Whats wrong with Skylon?

[blowfish]

17 minutes ago, sevenperforce said:

SABRE depends on slowing down the airflow to allow turbocompression even at prohibitively high speeds where ram compression (while generally less efficient) would be more realizable and result in less drag. But if you can put the turbocompressor in the sidewalls then you can have both. An additional supersonic-flow scramjet/scramrocket mode would be nice, I suppose, but it's not necessary; the gains are rather low.

I think you're still confused.  The only reason why slowing the air down results in thrust loss is because of total pressure loss (mostly in the intake).  Even at mach 5, it's not terrible - the only reason turbojets can't operate at such speeds is because the compressor would overheat (and the turbine would also need to be able to withstand higher temperatures).

The transition to scramjet mode would happen at a single instant, not instantaneously (either the terminal shock exists or it doesn't).  Which is fine, because scramjets aren't really effective below mach 5 anyway.  The engine core would transition from air-breathing to closed cycle mode and the bypass jets would transition to scramjet mode with the intake geometry adjusting as necessary to accommodate this and the airflow to the (now inactive) compressor closed off.

47 minutes ago, sevenperforce said:

I'm not averse to the thought of a lightweight sliding cover to seal off the precooler inlet, but if bleed air (which already would be used to run the compressor) can be used to form a smooth layer without moving parts, that ought to be simpler.

It's difficult to use bleed air for this purpose with supersonic flow.  The flow you're bleeding in really needs to be at the same speed as the same speed as the main flow, but then you're pushing it through the precooler too and you get the same pressure loss problem.

In the SABRE at least, the compressor is actually run by the helium loop.  I think this makes sense for a number of reasons, including reducing long-term wear on the turbines due to high temperature gas flowing through them.

51 minutes ago, sevenperforce said:

What do you think, generally, about having a cylindrical turbocompressor mounted into the sidewalls of the intake, to allow mechanical compression of the airstream at subsonic and low-supersonic speeds without impeding ram compression at mid-to-high-supersonic speeds?

I honestly don't see the advantage of having the bypass down the center rather than along the perimeter.  For a shock cone intake, the intake throat is a ring near the edge of the nacelle anyway, so the path between that and the bypass is basically straight.

[sevenperforce]

29 minutes ago, blowfish said:

I think you're still confused.  The only reason why slowing the air down results in thrust loss is because of total pressure loss (mostly in the intake).  Even at mach 5, it's not terrible - the only reason turbojets can't operate at such speeds is because the compressor would overheat (and the turbine would also need to be able to withstand higher temperatures).

I honestly don't see the advantage of having the bypass down the center rather than along the perimeter.  For a shock cone intake, the intake throat is a ring near the edge of the nacelle anyway, so the path between that and the bypass is basically straight.

The purpose of having a central bypass rather than an annular bypass is to keep the compressor out of the flow path. That way it can still massively augment thrust at low speeds, but does not produce drag at high speeds.

I also suspect that its maximum operating speed will be slightly higher, since it is pulling in air perpendicular to the airflow vector rather than bearing the full brunt of stagnation pressure.

37 minutes ago, blowfish said:

The transition to scramjet mode would happen at a single instant, not instantaneously (either the terminal shock exists or it doesn't).  Which is fine, because scramjets aren't really effective below mach 5 anyway.  The engine core would transition from air-breathing to closed cycle mode and the bypass jets would transition to scramjet mode with the intake geometry adjusting as necessary to accommodate this and the airflow to the (now inactive) compressor closed off.

On the topic of (sc)ramjets vs (sc)ramrockets...

We have a few choices for a partially airbreathing rocket engine capable of vacuum operation. You can put the fuel, LOX, and compressed-air inputs directly into the combustion chamber, as SABRE has done. Alternatively, you can put the fuel and LOX inputs into the combustion chamber and put the compressed-air input in the exhaust bell. The former choice has the attractive option of switching between pure airbreather and pure rocket, which seems like a major advantage. However, since it's possible to run the combustion extremely fuel-rich, comparable performance should be attainable with the latter choice, particularly because the use of pure LOX will result in a more efficient reaction than if your airbreathing mode is damping combustion with nitrogen. The latter choice also allows a much higher air flow than the former, which should significantly offset any of the former's advantages.

So there needn't be bypass jets, just a bypass flow path opening into the exhaust bell.

 

[mikegarrison]

I don't understand your infatuation with the idea that "the flowpath" must go right down the middle of the engine. Directing it into an annular path around the core engine is not a big source of losses.

Over and over I keep getting the same impression from you -- you are suboptimizing on specific parts of the engine and ignoring the whole picture.

Edited March 8, 2016 by mikegarrison

[sevenperforce]

23 minutes ago, mikegarrison said:

I don't understand your infatuation with the idea that "the flowpath" must go right down the middle of the engine. Directing it into an annular path around the core engine is not a big source of losses.

Over and over I keep getting the same impression from you -- you are suboptimizing on specific parts of the engine and ignoring the whole picture.

Eh, just trying to solve the problem of being able to use ram compression and turbocompression in the same engine.

[AngelLestat]

23 hours ago, magnemoe said:

Why? Yes Mach 5 is probably a bit slow, however you already has to load the fragile satellite, second stage will be larger but lighter then empty and designed to fitt

They need 25 mach to reach orbit speed, so they achieve 20% of the speed at 28km altitude which is most of the atmosphere drag.
It is a substantial help for a second stage, but...  the development cost and mostly the operational cost of a 2 stage launcher is higher than 1 stage to orbit.
Imagine a design in which a first reusable stage vehicle can reach 5 MACH  with something else attached to it, no way...  The image that was post it few pages back would not work.
You can make it as a rocket in front, but when this detach it will change all the aerodynamic and mass equilibrium of each of its parts.

[blowfish]

41 minutes ago, sevenperforce said:

The purpose of having a central bypass rather than an annular bypass is to keep the compressor out of the flow path. That way it can still massively augment thrust at low speeds, but does not produce drag at high speeds.

I also suspect that its maximum operating speed will be slightly higher, since it is pulling in air perpendicular to the airflow vector rather than bearing the full brunt of stagnation pressure.

On the topic of (sc)ramjets vs (sc)ramrockets...

We have a few choices for a partially airbreathing rocket engine capable of vacuum operation. You can put the fuel, LOX, and compressed-air inputs directly into the combustion chamber, as SABRE has done. Alternatively, you can put the fuel and LOX inputs into the combustion chamber and put the compressed-air input in the exhaust bell. The former choice has the attractive option of switching between pure airbreather and pure rocket, which seems like a major advantage. However, since it's possible to run the combustion extremely fuel-rich, comparable performance should be attainable with the latter choice, particularly because the use of pure LOX will result in a more efficient reaction than if your airbreathing mode is damping combustion with nitrogen. The latter choice also allows a much higher air flow than the former, which should significantly offset any of the former's advantages.

So there needn't be bypass jets, just a bypass flow path opening into the exhaust bell.

 

The SABRE's operating speed is limited by compressor heating, not airflow, or TPR losses, or anything else.

Again, airflow is mostly limited by the intake.  The nozzle throat should be designed to admit the same amount of air as the intake at the design point.  If off-design conditions become an issue, you can either direct more air to the bypass or have a variable-sized nozzle throat (this is actually pretty easy on the expansion-deflection type nozzles planned for the SABRE - the nozzle centerbody just has to move forward and backward).  Mixing when the flow is already supersonic isn't really going to do much.

You seem to be under the impression that turborockets can be more efficient than other air-breathing engine types.  Didn't we disprove this based on the performance of real engines?

[mikegarrison]

31 minutes ago, sevenperforce said:

Eh, just trying to solve the problem of being able to use ram compression and turbocompression in the same engine.

A problem that was addressed pretty successfully 50 years ago in the SR-71. Granted, only to Mach 3.2ish, but that's where I would start working from.

Edited March 8, 2016 by mikegarrison

[sevenperforce]

2 hours ago, blowfish said:

The SABRE's operating speed is limited by compressor heating, not airflow, or TPR losses, or anything else.

If the SABRE's operating speed is limited by compressor heating, then it would benefit by gradually shifting compression off of the turbocompressor and onto a ram compressor. 

But it can't do that, because it can only accept air in the combustor (spill ramjets aside).

2 hours ago, blowfish said:

Airflow is mostly limited by the intake.  The nozzle throat should be designed to admit the same amount of air as the intake at the design point.  If off-design conditions become an issue, you can either direct more air to the bypass or have a variable-sized nozzle throat (this is actually pretty easy on the expansion-deflection type nozzles planned for the SABRE - the nozzle centerbody just has to move forward and backward).  Mixing when the flow is already supersonic isn't really going to do much.

You seem to be under the impression that turborockets can be more efficient than other air-breathing engine types.  Didn't we disprove this based on the performance of real engines?

Well, "real engines" is a bit of a tricky consideration; there haven't been any airbreathing SSTO engines of any kind ever successfully built. But that depends on what constitutes a turborocket. A rocket engine which runs fuel-rich and accepts compressed air solely in the nozzle may well be more efficient than an engine which does not use its own oxidizer and accepts compressed air solely in the combustion chamber, simply because the former can accept a greater air/propellant ratio.

The central bypass approach seemed to be the best way to get "high-bypass turbofan" performance. You can have a large intake because you can have a really, really ridiculously large effective nozzle. 

2 hours ago, mikegarrison said:

A problem that was addressed pretty successfully 50 years ago in the SR-71. Granted, only to Mach 3.2ish, but that's where I would start working from.

In a craft which was too heavy to take off under its own weight unless its fuel tanks were all but emptied. 

[mikegarrison]

1 minute ago, sevenperforce said:

In a craft [SR-71] which was too heavy to take off under its own weight unless its fuel tanks were all but emptied. 

That's got nothing to do with the engine technology, other than that pure jets (or rockets) are lousy at providing thrust at low speed.

It seems like where you should start is by asking what non-optimal choices they had to make because of the available technology of the day, and then whether current technology could find a better solution. Also, what solutions that worked at Mach 3+ would not work at Mach 6+. Etc.

[sevenperforce]

Just now, mikegarrison said:

That's got nothing to do with the engine technology, other than that pure jets (or rockets) are lousy at providing thrust at low speed.

It seems like where you should start is by asking what non-optimal choices they had to make because of the available technology of the day, and then whether current technology could find a better solution. Also, what solutions that worked at Mach 3+ would not work at Mach 6+. Etc.

Well, the goal is markedly different. Their engine wasn't intended to operate in a vacuum. The beautiful engine on the sr-71 answered a different question: how fast can an air breather go?

That isn't the same question that SABRE is trying to answer. SABRE is asking: how long can an orbit-capable airbreathing rocket use a turbocompressor?

The question I'm asking is different still: how much of an orbital rocket's propellant mass fraction can be air? Because that is the question that will get you to SSTO.

[blowfish]

23 minutes ago, sevenperforce said:

If the SABRE's operating speed is limited by compressor heating, then it would benefit by gradually shifting compression off of the turbocompressor and onto a ram compressor. 

Ram compression happens regardless.  It's not something you can control.  You could possibly reduce turbocompression while beginning to inject LOX into the combustion chamber to keep thrust up ... maybe it could gain you another half a mach or so.

35 minutes ago, sevenperforce said:

Well, "real engines" is a bit of a tricky consideration; there haven't been any airbreathing SSTO engines of any kind ever successfully built. But that depends on what constitutes a turborocket. A rocket engine which runs fuel-rich and accepts compressed air solely in the nozzle may well be more efficient than an engine which does not use its own oxidizer and accepts compressed air solely in the combustion chamber, simply because the former can accept a greater air/propellant ratio.

Sorry, but did you read anything I said?  Nothing about this engine cycle makes sense.  There's no reason to inject into the nozzle.  The fuel/air ratio, even if you burn all the amospheric oxygen, is only a few % with hydrogen.  Giving yourself more airflow without burning its oxygen works fine at low speeds, but it ceases to be an effective strategy once you get supersonic - turbofans loose thrust a lot faster than turbojets, even with the same core.

40 minutes ago, sevenperforce said:

The central bypass approach seemed to be the best way to get "high-bypass turbofan" performance. You can have a large intake because you can have a really, really ridiculously large effective nozzle. 

I still don't see the benefit.  Remember that in a shock cone intake, the capture area is not the area between the cone and the lip, but, at least at the design condition, is the entire frontal area of the intake up to the lip (by design condition I mean that the shockwave intercepts the lip).  You're not getting any more air with your geometry.

42 minutes ago, sevenperforce said:

In a craft which was too heavy to take off under its own weight unless its fuel tanks were all but emptied. 

It's a design tradeoff that the Skylon needs to make too.  The Skylon only needs to be able to operate from specialized runways, so it can accept a 4 km takeoff roll and high (~160 m/s) takeoff speed.  These sorts of compromises don't really make sense for the SR-71.

[mikegarrison]

1 hour ago, blowfish said:

It's a design tradeoff that the Skylon needs to make too.  The Skylon only needs to be able to operate from specialized runways, so it can accept a 4 km takeoff roll and high (~160 m/s) takeoff speed.  These sorts of compromises don't really make sense for the SR-71.

Reminds me of the A340-600, for which one of my colleagues once commented, "it only has a measurable rate of climb because the Earth is curved."

[magnemoe]

12 hours ago, AngelLestat said:

They need 25 mach to reach orbit speed, so they achieve 20% of the speed at 28km altitude which is most of the atmosphere drag.
It is a substantial help for a second stage, but...  the development cost and mostly the operational cost of a 2 stage launcher is higher than 1 stage to orbit.
Imagine a design in which a first reusable stage vehicle can reach 5 MACH  with something else attached to it, no way...  The image that was post it few pages back would not work.
You can make it as a rocket in front, but when this detach it will change all the aerodynamic and mass equilibrium of each of its parts.

You will have to go a bit faster than mach 5 I think as you will do an suborbital trajectory, separation will be in freefall and vacuum yes it will have to be done faster than on space shuttle as you have an time constrain, cold gass trusters rails out of cargo hold then continuing a bit out from spaceplane until you are safe to start its engine. 
Note that Skylon will need an second stage for everything not into LEO anyway. Yes it might use an reuseable tug to GTO but this does not work for other orbits who is not common. 
Agree that doing an hypersonic separation is too hard. 

[AngelLestat]

6 hours ago, magnemoe said:

You will have to go a bit faster than mach 5 I think as you will do an suborbital trajectory, separation will be in freefall and vacuum yes it will have to be done faster than on space shuttle as you have an time constrain, cold gass trusters rails out of cargo hold then continuing a bit out from spaceplane until you are safe to start its engine. 
Note that Skylon will need an second stage for everything not into LEO anyway. Yes it might use an reuseable tug to GTO but this does not work for other orbits who is not common. 
Agree that doing an hypersonic separation is too hard. 

No sure what you want to said.. you mean mach 5 or mach 25?  Take into account that you can even attach a small and expandable h2 tank (cryoballoon maybe) to reach the equator or the best location to launch in case you want to maximize your payload to geo or other special orbits.
No sure if they said you need 25 mach more of speed after you achieve 5 mach or you just need 25 mach in total..  maybe the first one.

It would no need the tug for different kind of low orbits, because it will depend on the payload mass, less mass it means you have extra fuel to spent in space without the need of a tug, but the tug is a great idea because it needs much less propellent to achieve an orbit and go back instead push the whole skylon and then deorbit.
It also can use a ion tug to place many different small sats in their orbits, but with this tug the skylon would not wait until it complete the task, it will go back to earth, land, then many days after launch other sats and recover the ion tug. 

But I want to remark that the biggest issue of a 2 stage vehicle fully reusable (using sable in the first stage) is aerodynamics, take a look to skylon:
  

 

You think someone can achieve a similar drag coefficient maintaining aerodynamics and reusability of each stage the same as a whole?
Also the benefit of skylon size is re entry heat, in the reentry is emptly and light, the overall density is very low, so it reach only few hundreds of degrees vs 1200c of the space shuttle, this mean a lot of weight save in thermal insulation (going up and down).

I will bet all my reputation that we will never see a 2 stage fully reusable vehicle using sable in the first stage, it does not have sense by these and many other reasons.

Edited March 9, 2016 by AngelLestat

[magnemoe]

25 minutes ago, AngelLestat said:

No sure what you want to said.. you mean mach 5 or mach 25?  Take into account that you can even attach a small and expandable h2 tank (cryoballoon maybe) to reach the equator or the best location to launch in case you want to maximize your payload to geo or other special orbits.
No sure if they said you need 25 mach more of speed after you achieve 5 mach or you just need 25 mach in total..  maybe the first one.

It would no need the tug for different kind of low orbits, because it will depend on the payload mass, less mass it means you have extra fuel to spent in space without the need of a tug, but the tug is a great idea because it needs much less propellent to achieve an orbit and go back instead push the whole skylon and then deorbit.
It also can use a ion tug to place many different small sats in their orbits, but with this tug the skylon would not wait until it complete the task, it will go back to earth, land, then many days after launch other sats and recover the ion tug. 

But I want to remark that the biggest issue of a 2 stage vehicle fully reusable (using sable in the first stage) is aerodynamics, take a look to skylon:

You think someone can achieve a similar drag coefficient maintaining aerodynamics and reusability of each stage the same as a whole?
Also the benefit of skylon size is re entry heat, in the reentry is emptly and light, the overall density is very low, so it reach only few hundreds of degrees vs 1200c of the space shuttle, this mean a lot of weight save in thermal insulation (going up and down).

I will bet all my reputation that we will never see a 2 stage fully reusable vehicle using sable in the first stage, it does not have sense by these and many other reasons.

Cargo and second stage would be in the cargo hold just like in Skylon until separation, this is an important point as it eliminate the need of protection or aerodynamic for second stage. Cargo hold would be larger but this will not affect the overall shape except perhaps making the cylindrical part a bit longer. 
You might need some mechanical system to bring it outside the hold to make this faster and safer. 
Speed of first stage would be as high as plausible say march 5 or up past 10 dependent of mass factions and then reentry heat start becoming an problem. 
Second stage would have to do the rest of the work, benefit is that separation speed is faster and higher than most 2 stage rockets. 

Benefit of an suborbital spaceplane is that you don't have the SSTO strict limits you can manage with an heavier plane and lower performance. 
Say you do mach 13, this is halfway to orbital speed but mote than enough to reach space, you will here only need an small upper stage. 

[AngelLestat]

mmm, I did not think about that possibility, but we are talking of a second stage with wings?  is reusable?  How it will land again?
No sure about how a mechanism of big doors and deployment + heavy heat shields and extra operational cost can have more sense than 1 stage to orbit.
Because I dont really find more hard steps in the skylon design than in this 2 stages launcher.

You can find a design that might work, but it really worth it? 

[sevenperforce]

If you're going to use an airbreather as your first stage then eschew LOX altogether and just have it accelerate in-atmo on a trajectory that will take it above the Karman line. Open internal bay, release vacuum-optimized second stage+payload, and then re-enter. 

Should be able to beat the Falcon on a very narrow range of flight profiles. 

In-air refueling is by far the simplest way to get an SSTO spaceplane though. 

[magnemoe]

11 hours ago, AngelLestat said:

mmm, I did not think about that possibility, but we are talking of a second stage with wings?  is reusable?  How it will land again?
No sure about how a mechanism of big doors and deployment + heavy heat shields and extra operational cost can have more sense than 1 stage to orbit.
Because I dont really find more hard steps in the skylon design than in this 2 stages launcher.

You can find a design that might work, but it really worth it? 

Upper stage could either be lifting body or use the planned falcon 9 upper body recovery method where it has heat shield in front, having an disposable fuel tank between second stage and cargo is another option. This would let you do an air capture of stage as it would be light, this might be relevant for GTO missions unless you use an tug, 
For deep space you will use an disposable stage unless you use an 3rd stage. 

The problem with an SSTO is that it has to have an mass faction to reach orbit, while being able to land again, any cargo is an bonus. Skylon get the first part cheap because of air breathing, the rest of the trip it has to carry its heavy aircraft body. 
Skylon is an realistic design for an SSTO but its still hard, with the typical airframe weight increase its unlikely to bring much cargo to orbit. 
Note that even the SSTO version would be able to bring far more cargo going suborbital with second stage. Can do some calculations on this later. 

[fredinno]

15 hours ago, AngelLestat said:

mmm, I did not think about that possibility, but we are talking of a second stage with wings?  is reusable?  How it will land again?
No sure about how a mechanism of big doors and deployment + heavy heat shields and extra operational cost can have more sense than 1 stage to orbit.
Because I dont really find more hard steps in the skylon design than in this 2 stages launcher.

You can find a design that might work, but it really worth it? 

Here. The 1st and 2nd stage would be smaller though due to the better ISP of air-breathing.

9 hours ago, sevenperforce said:

If you're going to use an airbreather as your first stage then eschew LOX altogether and just have it accelerate in-atmo on a trajectory that will take it above the Karman line. Open internal bay, release vacuum-optimized second stage+payload, and then re-enter. 

Should be able to beat the Falcon on a very narrow range of flight profiles. 

In-air refueling is by far the simplest way to get an SSTO spaceplane though. 

In-air refueling at Mach 5 sounds really difficult

[magnemoe]

2 hours ago, fredinno said:

Here. The 1st and 2nd stage would be smaller though due to the better ISP of air-breathing.

In-air refueling at Mach 5 sounds really difficult

Yes, refueling is always subsonic, main benefit of refueling is that you don't need engines and landing gear able to take of with an fully loaded plane. 
You can also cruise from base to the launch area and top of the tanks before the orbital burn. 

[sevenperforce]

2 hours ago, fredinno said:

In-air refueling at Mach 5 sounds really difficult

Yeah, it's not supersonic. Basically the same concept as the SR-71: takeoff with empty tanks, refuel at modest speed, then burn to orbit from a much more optimal launch location. 

It's basically the same as air-launching, but you can use existing tanker aircraft rather than designing an entirely new mothership or dealing with decoupling. 

[magnemoe]

Did an few calculations on Skylon, it can reach 1700 m/s air breathing, isp of engines in rocket mode is 465s and will mostly run in vacuum.
Empty weight is  fully loaded is 325 ton, empty is 53 ton, payload is 15 ton, 
This gives ln(325/68)=1.56
1.56*465*9.8=7109, now add 1700 m/s and you start at the rockets at 30 km attitude, 

Now lets try an two stage version, assuming an total upper stage mass of 110 ton, its the same as falcon 9 upper stage with payload. 
ln(325/(53+110)=0.7, 0.7*465*9.8=3190, we end up 4km/s short with an maximum speed of 5900 m/s 

Assuming we use hydrogen for second stage to and an isp of 450, we will need an mass fuel faction of around 2.6 for upper stage. (ln(2.6)*450*9.8=4200 m/s
110/2.6=42 ton to orbit, assume second stage dry weight is 4 ton or the same as falcon 9 second stage, it can be fragile as most structure stays in the skylon.

Edited March 10, 2016 by magnemoe

[sevenperforce]

On 3/9/2016 at 7:18 PM, blowfish said:

On 3/8/2016 at 6:33 PM, sevenperforce said:

If the SABRE's operating speed is limited by compressor heating, then it would benefit by gradually shifting compression off of the turbocompressor and onto a ram compressor. 

But it can't do that, because it can only accept air in the combustor (spill ramjets aside).

Ram compression happens regardless.  It's not something you can control.  You could possibly reduce turbocompression while beginning to inject LOX into the combustion chamber to keep thrust up ... maybe it could gain you another half a mach or so.

I mean that there is a range of airspeeds where ram compression could compensate for gradual losses in turbocompression IF there was a way to inject ram-compressed air.

On 3/9/2016 at 7:18 PM, blowfish said:

There's no reason to inject into the nozzle.  The fuel/air ratio, even if you burn all the amospheric oxygen, is only a few % with hydrogen.  Giving yourself more airflow without burning its oxygen works fine at low speeds, but it ceases to be an effective strategy once you get supersonic - turbofans loose thrust a lot faster than turbojets, even with the same core.

But injecting airflow into the nozzle still allows you to burn its oxygen. It's the same principle as an afterburner, but in reverse; instead of injecting fuel into air-rich exhaust in the nozzle, you're injecting air into fuel-rich exhaust in the nozzle.

Turbofans lose thrust rapidly as you reach supersonic because they rely on mechanically forcing air around the engine using the fan. But if you inject directly into the nozzle, you don't have to rely on mechanical force to pull the air past; you're using mechanical force to compress and inject the air but the hot exhaust does the work of pushing the air out. So it's not a pure turborocket; it's a turborocket with reheat. Kinda like this.

On 3/9/2016 at 7:18 PM, blowfish said:

On 3/8/2016 at 6:33 PM, sevenperforce said:

The central bypass approach seemed to be the best way to get "high-bypass turbofan" performance. You can have a large intake because you can have a really, really ridiculously large effective nozzle. 

I still don't see the benefit.  Remember that in a shock cone intake, the capture area is not the area between the cone and the lip, but, at least at the design condition, is the entire frontal area of the intake up to the lip (by design condition I mean that the shockwave intercepts the lip).  You're not getting any more air with your geometry.

Yeah, I recognize that I'm not getting any additional air...but I'm also allowing a lot of bypass to compensate for densities or speeds which would cause extreme drag.

The purpose of turning it inside out was essentially to do for intakes what an annular aerospike nozzle does for nozzles. With the benefit of having a virtual aerospike on the back end.

 

[Val]

6 hours ago, magnemoe said:

Skylon is an realistic design for an SSTO but its still hard, with the typical airframe weight increase its unlikely to bring much cargo to orbit. 

93143 over at the nasaspaceflight did some maths on how much wrong the estimates for Skylon needs to be before it has no payload fraction left.

Quote

We've been over this before.  It's not a matter of a "slight shortfall"; this is not an all-rocket SSTO and is nowhere near as sensitive to either mass growth or engine underperformance.

According to my calculations, based on the numbers in the 2014 NISSIG presentation*, total payload loss would require one of:  A) an 11.6% Isp loss across both engine modes, B) a 12.2% rocket Isp loss with no airbreathing performance loss, C) a 72% loss of airbreathing Isp with no rocket performance loss, D) 30.5% dry mass growth, or E) some combination of the above.

But that's assuming no mass margin is being carried in the design, which is not accurate; they claim they're using mass growth margins "consistent with AIAA guidelines" for Skylon D1, which to me means at least 15% (IIRC their structural calculations for Skylon C1 back in the day used 15%).  Recalculating with dry mass divided by 1.15 results in the scenarios becoming:  A) a 16.8% Isp loss across both engine modes (for perspective, this would be like the SSME coming in at 379 s vac), B) a 17.5% rocket Isp loss with no airbreathing performance loss, C) an 80% loss of airbreathing Isp with no rocket performance loss, D) 50% dry mass growth, or E) some combination of the above.

* _{325,000 kg at start of roll (from which I have subtracted 1,418 kg of brake water, under the assumption that not carrying it through the takeoff roll won't affect the numbers much), 299,819 kg at transition, 73,435 kg at MECO, 52,347 kg dry.  Payload to standard orbit is 15,000 kg with 986 kg performance margin.  For the purpose of these calculations, I have assumed that lost payload due to underperformance is subtracted from GTOW - ie: the vehicle is not being redesigned either to have lower dry mass or to carry more propellant in place of the lost payload.}

If that is correct, it looks pretty promising to me.

[blowfish]

2 hours ago, sevenperforce said:

I mean that there is a range of airspeeds where ram compression could compensate for gradual losses in turbocompression IF there was a way to inject ram-compressed air.

All the air is ram-compressed.  The question is how much additional turbocompression you apply.

2 hours ago, sevenperforce said:

But injecting airflow into the nozzle still allows you to burn its oxygen. It's the same principle as an afterburner, but in reverse; instead of injecting fuel into air-rich exhaust in the nozzle, you're injecting air into fuel-rich exhaust in the nozzle.

Okay, so

1. An afterburner is entirely in front of the nozzle
2. Combustion takes a nontrivial amount of time.  Afterburners are very long tubes with fuel injection at the front for that reason.  Flow is entirely subsonic over the course of combustion.
3. Getting good combustion requires very careful mixing of the air and fuel - afterburners have spray bars that distribute the fuel in an even mist throughout the flow.  Rockets have arrays of fuel injectors which serve a similar purpose.  Supersonic flows are basically not going to mix at all.

2 hours ago, sevenperforce said:

The purpose of turning it inside out was essentially to do for intakes what an annular aerospike nozzle does for nozzles. With the benefit of having a virtual aerospike on the back end.

The inlet geometry still doesn't make sense.  You're certainly not going to get any better performance than a shock cone intake.