X-106 "Christa", the Hyper Dart

[gunsandrockets]

"The 106 has a preexisting 1950's era environmental control system to pressurize the cockpit and provide oxygen to crew up to very high altitude."<br /><br />The F-106 has a pressurized cockpit? You mean a shirtsleeve environment? As far as I was aware all the high flying single-seat aircraft of the 1950's and 1960's such as the U-2, F-104, F-4, Foxbat required the pilots to wear pressure suits for high altitude operations.<br /><br />"This will likely be replaced with either a newer system, possibly something borrowed from the Russians, or else just augmented by peroxide surplus providing both O2 and power for the ECS."<br /><br />Mass estimates? Life support duration estimate?<br /><br />And that doesn't answer my question about what provides electrical power to operate all the spacecraft systems and how much it masses. What is the power source to operate the aerodynamic flight controls? Is it a hydrogen peroxide fueled generator?

 

[mlorrey]

"44 kg/m^2"??? That is an area measure, not a volume density measure (which would be kg/m^3). <br /><br />44kg/m^3 is actually extremely light. For example, water is 1000 kg/m^3. If your number is accurate, C-C should be less dense than liquid hydrogen.<br /><br />In fact, not only are you wrong here, you are wrong in claiming that C-C is heavy:<br />http://www.ms.ornl.gov/researchgroups/cmt/CFCMS/TMS.HTM<br />http://www.ultramet.com/u2000a.htm<br /><br />1.6 g/cm^3 appears to be a good number for reinforced C-C, which is 60% denser than water. By contrast, Inconel is over 8 g/cm^3 and titanium is 4.5 g/cm^3. Aluminum is 2.7 g/cm^3.<br /><br />Steel was used in the leading edges, control surface structures, and around the turbojet, as well as other structural elements. It was used at the time because titanium was not yet available in sufficient affordable quantities for a 350 aircraft construction contract in the mid 1950's.<br /><br />Actually, looking at this, I may consider using C-C in a lot more of the vehicle surface and dump Ti entirely.<br /><br />As for 'blankets', that is so 1970's. As I stated previously, which you apparently missed, was I was expecting to coat the Ti with a boron carbide or boron nitride coating, which offers extremely high thermal protection.<br /><br />I would suggest that:<br /><br />a) a fixation upon silica bricks is unhealthy for the imagination<br />b) there are far better TPS' than silica<br />c) I'm not going to fund a jobs program for a bunch of silica bricklayers.<br />d) I'd rather make the entire aircraft of low density C-C than hire one silica bricklayer.

 

[mlorrey]

The F-106 was "The Ultimate Interceptor". Of course it had a pressurized cockpit. It was designed to intercept soviet bombers at high altitude. There are, however, limits to normal fighter aircraft pressurization capabilities.<br /><br />Typically, such pressurization systems I worked on (F-106, F-111, F-15) seal up at 8500' alt, and hold that pressure to 25,000'-35,000'. Above that level, the pressure scales and O2 levels are increased to keep the O2 partial pressure constant.<br /><br />I haven't decided yet what works best for the power. I'd like to look at developing a type of fuel cell that generates power from decomposing the peroxide, without having moving parts involved. This would provide power, O2 for the crew, and retain the spent peroxide as water which can be reused to cool the ramscoop during reentry.<br /><br />As I said, I haven't figured this all out yet, but I'm certain with modern technology it can be done with less mass than a 1950's era fighter ECS. Nor will asking this question again in two days get you a valid answer, this is a significant engineering question that I intend to resolve properly. I have already budgeted what I view as sufficiently excessive mass to cover all the bases on this to provide at the very least enough power and life support for a few orbits. As the mass is there, the details can be worked out.

 

[mlorrey]

<blockquote><font class="small">In reply to:</font><hr /><p>"I expect takeoff to 450 kt would be accomplished either by launch sled, tow-launch (as Kellyspace demonstrated with the 106 already), airlaunch (a la WK2 or An-225) or through use of the Merlin engine at takeoff as well, or possibly just using some JATOs." <br /><br />The Kellyspace tow launch concept always seemed like an excellent idea to me. <br /><br />You should stop refering to your vehicle as a SSTO however, unless the Merlin gets the vehicle up to ramjet speed. And if you do use the Merlin for takeoff, serious recalculation of performance will be needed. For one thing the Merlin will lose a lot of ISP operating at low altitude and use up a lot of fuel to accelerate the vehicle up to ramjet speed/altitude. <p><hr /></p></p></blockquote><br /><br />This is a valid criticism, though I don't consider a small amount of help at take-off to violate the SSTO status. I'm not looking to tow it to 45,000 feet, though that might improve performance further. Getting it to 450 kt at low altitude is fine to me. Given the T/W ratio of the fully loaded vehicle, the Merlin should be able to get the vehicle to 450 kt with less than 30 sec of 100% burning, from a dead stop. Less time if a few jatos are used to up the T/W on takeoff. However, use of Merlin for take-off to 450 kt still lies entirely within the allowable performance of the given mass fraction. <br /><br />If I went to an air launch like Kellyspace proposed, I'd save 5% on the mass fraction requirements.<br /><br />As for someone else's assertions that this sort of vehicle was not previously proposed, this is also false. The X-24C, also known as the L-301, was of similar design, and was cancelled by Carter in 1977 along with the B-70 and a slew of other hypersonics programs. Test articles had already been constructed and flown. My inside sources refuse to confirm or deny that this vehicle ever made it to orbit under a black classification...<br />

 

[mlorrey]

<blockquote><font class="small">In reply to:</font><hr /><p>Assuming your modified Merlin engine would achieve the higher ISP you estimate, you can't achieve that ISP with the same rating of 80.000 lb of thrust. A 40% increase of ISP while maintaining the same thrust would subject the Merlin engine to double the energy it normally operates under. Operating under the normal amount of energy plus increased ISP means a thrust reduction of 30% to about 56,000 lb of thrust. <br /><br />It's the pitiless equations of rocketry at work. Higher ISP engines have lower thrust to weight ratios. You yourself noted one such example, "While H2 gets you higher Isp, its thrust is lower for a given engine size." <p><hr /></p></p></blockquote><br /><br />A fair criticism. H2's problem is its density, IMHO, and the resulting poor exhaust pressures and densities. Kerosene offers higher thrust because it is denser, and its exhaust is denser. Boron slurried kerosene should provide an exhaust density similar to kerosene, while at a higher temp and pressure. This may require a redesign of the ablative nozzle and thus may add some weight to the nozzle. Fair enough. It isn't a question that I think can be answered either way at this point.

 

[gunsandrockets]

I got the weight figure from this link, the same link I posted before.<br /><br />http://www.centennialofflight.gov/essay/Evolution_of_Technology/TPS/Tech41G2.htm<br /><br />Which has a table showing that 38.0 square meters of the Shuttles surface is protected by RCC, and this RCC has a mass of 1697.3 Kg.

 

[gunsandrockets]

"In fact, not only are you wrong here, you are wrong in claiming that C-C is heavy:"<br /><br />I claimed that RCC TPS is heavy. ANY TPS is going to be heavy in comparison with normal construction, that's the nature of the problem of TPS.

 

[mlorrey]

Okay, but , as I said, that is an area measure, with no regard to material thickness. 44 kg/m^3 would be less dense than hydrogen. The material could be 44kg/m^2 at a density of 1.6 g/cm^3, which would specify a certain material thickness.<br /><br />RCC, by its density, demonstrates that it is less dense than inconel, titanium or even aluminum. The links I provided also show superior structural strength at higher temps than any of the above. Furthermore, eliminating the aluminum skin AND the silica TPS, which has no structural strength, with ONE material that does both will save significant weight.<br /><br />RCC was not used extensively on the shuttle because it was built in the 1970's, when RCC was in its infancy, very expensive, both in materials, labor, as well as tooling. The budget constraints we've all talked about that constrained its design from the Starclipper to the present abortion also limited the ability of engineers to spend money on RCC.

 

[mlorrey]

The thickness of the material used in the shuttle's RCC components is dictated by two things: amount of internal structure, and reentry dynamic pressure. Given the X-106 is half the length and 1/10th the empty mass of the orbiter, both the dynamic pressure regime, and the internal structural requirements for the 106 is much less than the orbiter. I am expecting a peak dynamic pressure of approximately 1000-1200 psf at the nose and leading edges.

 

[josh_simonson]

The CC must be thick enough that it sufficiently insulates the craft to avoid damaging the internal parts during re-entry. That's probably the main issue that determines the TPS thickness, not structural strengh.

 

[mlorrey]

On the contrary, RCC is so effective a thermal shield, and becomes stronger and more rigid with higher temp, that it is quite clear that its thickness requirements on the shuttle leading edge and nose cone are clearly structural. This is made even clearer by looking at cross sections of these areas, to see that there is virtually no reinforcing structure underneath the RCC panel. Nothing but empty space, so each RCC panel is treated as if it is also its own hot structure. The nose cone of the shuttle is similar: virtually no structure backing up the nose cone.

 

[gunsandrockets]

"Nothing but empty space, so each RCC panel is treated as if it is also its own hot structure. The nose cone of the shuttle is similar: virtually no structure backing up the nose cone. "<br /><br />Nothing, aside from insulative material to shield the interior of the vehicle from heat radiating from the interior surface of the hot RCC. From NASA comes this...<br /><br />"Since carbon is a good thermal conductor, the adjacent aluminum and the metallic attachments must be protected from exceeding temperature limits by internal insulation. Inconel 718 and A-286 fittings are bolted to flanges on the RCC components and are attached to the aluminum wing spars and nose bulkhead. Inconel-covered cerachrome insulation protects the metallic attach fittings and spar from the heat radiated from the inside surface of the RCC wing panels."<br /><br />"The nose cap thermal insulation ues a blanket made from ceramic fibers and filled with silica fibers. HRSI or FRCI tiles are used to protect the forward fuselage from the heat radiated from the hot inside surface of the RCC."<br /><br />http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_sys.html

 

[mlorrey]

One thing to consider is to ask how hot this vehicle is going to get vs the Orbiter. You can tell by the wing loading. They both have a similar planform, but the Orbiter is about twice the length and wingspan, so it should have four times the surface area as the X-106. Max landing weight of the orbiter is 230,000 lb, while the X-106 should be I believe 18,600 lb. Even if this weight estimate is low and a TPS is going to add a few thousand pounds to it, the reentry weight will never exceed 1/10th that of the Orbiter, while having 1/4 the surface area. This means the Orbiter has 2.5 times the wing loading, at a minimum, of the X-106, and should have a proportionately higher dynamic pressure on the nose and wing surfaces on reentry. This means the X-106 is going to have significantly lower thermal issues than the Shuttle, more in line with that of the X-33 and the SR-71.

 

[spacefire]

no way on that. It has to slow down from orbital speeds unlike both the X33 and the Sr71.<br /><br />I told you, it's gonna be way heavier than you think, but with inflight refueling you might make it work. <div class="Discussion_UserSignature"> <p>http://asteroid-invasion.blogspot.com</p><p>http://www.solvengineer.com/asteroid-invasion.html </p><p> </p> </div>

 

[mlorrey]

It has plenty of fuel to not need in flight refuelling. That is not an issue. The issue is the loading. This is essentially a measurement of the 'density' of the vehicle. Less dense vehicles slow down faster in the thinnest part of the atmosphere and do not experience the same stresses and heating that denser objects do. This is why meteorites burn up but the shuttle does not even melt (most of the time): the shuttle is not solid metal or rock, it is less dense due to all the empty space in it.<br /><br />The X-33, Venturestar, and the X-106 all have far more empty space for the amount of mass than the shuttle orbiter does on reentry. The 106 isn't as low as the Venturestar because it doesn't waste space on using LH2, but it is still, unlike the shuttle orbiter, bringing home all of its fuel tanks, empty. Due to this, the wing loading is 1/5th that of the shuttle orbiter, it will therefore slow down much faster. I've also designed it with much larger flaps and Rutan-style feathers, so the X-106 will be able to begin its reentry at a much higher angle of attack, and with thus slow down even faster. With the lower wing loading, no only will it slow down faster, in the thinner part of the atmosphere, it will be more stable in that high AOA.

 

[mlorrey]

All this inconel structure is the real problem: how to mate the RCC to the low temp aluminum frame without melting it. This is a significant and serious issue, but RCC is considered 'dense' compared to the other shuttle orbiter TPS only because the other TPS items are not structural components as well as TPS. If you figure in the weight of the materials being used as structure to support the non-RCC TPS, you'd likely find they are in the end heavier.<br /><br />So, the question is: how to mate aluminum frame to RCC panels, leading edges, and control surfaces? Would TitanAl be good enough? Its much less dense than Inconel and has high temp properties, depending on the mixture of the alloy.

 

[mlorrey]

Okay, from here on, I'd like to see people who have objections be posting some real analysis: numbers, math, etc to support arguments, and not vague suppositions based on the shuttle, since this vehicle will not in any way be "like" the shuttle. Its volume density is much lower, its mass fraction and average Isp are much better, among other things.

 

[mcbethcg]

Too bad.<br /><br />The GTX design used as an example incidentally had 3 ramjets that were much larger. I still say he is overstating the thrust of his ramjet.

 

[mlorrey]

Did you even look at the thrust of the ramjets on the GTX? What was their thrust? Stop making unsupported statements without any facts.<br /><br />The GTX reference vehicle is 236,000 lbs. Its GLOW is 690,000 lbs. It is 211 feet long. With a liftoff T/W ratio of 1.8, this means the GTX engines will have a thrust of about 1.25 million lb. This means over 400,000 lbs of thrust per engine on the GTX. Each engine has a radius of 7.33 feet, or or a cross section area of about 42 square feet, which means about 10,000 lbs of thrust per square foot of cross sectional area.<br /><br />In comparison, the ramjet engine of the proposed X-106 will have half the diameter, therefore 1/4 the cross sectional area. If it were built according to the standard of the NASA RBCC engine, it would produce 100,000 lbs of peak thrust, which is twice what I am predicting as an average. Thus, my estimate is entirely within the realm of completely realistic and conservatively reasonable possibility, once again. I will refrain from snide comments about doing one's homework... <br /><br />That my design is completely conservative yet should orbit four times the payload of the GTX reference vehicle, at a minimum, while being a small fraction of the GTX GLOW, is clear evidence of the absurd lengths that NASA will go to use LH2 fuel for absurd reasons that have absofreakinlootly nothing to do with the alleged so-called 'superior performance' of LH2. I think the comparison on this thread clearly demonstrates the inferiority of LH2 as a rocket fuel.<br /><br />That NASA continues to be obsessed with LH2 is clearly NOT for performance reasons, it has to be for other reasons that have nothing to do with space flight.

 

[mlorrey]

As for concerns about the TPS:<br /><br />a) the STS Orbiter's reentry wing loading is a minimum of 95.6 lb/ft^2. The X-33, and Venturestar, which were designed to use metallic TPS and operate at a much lower peak temperature, had a design wing loading of about 30 lb/ft^2. The X-106 has a wing loading of 18-19 lb/ft^2. Even if I loaded it up with a 10% mass increase for additional TPS (at which time it would still have a mass fraction in excess of that needed to reach orbit), it would still be about 2/3 of that planned for the X-33 and Venturestar.<br /><br />For this reason, I am expecting a very tolerable thermal regime for this vehicle. Nose cone and leading edges will remain RCC, given their issues with both boundary layer and shock wave thermal effects, but there will be a more extensive study as to whether RCC is needed for windward or undersurface body panels, or whether titanium aluminide panels mounted to the airframe with Ti fasteners and ceramic isolation washers would be sufficient. Given this information, I suspect that the Ti/AL will win out, but I'll need to make a mass comparison with RCC. If RCC is kept on windward surfaces, it will likely be attached similarly, with titanium fasteners and ceramic isolation washers, though airframe stringers and bulkhead flanges may be replaced with titanium as well. While the titanium is denser than aluminum, its strength/weight ratio is better, so Ti parts would be smaller for the same strength. <br /><br />As for concerns about vehicle loading:<br />Wing loading at launch is just over 83 lb/ft^2, significantly less than the max allowed wing loading of a fully loaded F-15 Eagle. Thus this vehicle is neither overloaded at takeoff, yet has high enough mass fraction (since it isn't carrying ordnance or a turbine engine or a weapons system) to achieve orbit for the designed average effective Isp.

 

[gunsandrockets]

"Okay, from here on, I'd like to see people who have objections be posting some real analysis: numbers, math, etc to support arguments, and not vague suppositions based on the shuttle,"<br /><br />Here's the problem. Your vehicle airframe is larger than the F-106, and in addition has a passive-reusable-TPS. Despite this you claim the the weight of the vehicle airframe would be LESS than the the weight of the original F-106. That's an astonishing claim that you need to prove with hard numbers. Because a passive-reusable-TPS that makes a vehicle LIGHTER sounds like magic TPS to me.

 

[gunsandrockets]

"RCC is considered 'dense' compared to the other shuttle orbiter TPS only because the other TPS items are not structural components as well as TPS."<br /><br />Shuttle tiles and FRSI blankets are light in comparison with RCC TPS because they have insulative qualities in addition to being temperature resistant.<br /><br />RCC not only conducts heat very well, it radiates heat very well. Which means behind RCC structure there must be placed tiles or other material to insulate the interior spacecraft-structure from the heat radiated from the interior-surface of the RCC. <br /><br />http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_sys.html<br /><br />"The nose cap thermal insulation ues a blanket made from ceramic fibers and filled with silica fibers. HRSI or FRCI tiles are used to protect the forward fuselage from the heat radiated from the hot inside surface of the RCC."

 

[gunsandrockets]

The claimed maximum speed of the X-106 under ramjet power is Mach 6. But I wonder if the claimed thrust of 50,000 lb is enough to push an aircraft with a GLOW of 86,000 lb and a span of 38 feet as fast as Mach 6.<br /><br />The Bomarc B ramjet powered missile had a maximum speed of Mach 3, a T/W of 1.5, and a span of 18 feet.<br /><br />The YF-12 had a span of 55 feet, a GLOW of 127,000 lbs, 64,000 pounds of thrust and a maximum speed above Mach 3.<br /><br />The RASCAL which would use MPICC, have a span of 81 feet, a GLOW of 80,000 pds, and 100,000 lbs of thrust, yet a maximum speed only slighter greater than Mach 3.<br /><br />http://www.aviationnow.com/avnow/news/channel_awst_story.jsp?id=news/09223top.xml<br /><br /><br />From these examples I can't see how the X-106 can reach Mach 6.<br /><br />

 

[gunsandrockets]

"However, use of Merlin for take-off to 450 kt still lies entirely within the allowable performance of the given mass fraction." <br /><br /><br /><br />In an earlier post you said the ramjet would consume 3/5 of the fuel. So I ran the numbers; using the remaining fuel and all the LOX and assuming the vehicle starts the Merlin burn at Mach 8, a final weight at burnout of 18,300 lbs, and an ISP of 457, I get a final result a little better than 8.5 km/s.<br /><br />I don't see any excess margin for using the Merlin below speed Mach 8.

 

[mlorrey]

Lets see here: the YF-12 had a GLOW T/W of .5 yet reached above Mach 3?<br /><br />The RASCAL had a much higher GLOW in the reference design than 80k, more lik 180k.<br /><br />Try looking at the sweep angle of the respective vehicles, also. The X-106 will have a sweep from nose to wingtip of 74 degrees, significantly less than the RASCAL, which is about 45-55 deg. At 74 degrees sweep, there will be no nose shock impingement on the wings until well above mach 5.<br /><br />The difference is that the 50,000 original estimate of mine is low. Going by the RBCC performance, one should figure about 10,000 lbs thrust per square foot of ramjet cross sectional area, which comes out to 100,000 lbs thrust.<br /><br />The difference between the X-106 and these other vehicles is that it combines a pure ramjet with MIPCC, which AFAIK has not been done. This would lower the Isp of the ramjet from above 2000 sec to about 1500 sec, but expand its thrust range. The MIPCC would make the ramjet believe it was operating at mach 3 when it was actually at mach 6.<br /><br />The RASCAL was designed to use four turbofan engines, which have a much lower operating range than a pure ramjet, so MIPCC would expand the lower operating range of the turbojet, but you are also wrong on the RASCAL vehicle, it was intended to reach mach 5-6.