Project Nova HLV

[K_Jameson]

Perhaps the classic approach of dividing the spacecraft into modules to reduce the reentry volume, as in the Soyuz or the Dark Energy add-on, would have been preferable because it would have made the spacecraft more easily reconfigurable, for example with different service modules (also reducing the problem of the heat shield weight). But we have already worked on this in the past, with the Antares capsule. This time I strongly wanted a fully reusable vehicle, albeit with a more reasonable size than Starship.

 

[Urwumpe]

Perhaps the classic approach of dividing the spacecraft into modules to reduce the reentry volume, as in the Soyuz or the Dark Energy add-on, would have been preferable because it would have made the spacecraft more easily reconfigurable, for example with different service modules (also reducing the problem of the heat shield weight). But we have already worked on this in the past, with the Antares capsule. This time I strongly wanted a fully reusable vehicle, albeit with a more reasonable size than Starship.

I think modularity and a (mostly) reuseable approach are not automatically opposites. Since you want a larger volume with low mass during reentry, you can still use the payload and maybe the intertank volume for installing modules and maybe even expendable modules, that are not returned to Earth.

Also, you could make use of docked spacecraft modules, like for example a special propulsion or habitation module, that does also not break the fully reusable scheme.

 

[K_Jameson]

Back to the heat shield, I don't really understand how Starship handles the heat of a BLEO reentry. Since my vehicle is modeled after Starship, I pay special attention to it (even though I'm not a 100% fan of the concept). But the reentry profile still escapes me. I don't think the vehicle has enough fuel to slow down before atmospheric entry. My Argo ship surely doesn't!

 

[Urwumpe]

Back to the heat shield, I don't really understand how Starship handles the heat of a BLEO reentry. Since my vehicle is modeled after Starship, I pay special attention to it (even though I'm not a 100% fan of the concept). But the reentry profile still escapes me. I don't think the vehicle has enough fuel to slow down before atmospheric entry. My Argo ship surely doesn't!

I also don't think that a purely propulsive slow down before reentry is really an effective way of reducing heat shield mass. While SpaceX might use lunar refueling eventually to slow down by about 2000 m/s, it would either mean a long stay on the lunar surface for the first missions or early missions launch with heavier heatshield and only partial refueling.

I could imagine a sequence of aerobraking and perigee corrections before final reentry takes place from LEO, this could also permit landing near the launch site for better reusablility. This also consumes only a little bit of fuel, but means returning to Earth could take 2-3 days longer, depending on the maximum heat load.

 

[K_Jameson]

One problem I see with a repeated aerobraking sequence is that the vehicle probably has to cross the Van Allen belts multiple times in a relatively short time. This is concerning

 

[Urwumpe]

One problem I see with a repeated aerobraking sequence is that the vehicle probably has to cross the Van Allen belts multiple times in a relatively short time. This is concerning

Might still be better than burning up. More important would be the time that it spends there. Any apogee in the more active regions of the belts would be bad, too much time. Maybe I can program a small "dosimeter" MFD next week to test how much radiation has to be expected.

 

[K_Jameson]

This would be extremely interesting!

 

[K_Jameson]

I am currently looking at the dry weights of all stages and vehicles. I'm considering saving some weight by switching from aluminum to carbon fiber for the interstages. This would also have a beneficial effect on heat management on reentry, for the LEO stack (Nova IB). In the lunar stack (Nova VB), on the other hand, the interstage will be lost along with the expendable core stage, which has been deliberately kept as low-tech as possible so far. So I wonder how much this component would cost, if it were made of carbon fiber.

Also, to fight the increasing weight of the Argo spacecraft, I'm wondering about replacing the tank domes, which are actually aluminum, with carbon fiber ones (the sidewalls, which are exposed to the external environment during reentry, are stainless steel). I read that you need some epoxy resin if you want to join carbon fiber and steel, to avoid galvanic corrosion.

 

[K_Jameson]

I found several good references to confirm my previous calculated weights. In stark contrast, I found that the SLS core stage looks ridiculously overweight for its size. And I wonder why.

 

[Urwumpe]

I found several good references to confirm my previous calculated weights. In stark contrast, I found that the SLS core stage looks ridiculously overweight for its size. And I wonder why.

Likely because it is a) a very conservative design based on Space Shuttle era tooling and b) has a pretty massive engine section that mounts a lot of more subsystems as needed for other directly expendable launchers.

 

[DaveS]

Likely because it is a) a very conservative design based on Space Shuttle era tooling and b) has a pretty massive engine section that mounts a lot of more subsystems as needed for other directly expendable launchers.

Yes, it's actually based on the older Lightweight External Tank (LWT) which was used from STS-7 through STS-107. The LWT was some 7,000 lbs (3,178 kg) heavier than the Super Lightweight External Tank (SLWT) that replaced it (STS-91 through STS-135). It does have some Delta IV stuff thrown in, mainly avionics. It also uses modified Auxiliary Power Units (APUs) to generate the hydraulic pressure required to move the main engines from the shuttle orbiters as well. But instead of running on anhydrous hydrazine like they during the shuttle days, they run on tapped off gaseous hydrogen from the Core Stage hydrogen tank. And the Core Stage Main Engines themselves are modified stock RS-25 Block II's from the shuttle days with each engine having a rich flight history. The new build RS-25E's were flight certified just a year ago.

 

[K_Jameson]

The official weight I found everywhere is 85,275 kg. This, after some rough calculations, seems quite reasonable.

BUT

in the 2022 SLS reference guide, page 8, the weight is stated at 85.3 tons without the engines. This would bring the total dry weight to the ludicrous value of around 97.7 tons! Even with all your fair considerations, this figure still seems like nonsense to me.

 

[Urwumpe]

in the 2022 SLS reference guide, page 8, the weight is stated at 85.3 tons without the engines. This would bring the total dry weight to the ludicrous value of around 97.7 tons! Even with all your fair considerations, this figure still seems like nonsense to me.

10% structural mass is not too bad for a first stage of the 1960s. Of course, today, expectations were a lot different.

 

[K_Jameson]

Do you think the stage could have been lighter without being too expensive? (well, costs got out of control anyway...)
Also I wonder if for the first flight the stage carried test equipment that can be removed for the next launches - admitting there will be next launches...

 

[Urwumpe]

Do you think the stage could have been lighter without being too expensive? (well, costs got out of control anyway...)

Yes, absolutely. Even a 97.7 tons of gold could be cheaper than the SLS core stage. The gold costs just $8.5 billion. In 2021, the estimated costs of one SLS Core Stage exceeded 2.2 billions, assuming enough are build....

I think by using designed-to-be-expendable RS-68 engines, the SLS Core Stage could have been made much lighter and simpler in first place. The decision to use RS-25 engines there was very unfortunate and mostly driven by political pressures.

 

[DaveS]

I think by using designed-to-be-expendable RS-68 engines, the SLS Core Stage could have been made much lighter and simpler in first place. The decision to use RS-25 engines there was very unfortunate and mostly driven by political pressures.

Not quite. The decision to use the RS-25 is rooted in sound engineering rationale from the Ares V days, when they wanted to use RS-68's. But they quickly found out that the ablative nozzle used by the RS-68 would not survive the high base heating conditions created by the SRBs (just look at severe charring of the aft LH2 tank dome of the STS External Tank, that's trapped heat from the SRBs).

So, they would have to come up with brand new "RS-68B" design that incorporated the same regenerative nozzle cooling that the RS-25 already used. And the Air Force who had paid for much of the RS-68 R&D costs didn't want NASA mucking about with "their" engine which was used on critical NSSL launches by the Delta IV. So, the RS-68 for NASA CxP died right then and there, and the switch to the RS-25 was done.

 

[K_Jameson]

RS-68 is much heavier than RS-25 and has much lower performance, apart from raw thrust; unless you use a significantly stretched core stage, it cannot work for SLS.

 

[Urwumpe]

So, they would have to come up with brand new "RS-68B" design that incorporated the same regenerative nozzle cooling that the RS-25 already used. And the Air Force who had paid for much of the RS-68 R&D costs didn't want NASA mucking about with "their" engine which was used on critical NSSL launches by the Delta IV. So, the RS-68 for NASA CxP died right then and there, and the switch to the RS-25 was done.

As I said: Politics.

Also the Ares V back then used a smaller stage diameter than SLS today, which is also a way to reduce the heat load.

 

[K_Jameson]

OK, I have now pretty solid weight figures for my stages. My previous estimates were more or less correct. The problem remains the Argo spacecraft, that grew heavier. In order to assure the required performances without redesigning the stages, I'm forced to swich to subcooled propellants; until now I was hesitant to do so because I was unable to find exact density values, but I'm working on it...

 

[K_Jameson]

OK, found these reference values.

densified LH2: 0.0753 g/cm3 (+ 6%)
densified LCH4: 0.439 g/cm3 (+ 3.6%)
densified LOX: 1.250 g/cm3 (+ 9.5%)

I find them quite good. Can anyone confirm them?