Agreed, re: the first stage. The biggest bang for the buck would have been to change the second stage, rather then mussing with the S-IB stage. Shifting to a LOX/RP second stage powered by three Atlas sustainer engines, designed for Atlas sustainer type propellant mass fractions, could have reduced launch costs with minor loss of performance to LEO. A benefit would have been that the entire machine would have used, essentially, the same engines as NASA's Atlas-Centaur.But, as we've discussed, the first stage would had to have been substantially redesigned and re-engined to achieve Falcon 9 like recovery - and recovery would have bit hard into performance. - Ed Kyle
A 9th center engine would be necessary for any future potential attempt at propulsive landing), so just add it in right there),
So yes, a 1st stage with significantly better characteristics, for sure. Otherwise, just going to a mono core and changing nothing makes little sense
Quote from: Lobo on 03/09/2018 05:08 pmA 9th center engine would be necessary for any future potential attempt at propulsive landing), so just add it in right there),I'm thinking that the 9th engine would have to have been of less thrust than the H-1, unless the H-1 could have been made to throttle down to about 100,000 lb. The reason is that I suspect the hover-slam landing would have been beyond the capability of circa 1970 avionics. On the other hand, without hoverslam, the whole landing is more propellant intensive, so maybe that's just another reason the concept would not have been practical.
Quote from: Lobo on 03/09/2018 05:08 pmSo yes, a 1st stage with significantly better characteristics, for sure. Otherwise, just going to a mono core and changing nothing makes little senseCome to think of it, even though the stage's relatively high structural fraction didn't hurt its performance as a launch vehicle very much, it would have made rocket-powered boost-back and recovery very propellant thirsty. So I suppose a first stage redesigned for much less weight was essential. That then means, though, that the H-1 must throttle even deeper. If you got the stage's structural fraction down to an F9-like 4%, that would be about 40,000 pounds' dry weight, giving the H-1 a very long way to throttle.Another issue occurs to me with the whole concept, namely cost savings. I don't know the ratio of the costs of F9's two stages, but for a first guess I'd start with the ratio of Merlins on each, making the second stage about 1/9 the cost of the first. I'm sure it's actually more expensive than that (for one thing, it's got an Mvac, with a long nozzle that's produced in relatively low numbers). But still, it seems likely that the first stage represents most of the cost of the rocket.With the classic Saturn IB, however, studies of various improved versions show that the two stages are similar in cost. The precise numbers depend on production rates for both the Saturns IB and V, since making S-IVB-500s for the Saturn V helps keep unit costs down for the S-IVB-200s flown on the IB. But, the point is, you go to all this trouble and accept some loss of payload to recover only about half the value of the launch vehicle. Probably not viable.
Bloke,Interesting. Seems like it'd make more sense to land engine down though, even if not actually using the J2 to land, having some landing engines on the MTS would help keep a lower center of gravity than with the engine up top.
The S-1C would makes sense to water land nose first, if they could keep the engines up out of the water, and to keep them from impacting first which could risk impact damage to the engines, as well as the salt water exposure.
Blowing the O2 dome would be the attempt to create a piston deceleration on impact and ballast the heavy engines to keep them up out of the water, as I understand.
Although I then wonder how much of the stage itself would be adequately reusable given the top dome needs replaced and the tanks are full of water. And there could be damage of the barrel walls or tanks from that impact.
The jettisonable 4-engine ring concept of the S-1D seems more simple, and the engines are the most expensive bit of the whole stage. They would hit the water, and be exposed to it, but with simple gas generator type engines, I don't know that that would be too hard to clean the up after. Especially if there was a vessel downrange that could pluck them out of the water in short order, limited their salt water exposure to the bare minimum.
Plus then you'd have the Saturn VB 1.5 stage to orbit booster that could be used stand alone.
Some pretty interesting concepts if history had taken a different tact and the Saturn hardware had been stuck with and continued funding.
Ironically, once you calculate up all of the money it took to develop STS, including modifications to KSC which was already set up for Saturn, they probably could have kept Saturn V and/or some INT derivatives and just flown that. While adding in some limited reusability like the reusable S-1C, along with upgrades to the hardware and manufacturing process as they'd be building long term, rather than just in a limited batch or two...And saved money.
The whole 'landing' kit was supposed to mass around 6,500lbs with about half that for the 'legs' (deployable but of fixed length with no 'shock' absorption) and the 'crushable' landing segment. Which would pretty well 'balance' out the mass of the J2 and thrust structure or real close. Couple that with the reentry shield and the fact you're trying to keep it 'nose-forward' for most of the flight anyway landing on the 'nose' is actually easier than trying to transition the stage in flight. (Something SpaceX has learned quite well)
Note: The engines still get 'dunked' they just don't have to deal with a lot of 'impact force' initially. As it is the stage "gradually" (over about three minutes was the value IIRC) rolls over to an 'engine-down' and the LOX tank drains of water. The empty RP1 tank keeps the whole thing afloat.
Well they were 'assuming' the testing of the H1 would be applicable to the F1 but since the never actually tried it... As for being 'simpler' you're right but... Recovery of the 'ring' was problematical as Michel Van noted in the other thread:https://forum.nasaspaceflight.com/index.php?topic=45341.msg1810605#msg1810605
And you'd need some sort of 'impact attenuation' system which you don't have with the ring but do with the stage itself. Plus the engines ARE going to is the water bell first which is a problem and maybe a big one without MORE weight in a set of radar activated retro-rockets, (which was initially suggested for the Saturn-1 recovery for the same reason; it landed engines down) and all that implies.
QuotePlus then you'd have the Saturn VB 1.5 stage to orbit booster that could be used stand alone.Well yes, but for how much cargo on-orbit compared to the overall cost? Between 49Klbs and 63Klbs according to Boeing as long as they used stretched tanks. Otherwise it couldn't make orbit with any payload. Simpler but is it cost effective?*snip*Well, actually the answer at the time wasn't as clear. The STS was supposed to be fully reusable from the start and that it wasn't nor was it as 'cost-effective' as promised wasn't an 'engineering' issue but a political and policy issue.
The propulsion system selected as the prime candidate was the Rocketdyne LR-105 rocket engine with twelve Rocketdyne LR-101 Atlas vernier engines, due to the mass growth potential it provided. As far as the basic thrust engine and vehicle performance are concerned, the LR-91 (Titan 2nd stage), and LR-105, Atlas sustainer, are essentially the same. The LR-81 Agena engines were ruled out because of the highly toxic IRFNA oxidizer. The LR-99 was specified for study because it is man rated, throttleable, and available. Due to the limited throttling capability of the basic LR-99 and LR-105 (132.6 kN and 204.6 kN, respectively, compared to 71.2 kN required) other cruise modes needed consideration. Excess thrust during cruise is detrimental in two respects. Excess propellant is burned and the speed brake requirements add considerable vehicle mass. A derated LR-105 was included with a throttling capability down to 57.8 kN.Several cruise engine options were included. To allow flexibility between rocket cruise time and boosted Mach number, a common propellant is required. LR-101's are compatible with the LR-105 and, in fact, can utilize the LR-105 propellant pumping system. This gives a high degree of flexibility insofar as number of chambers used in combination with the potential throttling capability of the LR-105 pumps. The LR-101 thrust can vary from 2. 74 to 6. 23 kN per chamber.
Sooo, since the H-1s were not able to throttle, I was wondering about the following alternative: replace the four outer H-1s with four "mix" of derated LR-105 + verniers, for a smooth landing. Later the four H-1 in the center could be replaced by a F-1 but that's another story. Thoughts ?
I've cleaned up the Bellcomm document "Propulsive landing of ballistic vehicles" dated June 30, 1971
http://americanrocketnews.com/19790072472_1979072472_Propulsive_Landing_of_Ballistic_Vehicles.pdfAnd attached.
...The reason is that I suspect the hover-slam landing would have been beyond the capability of circa 1970 avionics. ...
Even 1950's electronics would have been able to handle it.
Quote from: libra on 06/10/2021 11:00 amhttp://americanrocketnews.com/19790072472_1979072472_Propulsive_Landing_of_Ballistic_Vehicles.pdfAnd attached. Thanks! And bravo for attaching the paper, since we've all learned the hard way that links don't always last.Next question: how did you find that link? I've gone to americanrocketnews.com but see no evidence of documents available on the site. Are there more nuggets to be downloaded from that site?