Author Topic: Aerojet Rocketdyne's AR-1 engine (aka AJ-1E6)  (Read 230042 times)

Offline ncb1397

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I think you are making this too complicated. The Atlas V weighs about 724,000 pounds and lifts ~21,500 pounds to orbit(or 3% payload). This is a 500,000 pound-force engine, say the stage masses 400,000 pounds. 3% of 400,000 pounds is 12000 pounds or ~ 5500 kg. They would likely have to get the cost down to about $60 million for a government customer to come in lower than F9R would cost and then steal some of the lighter payloads it would be launching mostly empty for. $10 million for the AR-1 and $10 million for the RL-10 would make the business case roughly close which is a bit less than what NASA pays for an RL-10 (but this includes some likely healthy profit margins that can be negotiated against and may not include all the modernized manufacturing savings). Assuming first stage recovery/reflight, you could presumably get 2,500 kg+ to orbit and save on buying a lot of AR-1s.
« Last Edit: 04/18/2019 12:33 am by ncb1397 »

Online ZachS09

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I don't know if this question was asked:

Is there a source that tells how many kilograms the AR-1 weighs?
Liftoff for St. Jude's! Go Dragon, Go Falcon, Godspeed Inspiration4!

Offline Steven Pietrobon

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Dr Steven; if a heavy lifter like SLS (or similar) had a pair of expendable strap on boosters, each powered by 4x AR-1s - what sort of lifting performance could we expect instead of big solids?

Using three AJ1E6 (equivalent to six AR-1) performance is 136.2 t using dual J-2X upper stage. Note that AJ1E6 has 10% greater thrust than two AR-1s, so performance would be a little less with AR-1s. ATK Advanced Boosters (Dark Knights) with dual J-2X has a payload 124.8 t.

Using four AR-1's would probably have a big performance hit, since the thrust from the boosters would be reduced by 33%.

Is there a source that tells how many kilograms the AR-1 weighs?

Not that I know of. The only number we know is the sea level thrust!
« Last Edit: 04/18/2019 05:50 am by Steven Pietrobon »
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline brickmack

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has contract for RL10 with NASA that averages $14.5 million per engine

RL10 is 6.5-11.5 million a piece, depending on RL10A vs B (RL10C-3 for EUS doesn't count, small production run of a custom variant and for a government customer, both mean heavily inflated price). Next-gen RL10C variants with 3d printing should be ~half that.

Not that I know of. The only number we know is the sea level thrust!

AR-1 weighs 17200 pounds for a pair (so like 7.8 tons). ISP is 304.7-337.5 seconds. Mix ratio is 2.71:1 at full throttle, preburner mix ratio is 49:1. Max throttle thrust is 582.5 klbf vac, 526 klbf SL. Chamber pressure is 3100 psi. Expansion ratio is 38.8:1. It can throttle down to 40%. Any other information you require for simulation purposes?
« Last Edit: 04/23/2019 02:26 pm by brickmack »

Offline Patchouli

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Aerojet Rocketdyne is pitching that someone could build a Delta II class rocket using a single AR-1 and single RL-10 as the upper stage. I'd assume some vernier thrusters for roll control as well.
This should outperform Delta 2 by a fair margin.  It would outperform Antares 230+, etc.  It is closer to Atlas 3B and Delta 4 Medium performance.  I figure that 9 metric tons to LEO/ISS should be possible, 8 tonnes to LEO/S, almost 4 tonnes to GTO, and nearly 3 tonnes to escape velocity.   

 - Ed Kyle

Bring back the LR-101 engine for that.

It could be a good fit for a smaller LV for ULA if they wanted to keep the Delta II payload class as even the smallest version of Vulcan is a medium heavy.

Too bad all the crew vehicles are around 10 tons or more unless you want to launch a Soyuz or Shenzhou as with only one main engine in each stage and a single staging event it could in theory be very reliable.
« Last Edit: 04/22/2019 06:28 pm by Patchouli »

Offline Jim

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Bring back the LR-101 engine for that.


No, hydrazine thrusters would be better for that.  See Atlas II

Offline Steven Pietrobon

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AR-1 weighs 17200 pounds for a pair (so like 7.8 tons). ISP is 304.7-337.5 seconds. Mix ratio is 2.71:1 at full throttle, preburner mix ratio is 49:1. Max throttle thrust is 1165 klbf vac, 1052 klbf SL. Chamber pressure is 3100 psi. Expansion ratio is 38.8:1. It can throttle down to 40%. Any other information you require for simulation purposes?

Where did this data come from? This looks like the spec for AJ1E6, not AR-1. The dry mass looks very high. RD-180 is 5.48 t.
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline brickmack

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Where did this data come from? This looks like the spec for AJ1E6, not AR-1. The dry mass looks very high. RD-180 is 5.48 t.

A ULA presentation. Numbers are close to, but not the same as, the last ones I saw for AJ1E6/AJ500 (I edited my previous post, thrust figure was for the dual engine configuration, halved that for single-engine). Which makes sense given AR-1 is the direct continuation of that program (note how Aerojet's bid for the SLS Block 2 boosters changed from AJ1E6 to AR-1 with effectively no performance or external design change also)

Offline Steven Pietrobon

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A ULA presentation. Numbers are close to, but not the same as, the last ones I saw for AJ1E6/AJ500 (I edited my previous post, thrust figure was for the dual engine configuration, halved that for single-engine).

Thanks for the clarification. Is there a link to this presentation? What about the dry mass? Is it 3.9 t for AR-1? That would be very high. RD-191 with similar thrust (1.92 MN vs 2.34 MN for AR-1) is only 2.29 t.
« Last Edit: 04/24/2019 04:22 am by Steven Pietrobon »
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline hkultala

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AR-1 has about 220 tonnes of sea level thrust. Antares and Atlas IIIB have about 380 tonnes of sea level thrust. I don't see single AR-1 outperforming those, even though it would have more efficient upper stage than antares has.

Well, for a very rough calculation...

220 tonnes thrust with a 1.2 liftoff T/W gets you a ~175t rocket with a 9t payload.

If the first stage is 75% of the rocket (130t) and has a 0.92 propellant mass fraction (a PMF a bit worse than F9), then you have ~120.3t prop and ~10.5t structure.

Using RD-180's 311/338s ISP as a stand-in, and using the rule-of-thumb that says the average ISP over a booster flight is the Sea Level ISP + 2/3rds of the SL/Vacuum difference, gets you an average ~329s ISP over the flight.

So: 329*9.806*ln((175+9)/((175+9)-120.3))) = ~3420m/s for first stage flight.


The second stage is 25% of the rocket (~43.6t) and, if it has a .89 PMF (a bit worse than Centaur), then ~38.8t of it is prop. RL-10-C's ISP is ~450s.

So: 450*9.806*ln((43.6+9)/((43.6+9)-38.8))) = ~5900m/s for second stage flight.

Which gets you a total ~9320m/s dV, which does seem like enough to put something into LEO.



You are ignoring one very important point: The terrible gravity drag of the second stage. The second stage would simply drop from the sky before reaching orbital speed.

450s isp and 110 kN of thrust means it's burning 25 kg of propellant per second.
This means that in order to burn 38.8 tonnes of propellant, it's burn would last about 1550 seconds.



Centaur can survive with (but suffers from) a low T/W ratio because it stages very high in all rockets where it's used, but you are proposing much lower staging for upper stage with even lower T/W ratio.

Lets say our 3.2 km/s for lower stage contains 1 km/s of gravity losses(as initially 83% of the thrust is used for fighting gravity). Then our velocity is 2.2 km/s at staging.

At 30 degrees angle this means 1.1km/s vertical velocity 1.9km/s horizontal velocity.

This 1.1km/s vertical velocity means that the first stage has kinetic energy to keep ascending for 60.5 kilometers, 110 seconds.

Lets take the second stage into account:



Your rocket would have initial second stage T/W of only 0.244. If initially rising at 30 degree angle, this would  mean only about 1.23 m/s^2 of vertical thrust component. Assuming the 1.9km/s horizontal velocity reduces effective gravity by 25% (not sure if this is linear in reality) we are losing about 6.1 m/s^2 vertical velocity.

So, with it's initial weight and T/W ratio the second stage would continue rising for only 180 seconds.

But it keeps getting lighter?

After 200 seconds it would have burned only 5 tonnes of propellant. It would still weight 38.6 and have T/W of only 0.285. At 30 degree angle this gives about 1.4 m/s^2 of vertical thrust component.

It would have gained only about 465 m/s of vertical velocity, going at about 2.47km/s. Assuming this reduces effective gravity to about 6.5 m/s^2, it would still have 5.1 km/s of downwards acceleration. So with this weight, it would keep rising for 215 seconds.

So, in practice it would reach it's apogee at about 200 seconds, and then it would fall down.

No amount of delta-v helps when there is no time to use it.


What if trying to use more lofted trajectory? Make the staging at 45 degrees.

Then at staging it has 1.55 km/s vertical velocity and 1.55 km/s horizontal velocity. About 155 seconds of rise time for the first stage after stage separation.

Now the horizontal speed is 350 m/s less and we get intial effective dravity drop of 20% instead of 25%.

Second stage burning at 45 degrees also? Now our vertical thrust component is 1.7 km/s^2 instead of 1,2 km/s^2. We are still losing vertical velocity 6.1 km/s^2 , which means rise time of about 255 seconds.

After burning for 240 seconds, 6 tonnes of propellant is burned.  We have now T/W of about 2.9. The vertical component of this is about 2 m/s^2.

It has gained about 460 m/s of horizontal velocity, which is now about 2.01 km/s. Effective gravity is still about 7.2 m/s^2 , so total downwards forces 5.2 m/s^2, rise time ~300 seconds.

After burning for 480 seconds, 12 tonnes of propellant is burned. We have now T/W of about 0.35.
The vertical component of this is about 2.4 m/s^2.

It has gained about 990 m/s of horizontal velocity, which is not about 2.54 km/s. Effective gravity about 6.5 m/s^2, so total downwards acceleration is about 4.1 m/s^2 , rise time ~380 seconds.

So, in practice it would start falling down after something like 320 seconds.


No amount of lofting will make this work. Even if there was no gravity losses in the first stage and it would burn vertically, the first stage would have ballistic ascent time of only 320 seconds after stage separation, but the seconds stage needs 1550 seconds to burn it's fuel.


The first stage would need to be longer and second stage much smaller.

Aim for 7.5-tonne payload, make the first stage weight 150 tonnes and second stage 17.5 tonnes. Then your first stage has ~5 km/s of delta-v and second stage has T/W of 0.44, delta-v of ~4.2 km/s and burn time of ~612 seconds. This might work.

... but this is much less than the payload of Atlas IIIB and slightly less than the payload of Antares 230.


With DEC(Dual-Engine-Centaur) upper stage and about 145-tonne first stage it might beat Antares 230 to LEO, but would still easily lose to Atlas IIIB



« Last Edit: 04/24/2019 09:52 am by hkultala »

While the presence of OmegA has made most of us (including me initially) dismiss the idea of using AR1 on Antares, reusability could be the thing that makes Antares a viable commercial launcher. If the AR1 really can power a reusable launch vehicle, it's something that should be looked into.

Imagine for a moment that the Antares 300 first stages are stretched stages powered by 2 AR1s. The stretch is to accommodate carrying whatever method of RTLS NGIS would pick. Boost-back is the most obvious now, but I personally still hold out hope we'll see a glide-back booster one day.

How much cheaper would a reusable Antares be to operate, and what sort of second stage would be needed. It would be easiest (and cheapest) to use the Castor 30XL, but I worry that might not have enough delta-V to be practical on an RLV.
Wait, ∆V? This site will accept the ∆ symbol? How many times have I written out the word "delta" for no reason?

Offline hplan

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While the presence of OmegA has made most of us (including me initially) dismiss the idea of using AR1 on Antares, reusability could be the thing that makes Antares a viable commercial launcher. If the AR1 really can power a reusable launch vehicle, it's something that should be looked into.

Imagine for a moment that the Antares 300 first stages are stretched stages powered by 2 AR1s. The stretch is to accommodate carrying whatever method of RTLS NGIS would pick. Boost-back is the most obvious now, but I personally still hold out hope we'll see a glide-back booster one day.

How much cheaper would a reusable Antares be to operate, and what sort of second stage would be needed. It would be easiest (and cheapest) to use the Castor 30XL, but I worry that might not have enough delta-V to be practical on an RLV.

Reusability on two engines? I don't see how that could work, at least if you are thinking powered vertical landing. To have symmetrical thrust about the center axis, you'd need to have both engines powered on descent. But then you'd have a really hard time powering down enough to actually land, unless the engines can power down to something like 10%...

Online edkyle99

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While the presence of OmegA has made most of us (including me initially) dismiss the idea of using AR1 on Antares, reusability could be the thing that makes Antares a viable commercial launcher. If the AR1 really can power a reusable launch vehicle, it's something that should be looked into.

Imagine for a moment that the Antares 300 first stages are stretched stages powered by 2 AR1s. The stretch is to accommodate carrying whatever method of RTLS NGIS would pick. Boost-back is the most obvious now, but I personally still hold out hope we'll see a glide-back booster one day.

How much cheaper would a reusable Antares be to operate, and what sort of second stage would be needed. It would be easiest (and cheapest) to use the Castor 30XL, but I worry that might not have enough delta-V to be practical on an RLV.
While AR1 could conceivably be adapted for Antares, a reusable Antares is simply not going to happen, certainly not with AR1 propulsion, nor would it cost less than the current expendable Antares.  Northrop Grumman's goal is steady U.S. Government launch services work via. Minotaur, Antares, and if they can sell it Omega.  In my opinion. 

 - Ed Kyle 
« Last Edit: 05/07/2019 09:51 pm by edkyle99 »

Offline TrevorMonty

While the presence of OmegA has made most of us (including me initially) dismiss the idea of using AR1 on Antares, reusability could be the thing that makes Antares a viable commercial launcher. If the AR1 really can power a reusable launch vehicle, it's something that should be looked into.

Imagine for a moment that the Antares 300 first stages are stretched stages powered by 2 AR1s. The stretch is to accommodate carrying whatever method of RTLS NGIS would pick. Boost-back is the most obvious now, but I personally still hold out hope we'll see a glide-back booster one day.

How much cheaper would a reusable Antares be to operate, and what sort of second stage would be needed. It would be easiest (and cheapest) to use the Castor 30XL, but I worry that might not have enough delta-V to be practical on an RLV.
VTOHL RLV is quite within NG capabilities, they even have design from XS-1 competition. A 1xAR1 RLV would fit nicely in their design as it was RP1 LOX powered, size would need scaling to match AR1 but most likely similar size to XS1 design. Ideal for competiting against new 1000-1500kg class ELVs being developed.

Omega is probably better option for Antares replacement.







Offline brickmack

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Reusability on two engines? I don't see how that could work, at least if you are thinking powered vertical landing. To have symmetrical thrust about the center axis, you'd need to have both engines powered on descent. But then you'd have a really hard time powering down enough to actually land, unless the engines can power down to something like 10%...

A spaceplane could work, though it'd be heavy (though Antares core stage is pretty overweight as it is, both because of the heavy Ukrainian construction and the lack of a stretch to take maximum advantage of RD-181. A clean-sheet American design built around AR-1, even with reuse, might not perform much worse) and would have to be an almost complete redesign of the whole vehicle. Doesn't make sense unless the flightrate improves a bunch, and unlike [nameless other reusable vehicles] I don't think thats likely to happen as long as Antares or its derivatives have a solid upper stage. Its pretty much only useful for LEO, both because of the terrible ISP and the lack of restartability. Antares has always been cheaper than Atlas V 401 for not much worse LEO performance, but has failed to get actual non-Cygnus missions anyway. Further Antares prospects died with the High Energy Second Stage

SMART-style reuse would be much easier, with almost no development needed, minimal mass impact, and it can probably break even in only 1 or 2 flights.

In either case though, I don't see that AR-1 makes sense to power this hypothetical vehicle.  RD-181 has effectively identical performance, the RD-170 family as a whole is reusable to a comparable degree as AR-1 targets, both RD-181 and Antares-RD-181 already exist and are flying, and with reuse the geopolitical issues around RD-180/181 largely go away (Aerojet/Pratt & Whitney has the ability to service them and even build some parts. RD-180 coproduction may not have made financial sense, but maintenance is an entirely different matter). Little to no money going back to Russia, little to no risk of interrupted launch capability. What does AR-1 bring to the table, except hundreds of millions in extra dev/qualification costs before you can even *consider* the development needed to reuse anything? I want to see Antares reach something along the lines of its original goals, but this ain't it. If anything, replacing the Zenit-derived tanks is probably a higher priority both technically and geopolitically. But thats a separate thread

Offline Tomness

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https://news.lockheedmartin.com/2020-12-20-Lockheed-Martin-to-Acquire-Aerojet-Rocketdyne-Strengthening-Position-as-Leading-Provider-of-Technologies-to-Deter-Threats-and-Help-Secure-the-United-States-and-its-Allies

Quote

**Snip**
*-Thread bump**
 in light of this breaking news, could AR-1 be  resurrected for the Atlas V have Atlas V and Vulcan continue to launch in tandem if this merger goes through?
« Last Edit: 12/28/2020 03:54 pm by zubenelgenubi »

Offline DreamyPickle

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https://news.lockheedmartin.com/2020-12-20-Lockheed-Martin-to-Acquire-Aerojet-Rocketdyne-Strengthening-Position-as-Leading-Provider-of-Technologies-to-Deter-Threats-and-Help-Secure-the-United-States-and-its-Allies

In light of this breaking news, could AR-1 be  resurrected for the Atlas V have Atlas V and Vulcan continue to launch in tandem if this merger goes through?
ULA is pressured to decrease costs, running Atlas V and Vulcan in parallel would achieve the opposite. They definitely want to get down to a single rocket on 2 pads.

Vulcan itself is very far along in development and would be ready much sooner than any rocket powered by AR-1.

Could an AR-1 powered Atlas even fly the "class C" direct-to-GEO mission?

Offline Sam Ho

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https://news.lockheedmartin.com/2020-12-20-Lockheed-Martin-to-Acquire-Aerojet-Rocketdyne-Strengthening-Position-as-Leading-Provider-of-Technologies-to-Deter-Threats-and-Help-Secure-the-United-States-and-its-Allies

In light of this breaking news, could AR-1 be  resurrected for the Atlas V have Atlas V and Vulcan continue to launch in tandem if this merger goes through?
ULA is pressured to decrease costs, running Atlas V and Vulcan in parallel would achieve the opposite. They definitely want to get down to a single rocket on 2 pads.

Vulcan itself is very far along in development and would be ready much sooner than any rocket powered by AR-1.

Could an AR-1 powered Atlas even fly the "class C" direct-to-GEO mission?
ULA is already planning to fly Atlas V and Vulcan in parallel, sharing SLC-41, because only Atlas V is certified for crew and nuclear.  That said, I don't see ULA investing in a reengined Atlas V.  Also, crew and nuclear are both for civil space; ULA is free to buy more RD-180s for civil use.

Offline russianhalo117

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https://news.lockheedmartin.com/2020-12-20-Lockheed-Martin-to-Acquire-Aerojet-Rocketdyne-Strengthening-Position-as-Leading-Provider-of-Technologies-to-Deter-Threats-and-Help-Secure-the-United-States-and-its-Allies

Quote

**Snip**
*-Thread bump**
 in light of this breaking news, could AR-1 be  resurrected for the Atlas V have Atlas V and Vulcan continue to launch in tandem if this merger goes through?

This is completely targeted at NG in response to NG buying OATK to position them favourably against Boeing et al while also acting as a check against certain Boeing led projects.
« Last Edit: 12/28/2020 03:54 pm by zubenelgenubi »

Offline Kryten

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https://twitter.com/AerojetRdyne/status/1348992893051154432
Quote
The first #AR1 engine is complete the first American-made liquid oxygen/ kerosene staged-combustion engine. AR1 is the ideal engine for many possible solutions; it brings the right thrust level, size, and performance to a wide variety of launch vehicles.

(We should probably take 'proposed' out of the thread title)

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