ESA ATL Plausibility Checks and Development

I looked elsewhere for info about Jupiter Injections. Other sources give a more consistent delta-v requirement of ~9,300 m/s - mainly by a simple Googling of Earth-Jupiter Delta-V.

For that, and having recalculated the maximum payload for the Argo-HU at 33,420Kg for a 225x225Km Orbit at 51.6 degrees. I get a payload of 36,230Kg for a 185x185Km orbit at 7 degrees - the one I find most likely for GTO and Trans-Anywhere-Injection.

That means for a TJI assuming I actually need 9,300 m/s of delta-v, 31,700Kg of the Parking Orbit Payload must be propellant. Leaving me with 4,530Kg for not only the payload, but the dry mass of the upper stage. The above using In(TM/DM)*9.8*448.

If the above is incorrect in any way, could someone alert me to how it's actually done, and the correct number.

In any case. Even if it is correct, that would suggest that the Shuttle-C planned for TTL would have the necessary payload budget to accomplish sending a Juno-sized (3.635Kg) probe straight to Jupiter without the need for Planetary Flyby Assist. Though it also suggests that a Venus-Venus-Earth-Jupiter Trajectory would still be required for the 5,900Kg - IOTL - Cassini/Huygens mission.

Clearly a lot of work to be done if I wanna get this right.
 
That means for a TJI assuming I actually need 9,300 m/s of delta-v, 31,700Kg of the Parking Orbit Payload must be propellant. Leaving me with 4,530Kg for not only the payload, but the dry mass of the upper stage. The above using In(TM/DM)*9.8*448.

If the above is incorrect in any way, could someone alert me to how it's actually done, and the correct number.
Math checks out with the 9.3 km/s number, though I can't vouch for if it's right or not. However, that's with Argo, right? So this means the Argo-HU would be able to launch a Galileo (2,350 kg) or Juno (3,635 kg) directly to Jupiter. However, you appear to have gotten a bit confused here, using the Argo numbers for the nearly 2.5x larger Shuttle-C:
In any case. Even if it is correct, that would suggest that the Shuttle-C planned for TTL would have the necessary payload budget to accomplish sending a Juno-sized (3.635Kg) probe straight to Jupiter without the need for Planetary Flyby Assist. Though it also suggests that a Venus-Venus-Earth-Jupiter Trajectory would still be required for the 5,900Kg - IOTL - Cassini/Huygens mission.
Shuttle-C has a LEo payload of about 77 tons, right? So that a maximum payload through 9.3 km/s injection of 77000/(exp(9300/(9.8*448)) tons, or 9,278 kg. Titan V, with a maximum payload of 44 tons is about 5300 kg. this is only about 400 kg short of the OTL Cassini/Huygens, if they want to save a bit and use it instead (assuming it's ready in time). So they can either save some cash and design around a 5400 kg direct probe, go bigger and use some transfers (even just a single Earth flyby like OTL Juno should give the necesary boost, I'd think), or go really big and use Shuttle-C with direct transfer.
 
Math checks out with the 9.3 km/s number, though I can't vouch for if it's right or not. However, that's with Argo, right? So this means the Argo-HU would be able to launch a Galileo (2,350 kg) or Juno (3,635 kg) directly to Jupiter.

Alright, but you are aware that the Mass I gave was the total Dry Mass - Payload and Propellant Tankage? I would think 9-11% Dry Mass for the TJI stage is about reasonable. And that leaves me maybe 1,400Kg for the payload.


Shuttle-C has a LEo payload of about 77 tons, right? So that a maximum payload through 9.3 km/s injection of 77000/(exp(9300/(9.8*448)) tons, or 9,278 kg.

And the "exp" in that calculation is what, exactly? So I know what to use on my calculator for future use.


Titan V, with a maximum payload of 44 tons is about 5300 kg. this is only about 400 kg short of the OTL Cassini/Huygens, if they want to save a bit and use it instead (assuming it's ready in time). So they can either save some cash and design around a 5400 kg direct probe, go bigger and use some transfers (even just a single Earth flyby like OTL Juno should give the necesary boost, I'd think), or go really big and use Shuttle-C with direct transfer.

Shedding the necessary 400-600Kg may be doable with some good design. If not, using Earth and/or Mars for the extra delta-v could always be done. Failing those, there is always the 'Cop-Out' option of simply upping the mass and repeating the OTL travel path.
 
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And the "exp" in that calculation is what, exactly? So I know what to use on my calculator for future use.
exp(x) is often used for e^(x), the inverse operation of the natural log ln(x). If ln(X)=y, then exp(y) or e^(y)=x. If your calculator can't do that, the google web one can--that calculation I copied is the one I was using in the google calculator.
Shedding the necessary 400-600Kg may be doable with some good design. If not, using Earth and/or Mars for the extra delta-v could always be done. Failing those, there is always the 'Cop-Out' option of simply upping the mass and repeating the OTL travel path.
Well, it's less that they'd shed 400kg than that they'd just design a probe that's 400 kg lighter to start. Anyway, if you're repeating the OTL launch path, then it can also probably be larger than 5900kg even using Argo--after all, the stack IMLEO couldn't have exceeded the 21,000 throw of the Titan IVB launcher, so with a larger initial mass and the same boost, a proportionally larger probe could be sent.

Something I'm thinking about is the "Big Centaur" stage we were discussing previously for the Shuttle-C, and whether it'd offer anything as a replacement for Centaur on at least the larger Titan V. Seems like liftoff thrust should be okay thanks to the boost of the solids, the question is if at staging of the solids the core's burned enough for it to muddle along on the single SSME even with a larger upper stage and payload on the top, for which I really need the core's dry mass and initial fuel load.
 
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exp(x) is often used for e^(x), the inverse operation of the natural log ln(x). If ln(X)=y, then exp(y) or e^(y)=x. If your calculator can't do that, the google web one can--that calculation I copied is the one I was using in the google calculator.

Well it's not working in my own calculator - for some reason - but the Google one manages fine.


Well, it's less that they'd shed 400kg than that they'd just design a probe that's 400 kg lighter to start. Anyway, if you're repeating the OTL launch path, then it can also probably be larger than 5900kg even using Argo--after all, the stack IMLEO couldn't have exceeded the 21,000 throw of the Titan IVB launcher, so with a larger initial mass and the same boost, a proportionally larger probe could be sent.

I get 12,500Kg for Argo-HU when using a delta-v of 3,900 m/s. And 15,300Kg for the Titan V 12004 - the top payload version. I know I got a number of 3,500 m/s for an Earth-Venus Hohmann Transfer, but it's a standardised Upper Stage in both cases for the sending of probes to multiple targets on one common design. Should be well worth some small loss in maximum payload.

By the way E. I checked. The Titan IVB that was used for Cassini/Huygens did not send it all the way into IMLEO. The Centaur Upper Stage actually performed a short 2 minute burn to complete the Orbit prior to TVI.


Something I'm thinking about is the "Big Centaur" stage we were discussing previously for the Shuttle-C, and whether it'd offer anything as a replacement for Centaur on at least the larger Titan V. Seems like liftoff thrust should be okay thanks to the boost of the solids, the question is if at staging of the solids the core's burned enough for it to muddle along on the single SSME even with a larger upper stage and payload on the top, for which I really need the core's dry mass and initial fuel load.

Well the Core Stage for the Titan V carries a Propellant Mass of 200,000Kg with a Dry Mass of 25,000Kg. A Big Centaur I'd believe to be 42,000-43,000Kg so it should be possible to place it on top of the Titan V Core Stage. It appears to be ~0.8g following SRB separation, needing 60-75 seconds to return to >1g acceleration.
 
E. I ran the numbers concerning using a 'Big Centaur' on the Titan V LV - using the 4 7-seg SRB version. I get a number of between 53,800Kg and 60,000Kg for a 51.6 degree orbit at 225x225 Km. The variation being due to using either one or two 12,500Kgf RL-10 derived engines assuming a Vacuum Isp of either 462 or 470 seconds.

Both of which appear to allow for a payload of up to 19,280Kg and 21,900Kg for payloads fired to Mars and Venus at 3,900m/s.

So I suppose they could find a way of justifying the use of the Big Centaur on the Titan V, to which NASA may be willing to support - the resultant "cost-split" should make it an easier sell to Congress.

As for its diameter? I'm currently looking at 660cm - which would let it fit on both the Titan V and inside the Shuttle-C payload bay.
 
E. I ran the numbers concerning using a 'Big Centaur' on the Titan V LV - using the 4 7-seg SRB version. I get a number of between 53,800Kg and 60,000Kg for a 51.6 degree orbit at 225x225 Km. The variation being due to using either one or two 12,500Kgf RL-10 derived engines assuming a Vacuum Isp of either 462 or 470 seconds.
That's...very impressive, I'd say!
So I suppose they could find a way of justifying the use of the Big Centaur on the Titan V, to which NASA may be willing to support - the resultant "cost-split" should make it an easier sell to Congress.
Yeah, though thinking about it more I'm not sure how many uses the DoD will have for the kinds of payloads in that kind of range. A 60 ton thing in LEO is pretty big, and even if that's down to 20 tons in GEo that's pretty fricking huge, even double or tripple manifested. It may end up being used more by NASA probes (or, heck, small BEO manned missions!) than by anything the DoD itself is interested in.

As for its diameter? I'm currently looking at 660cm - which would let it fit on both the Titan V and inside the Shuttle-C payload bay.
Ah, 6.6m. Good pedigree for an EDS in my book. :)
 
Yeah, though thinking about it more I'm not sure how many uses the DoD will have for the kinds of payloads in that kind of range. A 60 ton thing in LEO is pretty big, and even if that's down to 20 tons in GEo that's pretty fricking huge, even double or tripple manifested. It may end up being used more by NASA probes (or, heck, small BEO manned missions!) than by anything the DoD itself is interested in.

It translates into about 23,500Kg to 26,000Kg for GTO, and unless you got a lot of payloads you want to send at once or wanna put a small station there, it's not gonna see too much use there. So large NASA probes may be the most likely use of such a system.

In any case, expect the DoD to use the smaller variants of the Titan V - the Base to no more than 2 7-seg SRB. With NASA tending towards the larger ones. Seems the most likely scenario.


Ah, 6.6m. Good pedigree for an EDS in my book. :)

Well I wonder why that is?:rolleyes:


In another related point. Having that 660cm 'Big Centaur' seems to add credibility to the use of a 570cm Core Stage for Titan V. Maybe it's just my mind playing games with me, but it seems to look that way. Nether the less, it's making Mr. 570cm look the more appealing candidate ATM.
 
So, the same rocket engine for a TJI requiring 6700 meters/sec delta-V from an Earth parking orbit could instead send a larger mass toward Venus; achieving a multiple set of passes by Earth and eventually on to Jupiter that way. The question was how much larger a mass? If injection to Venus requires only 3900 m/sec then I figure, about 1.87 times as much, using e of pi's implied ISP of 454 for the hydrogen engine.

The better figure of merit to use here is "C3," the orbit specific energy; roughly speaking, the square of the delta-V. Anyways, the benefits can be quite a bit larger than that, depending on the particular path chosen. For example, a flight to Saturn via Jupiter, in the fashion of Pioneer 11 or Voyager 1/2 will require a minimum C3 of 85 km^2/s^2 (that being the minimum C3 for getting to Jupiter directly), provided any very clever tricks like the "broken-plane maneuver" they were going to do during the 1986 Galileo opportunity don't exist/can't be taken advantage of. A flight to Saturn via Venus, Earth, and/or Jupiter can be as low as 11 km^2/s^2 (source: this is a pdf, by the way), using a September 1994 launch with a VEEGA trajectory. Similarly easy trajectories can be found for Jupiter, as can be seen in that document. Many of these do not require significant deep space maneuvers, which reduces the delta-V budget significantly. Course corrections will not generally take more than 1-200 m/s of delta-V, so the biggest drains are orbit insertion (which you can see from the source are generally around 700-1000 m/s), any deep space maneuvers (which range from a shade under 100 m/s in one of the Jupiter trajectories to just over 800 m/s in a different trajectory; Cassini OTL used a ~400 m/s DSM), and the necessary allowance for orbital tours (Galileo, in 1986, had a budget of ~200 m/s; Cassini had a budget of ~500 m/s).

So, what's the advantage of going from ~85 km^2/s^2 to ~15 km^2/s^2? Well, an Atlas V 551 using the 5-meter medium fairing and launching from Cape Canaveral into a trajectory with a perigee of 185 kilometers and a declination of 0 degrees* can put 1728 kg onto a trajectory with a hyperbolic C3 of 85 km^2/s^2, and 5204 kg onto a trajectory with a hyperbolic C3 of 15 km^2/s^2. So, at the price of roughly doubling your flight time to Jupiter (a direct trajectory would take about 3 years) you've tripled your payload, more or less. Seems fair to me, especially with how reliable probes are these days.

* And using the two-burn mode, GCS shutdown, and User's Guide calibration, for those of you who want to check my numbers in Schilling's.

Well, that sounds great, but of course out of that extra 87 (plus) percent has to come maneuvering propellant for tweaking the passages past Venus and Earth, not to mention any adjustments for slingshotting on past Jupiter. And assuming the planets are aligned correctly it takes many many years to get where it's going. And the planets are rarely aligned conveniently.

This is true and not so true. The advantage of the inner planet (Earth-Venus) encounters is that they actually are in launch position pretty frequently and have pretty frequent windows with Jupiter (about every 13 months for Earth). It's when you get Jupiter-transJupiter windows that everything starts going to hell in terms of launch frequency, because Jupiter and the planets beyond have decadal orbits. So if you stick to V(VV...)E(EE...)GA trajectories, you're fine, you'll have plenty of opportunities.

Of course, Titan V and Shuttle-C have the raw power to send extremely large payloads to Jupiter directly, so the question is more a trade of flight time versus launch cost than flight time versus payload. There...well, I'm a physicist, not an economist, so I can't tell you how likely they are to buy a really big expensive rocket to send a really big expensive probe to Jupiter or Saturn, versus using a smaller, cheaper rocket to send the same big expensive probe slowly versus using a smaller, cheaper rocket to send a smaller probe directly, etc. etc.

You could also use an SEP transfer stage to make things really ridiculous(ly big), but that's far more difficult to analyze than nice, simple ballistic trajectories, so...
 
Back to USAF LVs

Since from the 1990s onwards, V(V...)E(E...)(J)GA trajectories will be favoured on account of improved spacecraft reliability and superior payload capability, I think I'll stick to it and now turn my attention back to the USAF LVs.

OTL the Delta II was the one of the first LVs to be re-introduced to service following the Challenger Disaster. ITTL, it will be back a little sooner, since my planned STS failure occurs a little sooner than IOTL.

As for design, initially - at least at first - it will be largely like OTL Delta II. I decided on this since with the Titan V LV in development, and only so much funding available, combined with the need to get an LV in service quickly, it appears to me to be the only truly plausible scenario. Improvements can come later, but not right at the start.

As for what upgrades it can receive. The swapping out of the older Castor SRBs in favour of the GEM-40 SRBs is the obvious first step. Already being developed for the Titan V, seeing the Delta II adapted for use of them should only be as difficult as IOTL - because they did get GEM-40 SRBs IOTL.

The only other upgrades I can see it receiving is in the upper stages, where the old Delta-K 2nd stage is replaced by a LOX/Kerosene upper stage and/or a LOX/LH2 Centaur upper stage. All of which can close the payload gap between the Titan V 22000 and the best - payload-wise - Delta II.

The only other item I can see occurring here is the running-down of the remaining Titan IIIs and older Delta LVs to fill the gap between the STS disaster and the new LVs coming online plus resumption of STS flights - expect a lengthy wait though.
 
Since from the 1990s onwards, V(V...)E(E...)(J)GA trajectories will be favoured on account of improved spacecraft reliability and superior payload capability, I think I'll stick to it and now turn my attention back to the USAF LVs.
There are several advantages to direct injection. 1) the transit time is a lot shorter, you don't have to carom off the inner planets first, 2) you can use high energy propellants (LH2) for the whole thing.

OTL, the Venus/Earth flybys were makeshift because they COULDN"T use the LH2 Centaur.

I haven't done the math, but I wouldn't be at all surprised if you could get at least as big a payload with a Centaur on direct injection.
 
There are several advantages to direct injection. 1) the transit time is a lot shorter, you don't have to carom off the inner planets first, 2) you can use high energy propellants (LH2) for the whole thing.

An Earth/Jupiter Hohmann Transfer Orbit has an orbital period of 5.5 years IIRC, meaning ~2.75 years to get to Jupiter.

And you only really need LOX/LH2 for the really big delta-v burn. The one that gets you to your first target. After that, storable propellant can be used.


OTL, the Venus/Earth flybys were makeshift because they COULDN"T use the LH2 Centaur.

I haven't done the math, but I wouldn't be at all surprised if you could get at least as big a payload with a Centaur on direct injection.

Actually they could. But they wanted to send a very large payload to Jupiter/Saturn/Anywhere past those two. That meant Gravity Assists from the inner planets first.

In any case, what you're looking at is ~3,900 m/s vs. ~9,220 m/s. That is going to translate into a tremendous difference in spacecraft mass and capability.
 
An Earth/Jupiter Hohmann Transfer Orbit has an orbital period of 5.5 years IIRC, meaning ~2.75 years to get to Jupiter.
VEEGA was over 6 years, so that's a bit of a difference.
And you only really need LOX/LH2 for the really big delta-v burn. The one that gets you to your first target. After that, storable propellant can be used.
Of course. It's only available soon after launch, anyway, due to the LH2 boiling off. As you know.

In any case, what you're looking at is ~3,900 m/s vs. ~9,220 m/s. That is going to translate into a tremendous difference in spacecraft mass and capability.
Wow. I had no idea it was that much of a difference. Especially since you have to start going in the wrong direction (LOSING solar orbital velocity).

Ja, that's HUGE. Especially given the exponential nature of the rocket equation.
 
Of course. It's only available soon after launch, anyway, due to the LH2 boiling off. As you know.
Yeah, but with a VEGA or such, you're doing almost all the burn at the same time--right after launch into an injection trajectory. Most of the delta-v at flyback will come from the flyby itself, hence it's not really using a hydrolox stage vs not using one, it's using a hydrolox stage for a big burn (direct injection) or a smaller burn, then making up the difference with the extra gravity assist flybys. Adds flight time, but lets you get more ability from the same injection stage.
 
Delta 6/7000 Series LV

Boy did this thing go into the Cold Sleep! :eek:

And now it's time for this TL to thaw out and start breathing again.

In any case. It's time to work on the Delta 6/7000 series which should complete the US side of things insofar as the LVs are concerned. After that, a loose timeline can be sketched out for them.

As I stated earlier, with the Shuttle-C and Titan V LVs coming into existence ITTL, something is going to have to give, and that's here, where upgrade options for the Delta series will be rather limited. As such, expect it to track largely like OTL for the early years. That is, the Delta 6000 is built as an interim, while the 7000 series is developed. Now there does lie, one piece of help. OTL, the GEM-40 SRBs were initially developed for the Delta 7000, ITTL, they are being developed for the Titan V as well, that could be of help to the both of them, owing to a piece of hardware commonality.

As for the main engine on the 7000. I'd expect it to be the same RS-27A as OTL, NIHS(1) being a key factor. That and I find the performance figures for the RZ.4 and RS-27 are extremely similar, really not much of a gap.

If it's decided to try and uprate the performance of the Delta 7000, I see only two key scenarios:


1. Augmented SRBs. IOTL, the Delta 7000 could be fitted with larger GEM-46 SRBs for improved performance, but that takes their diameter to the largest they can be while still squeezing nine of them onto the core stage.

2. New upper stage. They actually tried this with the Delta 8000 series IOTL, it failed rather spectacularly - only the GEM-46 SRBs survived. In any case, a 305cm LOX/Kerosene upper stage or a 400cm LOX/LH2 upper stage is doable so long as they get it right. But with everything else happening, funding may not be available.


Definitely not gonna be easy to keep on the right side of plausible, but it can be done. Thoughts?


(1) - NIHS Not Invented Here Syndrome
 
Boy did this thing go into the Cold Sleep! :eek:

And now it's time for this TL to thaw out and start breathing again.

In any case. It's time to work on the Delta 6/7000 series which should complete the US side of things insofar as the LVs are concerned. After that, a loose timeline can be sketched out for them.

As I stated earlier, with the Shuttle-C and Titan V LVs coming into existence ITTL, something is going to have to give, and that's here, where upgrade options for the Delta series will be rather limited. As such, expect it to track largely like OTL for the early years. That is, the Delta 6000 is built as an interim, while the 7000 series is developed. Now there does lie, one piece of help. OTL, the GEM-40 SRBs were initially developed for the Delta 7000, ITTL, they are being developed for the Titan V as well, that could be of help to the both of them, owing to a piece of hardware commonality.

As for the main engine on the 7000. I'd expect it to be the same RS-27A as OTL, NIHS(1) being a key factor. That and I find the performance figures for the RZ.4 and RS-27 are extremely similar, really not much of a gap.

If it's decided to try and uprate the performance of the Delta 7000, I see only two key scenarios:


1. Augmented SRBs. IOTL, the Delta 7000 could be fitted with larger GEM-46 SRBs for improved performance, but that takes their diameter to the largest they can be while still squeezing nine of them onto the core stage.

2. New upper stage. They actually tried this with the Delta 8000 series IOTL, it failed rather spectacularly - only the GEM-46 SRBs survived. In any case, a 305cm LOX/Kerosene upper stage or a 400cm LOX/LH2 upper stage is doable so long as they get it right. But with everything else happening, funding may not be available.


Definitely not gonna be easy to keep on the right side of plausible, but it can be done. Thoughts?


(1) - NIHS Not Invented Here Syndrome

There enough variants of the centaur out there, that id guess a single engine kidsize version would be ,,easy,, to develop. Or is it too big in some dimension?

I dont remember what youve done with altariane here, but something like the european ariane4 third stage of otl, if an equivalent exists, would make a lot of sense. They could claim ,,its just temporary until we get the proper stage built,,.
 
There enough variants of the centaur out there, that id guess a single engine kidsize version would be ,,easy,, to develop. Or is it too big in some dimension?

In terms of length, a usable LOX/LH2 upper stage could be dangerous as the Delta II first stage is already ~28 metres with a diameter of only 2.44m, so they tended to want the total length to be less than 40 metres to help keep it controllable during ascent. Maybe if the Kerosene Tank in the first stage is widened, they could accommodate a small LOX/LH2 stage. But I don't see it being much greater than 6 or 7 metres in length.


I dont remember what youve done with altariane here, but something like the european ariane4 third stage of otl, if an equivalent exists, would make a lot of sense. They could claim ,,its just temporary until we get the proper stage built,,.

Actually, since Europa was successful, there is no Ariane at all. And the LOX/LH2 upper stages they have for Europa III and Argo upper stage are wholly unsuitable for the Delta.
 
Could we get a common core variant, or similar for a delta? Especially if you could fit a couple of engines?

Too expensive to develop, i suppose.


Btw, if part of the problem was that the first stage was too long and skinny, how did it get that way? Incremental length stretches of the original thor, eh?
 
I see no Problem in uprating the Delta in this TL

There is need by NASA, USAF, USNavy and NRO to launch medium satellite or space probes to Mars, Moon, Venus
in OTL the Transformation from Delta 7000 series to Delta 8000 aka Delta 3 was so a step
the RP-1 tank was replace by new one but instead 2.44 meters ø it is now 4 meters ø what reduce the stage length to 20 meters
this Hammerhead configuration is not new, the Titan III-IV and Ariane 1-4 used it years before.

the use of new LOX/LH2 upper stage with RI-10 is only question of financing by NASA or Pentagon.
Alternative MDC could buy Ariane 4 third stage H-10 III who is similar to second Stage of Delta 8000 series.

Why was Delta 8000 aka Delta 3 not a success ?
Delta 8000 was the next upgrade by his manufacture McDonnell/Douglas but during the Program, MDC was take over by Boeing in 1997
and the Boeing Management had no clear concept what to do with this rocket. Next to that were usual problems during R&D

The First Delta Flight 259 had take the Software of Delta II series unchanged,
in result the rocket behaved like in the first Ariane 5 flight and had to be destroy by the safetyrange officer.
On second Delta Flight 269, on his first use in space, the new developed RL-10B engine start to overpressure and exploded on second burn.
The third Delta Flight 280, they made a unpleased discovery that rocket work perfectly,
but brought the payload 3000 km lower orbit as planned (180.76 x 20,694 km x 27.5 deg. versus a planned 185 x 23,404 km)

this Series of misfortune and upcoming EELV program, the Boeing Management take the decision to stop Delta III program in favor of Delta IV

Those Errors had be forestall.
The first launch of Ariane 5 was in 1995, three years before Delta Flight 259! so rewriting the Software
The use of RL-10A in begin in program and later change to RL-10B or buy H-10III stage
last one had brought the Payload into the right Orbit
 
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