totm dec 2023 Artemis Discussion Thread

[Exoscientist]

2 hours ago, tater said:

The mission goal for Artemis is a sustainable, and sustained human presence on the Moon, and specifically in a polar site where water is possibly available. It is not, "replicate Apollo." The mission goals were profoundly different.

Any mission architecture that does not deliver that is wrong from the start. I could see using a frozen LLO for access, but the need for a habitat for at least 2 weeks is part of the equation and over time, I would assume something that can survive night (however many days that is at the chosen landing site given vehicle height, etc).

Someone on Reddit made this observation, “SLS is fat. Orion is not Apollo on steroids. It is an aged overweight Apollo.”

  Robert Clark

[tater]

2 hours ago, Exoscientist said:

Someone on Reddit made this observation, “SLS is fat. Orion is not Apollo on steroids. It is an aged overweight Apollo.”

Yeah, Orion was designed for a different mission—Constellation.

SLS was designed for NO mission.

[tater]

SpaceX wants to have their HLS contract move the ball down the field for SS. They pitched it, NASA accepted. Had they failed, then maybe they would have bid some smaller lander for the next phase, who knows.

3 hours ago, Exoscientist said:

Either the Falcon 9 first stage or second stage could be used for the purpose.

No. They would not be doing that, makes no sense for them, complicates stage 0, etc.

[StrandedonEarth]

20 minutes ago, tater said:

No. They would not be doing that, makes no sense for them, complicates stage 0, etc.

Certainly not worth the hassle. But it might be worth the hassle for a Centaur V…

[tater]

52 minutes ago, StrandedonEarth said:

Certainly not worth the hassle. But it might be worth the hassle for a Centaur V…

But a waste unless even larger than that. Shame to have 9m to play with and put some tiny stage there. A large Centaur probably costs more than the entire SS stack, lol.

[Exoscientist]

 I’m estimating the payload with the Silverbird Astronautics estimator for the SLS 1B if using the more powerful J-2X engine on the Boeing EUS rather than the four RL-10 engines. The original plan was to use the J-2X but it was decided to use the 4 RL-10’s instead because the greater thrust could damage the solar panels on the Orion. 

 But that was for the ~26.5 tons for the Orion+Service Module. I’m envisioning adding an additional 10 tons to the Service Module propellant load and a 13 ton Apollo-sized lander, bringing the total mass sent to TLI in the range of 49.5 tons. Then the higher thrust of the J-2X should not be an issue. 

 Then I want to know what would be the LEO and TLI payload with the more powerful J-2X engine. First, I’m going to check how accurate Silverbird is now for the payload using the 4 RL-10 engines:

 

 Using the data here:

 

 The numbers for the payload fairing field of 8,000 kg and 200 seconds jettison time are taken from the numbers for the launch abort tower.

 The calculated payload to LEO of 109.7 tons is bit higher than the currently quoted value of 105 tons by NASA. But NASA has acknowledged the actual payload value will be somewhat higher than 105 tons. So ~110 tons may actually be an accurate estimate for the SLS 1B payload to LEO.

 Then I’ll calculate the payload using the J-2X:

 This based on the input data here:

 

 Note, I increased the vacuum ISP to 482.5s  A remarkable report suggested a hydrolox upper stage could get a 482.5 s vacuum ISP with a quite high expansion ratio of ~625.

 

The large diameter of the SLS does make that expansion ratio  possible. 

 Note this payload is in the range of SLS Block 2’s 130 tons to LEO without having to use the advanced carbon-fiber upgrades to the boosters.

 In the next post, what’s really important is the payload to TLI. I’ll calculate that next.  

Bob Clark

 

 

Edited December 7, 2023 by Exoscientist

[tater]

1 hour ago, Exoscientist said:

 Note, I increased the vacuum ISP to 482.5s  A remarkable report suggested a hydrolox upper stage could get a 482.5 s vacuum ISP with a quite high expansion ratio of ~625.

That's not a J-2. Developing a new engine puts whatever upper stage you're talking about out at least a decade unless some fast company does it just because.

I'm fine with considering alternate architectures for cislunar missions, but why make up an engine that doesn't exist then put it on a overpriced system that will likely be obviated before the whitepaper level work on the new engine has even been submitted?

[sevenperforce]

18 minutes ago, Exoscientist said:

Note, I increased the vacuum ISP to 482.5s  A remarkable report suggested a hydrolox upper stage could get a 482.5 s vacuum ISP with a quite high expansion ratio of ~625.

This approach makes no sense. If you think that you can simply arbitrarily increase the specific impulse of any hydrolox engine, then why do it with the J-2X but not with the RL-10s? Also, your numbers were wrong to begin with: the RL10C-3 for the EUS develops 460.1 seconds I_sp, not the 448 seconds you gave it.

In reality, you've got it upside down. That particular report is featuring an expander cycle engine, which has an intrinsically high specific impulse because it doesn't burn any of the propellant to run the pumps. The J-2X, like the J-2 before it, is a gas generator engine which (as I've told you before) cannot achieve anything near the maximum attainable specific impulse of an expander cycle.

Comparing the J-2X to the J-2 will illustrate that pretty cleanly. The J-2X has 3.35x the expansion ratio of the J-2 and 1.57x the chamber pressure of the J-2, but only increased the vacuum specific impulse by a paltry 6.4%. Specific impulse is roughly proportional to the square root of (1 - R_P^0.26), where R_P is the pressure ratio between the exit and the chamber, so if you increase the expansion ratio by from 27.5:1 to 92:1, you're going to get a MAXIMUM of 9.4% increase in vacuum specific impulse...obviously the actual number is slightly lower due to built-in inefficiencies and other factors. If you use a massively oversized nozzle to go ahead and increase the expansion ratio from 92:1 to 625:1, you're going to get a probable increase of ~5.8% which is going to get you no higher than 474 seconds.

And that would be for a nozzle diameter 2.61 times greater than the already massive 3-meter nozzle of the J-2X: nearly 8 meters. Adding a nozzle extension onto the J-2 to create the J-2X already increased its mass by 680 kg (actually more, because the other J-2X upgrades decreased the mass of the powerplant and chamber). Approximating the J-2X nozzle extension as a truncated cone with a lower diameter of 3 meters, a height of 1.3 meters, and an upper diameter of 2.1 meters, we get a lateral surface area of 11.02 square meters. Adding an additional nozzle extension to increase the exit diameter to 7.83 meters would require the new extension to be 6.98 meters in length and have a lateral surface area of 125.65 square meters, adding something like 7.98 tonnes.

Which brings us to additional problems. All four RL10C-3 engines together have a mass of 920 kg and occupy 2.17 meters of vertical space with the nozzle extension stowed. The J-2X has a mass of 2.47 tonnes and a fixed length of 4.7 meters, while your super-nozzle-extension J-2 (I'll call it the J-2Z) would have a minimum stowed length of 5.84 meters. So you'd need to remove 3.67 meters of length from the EUS, reducing its propellant capacity by something like 24%. Remember that you can't just make the EUS longer because this single-launch architecture of yours ALREADY doesn't have enough vertical space to fit both a capsule and a sortie lander. Your math didn't account for increased mass OR increased length.

Finally, there were other errors in what you plugged into Silverbird. For example, the dry mass of an SLS booster is 102 tonnes, not 110 tonnes. The dry mass of the SLS core is 85.3 tonnes but you have to add 7.4 tonnes for the interstage (Block 1's LVSA interstage massed 4.5 tonnes with a lateral surface area of 177 square meters so scale up for Block 1B's longer interstage with its cylindrical surface area of 292 square meters). Similar errors have been corrected throughout. 

Here are the correct calculations (spoilered for length). First, we create a baseline using the Silverbird calculator with the existing SLS Block 1B to gauge overestimation:

Spoiler

It gives just under 120 tonnes, which is significantly more than the 105-110 tonnes claimed by NASA. As I've explained before, Silverbird tends to overestimate both T/W ratio and sea level specific impulse when dealing with altitude-compensating sustainer engines like the RS-25. So we know we need to reduce whatever Silverbird tells us by ~11.4%.

Next, let's use the J-2X. Note that I have added the increased mass of the bigger engine but also reduced the length, propellant capacity, and dry mass of the EUS to accommodate the greater length of the engine.

Spoiler

The J-2X is 2.53 meters longer than the stowed RL10C-3s, so I added 2.53 meters to the length of the interstage and thus 1.67 tonnes to the weight of the core. The tank length of the EUS is 15.39 meters so we reduce its propellant capacity by approximately 16.4%. The tank mass of the EUS (without engines) is 13,190 kg, so we reduce this by 16.4% as well and then add in the 2.47 tonnes of the J-2X.

With the overestimation reduction, that's 119.6 tonnes to LEO: certainly an improvement, but not enough to get us the kind of performance we'd need for a single-launch architecture. And now finally let's use your imagined J-2Z, adding 7.98 tonnes for the nozzle extension and reducing the length and propellant capacity of the EUS accordingly:

Spoiler

The J-2Z is 3.67 meters longer than the stowed RL10C-3s, so we add 2.42 tonnes to the original mass of the core stage. We must now reduce the length of the original EUS to 76.2%, giving it a tank dry mass of 10.05 tonnes and a prop load of 96 tonnes. Finally we add the total 10.45 tonnes of J-2Z to get the true dry mass of the shrunken EUS. We do get a slight bump up of vacuum thrust from the longer nozzle.

And applying the same reduction, that gives us 116.3 tonnes to LEO. So this super-high-expansion-ratio idea is sending us in the wrong direction. That makes intuitive sense, after all: if it made sense to get higher specific impulse by attaching a bigger nozzle extension, that's what rocket engineers would do! But it doesn't, because nozzle extensions are really quite heavy, so there's always a balancing test.

3 hours ago, Exoscientist said:

In the next post, what’s really important is the payload to TLI. I’ll calculate that next.

You need around 3.15 km/s for TLI. Assuming a 185x185 km parking orbit, your parking orbit velocity is 7.793 km/s from the vis-a-vis equation. Adding 3.1 km/s at perigee raises your "apogee" to 451,517 km (which makes sense given that the moon is a little under 400,000 km away and you need to swing past it for an efficient lunar orbit insertion).

So if you plug my baseline SLS figures above into Silverbird just as I did before but setting a desired apogee of 451,000 km, you get an estimated TLI payload of 46.7 tonnes, 11.2% more than the 42 tonnes that NASA claims can be sent to TLI by SLS Block 1B Cargo. (Note: this should give us some confidence in our method because the overestimate remains consistent.) Using the J-2X figures from above, we get a TLI payload of 48 tonnes, which comes to 43.2 tonnes when you apply the overestimate. So it's really not a significant increase once you are looking at TLI deliveries. Could it do an Apollo-style mission? Sure: if you gave it the exact same capsule, service module, and lunar module of Apollo. But so could ordinary SLS Block 1B.

[tater]

From reddit:

Interesting visualization (since stages have finite cross-sectional area to put engines on). Shows the trade off of large nozzle engines in terms of thrust per unit area dropping.

 

[tater]

FWIW, it seems like any suggestion of a new frankenrocket—which is fine to do, rockets have had (and still do have) stages/engines by different contractors before—should start with NOT the SLS as the rocket to alter.

SLS is already not cost-effective.

SLS is already not in a position to have schedule margin to alter it—changes that would result in VAB changes, ML changes, etc, into the bargain.

There are up and coming vehicles that would in fact be better candidates for novel upper stage concepts.

Vulcan.

New Glenn.

Starship.

Any of those 3 would be better fodder for some new concepts. Also, as has been said above, perhaps an EOR architecture might be considered, and you build a larger vehicle out of parts launched with existing systems and soon to exist systems (Vulcan/NG, ignoring SS).

[DAL59]

Would it be technically feasible to make a Falcon-Heavy derived HLS? One launch to send a small lander to LEO, a second launch to send a kicker stage to LEO to dock with the lander and take it to NRHO, no superheavy launch vehicles or orbital refueling required. If so, why aren't they doing this if it would be much cheaper and faster than either BM or SHLS?

[SunlitZelkova]

2 hours ago, DAL59 said:

Would it be technically feasible to make a Falcon-Heavy derived HLS? One launch to send a small lander to LEO, a second launch to send a kicker stage to LEO to dock with the lander and take it to NRHO, no superheavy launch vehicles or orbital refueling required. If so, why aren't they doing this if it would be much cheaper and faster than either BM or SHLS?

Because Artemis is going to the Moon to stay. It would be far less capable of wouldn’t be able to deliver the cargo necessary for a lunar base like Starship HLS can.

In addition, Starship HLS, and to a lesser extent, Blue Moon are already under development. Starting on a new lander would probably take so much time it would be easier to continue with what there is now.

[sevenperforce]

4 hours ago, DAL59 said:

Would it be technically feasible to make a Falcon-Heavy derived HLS? One launch to send a small lander to LEO, a second launch to send a kicker stage to LEO to dock with the lander and take it to NRHO, no superheavy launch vehicles or orbital refueling required.

This was the original plan for HLS as an accompaniment to SLS, but then Starship came along and obviated that.

[Exoscientist]

11 hours ago, sevenperforce said:

This approach makes no sense. If you think that you can simply arbitrarily increase the specific impulse of any hydrolox engine, then why do it with the J-2X but not with the RL-10s? Also, your numbers were wrong to begin with: the RL10C-3 for the EUS develops 460.1 seconds I_sp, not the 448 seconds you gave it.

 As shown in the image from the twitter post, for that 482.5 ISP, it is a matter of a nozzle extension. Those are already used on the highest performing RL-10 versions that bring their ISP’s  to the 465.5 range. You can’t get much higher than this on current rockets that use them such as the Delta IV Heavy and Atlas V because the expansion ratio is limited by the rocket diameter.
 Why not just use the ultra large extension nozzles on the RL-10’s on the SLS with its larger diameter? The high payload possible is due also to the high thrust of the J-2X engine. You would need 3 times more of the RL-10’s so twelve of them to match the J-2X. Too much added weight in addition to practicality issues. Remember it was NASA’s original plan to use the J-2X engine. It was decided not to use it when simulations showed its high thrust could damage the Orions solar panels once deployed. But I’m envisioning twice as much payload to TLI so the acceleration would be much reduced.

Edit: the payload would still be quite high if we added a nozzle extension to the J-2X to bring the ISP to the range of 465.5s already reached by the RL-10 with nozzle extensions:

 

 Giving a result:

 

  Still over 130 tons to LEO.

  Robert Clark

Edited December 7, 2023 by Exoscientist

[tater]

8 hours ago, Exoscientist said:

  Still over 130 tons to LEO.

So years of dev added in order to change nothing.

Orion is ~26.5t.

The min Artemis lander per NASA is ~36t (and is 3 stage, so height in single stack very likely a problem). Your stack in LEO with that min lander now has ~2400 m/s of dv. Not enough for TLI. If you want some other SM for Orion... that's more years of dev, also changing the stack. Even if it cuts lander mass, it won't cut the total stack mass enough.

Add 10t to Orion (SM). Literally use an Apollo LM. Total dv in LEO of your upper stage? 3069 m/s. Not enough, unless you shave more mass somehow.

Even if you shave the mass—now, at the cost of billions of dollars, and many years, you have a useless "not Artemis" mission. Why?

Serious question, why mess with SLS at all?

SLS was badly designed. It can never be improved enough to make it useful. The total mass to TLI needs to be on the order of 70+ tons with Orion—and that is bare bones Artemis I level stuff, meeting none of the later requirements.

Edited December 7, 2023 by tater

[sevenperforce]

7 hours ago, Exoscientist said:

the payload would still be quite high if we added a nozzle extension to the J-2X to bring the ISP to the range of 465.5s already reached by the RL-10 with nozzle extensions:

I don't know how else to tell you, man -- if you put garbage in, you're going to get garbage out. If we could get magic, weightless nozzle extensions for free, that don't take up any volume or mass, then rocket engineers everywhere would be adding them to every engine everywhere.

7 hours ago, Exoscientist said:

Still over 130 tons to LEO.

And no, it's still not over 130 tonnes to LEO, because not only are your numbers all wrong, but you're also not applying the overestimation factor.

7 hours ago, Exoscientist said:

for that 482.5 ISP, it is a matter of a nozzle extension. Those are already used on the highest performing RL-10 versions that bring their ISP’s  to the 465.5 range.

It's a matter of adding a nozzle extension to an already intrinsically-high-performing engine cycle. An expander cycle engine is fundamentally different than a gas generator engine.

Look at the RL10A-4-2. It has an expansion ratio of 84:1, lower than the expansion ratio of the J-2X, but it achieves 451 seconds I_sp because it is an intrinsically higher-performing engine. Gas generator engines simply do not have the same efficiency potential as expander cycle engines. The highest-efficiency GG engine flying today, the HM-7B, has an expansion ratio matching the RL10A-4-2 but it only pulls 446 seconds. The record-holding LE-5 gas generator engine had an expansion ratio of 140:1, more than 50% higher than the J-2X, and it was only able to reach 450 seconds I_sp.

And this is before you factor in issues like changes in O:F ratios, chamber pressure, chamber temperature, and so forth. Rocket science is a complex tradeoff balancing many different factors to decrease weight and size while increasing thrust, efficiency, and reliability. You can't just slap an expander-cycle nozzle extension onto a gas generator engine and expect to get expander-cycle efficiency any more than you can slap the straight pipes from a Mustang onto a Prius and expect the Prius to go 0-60 in 4 seconds.

7 hours ago, Exoscientist said:

You can’t get much higher than this on current rockets that use them such as the Delta IV Heavy and Atlas V because the expansion ratio is limited by the rocket diameter.

This just isn't true. The largest-diameter nozzle extension for an RL10 comes in at 2.15 meters, while the interstage of the Delta IV has a diameter of 5 meters. A 4.66-meter nozzle extension for the RL10 (the same size, relative to the Delta IV diameter, as your notional J-2Z) would bring its expansion ratio to 1,315:1. So why not do that? It should be obvious: the efficiency savings wouldn't outweigh the mass bloat.

As I discussed upthread, impulse scales approximately with the square root of (1 - R_P^(k-1)/k), where R_P is the inverse of the expansion ratio and k is a specific heat ratio. Let's pause and take a closer look. If the expansion ratio is infinite, the R_P term drops away, and the specific impulse scales to the square root of 1, which is just . . . 1. That's because every engine has an inherent maximum specific impulse set by the chemistry of the propellants and the thermodynamic constraints of the universe. Increasing the expansion ratio decreases the loss associated with underexpansion at the nozzle outlet, but it can never exceed that theoretical maximum.

However, that equation for impulse is only part of the picture. 

Imagine a rocket engine with a nozzle of zero length. If the expansion ratio is simply 1, then 1 raised to the power of anything is still 1, and so impulse would scale with the square root of (1-1) which is zero. Does a rocket engine with a zero-length nozzle have no thrust? Of course not -- intuitively we know that it would still have thrust: 

The thrust produced by this engine is simple: it's the chamber pressure times the throat area. Of course this is highly inefficient, because all of the heat from the exhaust is being lost to vacuum and all of the gas molecules are flying in every conceivable direction instead of lending their momentum properly. But it's still thrust. If you imagine chopping the nozzle off of an engine (ignoring that many engines would stop working because they use regenerative cooling in their cycle), this is the pressure thrust (F_Pressure) you would get from various engines:

Engine	Chamber Pressure (kPa)	Throat Area (m2)	FPressure (kN)
RL10B-2	4412	0.01297	57.2
RL10A-4-2	4200	0.01280	53.8
RL10C-1	4400	0.01270	55.9
J-2X	9515	0.07931	754.6
J-2	5260	0.10994	578.3
RS-25D	20640	0.05327	1099.4
RS-68A	10260	0.21571	2213.2

Note that because even similar engines have slightly different mass flows, F_Pressure is not going to line up exactly the same. In a pressure-thrust regime, the only way to increase thrust is to either increase chamber pressure or use a wider nozzle. Using a wider nozzle will of course require you to have higher mass flow to maintain chamber pressure. In order to better utilize some of that heat and pressure, we add a nozzle to our engine:

We still have some hot, underexpanded gases escaping around the edges, but now a bunch of our exhaust has cooled and is now pointing downrange, which helps us out tremendously. Now, our total thrust becomes the sum of two components: the original pressure thrust component and a new "momentum thrust" (or "dynamic pressure") component: 

F_Total = F_Pressure + F_Momentum

 

The difficulty is that the contributing proportions of these two components change nonlinearly with chamber pressure, expansion ratio, combustion temperature, mixture ratio, and more -- certainly beyond the scope of this thread. But this should illustrate why you can't just keep imagining a longer and longer nozzle (and apparently a weightless one at that?) and conclude that arbitrarily high specific impulses are possible.

12 hours ago, Exoscientist said:

if we added a nozzle extension to the J-2X

The J-2X already has a nozzle extension. A J-2X with a different nozzle extension would be a different engine.

Even if we imagined that the J-2X was not a gas generator and was actually just a really big uprated RL10, getting up to 465.5 seconds of specific impulse would make it prohibitively heavy.

Since we keep running into the same issues with Silverbird, I went ahead and did a little tweaking. As I said in my last post, setting up the actual values for SLS with a TLI trajectory gives about 11.2% more than what Block 1B can actually send to TLI:

Spoiler

This is primarily from overestimation due to how Silverbird handles altitude-compensating engines like the RS-25.

To fix this, just downgrade the efficiency of the RS-25s to 95.95%, setting total core thrust to 8,747 kN and vacuum specific impulse to 434 seconds:

Spoiler

This allows Silverbird to better characterize the performance of the RS-25s and yields 42.2 tonnes to TLI, right about where you want it to be.

Note that setting this to a 185 x 185 km nominal parking orbit with direct ascent yields 111.1 tonnes to LEO, again right about where NASA says the max is for Block 1B. 

[Minmus Taster]

Orion reentry as seen from onboard the spacecraft:

https://images.nasa.gov/details/art001m1203451716

So THIS is what Jeb goes through when he reenters from Eeloo at twice the solar escape velocity.

[PakledHostage]

Speaking of Orion and Artemis, this popped up in my YouTube feed today. And no matter what one may think about the program, it is cool footage.

Edit: Apologies @Minmus Taster. I thought your link was just an image, but it's actually the source footage for the YouTube clip.

Edited December 13, 2023 by PakledHostage

[cubinator]

1 hour ago, Minmus Taster said:

Orion reentry as seen from onboard the spacecraft:

https://images.nasa.gov/details/art001m1203451716

So THIS is what Jeb goes through when he reenters from Eeloo at twice the solar escape velocity.

I watched the whole thing lamenting that there wasn't any audio included, and then realized my computer speakers were muted. Time to enjoy again!

[AckSed]

Two things I noticed: one, the pieces flying off the heatshield were a bit alarming, but it was an ablative heatshield. Two, a few bits of paper or something clung to the window and seemed to scorch an outline on the window during the first re-entry. It wasn't on the outside as the ocean landing essentially washed the window, and the scorch mark was still there. How hot was that window inside?

[cubinator]

15 minutes ago, AckSed said:

Two things I noticed: one, the pieces flying off the heatshield were a bit alarming, but it was an ablative heatshield. Two, a few bits of paper or something clung to the window and seemed to scorch an outline on the window during the first re-entry. It wasn't on the outside as the ocean landing essentially washed the window, and the scorch mark was still there. How hot was that window inside?

I think the papery stuff could still have been on the outside, as the ocean 'washed' but didn't scrub. Probably bits of ablative material that got sucked back onto the nose end of the spacecraft and toasted as the air has to follow the vehicle to fill the gap left behind it.

This window is on the front facing up toward the nose, and on the inside of the spacecraft things should have been getting pulled toward the other side, away from the window and behind the camera by the g's.

Edited December 13, 2023 by cubinator

[sevenperforce]

13 hours ago, cubinator said:

I watched the whole thing lamenting that there wasn't any audio included, and then realized my computer speakers were muted. Time to enjoy again!

It's so cool to watch with the audio on, because the plasma stream starts to heat up visibly before you get any significant wind noise. It's really something.

All those burning bits from the ablatives are crazy to watch too.

[darthgently]

38 minutes ago, sevenperforce said:

It's so cool to watch with the audio on, because the plasma stream starts to heat up visibly before you get any significant wind noise. It's really something.

All those burning bits from the ablatives are crazy to watch too.

It looked like they may have gotten a good test of cross-range capability with those roll maneuvers also.  Great vid

[PakledHostage]

I went looking for the re-entry path and found these on space.com:

 

 

I was trying to figure out what those snow covered mountains were early in the clip. I thought maybe the South Island of New Zealand, but it doesn't fit. Maybe Antarctica? But that seems too far away from the entry interface? Maybe they weren't mountains?

Edit: I just watched it again on the livingroom TV with my 5 year old son, and we agree that those aren't mountains at the start. And I'm proud to say that my little guy was just as enthralled by the video as I was... But then he's also the kid that, in his kindergarten class, drew a closeup of a rocket nozel with orange exhaust coming out and surrounded by clouds. The teacher didn't have enough technical understanding to realize what he was drawing, but I suspect most here do.

Edited December 16, 2023 by PakledHostage

[Exoscientist]

  Another attempt to get an estimate using more accurate numbers of the payload to LEO of the SLS 1B with the RL-10’s on the Boeing EUS swapped out for the J-2X engine.

 First, note a short-coming of the SilverbirdAstronautic.com is that it overestimates the payload for rockets with large side boosters, such as the Space Shuttle, Ariane 5, and SLS. This is likely due to the fact these large SRB’s greatly throttle down even well before burnout.

 

 So instead of using the vacuum thrust for the SRB’s we’ll input the smaller sea level thrust instead. In regards to the parameters of the SRB, Core stage, and Boeing EUS, we’ll use the data in reports by Boeing engineers:

 

 

  Note a key modification to the input numbers will have to be in the residual fields for each stage. The SilverbirdAstronautics estimator as default takes the residuals as 0.5%. But in the image above with the SLS parameters the residual for the SRB’s amounts to only 0.14% but more importantly the residual for the core is about 1%, and the residual for the ICPS stage is about 3%.

 Boeing EUS residuals aren’t specified but we’ll take it also as about 3% since it also uses the RL-10 and has similar structural form as the ICPS. 

 Note also the mass of the adapter for the ICPS called the LVSA, “launch vehicle stage adapter”, is about 5 tons. We’ ll take the adapter, more accurately, interstage, for the Boeing EUS as also about 5 tons and add this onto the dry mass of the core. Then the input to the estimator is:

 With an estimated payload to LEO as:

 

 The payload value of 105.7 tons is close to the NASA value of 105 tons. 

 In the next post I’ll do the estimate with the four RL-10’s swapped out for the J-2X.

  Robert Clark