Lorrey Aerospace planning X-106 spaceplane

[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.