Lorrey Aerospace planning X-106 spaceplane

[mlorrey]

Those of you with an interest in spaceplane concepts, advanced propulsion and TPS, are welcome to review my ongoing design development at http://www.lorrey.biz of my X-106 Hyper Dart concept.

Discuss...

[Chris Bergin]

This will be interesting if you're willing to take questions?

[mlorrey]

Chris Bergin - 28/5/2006  7:31 AM

This will be interesting if you're willing to take questions?

Certainly.

[kevin-rf]

From the Website :

http://www.lorrey.biz/x-106/vision.html

NASA uses hydrogen wherever they can for a few reasons: one is that their budget is determined by political committees of senators and congressmen, whose election is often dependent upon the environmentalist vote, which tends to support a shift to a hydrogen economy; another reason is that a vehicle that is more difficult to reach orbit means it is more expensive, with more money to go to more contractors, which can be located in more congressional districts, which also serves the political ends of congressmen seeking pork for their constituencies. Additionally, NASA's mandate is about developing advanced technologies, which is used to spin off into the private sector to maintain US technological and economic supremacy. This also serves the political ends of congressmen; lastly, I believe that NASA engineers just haven't deeply analysed the issue, and are willing to go with the politically expedient choice.

Ouch, now I am no friend of Hydrogen as a first stage fuel due to the low density issue but that looks to be picking a fight where none is needed.

Still looks to be a fun project...

Have you done any research into Methane vs. Kerosene? It would be a lowwer density fuel than kerosene, but would have a higher ISP.

[Jim]

I am going the throw the B.S. card.  The website is nothing but rants on supposed conspiracies.  

The hydrogen rant quoted in the last post.  Atlas and Delta are not NASA programs.  They chose their propellants based on their own trade studies.

Here is another one

"Boeing and LockMart, run a duopoly both in satellites and launchers, selling launch services to those they build satellites for"

Yes, sometimes the spacecraft manufacturer buys the launch service as part of the contract for the spacecraft services.  But both LM and Boeing have firewalls (per FTC direction) between the spacecraft and LV divisions.  They can not share any info that it outside of the few missions that fly on the same LV.

The attached chart clearly shows that Boeing and LM do not fly on their own LV's and there is no "duopoly".    Boeing and LM only have approx 40% of the spacecraft market (mostly because Boeing bought Hughes), yet only 20% of the LV market (Sealaunch not included due to previous Hughes contract).

Additionally, It is the spacecraft driving the LV's to carry more and not by the choose of the LV contractors.  
The Atlas II - III growth path; the addition of solids to the EELV's, the Ariane ECA development  and Sealaunch Zenit uprating were all driven by spacecraft mass increases.  Not the reverse, the LV's just reacted to market pressure.

[mlorrey]

Jim - 29/5/2006  8:55 AM
Here is another one

"Boeing and LockMart, run a duopoly both in satellites and launchers, selling launch services to those they build satellites for"

Yes, sometimes the spacecraft manufacturer buys the launch service as part of the contract for the spacecraft services.  But both LM and Boeing have firewalls (per FTC direction) between the spacecraft and LV divisions.  They can not share any info that it outside of the few missions that fly on the same LV.

The attached chart clearly shows that Boeing and LM do not fly on their own LV's and there is no "duopoly".    Boeing and LM only have approx 40% of the spacecraft market (mostly because Boeing bought Hughes), yet only 20% of the LV market (Sealaunch not included due to previous Hughes contract).

Additionally, It is the spacecraft driving the LV's to carry more and not by the choose of the LV contractors.  
The Atlas II - III growth path; the addition of solids to the EELV's, the Ariane ECA development  and Sealaunch Zenit uprating were all driven by spacecraft mass increases.  Not the reverse, the LV's just reacted to market pressure.

Since Sealaunch is owned by the two (plus Energia), its launches should be included, so you are fudging the numbers. 62% of Boeing/LM satellites launch on Boeing/LM launchers, and B/LM launches 44% of all launches. That is pretty duopolistic.

Also: what time span is this chart supposed to represent? Cherry picking a time frame beneficial to one's argument isn't appropriate, so lets see what time frame this refers to.

[mlorrey]

kevin-rf - 28/5/2006  9:56 PM

From the Website :

http://www.lorrey.biz/x-106/vision.html

Ouch, now I am no friend of Hydrogen as a first stage fuel due to the low density issue but that looks to be picking a fight where none is needed.

Still looks to be a fun project...

Have you done any research into Methane vs. Kerosene? It would be a lowwer density fuel than kerosene, but would have a higher ISP.

Methane is a good performer, though its handling issues are not something I want to deal with. If you look at Dr. Dunn's paper on alternative fuels for SSTO, the numbers just don't work. Also, since the fuel is being stored in tanks formed from airframe bulkheads and skins, pressurization is also an issue. I only want to have to deal with, at most, one pressurized and/or cryogenic propellant. Even there, I'm looking at dumping the pressurized LOX tank in favor of a larger volume unpressurized H2O2 tank. This would increase oxidizer density as well (leading to CG as well as wing loading issues) but will allow more fuel to be carried for greater dv. This is the route that Clapp went with his Black Horse project. I'm not convinced as to the advantages of peroxide yet.

As for my harshness towards hydrogen and those at NASA and in the NASA amazing people circuit who promote it, I admit to some irritation at the constant harping of the Green contingent for so-called Green rocket fuels. The facts are that lots of fossil fuels get burned producing so called "green" fuels, which makes them not so green. Nor is LH2 that great for SSTO payload to orbit, as numerous NASA SSTO studies have proven (which rarely, if ever, study side by side the performance of launching SSTOs with hydrocarbon fuels.)

LH2 is fine for upper stages, and for in-space propulsion (though it has issues of long term storage there as well), where high Isp is paramount.

My primary arguments are that air augmentation is needlessly being given short shrift by NASA, that NASA, as seems sadly typical, picked two wrong horses (scramjets and hydrogen, rather than ramjets and hydrocarbons) when it did study air augmented launching. The pattern of seeming to pick the wrong horse in so many key decisions (Thiokol over Aerojet on SRB contract, LH2 vs Hydrocarbons, Scramjet vs Ramjet, aluminum airframe vs hotframe on Shuttle, Blunt body vs SHARP TPS, composite vs metal tank on X-33, etc) just seems too glaring to be an accident to me.

Does my writing sound ranty? Maybe some of it does, but it merely reflects how much we in the space access community have become frustrated with NASA decisions, and why we are motivated to do it privately.

Now, you can complain that our negative attitudes toward NASA hurt the space program, you can complain I may be obsessed with ramjets (as some have said), or alternative fuels (as others have said), but it takes dissatisfaction and frustration with the status quo and obsession with solutions to create breakthroughs in any field of endeavor. If my work results in breakthroughs, by me or by others, then it is for the good of the greater space program, even if it may prove embarrassing to NASA in particular.

FWIW, I will recognise that NASA does seem honest enough to openly admit when it has been seriously wrong, when it has been proven by others. The openness of NASA recently to private launch services to ISS is IMHO welcome and refreshing, and I hope they follow through on their statements of support. If X-106 is someday contracted to launch to ISS, I will be both vindicated, and at the same time, NASA will prove it is capable of learning to make better decisions, which I sincerely hope for.

[Jim]

mlorrey - 29/5/2006  4:03 PM

Jim - 29/5/2006  8:55 AM
Here is another one

"Boeing and LockMart, run a duopoly both in satellites and launchers, selling launch services to those they build satellites for"

Yes, sometimes the spacecraft manufacturer buys the launch service as part of the contract for the spacecraft services.  But both LM and Boeing have firewalls (per FTC direction) between the spacecraft and LV divisions.  They can not share any info that it outside of the few missions that fly on the same LV.

The attached chart clearly shows that Boeing and LM do not fly on their own LV's and there is no "duopoly".    Boeing and LM only have approx 40% of the spacecraft market (mostly because Boeing bought Hughes), yet only 20% of the LV market (Sealaunch not included due to previous Hughes contract).

Additionally, It is the spacecraft driving the LV's to carry more and not by the choose of the LV contractors.  
The Atlas II - III growth path; the addition of solids to the EELV's, the Ariane ECA development  and Sealaunch Zenit uprating were all driven by spacecraft mass increases.  Not the reverse, the LV's just reacted to market pressure.

Since Sealaunch is owned by the two (plus Energia), its launches should be included, so you are fudging the numbers. 62% of Boeing/LM satellites launch on Boeing/LM launchers, and B/LM launches 44% of all launches. That is pretty duopolistic.

Also: what time span is this chart supposed to represent? Cherry picking a time frame beneficial to one's argument isn't appropriate, so lets see what time frame this refers to.

Sealaunch is only 40% Boeing and ZERO percent Lockheed interest.  SDO Yuzhnoye/PO Yuzhmash, Energia and Aker ASA are 60%

The 10 Boeing satellites on Sealaunch are from a previous Hughes block buy contract.

As the chart says since 2000, which covers the entry of Sealaunch, Atlas V and Delta IV into the commercial market.

[Jim]

Since there is only 4 launch service providers in the same class, Sealaunch, Atlas, Ariane, and Proton (Delta is not available commercially).  2 of them would be 50%.  So it makes sense that Boeing and lM have 50% of the market, but Sealaunch is only 40% Boeing, so Boeing and LM only have 35% of the LV market.

Boeing Satellite Systems is just Hughes Space and Comm, most of the comsat and LV contracts were done before the acquistion.  HSC was already the largest comsat manufacturer in the world, with no help from Boeing, so putting them as "Boeing" is misleading.

Proton and Atlas have marketing agreements, as well as Arianespace and Sealaunch, this pits LM against Boeing.    Even as LM and Boeing form ULA, commercial launches are not part of the agreement and the moneys not shared

[mlorrey]

Jim - 29/5/2006  6:54 PM

Since there is only 4 launch service providers in the same class, Sealaunch, Atlas, Ariane, and Proton (Delta is not available commercially).  2 of them would be 50%.  So it makes sense that Boeing and lM have 50% of the market, but Sealaunch is only 40% Boeing, so Boeing and LM only have 35% of the LV market.

Boeing Satellite Systems is just Hughes Space and Comm, most of the comsat and LV contracts were done before the acquistion.  HSC was already the largest comsat manufacturer in the world, with no help from Boeing, so putting them as "Boeing" is misleading.

Proton and Atlas have marketing agreements, as well as Arianespace and Sealaunch, this pits LM against Boeing.    Even as LM and Boeing form ULA, commercial launches are not part of the agreement and the moneys not shared

Ah, commercial launches are not shared? How are they going to work that? What is the point of ULA if they don't get economies of scale? sounds to me like 12 of one, dozen of the other.

[Tap-Sa]

Have you studied more about the Merlin-burning-boron-slurry concept? IMO it's the most probable showstopper for this. To make Merlin exhibit ~50% Isp increase means either much higher temps in the chamber or exhausting something much lighter than carbondioxide/water.

[Jim]

mlorrey - 29/5/2006  10:21 PM

Ah, commercial launches are not shared? How are they going to work that? What is the point of ULA if they don't get economies of scale? sounds to me like 12 of one, dozen of the other.

ULA is for gov't launches only.  Atlas for commercial launches will be marketed and managed by ILS/LM.

[Spirit]

Jim - 30/5/2006  1:54 AM

Since there is only 4 launch service providers in the same class, Sealaunch, Atlas, Ariane, and Proton (Delta is not available commercially)

What about the Japanese, Chinese and Indian launchers?

[aero313]

mlorrey - 29/5/2006  4:25 PM

My primary arguments are that air augmentation is needlessly being given short shrift by NASA, that NASA, as seems sadly typical, picked two wrong horses (scramjets and hydrogen, rather than ramjets and hydrocarbons) when it did study air augmented launching. The pattern of seeming to pick the wrong horse in so many key decisions (Thiokol over Aerojet on SRB contract, LH2 vs Hydrocarbons, Scramjet vs Ramjet, aluminum airframe vs hotframe on Shuttle, Blunt body vs SHARP TPS, composite vs metal tank on X-33, etc) just seems too glaring to be an accident to me.

Our company was one of the DARPA/RASCAL Phase 1 teams.  We studied airbreathing ascent vs. rocket (albeit with mythical MIPCC performance vs. mythical ramjet or scramjet performance).  The bottom line was that to get enough velocity to make the airbreather useful, you needed to fly very fast while down low in the atmosphere.  Aero heating was a perpetual problem.  More to the point, once you attained that useful velocity, your velocity vector was in the wrong direction, necessitating either the use of a lot of delta v to turn the velocity vector or else massive wings to do it aerodynamically.  Of course, the wings have strutural, aeroheating, and drag problems.  This hurts performance, which necessitates more velocity in the wrong direction, which requires bigger wings, etc, etc...

[mlorrey]

Boron is element 5, so boron oxide and boron carbides are both lighter molecules than carbon dioxide, so we already have that.

The Merlin/boron-slurry issue is acknowledged to be something requiring study. I'm proposing building and flight testing first with RP-1, then experimenting with boron slurrying.
Some options:
One thing I've been looking at recently is to carry a third propellant, hydrazine, which would power the turbopump with the LOX (thus eliminating the particulates issue with boron).
Another concept is precombusting hydrazine with the boron/RP-1 slurry. This binds the boron to nitrogen from the hydrazine, and frees up all the hydrogen for LOX combustion. There have been hydrazine/borane precombustion tests done in the past.
Assuming that we can use a nanoparticle boron in the slurry, the particulate issue will be eliminated, then is merely the thermal issues. Thermally, the Merlin engine has an ablating nozzle. If the current nozzle is insufficient for the task, this will merely necessitate a heavier nozzle with thicker ablator. I'm not sure what material the Merlin is using in its nozzle, but if it's not using hafnium diboride, that may be something to consider.

[mlorrey]

aero313 - 30/5/2006  4:10 PM

mlorrey - 29/5/2006  4:25 PM

My primary arguments are that air augmentation is needlessly being given short shrift by NASA, that NASA, as seems sadly typical, picked two wrong horses (scramjets and hydrogen, rather than ramjets and hydrocarbons) when it did study air augmented launching. The pattern of seeming to pick the wrong horse in so many key decisions (Thiokol over Aerojet on SRB contract, LH2 vs Hydrocarbons, Scramjet vs Ramjet, aluminum airframe vs hotframe on Shuttle, Blunt body vs SHARP TPS, composite vs metal tank on X-33, etc) just seems too glaring to be an accident to me.

Our company was one of the DARPA/RASCAL Phase 1 teams.  We studied airbreathing ascent vs. rocket (albeit with mythical MIPCC performance vs. mythical ramjet or scramjet performance).  The bottom line was that to get enough velocity to make the airbreather useful, you needed to fly very fast while down low in the atmosphere.  Aero heating was a perpetual problem.  More to the point, once you attained that useful velocity, your velocity vector was in the wrong direction, necessitating either the use of a lot of delta v to turn the velocity vector or else massive wings to do it aerodynamically.  Of course, the wings have strutural, aeroheating, and drag problems.  This hurts performance, which necessitates more velocity in the wrong direction, which requires bigger wings, etc, etc...

What company was that?

With heating, I'm looking at implementing the SHARP materials that nobody has yet been looking at using in their proposals. Its performance used as leading edge and nosecone materials allows mach 7 at sea level and mach 11 at 100,000 ft.

As I noted in my website, part of the problem with the other proposals I've seen out there is that they try to go too far with air breathing, which IMHO seems intended to prove its inadequacy. I have little to no interest in operating at speeds and dynamic pressures necessary for supersonic combustion. I am entirely convinced by the numbers that conventional ramjets, augmented with MIPCC at hypersonic speeds, allow for acceleration up to mach 8-9 by 100k-120k ft altitudes. We don't need 1800 lb/ft^2 dynamic pressures, 1000-1200 at most would be sufficient.

Deltav requirements: it is a bigger dv drain to waste lots of fuel and oxidizer accelerating straight up, then changing direction by 90 degrees, than to air breath up to mach 8 and 100k ft, pop a zoom maneuver to 250kft, then keep on trucking downrange. A third of a pure rocket launchers fuel gets consumed below mach 2 and 100kft.

I've looked at DARPA's RASCAL program, and IMHO it was designed to fail with both obsolete turbine engine technology (F-100's with terrible T/W ratios), burdensome low Isp/high mass upper stage technology (hybrids), as well as demands for unrealistic mission requirements like 30 minutes of loiter time before zoom, 300 mile ferry before launch, etc. As far as I'm concerned, DARPA duped a lot of companies with RASCAL with a Sisyphusian chore.

Nor is ramjet performance "mythical", though it may as well be, given that most all of the real experts from Marquardt are retired or dead, and engineering schools ignore the technology. Ramjets have a rather well established history of performance that some engineers consider mythical only because they've never been exposed to them.

[Jim]

Spirit - 30/5/2006  4:08 PM

Jim - 30/5/2006  1:54 AM

Since there is only 4 launch service providers in the same class, Sealaunch, Atlas, Ariane, and Proton (Delta is not available commercially)

What about the Japanese, Chinese and Indian launchers?

They are not commercially active.

[aero313]

mlorrey - 30/5/2006  5:54 PM

With heating, I'm looking at implementing the SHARP materials that nobody has yet been looking at using in their proposals. Its performance used as leading edge and nosecone materials allows mach 7 at sea level and mach 11 at 100,000 ft.

Never heard of SHARP.  Is this material real?  Is there flight data?

Deltav requirements: it is a bigger dv drain to waste lots of fuel and oxidizer accelerating straight up, then changing direction by 90 degrees, than to air breath up to mach 8 and 100k ft, pop a zoom maneuver to 250kft, then keep on trucking downrange. A third of a pure rocket launchers fuel gets consumed below mach 2 and 100kft.

Have you actually looked at a trajectory for a conventional ground-launched vehicle?  Launch vehicles don't accelerate "straight up" then hang a hard right.  They follow a gravity turn that minimizes gravity loss and aero drag while gaining altitude and providing tangential acceleration.  Vehicles with high T/W (typically solids like Athena and Taurus) will start to pitch over almost immediately.  Liquid vehicles with low T/W need to get some altitude first.

The zoom maneuver is the problem - it's the definition of "gravity loss".  It doesn't matter if you get your acceleration from mythical MIPCC turbine engines or mythical ramjet engines.  You still need to turn the velocity vector.  If you're starting at Mach 8 in somewhat horizontal flight, you need a lot of wing area and very high dynamic pressure to do that zoom.  By the way, flight path angle is a bigger driver on system performance than either altitude at the end of the zoom or velocity (which, by definition, you lose as you zoom).  Every one of the RASCAL teams (except Space Launch) told DARPA that it made more sense to release the rocket upper stages lower and at higher dynamic pressure than what Program Manager Preston Carter wanted.  Space Launch told him what he wanted to hear (instead of what was real) and therefore got selected for Phase 2.   That's also why the program augered in.

So what is your velocity and flight path angle at 250kft?

I've looked at DARPA's RASCAL program, and IMHO it was designed to fail with both obsolete turbine engine technology (F-100's with terrible T/W ratios), burdensome low Isp/high mass upper stage technology (hybrids), as well as demands for unrealistic mission requirements like 30 minutes of loiter time before zoom, 300 mile ferry before launch, etc. As far as I'm concerned, DARPA duped a lot of companies with RASCAL with a Sisyphusian chore.

Not all the teams proposed F-100s or hybrid upper stages.  As I said, Space Launch was the only team that parroted back what Preston Carter wanted to hear.   As I noted above, none of that (nor the ferry requirement, nor the loiter time) changes the fact that the high velocity turn and zoom are very difficult to do in a way that doesn't cause more harm than good.

Nor is ramjet performance "mythical", though it may as well be, given that most all of the real experts from Marquardt are retired or dead, and engineering schools ignore the technology. Ramjets have a rather well established history of performance that some engineers consider mythical only because they've never been exposed to them.

Any airbreather that can go from the ground to Mach 8 or 11 (or whatever it is) at 100kft is mythical.  

Frankly, contrary to what the conspiracy theorists seem to believe, those of us who have actually built and operated space launch vehicles aren't in it just to screw the government.  Physics makes putting something in orbit damn hard.  For some reason, only people who haven't done it seem to be the ones who think there's some magic technology that makes it trivial.

[yinzer]

mlorrey - 30/5/2006  2:54 PM
Deltav requirements: it is a bigger dv drain to waste lots of fuel and oxidizer accelerating straight up, then changing direction by 90 degrees, than to air breath up to mach 8 and 100k ft, pop a zoom maneuver to 250kft, then keep on trucking downrange. A third of a pure rocket launchers fuel gets consumed below mach 2 and 100kft.

Really?  The Atlas V 401 (nearly the simplest launch vehicle possible) reaches the 1-2 staging event with a total energy (velocity + altitude) of around 4740 m/s, from a first stage that has a vacuum delta-v capability of 6200 m/s, for total aerodynamic, back pressure, and gravity losses of 1460 m/s.

It's not entirely clear from your web page, but assuming that all the LOX goes to feed the Merlin, along with enough RP-1 to keep the stock mixture ratio, leaves you with 28,000 pounds or so of RP-1 and peroxide for the MIPCC ramjet.  With an average ISP of 1500 seconds, this would give a theoretical delta-V of nearly 5000 m/s, but Mach 8 is only about 2500 m/s.  So the losses are large.  You are also clearly not going to get 450 seconds of Isp out of the Merlin as shown on your diagram.

The "air breather's burden" - that is the integrated drag due to lift over the longer acceleration period that comes with low installed T/W of airbreathing engines, is a very serious problem that does not show up in many first-cut conceptual studies.  I remain unconvinced that high performance air breathing (as opposed to subsonic first stages) really gives you any detectable advantage over just making the rocket bigger.

[bad_astra]

I couldn't find the expected mass fraction of your Hyperdart vehicle from your website. Curious to know what you expect it to be.

[mlorrey]

aero313 - 30/5/2006  5:52 PM

mlorrey - 30/5/2006  5:54 PM

With heating, I'm looking at implementing the SHARP materials that nobody has yet been looking at using in their proposals. Its performance used as leading edge and nosecone materials allows mach 7 at sea level and mach 11 at 100,000 ft.

Never heard of SHARP.  Is this material real?  Is there flight data?

There is classified flight data from the X-24C, from the classified test articles from Copper Canyon/HAVE SPACE, but little of that is public, though various reports, including one from RAND, discuss it tangentially. NASA sponsored a public domain project at U of Montana in 2000-2001, which built a suborbital test article, but had further funding cancelled by NASA for one bogus reason (claiming there was a lack of adequate sounding rockets, which is wrong, they had one lined up from Wickman Aerospace) and one halfway legit reason (a claim that it was "duplicate research", but duplicate of what? Classified research?).

The only openly acknowledged use of SHARP materials I've seen was on small fins for a highly maneuverable MIRV design. I'll dig up a reference to it, though I may have referenced it on my website.

Deltav requirements: it is a bigger dv drain to waste lots of fuel and oxidizer accelerating straight up, then changing direction by 90 degrees, than to air breath up to mach 8 and 100k ft, pop a zoom maneuver to 250kft, then keep on trucking downrange. A third of a pure rocket launchers fuel gets consumed below mach 2 and 100kft.

Have you actually looked at a trajectory for a conventional ground-launched vehicle?  Launch vehicles don't accelerate "straight up" then hang a hard right.  They follow a gravity turn that minimizes gravity loss and aero drag while gaining altitude and providing tangential acceleration.  Vehicles with high T/W (typically solids like Athena and Taurus) will start to pitch over almost immediately.  Liquid vehicles with low T/W need to get some altitude first.

The zoom maneuver is the problem - it's the definition of "gravity loss".  It doesn't matter if you get your acceleration from mythical MIPCC turbine engines or mythical ramjet engines.  You still need to turn the velocity vector.  If you're starting at Mach 8 in somewhat horizontal flight, you need a lot of wing area and very high dynamic pressure to do that zoom.  By the way, flight path angle is a bigger driver on system performance than either altitude at the end of the zoom or velocity (which, by definition, you lose as you zoom).  Every one of the RASCAL teams (except Space Launch) told DARPA that it made more sense to release the rocket upper stages lower and at higher dynamic pressure than what Program Manager Preston Carter wanted.  Space Launch told him what he wanted to hear (instead of what was real) and therefore got selected for Phase 2.   That's also why the program augered in.

So what is your velocity and flight path angle at 250kft?

I'm expecting to be at mach 8 at 100k-120k ft in a 20-30 degree climb. I'm not looking for a huge zoom angle. The NF-104 had to zoom at 70 deg+ because it entered the zoom at 30,000ft. At 120kft, there is a lot less atmosphere in the way, so a shallower zoom angle can be used, essentiall transitioning from horizontal flight into the upper phase of a traditional gravity turn trajectory. Its more of an s curve than a real zoom. One reason is that with SHARP materials, hypersonic drag is much less due to leading edge sharpness.

I'm not doing flight simulations yet to determine optimum flight path by Monte Carlo methods. I'm currently working on the design in X-Plane.

I've looked at DARPA's RASCAL program, and IMHO it was designed to fail with both obsolete turbine engine technology (F-100's with terrible T/W ratios), burdensome low Isp/high mass upper stage technology (hybrids), as well as demands for unrealistic mission requirements like 30 minutes of loiter time before zoom, 300 mile ferry before launch, etc. As far as I'm concerned, DARPA duped a lot of companies with RASCAL with a Sisyphusian chore.

Not all the teams proposed F-100s or hybrid upper stages.  As I said, Space Launch was the only team that parroted back what Preston Carter wanted to hear.   As I noted above, none of that (nor the ferry requirement, nor the loiter time) changes the fact that the high velocity turn and zoom are very difficult to do in a way that doesn't cause more harm than good.

I know that some teams proposed things outside the DARPA reference design (even one that proposed a less drastic mod of an F-106 than mine). The problem is that when an agency issues a "reference design", it means that NIH is in charge and you'd better tell them what they wanted to hear to win the contract, whether or not it will actually work. This is why I concluded that RASCAL was intended to fail.

Nor is ramjet performance "mythical", though it may as well be, given that most all of the real experts from Marquardt are retired or dead, and engineering schools ignore the technology. Ramjets have a rather well established history of performance that some engineers consider mythical only because they've never been exposed to them.

Any airbreather that can go from the ground to Mach 8 or 11 (or whatever it is) at 100kft is mythical.  

Frankly, contrary to what the conspiracy theorists seem to believe, those of us who have actually built and operated space launch vehicles aren't in it just to screw the government.  Physics makes putting something in orbit damn hard.  For some reason, only people who haven't done it seem to be the ones who think there's some magic technology that makes it trivial.

I wasn't aware that X-43 was mythical.

[mlorrey]

bad_astra - 30/5/2006  7:38 PM

I couldn't find the expected mass fraction of your Hyperdart vehicle from your website. Curious to know what you expect it to be.

It's plainly posted in the design section. http://www.lorrey.biz/x-106/conversion.html#specs

.84 is quite sufficient for a high average Isp vehicle.

[hop]

I wasn't aware that X-43 was mythical.

The X-43 got there by way of a large aircraft and a solid rocket booster, not airbreathing, so it doesn't really support your point.

[meiza]

Interesting concept. I feel it's trying to push too many boundaries at once to be succesful though. Maybe a suborbital demonstrator first would be useful, now with the suborbital tourism and all?

I have two questions at first:
What is the principal advantage of a hot structure compared to a cold one? What's the stiffness per weight difference between cold aluminium and hot titanium?

Are the propellants and exhaust products toxic? If yes, don't they increase operating costs a lot? They also scale with operation, meaning more frequent flying causes more problems. Also, don't additives clog up the engine, which is bad for a reusable craft?

I've read before about the SHARP materials and they seem useful to your case. For the readers, there's some material here at Wickman Aerospace: http://www.space-rockets.com/sharp.html.

[mlorrey]

meiza - 1/6/2006  6:07 AM

Interesting concept. I feel it's trying to push too many boundaries at once to be succesful though. Maybe a suborbital demonstrator first would be useful, now with the suborbital tourism and all?

I fully expect to do a lot of suborbital flights. It is intended to be capable of both suborbital and orbital, following a progressive flight test program to determine the actual limits of the vehicle. I am hoping for it to become the T-38 or Cessna 172 of space flight. It may not make orbit, my concepts may be too grand for the technology, but one thing I know is that we'll never find out if we don't actually build it and fly it. Even if it is only a suborbital vehicle, it should be able to launch upper stages to orbit, and make money on suborbital travel as well, and not just tourism, but high speed transcontinental executive travel for people whose time is worth $10k a minute.

I have two questions at first:
What is the principal advantage of a hot structure compared to a cold one? What's the stiffness per weight difference between cold aluminium and hot titanium?

A hot structure allows the entire airframe to function as a heat sink for reentry generated thermal energy. A cold structure of aluminum limits this due to the low temperature tolerance of the metal. Thus, with an aluminum airframe (like the Shuttle has), you need your TPS to absorb and re-radiate as much of the heat as fast as possible to the outside.

One of the problems with the RCC on the shuttle is that it conducts heat better than it reradiates it, which means the shuttle needs heavy inconel mount points for any RCC structure to isolate it from the aluminum airframe, and silica tiles under the RCC as well. SHARP materials have much more effective re-radiating capability than RCC. Another problem is that the shuttle TPS is not a contributing part of the airframe structure, thus it is dead weight. It is essential that as much TPS be load bearing as possible.

Stiffness to weight for aluminum and titanium is essentially the same, the difference between the two being the much greater temperature range of titanium, so aluminum's stress strain limit is at a much lower temperature than that of Titanium.

Are the propellants and exhaust products toxic? If yes, don't they increase operating costs a lot? They also scale with operation, meaning more frequent flying causes more problems. Also, don't additives clog up the engine, which is bad for a reusable craft?

Boron oxide is not toxic (unless you go and swallow quantities of it, of course). When it falls in the ocean it combines with natural sea salts to form borax, a common natural compound.

Pure atomic boron is currently rather pricey, due to lack of a market (other than black spaceplanes that don't exist!), but fuel costs normally are less than 1% of the cost of an expendable launchers launch cost, but the cost of RP-1 and LOX is rather miniscule.

Exhaust precipitates are an issue that needs researching. Experimentation with borane fuels (boron-hydrogen compounds) in the 60's found problems with Boron Oxide depositing on the exhaust nozzle, however this may be a feature, not a bug, given the high thermal tolerance of boron oxide, it may help protect the engine from the higher combustion heat of the boron additive. The nozzle of the Merlin is of course ablative, so at some point, the material flakes off, carrying any depositions with it. It is an issue to study.

One thing I am concerned about is deposits on turbopump surfaces. For this reason, and given that we have several RP-1 tanks on the vehicle, I am considering carrying one tank of plain RP-1 without boron, which will fuel the turbopump, thus negating the possibility of depositions there.

However, as I said previously, and on the website, we'll start off with plain RP-1 combustion and run a test program with boron, and possibly other additives (I've learned that fluorine in the LOX negates the boron deposition problem that others have had in tests with ramjets).

I've read before about the SHARP materials and they seem useful to your case. For the readers, there's some material here at Wickman Aerospace: http://www.space-rockets.com/sharp.html.

[yinzer]

What's your total trajectory delta-V?  Where did you get the rocket specific impulse figure of 450 seconds from?  Are your ramjet weights installed, or bare?  How does the thrust / specific impulse change with increasing mach number?  What's the airbreathing T/W?  What hypersonic L/D ratios are you assuming for the horizontal acceleration and during the pitchup maneuver?

[mlorrey]

yinzer - 1/6/2006  5:24 PM

What's your total trajectory delta-V?

This is highly dependent upon the final Isp performance of the ramjets, among a large number of other things. By rough numbers, I'm convinced it can get near orbit at a minimum, and as Max Hunter once said, "given that, I can nickle and dime it into orbit".

Where did you get the rocket specific impulse figure of 450 seconds from?  

From USAF documents, also from a US patent on nanopowdered boron additive.

Are your ramjet weights installed, or bare?  How does the thrust / specific impulse change with increasing mach number?

Have not even started cutting metal yet, so these answers have a long ways yet to even be properly measured.

What's the airbreathing T/W?

Ramjet T/W historically ranges from 20 (twice that of turbojets/turbofans) to as high as 50 (about equal to that of LH2 rocket engines) or more. Given the current state of the art, I'm confident we can reach the top end of that range. MIPCC experiments show a thrust multiplier of as much as 2 could increase T/W of the ramjets potentially to the range of a kerosene rocket engine (given that is what it becomes in part, this should be unsurprising.) (see NASA/TM-2003-212023)

What hypersonic L/D ratios are you assuming for the horizontal acceleration and during the pitchup maneuver?

Have not yet done any hypersonic fluid dynamics simulations. Given NASA studies of winged vs lifting body vehicles, in which the X-15 had an L/D of about 4.2, while Enterprise showed L/D of 5 (subsonic). Given their much much higher wing loadings over that of the X-106 design, and the lack of true SHARP edging and nose cones on both of these, these should be considered beneath what should be expected of this design, since L/D at hypersonic speeds is largely determined by the size of the leading edge/nosecone...

RASCAL reference mission proposed a zoom angle of 55 degrees, significantly less than the NF-104 tests of the late 50's, with a peak g load of 3.5. Maximum AoA of 20 degrees during pitch up.

[publiusr]

Jim - 29/5/2006  8:55 AM

I am going the throw the B.S. card.  The website is nothing but rants on supposed conspiracies.  

For once--we are in perfect agreement.