Reusable Single Stage to Orbit Concept

[Sumontro Sinha]

Hi, I'm new to this site. I have skimmed the pages a bit to keep abreast of new ideas and new viewpoints. From all the threads and articles that I have read, I feel there are a lot of qualified people that post here, and can help with proofing concepts.

From all the literature in NASA, AIAA, Acta Astronautica, and on this forum, I know that a reusable rocket based SSTO is not possible due to the insanely small mass fraction needed. Also, both after experiencing it myself, and hearing it from all the engineers that designed the Shuttle, X-33, and NASP, I know that closing a design by using lots of new technology usually dooms it to not meeting its goals. 

So here's my idea for a reusable SSTO vehicle. The vehicle would have two sets of engines, mixed flow turbofans (Like a F100), and rocket engines (like a RS-68). It would take off and land horizontally. To ascent to orbit, it would use air breathing engines and aerodynamic lift to give the vehicle some velocity and altitude, before boosting into orbit using its rocket engines. It would use a similar lifting entry as the shuttle. Before landing it would reignite its engines, to allow runway misses in case of problems. For atmospheric controls, it would use electromechanical controls powered by on board fuel cells. For OMS and RCS, it would use heritage shuttle hardware, since it has been proven in flight. To keep fuel weight down, it would use LOX/LH2 for the rocket engines. To keep engine weight down, the wings would use the same type of airfoils used on the F-16 to allow high L/D operation at subsonic, transonic, and supersonic speeds.

The only advances in technology, that I think, would be needed, are in the heat shield. The heat shield would have to be more durable, and have a higher temperature resistance than the Shuttle tiles.

I realize the design might seem totally outlandish, but I would appreciate any and all feedback. 

[Rabidpanda]

From all the literature in NASA, AIAA, Acta Astronautica, and on this forum, I know that a reusable rocket based SSTO is not possible due to the insanely small mass fraction needed.

I am of the opinion that a rocket based SSTO RLV is possible if denser fuels are used in the place of hydrolox.  Methylacetylene and lox is an example of a promising combination.

[Sumontro Sinha]

From all the literature in NASA, AIAA, Acta Astronautica, and on this forum, I know that a reusable rocket based SSTO is not possible due to the insanely small mass fraction needed.

I am of the opinion that a rocket based SSTO RLV is possible if denser fuels are used in the place of hydrolox.  Methylacetylene and lox is an example of a promising combination.

Are there existing engines that can run on this combination?

[Rabidpanda]

From all the literature in NASA, AIAA, Acta Astronautica, and on this forum, I know that a reusable rocket based SSTO is not possible due to the insanely small mass fraction needed.

I am of the opinion that a rocket based SSTO RLV is possible if denser fuels are used in the place of hydrolox.  Methylacetylene and lox is an example of a promising combination.

Are there existing engines that can run on this combination?

No.  If you want to use existing engines kerolox would be the best choice.

[gospacex]

So here's my idea for a reusable SSTO vehicle.

Why exactly do you need your vehicle to be RLV and SSTO?

[IsaacKuo]

So here's my idea for a reusable SSTO vehicle. The vehicle would have two sets of engines, mixed flow turbofans (Like a F100), and rocket engines (like a RS-68). It would take off and land horizontally.

This is similar to Rocketplane XP, which would have been strictly suborbital.  The combination of jet engines and rocket engines results in too much dry mass to get to orbit.

[strangequark]

Hi, I'm new to this site. I have skimmed the pages a bit to keep abreast of new ideas and new viewpoints. From all the threads and articles that I have read, I feel there are a lot of qualified people that post here, and can help with proofing concepts.

From all the literature in NASA, AIAA, Acta Astronautica, and on this forum, I know that a reusable rocket based SSTO is not possible due to the insanely small mass fraction needed. Also, both after experiencing it myself, and hearing it from all the engineers that designed the Shuttle, X-33, and NASP, I know that closing a design by using lots of new technology usually dooms it to not meeting its goals. 

So here's my idea for a reusable SSTO vehicle. The vehicle would have two sets of engines, mixed flow turbofans (Like a F100), and rocket engines (like a RS-68). It would take off and land horizontally. To ascent to orbit, it would use air breathing engines and aerodynamic lift to give the vehicle some velocity and altitude, before boosting into orbit using its rocket engines. It would use a similar lifting entry as the shuttle. Before landing it would reignite its engines, to allow runway misses in case of problems. For atmospheric controls, it would use electromechanical controls powered by on board fuel cells. For OMS and RCS, it would use heritage shuttle hardware, since it has been proven in flight. To keep fuel weight down, it would use LOX/LH2 for the rocket engines. To keep engine weight down, the wings would use the same type of airfoils used on the F-16 to allow high L/D operation at subsonic, transonic, and supersonic speeds.

The only advances in technology, that I think, would be needed, are in the heat shield. The heat shield would have to be more durable, and have a higher temperature resistance than the Shuttle tiles.

I realize the design might seem totally outlandish, but I would appreciate any and all feedback. 

The first obvious critique is that your jet engines don't buy you all that much. To achieve LEO, including gravity and drag losses, you have to achieve a delta-V of about 9500 m/s. Your jet engines may buy you 600-700 m/s. With a rocket engine of 450s of Isp, this means you can have an inert fraction of 13.5% instead of 11.5%. To give a basis of comparison, a Delta IV is about 300,000 kg. 2% of the gross liftoff weight is 6000 kg. Two F100 engines already weigh almost 4000kg, and would be grossly insufficient for something the size of a Delta IV. Even if they weren't, you'd eat up any additional savings with the wings. The problem only gets worse with smaller vehicles, because the weight savings are less.

[Hauerg]

IF you really want to go with RLV (an that's a big IF, you need the market, flight rate, and what if you crash one of your few or even the only vehicle??) you should first consider going TWO stages. In an SSTO approach almost every minimal design/development issue will lead to negative payload.

[Cinder]

... Skylon?
http://forum.nasaspaceflight.com/index.php?topic=21530.0

[Sumontro Sinha]

Yes, if you take the Delta IV as the reference for mass fraction, the dry weight would be too much. But the mass fraction of 13.5% is right about that of the Space Shuttle stack, and this concept would get rid of all the staging problems. Because of using aerodynamic lift and lack of extra boost hardware, the fuel needed to overcome the drag and gravity losses at lower altitudes with rocket engines cuts out half the fuel needed by the shuttle. You are correct in saying the jet engines only give you a delta v of 600-700 m/s, and the wings add additional weight. However, unlike on a rocket, they help you both on ascent and entry. Also, the wings and jet engines give the vehicle the ability to test incrementally, so problems can be spotted in a more controlled manner.

In terms of payload to orbit, the design seems to work if the turbojet weight is reduced by increasing the craft L/D to around 15. Then the design seems to be able to take a moderate payload into orbit.

Also, I thought Rocketplane's suborbital concept was the size of a business jet, so it couldn't carry the amount of fuel needed to get to orbit.

[JohnFornaro]

So here's my idea for a reusable SSTO vehicle.

Why exactly do you need your vehicle to be RLV and SSTO?

Pay him no mind.  It's a great idea.  I think that the biggest problem is to build it and launch it for less money than current expendable vehicles, and do that on the first or second launch.  Not that I know how to do this, mind you, but if you don't do this, your vehicle will [ahem] never get off of the ground.

[Sumontro Sinha]

Skylon does not use current engines, and they seem to STILL be having enormous problems with their engines even after nearly 20 years of the development. So while great, it seems it might take a while longer for it to come to fruition.

[Sumontro Sinha]

So here's my idea for a reusable SSTO vehicle.

Why exactly do you need your vehicle to be RLV and SSTO?

Pay him no mind.  It's a great idea.  I think that the biggest problem is to build it and launch it for less money than current expendable vehicles, and do that on the first or second launch.  Not that I know how to do this, mind you, but if you don't do this, your vehicle will [ahem] never get off of the ground.

I agree with you. I reviewed a lot of the material on the Shuttle, and listened to the folks that designed it; they all said that if they hit the design flight rate, the cost savings per launch would have been realized. Unfortunately, for them, they did not know the problems/additional costs that the TPS, engines, and increased overhead would bring.

Now for this vehicle, assuming an advance in TPS technology, I feel could launch for less cost than expendables if it brings in a launch rate like that intended for the shuttle (~60 per year). Now are there customers for that amount of launches? That is yet to be seen. I was only looking at the technical feasibility of the approach.

But in terms of development cost, that is a tough nut to crack. I was hoping on some input on which areas would shoot up the development cost and why.

[Cinder]

Pardon me, but what enormous problems?  The engines are due for testing "in ~3 years" which isn't the order of magnitude you'd think of for something facing enormous problems.  IIRC the cooling system is expected to be the most difficult development part, and they reportedly have that one under control.   

Delays, I sure don't see why not.  But enormous problems? 

[Sumontro Sinha]

well, how far has the airframe of the Skylon come? Also, how about the TPS? Are these the same structures and materials used on current aircraft and spacecraft? or are they different? How far is the Skylon vehicle from actually launching into orbit?

[sandrot]

... and they reportedly have that one under control.   

And they have reportedly flown nothing.

[sandrot]

well, how far has the airframe of the Skylon come? Also, how about the TPS? Are these the same structures and materials used on current aircraft and spacecraft? or are they different? How far is the Skylon vehicle from actually launching into orbit?

For information on paper rockets you've got to follow the paper trail.

http://www.reactionengines.co.uk/downloads/SKYLON_User_%20Manual_rev1%5B3%5D.pdf

[Cinder]

Is there a thread I could be directed to that clearly explains for dummies what the difference is, not in hindsight but at the time of development, between a paper rocket that'll turn into flying hardware and one that'll turn into nothing beyond paper figures and line art?

[aero]

Is there a thread I could be directed to that clearly explains for dummies what the difference is, not in hindsight but at the time of development, between a paper rocket that'll turn into flying hardware and one that'll turn into nothing beyond paper figures and line art?

I think most will agree that you don't need a thread. The difference is political will and money. Look at Constellation and Jupiter Direct as examples. Both include paper rockets, and both could have been converted to flight hardware given the money. Of course the money comes as a result of the political will in government programs but a very similar "political" will is in effect for commercial efforts.

As for the subject of this thread, I suggest you read and study everything you can about the Skylon and its Sabre engines. Someone will correct me if I'm wrong but I think the Sabre powered Skylon vehicle comes about as close as we have ever gotten to a viable subject vehicle for this thread.

Note that Mark Hempsell posts here. If we are lucky he will post a list of available public documents that would provide a Skylon/Sabre familiarization and training course. As for me, I suggest starting with the Wikipedia article then following the links included there-in. Google Skylon/Sabre.

[Cinder]

If the Skylon formula works out, I haven't yet seen any arguments for it not being a very viable space access solution for smaller payloads.  One that'd garner enough interest to very possibly break that initial investment friction. 

I have read everything I've come upon, and to the limits of my technical understanding there aren't any show stoppers in sight.   If there are, to the point that the OP formula is clearly a better alternative, even if just considered in a political/economical vacuum (or that Skylon per se's a clear no-go), I'm curious to hear them.

Of course if the topic's "proven technology" tag is an absolutely strict criteria for discussion here, then I have no argument there -- but we're back to the same catch 22 that I can't get my head around as a reasonable premise for speculation on "Reusable single stage to orbit concepts".

And yes, M. Hempsell posted here, IIRC mostly in that thread I linked to.  S. Sinha might be interested in those posts.

[Sumontro Sinha]

I agree that the Skylon concept seems good. But how are they protecting both the engines and the airframe during reentry? If it is routing the liquid hydrogen through the airframe, how do they handle the large pressures required to prevent cavitation in the LH2?

Also, the discussion has gone a little tangent. This question is directed at anyone, but I want to ask Cinder, what is your critique of the SSTO concept posted here?

I know that it is essentially the Skylon concept with existing technology, but since it seems like you know a lot about the Skylon, how feasible do you think the SSTO concept that thread this thread is?

[Sumontro Sinha]

Also, I have read the user manual that sandrot is referring to. The TPS that they refer to, the multilayer metal TPS, and glass ceramic material, has it been flight proven? Or is it in the middle of testing?

Also, in terms of the SABRE engine, how is this different from the other air breathing rocket engines that were tried in the 60's by Marquedt and others? Does the SABRE have a mechanism for liquifiying only atmospheric oxygen when it is flying?

[IsaacKuo]

Also, I thought Rocketplane's suborbital concept was the size of a business jet, so it couldn't carry the amount of fuel needed to get to orbit.

How big is your proposed vehicle, and how much fuel does it need to carry in order to get to orbit?

By the way, you can't just add more and more fuel to a design in order to get to orbit.  Adding more fuel means adding more mass that needs to be lifted.

[marsavian]

Sumontro,

There already is a SSTO thread where a Skylon developer already posts. Perhaps you can ask Chris/Mods to merge this thread into that one.

http://forum.nasaspaceflight.com/index.php?topic=21530.0
http://forum.nasaspaceflight.com/index.php?action=profile;u=15949;sa=showPosts

[Sumontro Sinha]

Also, I thought Rocketplane's suborbital concept was the size of a business jet, so it couldn't carry the amount of fuel needed to get to orbit.

How big is your proposed vehicle, and how much fuel does it need to carry in order to get to orbit?

By the way, you can't just add more and more fuel to a design in order to get to orbit.  Adding more fuel means adding more mass that needs to be lifted.

Currently, (assuming I did the calculations correctly) it is a bit over the mass of a fully loaded 747. The fuel needed to reach orbit is a little under 550,000 kg. I realize that is a lot heavier than most expendable rockets, but the dimensions of the concept are around that of a 747. Also, to give the design some margin for weight increases, aerodynamic efficiency (I.E. L/D) is maximized, using a variable airfoil like that used on the F-16. This allows the plane to takeoff with much fewer jet engines, and makes a positive weight loss cycle. 

[IsaacKuo]

Currently, (assuming I did the calculations correctly) it is a bit over the mass of a fully loaded 747. The fuel needed to reach orbit is a little under 550,000 kg. I realize that is a lot heavier than most expendable rockets, but the dimensions of the concept are around that of a 747.

A 747 has an empty mass of about 180,000kg.  With 550,000kg of fuel, the mass ratio is only 4:1.  Even with a specific impulse of 455s, that would only get you 6km/s of delta-v--far below what's needed for orbit.

Also, to give the design some margin for weight increases, aerodynamic efficiency (I.E. L/D) is maximized, using a variable airfoil like that used on the F-16. This allows the plane to takeoff with much fewer jet engines, and makes a positive weight loss cycle. 

You've repeated this several times, so I know it's not just a typo.  What do you think is so special about the F-16, and why do you think it is not already used by the 747?

[Sumontro Sinha]

Because, the F-16 flies at a wide range of airspeeds, from subsonic to supersonic, and wide range of angles of attack without getting excessive amounts of drag. The 747 can neither fly at supersonic nor fly at large angles of attack. This concept would require a pitch up maneuver to be able to make orbit.

Also, when fully fueled, the concept has the structural mass fraction of the Space Shuttle, around 13% of total mass. I know that seems tenuous, so I was wondering why the structural mass would be more for this type of SSTO since all it does is bring the Shuttle ET inside the reusable airframe, while getting rid of the heavier boosters. If anyone can explain why this would not work, I would greatly appreciate it.

[IsaacKuo]

Because, the F-16 flies at a wide range of airspeeds, from subsonic to supersonic, and wide range of angles of attack without getting excessive amounts of drag. The 747 can neither fly at supersonic nor fly at large angles of attack. This concept would require a pitch up maneuver to be able to make orbit.

Large angles of attack absolutely always mean getting excessive amounts of drag.  This is true of the F-16, the same as any other aircraft.  An F-16 will bleed off its airspeed rapidly at a high angle of attack, but this is a necessary evil when engaging in a dogfight.  A high angle of attack is needed in order to make severe turns.

I do not know why you think a "pitch up maneuver" is required to make orbit.

Anyway, I need to clarify--what do you mean by a "variable wing"?  The F-16 has a fixed wing, and its wing is a design compromise.  It's not nearly as efficient as the 747 wing.

Also, when fully fueled, the concept has the structural mass fraction of the Space Shuttle, around 13% of total mass.

What is the dry mass of your proposed vehicle?  Is it supposed to be 82,000kg?  With a dry mass of 82,000kg and 550,000kg of fuel, the dry mass would be 13% of the total mass.

But that's less than half the mass of a 747!  In other words, your dry structural mass is far to small.

I know that seems tenuous, so I was wondering why the structural mass would be more for this type of SSTO since all it does is bring the Shuttle ET inside the reusable airframe, while getting rid of the heavier boosters. If anyone can explain why this would not work, I would greatly appreciate it.

You're not just bringing the Shuttle ET inside the reusable airframe.  The Shuttle ET won't fit inside the Shuttle's reusable airframe.  It's too big.  Nor are the Shuttle's wings big and strong enough to fly with the massive ET.  It's too heavy.  Nor does the Shuttle have any jet engines to fly horizontally.  Nor will the Shuttle's landing gear support the weight of the massive ET.

If you start from the other direction, a 747 is just plain a lot heavier than what you're hoping for.  Unlike the Space Shuttle, it needs big heavy wings to support a huge mass.  It needs beefy landing gear to take support all that mass.  It needs big heavy jet engines to take off and fly.

[strangequark]

Also, when fully fueled, the concept has the structural mass fraction of the Space Shuttle, around 13% of total mass. I know that seems tenuous, so I was wondering why the structural mass would be more for this type of SSTO since all it does is bring the Shuttle ET inside the reusable airframe, while getting rid of the heavier boosters. If anyone can explain why this would not work, I would greatly appreciate it.

Shuttle ET: 26.5 mT
Propellant: 735 mT
4 SSMEs: 12.8 mT
Thrust Structure (Shuttle Boattail Mass): 20 mT
12 F100 engines (0.2 g initial acceleration): 20 mT
Payload: 10 mT

Required dV: 8800 m/s (assume 700 m/s from F100s)
Remaining Mass Available for wings, etc: 4.5 mT

A 747 wing weighs in at 43 mT and lifts an aircraft that weighs less than half as much.

[Rhyshaelkan]

Aye, as of yet, big dumb boosters seem the best way to get material into space.

When space industry, based on asteroids or lunar materials, gets to a certain level then things might change. Who knows what questions might be answered once out there. But to get there we should use the cheapest most reliable big dumb boosters we can.

When volatiles, rare earth elements, and large amounts of raw materials start to be harvested in space for use in space. Then the need for launches from Earth should drop precipitously. The only real cargo from Earth being humans.

Not quite on topic. But is how I feel about spacecraft design at this moment.

[alexterrell]

Aye, as of yet, big dumb boosters seem the best way to get material into space.

When space industry, based on asteroids or lunar materials, gets to a certain level then things might change. Who knows what questions might be answered once out there. But to get there we should use the cheapest most reliable big dumb boosters we can.

When volatiles, rare earth elements, and large amounts of raw materials start to be harvested in space for use in space. Then the need for launches from Earth should drop precipitously. The only real cargo from Earth being humans.

Not quite on topic. But is how I feel about spacecraft design at this moment.

Does a big dumb booster need to be reliable? If you use it to reansport water, aluminium, glass etc, and you launch 1 per day, maybe you can accept a 10% failure rate?

Once volatiles etc are harvested in Space, the basic materials will be replaced by people and advanced goods, so I wouldn't expect the need for launches to fall. A relatively small Torus station might have 3,000 rooms, for 150,000 guests per year. Then you do need extreme reliability.

[gospacex]

So here's my idea for a reusable SSTO vehicle.

Why exactly do you need your vehicle to be RLV and SSTO?

Pay him no mind.  It's a great idea.  I think that the biggest problem is to build it and launch it for less money than current expendable vehicles

Bingo. That's what the goal should be: launch stuff to orbit for less $/kg. How it is done (TSTO? SSTO? RLV? Magic carpet?) is secondary.

Topic starter again tried to put the cart (SSTO RLV) before the horse (cost). We already did that with Shuttle.

[Rhyshaelkan]

Does a big dumb booster need to be reliable? If you use it to reansport water, aluminium, glass etc, and you launch 1 per day, maybe you can accept a 10% failure rate?

Once volatiles etc are harvested in Space, the basic materials will be replaced by people and advanced goods, so I wouldn't expect the need for launches to fall. A relatively small Torus station might have 3,000 rooms, for 150,000 guests per year. Then you do need extreme reliability.

Aluminum and silicon are in abundance on Luna. And as exploration has shown hydrogen, possibly in the form of water-ice, exists in usable amounts. If we would launch something to support ongoing lunar industry. It would be those things that we just cannot get yet. Petrochemical items, heavy metals, human comfort items, humans.

Further exploration of lunar asteroid impacts might show that even what once was thought rare on Luna, might be quite accessible. Heavy metals that are not in the normal regolith could be found in breccias. If so, that is one less group of items needed to import from Earth to support lunar industry.

With that industry comes the ability to launch asteroid harvesting missions straight from Luna. With the number of supposed dormant comets as Earth-crossers, we will then be able to eliminate volatiles from the list of imports.

Possible money makers to support the system. SPS construction and delivery, space going vessels(probes, OTVs, etc), hydrox fuel. Lunar based telescopes with time sold to universities across the Earth. All from lunar materials, with little to no need for an Earth launch, except some workers. If they are motivated. They could make their "commute" well worth it.

If you had an OTV run on lunar fuel. You could have it enter LEO to grab a Dragon capsule to take 7 folks to Luna for the commute. As the lunar industry demand increases you could even charter an F9H with a Dragon II to deliver 21 peoples to Luna for the cost of $100M.

It is a universe of infinite possibility. It just takes some investment to get over that hump./me shrugs

[gospacex]

Aluminum and silicon are in abundance on Luna. And as exploration has shown hydrogen, possibly in the form of water-ice, exists in usable amounts.

I doubt that it is in the form of ice. More likely as a hydrated regolith, approximately as "wet" as dried concrete. And there is really not that much of it.

With that industry comes the ability to launch asteroid harvesting missions straight from Luna. With the number of supposed dormant comets as Earth-crossers, we will then be able to eliminate volatiles from the list of imports.

This is not known. They all may be very dry. Even active comets proved to be rather dry on surface.

Lunar based telescopes with time sold to universities across the Earth. All from lunar materials, with little to no need for an Earth launch, except some workers. If they are motivated. They could make their "commute" well worth it.

The Moon is a bad place for a telescope - too much dust. Even LEO isn't ideal.

It is a universe of infinite possibility. It just takes some investment to get over that hump./me shrugs

Where "some" is an understatement.

[ukrocketman]

Does the SABRE have a mechanism for liquifiying only atmospheric oxygen when it is flying?

SABRE does not liquify atmospheric oxygen. This is a common misconception people make. The incoming air is deeply cooled, but it is not liquified.

[JohnFornaro]

Aluminum and silicon are in abundance ... you could even charter an F9H with a Dragon II to deliver 21 people to Luna...

Yep.  I think that the pull of the mineral resources of space is beginning to be felt.  Maybe history will suggest that the discovery of water ice on the lunar poles was the turning point.  If we can manufacture massive things up there, like spacecraft, habitats, fuel, air, and water then we can concentrate on launching high value cargo, people, and carbon to the lunar surface.  If we can make that happen, the other destinations will naturally become available.

The Moon is a bad place for a telescope...

An optical one, yeah.  But a radio one on the far side?

Magic carpet?

From what I've seen, this is actually one of the reasons the Iraq thing hasn't gone as smoothly as predicted.  These things are very maneuverable in the atmosphere and are very had to catch....

[Rhyshaelkan]

Aluminum and silicon are in abundance on Luna. And as exploration has shown hydrogen, possibly in the form of water-ice, exists in usable amounts.

I doubt that it is in the form of ice. More likely as a hydrated regolith, approximately as "wet" as dried concrete. And there is really not that much of it.

With that industry comes the ability to launch asteroid harvesting missions straight from Luna. With the number of supposed dormant comets as Earth-crossers, we will then be able to eliminate volatiles from the list of imports.

This is not known. They all may be very dry. Even active comets proved to be rather dry on surface.

Lunar based telescopes with time sold to universities across the Earth. All from lunar materials, with little to no need for an Earth launch, except some workers. If they are motivated. They could make their "commute" well worth it.

The Moon is a bad place for a telescope - too much dust. Even LEO isn't ideal.

It is a universe of infinite possibility. It just takes some investment to get over that hump./me shrugs

Where "some" is an understatement.

With a name like "gospacex" I would expect you to be more positive about space.

The page here says 600 million tons of water/hydrogen. By the radar you can also assume it is concentrated in the permanently shadowed craters.

The lunar surface is actually a pristine environment. No rain, no wind. The only dust will be when mankind disturbs the land. So to that end. Sinter a building pad using the method found here. Once the pad is sintered then drive the "truck" to the edge of the pad and unload the building materials. With the telescope assembled, given a final clean, hooked up to the Lunar Broadband Network and Lunar Power Grid. Then the humans retreat. The telescope should stay that way for eons. The only dust would come from the occasional maintenance that must be performed. But with proper steps taken when entering and leaving the area it should stay clean. Now you have a telescope able to train its eye on a section of the universe for 14 days at a time. At certain latitudes you could get constant surveillance.  Lather, rinse, and repeat as the astronomical community demands.

As for costs. It only costs as much as bureaucracy allows. Elon is doing well at starting the "cut the glut" process. A few more steps to cut the fat out of the process, and space could be orders of magnitude cheaper. Space will always be expensive, I grant you. However with a proper plan it should start to support itself.
/me shrugs emphatically

[93143]

Wasn't there a phenomenon observed during Apollo whereby electrical imbalances between the day side and the night side gave rise to dust flows near the surface of the moon?

[Rhyshaelkan]

Would be the first I heard of it. If it is so, just build that into your plan.


If lunar astronomy was desired still. You could pave a much larger area than what I had initially planned. To accommodate a cluster of telescopes instead of just one. Then around the edge of your sintered platform an electrostatic fence to keep out the dust.

Or perhaps a concept like this. Artwork by Pat Rawlings.

[savuporo]

Magic carpet?

From what I've seen, this is actually one of the reasons the Iraq thing hasn't gone as smoothly as predicted.  These things are very maneuverable in the atmosphere and are very had to catch....

MCTO. Mass fraction includes the hookah.

[gospacex]

The lunar surface is actually a pristine environment. No rain, no wind. The only dust will be when mankind disturbs the land.

Nope, there is also electrostatic effects and secondary impacts from crater formation elsewhere.

Without air, dust kicked up by impact (or rocket exhaust) elsewhere can travel hundreds or thousands of kilometers, without losing one iota of its quite respectable velocity.

So to that end. Sinter a building pad using the method found here. Once the pad is sintered then drive the "truck" to the edge of the pad and unload the building materials. With the telescope assembled, given a final clean, hooked up to the Lunar Broadband Network and Lunar Power Grid. Then the humans retreat. The telescope should stay that way for eons. The only dust would come from the occasional maintenance that must be performed. But with proper steps taken when entering and leaving the area it should stay clean. Now you have a telescope able to train its eye on a section of the universe for 14 days at a time.

Why bother with all this stuff when in certain orbits, a telescope is able to look at the same patch of sky for years? (hint: Kepler)

As for costs. It only costs as much as bureaucracy allows. Elon is doing well at starting the "cut the glut" process. A few more steps to cut the fat out of the process, and space could be orders of magnitude cheaper.

Orders (plural) of magnitude cheaper? As in "$100/kg in LEO"? The required amount of LOX/RP fuel alone will cost about that much!

I an a Musk amazing people allright, but that kind of efficiency is just not going to happen for many decades, if not centuries.

[mlorrey]

The Moon is a bad place for a telescope - too much dust. Even LEO isn't ideal.

The lunar dust problem is primarily an issue of electrostatic suspension at the day/night terminator

The only real utility of the back side of the moon for astronomy is with radio astronomy.

[colbourne]

Orders (plural) of magnitude cheaper? As in "$100/kg in LEO"? The required amount of LOX/RP fuel alone will cost about that much!

I an a Musk amazing people allright, but that kind of efficiency is just not going to happen for many decades, if not centuries.

I think the way to get there , is by use of a rotorvator. Hopefully within 20 years. 

[Hempsell]

I am Mark Hempsell  and I am sorry to be late in to this thread. We are having a rather busy time At Reaction Engines and I can only web-surf very occasionally and so I have only just caught up with this thread.

Let me start by commenting on Sumontro Sinha original outline which is the subject of this thread.  The problem with jet engines is that for the purposes of getting to orbit they have a poor thrust to weight ratio. So the advantage you gain from airbreathing  is lost because of the impact the the heavy jet engines have on the dry mass and hence the final speed as determined by the rocket equation.  In short to reach orbits you will have to somehow ditch the jet engines after they have done their job.

This is why viable proposals for the use of conventional jets in launch systems end up with a two stage solution. Subsonic separation (e.g. of the back of a 747 or An225) does make the problem easier than a pure rocket SSTO but if you use existing aircraft the payload is rather small,  and if you develop a new, much larger, aircraft it is very expensive.

If you go to supersonic or hypersonic separation you have a serious issue with the aerodynamics of the separation and you are probably best with the Bristol Spaceplanes solution of fitting the lower stage with rockets to boost the stack outside the dense atmosphere before separating. But a large hypersonic aircraft with both rocket and jet engines is a very expensive system to develop alongside the orbiter. 

So while many two stage proposal can be made to work they all have the double whammy development cost issue. The exception is the Siamese approach where the two stages are identical, that is you develop one vehicle that can perform both the booster and orbit roles, and this would be my favourite of the two stage approaches.

Where SKYLON differs is that the extra kit need to make the rocket airbreath is light, although it does achieve this by sacrificing performance - the SABRE 3 is only half as efficient as a jet engine, but still much better than a rocket.  Apart from the heat exchangers the components are all standard rocket and jet technology and the heat exchangers are the subject of a technology development programme to show flight standard heat exchangers produced in realistic production facilities.  It is not true to say we are having trouble with the engine the 20 years is a solely funding induced.

If you go to Reaction Engines website and look under “pdf downloads” which itself is under “”Media Library” on the home page menu

http://www.reactionengines.co.uk/pdf_documents.html

You will find several refereed journal papers (mostly JBIS)

The SKYLON Spaceplane - Progress to Realisation Richard Varvill, and Alan Bond
Is the best summary as to where we are although things have progressed rapidly in the last two years as our monthly news pages records.

The SKYLON Spaceplane  Richard Varvill and Alan Bond
Is the best technical summary of the SKYLON C1 configuration

A Comparison of Propulsion Concepts for SSTO Reuseable Launchers Richard Varvill and Alan Bond
Explains the logic behind the engine selection

Application of Carbon Fibre Truss Technology to the Fuselage Structure of the SKYLON Spaceplane Richard Varvill and Alan Bond
Explains the Structure

[tnphysics]

What is the T/W of an ATR, which can reach Mach 8?

Mach 8 is enough to leave the atmosphere with a pitch-up maneuver.

[Warren Platts]

If you could refuel in LEO, that would radically simplify the design of the SSTO, since you could do a mostly propulsive re-entry, and thus skip the TPS.

[M_Puckett]

Of course if you have an abort at high delta V, you are going to be cooked without a TPS.

You cannot get by without a native TPS.

[Warren Platts]

You mean like Challenger?!?

[Jorge]

You mean like Challenger?!?

No, I think he's referring to a TAL/AOA type scenario, where orbit insertion is not possible but the vehicle's velocity is too high to be slowed down propulsively. That would be a black zone if the vehicle lacks TPS.

[Warren Platts]

TAL = Transatlantic Abort Landing; AOA = Around Once Abort

OK, fair enough; although I'm unaware of such an abort mode ever being necessary in the history manned spaceflight. However, Challenger also shows that TPS's aren't perfectly failsafe. Choose your poison. I'm just sayin' if you didn't have to worry about a TPS, it would simplify the job of designing a SSTO.

[Jorge]

TAL = Transatlantic Abort Landing; AOA = Around Once Abort

OK, fair enough; although I'm unaware of such an abort mode ever being necessary in the history manned spaceflight.

That is only because the total number of manned spaceflights to date have been pitifully small. The Wright brothers probably made more airplane flights in their first year than all the spaceflights of all nations in the last 50 years.

However, Challenger also shows that TPS's aren't perfectly failsafe.

The Challenger breakup occurred so early in the mission that the presence or absence of TPS was irrelevant.

Choose your poison. I'm just sayin' if you didn't have to worry about a TPS, it would simplify the job of designing a SSTO.

Oh, I agree, it's easy to design an unacceptably unsafe spacecraft.

[Downix]

This is why viable proposals for the use of conventional jets in launch systems end up with a two stage solution. Subsonic separation (e.g. of the back of a 747 or An225) does make the problem easier than a pure rocket SSTO but if you use existing aircraft the payload is rather small,  and if you develop a new, much larger, aircraft it is very expensive.

My own twist on this is a two stage SSTO, jet engine to get to low Mach/high altitude, but staging is not till much later on.  The carrier also functions as a fuel tank, and is carried by the orbital element for a period before release, after which it flys back under its own power.

it just makes sense to me, you can have a much more efficient orbital element that way.

[kch]

... a two stage SSTO ...

... so, two-stage single-stage-to-orbit ... "two-stage single-stage" ... hmmm ... gonna have to ponder that one a bit ...

[sandrot]

... a two stage SSTO ...

... so, two-stage single-stage-to-orbit ... "two-stage single-stage" ... hmmm ... gonna have to ponder that one a bit ...

Steve Cook's Law of Staging: "After the 2nd to last stage has been dropped, they're all SSTO".

[kch]

... a two stage SSTO ...

... so, two-stage single-stage-to-orbit ... "two-stage single-stage" ... hmmm ... gonna have to ponder that one a bit ...

Steve Cook's Law of Staging: "After the 2nd to last stage has been dropped, they're all SSTO".

There ya go! 

[FinalFrontier]

I dont know how many times I have seen this discussed only to fall due to the same basic problems.

Let me lay out the things that any SSTO, let alone RESUSABLE SSTO designers will need to overcome in order to make this idea reasonable, let alone viable:

1. COST COST COST: If its not cost effective, nay, if its not COMMERCIALLY economic (i.e. as in for a commercial launch provider) then it won't work. Note: Don't expect government funding for this. You might get it (DOD side at least), but might isnt enough. Design it to be commercially feasible.

2. Reusability: Determine a low cost and effective system to protect the stage during rentry such that refurbishing for reuse is quick and easy.

3. Retriveal: Where are you going to land it and what are the consequences of landing there?

4. Feasibility: Is Reusability really worth it? Or is it too expensive??

5. Saftey: Make it safe

6. KG/$ to LEO economics: How much can it lift? Can it compete?

[JasonAW3]

TAL = Transatlantic Abort Landing; AOA = Around Once Abort

OK, fair enough; although I'm unaware of such an abort mode ever being necessary in the history manned spaceflight.

That is only because the total number of manned spaceflights to date have been pitifully small. The Wright brothers probably made more airplane flights in their first year than all the spaceflights of all nations in the last 50 years.

However, Challenger also shows that TPS's aren't perfectly failsafe.

The Challenger breakup occurred so early in the mission that the presence or absence of TPS was irrelevant.

     I think he was referring to the Columbia disaster.  Breakup during re-entry, not a good way to end a mission.
     I'm beginning to suspect an inflatable ballute TPS, would be a viable alternative to traditional systems and would allow a safe return mode.  Besides, increasing the surface area versus the total mass of a craft should lead to a lower tempreture re-entry, but at a higher Gee load.  It should also shorten the time it takes to slow down  a bit.

Jason

[alexterrell]

I am Mark Hempsell  and I am sorry to be late in to this thread. We are having a rather busy time At Reaction Engines and I can only web-surf very occasionally and so I have only just caught up with this thread.

Mark, Thanks for the profesional input, and best of luck at Reaction Engines.

I browsed the web site and have a general comment.

Any reusable SSTO is going to be expensive to develop. One way of making it cheaper is to share commonality with an air craft. For example, if you go for subsonic launch, design it round an existing aircraft - e.g A380.

That brings me to your Brussels to Sydney concept (though I laughed at "Brussels" - fill it with Eurocrats!). This seems too large for a first version. It's similar size to an A380, and carries more 1st and business class passengers, and makes more journeys. So BA would only every buy 2 or so.

Can it not be made the same size, with the same airframe, as Skylon?

[Hempsell]

I dont know how many times I have seen this discussed only to fall due to the same basic problems.

Let me lay out the things that any SSTO, let alone RESUSABLE SSTO designers will need to overcome in order to make this idea reasonable, let alone viable:

1. COST COST COST: If its not cost effective, nay, if its not COMMERCIALLY economic (i.e. as in for a commercial launch provider) then it won't work. Note: Don't expect government funding for this. You might get it (DOD side at least), but might isnt enough. Design it to be commercially feasible.

2. Reusability: Determine a low cost and effective system to protect the stage during rentry such that refurbishing for reuse is quick and easy.

3. Retriveal: Where are you going to land it and what are the consequences of landing there?

4. Feasibility: Is Reusability really worth it? Or is it too expensive??

5. Saftey: Make it safe

6. KG/$ to LEO economics: How much can it lift? Can it compete?

A very good list - needless to say that at Reaction Engines we believe that SKYLON ticks all those boxes.

[Hempsell]

I am Mark Hempsell  and I am sorry to be late in to this thread. We are having a rather busy time At Reaction Engines and I can only web-surf very occasionally and so I have only just caught up with this thread.

Any reusable SSTO is going to be expensive to develop. One way of making it cheaper is to share commonality with an air craft. For example, if you go for subsonic launch, design it round an existing aircraft - e.g A380.

That brings me to your Brussels to Sydney concept (though I laughed at "Brussels" - fill it with Eurocrats!). This seems too large for a first version. It's similar size to an A380, and carries more 1st and business class passengers, and makes more journeys. So BA would only every buy 2 or so.

Can it not be made the same size, with the same airframe, as Skylon?

The key issue on all these concepts is the airframe/engine integration. both must be matched and produce a trim-able flight system. The SABRE engine cannot be intergrated into any existing airframe, nor can its benefits be gained with lower Mach number separation. Even very small design changes from what we show rapidly erode the payload to zero.

A similar argument applies to the pure airbreather LAPCAT with its SCIMITAR engine. This is actually more difficult to engineer than SKYLON and would need an optimised airfame with a longer life, able to use existing airports and with active structural cooling to cope with around 4 hours at 1000 degrees.

LAPCAT is an EU technology programme and the Bussels to Sidney requirement came from them. Our A2 concept is one of three being used to explore the technologies required for hypersonic civil airflight - the other two concepts use scamjets. The A2 sizing was derived from an evaluation of the point where the journey costs match current business class fares.

(AND if I ever meet the author of the text editor on this site his or her  life expectancy will drop to seconds)

[JohnFornaro]

The Wright brothers probably made more airplane flights...

That was an interesting perspective.

3. Retrieval

I hadn't thought about that one so much.  You can land a Piper Cub on an interstate in a pinch.  You can land an Airbus in the water, if you're lucky. The shuttle has a few spots available for emergency landings also:

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

Part of the COST COST COST issue is SIZE SIZE SIZE.  To me, the size of this re-usable craft might ought to be limited to crew only, and maybe carry-on luggage.  Let the expendables be for cargo and propellant.  Just a general strategic sort of observation.

If I ever meet the author of the text editor...

Basically, it all boils down to one guy.  Name of Bill.  Lives in Redmond or Seattle?  Has a honkin' house?  Fixes software that doesn't need to be fixed; won't fix software that actually needs to be fixed.  I just use Notepad, FWIW.

[DLR]

What about VentureStar? The X-33 prototype was canceled largely because of the failure to construct a large liquid hydrogen tank out of composite materials. But I remember reading that a standard aluminum tank would have been sufficient or perhaps even lighter than the composite tank. I think the most straightforward path to SSTO would be to revive the X-33 programme.

[JohnFornaro]

I thought that the linear aerospike rocket engine worked on the x-33.  I don't understand why these engines haven't been developed further.  They certainly have contemporary use on upper stages, using conventional ullage.

[Hempsell]

John Fornaro

Quote “Part of the COST COST COST issue is SIZE SIZE SIZE.  To me, the size of this re-usable craft might ought to be limited to crew only, and maybe carry-on luggage.  Let the expendables be for cargo and propellant.  Just a general strategic sort of observation.”

There are 4 problems with going small.

1 Systems development and operational cost do not drop linearly with size. A system half the size is still three quarters the cost (very roughly) 

2 Smaller systems are more sensitive to mass variation and hence, while a bit cheaper, they are more technically risky.

3 People wanting to launch small payloads tend not to have much money. Any commercial system needs to grab the bulk of the market i.e. the cargo and the propellant, in order to generate the revenue to pay for its development.

4 Given the cargo and the propellant are by far the majority of what you want to launch, if you do not introduce the cost benefits to these payloads you do not change astronautics.



On the X-33


I believe it very likely that a single stage to orbit reusable based solely on pure rocket cannot be achieved with any foreseeable technology. But this aside, the X-33 VentureStar had several specific problems inherent to its approach.

1 The configuration was not inherently trim-able; look at the difference between the X-33 and the final form of VentureStar! If you do not see trim as a fundamental killer of almost all SSTO concepts you have some learning to do.

2 Plug nozzles and especially the linear form are considerable heavier than bell nozzle equivalents.  The flow field in advanced nozzles is far more complex than most of the literature would have you believe and it is not clear either aerospikes (or expansion deflection nozzles which is what we at Reaction Engines are looking at) can in real operation engines deliver atmospheric compensation, and thus justify their extra mass.

3 Structural composite tanks are a nightmare (we had them on HOTOL) – composite truss frames and aluminium tanks everytime.

[M_Puckett]

TAL = Transatlantic Abort Landing; AOA = Around Once Abort

OK, fair enough; although I'm unaware of such an abort mode ever being necessary in the history manned spaceflight. However, Challenger also shows that TPS's aren't perfectly failsafe. Choose your poison. I'm just sayin' if you didn't have to worry about a TPS, it would simplify the job of designing a SSTO.

Ask the Russians, they have done it.

[JohnFornaro]

A system half the size is still three quarters the cost...

Readily granting the 75% ratio as an informal heuristic, it still may be that $30B is affordable and $40B not.  The old saw about three quarters of a loaf, and all.   I don't know what the correct answer is, but the correct answer on size will have to be developed in the context of past wastefulness and program failure; smaller may be required, but not desired.  Just sayin'.

Removing the human from the cargo and propellant launches will certainly save on Cost and Size, generally speaking.

configuration was not inherently trim-able

I am aware enough about the necessity of trim, tho I may not be sufficiently learned of the subject to comment on x-33/VentureStar.  I probably don't need to remind you that this "fundamental killer" of that and other SSTO suggestions is not a result of my proposing.

Anyhow, I was thinking that the aerospike engine was a good idea, and wondered why the implementation was not developed further.  There's been no further NASA development on this engine, right?

[DLR]

What about reusable SSTO concepts which land vertically? They only need heavy TPS around the bottom and they are aerodynamically more stable.


http://www.astronautix.com/lvs/milpress.htm

[Downix]

What about VentureStar? The X-33 prototype was canceled largely because of the failure to construct a large liquid hydrogen tank out of composite materials. But I remember reading that a standard aluminum tank would have been sufficient or perhaps even lighter than the composite tank. I think the most straightforward path to SSTO would be to revive the X-33 programme.

Not quite.  It was failure to keep the hydrogen tank from cracking, due partly to materials, partly to shape, which was the problem. 

Aluminum tanks with composite structure would have likely been much better.  (IIRC, the Rockwell proposal for X-33 did just that, and used the SSME over the Aerospike)

[Warren Platts]

TAL = Transatlantic Abort Landing; AOA = Around Once Abort

OK, fair enough; although I'm unaware of such an abort mode ever being necessary in the history manned spaceflight. However, Challenger also shows that TPS's aren't perfectly failsafe. Choose your poison. I'm just sayin' if you didn't have to worry about a TPS, it would simplify the job of designing a SSTO.

Ask the Russians, they have done it.

Done what?

[Jorge]

TAL = Transatlantic Abort Landing; AOA = Around Once Abort

OK, fair enough; although I'm unaware of such an abort mode ever being necessary in the history manned spaceflight. However, Challenger also shows that TPS's aren't perfectly failsafe. Choose your poison. I'm just sayin' if you didn't have to worry about a TPS, it would simplify the job of designing a SSTO.

Ask the Russians, they have done it.

Done what?

An abort that was the distance/re-entry heating equivalent of a TAL (on Soyuz 18-1).

[mmeijeri]

If you could refuel in LEO, that would radically simplify the design of the SSTO, since you could do a mostly propulsive re-entry, and thus skip the TPS.

How would you lift the return propellant to LEO cost-effectively? And what's the point of an SSTO that cannot return to Earth without refueling?

[sewand]

2 Plug nozzles and especially the linear form are considerable heavier than bell nozzle equivalents.  The flow field in advanced nozzles is far more complex than most of the literature would have you believe and it is not clear either aerospikes (or expansion deflection nozzles which is what we at Reaction Engines are looking at) can in real operation engines deliver atmospheric compensation, and thus justify their extra mass.

Thinking back to some of the Philip Bono plug nozzle designs - can the extra weight of a plug nozzle be offset to an extent by having it double as the re-entry TPS?  I think Bono proposed active cooling by circulating liquid hydrogen through the plug.  Is that possible in practice?

[Warren Platts]

An abort that was the distance/re-entry heating equivalent of a TAL (on Soyuz 18-1).

OK, I guess we're even.

If you could refuel in LEO, that would radically simplify the design of the SSTO, since you could do a mostly propulsive re-entry, and thus skip the TPS.

How would you lift the return propellant to LEO cost-effectively? And what's the point of an SSTO that cannot return to Earth without refueling?

It would be shipped from the Moon of course! Where else would it come from?!? The point would be shipping that most valuable cargo of all: living meat.

There are 4 problems with going small.

1 Systems development and operational cost do not drop linearly with size. A system half the size is still three quarters the cost (very roughly) 

2 Smaller systems are more sensitive to mass variation and hence, while a bit cheaper, they are more technically risky.

3 People wanting to launch small payloads tend not to have much money. Any commercial system needs to grab the bulk of the market i.e. the cargo and the propellant, in order to generate the revenue to pay for its development.

4 Given the cargo and the propellant are by far the majority of what you want to launch, if you do not introduce the cost benefits to these payloads you do not change astronautics.

This is a good explication. I wish you the best of luck. But you are going to have a rough time competing with the Moon--once it gets going--for providing low-cost bulk materials like rocket propellant to LEO. I mean, a pure-rocket SSTO from the lunar surface is a piece of cake compared to the comparable problem from the surface of the Earth. I think we can all admit that.

What John--and I--like about Skylon, is that it would make an elegant way for a human to get into space and back--you could literally go to an airport. There is a market for this. Most freight shipped by air these days consists of living humans, if I'm not mistaken.

Thus my question to you is this: seriously, what if you could refuel a Skylon in LEO? Would that make your life as an engineer any easier?

[Hempsell]

My apologies to John Fornaro the wording on my exhortation on trim reads like a person point and I had not intended it to such. I had meant to strongly emphasise a general point to the world at large that generic ideas such as discussed in thread like this and even well into design projects in major companies miss the difference between a point mass analysis (which often looks quite good) and a mass budget after a vehicle that can be trimmed (which often looks very grim).  The rule of thumb we have is do not believe any spaceplane concept until it has gone through at least 2 major configuration changes. Delta Clipper and VentureStar both got through their first major change, but if they had progressed and assuming they were viable my betting money is on the need for another.

On flying humans into space; SKYLON can do that (with a suitable module) and we have a short movie (narrated by Prince Voltan himself) showing just that on our website with details in the User’s Manual. 

Warren - The comparison with air traffic being mostly people doesn’t work because aircraft are complimented by bulk cargo ships on the sea.  A space launch system has to do both jobs and this has been the assumption behind SKYLON, it needs to do both to capture enough market to be commercial viable.

Refuelling to lower re-entry heating would take in the order of 100 tonnes of which only 80% could be lunar oxygen leaving 20% hydrogen far more than the payload capability of the SKYLON.  And how dose the massive infrastructure (orbiting propellant storage and refulling stations – never mind the large lunar base and earth moon tankers - come from to do this? It’s a chicken and egg problem and the answer is you need the SKYLON first so you have to solve the TPS problem, which is not as difficult as many think once you lower the ballistic coefficient from the Shuttle value.

Refuelling and flying the earth to orbit system out to the Moon (including a landing) was something the Delta Clipper team were very keen on, especially Max Hunter - it makes more sense in their case than SKYLON because of the vertical landing strategy. I never saw it really working, if you have a large enough infrastructure to be able to refuel launchers having specialist orbit transportation is little extra and much more efficient. In the case of SKYLON something like 1/3 the vehicle dry mass would be useless dead weight on a lunar mission, and the extra kit we would need to extend the mission while probably not heavy would be expensive.

However I think it is possible that use of Lunar materials can reduce significantly the bulk cargo traffic if you have large projects going on.  Bob Parkinson back in the 1980s calculated the breakeven point was in the few 100 tonnes a year region which is as good a number as I have seen anywhere.  But activity on this scale requires low cost cargo and human transport to initiate it and even when in place there is still massive earth to orbit cargo requirements compared to what we have now, it is just that the balance between human and automatic flights alters.

Seward – Philip Bono was a good engineer and I am sure the basics of his ballistic plug nozzles designs were credible. However they are pure rocket (I am not sure that works), they had not been trimmed (see comment above), and aerospikes are unlikely to work in the simple manner that was assumed in the 1960s.  When MacDonnell Douglas (heirs to his work) got a second go with Delta Clipper, they saw Bono as the heritage, but decided nose first was best and preferred bell engines (I was on team in Huntington Beach the initial stages of this project).

[JohnFornaro]

shipping that most valuable cargo of all...

...Cheeseburgers, buffalo wings, T-bones... I don't get it.  That's all "dead meat".  Whaddaya mean "living meat"?  Are they gonna be shipping livestock up there?

A few basics about trim:

http://ma3naido.blogspot.com/2009/12/aircraft-trim-tabs.html

Anyhow, from:

http://www.pwrengineering.com/dataresources/X-33AttitudeControlUsingTheXRS-2200LinearAerospikeEngine.pdf

Since no launch vehicle has flown with this type of engine, the interactions between the aerodynamics of the vehicle and the plumes of the engine are largely unknown. A five percent scale model of the X-33 lifting body and aerospike engine with working nozzles was used in wind tunnel tests to determine some of these effects in the subsonic regime....

Attitude control is effected by engine TVC and eight aerosurfaces, using electromechanical and electro-mechanical/pneumatic assisted actuators.

Which I would have thought included some kind of trim tab arrangement, but hey.  My understanding was that trim in this vehicle would have been accomplished by thrust vector control, which the article confirms with the conclusion:

The total attitude control system is robust to dispersions, ensuring a high probability of successful flight from liftoff to landing.

The trim issue had certainly been considered, though even this article about what turned out to be a failed program features some of my pet peeve language of false arrogance: "The controller is easily tuned to provide good transient response..."  Too bad the fuel tanks weren't easily manufactured, or that the ensuing vehicle didn't easily get into orbit.  But that's just me.

If we only get to choose one point of failure for the X-33,  I thought it would have been the fuel tank issue.  "I thought", not "I know for an absolute fact", just to reiterate.  As to the trimmable shape of the vehicle being intrinsically difficult in the evolution from the one to the other; well, ya coulda fooled me.  So I just don't get: "look at the difference between the X-33 and the final form of VentureStar" at the moment.

So I guess I'm left musing about why the aerospike engine hasn't seen further development.

Moving on to the comparison of the terrestrial air/sea : passenger/cargo relationship to its space based analog:  At first, we'll be needing to launch a lot of prop and cargo from the Earth.  Lunar ISRU for prop might be the next thing on the horizon, but it's several years down the road, and smelting aluminum and other heavy industries is even further down the road.  Right now, I see SSTO for passengers only.

And about the "market" for passenger flight.  There's no doubt a market out there:

http://www.washingtonpost.com/wp-dyn/content/article/2010/08/06/AR2010080606053.html

Of course, I could only find one documented example, so I guess we can conclude that there's no market for space tourism.

[A_M_Swallow]

{snip}
And what's the point of an SSTO that cannot return to Earth without refueling?

Such an SSTO could be used to carry cargo and propellant into space.  Like current rockets it would be thrown away after its single flight.

[Hempsell]

John – your VentureStar image is what they started with. Attached is what they ended up with before it was cancelled.

I am not sure NASA ever had a programme of work to look at aerospikes seriously, the X-33 engine was a Lockheed / Rocketdyne initiative and I think died with the X-33. 

Even if there was as separate plug nozzle programme, NASA would have stopped work when President Bush started the Moon Mars Initiative. On this presidential directive NASA closed down all work on advanced reusable launcher technologies, excepting those technologies that could be directly justified by a Moon or Mars mission using expendables.

[sandrot]

{snip}
And what's the point of an SSTO that cannot return to Earth without refueling?

Such an SSTO could be used to carry cargo and propellant into space.  Like current rockets it would be thrown away after its single flight.

Just to have then a larger mass to dispose of from LEO. Where's the appeal? Where's the lower cost?

[JohnFornaro]

Look at that will ya?  Thanks for that newer image.  That is rather different.  What's that payload thingy on the back of it? 

Somebody quick: fix the Wiki page!

[Warren Platts]

Warren - The comparison with air traffic being mostly people doesn’t work because aircraft are complimented by bulk cargo ships on the sea.  A space launch system has to do both jobs and this has been the assumption behind SKYLON, it needs to do both to capture enough market to be commercial viable.

That's a good point.

Refuelling to lower re-entry heating would take in the order of 100 tonnes of which only 80% could be lunar oxygen leaving 20% hydrogen far more than the payload capability of the SKYLON.  And how dose the massive infrastructure (orbiting propellant storage and refulling stations – never mind the large lunar base and earth moon tankers - come from to do this? It’s a chicken and egg problem and the answer is you need the SKYLON first so you have to solve the TPS problem, which is not as difficult as many think once you lower the ballistic coefficient from the Shuttle value.

However I think it is possible that use of Lunar materials can reduce significantly the bulk cargo traffic if you have large projects going on.  Bob Parkinson back in the 1980s calculated the breakeven point was in the few 100 tonnes a year region which is as good a number as I have seen anywhere.  But activity on this scale requires low cost cargo and human transport to initiate it and even when in place there is still massive earth to orbit cargo requirements compared to what we have now, it is just that the balance between human and automatic flights alters.

Actually, the Chandrayaan mini-SAR circular polarization radar recently discovered many permanently shaded craters in the north polar region of the Moon that apparently have surface layers of relatively pure water ice that are at least two meters thick. This discovery was only announced in March of this year with the publication in Geophysical Research Letters of a paper describing the results by Paul Spudis and 30 other authors.

The upshot of the discovery--if it pans out--is that getting rocket propellant from the Moon will be far easier than ever before imagined. Rather than massive strip mines required to extract a few percent oxygen, a small operation on the scale of a mom and pop local gravel pit could produce hundreds to thousands of tons of propellant--both oxidizer and LH2 fuel. Probably as few as 6 EELV cargo flights could deliver the equipment necessary to make a Moon base self-sufficient (defined as the propellant necessary for landings from L2, ascents back to L2, and return to Earth). An aggressive Moon base that had 5 or 6 flights per year would need to produce maybe 800 tons of propellant per year, and thus have to process 1,200 tons of water per year (assuming a mixture ratio of 5)--an amount of water that would fill up about 1/2 of a standard Olympic-sized swimming pool.

Combine that with a reusable SSTO tanker/lander, and the marginal cost to deliver said propellant to L2 would be on the order of a few hundred dollars per kg or less, depending on the number of reuses that could be expected.

As for the depots and landers, ULA has proposed a family of such vehicles based on their Centaur upper stage combined with 5-meter diameter fuel tanks that seems reasonably affordable.

But, yes, 100 tons to slow down a Skylon seems like a lot--especially if you have a high flight rate. To get that much to LEO would require production of at least 500 tons on the lunar surface. I can see why you would want a TPS.
 

Refuelling and flying the earth to orbit system out to the Moon (including a landing) was something the Delta Clipper team were very keen on, especially Max Hunter - it makes more sense in their case than SKYLON because of the vertical landing strategy. I never saw it really working, if you have a large enough infrastructure to be able to refuel launchers having specialist orbit transportation is little extra and much more efficient. In the case of SKYLON something like 1/3 the vehicle dry mass would be useless dead weight on a lunar mission, and the extra kit we would need to extend the mission while probably not heavy would be expensive.

I see what you're saying here. (But I can also see how a beefy vehicle capable of 10 km/sec delta v would be nice to have on a Mars mission....) Maybe it would make more sense to go the other direction--i.e., use the lunar lander SSTO for transport from lunar surface to LEO and back. If propellant was cheap enough, fully propulsive injections into LEO without aerobraking is potentially feasible.

Anyway, good luck with the Skyon project, and thanks for your detailed comments here. Hopefully, under the new regime, you guys will be able to get some NASA $$$ to help get the prototype flying.

PS I, too, find the text editor on this forum to be a real pain in the arse! I usually do what John does: use Notebook instead....

[Hempsell]

John – Payload thingy? I am not sure why they did this. Where I have seen external cannisters before, they are to save weight – a somewhat desperate measure, the idea being you junk them in orbit so the return vehicle is smaller and lighter. The problem being that you lose both payload return from orbit capability and payload recovery during all aborts. But the picture seems to indicate the canister is covered with TPS so presumably is expected to return - I am left puzzled.

Warren -  Lunar resources: I like the idea of the Moon extending down over the Earth extending up. But I return to my point that this is not first generation solutions as you need to establish a major presence using earth centred systems and I cannot see how this could be managed with existing (or indeed even Saturn 5’s) but SKYLON can definitely hack it. Once you have a working Lunar infrastructure then second generation system using Lunar propellants can be deployed and then, I agree, magic happens.

NASA $$$ - sorry no US involvement with ITAR as currently enforced. At the moment SKYLON is an ITAR free product.

[JohnFornaro]

That payload thingy would certainly alter the launch characteristics, including, uhhhh... trimmability.   I do like SKYLON, tho.  Appreciate your comments.

[JohnFornaro]

OK, some questions:

Why does the Sabre engine have that gentle curve to it?  I usually visualize a jet engine being absolutely linear.

What are "shock-on-lip conditions", and why are they worthy of special mention?

I understand the necessity for the heat exchanger to cool the inlet air, but not liquify it.  Could you just discuss the heat exchanger further, since it has special mention on your website?

If ya don't mind.

[Warren Platts]

Warren -  Lunar resources: I like the idea of the Moon extending down over the Earth extending up. But I return to my point that this is not first generation solutions as you need to establish a major presence using earth centred systems and I cannot see how this could be managed with existing (or indeed even Saturn 5’s) but SKYLON can definitely hack it.

More power to you sir! An aggressive lunar presence could be done within anticipated budget levels if launch costs for propellant can be kept down to $5,000 USD per kg. If Skylon can do much better than that, that would be awesome! What are your current projections these days for bulk cargo launch prices, and when is the earliest do you think that Skylon can fly?

Once you have a working Lunar infrastructure then second generation system using Lunar propellants can be deployed and then, I agree, magic happens.

Getting to the point where the Moon could supply surplus propellant is definately a second generation project; but it could happen within about 10 years after first landing with the proper support and focus.

NASA $$$ - sorry no US involvement with ITAR as currently enforced. At the moment SKYLON is an ITAR free product.

I thought the US and the UK were allies! Whatever happened to the special relationship? I could understand it if Reaction Engines was a Chinese operation, but it's not. Have you considered relocating your headquarters to the USA? Maybe sell some stock on the NASDAQ--that might help.

[Jorge]

NASA $$$ - sorry no US involvement with ITAR as currently enforced. At the moment SKYLON is an ITAR free product.

I thought the US and the UK were allies! Whatever happened to the special relationship?

ITAR doesn't adequately distinguish. That's one of many reasons it needs to be reformed.

[Hempsell]

OK, some questions:

Why does the Sabre engine have that gentle curve to it?  I usually visualize a jet engine being absolutely linear.

What are "shock-on-lip conditions", and why are they worthy of special mention?

I understand the necessity for the heat exchanger to cool the inlet air, but not liquify it.  Could you just discuss the heat exchanger further, since it has special mention on your website?

If ya don't mind.

Why a Curved nacelle? – the most frequently asked technical question. The answer is: the air intake on the front of the nacelle needs to point directly into the incoming airflow whereas SKYLON’s wings and body need to fly with an angle of incidence to create lift, so the intake points down by 7 degrees to account for this. The rocket thrust chambers in the back of nacelle need to point through the centre of mass of the vehicle so are angled down; again by 7 degrees but it is a coincidence the angle is the same.

The Intake is “shock-on-lip”, it means the intake swallows the shockwave from the intake cone (and the cone moves to ensure this happens) so all the air passing through the shock enters the engine nacelle, which in turn means we have more air than the engine needs for most of the supersonic flight regime. This is why we have a spill duct to dump the excess air thorough the burners, where we also dump the excess hydrogen used for cooling but not needed by the engine, and thus get a bit more thrust

Most supersonic aircraft intakes spill the air round the outside of the engine but this adds drag and means we cannot to the ramjet trick.

The Pre-cooler heat exchangers are the main technology under development at Reaction Engines and will be our hardware contribution to the vehicle. They are arrays of very fine thin walled Inconel tubes with cryogenic cooled helium gas through them, while the air flows outside the tube through the tube matrix. At Mach 5 goes from around 950 degrees C to around -140 degrees C with a total heat transfer of around 400 MW. The SABRE heat exchanger of made up of spiralling modules nested within each other like a Swiss roll helium entering from the inside and flowing outwards while the air counter flows inwards. It means the temperature difference between the air and the tube is never more than 10 degrees.

[Hempsell]

A reply to Warren

QUESTION: What are your current projections these days for bulk cargo launch prices?

From the objectives laid out in the Requirement Specification

On entry into service cost per kg = $8000 (2004)
About 80% current prices but this is a true unsubsidised cost and is less than half the true cost of current expendables.

In Mature Service (the most meaningful number) cost per kg = $1000 (2004)

Not in the Req. Spec. but has been looked at is the cost if major activity like Solar Power Satellites cost per kg below $500 (2004).

QUESTION; When is the earliest do you think that Skylon can fly?

The programme we are working to has the production prototypes doing the qualification test flight programme from 2018 and into service in 2020

[mlorrey]

{snip}
And what's the point of an SSTO that cannot return to Earth without refueling?

Such an SSTO could be used to carry cargo and propellant into space.  Like current rockets it would be thrown away after its single flight.

Expendable SSTO is automatically cheaper to build and operate than a TSTO, allowing for more frequent launches and lower launch prices.

[mlorrey]

A reply to Warren

QUESTION: What are your current projections these days for bulk cargo launch prices?

From the objectives laid out in the Requirement Specification

On entry into service cost per kg = $8000 (2004)
About 80% current prices but this is a true unsubsidised cost and is less than half the true cost of current expendables.

In Mature Service (the most meaningful number) cost per kg = $1000 (2004)

Falcon 9 at current prices is $4667 per kg using latest pricing ($49 million per launch) and max payload estimates (10,500 kg). This is a true unsubsidized cost and is the true cost of a current expendable.

F9 Heavy single payload to LEO current price is $95 million, with a max payload of 32,000 kg = $2969 per kg.

Extrapolating this into future models by SpaceX (and yes, some SWAG is involved), the Falcon X price per kg I estimate will be around $1990/kg, the Falcon X Heavy will be $1330/kg, and the Falcon XX will be around $900/kg. I am not aware if SpaceX has a Falcon XX Heavy planned, if they did, it should price at about $600/kg.

[mlorrey]

NASA $$$ - sorry no US involvement with ITAR as currently enforced. At the moment SKYLON is an ITAR free product.

I thought the US and the UK were allies! Whatever happened to the special relationship?

ITAR doesn't adequately distinguish. That's one of many reasons it needs to be reformed.

Britain uses our Trident missiles in its subs. While Canada holds an exemption to ITAR, Rep Henry Hyde stymied an exemption for Britain for years (he is dead now). The US Government signed bilateral Defense Trade Cooperation Treaties with the UK and Australian Governments in June[103] and September 2007[104] respectively. The treaties are designed to allow the exchange of technical data, defense articles and defense services (including USML items) without the need for export authorizations such as TAAs and MLAs, subject to restrictions on the type and classification of the technology in question and its intended end-use.[105] In return, the UK and Australian Governments will be required to take steps to improve their export control laws (such as the Australian Customs Act).[106] As of September 2008, the US Senate Foreign Relations Committee had deferred consideration of the Treaties on the grounds that it had not had time to properly evaluate proposed amendments to ITAR required in order to implement the Treaties.[107] (from wikipedia)

[Hempsell]

A reply to Warren

QUESTION: What are your current projections these days for bulk cargo launch prices?

From the objectives laid out in the Requirement Specification

On entry into service cost per kg = $8000 (2004)
About 80% current prices but this is a true unsubsidised cost and is less than half the true cost of current expendables.

In Mature Service (the most meaningful number) cost per kg = $1000 (2004)

Falcon 9 at current prices is $4667 per kg using latest pricing ($49 million per launch) and max payload estimates (10,500 kg). This is a true unsubsidized cost and is the true cost of a current expendable.

F9 Heavy single payload to LEO current price is $95 million, with a max payload of 32,000 kg = $2969 per kg.

Extrapolating this into future models by SpaceX (and yes, some SWAG is involved), the Falcon X price per kg I estimate will be around $1990/kg, the Falcon X Heavy will be $1330/kg, and the Falcon XX will be around $900/kg. I am not aware if SpaceX has a Falcon XX Heavy planned, if they did, it should price at about $600/kg.

If Space-ex can really meet these numbers, wonderful, more power to there proverbial elbow, they will provide good competition and I am sure we will keep each other on our proverbial toes. My fear for them is I cannot see where the “magic” is that makes them so much better than everyone else, but that will be become clear as their programme progresses.

But it will not be un-subsidised they have had very substantial funding from NASA which will not be paid back.

[93143]

Expendable SSTO is automatically cheaper to build and operate than a TSTO, allowing for more frequent launches and lower launch prices.

An expendable SSTO has to be a lot bigger than an expendable TSTO with the same payload.  It's perfectly feasible with 1960s rocket technology; most likely no one's tried it because it's not worth it.

[JohnFornaro]

Mark, Thanks up there.

[mlorrey]

Expendable SSTO is automatically cheaper to build and operate than a TSTO, allowing for more frequent launches and lower launch prices.

An expendable SSTO has to be a lot bigger than an expendable TSTO with the same payload.  It's perfectly feasible with 1960s rocket technology; most likely no one's tried it because it's not worth it.

Actually, I think the major road bump is to do with ballistic rocket folks being somewhat separate from the hypersonic air breathing folks. I am 99% sure that if SpaceX were to experiment with a Falcon 1e first stage with an air ram ejector around the Merlin engine, running a bit more fuel rich than usual (it normally runs fuel rich), that this would be a feasible SSTO expendable with nearly identical payload as the Falcon 1 TSTO. Without the expense of a second stage, this would make the cost to orbit drop at least 25-30%, and enable a huge improvement in responsiveness in rapid launch.

[DLR]

A reply to Warren

QUESTION: What are your current projections these days for bulk cargo launch prices?

From the objectives laid out in the Requirement Specification

On entry into service cost per kg = $8000 (2004)
About 80% current prices but this is a true unsubsidised cost and is less than half the true cost of current expendables.

In Mature Service (the most meaningful number) cost per kg = $1000 (2004)

Falcon 9 at current prices is $4667 per kg using latest pricing ($49 million per launch) and max payload estimates (10,500 kg). This is a true unsubsidized cost and is the true cost of a current expendable.

F9 Heavy single payload to LEO current price is $95 million, with a max payload of 32,000 kg = $2969 per kg.

Extrapolating this into future models by SpaceX (and yes, some SWAG is involved), the Falcon X price per kg I estimate will be around $1990/kg, the Falcon X Heavy will be $1330/kg, and the Falcon XX will be around $900/kg. I am not aware if SpaceX has a Falcon XX Heavy planned, if they did, it should price at about $600/kg.

Zenit is even better than Falcon 9. On Zenit, it's about 3000 to 4000 $ per kg.

[Jim Davis]

I am 99% sure that if SpaceX were to experiment with a Falcon 1e first stage with an air ram ejector around the Merlin engine, running a bit more fuel rich than usual (it normally runs fuel rich), that this would be a feasible SSTO expendable with nearly identical payload as the Falcon 1 TSTO.

Do you have any numbers to share with us? What do you base your confidence on?

[mlorrey]

I am 99% sure that if SpaceX were to experiment with a Falcon 1e first stage with an air ram ejector around the Merlin engine, running a bit more fuel rich than usual (it normally runs fuel rich), that this would be a feasible SSTO expendable with nearly identical payload as the Falcon 1 TSTO.

Do you have any numbers to share with us? What do you base your confidence on?

Falcon 1e first stage:
Dry mass 5680 kg
Fuel mass 87000kg
payload mass 500 kg
Base Merlin Isp: 304 sec
Ram ejector average Isp: 1200 sec (mach 0.5-mach 8.5)
The degree the ram ejector boosts total average isp is related to the thrust of the ejector vs the thrust of the Merlin itself.
For any combined average flight Isp of 375 seconds or greater:
delta-v 9938 m/s

You should be able to reduce the dry mass a bit since you don't need the same amount of structural support that the TSTO needs to support the mass of the second stage.

[gospacex]

{snip}
And what's the point of an SSTO that cannot return to Earth without refueling?

Such an SSTO could be used to carry cargo and propellant into space.  Like current rockets it would be thrown away after its single flight.

Expendable SSTO is automatically cheaper to build and operate than a TSTO, allowing for more frequent launches and lower launch prices.

Why? SSTO is *simpler* that TSTO, but with the same GLOW it also puts less mass into the orbit.

Therefore "$ *per kg*" metric need not automatically be better.

IIRC there are 1st stages (Titan II?) which can be theoretically used as SSTOs, but they would have near-zero payload fractions.

[mlorrey]

{snip}
And what's the point of an SSTO that cannot return to Earth without refueling?

Such an SSTO could be used to carry cargo and propellant into space.  Like current rockets it would be thrown away after its single flight.

Expendable SSTO is automatically cheaper to build and operate than a TSTO, allowing for more frequent launches and lower launch prices.

Why? SSTO is *simpler* that TSTO, but with the same GLOW it also puts less mass into the orbit.

Therefore "$ *per kg*" metric need not automatically be better.

IIRC there are 1st stages (Titan II?) which can be theoretically used as SSTOs, but they would have near-zero payload fractions.

Actually, Titan II was a valid 1.5STO with a substantial payload fraction. It dropped two of its three 1st stage motors after 50% of fuel was consumed.

The cost of launching a TSTO is not simply 1+2=3. It is 1 plus 2 plus the time and expense of combining 1 and 2, plus the risk that 2 won't ignite and 3 will be lost. So it comes out to something like 1+2=4.3.... whereas 1 = 1....

Since you can produce and launch 1's all day long with a much higher chance each 1 will get to orbit, then even if your payload to orbit is less than half that of a TSTO, two or more SSTO's will still put the same amount of gross payload in orbit for less cost per kg.

[gospacex]

Actually, Titan II was a valid 1.5STO with a substantial payload fraction. It dropped two of its three 1st stage motors after 50% of fuel was consumed.

Titan-II had two engines on its 1st stage and it did not drop them. Did you mean "Atlas"?

[aero]

Since you can produce and launch 1's all day long with a much higher chance each 1 will get to orbit, then even if your payload to orbit is less than half that of a TSTO, two or more SSTO's will still put the same amount of gross payload in orbit for less cost per kg.

You guys tickle me to death. You write as though payload has infinite divisibility and one-half plus one-half is always one. Its especially comical when you compare the efficiency of projected future  heavy lift boosters to current vehicles. If you're launching fuel depots, then sure, you can count on a full load, but if you're launching a geosynchronous communications satellite, it masses what it masses, and using a big, efficient heavy lift rocket isn't necessarily the cheapest way to go. I guess my point is illustrated by the above quote, simply by noting that the TSTO can just barely orbit the astronaut, but it is easy to get him to orbit with two launches.  Only who will decide which half to launch first? Or is it, "Hold your breath and don't worry, your life support is already in orbit."

When talking launch vehicles, you can not ignore the mission.

[KelvinZero]

Could a single stage to orbit really be cost effective if the mission was a fuel depot? That would be fantastic.

[mlorrey]

Actually, Titan II was a valid 1.5STO with a substantial payload fraction. It dropped two of its three 1st stage motors after 50% of fuel was consumed.

Titan-II had two engines on its 1st stage and it did not drop them. Did you mean "Atlas"?

Ah yes, sorry I misspoke...

[Jim Davis]

Mike, thanks for the numbers.

Falcon 1e first stage:
Dry mass 5680 kg
Fuel mass 87000kg
payload mass 500 kg

The Falcon planners guide gives those numbers in pounds although that shouldn't affect your argument since we're concerned with ratios.

I do note that you seem to be suggesting that the added mass of the ejector ram jet assembly will be negligible. That is completely unrealistic. But let's assume for sake of argument that it can be done.

Ram ejector average Isp: 1200 sec (mach 0.5-mach 8.5)

This is also completely unrealistic. A hydrocarbon fueled ramjet could only make 1200 s over a small part of that speed range. Even that assumes very good inlet efficiencies which are hardly likely to be obtained from being added to an existing rocket stage that made no such provision for it. But let's assume for sake of argument that it can be done.

The degree the ram ejector boosts total average isp is related to the thrust of the ejector vs the thrust of the Merlin itself.

Indeed. Let us assume that half the thrust comes from the ejector ramjet which would be quite generous. The effective Isp while the ejector ramjet is operating becomes:

Ieff = 1/(0.5/304 + 0.5/1200) = 485 s

For any combined average flight Isp of 375 seconds or greater:
delta-v 9938 m/s

That's probably a generous effective delta v for a vehicle which has to spend so much time in the lower atmosphere to provide air for the ejector ramjet. But let's assume for sake of argument that it can be done.

Now let's see what the effective Isp is when the ejector ramjet is operating over 1/3 of the delta V range.

Ieff = 3/(1/485 +2/304) = 347 s

So even with generous assumptions you don't reach orbit.

You should be able to reduce the dry mass a bit since you don't need the same amount of structural support that the TSTO needs to support the mass of the second stage.

But on the other hand you're dealing with far greater aerodynamic loads that you normally would be and incomparably greater aerodynamic heating loads. I wouldn't count on reducing dry mass.

There are reasons why airbreathing schemes have yet to make it to the launch pad (or runway).

[mlorrey]

Since you can produce and launch 1's all day long with a much higher chance each 1 will get to orbit, then even if your payload to orbit is less than half that of a TSTO, two or more SSTO's will still put the same amount of gross payload in orbit for less cost per kg.

You guys tickle me to death. You write as though payload has infinite divisibility and one-half plus one-half is always one. Its especially comical when you compare the efficiency of projected future  heavy lift boosters to current vehicles. If you're launching fuel depots, then sure, you can count on a full load, but if you're launching a geosynchronous communications satellite, it masses what it masses, and using a big, efficient heavy lift rocket isn't necessarily the cheapest way to go. I guess my point is illustrated by the above quote, simply by noting that the TSTO can just barely orbit the astronaut, but it is easy to get him to orbit with two launches.  Only who will decide which half to launch first? Or is it, "Hold your breath and don't worry, your life support is already in orbit."

When talking launch vehicles, you can not ignore the mission.

Of course, I never said you didn't have to. Didn't John Glenn get into orbit with his life support on one launch of an Atlas 1.5STO? Of course he did.

Now lets examine the Falcon 9 first stage as a potential expendable SSTO with a ram air ejector cowling.

Gross mass: 270 tonnes
Dry mass: 14.73 tonnes
payload: 5 tonnes
Main engine Isp: 304 sec
Ram air ejector Isp: 1200 sec
For any flight average Isp above 375 seconds, delta-v is 9621 m/s, with dV increasing as avg Isp rises.

Furthermore, should the first stage thrust assembly be reconfigured to allow the outer four engines to detach and be recovered with the ram ejector cowling, then payload mass to orbit rises on the order of 4-5 tonnes, nearly the full payload of the TSTO Falcon 9 rocket. Certainly enough to put a Dragon capsule into orbit.

[RanulfC]

Mike Lorry:

Jim Davis had/has a point: The ram air ejector cowling would only generate the given ISP boost for a very narrow amount of time with most operation being off-design and FAR less effective for the majority of it's operation.

It would "seem" to be a better option to simply design the to stage half the engines later in the flight to increase payload capacity.

Having noted that, Jim Davis:
You're assuming a greater heating and aerodynamic load for an ejector-ramjet cowl because it is an "air-breather"? Because of a more depressed trajectory?

This usually isn't the case with ejector-cowlings. While they impose some design loads the trajectory isn't effected because they are assumed to be used only for a short time at relativly low altitude, usually below 80,000ft. (At least in the majority of studies I've seen)
Like the boost-assist SRMs used on many expendable launchers today the ejector cowling would boost the ISP and thrust of a launch vehicle when it's deep in the atmosphere and moving relativly slowly and would be staged relativly soon after launch.

Randy

[Jim Davis]

Having noted that, Jim Davis:
You're assuming a greater heating and aerodynamic load for an ejector-ramjet cowl because it is an "air-breather"? Because of a more depressed trajectory?

Mike's scenario involved the ejector ramjet to provide substantial thrust augmentation at high Isp over the velocity range of M = 0.5 to 8.5. The dynamic pressures necessary would require a more depressed trajectory.

So the answer is yes.

Like the boost-assist SRMs used on many expendable launchers today the ejector cowling would boost the ISP and thrust of a launch vehicle when it's deep in the atmosphere and moving relativly slowly and would be staged relativly soon after launch.

Mike was very much more ambitious than that.

[RanulfC]

Jim Davis wrote:
Mike was very much more ambitious than that.

Which is why we both questioned his assumptions

I'm not making QUITE that much of a leap, only correcting for the information that "I" know

I've noted that there are quite a number of "assumptions" on the operating parameters for air-breathing propulsion as both Launch Assist and Orbital Launch that are based on contextual and situational criteria that don't actually apply in every single case.

Just an FYI.

Randy

[mlorrey]

Having noted that, Jim Davis:
You're assuming a greater heating and aerodynamic load for an ejector-ramjet cowl because it is an "air-breather"? Because of a more depressed trajectory?

Mike's scenario involved the ejector ramjet to provide substantial thrust augmentation at high Isp over the velocity range of M = 0.5 to 8.5. The dynamic pressures necessary would require a more depressed trajectory.

So the answer is yes.

Like the boost-assist SRMs used on many expendable launchers today the ejector cowling would boost the ISP and thrust of a launch vehicle when it's deep in the atmosphere and moving relativly slowly and would be staged relativly soon after launch.

Mike was very much more ambitious than that.

The Falcon 1 first stage already separates from its second stage at mach 8.5 at 130,000 ft. An F1 based SSTO would accelerate faster than this (lower GLOW and higher thrust) and thus achieve this speed at a lower altitude in the same trajectory, so I don't see Jims flatter trajectory claim as valid. Furthermore, given the lack of second stage mass, the structure of the first stage would be capable of handling a significantly higher max Q pressure.

[Jim Davis]

An F1 based SSTO would accelerate faster than this (lower GLOW and higher thrust) and thus achieve this speed at a lower altitude in the same trajectory, so I don't see Jims flatter trajectory claim as valid.

If it's accelerating faster it will not be following the same trajectory. It still needs a certain vertical and a certain horizontal velocity component to achieve orbit. If it is accelerating faster it will gain the vertical component faster and will have to make its turn to the horizontal sooner, i.e. at a lower altitude.

But even setting this aside your scheme requires much thrust augmentation from the ejector ramjet from M=0.5 to M=8.5. This requires following a much higher dynamic pressure profile, i.e. lower profile.

Furthermore, given the lack of second stage mass, the structure of the first stage would be capable of handling a significantly higher max Q pressure.

Not at all obvious. True, it doesn't have the second stage mass but it will still have the propellant mass needed for ascent to orbit. But even setting that aside there is the much higher thermal loads to consider.

There are reasons why these ideas, which have been around a very long time, never make it to the launch pad.

[rusty]

(reconfigured for clarity)
I dont know how many times I have seen this discussed only to fall due to the same basic problems.
Let me lay out the things that any SSTO, let alone RESUSABLE SSTO designers will need to overcome in order to make this idea reasonable, let alone viable:

1. COST COST COST: If its not cost effective, nay, if its not COMMERCIALLY economic (i.e. as in for a commercial launch provider) then it won't work. Note: Don't expect government funding for this. You might get it (DOD side at least), but might isnt enough. Design it to be commercially feasible.
6. KG/$ to LEO economics: How much can it lift? Can it compete?

2. Reusability: Determine a low cost and effective system to protect the stage during rentry such that refurbishing for reuse is quick and easy.
3. Retriveal: Where are you going to land it and what are the consequences of landing there?
5. Saftey: Make it safe

4. Feasibility: Is Reusability really worth it? Or is it too expensive??

A very good list - needless to say that at Reaction Engines we believe that SKYLON ticks all those boxes.

There is a lot to like about SKYLON, but wouldn't a smaller HTHL vehicle be more robust and economically viable? A payload of 1,500-2,500kg is all that's needed for small payloads, experiments or crew rotation.

[93143]

wouldn't a smaller HTHL vehicle be more robust and economically viable? A payload of 1,500-2,500kg is all that's needed for small payloads, experiments or crew rotation.

This has been answered already.  A smaller vehicle is more developmentally risky, not less, because the required mass fraction is harder to get (which also leaves the final product with smaller margins, if it works at all).  Also, there are issues with scaling down the SABRE engine, which is what makes the whole thing work.  So no, it wouldn't be more robust.

The development cost doesn't decrease in direct proportion to the size, and neither do the manufacturing or operational costs, so the value-for-money is less, and a smaller vehicle would be shut out of markets for hardware above its capacity, meaning less business to offset the considerable development costs.  So it wouldn't be more economically viable either.

There's a reason the design iterations are actually making the vehicle bigger...

[mlorrey]

An F1 based SSTO would accelerate faster than this (lower GLOW and higher thrust) and thus achieve this speed at a lower altitude in the same trajectory, so I don't see Jims flatter trajectory claim as valid.

If it's accelerating faster it will not be following the same trajectory. It still needs a certain vertical and a certain horizontal velocity component to achieve orbit. If it is accelerating faster it will gain the vertical component faster and will have to make its turn to the horizontal sooner, i.e. at a lower altitude.

But even setting this aside your scheme requires much thrust augmentation from the ejector ramjet from M=0.5 to M=8.5. This requires following a much higher dynamic pressure profile, i.e. lower profile.

Again, you don't seem to be reading anything I am saying. 130,000 ft altitude is perfectly valid flight envelope for ram/scram air combustion at mach 8.5, therefore, since the Falcon 1 first stage separates at that altitude and speed, it already follows a proper trajectory for optimum use of a ram ejector. There is no need for a lower profile.

Furthermore, given the lack of second stage mass, the structure of the first stage would be capable of handling a significantly higher max Q pressure.

Not at all obvious. True, it doesn't have the second stage mass but it will still have the propellant mass needed for ascent to orbit. But even setting that aside there is the much higher thermal loads to consider.

Not at all, since we are saving so much mass by combusting atmospheric oxygen, all you have is the orbital ascent propellant mass and that mass is sitting at the BOTTOM of the propellant tanks, not above them in the structure as the second stage would.

Secondly, as I have previously shown that the current flight profile is perfectly valid for optimum use of the ram ejector, there will not be much higher thermal loads to consider.

There are reasons why these ideas, which have been around a very long time, never make it to the launch pad.

Actually, the GNOM concept was tested on a scale prototype and proved such a large performance increase that it could throw the same payload as a US missile 50% more massive.  That was the ONLY time that this concept has been actually tested, the US has never tested the idea.

[aero]

For those of us not familiar with the GNOM project, here is an informative link:http://www.astronautix.com/lvs/gnom.htm
It seems there is a lot to be gained by air augmentation.

[mmeijeri]

The development cost doesn't decrease in direct proportion to the size, and neither do the manufacturing or operational costs, so the value-for-money is less, and a smaller vehicle would be shut out of markets for hardware above its capacity, meaning less business to offset the considerable development costs.  So it wouldn't be more economically viable either.

Surely this is not a general principle? You are not suggesting that the participants in the Northrop Grumman Lunar Lander Challenge could just as easily have developed vehicles five times their current sizes? Or that developing a small business jet takes about as much money as a 787?

[93143]

The development cost doesn't decrease in direct proportion to the size, and neither do the manufacturing or operational costs, so the value-for-money is less, and a smaller vehicle would be shut out of markets for hardware above its capacity, meaning less business to offset the considerable development costs.  So it wouldn't be more economically viable either.

Surely this is not a general principle?

No, it applies specifically when considering SSTO.  Very specifically when considering Skylon, but I believe it applies to every SSTO concept that actually has a chance of working.

You are not suggesting that the participants in the Northrop Grumman Lunar Lander Challenge could just as easily have developed vehicles five times their current sizes?  Or that developing a small business jet takes about as much money as a 787?

Straw man.  I didn't say development cost was unrelated to size, just that the relationship isn't 1:1 linear.  Pay attention.

[mmeijeri]

No, it applies specifically when considering SSTO.  Very specifically when considering Skylon, but I believe it applies to every SSTO concept that actually has a chance of working, if the goal is to get costs down.

Can you say more? I can see that not being able to lift current commercial payloads doesn't help, but if you get incremental costs low enough you can tap new markets. It's a multidimensional thing.

Straw man.  I didn't say development cost was unrelated to size, just that the relationship isn't 1:1 linear.  Pay attention.

I didn't say anything about linear. You appeared to me to be implying there wasn't a strong monotonicity, which struck me as odd.

[93143]

No, it applies specifically when considering SSTO.  Very specifically when considering Skylon, but I believe it applies to every SSTO concept that actually has a chance of working, if the goal is to get costs down.

Can you say more? I can see that not being able to lift current commercial payloads doesn't help, but if you get incremental costs low enough you can tap new markets. It's a multidimensional thing.

An SSTO needs to be very careful with its mass fraction.  Basic geometry and physics says it's easier to achieve a good mass fraction with a larger vehicle.  There's probably a minimum size below which SSTO isn't feasible at all.  Skylon doesn't need as aggressive a mass fraction as an all-rocket vehicle would, but it has extra difficulties because the engines can't be scaled down easily.

Reaction Engines seems to have done the analysis, and apparently the prospects for new markets in the low-mass payload range aren't good enough to justify passing up the big-ticket satellite market and trying to shrink their vehicle enough to be able to take smallsats on dedicated flights.

As it is, there's no reason you couldn't piggyback smallsats on a not-quite-maxed Skylon launch - the lab next door to mine launches their stuff that way all the time.  Granted, Canada seems to be getting tired of doing everything this way; hence the CSLV project...

Straw man.  I didn't say development cost was unrelated to size, just that the relationship isn't 1:1 linear.  Pay attention.

I didn't say anything about linear. You appeared to me to be implying there wasn't a strong monotonicity, which struck me as odd.

It's cheaper to develop a small vehicle, but not by as much as the reduction in size.  This is a general principle, and doesn't take into account the additional difficulties in scaling down an SSTO.  Cut the size in half, and you might cut development costs by 40%.  Go to 1/10, and you might save 80%.  It's similar for manufacturing and operations - you save, but not enough to make up for the loss of capacity, so cost per kg goes up.

[mmeijeri]

An SSTO needs to be very careful with its mass fraction.  Basic geometry and physics says it's easier to achieve a good mass fraction with a larger vehicle.

Do you mean because of cube-square effects?

There's probably a minimum size below which SSTO isn't feasible at all.  Skylon doesn't need as aggressive a mass fraction as an all-rocket vehicle would, but it has extra difficulties because the engines can't be scaled down easily.

What is it about the engines that precludes scaling them down easily? I thought that engines typically scaled up badly, twice as large being more than twice as difficult.

Reaction Engines seems to have done the analysis, and apparently the prospects for new markets in the low-mass payload range aren't good enough to justify passing up the big-ticket satellite market and trying to shrink their vehicle enough to be able to take smallsats on dedicated flights.

That may be true, but I was thinking of manned suborbital hops. More than one way to skin a cat and many more plausible ways that won't work out - if only we knew in advance which was which.

It's cheaper to develop a small vehicle, but not by as much as the reduction in size.  This is a general principle, and doesn't take into account the additional difficulties in scaling down an SSTO.  Cut the size in half, and you might cut development costs by 40%.  Go to 1/10, and you might save 80%.  It's similar for manufacturing and operations - you save, but not enough to make up for the loss of capacity, so cost per kg goes up.

I already mentioned engines which I believe contradict this rule. But I can believe there may be systems for which this is true. Even so cost/kg isn't the only variable. ROI demands by investors may be a tougher constraint to live with. And you don't have to achieve a reduction in price by an order of magnitude all at once, merely enough to build up enough market share. I think the suborbital RLV people are being very sensible in how they go about their work which they ultimately want to lead to orbit.

[Jim Davis]

Again, you don't seem to be reading anything I am saying. 130,000 ft altitude is perfectly valid flight envelope for ram/scram air combustion at mach 8.5, therefore, since the Falcon 1 first stage separates at that altitude and speed, it already follows a proper trajectory for optimum use of a ram ejector. There is no need for a lower profile.

You don't get it, Mike. Mach 8.5 at 130,000 ft equates to a dynamic pressure of about 350 psf. Yes, a ramjet can operate at this dynamic pressure but in order to boost the performance of the vehicle by a significant amount it has to develop thrust at least comparable to the existing rocket engine. This is a very demanding requirement and demands operation at much higher dynamic pressures, typically between 1000 and 2000 psf.

Not at all, since we are saving so much mass by combusting atmospheric oxygen, all you have is the orbital ascent propellant mass and that mass is sitting at the BOTTOM of the propellant tanks, not above them in the structure as the second stage would.

Mike, your notional vehicle is going to be following a high dynamic pressure profile from M=0.5 to M=8.5. The Falcon 1-e first stage has about ten times the mass of the second stage. Your notional vehicle will have a mass of about 39,000 lbs at M=8.5. The standard Falcon 1-e will have a mass of about 16,000 lbs. I do not think it immediately obvious that a twice as massive vehicle following a much higher dynamic pressure profile is going to be subjected to lower stresses. It is certainly going to be subjected to much higher thermal loads regardless of how the loads are distributed.

Secondly, as I have previously shown that the current flight profile is perfectly valid for optimum use of the ram ejector, there will not be much higher thermal loads to consider.

Sorry, Mike you're just not getting it. The ejector ramjet is useless if it doesn't contribute the bulk of the thrust. It can't do that while flying the same low dynamic pressure profile as a pure rocket. A very efficient ramjet which only contributes a minute amount of thrust will have negligible impact on vehicle performance. You might as well add a Hall thruster to the Falcon 1-e.

Actually, the GNOM concept was tested on a scale prototype and proved such a large performance increase that it could throw the same payload as a US missile 50% more massive.  That was the ONLY time that this concept has been actually tested, the US has never tested the idea.

But note that GNOM was a 3 stage concept - nowhere near the single stage performance that you claim.

You're claiming that a negligible mass ejector ramjet can built which will develop high thrust at low dynamic pressures and high specific impulse over a very large Mach number range. Sorry, you haven't made the case for any of this.

[93143]

Do you mean because of cube-square effects?

Pretty much.  They don't hit everything equally, but they are a factor.

What is it about the engines that precludes scaling them down easily? I thought that engines typically scaled up badly, twice as large being more than twice as difficult.

I don't know exactly, but Hempsell has mentioned this more than once, and it's in their FAQ.  Whatever it is, it's specific to SABRE.

...How much did SpaceX spend on the Merlin 1?  They seem to think they can do an engine with about 15 times the thrust for $1B.

I was thinking of manned suborbital hops.

Not worth it.  SABRE is an SSTO engine.  Suborbitals don't need that kind of technology, so it would be underutilized and overpriced for that market.

I can believe there may be systems for which this is true. Even so cost/kg isn't the only variable. ROI demands by investors may be a tougher constraint to live with. And you don't have to achieve a reduction in price by an order of magnitude all at once, merely enough to build up enough market share. I think the suborbital RLV people are being very sensible in how they go about their work which they ultimately want to lead to orbit.

Actually, you do pretty much have to achieve an order of magnitude all at once.  Otherwise SpaceX will eat your lunch...  What do you suppose a Falcon 9 Heavy will cost by the time Skylon flies?

[marginally-relevant rant about TSTO RLV vs. SSTO RLV deleted]

You may argue that Reaction Engines should be concentrating on simpler technology and trying to build up market share before tackling the big one.  I say that would just distract them.  The technological commonality wouldn't be high enough to justify it in an Ares I/Ares V sense, and it's not like these guys need the engineering experience.  Besides, there's no way operating a suborbital vehicle in this soon-to-be-highly-competitive market would net them a significant fraction of the 12 billion dollars they need to bring Skylon to market in any reasonable timeframe...

[RanulfC]

Seward – Philip Bono was a good engineer and I am sure the basics of his ballistic plug nozzles designs were credible. However they are pure rocket (I am not sure that works), they had not been trimmed (see comment above), and aerospikes are unlikely to work in the simple manner that was assumed in the 1960s.  When MacDonnell Douglas (heirs to his work) got a second go with Delta Clipper, they saw Bono as the heritage, but decided nose first was best and preferred bell engines (I was on team in Huntington Beach the initial stages of this project).

Couple of points on plug-nozzle engines;

Aerospikes work, they work well and they have been extenisivly tested using various fuel combinations and nozzle types. There was some question of the "round" (annular) design having thrust fall offs between Mach3 and Mach5 but that hasn't appeared to be an actual issue outside a wind-tunnel. The "flat" (linear) aerospike engine worked as advertised there was an issue however in that instead of planning on thrust-vector control it was decided that actual actuators would be used to move the entire engine as per a normal engine gimbeling system. (Why? Not a clue on where that came from)

This greatly increased the already heavy engine structure, which was critical on the X-33 design as it made the tail structure too heavy to balance properly on glide back.

Mr. Hempsell; On the Delta Clipper design, the "nose-first" entry wasn't a "design" decision that was a requirement to meet the Air Force specified cross-range ability so there was no reason to use a plug-nozzle even though it would have been easier all around to use for both engine and heat-shield. The Air Force had done a lot of the original work on plug-nozzle engines as both propulsion and heat-shielding and a purely ballistic "Delta-Clipper" would have met the requirements except for the stated cross-range ability.

Since neither DARPA nor the Air Force had actually intended to pursue the Delta Clipper to a flight testing stage there wasn't even a need for the stated cross-range requirement but it's interesting that General Dynamics had already submited a study of an SSTO ballistic vehicle using a plug-nozzle engine/heat-shield combination requested by DARPA and the Air Force in conjunction with NASA. Alternative engine lay-outs and types were studied but the combination engine/heat-shield was by far the suprior design in terms of payload to orbit.

Randy

[mmeijeri]

Question for Hempsell (or anyone else who might know): could Skylon be profitably combined (or even at all) with mass injection precooling?

[RanulfC]

Question for Hempsell (or anyone else who might know): could Skylon be profitably combined (or even at all) with mass injection precooling?

The concept is already "deep-cooling" the incoming air so I don't see precooling as working for the engine.

Randy

[mmeijeri]

Couldn't it allow higher mach numbers in air breathing mode, or maybe less sophisticated heat exchangers?

[RanulfC]

mmeijeri wrote:
Couldn't it allow higher mach numbers in air breathing mode, or maybe less sophisticated heat exchangers?

MIPCC (Mass Injection Pre-Compressor Cooling) is mainly used to extend the Mach envelope of a standard turbo-fan/jet engine by allowing the turbo-machinery to operate closer to normal temperatures even when at high Mach and altitude.

The injected water cools and increase the air density while also increasing the mass flow, while the injected LOX stabilizes the combustion in both the combustion chamber and the afterburner. The SABRE engine already densifies the air by deep-cooling the incoming air and it doesn't use the same temperature limited turbomachinery that regular jet engines do so it never reaches it temperature limit. In order for water injection to work at all the MIPCC duct has to be long enough to allow the water to vaporize, but with the deep-cooling of the SABRE engine the water would re-condense out of the air stream and turn to ice which would be VERY BAD on the mechanicals of the engine system.
(Note that one of the mentioned "show-stoppers" was issues with icing of the heat exchangers )

SABRE (and SKYLON) wouldn't get any benifit from using MIPCC so the extra equipment and such would be dead-weight.
Here is a High Speed Propulsion Cycles report:
http://ftp.rta.nato.int/public//PubFullText/RTO/EN/RTO-EN-AVT-150///EN-AVT-150-02.pdf

This one shows the majority of the mentioned types of propulsion and their good points and bad with a focus on MIPCC. The majority of the cycles shown DO NOT WORK WELL TOGETHER with only a few exceptions because of the mechanics of the cycles themselves. (About the only ones are MIPCC and the Turbine Engine With Rocket Augmentation cycles. This since the MIPCC does a good job fooling the engine into thinking it's operating at Mach-2 and 50Kft when it's actualy Mach-4 at 80Kft, and TEwRA basiclly turns the After-Burner into a crude rocket engine from there.)

Now... Were this thread opened up a bit more to allow 1.5 or TSTO ideas I'd be pushing my latest line of research, the Spacejet concept.
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19810020560_1981020560.pdf
(There's a low speed subsonic study someplace too but it gives the idea)

Add MIPCC to the turbojet packages and TEwRA and you can release the Spacejet at higher Mach and altitude for a more effcient spaceplane.
But that ain't this thread )

Here is some other high speed air-breathing and such reports to take a look at:

Innovative Airbreathing Propulsion Concepts for Access to Space Report, 2001
http://gltrs.grc.nasa.gov/reports/2001/TM-2001-210564.pdf

Populsion Systems For Hypersonic Flight:
http://ftp.rta.nato.int/public//PubFullText/RTO/EN/RTO-EN-AVT-116///EN-AVT-116-09.pdf

Here is a LAPCAT (Mach-5 cruiser version of SKYLON, sort of) report:
http://ftp.rta.nato.int/public//PubFullText/RTO/EN/RTO-EN-AVT-150///EN-AVT-150-12.pdf

Paper on Advanced Rocket Motors:
http://ftp.rta.nato.int/public//PubFullText/RTO/EN/RTO-EN-AVT-150///EN-AVT-150-06.pdf

And a download site for a series of breifings and reports given at NATO on high speed flight and propulsion:
http://www.rta.nato.int/pubs/rdp.asp?RDP=RTO-EN-AVT-150

Randy

[lkm]

So why not Lox injection? Couldn't you inject Lox from Mach 5 to 8 with greater efficiency than the SABRE rocket mode? It doesn't introduce new tankage, it won't freeze in the heat exchanger, what's wrong with it?

[mmeijeri]

Thanks for the links. I had read a number of them already and I don't really understand your explanation. I thought that Skylon was limited at M5.5 by a compressor temperature limit and I was wondering if MIPCC couldn't help with that. Water would obviously not work (unless you precool less deeply, which might require less advanced heat exchangers), but LN2 or LOX could. I have no good feel how much this would help as I'm not sure what the limiting factor for Mach number is: skin heating, drag losses, air-breathing/rocket Isp or T/W.

[RanulfC]

lkm wrote:
So why not Lox injection? Couldn't you inject Lox from Mach 5 to 8 with greater efficiency than the SABRE rocket mode? It doesn't introduce new tankage, it won't freeze in the heat exchanger, what's wrong with it?

Actually would it be more 'efficient' than the rocket mode? By Mach-5 your aeroheating on the hull itself is becoming an issue so going straight to rockets and climbing out of the atmosphere would probably be better.

(And wouldn't it add tankage? The LOX you use would no longer be available for use in the rocket and you're haveing to add to densify and augment the incoming over-heated air? I'd think it would be less effective. YMMV )

mmeijeri wrote:
Thanks for the links. I had read a number of them already and I don't really understand your explanation. I thought that Skylon was limited at M5.5 by a compressor temperature limit and I was wondering if MIPCC couldn't help with that.

That may be my bad then because I was under the impression Mach-5+ was getting to hot for the airframe for extended aeroheating.

Water would obviously not work (unless you precool less deeply, which might require less advanced heat exchangers), but LN2 or LOX could. I have no good feel how much this would help as I'm not sure what the limiting factor for Mach number is: skin heating, drag losses, air-breathing/rocket Isp or T/W.

I'll look around but I did find a paper on LN2 addition to UAV engines for added thrust. The issue with less pre-cooling (and simpler heat exchangers) runs into (as I understand it, again YMMV ) is that less deeper cooling effects your overall engine efficincy, which effects your overall payload to orbit. The added mass would certainly help the thrust, but it reduces your overall ISP by about the same amount that it helps the thrust. One study I've seen shows a linear progression, added thrust from mass injection equals the overall loss in ISP over the studied range.

You could probably get away with it easier if you DON'T go SSTO but that need for single-stage tends to make ISP over the whole envelope pretty damn critical.

Randy

[lkm]

For SABRE the energy recovered from the precooling drives the compressor, less precooling means less pressure in the rocket engine. The engine is limited by the temperature limit of 1350 K of the compressor and helium cycle. If Lox was bled in at the shock cone to keep to that limit it should require less Lox than the pure rocket mode requires up to around mach 8, I think. As to the Aeroshell I don't know, doesn't  it already face greater temperatures from re-entry?

[Hempsell]

RanulfC  “Aerospikes work?”
Well yes. They work well on static test stands but once on flying vehicle complications arise like base drag and the interaction of the two supersonic flows they get more complicated. That is not to say they cannot be made to work it is just a lot more complex than simple theory suggests. There are similar complexities with Expansion/Defection nozzles and we may yet still use these, so it is not that we are anti-advanced nozzles .

RanulfC  “On the Delta Clipper design, the "nose-first" entry wasn't a "design" decision that was a requirement to meet the Air Force specified cross-range ability”
I agree it was not driven by the UASF requirements my point was that the team most motivated to follow Bono’s design decided it was more difficult to do than the nose first approach. (Again I would point out I was in Huntingdon Beach working with McDAC working on this programme and saw these decision being made). I am sure “Aerospike as a heat shield” might be made to work but, once the realities have been engineered in, it might not be as attractive as Bono initially thought.

LOX injection: This was actually used on HOTOL as the frost control system (there is a UK patent on this). The problem with using it to much beyond Mach 5 is not temperature but Momentum drag;  that is the force needed to stop the air so it can be fed to engine is getting comparable to the force the engine provides and acceleration slows down to the point it is not worthwhile.  If you want to go much beyond the Mach 5 or 6 region you need to go to Scramjets where the air is not stopped.

Also RanulfC is correct that heating on the way up is also an issue it is why the medium temperature TPS extends over the whole airframe rather than just the re-entry areas as on the Shuttle. This is another reason why Mach 5 is looking like a good transition point.  In the end it is difficult to firmly establish whether it is the journey up or down that is driving the SKYLON temperature control.

Another point RanulfC is correct to point out is that once you start to use LOX the effective Isp starts drop dramatically.

[mmeijeri]

To what speed do you need to slow down the air before you can feed it to the engine? I recall reading somewhere that there were concepts for high subsonic combustion, but I don't remember if it was in the Skylon documentation or somewhere else.

[RanulfC]

RanulfC  “Aerospikes work?”
Well yes. They work well on static test stands but once on flying vehicle complications arise like base drag and the interaction of the two supersonic flows they get more complicated. That is not to say they cannot be made to work it is just a lot more complex than simple theory suggests. There are similar complexities with Expansion/Defection nozzles and we may yet still use these, so it is not that we are anti-advanced nozzles .

Actually FLIGHT testing has proven the "complications" aren't as large an issues as was suspected. The test-stand testing didn't address some of those concerns that's WHY they were flight tested. Flight tests have been done with both liquid bi-propellant, mono-propellant,and solid aerospike engines and they do work.

I don't think you're "anti-advanced-nozzles" and don't mean to give that impression. I just am tossing the information out there because it seems that many people aren't aware of how far the research has gone because the majority of the work hasn't been done under large, big-budget programs run by the major space agencies. The current research programs are private industry and collage/university level endevours for the most part. Current research is focusing on multi-nozzle "plug" and "dynamic" aerospike engines scalable for larger launch vehicles.

RanulfC  “On the Delta Clipper design, the "nose-first" entry wasn't a "design" decision that was a requirement to meet the Air Force specified cross-range ability”
I agree it was not driven by the UASF requirements my point was that the team most motivated to follow Bono’s design decided it was more difficult to do than the nose first approach. (Again I would point out I was in Huntingdon Beach working with McDAC working on this programme and saw these decision being made). I am sure “Aerospike as a heat shield” might be made to work but, once the realities have been engineered in, it might not be as attractive as Bono initially thought.

Uhm, the "Nose-First" reentry WAS an Air Force requirement because a base first ballistic entery did not have the necessary "cross-range" as required by the Air Force. Sorry if that was confusing the issue. I understand that you were there when the decsions were made but the REASON the base-first entry wasn't considerd was that the Air Force had set a required "cross-range" ability for the vehicle and there was no way to make that requirement with a base-first ballistic entry. So a lifting "Nose-First" entry was required to meet the specified cross-range distance.

While there may have been 'motivation' to use the previous work related to "plug-nozzle aerospike as engine and heat-shield" from the 60s (which, BTW was mostly funded by the Air Force) the given requirements would not allow consideration of that method.
As part of the problems with Delta Clipper was the added TPS requirements that the "nose-first" entry required which added much more mass to the system than the simpler "ballistic entry" with the plug-nozzle and base activly cooled did. The neccessary manuevers (end-over-end flip required to go from "nose-forward" glide to "base-first" for landing) also added complexity and increased the danger, and difficulty of recovery versus the more straight forward base-first all the way landing that the Bono style engine/heat-shield system would have used.

It's not that MacDac didn't want to use the system or that they didn't feel it was technically sound, the given requirements for the flight parameters diss-allowed the idea of ballistic "base-first" entry from the design.

I hope this clears up what I was meaning.

Randy

[Hempsell]

mmeijeri “To what speed do you need to slow down the air before you can feed it to the engine?”

To all intensive purposes it is brought to a stop in the Nacelle before it is drawn through the heat exchanger.

RanulfC The Delta Clipper  "Nose-First" reentry

As I recall it (and it was 20 years ago and my notes are no longer with me - property of what is now Astrium) the underlying motivation of the team for the nose first was they were not convinced about things like using a low level burn to protect the combustion chambers and on other work I have found these low level protective burns consume more propellant than I for one first thought. Although the Delta Clipper did baseline an aerospike plug several team members, including Max Hunter if I remember correctly, did not really like the engines and were arguing for conventional bells.

I have to admit I joined the team after the nose first decision had been made and they may have been more influenced by the Air forces requirements than I was aware, but I am certain they saw it as a better technical approach to “engine first” regardless of meeting the Air Force’s Cross Range requirements which were so modest I suspect an “engine first” configuration could have been found to meet them.

My job on the team was to explore a HOTOL like alternative, which of course had massive cross range but that feature gave us no brownie point whatsoever!

[Warren Platts]

Mr. Hempsell, I'd just like to say that we really appreciate you taking the time to keep us informed on this discussion forum. Thank you sir!

[Cinder]

I'll second that.  A real privilege, thank you.

[RanulfC]

RanulfC The Delta Clipper  "Nose-First" reentry

As I recall it (and it was 20 years ago and my notes are no longer with me - property of what is now Astrium) the underlying motivation of the team for the nose first was they were not convinced about things like using a low level burn to protect the combustion chambers and on other work I have found these low level protective burns consume more propellant than I for one first thought. Although the Delta Clipper did baseline an aerospike plug several team members, including Max Hunter if I remember correctly, did not really like the engines and were arguing for conventional bells.

Well I actaully 'see' an issue with something in the above information that doesn't bode well for the design team ACTUALLY checking the work on a plug-nozzle/Aerospike engine!

Plug-nozzle/Aerospike engines don't DO a "burn" to protect the combustion chambers of the rocket nozzles. You ONLY use that with more conventional bell nozzles and it is very much less for protecting the combustion chambers as it is for generating a pressure bubble to defelect heat from the nozzles themselves.

In "heat-shield" mode Liquid Hydrogen (in an LH2/LOX vehicle, you can also use Lox, fuel, even water) is circulated behind the "plug" base where the majority of heating is taking place and then flows around the engine system and finally is dumped into the combustion chambers and into the vehicle slipstream. As noted if you're not using LH2 other cryogenic and some non-cryogenic propellants can be used instead. It was shown that in some cases (mostly the low-density LH2 designs) that it might actually be better to carry a volume of water instead of more LH2.

I'm wondering if the design baseline wasn't actually a "plug-cluster-nozzle" system instead of an aerospike design?

I have to admit I joined the team after the nose first decision had been made and they may have been more influenced by the Air forces requirements than I was aware, but I am certain they saw it as a better technical approach to “engine first” regardless of meeting the Air Force’s Cross Range requirements which were so modest I suspect an “engine first” configuration could have been found to meet them.

I don't think an 'engine-first' ballistic design would have made the cros-range because by thier nature ballistic's don't have good cross-range. And especially in the case of the Air Force even "modest" cross-range can be rather a 'broad' defninition depending on the bias of the Air Force people involved. They have a rather 'bad' habit of waiting till a team settles on a  good design and then asking that the cross-range be increased... A lot!

I have a suspicion that there was an overall in-group opposition to the possible use of a plug-nozzle aerospike no matter what the original "baseline" design called for. As I recall Max Hunter for one, though being a SSTO-VTOL believer was (IIRC) unconvinced that ANY aerospike engine was a technically sound idea. I seem to recall him being quoted at one point saying a working "aerospike-rocket-engine" of any type would never fly because the theory and design concept were inheirently flawed.

Which ends up injecting what I generally call (after a Sci-Fi short story of the same name, synopsis below*) "Paper Virus" into the mix.
When an "expert" or respected manager/engineer/whatever passes judment on a system, concept, or idea there is a tendency for those around (even if they have another opinion) to accept that judment at face value with no questions asked.

This is similar to an issue with Space Suits that I have long been fighting, with some success. I have studied all the available data on, and confured with the designer and original researcher on a concept space suit called: The Space Activity Suit which is also called a "Mechanical Counter-Pressure Space Suit. This was designed in the early 70s as an alternative to the heavy and difficult to move Apollo and previous space suits and rather than trying to surround the body with small space ship filled with air it used the fact that the human skin, as long as it is reinforced to equalize lung and internal pressure with pressure from the outside is immune to the effects of vacuum.

Very basiclly it is a leotard garment made up of layers of stretchable cloth that squeezes the body and reinforces the skin to match the helmet and internal organs pressure. The suit is extremly light weight, easier to move and do activities in by orders of magnitude than any "pressure" suit ever designed. Thermal control is regulated by already existing body function such as sweating so this does not have to be included in the life support for the suit. If you get cold you put on an inulated set of coveralls. Simple, easy and pretty cheap. Once a way is found to allow the mechanical pressure to 'relax' when out of a vacuum, (current work is looking at air-tubes or memorty metal strips used in a manner similar to the 1950s "capstan" emergency pressure suit) it can be work comfortably as normal "clothing" until needed.

I had a chance to call into the Space Show at one point and ask some questions of the man who is considerd to be the "foremost" expert on Space Suit design and asked him about commercial and NASA work on the Mechanical Counter-Pressure suit. He quickly informed me that the entire concept was UNWORKABLE and had never passed any research test. When I mentioned by research on the subject and that it HAD passed NASA testing he snapped that the "astronauts" had been to "macho" during testing to report any problems and that the information I had was totally false. He then repeated that the concept was a dead and an would never work and went on to another question.

I should note that I have emailed him, tried asking on forums or message boards I happen to find him on to name WHICH astronauts participated in the test program as I could find none listed in the data as having even been asked about participation in the testing let alone having worn the suits at all. The only response I've ever recieved basiclly told me all the research I've mentioned is a lie, that the concept was never successfully tested and that he had no time to argue with someone who would obviously not listen to the 'facts' as given by an expert.
I can understand that he's busy, since he is now designing space suits for the Indian Human Space Flight program

The main point is this: Despite his being the worlds "foremost" expert on Space Suit design and having written all the text books used for teaching everyone ELSE how to design a space suit, the "concept" he does not believe will ever work HAS been tested extensivly (even in vaccum) and has passed every test with much higher marks than any "balloon" pressure suit can possibly achieve. Even more important is the fact that NASA and Russia are using more and more "low-pressure" balloon suits with bits and pieces of mechanical counter-pressure encorperated into it. The last winner of the NASA sponsored "glove" competition walked away easily with a design that surpassed every checkpoing and was a Mechanical Counter-Pressure glove!

Yet there are people being trained professionally to design the next generation space suit being told (and taught!) specifically that MCP does not work!

Similarly if "experts" are against using aerospike rocket engines they tend to get listened to, and believed even if they have no evidence to back them up. It's not a conspiricy, not a plot, nothing more sinister than common human emotionalism

My job on the team was to explore a HOTOL like alternative, which of course had massive cross range but that feature gave us no brownie point whatsoever!

Of course not... It's not like we actaully WANT anyone to win now is it? )

Randy

*Story was: People on space station that were flying probes around he sun to gather data. One "visiting" proffessor is trying to prove/disprove a theory that might allow for FTL flight. Most of the probes with data for this experiment come back damaged and the majority of what little data comes through seems to indicate that the theory is wrong.

Proffessor is going to declare the theory wrong and use the data gathered to refute the theory which he's "sure" will put him in the text-books for his efforts.

Turns out the "proffessor" is no such person and is an alien sent to discredit the suggested theory because it leads to the discovery of an FTL drive. By 'proving' the theory wrong and getting that conclusion entered into educatioin, reference materials, and databases the "wrong" conclusion becomes in fact a "paper-virus" whos main purpose is to stunt and marginalize any work done which might prove more succesful and therefor lead to a working FTL theory and drive!

[Hempsell]

Thanks for the comments on my participation I am flattered and very glad it helps out.  This also helps us out by reflecting back how our project is perceived externally.

Randy

You are right about just hydrogen bleeding rather than burning I should have been more precise.  I have never looked at this in detail' my experience with real burning of a Hydrazine NTO engine to protect it during reentry and was a surprise at the amount it consumed, but I accept the hydrogen case may be different.

The nozzle on Delta Clipper was segmented (you can see the fences on the plug skirt) and it was expected that on landing only some of the segments would be used. To be honest I cannot remember if the nozzle was fed with a full ring chamber (also segmented of course) or a cluster of conventional chambers.

You are right about the inherent suspicion of plug nozzles by some of the team especially (as I have ready pointed out) Max Hunter.  I was not one of those sceptics, however research work I started when at the University of Bristol has sobered me up a bit.  As I have said before I am sure they can be made to work, but they are complex and small changes to parameters can massively change the nozzle behaviour.  Anyone can design a working bell nozzle with a pencil, piece of paper and a calculator. Advanced nozzles must have advanced CFD verified by extensive experimental work on the precise configurations you expect to use in the real external flow regimes.

I agree with you on the potential of Mechanical Counter-Pressure Space Suits.

[Jim Davis]

Mr. Hempsell,

If memory serves one of the problems of air liquefaction and/or precooling schemes has been the icing up of heat exchangers because of moisture in the air.

How does Skylon/SABRE deal with this issue?

[Hempsell]

Mr. Hempsell,

If memory serves one of the problems of air liquefaction and/or precooling schemes has been the icing up of heat exchangers because of moisture in the air.

How does Skylon/SABRE deal with this issue?

You are right it is a big problem.

On HOTOL we injected liquid oxygen to take the air below the critical temperature so the ice crystals acted like dust that blew through rather than sticky particles that attached to the tubes.

Both Japan and USA have successfully tested ethanol (and methanol) with glycol sprayed on the tubes.

How it is done in SABRE is a commercial secret – Sorry to be evasive.

[RanulfC]

Hempsell;
I'll agree with your assesment since you've probably had more access to materials than I have. (I still have a 150 page print out listing a total DoD/NASA/AF specific paper search on "plug-nozzle engines" 90+ percent of which I STILL can't access... :::sigh::: )
And I won't disagree about the amount of work needed to get them to work but I WILL point out that so far while the majority of static testing was done with big-bucks and work by NASA and the Air Force, the only folks who are doing actual flight testing is Garvey Aerospace and the University of California.

(Of interesting note is that both GA and UCLB have become less forthcoming on current work on a multi-chamber engine because of NDA's on the technology due to private funding of the research )

Randy

[Jim Davis]

How it is done in SABRE is a commercial secret – Sorry to be evasive.

Mr. Hempsell,

No need to apologize; I quite understand.

At the risk of treading on other commercial secrets I'll ask another question if you'll indulge me.

It appears at first glance that Skylon, with its widely separated engine nacelles, would be difficult to control if there was a thrust imbalance. Specifically, an inlet unstart would subject the spacecraft to a severe yaw, would it not? Can that inadequate looking vertical tail suffice to control Skylon in those conditions? Or does the other inlet unstart to maintain a thrust balance? Would the mission have to be aborted in such a case?  Or does Skylon have such large static margins to make unstarts very unlikely?

Let me add my thanks to all the others for taking the time to post here.

[Hempsell]

How it is done in SABRE is a commercial secret – Sorry to be evasive.

Mr. Hempsell,

No need to apologize; I quite understand.

At the risk of treading on other commercial secrets I'll ask another question if you'll indulge me.

It appears at first glance that Skylon, with its widely separated engine nacelles, would be difficult to control if there was a thrust imbalance. Specifically, an inlet unstart would subject the spacecraft to a severe yaw, would it not? Can that inadequate looking vertical tail suffice to control Skylon in those conditions? Or does the other inlet unstart to maintain a thrust balance? Would the mission have to be aborted in such a case?  Or does Skylon have such large static margins to make unstarts very unlikely?

Let me add my thanks to all the others for taking the time to post here.

The answer is we can cope with an engine out throughout the whole but it does mean a mission abort.

The exact procedure depends on the nature of the failure and when in the ascent it happens.  But the tail fin does play a significant part in the yaw control in many of the cases.  Both it and the forward fin may look intuitively small compared with the body but it is because the main body is abnormally large for the vehicle mass (and hence moments) being about half way to a 1930s airship when empty.

[Hermit]

RanulfC  “Aerospikes work?”
Well yes. They work well on static test stands but once on flying vehicle complications arise like base drag and the interaction of the two supersonic flows they get more complicated. That is not to say they cannot be made to work it is just a lot more complex than simple theory suggests. There are similar complexities with Expansion/Defection nozzles and we may yet still use these, so it is not that we are anti-advanced nozzles .

Actually FLIGHT testing has proven the "complications" aren't as large an issues as was suspected. The test-stand testing didn't address some of those concerns that's WHY they were flight tested. Flight tests have been done with both liquid bi-propellant, mono-propellant,and solid aerospike engines and they do work.

There seem to be quite a few University teams now going for flight tests of aerospikes but I don't remember many that have actually flown.
Could you provide some links regarding the aerospike flight tests.
I'm aware of: http://pdf.aiaa.org/preview/CDReadyMJPC2005_1177/PV2005_3797.pdf and the work of Garvey and CSULB http://pdf.aiaa.org/preview/CDReadyMJPC2004_946/PV2004_3354.pdf.

[Xinvoker]

REL's August update:

During August the first tube manipulation machine was commissioned. Its installation at our manufacturing facility has enabled the assembly of the first production heat exchanger module to begin.
The lack of published news this month is due to internal preparations for significant events, which will be covered in next month's Update.
Be sure to check us out in October!

A teaser!? That's a first i think, i wonder what these events might be.

[RanulfC]

RanulfC  “Aerospikes work?”
Well yes. They work well on static test stands but once on flying vehicle complications arise like base drag and the interaction of the two supersonic flows they get more complicated. That is not to say they cannot be made to work it is just a lot more complex than simple theory suggests. There are similar complexities with Expansion/Defection nozzles and we may yet still use these, so it is not that we are anti-advanced nozzles .

Actually FLIGHT testing has proven the "complications" aren't as large an issues as was suspected. The test-stand testing didn't address some of those concerns that's WHY they were flight tested. Flight tests have been done with both liquid bi-propellant, mono-propellant,and solid aerospike engines and they do work.

There seem to be quite a few University teams now going for flight tests of aerospikes but I don't remember many that have actually flown.
Could you provide some links regarding the aerospike flight tests.
I'm aware of: http://pdf.aiaa.org/preview/CDReadyMJPC2005_1177/PV2005_3797.pdf and the work of Garvey and CSULB http://pdf.aiaa.org/preview/CDReadyMJPC2004_946/PV2004_3354.pdf.

So far the NASA solid-aerospike, NASA X-33 LASR tests, and the Garvey/CSULB flight tests are the only non-DoD stuff I've seen and they are "true" aerospike's rather than the plug-nozzle's and clustered plug-nozzles which are the "duel-use" designs.

I've had issues finding the various data on my usual searches (DTIC/NTRS/etc.) in the last few years as all of my usual "key-words" now bring up very little OTHER than the X-33 linear aerospike work. The DoD/Air Force had gotten as far as building and static testing several different types of flight-hardware pressure-fed, and pump-fed plug-nozzle engines as well as plug-cluster and "true" aerospike engines for various uses.

No actual flight-tests were conducted though and the Air Force had to shut the program down and transfer the data to NASA. Since NASA didn't have funding for it either nothing much has come of the research until the X-33 engine program.
What the USAF DID do is a huge amount of flight configuration wind tunnel work at subsonic, transonic, and medium to high supersonic speeds. But there was a 'gap' in most of the data for low-supersonic speeds and at one point it was reported (during the run-up to the X-33 program) that CFD modelling showed the aformentioned "base-drag" and "supersonic-flow" issues to be pretty daming for the performance of the aerospike design.

From what I read at the time the CFD was actually flawed in that it was based on data from pressure-fed and plug-cluster engines without either the "true" aerospike compression ramp or the "aerodynamic" aerospike of the standard plug-nozzle where turbopump exhaust is directed out the base of the "plug" and aerodynamically produces the elongated "spike" expansion ramp effect. Further CFD's have shown a slight loss of performance around Mach-3 on a nominal rocket flight trajectory due to reduced air-pressure on the outer plume but constriction by the Mach-3 shockwave but that seems to be the only operational issue with the design. (If it actually exists and isn't a "bug" in the CFDs but only actual flight testing to Mach-3 and beyond will tell)

The "duel-use" of a plug-nozzle engine as a heat shield was also well developed during the DoD program even as far as heat soak testing of water, LOx, and LH2 flow systems at reentry temperatures. No money for flight testing though.

Randy

[mlorrey]

So far the NASA solid-aerospike, NASA X-33 LASR tests, and the Garvey/CSULB flight tests are the only non-DoD stuff I've seen and they are "true" aerospike's rather than the plug-nozzle's and clustered plug-nozzles which are the "duel-use" designs.

Um, X-33 LASR was a true aerospike?

[mboeller]

Hi;
I’m rather late to the party so I have only a few comments regarding a few of the discussed designs and ideas.
I personally have 3 favorites regarding SSTO or TSTO concepts, because IMHO you have to learn to crawl before you can walk or even run. But it seems, that always the “feature-creep” seems to demand that the concepts are improved beyond reason so the best concepts, which are for me always the simplest designs, fall out of favor. Therefore my favorite designs are rather simple and straightforward compared to the more elaborate designs favored by the “customers” like NASA and/or the US Air Force:
1) Gomersall’s SSTO concept
2) GD’s Millennium Express
3) LNLL’s Space-Jet concept

1) Gomersall’s SSTO concept is IMHO really good because it had realistic mass properties and technology, therefore it was “suppressed”. The only information about this SSTO which is available is, AFAIK, the document from Scott Lowther. See:
Space-Doc 42 on this page:  http://www.up-ship.com/drawndoc/drawndocspaceshuttle.htm
preview: http://www.up-ship.com/drawndoc/sdoc42ani.gif

2) My next favorite is the Millennium Express from GD which was already mentioned here. The reasons that this concept is one of my favorites are the same as with the Gomersall’s concept: Simple axi-symmetrical structure, base-first reentry, “normal” technology etc. Here you can see a few information’s about the Millennium Express; unfortunately the only ones that I have found:
http://www.spacefuture.com/archive/single_stage_to_orbit_vertical_takeoff_and_landing_concept_technology_challenges.shtml
Also here is a overview of a lot, if not all VTVL SSTO concepts in the world:
http://www.spacefuture.com/archive/history_of_the_phoenix_vtol_ssto_and_recent_developments_in_single_stage_launch_systems.shtml

3) My third favorite is the Space-Jet concept from LNLL which again was already mentioned. Here my reasoning is different. The Space-Jet concept is definitely not simple but has a new twist which makes it IMHO worthwhile. The Space-Jet concept is AFAIK the only TSTO concept which uses small boosters and a large orbiter. All the other TSTO’s I know use a big booster and a rather small orbiter. This makes the Space-Jet concept unique. For me the Space-Jet is the only reasonable way to do it, because the orbiter has to cover by far the largest portion of delta-V to LEO regardless if the staging is at M3 or even M7 and therefore the orbiter has to have the best possible mass properties. A concept which uses a large orbiter like the Space-Jet allows this and improves the changes to succeed with a meaningful payload left.
Unfortunately as with the other concepts not a lot of information is available about this concept. The only real good information was included in the APR N5V5 from Scott Lowther. See:
http://up-ship.com/blog/?p=6010
http://up-ship.com/blog/wp-content/uploads/2010/04/art33ani.jpg

Unrelated to the topic of SSTO/TSTO designs IMHO you should not forget this recently discussed remarkable fuel/oxidizer-combination: Lithium AluminumHexahydride fuel with H2O2 as oxidizer. If the mentioned ISP's are correct (469sec! or 430sec when derived from table 2) then this really dense fuel/oxidizer combination would make SSTOs much easier.
Link:  http://www.sps.aero/Key_ComSpace_Articles/LibTech/LIB-035_Novel_Organometallic_Propellants_for_Hypergolic_Applications.pdf

One final off topic point is the mentioning of the against pressure space suit by RanulfC. IMHO you don’t have to develop such a space suit completely from scratch because IMHO the basic technology was already developed years ago and is even (sort of) available on the market.
The small innovative company prospective concepts had developed a fluid muscle based Anti-G suit nearly 20 years ago already which was then improved and marketed by the company Autoflug; the Libelle Anti-G suit. This Anti-G suit was developed for the german Eurofighter pilots but the german Luftwaffe decided that the suit is not necessary or sufficient. What makes this suit unique is that they use water as fluid muscle to pressurize the Anti-G suit and that the Anti-G effect covers the complete body of the pilot (hands and feet are excluded). IMHO you only have to pressurize the water in the Libelle suit with a different means and then it should work well enough as a basis for an against pressure suit. 
Here are a few links; unfortunately most are in german language:
http://de.wikipedia.org/wiki/Libelle_(Anzug)
http://www.airpower.at/news01/0625_libelle/libelle1.htm  (3 pages, libelle1-3)
http://www.autoflug.net/page.php?page_id=365&menu_id=&ebene=&lang=eng&bereich=p


kind regards
Manfred

[RanulfC]

So far the NASA solid-aerospike, NASA X-33 LASR tests, and the Garvey/CSULB flight tests are the only non-DoD stuff I've seen and they are "true" aerospike's rather than the plug-nozzle's and clustered plug-nozzles which are the "duel-use" designs.

Um, X-33 LASR was a true aerospike?

Yes actually in that it was designed with a "complete" expansion spike (ramp) for the exhaust. Unlike the 'trunicated' plug-nozzle's which use exhaust gases to form an "aerodynamic" spike the "true" aerospikes have a solid surface with which the engine exhaust interacts. In the X-33 engines case each nozzle had a ramp which acted as the "spike" for expansion on one side while ambient air pressure formed the outer surface of the exhaust.

Randy

[Proponent]

1) Gomersall’s SSTO concept is IMHO really good because it had realistic mass properties and technology, therefore it was “suppressed”. The only information about this SSTO which is available is, AFAIK....

Some info is available from NTRS:

http://ntrs.nasa.gov/search.jsp?R=502489&id=1&as=false&or=false&qs=Ntt%3Dgomersall%26Ntk%3Dall%26Ntx%3Dmode%2Bmatchall%26Ns%3DHarvestDate%257c1%26N%3D0

If that link doesn't work, just search for "gomersall".

[RanulfC]

Manfred; Welcome to the topic

Hi;
I’m rather late to the party so I have only a few comments regarding a few of the discussed designs and ideas.
I personally have 3 favorites regarding SSTO or TSTO concepts, because IMHO you have to learn to crawl before you can walk or even run. But it seems, that always the “feature-creep” seems to demand that the concepts are improved beyond reason so the best concepts, which are for me always the simplest designs, fall out of favor. Therefore my favorite designs are rather simple and straightforward compared to the more elaborate designs favored by the “customers” like NASA and/or the US Air Force:
1) Gomersall’s SSTO concept
2) GD’s Millennium Express
3) LNLL’s Space-Jet concept

The Gomersall's SSTO isn't actually "surpressed" as much as there doesn't seem to be a lot avaiable in the first place on the concept. As note by Scott Lowther it is of the same "class" as the Nexus, Bono-related, and Phoenix type VTVL-SSTOs and looks to have come out (1970) about the time Apollo was looking to ramp down so there was probably little if any follow up on the idea.
The GD-Millenium Express is also a good "minimalist" design for an SSTO though it seems to also suffer from the "normal" SSTO issues of not having a lot of margin to play around with during development. Adding boosters is always possible of course but like the "Spacejet" it makes the vehicle a TSTO if not at least a 1.5STO which (much as I like both concepts) doesn't exactly fit the topic unfortunatly

Unrelated to the topic of SSTO/TSTO designs IMHO you should not forget this recently discussed remarkable fuel/oxidizer-combination: Lithium AluminumHexahydride fuel with H2O2 as oxidizer. If the mentioned ISP's are correct (469sec! or 430sec when derived from table 2) then this really dense fuel/oxidizer combination would make SSTOs much easier.
Link:  http://www.sps.aero/Key_ComSpace_Articles/LibTech/LIB-035_Novel_Organometallic_Propellants_for_Hypergolic_Applications.pdf

Thanks I hadn't seen that one before I'll have to read up on it.

One final off topic point is the mentioning of the against pressure space suit by RanulfC. IMHO you don’t have to develop such a space suit completely from scratch because IMHO the basic technology was already developed years ago and is even (sort of) available on the market.

Actually there HAS been a lot of research done on the Mechanical Counter-Pressure Suit, and in fact the original inventor has a company that is taking the concept public. More can be found at "elesticspacesuit.com" if anyone is interested.

Randy

[mboeller]

Some info is available from NTRS:

http://ntrs.nasa.gov/search.jsp?R=502489&id=1&as=false&or=false&qs=Ntt%3Dgomersall%26Ntk%3Dall%26Ntx%3Dmode%2Bmatchall%26Ns%3DHarvestDate%257c1%26N%3D0

If that link doesn't work, just search for "gomersall".

Strange,

When I looked a few years ago on the net, and also specifically at ntrs.nasa.gov the report was not available.

The PDF from ntrs is the same as the PDF from Scott Lowther by the way.

kind regards

Manfred

[mboeller]

The Gomersall's SSTO isn't actually "suppressed" as much as there doesn't seem to be a lot avaiable in the first place on the concept. As note by Scott Lowther it is of the same "class" as the Nexus, Bono-related, and Phoenix type VTVL-SSTOs and looks to have come out (1970) about the time Apollo was looking to ramp down so there was probably little if any follow up on the idea.

the "suppressed" is in brackets for a reason. I was not really serious about the suppression. But others seem to think this proposal was really kind of suppressed :

The vehicle was promoted as a Space Shuttle, and because of the strong opposition within NASA headquarters and the two other principal NASA manned Spaceflight Centers, Marshall and Johnson, the concept was quickly suppressed and Gomersall was assigned to non-launch vehicle-related duties [Ref. personal communication].

source:  http://www.spacefuture.com/archive/history_of_the_phoenix_vtol_ssto_and_recent_developments_in_single_stage_launch_systems.shtml

The GD-Millenium Express is also a good "minimalist" design for an SSTO though it seems to also suffer from the "normal" SSTO issues of not having a lot of margin to play around with during development. Adding boosters is always possible of course but like the "Spacejet" it makes the vehicle a TSTO if not at least a 1.5STO which (much as I like both concepts) doesn't exactly fit the topic unfortunatly

True, but SSTO's are always on the edge of the technology. Therefore, at least for me, it should be mandatory to use a design with the highest possible mass margin, simplest structure and the least demanding systems and environment. This is IMHO always violated as far as I can see.

Thanks I hadn't seen that one before I'll have to read up on it.

I'm really curious about this fuel/oxidizer combo, because it is AFAIK something really new and the quoted performance is simply outstanding, if true.

Actually there HAS been a lot of research done on the Mechanical Counter-Pressure Suit, and in fact the original inventor has a company that is taking the concept public. More can be found at "elesticspacesuit.com" if anyone is interested.

WOW! Thanks for the link. I know that a lot or research was done in the 60th's but I was not aware of the recent developments.


Manfred

[Da5id]

http://www.bbc.co.uk/blogs/thereporters/jonathanamos/2010/09/it-is-one-of-those.shtml

Thought this might be interesting in this thread about Skylon

[93143]

Very interesting.  Apparently the D1 looks like it will have transatlantic ferry range on straight hydrogen.  I wonder if you could increase that by loading some LOX and going semiballistic...  well, obviously you could; it's an SSTO - but it might stress the vehicle too much and wind up counting as one of the 200 "flights" it was designed for...

Anyway, I hope this works...  good to hear the London Economics people think it's such an awesome idea...

I can see this cornering the market for 15 mT and below, while SpaceX expands upwards with their Merlin 2 - the launch cost per kg on a Falcon XX could get fairly close to the expected figures for Skylon, provided the market doesn't suddenly go nuts and start flying Skylons for close to the cost of propellant...

[mlorrey]

Very interesting.  Apparently the D1 looks like it will have transatlantic ferry range on straight hydrogen.  I wonder if you could increase that by loading some LOX and going semiballistic...  well, obviously you could; it's an SSTO - but it might stress the vehicle too much and wind up counting as one of the 200 "flights" it was designed for...

Anyway, I hope this works...  good to hear the London Economics people think it's such an awesome idea...

I can see this cornering the market for 15 mT and below, while SpaceX expands upwards with their Merlin 2 - the launch cost per kg on a Falcon XX could get fairly close to the expected figures for Skylon, provided the market doesn't suddenly go nuts and start flying Skylons for close to the cost of propellant...

Don't worry, British and EU taxes will make up for it.