Author Topic: Problems of SSTO and technologies to solve these  (Read 124704 times)

Offline AlanSE

  • Full Member
  • *
  • Posts: 153
  • N Cp ln(T)
    • Gravity Balloon Space Habitats Inside Asteroids
  • Liked: 54
  • Likes Given: 33
Re: Problems of SSTO and technologies to solve these
« Reply #40 on: 11/06/2014 03:59 pm »
My favorite "quasi-SSTO" approach would be:

1- Airlaunched "Assisted SSTO"
2- Use gamma maneuver (ie rocket assisted zoom-climb) with crossfeed from carrier aircraft to keep rocket topped until separation. This cuts the dV to orbit down into the ~8000m/s range.
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN to get both the thrust and the T/W ratio at ignition but high Isp in sustainer mode, with only two propellants and a good stage O/F ratio. The higher altitude start decreases the Isp penalty during TAN operations.

Probably with helecopter landing and a retractable engine nozzle.

I think that with LOX/LH2 + LOX-rich TAN, you could probably make a ground takeoff SSTO close.

~Jon

There was some post on this board and a blog post which I can remember making a similar point. I did not understand the case at all.

Particularly, why do you crossfeed the fuel? I mean, what is the design environment which leads to this making sense? As I understand it, this made sense for the shuttle. Why? Because you avoid putting any engines on the external tank. Makes sense. You're throwing it so obviously you don't want to use your mass putting engines on it.

But if we're trying to make reusable rockets, the argument seems to fall apart. Isn't it much easier to transfer thrust than to transfer liquid? If you have 2 stages and you plan on reusing them both, why doesn't the first one just push the 2nd one until separation instead of dealing with the difficulties of crossfeed?

Offline Nilof

  • Full Member
  • ****
  • Posts: 1177
  • Liked: 597
  • Likes Given: 707
Re: Problems of SSTO and technologies to solve these
« Reply #41 on: 11/06/2014 04:48 pm »

If you build the Startram, and basically become THE go-to place for launch (including human missions), how much is inclination an issue? What percentage of current low to mid earth orbits require a flexible inclination?
I'd imagine that almost all LEO and mid orbit industry and tourism would be captured at the megastructure's inclination, so flexibility wouldn't be much of an issue for that? Are missions to GEO, EM1 and lunar also affected substantially by inclination?
Not crash hot on the orbital dynamics - apologies in advance.

Edit: just finished reading the Startram book but inclination not really mentioned as a limiting factor in there except for their proposal for a polar orbit launcher in Antarctica (yikes!)

Doesn't change much, the concept is still viable at any latitude. A near-equatorial startram would mostly be neat since it'd minimize the inclination changes between the various orbits it launches into at different times, so you can launch into the same orbit several times a day and launch windows are much more forgiving. That is generally something that would be nice to have if your business case assumes 10 launches a day.
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline RanulfC

  • Senior Member
  • *****
  • Posts: 4595
  • Heus tu Omnis! Vigilate Hoc!
  • Liked: 900
  • Likes Given: 32
Re: Problems of SSTO and technologies to solve these
« Reply #42 on: 11/06/2014 07:02 pm »
Nuclear rockets are much less good for SSTOs than you might think. Very low Thrust-to-weight basically kills a pure NTR SSTO, and the incredibly low density of hydrogen propellant (though only thing that's really worth putting in a NTR) drives a stake through its heart.

People keep forgeting that Ammonia is a pretty good NTR propellant for Earth-to-Orbit use :) T/W can be improved as well but it remains to be seen if you can get over the MAJOR hurdle of using nuclear propulsion near Earth... The fear :)

LH2 is an all-around better propellant for an NTR but its not the "only" choice and especially not for the Earth-to-Orbit run.

Randy
From The Amazing Catstronaut on the Black Arrow LV:
British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

Offline RanulfC

  • Senior Member
  • *****
  • Posts: 4595
  • Heus tu Omnis! Vigilate Hoc!
  • Liked: 900
  • Likes Given: 32
Re: Problems of SSTO and technologies to solve these
« Reply #43 on: 11/06/2014 07:22 pm »
That's not what McDonnell Douglas said.  From a 1994 press release by McDonnell Douglas on NASA's web site:

I hope you're not suggesting that the press release for a military program could ever be deliberately misleading.

Quote from: ChrisWilson68
They weren't practising to launch expendable missiles quickly.

Where did I say they were? Read what I wrote or, ya know, go read one of the many great books which describe what Pete and the boys were doing.

You said:
"Nope. They were using a reusable vehicle to prove crews could launch missiles more responsively. The missiles weren't intended to "land" anywhere."

"Missiles" would in this case seem to mean "expendable-weapons" which if you DIDN'T mean that you should have been clearer :)

In reality those "books" clearly state that they WERE showing the operational possiblities of REUSABLE vertical take-off-and-landing rocket powered vehicles. They were NOT in fact doing anything that would have shown how launching "missiles" would be more or less "operationally" capable in fact using a LH2/LOX vehicle to do so would make no sense militarily. (No current operational "missiles" {other than subsonic "cruise" missiles} use liquid fuels and there is a very good and sound OPERATIONAL reason for that)

The miltary has no need to demonstrate "operationally responsive" missile launch, we HAVE that capability as a basis of our nuclear deterent force and have had since the late 1960s. Operationally responsive "space launch" is another thing all together and doing so would NOT require landing tests. It in fact hasn't since the idea has been studied. For military use expendable launch vehicles have always made more "sense" but DARPA (and SDIO) have been interested in reusable launch vehicles almost exclusivly BECAUSE they are reusable which fits the specific mission "profile" they favor.

There's other good reasons to aim for SSTO or even "assisted" SSTO. Especially as you start thinking about going beyond LEO.

I've heard the arguments but specifically which ones are you suggsting and given the low-payload-to-orbit of such how does it compare to a reusable TSTO "upper-stage" that is capable of being re-fueled on-orbit?

Randy
From The Amazing Catstronaut on the Black Arrow LV:
British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

Offline RanulfC

  • Senior Member
  • *****
  • Posts: 4595
  • Heus tu Omnis! Vigilate Hoc!
  • Liked: 900
  • Likes Given: 32
Re: Problems of SSTO and technologies to solve these
« Reply #44 on: 11/06/2014 07:25 pm »
My favorite "quasi-SSTO" approach would be:

1- Airlaunched "Assisted SSTO"
2- Use gamma maneuver (ie rocket assisted zoom-climb) with crossfeed from carrier aircraft to keep rocket topped until separation. This cuts the dV to orbit down into the ~8000m/s range.
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN to get both the thrust and the T/W ratio at ignition but high Isp in sustainer mode, with only two propellants and a good stage O/F ratio. The higher altitude start decreases the Isp penalty during TAN operations.

Pretty much the "Crossbow" concept IIRC?
Quote
Probably with helicopter landing and a retractable engine nozzle.

Man after my own heart :) Maybe plug, or cluster-plug nozzle?
Quote
I think that with LOX/LH2 + LOX-rich TAN, you could probably make a ground takeoff SSTO close.

"Close" maybe but wouldn't an assist still make it more "economical" as in giving a higher payload? Your problem is still in your margins allowable versus your payload size to orbit.

Randy
From The Amazing Catstronaut on the Black Arrow LV:
British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

Offline rusty

  • Full Member
  • *
  • Posts: 191
  • Liked: 14
  • Likes Given: 18
Re: Problems of SSTO and technologies to solve these
« Reply #45 on: 11/07/2014 07:06 am »
...
That's why people believe in SSTO - the numbers look to be just on the edge of the possible.
...
It's not that SSTO is undoable.  It's that it's SO marginal that we haven't got to where we've managed to make it work even as a demonstrator, never mind as an operational reusable vehicle.

Being very light and being very structurally robust at the same time is hard.
Good post and points out SSTO is simply an engineering issue, not pie in the sky. Many concepts get lost in the numbers and forget the design must be a whole. Structurally, being a whole is called "fusion" where the tanks and TPS are partially load-bearing.
Numbers can suggest the high isp of hydrogen (Skylon), but as a whole, denser and non-cryogenic propellants win out. The simplicity and perceived weight savings of VTVL may stand out, but as was mentioned in the OP, HTHL wins out. And of course there's the pure engineering advances like one all-encompassing engine to keep mass low and aerodynamic tricks for low-speed through hypersonic efficiency that also provide structural fusion of the wings (as in the SR-71).

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2430
  • Likes Given: 13606
Re: Problems of SSTO and technologies to solve these
« Reply #46 on: 11/07/2014 09:26 am »
Good post and points out SSTO is simply an engineering issue, not pie in the sky. Many concepts get lost in the numbers and forget the design must be a whole. Structurally, being a whole is called "fusion" where the tanks and TPS are partially load-bearing.
The fact that it has not been done in 60+ years (except on the Moon  :) ) suggests there is a little more to it that.
Quote
Numbers can suggest the high isp of hydrogen (Skylon), but as a whole, denser and non-cryogenic propellants win out. The simplicity and perceived weight savings of VTVL may stand out, but as was mentioned in the OP, HTHL wins out. And of course there's the pure engineering advances like one all-encompassing engine to keep mass low and aerodynamic tricks for low-speed through hypersonic efficiency that also provide structural fusion of the wings (as in the SR-71).
Actually SABRE is such an all-encompassing engine, not just up to M3 but M23.

[EDIT The clever part about SR71 aerodynamics is in fact the creation of the chines on the side of the fuselage. Like the carefully shaped wing of the Concorde (A feature Tupulev failed to spot from the data stolen by the KGB, which is why the Tu144 had canards) these provided critical low speed lift that allowed the design to be landed relatively easily. I'll also note that Concorde's aerodynamics allowed it to fly in "supercruise" for over 30 years without people making any drama about it. ]
« Last Edit: 11/08/2014 08:58 am by john smith 19 »
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline Oli

  • Senior Member
  • *****
  • Posts: 2467
  • Liked: 605
  • Likes Given: 60
Re: Problems of SSTO and technologies to solve these
« Reply #47 on: 11/07/2014 09:38 am »
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN...

Expander? You'd need a lot of engines...

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2430
  • Likes Given: 13606
Re: Problems of SSTO and technologies to solve these
« Reply #48 on: 11/08/2014 09:05 am »
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN...

Expander? You'd need a lot of engines...
IIRC P&W were  were looking at the RL60 (60Klbs?) and up to 100 Klbs (Aerojet?).

The size issues are to do with the surface area to volume ratio of the chamber as the SA limits the extractable head you have to drive the turbopumps unless you a) have multiple chambers driven by a single set of turbo machinery Russian style or b) stick an extra HX in the chamber (Aerojet proposal).

The real limit seems to be getting chamber pressures in the stage combustion class (2000psi +) is quite difficult at large scale.

Why you would feel you would need such a high chamber pressure (other than to show you can do it) is another matter.  :(

MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline Proponent

  • Senior Member
  • *****
  • Posts: 7277
  • Liked: 2782
  • Likes Given: 1462
Re: Problems of SSTO and technologies to solve these
« Reply #49 on: 11/08/2014 09:18 am »
It seems to me that high chamber pressure leads to high thrust density which leads to low pressure losses for a conventional nozzle.  If you can easily compensate for pressure losses by other means (TAN, aerospike, etc.), then chamber pressure probably isn't so important.

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2430
  • Likes Given: 13606
Re: Problems of SSTO and technologies to solve these
« Reply #50 on: 11/08/2014 07:44 pm »
It seems to me that high chamber pressure leads to high thrust density which leads to low pressure losses for a conventional nozzle.  If you can easily compensate for pressure losses by other means (TAN, aerospike, etc.), then chamber pressure probably isn't so important.
That's the major issue. Note that historically the SSME was going to be a two position nozzle with an extension that would lock into place later in the flight.

Note that the J-2X (original in the late 60's) was working on improving the flow separation at around 1500psi.

A dual bell nozzle is an option.  The work from REL suggests people should also be re visiting the E/D nozzle.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline nadreck

Re: Problems of SSTO and technologies to solve these
« Reply #51 on: 11/08/2014 07:53 pm »
As has been pointed out above, most SSTO projects do not stack up against TSTO based on how much vehicle is required given the same technology of SSTO vs TSTO.

So what would change that picture?  Well changing the ISP is the first thing that comes to mind, no mater what we do it is the place where you get the biggest bang for your buck. For the sake of argument imagine a mythical system where you had an engine TW ratio of 50 and a non cryogenic fuel with an ISP of 500 with a fuel density of 1 (water).  Suddenly an SSTO, winged, engineered for 5 flights between overhauls is an easy possibility. The one drawback is there is no good chemical way to do this. NTR, well that blows the TWR ratio away though that ISP is possible with water, and of course the radiation is a no go. FTR (Fusion), ok, postulate that we have B11 fusion working what are the limits of this?  At 100% efficiency to produce an ISP of 500 you would need 50KW per kilogram of thrust which means at an engine 10:1 TW ratio you need an energy density of 500KW per kilo of engine.  Imagining a 100,000 kilo GTW winged SSTO with say 10,000 kilos thrust  you have an engine weighing in at 1000kilos, running at least 500MW, 87,500kg water for reaction mass, and payload and structure (including wings) at 11,250 kg.  What is really neat with this idea is that if want an isp of 5000, you get it in theory by reducing thrust in a square ratio so at 100kg of thrust you now have engines with the ISP needed to open up the solar system (though why you would want to carry the wings anywhere . . .).

Until we can do some sort of high energy density FTR SSTO doesn't make sense. And to me winged upper stages don't either, however shaping upper stages for better control and some cross range ability in re-entry does.

Using the proposed focus fusion device we are looking at an energy density of  about 2KW per kilo, similar numbers for the Lougheed one, this makes low thrust deep space engines possible at high ISP, it does not help with earth launch.

TSTO though gives you the opportunity to optimize each stage for its operating environment and I suggest, as a good approximation, going for a winged first stage that is capable of mach 3 briefly (using its own jet engines and the upper stage's engines running on cross fed fuel) you could separate the upper stage at about 20km altitude having gone from level subsonic flight at 15,000 meters to mach 3 and eventually a 30 degree 'gone balistic' maneuver. If we imagine the SSTO weighing in at 100,000kg and running hydrolox the fuel mass fraction would have to be 82,000kg, we could get by with an engine that has 120,000kg thrust. Given that the engine does not need to be perfectly efficient at 15,000 meters, and at 20,000 meters and above it will perform much like it does in vacuum it can be mostly tuned for vacuum and lets say weigh in at just less than 1000kg. That still leaves 17,000kg for structure, controls and payload. 

Mechanical and aerodynamic issues of the above concept would be based on having the upper stage mated to a shortened and far forward central fuselage of the first stage. Regular jet engines on the first stages wings that probably attach to twin secondary fuselages running the length of the upper stage and connecting for fuel and oxidizer at the rear of that stage.  The upper stage needs a volume around 500m3 in tankage so lets say a cylinder with a diameter of about 4 meters and a length for fuel tank portion of about 35m and maybe 50m overall. I would suggest that the upper stage engines have to shut down before separation and the lower stage must be able to nose down and pull away before re-ignition. The upper stage continues with the 30degree angle until it can be reduced. Max acceleration if there were no throttling would be just over 6G.  Could something this sized still have a useful payload after allowing for making it reusable?  Probably, but not a heck of a lot easier than an F9 upper stage. What about it's economics as an expendable? There it might be better. Given the much smaller volume required for a Kerolox upper stage or a Methalox one, those might be worth examining for the reusable case, however the carrier aircraft would have to handle a higher weight.
It is all well and good to quote those things that made it past your confirmation bias that other people wrote, but this is a discussion board damnit! Let us know what you think! And why!

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2430
  • Likes Given: 13606
Re: Problems of SSTO and technologies to solve these
« Reply #52 on: 11/08/2014 11:36 pm »
<long post full of speculation snipped.
You're correct that nuclear systems have very poor T/W ratio's and solid core systems cost a lot of money to give basically a x2 increase in readily achievable Isp.

Once you know that an HTOL system needs a thrust of 1/3 GTOW to get it's payload off the ground and a VTOL needs (at least) 1.2x GTOW wings start to look pretty good if you are OK with needing a runway.

The big one turns out to be air breathing. This trades engine T/W  for greatly expanded reaction mass.

In the case of SABRE the T/W is around 14:1 (excellent by turbo fan standards but very bad by rocket standards) however during air breathing it hits around 3-4000 secs Isp.

While that's only up to about M5.5 the effect on the average Isp for the flight is substantial, as it would be for any air breathing system.

The trouble is it needs some kind of trajectory analysis program (or a spreadsheet) to give you a real feel for what's going on as things like air density, pressure and temperature are all constantly changing during the ascent (which is sort of the point)

MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline Jim Davis

  • Full Member
  • ****
  • Posts: 560
  • Liked: 124
  • Likes Given: 2
Re: Problems of SSTO and technologies to solve these
« Reply #53 on: 11/08/2014 11:55 pm »
While that's only up to about M5.5 the effect on the average Isp for the flight is substantial, as it would be for any air breathing system.

It's substantial but it tends to be overstated. For example, for a vehicle operating at 4000s over M=0-5 and 400s over M=5-25 the effective Isp over M=0-25 is 488s. A substantial improvement as you say but it is not obvious that the improvement in Isp makes up for the lower engine T/W, lower propellant density, and the rest of the air breathing requirements.

Offline nadreck

Re: Problems of SSTO and technologies to solve these
« Reply #54 on: 11/09/2014 12:24 am »
While that's only up to about M5.5 the effect on the average Isp for the flight is substantial, as it would be for any air breathing system.

It's substantial but it tends to be overstated. For example, for a vehicle operating at 4000s over M=0-5 and 400s over M=5-25 the effective Isp over M=0-25 is 488s. A substantial improvement as you say but it is not obvious that the improvement in Isp makes up for the lower engine T/W, lower propellant density, and the rest of the air breathing requirements.

and carrying the wings from M5.5 to M25, then the TPS for the wings and the jet engines coming back down, for the benefit of being able to fly around a bit at both ends. I am sorry unless we have high energy density FTR on the order of 500KW per kilo of engine, SSTO will not make sense when compared to the same technologies in TSTO.
It is all well and good to quote those things that made it past your confirmation bias that other people wrote, but this is a discussion board damnit! Let us know what you think! And why!

Offline john smith 19

  • Senior Member
  • *****
  • Posts: 10351
  • Everyplaceelse
  • Liked: 2430
  • Likes Given: 13606
Re: Problems of SSTO and technologies to solve these
« Reply #55 on: 11/09/2014 12:35 am »
It's substantial but it tends to be overstated. For example, for a vehicle operating at 4000s over M=0-5 and 400s over M=5-25 the effective Isp over M=0-25 is 488s. A substantial improvement as you say but it is not obvious that the improvement in Isp makes up for the lower engine T/W, lower propellant density, and the rest of the air breathing requirements.
Then you're missing the biggest factor in the

The air.

If you haven't done a trajectory simulation you're very off in your estimates. One of the Skylon design team commented that air breathing adds about 220 tonnes of reaction mass to their vehicle.
Without air breathing that would have to be included on the vehicle, as would it's tankage, and the loads it would impose on the structure during ascent. That's why simple minded substitutions of SABRE for, say the SSME don't work.

So an engine nacelle weighing c5 tonnes "produces" maybe 110 tonnes of reaction mass and oxidizer literally out of thin air.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline jongoff

  • Recovering Rocket Plumber/Space Entrepreneur
  • Senior Member
  • *****
  • Posts: 6807
  • Lafayette/Broomfield, CO
  • Liked: 3987
  • Likes Given: 1681
Re: Problems of SSTO and technologies to solve these
« Reply #56 on: 11/09/2014 02:36 am »
My favorite "quasi-SSTO" approach would be:

1- Airlaunched "Assisted SSTO"
2- Use gamma maneuver (ie rocket assisted zoom-climb) with crossfeed from carrier aircraft to keep rocket topped until separation. This cuts the dV to orbit down into the ~8000m/s range.
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN to get both the thrust and the T/W ratio at ignition but high Isp in sustainer mode, with only two propellants and a good stage O/F ratio. The higher altitude start decreases the Isp penalty during TAN operations.

Probably with helecopter landing and a retractable engine nozzle.

I think that with LOX/LH2 + LOX-rich TAN, you could probably make a ground takeoff SSTO close.

~Jon

There was some post on this board and a blog post which I can remember making a similar point. I did not understand the case at all.

Particularly, why do you crossfeed the fuel? I mean, what is the design environment which leads to this making sense? As I understand it, this made sense for the shuttle. Why? Because you avoid putting any engines on the external tank. Makes sense. You're throwing it so obviously you don't want to use your mass putting engines on it.

But if we're trying to make reusable rockets, the argument seems to fall apart. Isn't it much easier to transfer thrust than to transfer liquid? If you have 2 stages and you plan on reusing them both, why doesn't the first one just push the 2nd one until separation instead of dealing with the difficulties of crossfeed?

Crossfeed from the aircraft is little different from a T-0 disconnect at the launch pad. It's just a way of making sure you keep the rocket tanked up until you've checked out the engines, and give the aircraft enough thrust to pitch up into the right flight path angle. If you didn't, you'd get a lot less delta-V out of that first 8-12 seconds of propellant. Crossfeed doesn't have to be that complicated.

~Jon

Offline jongoff

  • Recovering Rocket Plumber/Space Entrepreneur
  • Senior Member
  • *****
  • Posts: 6807
  • Lafayette/Broomfield, CO
  • Liked: 3987
  • Likes Given: 1681
Re: Problems of SSTO and technologies to solve these
« Reply #57 on: 11/09/2014 02:38 am »
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN...

Expander? You'd need a lot of engines...

Depends entirely on the scale of the vehicle. I'm thinking relatively smallish. Plus, expander w/ TAN could still likely get you up into the 100-300klbf/engine range. Also, if you're doing TAN, an expander bleed cycle could work too, with the "bleed" GH2 being injected into the TAN system.

~Jon
« Last Edit: 11/09/2014 02:40 am by jongoff »

Offline jongoff

  • Recovering Rocket Plumber/Space Entrepreneur
  • Senior Member
  • *****
  • Posts: 6807
  • Lafayette/Broomfield, CO
  • Liked: 3987
  • Likes Given: 1681
Re: Problems of SSTO and technologies to solve these
« Reply #58 on: 11/09/2014 02:44 am »
My favorite "quasi-SSTO" approach would be:

1- Airlaunched "Assisted SSTO"
2- Use gamma maneuver (ie rocket assisted zoom-climb) with crossfeed from carrier aircraft to keep rocket topped until separation. This cuts the dV to orbit down into the ~8000m/s range.
3- Use LOX/LH2 expander cycle engines with LOX-rich TAN to get both the thrust and the T/W ratio at ignition but high Isp in sustainer mode, with only two propellants and a good stage O/F ratio. The higher altitude start decreases the Isp penalty during TAN operations.

Pretty much the "Crossbow" concept IIRC?

Similar. Crossbow was definitely one of the inspirations for this concept. Though I'd prefer to find a way to make the design work without needing to roll your own custom carrier plane.

Quote
Quote
Probably with helicopter landing and a retractable engine nozzle.

Man after my own heart :) Maybe plug, or cluster-plug nozzle?

Maybe. I've always thought aerospikes were oversold (says the guy with TAN on the brain).

Quote
Quote
I think that with LOX/LH2 + LOX-rich TAN, you could probably make a ground takeoff SSTO close.

"Close" maybe but wouldn't an assist still make it more "economical" as in giving a higher payload? Your problem is still in your margins allowable versus your payload size to orbit.

You run into scalability issues with the air-launched concept. And the aircraft assist is a non-trivial part of your operating costs. I like airlaunched SSTO, but would also love to see a "Roton done right" fly some day.

~Jon

Offline 93143

  • Senior Member
  • *****
  • Posts: 3054
  • Liked: 312
  • Likes Given: 1
Re: Problems of SSTO and technologies to solve these
« Reply #59 on: 11/09/2014 07:08 am »
While that's only up to about M5.5 the effect on the average Isp for the flight is substantial, as it would be for any air breathing system.
It's substantial but it tends to be overstated. For example, for a vehicle operating at 4000s over M=0-5 and 400s over M=5-25 the effective Isp over M=0-25 is 488s. A substantial improvement as you say but it is not obvious that the improvement in Isp makes up for the lower engine T/W, lower propellant density, and the rest of the air breathing requirements.

Two things.  First, you've neglected the losses.  The airbreathing trajectory from zero to Mach 5 soaks up virtually all the drag and gravity losses; I calculate that Skylon C1 actually went through about 3.7 km/s in airbreathing mode, and I expect D1 is similar.  Once in rocket mode, only a couple hundred m/s of losses remain, for a total of about 6.3 km/s.  So it's not a delta-V ratio of 1:4 - it's more like 4:7.

Second, you've used a very low rocket Isp, at least for a hydrogen engine.  If you'd used a reasonable value, you'd have come up with a trajectory average comfortably over 500 seconds even with your oversimplified analysis.

Using the SABRE's predicted rocket Isp (459 s) and my delta-V numbers above, along with john_smith_19's 4000 s ballpark airbreathing number, I get an average Isp of 683 s.  Using Skylon D1.5's actual projected mass ratio of 4.406 with my calculated delta-V from the C1 spreadsheet of 9,991 m/s, I get 687 s.

...

Also, oddly enough the low propellant density doesn't seem to be an issue for Skylon.  During the D1 redesign, sizing the vehicle by payload bay diameter and shaping it for minimum drag produced a design too long for the tanks...

and carrying the wings from M5.5 to M25, then the TPS for the wings and the jet engines coming back down, for the benefit of being able to fly around a bit at both ends. I am sorry unless we have high energy density FTR on the order of 500KW per kilo of engine, SSTO will not make sense when compared to the same technologies in TSTO.

Have you seen the numbers for Skylon?  It manages nearly 5% payload fraction - better than any expendable TSTO - while carrying industry-standard mass growth allowances.  Even the dry payload fraction is similar to that of the Delta IV.

As far as I can tell, turning Skylon into a TSTO would make a mess of it.  You'd ruin the elegance of the structural and aerodynamic design, ending up with a bunch of compromises and duplication of functionality.  It would have extra maneuvers to do and extra flight regimes to deal with, the first stage would have to handle roughly twice the heat soak, and the whole breaking in half at speed and reassembling on the ground thing couldn't help but put a dent in the system's reliability, never mind maintenance costs.  Plus if you wanted a GTO mission, you'd probably need a third stage...

Staging is only good because it's so much harder to reach orbit without it.  Everything else about it is worse.  Given a technology that can do orbit in one leap without having to try ridiculously hard for a ridiculously small payload, why would you bother designing two vehicles when you could do the job with one?

Granted, Skylon is not a known quantity quite yet, but you seem to have already decided what the answer is without much real analysis, never mind enough real world operational experience with various types of RLVs for hindsight to kick in...
« Last Edit: 11/09/2014 07:49 am by 93143 »

Tags:
 

Advertisement NovaTech
Advertisement Northrop Grumman
Advertisement
Advertisement Margaritaville Beach Resort South Padre Island
Advertisement Brady Kenniston
Advertisement NextSpaceflight
Advertisement Nathan Barker Photography
1