations, perhaps worse applies to the 1B, but if you reduce the payload enough, you can always put a vehicle that goes to LEO on a GTO trajectory. The Saturn 1B has never gone to GTO, not because it can't but because it was flown specifically with the Apollo capsules on top and those mission have no business in GTO or GEO. The 1B was flown exactly 9 times that way and it was never flown any other way.


edkyle99":1yu8xjhc said:
Your design, BTW, produces fairly high g-forces at the end of the first stage burn. A more realistic design would probably have a heavier upper stage, perhaps 100 tonnes.

The biggest problem with this idea is that it would have required the Air Force and NASA to share a human-rated rocket, which Shuttle showed does not work.
The J-1 makes about 104 tons of thrust. Let's assume a 20 ton capsule and a 90% propellant fraction on that 80 ton upper stage. Let's also assume that the engine cannot be throttled and burns at 100% thrust all the way. This translates into a 1.04G initial acceleration and 3.7G at burn out. That's not particularly high, and definitely lower than the 6.5Gs or so experienced by the Mercury-Atlas astronauts put up with.
 
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edkyle99

Guest
dwight_looi":305j5izr said:
edkyle99":305j5izr said:
STS and R-7/Soyuz have equivalent safety records. Soyuz would have had a worse crew-loss record were it not for one successful abort system performance.
The Soyuz had no in flight failures since 1975. It had no loss of life since 1971. The Soyuz over 1000 flights to the STS's 100+ flights. The Soyuz has much lower pad times and schedule reliability.
R-7 has flown 1,761 times, but it failed on about 5% of those launches. The manned Soyuz launcher has posted a better record, but it still has a about a 3% failure rate - statistically equivalent to STS. Although "no one has died" on Soyuz since 1971, it has only flown about 96 times since then - and it has only flown about 59 times since the Soyuz T-10-1 pad abort in 1983. Soyuz is as reliable as it gets, but it is no sure thing. Every cosmonaut/astronaut is, I'm sure, keenly aware of the very real risks.
edkyle99":305j5izr said:
Saturn IB was a two-stage human LEO launcher. It could not go to GTO or GEO or deep space or sun synchronous orbit, all of which the EELVs can do. Both EELVs can outperform Saturn IB to LEO from the Cape - and both can do it for less money (inflation adjusted).
... Saturn 1B ... to say that it cannot go to GTO or GEO is also misleading. You can always go to GTO or GEO if you reduce the payload;
Saturn IB's S-IVB stage was far to big to act as an effective GTO stage. With a nearly 14 tonne dry mass, S-IVB would only be able to boost perhaps 1 tonne to GTO - less than much cheaper alternatives of the day. The only way to make Saturn IB into a GTO launcher would have been to add a third stage. A proposal was made to develop a Saturn IB/Centaur, which would have been able to lift about 6 tonnes to GTO, but the plan was canceled due to funding cutbacks in 1965.
edkyle99":305j5izr said:
Your design, BTW, produces fairly high g-forces at the end of the first stage burn. A more realistic design would probably have a heavier upper stage, perhaps 100 tonnes.
The J-1 makes about 104 tons of thrust. Let's assume a 20 ton capsule and a 90% propellant fraction on that 80 ton upper stage. Let's also assume that the engine cannot be throttled and burns at 100% thrust all the way. This translates into a 1.04G initial acceleration and 3.7G at burn out. That's not particularly high, and definitely lower than the 6.5Gs or so experienced by the Mercury-Atlas astronauts put up with.[/quote]

The G-force problem is at the end of the first stage burn, when F-1 would be making nearly 793 tonnes of vacuum thrust, resulting in G-forces in excess of 5.5 G. A slightly heavier upper stage (balanced by a slightly lighter first stage) would ease this problem, and would create a more efficient vehicle for LEO work.

Now, if you had F-1 and J-2 engines that could *throttle down*, it would be possible to create an effective two-stage GTO rocket. A 500+ tonne (gross) first stage topped by a 60 tonne (gross) second stage would be able to lift about 5.5 tonnes to GTO. This would have cut a couple of tonnes off LEO performance, but most payloads don't go to LEO. Later improvements would have allowed improved performance, like today's Proton.

- Ed Kyle
 
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edkyle99

Guest
oldAtlas_Eguy":3rggynfy said:
edkyle99":3rggynfy said:
Here's a third option: Rather than develop an all new rocket, or use an existing costly rocket, why not ask NASA to use the existing Titan IIIC system? This would have caused the same NASA/USAF sharing issues, but if money-saving was the goal, why not use a proven rocket that already existed?

- Ed Kyle
Ever hear of the Atlas The pratically same vehicle except for engine upgrades and tank streches flew from 1957 through the mid 1990's
Atlas Centaur of the early 1970s (e.g. SLV3D Centuar D1A) was a much less capable launch vehicle than the Titan IIIC then available. That Atlas Centaur could lift 5.1 tonnes to LEO or 1.8 tonnes to GTO. Titan IIIC could boost 11.5 tonnes to LEO or 1.4 tonnes directly to GEO. With a Centaur stage, a Titan IIIE (which was developed for NASA) could lift 17 tonnes to LEO or 6.8 tonnes to GTO.

- Ed Kyle
 
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dwight_looi

Guest
edkyle99":3vc8wfi2 said:
Saturn IB's S-IVB stage was far to big to act as an effective GTO stage. With a nearly 14 tonne dry mass, S-IVB would only be able to boost perhaps 1 tonne to GTO - less than much cheaper alternatives of the day. The only way to make Saturn IB into a GTO launcher would have been to add a third stage. A proposal was made to develop a Saturn IB/Centaur, which would have been able to lift about 6 tonnes to GTO, but the plan was canceled due to funding cutbacks in 1965.
Given the high upper stage dry mass to payload ratio, we can expect rather poor GTO performance since reducing the payload mass will have a smaller effect on the vehicle's total mass. However, I don't think it's as bad as you claim. Here's why...

Regardless of how big or how small the upper stage is, what matters is burn out velocity. The higher the velocity, the higher of an orbit you end up with. The gain in velocity is directly proportional to the reduction in total vehicular mass at burn out. Given that we know that LEO payload and burn out mass, we can estimate the GTO fraction.

The Falcon 9 carries 10.4 tons to LEO and 4.5 tons to GTO. Let's assume that the 45 ton Falcon nine upper stage has a very good propellant fraction of 92%* compared to the 120 ton Saturn IB upper's rather marginal 88%. At upper stage burn out the Falcon 9 will have put a total mass of 3.6 + 10.4 = 14 tons into LEO. Reducing the payload to Falcon 9's GTO claim of 4.5 tons reduces that to 3.6 + 4.5 = 8.1 tons. That is to say that a 42% reduction in total burn out mass (14 tons -> 8.1 tons) is sufficient to raise velocity from LEO to GTO numbers.

A similar deduction can be done using the Delta IV upper stage as an example. The Delta IV Upper stage has a mass of 2.85 tons empty. In a Delta IV Medium this is saddled with a 8.6 tons payload and both reach LEO velocities. That is a total of 2.85 + 8.6 = 11.45 tons being put in LEO. The same vehicle also puts 3.9 tons to GTO. This means that total mass of 2.85 + 3.9 = 6.75 tons can reach GTO. Again, shows that a reduction in total mass of (11.45 - 6.75) / 11.45 = 41% is needed.

A Saturn 1B carries 20 tons to LEO and will have a 14 + 20 = 34 tons of mass at LEO burn out. A 42% reduction pegs the total mass at 19.7 tons. Given that 14 tons of that is the upper stage itself, this leaves a payload of 5.7 tons to GTO. This represents a 71.5% reduction in payload when you want to do a GTO mission as opposed to an LEO one. This is much worse than the Falcon 9's 56.7%. But it is not 1 ton and it does not amount to a 95% reduction as you have suggested.

* The 92% number is consistent with Isogrid Tanks being good for about 90%, whereas Balloon Tanks are about 94%, a Flight Pressurized Tank should be somewhere in between. 92% is hence a prudent guess.

edkyle99":3vc8wfi2 said:
The G-force problem is at the end of the first stage burn, when F-1 would be making nearly 793 tonnes of vacuum thrust, resulting in G-forces in excess of 5.5 G. A slightly heavier upper stage (balanced by a slightly lighter first stage) would ease this problem, and would create a more efficient vehicle for LEO work.

Now, if you had F-1 and J-2 engines that could *throttle down*, it would be possible to create an effective two-stage GTO rocket. A 500+ tonne (gross) first stage topped by a 60 tonne (gross) second stage would be able to lift about 5.5 tonnes to GTO. This would have cut a couple of tonnes off LEO performance, but most payloads don't go to LEO. Later improvements would have allowed improved performance, like today's Proton.

- Ed Kyle
Agreed*. But don't you think 5.5Gs is acceptable, and well within human tolerance, even if it is higher than the STS's rather gentle 4Gs? Or, if you are really hard up on the G loading, you can increase the 1st stage lift off mass to say 600 tons. This will earn only a very small increase in payload capacity, but it'll reduce the 1st stage burn out Gs by making the dry mass of the 1st stage heavier -- mass that you'll dump shortly after anyway.

* BTW, it is actually it is not "in excess of" 5.5Gs but around 5.29 Gs (not that it changes essence of the argument). Here's the simple math:-

500 tons with a 0.90 propellant fraction = 50 ton of 1st stage empty mass.
Total Vehicle Mass at burn out = 50 (1st stage) + 80 (2nd Stage) + 20 (Payload) = 150 tons
Acceleration with 793 tons of F-1 Vacuum thrust = 793 / 150 = 5.29 G
 
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edkyle99

Guest
dwight_looi":2k6ij8gl said:
A Saturn 1B carries 20 tons to LEO and will have a 14 + 20 = 34 tons of mass at LEO burn out. A 42% reduction pegs the total mass at 19.7 tons. Given that 14 tons of that is the upper stage itself, this leaves a payload of 5.7 tons to GTO. This represents a 71.5% reduction in payload when you want to do a GTO mission as opposed to an LEO one. This is much worse than the Falcon 9's 56.7%. But it is not 1 ton and it does not amount to a 95% reduction as you have suggested.
Saturn IB could only lift about 18 tonnes (metric tons) to LEO (which, I suppose is about 20 tons).

The following is based on NASA's official post-flight mission report.
The SA-208 first stage weighed 448.62 tonnes at liftoff and 45.29 tonnes at burnout. Its average ISP was 296 seconds.
The SA-208 second stage weighed 119.42 tonnes at liftoff and 15.42 tonnes at burnout. Its ISP was 421 seconds.
With a payload fairing weighing roughly the same as the Apollo 8 LES and S-IVB SLA (total 5.26 tonnes), the first stage would, for a 1.2 tonne payload, produce 3,528 m/s delta-v and the second stage would make 8,180 m/s for a total 11,708 m/s, enough for GTO.
The same rocket would only be able to boost a 5.7 ton (5.17 tonne) payload to 10,911 m/s delta-v, only good for a transfer orbit apogee of a bit more than 12,000 km, far short of the needed 35,500 km.
* BTW, it is actually it is not "in excess of" 5.5Gs but around 5.29 Gs (not that it changes essence of the argument). Here's the simple math:-


I figured a 0.92 first stage propellant mass fraction, which is a conservative guess for a kerosene first stage.

The EELVs max at 5Gs and throttle to that maintain that level.

- Ed Kyle
 
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dwight_looi

Guest
edkyle99":3gx4ceul said:
Saturn IB could only lift about 18 tonnes (metric tons) to LEO (which, I suppose is about 20 tons).

The following is based on NASA's official post-flight mission report.
The SA-208 first stage weighed 448.62 tonnes at liftoff and 45.29 tonnes at burnout. Its average ISP was 296 seconds.
The SA-208 second stage weighed 119.42 tonnes at liftoff and 15.42 tonnes at burnout. Its ISP was 421 seconds.
With a payload fairing weighing roughly the same as the Apollo 8 LES and S-IVB SLA (total 5.26 tonnes), the first stage would, for a 1.2 tonne payload, produce 3,528 m/s delta-v and the second stage would make 8,180 m/s for a total 11,708 m/s, enough for GTO.
The same rocket would only be able to boost a 5.7 ton (5.17 tonne) payload to 10,911 m/s delta-v, only good for a transfer orbit apogee of a bit more than 12,000 km, far short of the needed 35,500 km.
Ah... I see where our assumptions differ. I calculated the GTO payload capacity without any Payload Fairing weight instead of 5.28 tons worth of it. With around 500 tons of lift-off mass and a 120 ton upper stage, the presence of a 5 ton fairing is immaterial during the 1st stage ascent. The fairing is typically discarded almost immediately after stage separation and 2nd stage ignition anyway so it won't affect upper stage delta V much. In general, nobody carries the fairing all the way to GTO. The Falcon for instance dumps the fairing a couple of seconds into a 380 second 2nd stage burn.

The assumption here of course is that the rocket is configured for satellite delivery with a clam shell fairing. The Apollo capsule is a different case, but it also doesn't have any business in GTO.
 
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edkyle99

Guest
dwight_looi":g1uhm7gu said:
Can you imagine what we could have done and how many more missions we could have flown if we simply rolled out a simple 2-stage vehicle using a single F-1 engine on the first stage and a single J-2 engine on the upper stage? ]
I've been thinking about this a bit. I've always thought that NASA could have built a sound program on Saturn IB in lieu of Shuttle. Saturn IB, flown with a Centaur third stage, could have performed a variety of deep space missions. The rocket could have served as a basis for future improvement. Perhaps F-1 could have been employed as one of these improvements, but the marriage would have been difficult and costly.

Consider, for example, NASA's thinking when the final Saturn IB vehicles were being built. The Agency was planning to steadily shave dry mass from future vehicles. It was also planning, notably, to replace J-2 with the more-efficient J-2S engine. A two-stage rocket employing these changes would have been able to lift 20,396 kg to a 186 km x 28.5 deg LEO. Adding a Centaur third stage, powered by a pair of RL10A-3-3 engines, would have resulted in a rocket able to lift 25,000 kg to LEO, 10,400 kg to GTO or 5,500 kg to GEO (if a 3-burn Centaur could be used). [see note 1]

This capability would not have come cheaply, however. Each launch would, for example, have cost more than today's Atlas V-551. The cost of a crewed flight atop the rocket, using an Apollo spacecraft, would have rivaled the cost of a Shuttle mission. On the other hand, NASA would not have had to pay the cost of developing Shuttle.

Saturn IB would have supported the Rocketdyne manufacturing base for Thor/Delta and Atlas engines, as well as for Pratt's RL10 and General Dynamic's Centaur. J-2S, however, would have been Saturn-specific.

- Ed Kyle



Note 1: Saturn IB/Centaur "Advanced" Details

Stage 1 + Interstage: Liftoff weight 440.142 tonnes. Burnout weight 44.7 tonnes. Liftoff Thrust 743.9 tonnes. Average ISP 285 seconds.

Stage 2 + IU: Liftoff weight 115.8 tonnes. Burnout weight 12.31 tonnes. Nominal Thrust 109.3 tonnes. Nominal ISP 433.5 seconds.

Stage 3: Liftoff weight 16.648 tonnes. Burnout weight 2.877 tonnes. Thrust 13.608 tonnes. ISP 444 seconds.

Payload Fairing mass ~5 tonnes.
 
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dwight_looi

Guest
I wasn't suggesting that the Saturn IB be used as the universal space launch vehicle. The IB has a relatively mass inefficient and complex eight-engine, eight tank 1st stage bundle.

I was thinking that a vehicle of the Saturn IB's lift capacity (perhaps slightly more) can be had with a 500 ton, F-1 powered, 1st stage and an 80 ton J-2 powered upper stage. 2-engines, 2-stages -- an orbital launch vehicle doesn't get simpler than that. I'll also use a 5m ~ 5.5m diameter so you have a straight, uni-diameter vehicle. We could have done it without developing any new engines.

If you are willing to spring for improved engines. It can be a 25 ton to LEO vehicle with a 600 ton F-1A 1st stage and a 100 ton J-2S upper. For earth escape trajectories, a third engine can be added as part of the payload. This will be compliant with the standard payload adapter, be enclosed in the standard fairing and be separated in the same manner as a satellite. The vehicle will reach escape velocity as a unit. A single RL10-A-2 will work or perhaps even an AJ10-118 if you want a storable propellant solution which can later be used for an extraterrestrial orbit insertion burn. The tank can be as big as 18 tons with a 2 ton instruments section if you want. Thrust-to-weight ratios won't manner much since you are already in orbit and can afford a long burn at a thrust level that is substantially less than the vehicle's mass. It can even be a 5 ton Star-63 (PAM-D) solid motor if it has to be umbilical free from integration to launch.

Whatever the cost of integrating the existing F-1 to a new 5m 1st stage tank, it'll be substantially less than the development cost of the STS -- new SRBs, new LH2/LOX tank, new SSME engines, new orbiter. With proper cost management, the F-1 1st stage will probably cost no more than the RD-180 1st stage in the Atlas V. Afterall, while the F-1 has some hard to build features like a brazed tube bell, the RD-180 has the much more complex turbopump assembly of a staged-combustion engine, two chambers and two bells. Roughly speaking, the 500 ton 1st stage will be roughly the size of the Delta IV 1st stage (5m vs 3.8m of the Atlas V; 500 tons vs 335 tons).
 
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edkyle99

Guest
dwight_looi":1iktzw4c said:
I wasn't suggesting that the Saturn IB be used as the universal space launch vehicle. The IB has a relatively mass inefficient and complex eight-engine, eight tank 1st stage bundle.

I was thinking that a vehicle of the Saturn IB's lift capacity (perhaps slightly more) can be had with a 500 ton, F-1 powered, 1st stage and an 80 ton J-2 powered upper stage. 2-engines, 2-stages -- an orbital launch vehicle doesn't get simpler than that. I'll also use a 5m ~ 5.5m diameter so you have a straight, uni-diameter vehicle. We could have done it without developing any new engines.

If you are willing to spring for improved engines. It can be a 25 ton to LEO vehicle with a 600 ton F-1A 1st stage and a 100 ton J-2S upper. ...

Whatever the cost of integrating the existing F-1 to a new 5m 1st stage tank, it'll be substantially less than the development cost of the STS ...
Setting ourselves back to 1972 or so, the basic problem with F-1 was cost. Here was an engine developed for one rocket - a rocket that had been canceled. H-1, on the other hand, was a basic design also built in large numbers for other rockets. Eight H-1 engines cost less, substantially less I believe, than one F-1. While the S-IB stage was not perfectly optimum, it got the job done. Better yet, it was bought and paid for. Consider, for example, Russia's less-than-optimum R-7. Much better designs are possible, and have been for decades, but R-7 continues to fly - more than 1,760 flights worth.

- Ed kyle
 
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oldAtlas_Eguy

Guest
The H-1 is shown to only have flown on the Saturn 1B but other engines of this size flew on other rockets. The booster engines on the Atlas were 180,000lbf but were later upgraded. The H-1 were 205,000 lbf engines and are constructed using the more expensive shower head injector design. The Merlin for example is a 125,000lbf engine that is much cheaper. A Falcon X with its larger core 6m possibly 6.6 m diameter could use a Saturn IVB upperstage. This would increase the Falcon X’s LEO capability but could also increase its cost because of the additional fuel types handling during pad operations. But the use of a IVB which is already designed using J-2X which has most of its development already paid for could be a development cost savings for Space X.
 
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vulture4

Guest
If you look at one of those H-1 or F-1 engines and try to count all the welded and brazed seams and precisely fitted joints you will loose your mind. These engines are from an era before CNC. In fact they are from an era before computers. Relying on legacy designs in launch vehicles and spacecraft is often a serious error. It results in excessive manufacturing and servicing costs, the acceptance of hazards that remain very costly to mitigate (i.e. hypergol RCS) decades after alternatives could have been implemented, and the failure of spacecraft due to designs for components as simple as propellant valves with fundamental materials incompatibilities. As to the question of injector plate design, in this era of CNC it's hard for me to believe a multiple-pintle injector cannot be machined out of three or four blocks of metal with each component of the pintles integral to one of the injector plates; isn't the RL-10 injector plate made this way? Similarly, most modern engine designs use channel-wall cooling rather than nozzles fabricated from tubes seen in these legacy designs.
 
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