SSTO challenge

[Cinder]

Skylon?

Seconded.

I know it's bad for an engineer to get on a bandwagon and stop thinking, but sometimes I'm amazed at how much effort it takes to get people to even acknowledge the Skylon concept's existence in an SSTO thread...

In fairness the OP specified rocket power.  I mentioned it because all things considered, getting to orbit cheaply is what really matters. And less how you do it.  Also because I'm curious what other credible scheme(s), paper or not, that compares.

[93143]

In fairness the OP specified rocket power.

No it didn't.  All it said was single-stage-to-orbit reusable launch vehicle, possibly with nuclear thermal rocket propulsion, but with air launch excluded.

Skylon isn't air launched.  It's airbreathing, which isn't mentioned anywhere in the original post, either yea or nay.  Considering that NTRs capable of doing SSTO tend to be at a similar TRL to SABRE, it seems to me that your suggestion was a valid answer to the thread challenge (except for the part about flying several times per day).  It's not BOTE from REL's perspective, but from ours it is...

[Cinder]

Sorry, I thought RLV means Rocket Launch Vehicle.  R for reusable, not rocket.  Actually I already knew that, but somehow still read it as "rocket LV" half the time.

[KelvinZero]

Skylon sounds really interesting but it is UK based isnt it?

Any hope that all this new money that is meant to go into research will lead to more money for Reaction Engines to prove their idea somehow?

[meiza]

I'll play, using existing engines and flown tank mass ratios (the Lego method):

SSTO, VTVL, truncated cone.

Mass budget in metric tons:

GLOW:  185
IMLEO:  14.7

Propellant: 171
  LOX:  125
  RP-1:  40
  LH2:     6

Tankage:  2.5
  LOX:   1.3
  RP-1:  0.5
  LH2:   0.7

Good concept with the LEGO method! Nice high tankage ratios. About 100 for LOX and RP and 10 for LH2.

[kkattula]

Good concept with the LEGO method! Nice high tankage ratios. About 100 for LOX and RP and 10 for LH2.

96, 80 and 8.6. All have been bettered in existing flight vehicles.

Could have gone with higher performance engines too, but the ones I selected could do the job and appear robust enough for multiple flights.

I used a spread sheet to optimize the kerolox/hydrolox ratio for delta v, based on tank & engine mass.

Interestingly, the pure RP-1 case requires RD-180 like Isp, and loses so much performance it would not be reusable. The pure LH2 would also not be reusable. But a kerolox/hydrolox ratio of anywhere from 2.5 to 10 works, with a maximum at about 4.2 for my assumptions.

You could probably get a small payload increase by carrying the RP-1 in drop tanks, but it probably wouldn't be enough for even one extra person in a manned vehicle.

[Downix]

Good concept with the LEGO method! Nice high tankage ratios. About 100 for LOX and RP and 10 for LH2.

96, 80 and 8.6. All have been bettered in existing flight vehicles.

Could have gone with higher performance engines too, but the ones I selected could do the job and appear robust enough for multiple flights.

I used a spread sheet to optimize the kerolox/hydrolox ratio for delta v, based on tank & engine mass.

Interestingly, the pure RP-1 case requires RD-180 like Isp, and loses so much performance it would not be reusable. The pure LH2 would also not be reusable. But a kerolox/hydrolox ratio of anywhere from 2.5 to 10 works, with a maximum at about 4.2 for my assumptions.

You could probably get a small payload increase by carrying the RP-1 in drop tanks, but it probably wouldn't be enough for even one extra person in a manned vehicle.

a TAN approach, perhaps, or something which utilizes air for part of the flight, such as a scramjet?

[tnphysics]

What about an RBCC?

Use air (compressed by a compressor) to replace the LOX oxidizer. Same thrust chamber and fuel pump. Use a heat exchanger in the inlet, so that the excess heat (from the shock-heated air) can be transferred to the fuel just prior to injection.

Note that the power turbine (which drives the pump) would be the SAME for both configurations. It would be an oxidizer-rich staged combustion cycle; O₂ and N₂ have close enough molecular masses that combustion gasses from either air or LOX and fuel could drive the turbine.

[tnphysics]

I'll play, using existing engines and flown tank mass ratios (the Lego method):

SSTO, VTVL, truncated cone.
4 x Merlin 1c
4 x RL-10-A-2

Mass budget in metric tons:

GLOW:  185
IMLEO:  14.7

Propellant: 171
  LOX:  125
  RP-1:  40
  LH2:     6

Engines (inc plumbing and thrust structures)
4 x Merlin 1c:    4.5
4 x RL-10-A-2:  1.5

Tankage:  2.5
  LOX:   1.3
  RP-1:  0.5
  LH2:   0.7

TPS: 0.5
(Bolt on PICA replaced after each flight)

Cabin: 1
Other: 1  (avionics, ECLS, batteries, RCS, etc)

Leaves 500 kg for crew/passengers.

The only armwavium is an assumption of double bell nozzles on the RL-10s so they can be used at sea level for landing. Propellant includes 1 t of LOX/LH2 for landing. Boil-off GOX/GH2 used for RCS.

Launches using the Merlins, which provide 3600 m/s dv, then switches to the RL-10s which provide another 5840 m/s.

Further details left as an exercise to the reader...

Have you considered methane?

The RL-10 engines can use it without modifications. They could probably switch from LCH₄ to LH₂ mid-flight (would require shutdown and restart)!! The engines could run on a mixture of the two fuels, provided cross-plumbing (cheap and easy to develop) was provided. Each engine's pump can only handle one fuel at a time, as methane freezes at LH₂ temps, but some could pump one while the others pumped the others. The extra plumbing would allow the thrust chambers to be fed a mixture.

Also, running on LCH₄ probably gives a thrust boost, assuming constant pump power.

Finally, assume one uses an RD-191 (instead of the Merlins). What is the payload?

[kkattula]

What about an RBCC?

Use air (compressed by a compressor) to replace the LOX oxidizer. Same thrust chamber and fuel pump. Use a heat exchanger in the inlet, so that the excess heat (from the shock-heated air) can be transferred to the fuel just prior to injection.

Note that the power turbine (which drives the pump) would be the SAME for both configurations. It would be an oxidizer-rich staged combustion cycle; O₂ and N₂ have close enough molecular masses that combustion gasses from either air or LOX and fuel could drive the turbine.

If you wanted to use air as the initial oxidixer you need much bigger chambers to get the same thrust, because there's only 20% of the oxygen to burn with the fuel. You get some benefit from the N2 reaction mass, but not enough. That's why rocket engines have far higher T/W ratio than jet engines.

I wouldn't worry about a heat exchanger, just stay sub-sonic until the air start to thin out at about 15 km.

Ducting the airflow into the rocket exhaust for extra thrust is probably more effective than trying to burn it.

[kkattula]

RL-10s are too small. you's need 25 of them. And their T/W is only about half that of a Merlin. So engine mass is almost doubled.

Methane is not dense enough for a first stage. With an O/F of about 1, you've increased the tankage mass by about 50%.

Congratulations, you now have negative payload as an SSTO, let alone RLV.


RD-191 is a litlle heavy, but with the higher Isp would provide a small increase in payload. Maybe enough for a third person.

[tnphysics]

What about NK-33? or plug-nozzle version of said engine?

Also, what about a higher performance LH₂/LOX engine? Use SSME T/W; 480sec vacuum I_sp.

As for methane, do not forget the thrust bonus when using it. This is because it is volume flow that limits thrust. For LOX/methane, stoichiometric, one gets a thrust bonus of 44% if the thrust chamber is the limiting factor, or 64% if one is limited by the turbopumps. These numbers are over stoichiometric LH₂/LOX, which itself has a 13% thrust bonus over the typical fuel-rich mixture ratio of 6 (if one is pump-limited). Use stoichiometric O/F ratios in the sustainer engine.

Don't forget that LOX/methane can reach 435sec vacuum I_sp!! This is for SSME- or NK-33- caliber designs.

[Downix]

What about NK-33? or plug-nozzle version of said engine?

Also, what about a higher performance LH₂/LOX engine? Use SSME T/W; 480sec vacuum I_sp.

As for methane, do not forget the thrust bonus when using it. This is because it is volume flow that limits thrust. For LOX/methane, stoichiometric, one gets a thrust bonus of 44% if the thrust chamber is the limiting factor, or 64% if one is limited by the turbopumps. These numbers are over stoichiometric LH₂/LOX, which itself has a thrust bonus over the typical fuel-rich mixture ratios. Use stoichiometric O/F ratios in the sustainer engine.

Don't forget, there was a proposal for an SSME based SSTO proposed by Rockwell for the X-33 program. 
http://www.astronautix.com/graphics/x/x33rock.jpg

[tnphysics]

In retrospect, it appears that I should have allowed air-launch, provided that it is done in an inexpensive way. In other words, no major ground equipment or crew for the assembly (too expensive).

One inexpensive method would be to tow the RLV up to altitude with a cable, using a cargo jet that happens to be taking off from the same location at the same time. This requires a winged vehicle, but could save sizable amounts of delta-V.

[Rabidpanda]

Also, what about a higher performance LH₂/LOX engine? Use SSME T/W; 480sec vacuum I_sp.

Are you sure that's possible?  I thought the maximum Isp for hydrolox was around 469 seconds.

Don't forget that LOX/methane can reach 435sec vacuum I_sp!! This is for SSME- or NK-33- caliber designs.

According to these two sources lox/methane can't go any higher than around 386 seconds.

http://en.wikipedia.org/wiki/Liquid_rocket_propellants

http://www.dunnspace.com/alternate_ssto_propellants.htm

[Patchouli]

RL-10s are too small. you's need 25 of them. And their T/W is only about half that of a Merlin. So engine mass is almost doubled.

Methane is not dense enough for a first stage. With an O/F of about 1, you've increased the tankage mass by about 50%.

Congratulations, you now have negative payload as an SSTO, let alone RLV.


RD-191 is a litlle heavy, but with the higher Isp would provide a small increase in payload. Maybe enough for a third person.

Replace the RL-10s with RL-60s redesigned to have TAN using hydrocarbon fuel spiked with boron.

The turbopump gas generator never sees the hydrocarbon or the boron so no coking issues.

TAN is not a free lunch as you'll need a heavier nozzle and a third pump if going with tripropellant but it's still far less mass then having Merlins or an RD-191 for take off and them be dead weight for 60% of the flight.

I'm not 100% sure on this but I think with TAN the nozzle could be over sized for non TAN operation and act like a high expansion nozzle in conventional mode.

Better yet just reuse the RL-60 turbo pump hardware with a plug nozzle engine not sure you can do TAN with it but the plug nozzle compensates for altitude.
Plus you get a free heat shield if there is left over hydrogen to cool the plug nozzle.

[mlorrey]

What about an RBCC?

Use air (compressed by a compressor) to replace the LOX oxidizer. Same thrust chamber and fuel pump. Use a heat exchanger in the inlet, so that the excess heat (from the shock-heated air) can be transferred to the fuel just prior to injection.

Note that the power turbine (which drives the pump) would be the SAME for both configurations. It would be an oxidizer-rich staged combustion cycle; O₂ and N₂ have close enough molecular masses that combustion gasses from either air or LOX and fuel could drive the turbine.

If you wanted to use air as the initial oxidixer you need much bigger chambers to get the same thrust, because there's only 20% of the oxygen to burn with the fuel. You get some benefit from the N2 reaction mass, but not enough. That's why rocket engines have far higher T/W ratio than jet engines.

I wouldn't worry about a heat exchanger, just stay sub-sonic until the air start to thin out at about 15 km.

Ducting the airflow into the rocket exhaust for extra thrust is probably more effective than trying to burn it.

I agree with the last statement. Use the GNOM concept of an ejector cowling around the rocket engine. The increase in average Isp would put it far ahead of LH2.

As for T/W, the ejector will provide ramjet-like T/W which approaches that of an LH2 rocket, about 35-40 vs 50 for LH2. Only turbine engines have terrible T/W.

[tnphysics]

Be sure to run very fuel rich if using air augmentation. The air burns your fuel, much less LOX needed.

Also, I would spike the LH₂ with boron and aluminum. This increases I_sp and thrust. The Al is to form particles of Al₂O₃, which act as condensation nuclei, reducing the problems that the B₂O₃ particles create.

One problem with this last approach is that the turbine may need be either hot enough that B₂O₃ vaporizes (over 2100K) or cool enough that it solidifies to prevent particle deposition on the turbine blades (bad).

Also, I would use a tripropellant (only 2 pumps-start with LOX in pump 1 and RP-1 in pump 2, then switch to LH2 in pump 1 and LOX in pump 2, thus making good use of both pumps regardless of which propellant is being used).

[mlorrey]

Be sure to run very fuel rich if using air augmentation. The air burns your fuel, much less LOX needed.

Also, I would spike the LH₂ with boron and aluminum. This increases I_sp and thrust. The Al is to form particles of Al₂O₃, which act as condensation nuclei, reducing the problems that the B₂O₃ particles create.

One problem with this last approach is that the turbine may need be either hot enough that B₂O₃ vaporizes (over 2100K) or cool enough that it solidifies to prevent particle deposition on the turbine blades (bad).

Also, I would use a tripropellant (only 2 pumps-start with LOX in pump 1 and RP-1 in pump 2, then switch to LH2 in pump 1 and LOX in pump 2, thus making good use of both pumps regardless of which propellant is being used).

I'm quite aware of boron doped fuels, I've written a few articles on it, as well as the requirements of air augmented rockets.

GNOM would have exhibited a trajectory performance equivalent to a US missile that was double its mass. A scaled prototype bore this out as well.

An experiment I'd like to try is a Falcon 1 first stage with an ejector as a SSTO ELV. Eliminating the cost of the second stage, and the increased ground operations costs that go with it, should be rather significant, and should allow a significantly larger payload than F1 currently is capable of.

Any SpaceX people listening?

[hop]

In retrospect, it appears that I should have allowed air-launch, provided that it is done in an inexpensive way. In other words, no major ground equipment or crew for the assembly (too expensive).

So how about launching from a scaled up Armadillo / Masten style VVTL ?  It appears that such a vehicle could meet your requirements, and be operated for a modest multiple of consumable cost with less ground crew than many aircraft.

Of course, like air launch, it would actually be a TSTO...