Arianespace launch thread

[sevenperforce]

22 minutes ago, tater said:

Maybe a handful of Arianespace employees can work some of these issues—it will give them something to do for the next year/whatever while they are not launching rockets.

Course the ones who would actually do this work are working on a slow-follower version of the F9, and no such changes to Ariane 6 will ever happen or even be considered—just as frankenrocket changes to SLS will never happen.

If you wanted to REALLY frankenrocket things up, you could use the Vulcain 2 engine to get a turbineless scramjet-based HTOL vehicle off the ground and up to ram compression speed, let the airbreathers get you moving high and fast without sucking up too much of your hydrogen, then switch the Vulcain 2 back on to provide the rest of the push to a high-suborbital trajectory, like 4 km/s. Then release a cryogenic upper stage from inside a fairing to deliver a comsat payload to GTO. The HTOL vehicle would be able to handle the re-entry with minimal shielding and coast back home.

[tater]

Just now, sevenperforce said:

If you wanted to REALLY frankenrocket things up, you could use the Vulcain 2 engine to get a turbineless scramjet-based HTOL vehicle off the ground and up to ram compression speed, let the airbreathers get you moving high and fast without sucking up too much of your hydrogen, then switch the Vulcain 2 back on to provide the rest of the push to a high-suborbital trajectory, like 4 km/s. Then release a cryogenic upper stage from inside a fairing to deliver a comsat payload to GTO. The HTOL vehicle would be able to handle the re-entry with minimal shielding and coast back home.

I would love to see some really out of the box ideas. Maybe at some point one of these other rocket outfits will do this to try and make themselves relevant. Stoke springs to mind as trying something really novel.

The rest is just slow-following Falcon 9 (New Glenn, Neutron, Terran R, etc).

[sevenperforce]

1 minute ago, tater said:

I would love to see some really out of the box ideas. Maybe at some point one of these other rocket outfits will do this to try and make themselves relevant. Stoke springs to mind as trying something really novel.

I like altitude-compensating nozzles, so I really want to see a better sustainer architecture.

 So far we have three main sustainer architectures:

• Wimpy core hydrolox stage with beefy boosters to get T/W over 1 (Ariane 5, SLS, STS, LM-5)
• Beefier core stage with large upper stage and optional small solid boosters (Atlas V, Delta IV M, Ariane 4)
• Two-stage architecture with multi-stick first stage (Delta IV Heavy, Falcon Heavy, Soyuz)

I want something more interesting. 

[wumpus]

31 minutes ago, sevenperforce said:

I like altitude-compensating nozzles, so I really want to see a better sustainer architecture.

 So far we have three main sustainer architectures:

• Wimpy core hydrolox stage with beefy boosters to get T/W over 1 (Ariane 5, SLS, STS, LM-5)
• Beefier core stage with large upper stage and optional small solid boosters (Atlas V, Delta IV M, Ariane 4)
• Two-stage architecture with multi-stick first stage (Delta IV Heavy, Falcon Heavy, Soyuz)

I want something more interesting. 

The gotcha is that 3 launcher platforms that solved altitude compensation all failed.

Pegasus (+XL) is basically retired
Stratolaunch isn't interested in orbital launches
Virgin Galactic (the air-launch to orbit company) is bankrupt and nobody is interested in the rocket

Not to mention, using a 2 stage rocket gives you ideal vacuum Isp  for most of the delta-v needed.  Didn't somebody slap a moveable extension onto a sustainer already, or was the proposal tabled before construction/launch?  I think China has a launch facility at altitude, but no idea how that will affect nozzles and pressure design.

Edited July 10, 2023 by wumpus
single spacing different from excell behavior.

[sevenperforce]

13 minutes ago, wumpus said:

The gotcha is that 3 launcher platforms that solved altitude compensation all failed.

Pegasus (+XL) is basically retired
Stratolaunch isn't interested in orbital launches
Virgin Galactic (the air-launch to orbit company) is bankrupt and nobody is interested in the rocket

Yeah, air-launch is not my favorite. Not unless the separation is exoatmospheric.

13 minutes ago, wumpus said:

Not to mention, using a 2 stage rocket gives you ideal vacuum Isp  for most of the delta-v needed. 

The (perhaps deceptively) attractive element of parallel staging is that you can get better T/W at liftoff without needing more engine mass. Of course, crossfeed makes it much more advantageous.

16 minutes ago, wumpus said:

Didn't somebody slap a moveable extension onto a sustainer already, or was the proposal tabled before construction/launch? 

I don't believe any vehicle has ever launched with a first-stage moveable nozzle extension, except maybe some of the earliest solid rocket ICBM designs.

16 minutes ago, wumpus said:

I think China has a launch facility at altitude, but no idea how that will affect nozzles and pressure design.

If the air pressure is high enough to support breathing, it's essentially sea level for the purposes of nozzle design.

[Beccab]

12 minutes ago, sevenperforce said:

I don't believe any vehicle has ever launched with a first-stage moveable nozzle extension, except maybe some of the earliest solid rocket ICBM designs.

Nobody actually built anything with that architecture, but it was somewhat heavily proposed during the early shuttle program for drop tank designs, such as Starclipper. The relative engine is the XLR-129

[wumpus]

6 minutes ago, sevenperforce said:

The (perhaps deceptively) attractive element of parallel staging is that you can get better T/W at liftoff without needing more engine mass. Of course, crossfeed makes it much more advantageous.

This seems to be one of those things that KSP <1.0 taught that wasn't quite right.  Of course it also taught (and still teaches) that a RL-10 engine is cheap.  Aerodyne says otherwise.  But you need a lot less thrust on a second stage than you'd put in a sustainer, and that weight comes out of your cargo capacity (much worse than forcing the first stage to lift the dead weight of the second stage engine).

I think the most important consideration is just how confident you are that the second stage will light.  If you can get it to light, then you want a two stage rocket.  If you aren't so confident, then a sustainer is ideal.  Another thing to unlearn from KSP.

And it looks like the highest Chinese launch site  gets .8 bar at launch.  Helps a little, but might not be worth redesigning a nozzle.  You'd be surprised how thin air can be and still be able to breathe (a lot higher than that).

[sevenperforce]

1 hour ago, Beccab said:

Nobody actually built anything with that architecture, but it was somewhat heavily proposed during the early shuttle program for drop tank designs, such as Starclipper. The relative engine is the XLR-129

The XLR-129 would have been a variable-mixture-ratio fuel-rich staged combustion engine, similar to the RS-25 but with only a single preburner feeding two pumps rather than two separate preburners.

I'm reminded of the RD-701 which also would have been used for an SSTO-like or SSTO-lite application, although it was an ORSC architecture that used only kerosene for the preburners but switched between kerosene and liquid hydrogen in the main chamber.

Sustainer architectures lend themselves to inventive concepts for increasing thrust at liftoff and increasing specific impulse at altitude, and so they really trigger my weakness for clever engineering. It would be cool to have a tripropellant engine with an ORSC kerosene-based LOX pump and a split hydrogen-based expander cycle fuel pump, for example...especially one with a nozzle which would be sea-level expanded in mostly-kerosene-mode and vacuum-expanded in mostly-hydrogen-mode, so that changing the mixture ratio and propellant during ascent would automatically provide altitude compensation. Cooler still if you could use crossfeed from a strap-on booster assembly (maybe even a fly-back booster) to provide the kerosene, minimizing tankage on the main stage.

Of course if you're going for full reusability this doesn't help EDL.

1 hour ago, wumpus said:

But you need a lot less thrust on a second stage than you'd put in a sustainer, and that weight comes out of your cargo capacity (much worse than forcing the first stage to lift the dead weight of the second stage engine).

Then again, the weight of a second-stage engine is typically much larger than the weight of a similar-flow-rate first stage engine (e.g. Merlin, Rutherford, Raptor), so if you can use a larger engine bell with some sort of thrust augmentation (especially crossfed) then you're winning.

1 hour ago, wumpus said:

And it looks like the highest Chinese launch site  gets .8 bar at launch.  Helps a little, but might not be worth redesigning a nozzle.

I believe most sea-level engines are actually optimized for around 2-3 km altitude: a little overexpanded at launch in order to be less underexpanded at altitude. The optimal expansion ratio depends not only on the integration of specific impulse across the flight regime to maximize actual delta-v but also on reducing delta-v lost to gravity drag from low thrust. I suppose they might be designed for a slightly higher altitude if they are launching from a higher elevation, but I don't think it would be worth a redesign unless those rockets are ONLY going to be launching from high elevation and never from closer to sea level.

1 hour ago, wumpus said:

You'd be surprised how thin air can be and still be able to breathe (a lot higher than that).

Fair point.

[Exoscientist]

On 7/9/2023 at 2:09 PM, AckSed said:

Science fight! Science fight!

Edit: Say we have a thrust-augmented nozzle that injects extra oxygen to increase thrust. Not looking for anything special, just a bit more mass-flow at the start for a O/F ratio of say 13:1 for 30 seconds. Does that do anything for the lift-off mass?

 An increased thrust will help with the take-off thrust weight ratio(TWR) so reduce the gravity drag. That would certainly help with the two Vulcain case. The three Vulcain case probably already has sufficient thrust so wouldn’t need it.

By the way, there might be a simpler way of achieving the same thing as the thrust-augmented increased oxygen flow. That is if the Vulcain can be made to be of variable mixture ratio. On take-off you would increase the oxygen rate to increase thrust. This though would decrease Isp so you would only use it for a few seconds during take-off for the added thrust. Thereafter when the propellant burn has made the rocket lighter you would switch back to the usual mixture ratio to get the high Isp.

 The thrust-augmented nozzle(TAN) does have a key advantage though that it makes it useful even for cases of high take-off TWR. It has altitude compensation effects. This refers to the fact the efficiency of a rocket engine is reduced by the dense air pressure at sea level. The large nozzles you would like to use to maximize Isp can be dangerous at sea level and can even destroy an engine. Various techniques known as altitude compensation have been developed to get high efficiency in vacuum and get the high thrust desired at sea level. TAN is one such technique.

Altitude compensating nozzles, or adaptive nozzles, is one of those methods that have been known about for years to be able to increase payload but hasn’t been used. I mentioned that an all-liquid Vulcain at 7.5% payload fraction would mean other launch companies such as SpaceX would be forced to catch up to this, and most simply could not. 

 I thought maybe the Falcon 9 could do it by using the same ultra efficient lightweighting used on the Ariane 5 core. Since kerolox is 3 times denser than hydrolox, the 16.3 to 1 mass ratio of the Ariane 5 core would correspond to a mass ratio for the Falcon 9 first stage of 50 to 1(!). But after running the numbers this still would not be enough for the Falcon 9 to get to 7.5% payload fraction. What might work though is if you also used altitude compensation on the first stage of the Falcon 9.

 The problem with the Falcon 9 being able to get such a high payload going from 22 tons to the 40 tons of a 7.5% payload fraction, is the Merlin on the first stage has such poor vacuum Isp at ca. 312s. But note with altitude compensation its vacuum Isp could be raised to > 360 s.

 But then the Falcon 9 using altitude compensation, would lead to ArianeSpace applying the same to the Ariane 5 core. In that case the vacuum Isp of the Vulcain could be raised to > 470s(!) I estimate the all-liquid Ariane 6 could then get a payload fraction of 10%. And this I don’t think the Falcon 9 could reach even using an ultra lightweighted first stage and using altitude compensation.

   Robert Clark

 

[sevenperforce]

22 minutes ago, Exoscientist said:

By the way, there might be a simpler way of achieving the same thing as the thrust-augmented increased oxygen flow. That is if the Vulcain can be made to be of variable mixture ratio. On take-off you would increase the oxygen rate to increase thrust. This though would decrease Isp so you would only use it for a few seconds during take-off for the added thrust. Thereafter when the propellant burn has made the rocket lighter you would switch back to the usual mixture ratio to get the high Isp.

You wouldn't want to only use it for a few seconds. You'd want a smooth transition as you climbed, balancing the need for higher specific impulse against the need to avoid gravity drag.

Gravity drag is a REALLY big deal.

22 minutes ago, Exoscientist said:

Altitude compensating nozzles, or adaptive nozzles, is one of those methods that have been known about for years to be able to increase payload but hasn’t been used.

Usually because altitude-compensating nozzles won't actually increase payload.

22 minutes ago, Exoscientist said:

I mentioned that an all-liquid Vulcain at 7.5% payload fraction would mean other launch companies such as SpaceX would be forced to catch up to this

There's no reason that other launch companies would need to catch up to any particular payload fraction. Launch companies try to catch up to cost; they don't care about catching up to a particular payload fraction.

22 minutes ago, Exoscientist said:

I thought maybe the Falcon 9 could do it by using the same ultra efficient lightweighting used on the Ariane 5 core. Since kerolox is 3 times denser than hydrolox, the 16.3 to 1 mass ratio of the Ariane 5 core would correspond to a mass ratio for the Falcon 9 first stage of 50 to 1(!).

The mass ratio of the Ariane 5 core is based on a tank which is partially in tension rather than compression for the highest-stress portions of the flight and thus can be very lightweight. So translating to the Falcon 9 tanks won't work for that reason. It also is not a 16.3:1 mass ratio. Additionally, the LOX tank is not significantly different in weight than a typical LOX tank; the difference is exclusively in the hydrogen tank. You can't just take the (incorrect) 16.3:1 mass ratio and multiply by the difference in density.

22 minutes ago, Exoscientist said:

But after running the numbers this still would not be enough for the Falcon 9 to get to 7.5% payload fraction.

Which it doesn't have any reason to do anyway.

22 minutes ago, Exoscientist said:

What might work though is if you also used altitude compensation on the first stage of the Falcon 9.

The Merlin 1D already uses a nozzle that is designed for optimal performance at all altitudes in its trajectory.

22 minutes ago, Exoscientist said:

the Merlin on the first stage has such poor vacuum Isp at ca. 312s. But note with altitude compensation its vacuum Isp could be raised to > 360 s.

It absolutely could not. The Merlin 1D Vacuum has a specific impulse of 348 seconds; there is no nozzle in existence that could raise the vacuum specific impulse of ANY Merlin engine to 360 seconds, let alone a nozzle which could be fired at sea level.  

22 minutes ago, Exoscientist said:

the vacuum Isp of the Vulcain could be raised to > 470s(!)

It absolutely could not. The Vulcain engine already has near its maximum vacuum specific impulse for its engine cycle. A gas generator hydrolox engine is not going to be able to get above 445-450 seconds. We have never ever had a hydrolox engine with over 470 seconds of specific impulse. 

[Exoscientist]

4 hours ago, sevenperforce said:

There's no reason that other launch companies would need to catch up to any particular payload fraction. Launch companies try to catch up to cost; they don't care about catching up to a particular payload fraction.

 The high cost of the Ariane 62 and Ariane 64 compared to the Falcon 9 is definitely largely due to the fact ArianeSpace is Old Space, which means their costs are always inflated. But it is notable and relevant that Falcon 9 has one of the highest payload fractions now available at ~4%, and the current Ariane 6 versions with SRB’s will have one of the worst at ~2%.

  Bob Clark

Edited July 11, 2023 by Exoscientist

[sevenperforce]

23 hours ago, Beccab said:

The clear disadvantages in this design are (a) no obvious abort mode and (b) overweight engines carried all the way to orbit. Every kg that can be moved off the orbiter is an extra kg of payload, so if there was a way to offload two of those three engines it would be a significant benefit.

Obviously reducing dead weight carried by the first stage isn't a 1-to-1 benefit to payload, but it does something. If you can do crossfeed, it's absolutely worth doing a sustainer architecture IF it doesn't result in a less-efficient upper stage engine and IF it doesn't force other suboptimal design choices.

A three-core crossfed design with the side boosters returning Falcon Heavy style and the center core going to orbit is the straightforward choice, but then EDL of the orbital stage (for the sake of reusability) becomes the challenge. With crossfeed and a potential tripropellant sustainer, the disadvantages of hydrolox are ameliorated because the nice thrusty boosters carry the fluffy hydrolox tanks out of the atmosphere where drag is no longer an issue, and drop them off full. Then the fluffy tanks help reduce heat load on EDL.

I wonder how large and draggy a hydrogen tank would need to be in order to reduce heat load to the "hot structure" regime.

Of course you still need descent control and a landing mode. You can do a standard biconic entry with split flaps but that doesn't provide a straightforward landing mode.

[tater]

7 minutes ago, sevenperforce said:

The clear disadvantages in this design are (a) no obvious abort mode and (b) overweight engines carried all the way to orbit. Every kg that can be moved off the orbiter is an extra kg of payload, so if there was a way to offload two of those three engines it would be a significant benefit.

As a crew vehicle, I think payload mass is less important (past the crew+vehicle). At the time engine cost likely dominated launch cost, so "save the engines" like Shuttle.

[tater]

https://www.esa.int/Newsroom/Press_Releases/Media_invitation_Ariane_6_test_campaign_update

Quote

The 4 September media briefing will also provide an update on the next important milestones, these being the upper stage Hot Firing Test 3 (HFT3) to be held in Lampoldshausen on 1 September (*) followed later that month, by a long firing test of the core stage and its engine, the Vulcain 2.1, in Kourou, tentatively scheduled on 26 September (*).

The latter will also give engineering teams all the results needed to define a launch period for the Ariane 6 inaugural flight in 2024.

So at some point after testing in September, they will start talking about NET dates in 2024.

[tater]

 

[tater]

https://twitter.com/SpaceNews_Inc/status/1708863227067527516

Connection sucks here.

Vega C return to flight moved to late 2024

Edited October 5, 2023 by tater

[tater]

 

 

[Exoscientist]

On 10/2/2023 at 11:37 AM, tater said:

https://twitter.com/SpaceNews_Inc/status/1708863227067527516

Connection sucks here.

Vega C return to flight moved to late 2024

ESA delays Vega C return to flight to late 2024
Jeff Foust
October 2, 2023
https://spacenews.com/esa-delays-vega-c-return-to-flight-to-late-2024/

 Large solids like on the Vega and as used on the Ariane 5 and 6 are not price competitive. Note this is true for large solids. Small solid side boosters like used on the Atlas V and Delta IV might be only 1/8th the size of the core stage, with a concomitant small increase in cost. But when the solids are large size such as being as much or more than the size of the core such as on the Ariane 5 and 6 or actually being the core like on the Vega, the bulk of the expense of the rocket comes from the solids.

See discussion here:

Friday, May 19, 2023
Who in European space will ask the impertinent question: How much would it cost to add a second Vulcain to the Ariane 5/6?
https://exoscientist.blogspot.com/2023/05/who-in-european-space-will-ask.html

 The cost of the two SRB’s on the Ariane 62 cost €40 million out of the €75 million cost. So the rest of the two-stage rocket is only €35 million. Then those two large SRB’s cost more than the entire rest of the rocket.

 As I argued there it would be cheaper just to put additional Vulcain(s) on the core and dispense with the SRB’s entirely. An additional Vulcain would add €10 million to the price to bring it to €45 million.

 Using all liquid propulsion also results in a cheaper rocket than the Vega. To see what such an all-liquid replacement for the Vega would look like see discussion here:

Saturday, November 29, 2014
A half-size Ariane for manned spaceflight.
https://exoscientist.blogspot.com/2014/11/a-half-size-ariane-for-manned.html

 By cutting down the core’s propellant size to a bit less than half and using a smaller ca. 10 ton upper stage, so it could be launched by a single Vulcain, you get an all-liquid two-stage rocket capable of about 5,000 kg to LEO. This compares to the 2,000 kg payload to LEO of the Vega.

 Quite important is the better cost per kilo for the all-liquid case. The Vega costs about €35 million for that 2,000 kg to LEO. But taking into account our all-liquid replacement to the Vega is half-size to the all-liquid Ariane 6, the cost conceivably could be in the range of only half the €45 million estimate of the all-liquid Ariane 6, so only ca. €22 million for a 5,000 kg to LEO launcher(!)

 And what about reusability? The Space Shuttle abundantly showed you don’t save on reusing solids. But SpaceX has abundantly showed you do save significantly on reusing a liquid-fueled booster. SpaceX reduces the price on the Falcon 9 from $60 million to $40 million, by reusing the booster only, so a price reduction of about one-third. If the same price reduction would apply for reusing the booster only for our half-sized Ariane, that would be a price of only €15 million for a 5,000 kg launcher(!)

  Bob Clark

 

[sevenperforce]

7 hours ago, Exoscientist said:

It would be cheaper just to put additional Vulcain(s) on the core and dispense with the SRB’s entirely. An additional Vulcain would add €10 million to the price to bring it to €45 million.

Both the Ariane 5 and the Ariane 6 have a 5.4-meter core. Vulcain's gimbal range doesn't appear to be publicly available, but assuming it's in the range of 8°, a Vulcain 2.1 engine (at 3 meters high and 1.76 meters wide) will need gimbal clearance of 0.42 meters. So a pair of Vulcains have a 3.9-meter footprint that will fit underneath the core just fine; doing three in a line won't work (that would take up 6.12 meters). You can do three if you put them in a triangular cluster, though.

You'll need all three, because Vulcain 2.1 has a sea level thrust of 970 kN. Ariane 6 has a gross liftoff weight of around 190 tonnes sans boosters, which means two Vulcains gets you an unacceptable T/W ratio of 1.04. Factoring in the weight of two additional Vulcain 2s, the trio would get a liftoff T/W ratio of 1.53 which is good enough.

Not necessarily good enough to get to GTO with any meaningful payload, though.

[Kermann Nolandung]

On 10/5/2023 at 5:29 AM, Exoscientist said:

and using a smaller ca. 10 ton upper stage...

I won't help you much but I read your interrogation regarding ESC-A upper stage.
I think someone has mixed up kg and lbs values of the empty mass (2100 kg / 4600 lbs) and spead the error on internet (at least on the English, French and German Wikipedia pages).  

I guess we should consider for ESC-A
Empty mass            2100 kg (instead of 4600 kg)
Fuel                          14400 kg
Gross mass          16500 kg (instead of 19 t)

As stated here : Ariane 5-2 ESC A (astronautix.com)

Edited October 8, 2023 by Kermann Nolandung
The mix-up seems to be on astronautix page.

[Exoscientist]

On 10/6/2023 at 12:13 AM, Kermann Nolandung said:

I won't help you much but I read your interrogation regarding ESC-A upper stage.
I think someone has mixed up kg and lbs values of the empty mass (2100 kg / 4600 lbs) and spead the error on internet (at least on the English, French and German Wikipedia pages).  

I guess we should consider for ESC-A
Empty mass            2100 kg (instead of 4600 kg)
Fuel                          14400 kg
Gross mass          16500 kg (instead of 19 t)

As stated here : Ariane 5-2 ESC A (astronautix.com)

 It might be some discussions of the ESC-A are including the mass of the “VEB”, the vehicle equipment bay, an instruments package, in its dry mass.

Article on the VEB:

“The vehicle equipment bay (VEB)is often called the ‘brains’ of a launcher. Situated on top of the main cryogenic stage, it interfaces directly with the upper stage.
The VEB is a big cylindrical ‘basket’ 5.4 m in diameter. It stands 1.56 m tall and weighs 1300 kg without propellant; in the centre is the storable propellant stage (EPS).”
https://www.esa.int/Enabling_Support/Space_Transportation/Launch_vehicles/Vehicle_equipment_bay

 If the ESC-A really has a mass ratio of about 8 to 1 it might work for a half-size Ariane or for the proposed two-Vulcain version of the Ariane 6.

 In my calculations I used the earlier cryogenic stage the H10 from the Ariane 4 because of its small size and high mass ratio:

ARIANE 4 STAGE 3
Specifications are given in H10/H10+/H10-3 order.
Designation: H10/H10+/H10-3
Engine: single cryogenic open cycle SEP HM-7B
Length: 10.73 m/11.05 m/11.05 m
Diameter: 2.60 m
Dry mass: 1,200 kg/1,240 kg/1,240 kg, excluding interstage 2/3
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 10,800 kg/11,140 kg/11,860 kg
Thrust: 63 kN vac/63.2 kN vac/64.8 kN vac
http://www.braeunig.us/space/specs/ariane.htm

   Bob Clark

[Kermann Nolandung]

Well, I was wrong.
Ariane 5 User's manuel gives 4540 kg for the dry mass.

The illustration on this website gives aditional details. It includes 900 kg for the Inter Stage Skirt, but the VEB is excluded.

The VEB changed between A5 G and A5 ECA. Something from 1450 kg to 950 kg when the aluminium structure has been replaced by composite structure.

Acme Engineering (acme-engineering.nl)

[Exoscientist]

10 hours ago, Kermann Nolandung said:

Well, I was wrong.
Ariane 5 User's manuel gives 4540 kg for the dry mass.

The illustration on this website gives aditional details. It includes 900 kg for the Inter Stage Skirt, but the VEB is excluded.

The VEB changed between A5 G and A5 ECA. Something from 1450 kg to 950 kg when the aluminium structure has been replaced by composite structure.

Acme Engineering (acme-engineering.nl)

 Thanks for that. I don’t know why the ESC-A has such a poor mass ratio when the earlier H10 cryogenic upper stage on the Ariane 4 was much better at ca. 10 to 1. Perhaps because the Ariane 5 had to carry much greater payload at 20 tons that the ESC-A needed greater structural strengthening.

  Bob Clark

[tater]

 

[sevenperforce]

On 10/8/2023 at 9:20 AM, Exoscientist said:

 Thanks for that. I don’t know why the ESC-A has such a poor mass ratio when the earlier H10 cryogenic upper stage on the Ariane 4 was much better at ca. 10 to 1. Perhaps because the Ariane 5 had to carry much greater payload at 20 tons that the ESC-A needed greater structural strengthening.

The ESC-A (and ESC-B) stage included a full maneuvering package/system on the stage itself for precision maneuvering and orbital insertion, while this was all located on the VEB ring for the Ariane 4's H10 stage.