Virgin Galactic and Branson. Virgin Galactic Makes First Flight To Space From New Mexico Base As It Prepares To Carry Tourists Next Year https://www.forbes.com/sites/joewalsh/2 … next-year/
As of 2020, the first passenger flight was planned for no earlier than May 2021
https://en.wikipedia.org/wiki/Virgin_Galactic
Orbital Stage Landing
Event Timelines
Event: T+ time (seconds)
Liftoff: 0
MECO: 169
Stage Separation: 171
SES: 176
Booster Touchdown: 495
SECO: 521
Ship Splashdown: 5420 (90 minutes 20 seconds)I added colons, the FCC exhibit has a table. And I added minutes/seconds.
Interesting they use the word "Touchdown" for Booster, but "Splashdown" for Ship. The Booster will "Touchdown" 20 miles from the shore, does that mean they'll attempt a drone ship landing? But Starship will simply splash in the ocean?
If you stand the ride you save a huge amount of jet lag....
]]>SpaceX plans to send Starship to Hawaii via space
The Booster stage will separate approximately 170 seconds into flight. The Booster will then perform a partial return and land in the Gulf of Mexico approximately 20 miles from the shore. The Orbital Starship will continue on flying between the Florida Straits.
It will achieve orbit until performing a powered, targeted landing approximately 100km (~62 miles) off the northwest coast of Kauai in a soft ocean landing."So an upcoming Starship prototype, paired with a Super Heavy booster, will blast off from the SpaceX development site in Boca Chica, Texas. Super Heavy would then separate and land, perhaps on a ship, off shore while Starship continues on to orbit, flying east all the way to Hawaii for a splashdown in the Pacific. The entire flight, from liftoff to splashdown, is expected to last 90 minutes.
SpaceX reveals first orbital Starship flight plan, launching from Texas and returning near Hawaii
The rocket initially launches on a "Super Heavy" booster, which makes up the bottom half of the rocket and stands about 230 feet tall. Together, Starship and Super Heavy will be nearly 400 feet tall when stacked for the launch.
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The gee exposure for any given burn during ascent ranges from just over 1 (no more than about 1.5) to around 3 or 4. The exposure to gees in the 3-4 range is only several seconds long. The entire ascent is only about 5 minutes long. Usually, there are two big burns (two stages).
Reentry gee exposure ranges from under 1 to at most 3 or 4, in a well-designed system. The 3-4 gee exposure is only several seconds long. The entire entry sequence is only about 3-4 minutes long. You actually experience far higher gees as a shock load in the typical ocean splashdown, but that is a pulse only milliseconds long. The short exposure time mitigates the hazard.
People endure higher gees riding roller coasters. The peak exposure is shorter, only a few seconds long. The old wooden roller coaster with sequential loop-the-loops at 6 Flags Over Texas pulled 5 gees at the entry into the first loop, and around 4 gees entering the second.
Entirely untrained and physically-unfit people endured this just fine for many years, and had fun doing it. The bigger threat was actually rain. At 55 mph out in the open, rain in the face can really hurt. That will NOT happen in spaceflight. 6 Flags recommended not riding if you had heart trouble, but that was about the only consideration, other than being tall enough to fit the restraint bar properly.
GW
]]>Something you haven't mentioned is pasenger fitness. Probably only maybe 10% of normal airline passengers would be fit enough to undertaken the flights. They would need to complete a specialist training course that might take 2 or 3 days, I think this woukd become something of a status symbol for younger very wealthy people.
E2E is not going to supplant jet airliners. It is going to supplement them. It will be similar, to the 1930s when a small number of airships and intercontinental propeller planes drew traffic from the ocean-going liners .
Louis,
To be perfectly frank, electric airliners are much closer to engineering reality than a hypersonic rocket-powered VTOL airliner. I'm only pointing out that the "huge challenges" faced are fuel burn rate, engine count, inert mass fraction (which determines vehicle durability), and passenger evacuation are intractable problems. Rocket engines aren't going to become more fuel efficient unless they're no longer rocket engines, the engine count is what provides the necessary thrust to lift off to begin with, and the inert mass fraction is driven by the need for nearly all of the vehicle's weight to be propellant. Since you're not going to reduce the fuel burn rate, use fewer engines than needed to get off the ground, nor change the fact that Starship would be an airliner sitting on its tail, I don't see how this will ever be practical. It may be technologically feasible if every flight is perfectly executed, but no such success rate, relative to conventional airliners, has ever been demonstrated with rockets. Individual jet aircraft have performed more power cycling associated with takeoffs and landings in a single day than any rocket engine has ever demonstrated in terms of total engine starts / stops.
Beyond that, all of those airport runways have already been paid for. A Starship would require a large exclusion zone beyond the tiny pad footprint so that accidents don't throw shrapnel into homes a mile away, so the small size of the landing pad is another moot point. When landing, there's no such thing as "line up and wait" with a rocket, either. If one of these things ever crashes and throws debris onto the next closely-spaced landing pad, is the next inbound pilot supposed to land in a flaming wreckage pile? No significant diversion is possible unless you're very high in the atmosphere. These VTOL rockets come down like meteors.
If that wasn't enough, the window-destroying sonic booms would prevent it from ever diverting to alternate landing pads that aren't far away from population centers. That was the entire point of Boom, and one of two reason it's fundamentally "different" from Concorde. The other reason is a remarkable reduction in fuel burn, relative to lighting afterburners, which Concorde had to do to fly at Mach 2. Boom changes "the sonic boom" into an acoustic wave form that doesn't trash windows or otherwise scare the crap out of people on the ground. It sounds like a deep rumble, rather than the thunder that follows a lightning strike. For various reasons, that's impractical for Starship. Starship will never pass noise abatement ordinances on account of the fact that it's powered by six of the most powerful rocket engines on the planet.
Anyway, the military would probably have some kind of use for it, but they won't be going supersonic over CONUS with it, in the same way that existing regulation forbids them from breaking Mach unless there's some kind of inbound military threat.
What I was involved with, in the rocket and ramjet motors for missiles business, was usually one-shot designs, although I was able to get limited reuse out of some things for experimental purposes. Our stuff was all ablative.
For solid rockets, we nearly always used a fiber-reinforced rubber as the case insulation, to which the propellant was directly bonded (given the right primer) in the casting process. These liners were laid up on expandable mandrels, and pressure-cooked in place inside the motor case. They were usually an EPDM-type rubber, with initially asbestos fiber, and later kevlar fiber. It started charring about 300 F, and was fully charred through as it reached 600 F. We just used a thickness such that it didn't char through until the burn was over.
In the ramjets, the best choice was the far more erosion-resistant (by an order of magnitude) DC-93-104 loaded silicone rubber. That's an ITAR material, so I can't really tell you what's in it and how it works so well, but the Japanese offer a reasonable equivalent called Type 0 Shin Etsu. I've used both of them. This silicone system was necessary because the burn times were 10-100-or-more times longer. You make it thick enough to just char through by end of burn, or else you figure a way to retain the char, which also made a decent insulator. It charred at about 600 F or so. Silicones really do survive at much higher temperatures than hydrocarbon polymers.
Bear in mind that we had to control motor case temperatures when adjacent to flame in the 4000-6000 F range, at anywhere from 100 to 3000 psi pressures. If aeroheated on the outside, we would have to add an insulator or coating outside, too. It was a really tough heat protection problem to solve. But it was nearly always one-shot. Reusable, or steady state, is much harder still.
This silicone stuff you pressure-cast around a hard mandrel in the case, and then cook it to accelerate the catalyzed cure before extracting the mandrel. You need an impermeable separator sheet bonded to it before you can cast propellant on it, because the silicone in the liner is chemically incompatible with the hydrocarbon rubber in the propellant binder system. The correct primers were required on both sides of the separator sheet, which also had to have the right surface textures (because teflon sheet is otherwise a release agent).
We usually used martensitic stainless steels for our motor cases, like 4130, or maybe D6ac. It was more difficult and expensive to use alloy steels like 17-7PH, but they do have better strength to higher elevated temperatures. What we liked about 4130 was it was strong and sound and resistant to impact cracking even if soaked-out to -65 F cold, while still fairly strong at substantially-elevated temperatures in the 1000-1200 F range. Plus it was easier to work, to get high strength by cold-working, and to weld.
Some of the Sidewinder cases we made of aluminum, because the customer specified it. Most were steel, however.
We did make third stage eject motor cases out of 6-4 alpha-phase titanium for Poseidon, but these were literally carved out of big stock by machining away everything that didn't look like the case, because that stuff is not formable. Those were quite expensive.
We did make experimentally some other motor cases out of a beta-phase formable titanium, but these would age to useless weakness at room temperature in about 6 months. Nothing useful there.
Believe me, I understand about ceramic surface coatings. There was a mag oxide /mag silicate slag produced copiously by one of the gas generator-fed ramjet fuels I played with. That stuff would plate-out during the burn onto the charring surface of the silica phenolic liners in our heavyweight ramjet test hardware. Having that hard slag ceramic coat atop the slightly-charred phenolic made the liners last for dozens and dozens of firings, instead of just a single handful of tests. The trick is less about which ceramic, and more about how to actually install it to whatever you wish to protect. In our case, the slag was molten at near 2000 F, while the phenolic was fully charred at 600 F. It stuck to the char, not the virgin phenolic, which could only be char at the molten slag temperature. But, was it ever tough!
Not at the rocket plant, but while working at a countermeasures house, I had the opportunity to play with low density ceramics as a refractory, non-ablative motor insulator. My homemade stuff was similar to NASA's shuttle tiles, except that the successful version of my stuff was a ceramic composite, reinforced with ceramic fire curtain cloth. It had about the density of the heavier grades of industrial styrofoam, and a rather low thermal conductivity somewhere in the vicinity of 0.02 BTU/hr-ft-R. Being reinforced, it was structurally stronger than NASA's fragile tiles, although it was still rather fragile itself. I put several hours' burn at near 3500 F gas temperature on it, including dozens of excursions of the combustor/inlet into violent rich blow-out instability. A quarter-inch thickness of it held about a 12,000 F/inch thermal gradient for me.
That's my pedigree into heat protection and structural design, obtained in the school of hard knocks. I majored in aerothermo and propulsion, and aerodynamics, not really this stuff.
GW
]]>Apart from active cooling using the fuel, supersonic stealth aircraft have thermal barrier coatings on them to protect against radiant heat. They're also used in racing.
The following link shows what a blow torch does to a protected vs unprotected carbon fiber composite:
Ceramic Coating of Carbon Composite Heatshield -Zircotec ThermoHold Ceramic Coating Thermal Barriers
Incidentally, this is how F1 teams prevent hot engine exhaust headers from pyrolyzing their carbon fiber chassis. The ceramic thermal barrier coating is applied to both the composites and the exhaust headers to keep under-cowl temperatures manageable.
The point is, it's doable, its been done for some time now, and it obviously works well.
]]>A quote from the article you linked:
"However, as the skin temperature rises to 120 °C or above, thermo-oxidative degradation and its potential consequences can either prevent composites from being used at all or require them to be protected – usually through insulative layers – such that they operate at more benign temperatures. In many cases, the use of insulation is not permissible (such as on aerodynamic surfaces or parts exposed to high-speed airflow), but, recently [50, 51], some development on thin thermal-barrier coatings has been done to try to address this problem. The goal of thermal barrier coatings is to reduce the temperature of the substrate such that it operates at a more benign temperature where it is stable."
120 C as listed in the article equates to 248 F. That's right in the same ballpark as the failure temperature rule-of-thumb (290 F) that we used for military-grade asbestos-epoxy cement within rocket motors at the rocket plant. It failed by ceasing to be epoxy: it pyrolyzed into carbon char. The resulting char was relatively worthless at binding the asbestos fibers together. It was no longer structural once that happened.
The "bond strength at 300 C" (572 F) that you quoted is the shear strength of char still holding things together after the epoxy has long "died", and it is NOT a large number, said char resembling the charcoal in your BBQ grill. Most epoxy-type composite materials need to stay under about 250-300 F to maintain some structural strength, and are usually pyrolyzing into char above 300 F. There may be some that can go somewhat higher, but not 300 C = 572 F. Not and stay epoxy. There is a huge difference between one-shot use and usable-repeatedly.
The stuff I quoted, which you had a problem with, was for material soaked out to the full recovery temperature of the adjacent air, which is very nearly the total temperature of the adjacent air. Materials WILL soak out that hot, if aeroheated without any paths to cool by conduction, and without any significant emissivity for cooling by re-radiation to the environment. That's just heat transfer physics, you cannot argue with that.
In aircraft structures, the paths for conduction into interior items are usually few and far between. Skin panels are usually relatively isolated in that respect. So unless there is some kind of active cooling designed-in (and usually there is not), that kind of conduction is not a significant cooling path.
In aircraft structures, the surface coat texture and color drastically influence the thermal emissivity that controls re-radiative cooling. Dark or black surfaces, especially those not smooth and shiny in texture, usually (but not always) have rather high emissivities, usually at or above 80%. Shiny surfaces and white or light paint colors usually (but not always) have low emissivities, usually under 20%.
This is quite important to radiation physics, because heat flow rate per unit area is proportional to emissivity, and proportional to the difference of two temperature-to-the-fourth-power terms. One is the material equilibrium temperature ^ 4, the other is the effective environmental temperature ^ 4 (said temperature usually taken to be 300 K).
Skin panels with high emissivity can reradiate and stay cooler than the recovery temperature of the adjacent air, which is the driving temperature for the aeroheating to the panel. Skin panels that cannot effectively re-radiate because of low thermal emissivity will essentially soak out to the air recovery temperature. The equilibrium panel temperature reflects the balance between heat flow inputs and outputs, to and from the panel. Physics.
That re-radiative cooling mechanism is EXACTLY how the skins on the jets you named stayed cool enough not to fail, at the speeds you indicated. Camouflage coatings are usually high emissivity. And that is usually specified by contract. The equilibrium temperature can be quite a bit cooler than the recovery temperature, especially for emissivities of 80% and above.
The lack of that same cooling mechanism at low emissivity is why skin panels on the Concorde were 100 C (212 F) at Mach 2 in the stratosphere. The standard-day total temperature for Mach 2 flight in the (cold) stratosphere is 242 F = 117 C. Recovery would be a tad lower: just about 212 F = 100 C. Those were shiny aluminum skins with near-zero thermal emissivity. No surprises there.
Most of the radomes I was familiar with, those that were capable of sustained supersonic flight, were simply not fiberglass or epoxy composites. There are a number of specifically-developed radome materials for supersonic application. Some will go faster than others. But they are NOT epoxy composites, unless your speed is rather low supersonic or less. Both the aircraft and the missiles use the same technology.
Some of the material designations I found in AD-A007956 "Avionic Radome Materials", authored by R. H. Cary (for a NATO/AGARD document) are "alumina", "pyroceram", "silica", "cordierite" = "rayceram", "silicon nitride", "boron nitride", and "beryl oxide". These are nearly all ceramics of one kind or another, listed in multiple forms in the document. Some are reinforced with ceramic fiber or fabric. The missiles usually use the silicon nitride or pyroceram materials. Some of those (like Phoenix) peaked briefly at Mach 5, but were significantly supersonic all the way to target.
For lateral skin panels, the key to success is re-radiative cooling. That's how you get away with aluminum panels past Mach 2.2 in the stratosphere. It's how you get away with epoxy composites past Mach 1.4-ish in the stratosphere. It's how you get away with titanium past Mach 4 in the stratosphere. Etc.
The aeroheat rates are order-of-magnitude worse at stagnation points and lines. That's why the nosetip and leading edge problems are harder to solve. But you use the same basic mechanisms: primarily re-radiative cooling, plus significant conduction from the stagnation line into more aft regions where the aeroheating is lower, using the attach structures for the leading edge parts as your conduction paths. You can still avoid active cooling by doing that, for any practical supersonic jet aircraft design.
You can even add some heat-sinking to go hypersonic, as long as that is a short transient, not steady-state. That's how the missiles get away with going that fast.
GW
]]>If any part of this persistent claim of yours was actually true, then there's no way that fighter jets like the F-35 and F-22 wouldn't suffer from catastrophic structural failures, nor would any supersonic radar-equipped fighter or bomber with a fiberglass nose radome survive, for that matter. The stripped-down F-22 test articles could and did super cruise above Mach 2 for extended periods of time, meaning longer than 30 minutes. None of them suffered from catastrophic structural failures, and nearly the entirety of their skin and leading edges were radar-absorbing structural composites using epoxy and glass or carbon fiber. The skin temperature of the Concorde was 100C at Mach 2, 120C at Mach 2.2, and 150C at Mach 2.4. That falls very well within the temperature limitations of high temperature epoxies, some of which maintain enough bond strength to operate up to 300C. The epoxies that are available today are not the same as whatever was available in the 1970s.
Boom has already stated that the fastest speed they can achieve over land is Mach 1.4, due to not being able to pass noise restrictions regarding sonic booms. The speed restriction has nothing whatsoever to do with the structural integrity of the airframe or epoxies used in CFRP. We have epoxied composites that retain more tensile strength at elevated temperatures, up to consumer oven temperatures, than Aluminum does. The epoxies that can do that are really expensive, but readily available as commercial products, and are made by the likes of 3M (ScotchBond), Loctite (Hysol), Cotronics (DuraBond), and others, specifically for use in both civil and military aircraft, as well as structural composite parts used near engines or high-heat / high-pressure industrial applications.
]]>The standard today for comparison is long range transport with high-subsonic gas turbine-powered airplanes. That's for both people and cargo. It takes 12-14 hours to cross the northern Pacific this way, a little longer to reach Australia from North America, which is essentially the antipodes.
A supersonic transport can do this with a lower transit time, but a higher cost. That is inevitable, because drag forces are inherently higher in supersonic flight, as compared to subsonic flight, almost no matter what design innovations you might have. The wind pressures are just greatly higher. You have to burn fuel to overcome drag in steady-state flight, no matter what.
Similarly, the engines suitable for supersonic flight inherently have lower bypass ratios than engines that only fly subsonically. There are no innovations to get around ths, only those which affect the precise numbers. But supersonic bypass ratios are, always have been, and likely always will be, lower than subsonic bypass ratios. The thrust specific fuel consumptions are inherently higher at low bypass ratio. There is no way around that, so there will always be a higher fuel burn in supersonic flight.
The faster you intend to fly supersonically, the higher the drag (and resulting fuel burn), more-or-less quadratically increasing with speed. The shorter your flight time will be, only linearly changing with speed. Those physics plus the bypass effect on engine efficiency, are why all forms of supersonic flight with (reusable) airplanes has always been far more expensive than subsonic flight. It will likely remain that way, no matter what innovations get introduced. Those only affect how strong a quadratic model you have to use.
There is a second complicating issue that increases very greatly the costs of supersonic flight, and that is friction aeroheating, also greatly (and nonlinearly) increasing with increasing supersonic speed. Organic composites are ruled out for steady flight by equilibrium skin temperatures somewhere near Mach 1.5-1.7. Aluminum is ruled out by around Mach 2.2-2.4. Titanium is ruled out above about Mach 3.8 to 4. That leaves you stainless and alloy steels, and dense ceramics. Both are heavy, and that increases fuel burn due to weight.
My understanding of the organic composite Boom Supersonic prototype (and related ultimate design) are that they never fly faster than about Mach 1.3-1.4. The high strength/weight organic composites simply do not allow faster flight in the face of the aeroheat risks. The lower supersonic speed and composite construction to save weight, plus the better engine technology available today, is EXACTLY how they get the cost down to something perhaps affordable. You can cut about 1/3 off your flying time, for only a somewhat-higher ticket price, compared to a subsonic transport.
Transoceanic rocket transport avoids much of the aeroheating problem by going exoatmospheric, and doing the reentry as a transient, fundamentally-heat-sinking problem, at flight's end. Just as we have done with reentry for about 66 years now. That kind of transient heat protection problem is vastly different from the steady atmospheric flight heating problem. It is far easier, less expensive, and lighter in weight to solve.
If the rocket really is reusable, then the main cost factor is the truly enormous propellant burn required for rocket travel at all. The ticket price will always be higher than subsonic air travel ticket prices, that is inherent. It will likely be higher than ticket prices for modest supersonic air travel (that being under Mach 2 at most, and more likely under Mach 1.5 to use composites). It just gets down to the question of how much more are you willing to pay for cutting transoceanic travel time drastically (from a dozen hours to under half an hour)?
We've tried supersonic travel once before with Concorde at Mach 2. That was with much poorer engine technology, using an aluminum and titanium airframe. (The Boeing Mach 3 design that was never built, was a heavier all-titanium airframe.) Concorde cut times roughly in half. But it was not really a commercial success, because it was only a niche market due to its cost. Slowing down to Mach 1.3-1.5, with far better engines, and a composite airframe, might be a lot more attractive, even though you only shave off about a third of flying time. We'll soon see, I think, thanks to Boom Supersonic and maybe a few others.
Whether drastically shaving flying time at a much higher price yet, with rocket travel, will prove commercially attractive is something we might actually learn if Musk is successful. I think the answer is still way far out in the future on that. I suspect it's a niche market only. But maybe we'll see.
GW
PS - most people do not know the real genesis of the Saturn 5 rocket design, done by Von Braun for US Army at Huntsville, AL. That was only a paper design until NASA "bought" it (and him) for its Apollo effort. It was originally a rocket travel troop transport, one-way to Russia with about 100 men per rocket. The "third stage" was the troop transport lander.
NASA had him replace that with a real propulsion third stage to be its Saturn 5 moon rocket. This stuff actually was published as "gee look how good we are" ads in the predecessor to "Aviation Leak" magazine in the late 1950's and early 1960's.
PPS- most people also do not know that the original Apollo design landed the entire Apollo CM/service module on the moon. It took two Saturn 5 launches per moon mission to do this, refuelling in LEO from one to the other, in order to send this cluster direct to the moon.
The idea of lunar orbit rendezvous using a separate lander came from outside NASA, meeting very heavy "not invented here" resistance from NASA. Only the pressure of beating both the Russians, and JFK's timeline, forced them to adopt it. Which got them down to one Saturn 5 per moon trip. Which was something they could afford, and that both Congress and the public would buy.
]]>Sorry, but I did not read the article. I'm using what I know about what things cost, what people are willing to pay, therefore what's practical and what's not. I don't need to read a news articles to tell me what I already know. I did some basic math using freely available information on the cost of Jet-A / LOX / LCH4 and fuel burn rates. That told me pretty much everything I needed to know about the practicality of making New York to LA flights using the upper stage of a super heavy lift launch vehicle.
Nobody has done this yet because it's not feasible to do it with current technology. We have greatly improved aerodynamics (thanks to CFD optimization software), decreased structural weight of airframe materials (CFRP), and turbofan engine efficiency (high pressure ratio, ultra-lean burn at blow torch temperatures) to the point that a purpose-built turbofan (like the F-136 engine with ADVENT technology that wasn't quite ready for prime time when the F-35 program was in initial development) can burn about twice as much fuel to go about twice as fast as a subsonic airliner. Whenever you start trying to fly appreciably faster, aerodynamic heating rapidly increases airframe weight (CFRP and Aluminum won't cut it, so you'll need stainless steel), thus the lift-to-drag (bigger wing generating more induced drag to carry the increased weight of steel) and therefore thrust requirement (more thrust to overcome more drag), thus engine efficiency (at low to moderate subsonic speeds pistons or turboprops are most efficient, at high subsonic speeds turbofans become more efficient, the fan in a turbofan starts becoming an aerodynamic equivalent of a brick wall around Mach 3, so turbojets tend to get used for efficiency at Mach 2.5+ speeds, then ramjets start looking more attractive around March 4 or so, and by Mach 10 or so, even scramjets become impractical, so you switch to using rocket engines), and then fuel burn really starts to "take off" (the faster you go, the more fuel you need to burn).
There's never been a shortage of people willing to blow mad money on impractical ideas. I still think supersonic jets are playthings for rich people, but the relative fuel burn associated with a modern Mach 2 airliner design is no worse than a small business jet, on account of how business jets are made (generally from heavier but cheaper Aluminum alloys, rather than carbon fiber tape layups that airliners have mostly switched over to using these days) and the poor fuel economy of very small turbofans (biz jets) vs very large turbofans (intercontinental airliners). There might be an engineering reason why modern turboprops typically start around 550hp and go up from there- just sayin.
This kind of stuff falls squarely in the same realm as all of us cruising around in "flying cars" like the Jetsons without any pilot training or Star Trek level technology. It's fun to fantasize about, but when the rubber meets the road, it's absurdly impractical using current technology, with a total lack of pilot training. The entire reason air travel is so much less dangerous than highway driving is specifically because far fewer vehicles are present in far more physical space and most of the pilots at the controls have better training, recurring training, lots of experience flying, and better awareness of what's going on around them. Computers are a valuable aid in that cause, but not the primary reason why the overwhelming majority of airliners make it to their destination in one piece. If NHTSA mandated the use of multi-million dollar driving simulators to train new drivers, recurrent qualification to maintain a vehicle operator's permit, drug testing, yearly medical screenings, and licenses that can be revoked at any time without the need to obtain court orders, does anyone here believe we'd have a fraction of the motor vehicle accidents? Any significant mistake you make at the controls of an aircraft, and you will find yourself talking to a FAA rep who will determine on the spot whether or not you keep your license, because they exist to maintain the confidence of the flying public in their ability to evaluate and regulate every aspect of flying, with a keen eye towards absolutely minimizing the number of aviation related accidents.
Anyway, it's a cool concept that's also totally impractical, even if you can technically make it work, which has yet to be proven. After I see several hundred successful orbital flights, which include 24 hour or less turnaround times, then we can reevaluate the practicality of hypersonic point-to-point travel.
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