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This is true but incomplete:
"Whatever that velocity is, it needs to be subtracted from the Hohmann transfer velocity.
The net result is the dv that the vessel must apply to dock with Phobos."
The result is NOT what you need to dock with Phobos, but the velocity "far" from Phobos before 3-body effects of Phobos's gravity accelerate you further toward Phobos. With Phobos, that's not a large effect, but it is there! The accelerated speed "near" Phobos is your actual docking requirement with Phobos.
This is very well-covered complete with a simple estimate of the 3-body effect, in the "orbits+" course materials. The estimating equation is based on conservation of mechanical energy: Vfar^2 + Vesc^2 = Vnear^2. In the case of Phobos, its escape speed Vesc is quite small, which in turn is why the effect is quite small.
Going to orbit about Mars, the same sort of thing applies, except that the 3-body effect is due to Mars, not Phobos; so the Mars escape speed at orbit altitude is quite substantial. It is a large effect.
GW
I saw films taken of Russian tanks being air-dropped in battle exercises. These were shown on Walter Cronkite's old "20th Century" program in the early 1950's. The rockets are located at the junction where the parachute shroud lines connect to the strap or straps that hang onto the payload. They fire in a conically-downward spread around the periphery of the payload. They are ignited upon proximity to the ground (somewhere around 10-15 feet or thereabouts).
Blue Origin is doing something very similar with its New Shepard suborbital spaceflight vehicles. The booster lands very similar to a Falcon 1st stage core. The capsule comes down by chutes, but fires rockets to slow down at the last second. That's what the dust cloud is, around the periphery of the capsule, when it touches down. Surface proximity fires them at about 10 feet or so.
This is a very old technique for making parachute landings on land without bad touchdown shock. It has worked for about 7 decades now.
A variant of it has been what the smallish JPL probes do for landing on Mars. It's just that the rockets have to be bigger, and you fire them earlier, because they have to slow you to around 2-5 mph from about Mach 0.8, not 20-25 mph. It only works up to about a ton landed on Mars, because otherwise you come out of the hypersonics too low for the chute to work. What good is a chute that does not have enough time to actually slow you down subsonic? Or maybe not even enough time to deploy and inflate?
GW
Technically, the path is a very eccentric ellipse in space, yes, but from an observer on the ground it looks like a straight line up, and right back down. At apogee, there is no relative velocity in any direction, as viewed from the launch site on the ground. Spin launch has everything to do with reaching an altitude outside the atmosphere. It has nothing to do with having any speed in any direction, when you get there. That is why the plan for spin launch is a small payload and a small two-stage solid rocket, riding inside the carrier device that you spin up and fling upward. Literally all of the speed to go anywhere in space comes from the rocket that you carry up there with your payload.
GW
That kind of stuff is quite expensive. To my knowledge, whole airplanes are not made with it.
GW
Wing loading works out as a function of size and intended use. Light aircraft are usually in the vicinity of 10 psf. Model aircraft are nearer 1 psf, for the same speed ranges. The older propeller airliners near 20-30-something psf, along with the propeller bombers of that time, all being under 300 mph. Higher-performance supersonic jet fighters today are nearly all in the 100 psf class, give or take.
The delta wing is not easily characterized as "inefficient". It depends upon what you want to do with it. The delta wing has a very severe center-of-pressure travel going from subsonic to supersonic and back. That introduces very severe center-of-gravity issues into any supersonic delta wing fighter or bomber designs, something known since the late 1940's. See XF-92, our delta prototype back then. See also the losses of the B-58 until the in-flight center-of-gravity control was automated to respond to an engine-out deceleration-to-subsonic faster than a human pilot could respond.
Countering that is the advantage of being able to reach angles of attack in the 40+ degree class at landing speeds, without adding a plethora of leading edge devices to keep from stalling. That has also been known since the late 1940's. Swept wings cannot do that. Nor can straight wings. And we know it works: F-102, F-106, B-58.
And adding a sort of strake ahead of the main wing, often termed a "double delta", does help control the center-of-pressure travel better. Such was on the shuttle, and the SR-71/YF-12A, and is on the FA-18, although that one is not strictly a delta. Such is a good add-on for most stub-type wings with swept edges.
The real problem of using a delta wing to achieve a low landing speed without high lift devices is that high angle-of-attack: the pilot literally cannot see the runway directly in front of him. He really needs a window through the forward part of the cockpit floor. Something almost impossible to do, in turn leading to other bizarre design implementations. See Concorde vs Tu-144.
While delta wings can be flown "tail-less", they work with better safety if you also add a horizontal tail somewhere: XB-70, A-4 Skyhawk, Tu-144.
GW
PS: the swept wing also has serious low-speed instability troubles, requiring leading edge flaps, and/or fences, and/or a step in the planform about mid-span, to help control those instabilities. You pay a design price to get transonic drag reduction with sweep. This was extensively explored with the research craft at Muroc Dry Lake in the late 1940's and early 1950's, particularly with the D-558-2 Skyrocket.
Parachute landing a capsule on land (with retro thrust at the last second) at something in the 20+ mph vertical velocity class is too sudden a stop for crews (or delicate equipment) to easily survive. Apollo was capable of landing on land in an emergency, with crew couch supports that would crush. But the risk of injury was rather significant.
The last-second retro thrust slows the 20+ mph descent rate to around 5+ mph at impact, which makes the landing easily survivable. However, it is live ordnance. So you are trading one danger for another. The track record suggests the risk of the hard landing outweighs the risks posed by the live ordnance. The Russians first came up with this concept for air-dropping main battle tanks back around 1950 or so. It went immediately into their early spacecraft designs, enabling landings within their borders.
GW
From the “Daily Launch” for Monday 6-17-2024:
SPACEFLIGHT NOW
NASA, Boeing set new undocking, landing date for Starliner spacecraft
NASA and Boeing teams pushed back the target undocking and landing date for the Starliner spacecraft from the International Space Station by four days. They shifted from June 18 to now no earlier than June 22. More testing will be done on both the thrusters on the spacecraft’s aft section. The Starliner crew will also repeat some previously conducted test activities, like the ‘safe haven’ protocol.
With additional data if you follow the link to the article: the 5 problem thrusters on the capsule were physically not performing correctly, and it was the software that shut them down for misbehaving. Testing got 4 working well enough, and they were re-enabled in the software, which is what allowed docking with the ISS at all! As for the helium leaks, there were 5 in all, with the 3rd found being the largest, and by far. The oxidizer valve problem is an isolation valve that refused to close properly, meaning it failed to close all the way.
My take on it:
As old a technology as pressure-fed NTO-hydrazine (of any kind) thrusters are (some 6+ decades), there is simply no excuse for these kinds of thruster misbehavior and thruster plumbing problems in a spacecraft being considered for man-rating! Or even for routine unmanned use! Boeing simply cut too many corners on this design, and it has already bit them quite hard, with the repeat unmanned ISS flight that they had to pay for out-of-pocket.
This is what you get when you are led by “professional” managers at the corporate level who do NOT want to hear anything from their engineers except how much money was saved. And they are not the only outfit afflicted with that flaw. Here’s another one, from the very same newsletter issue:
BREAKING DEFENSE
Space Force boots RTX from MEO missile warning/tracking program
The Space Force has terminated its contract with RTX (formerly Raytheon) for development of the service’s new missile warning/tracking constellation in medium Earth orbit (MEO) due to cost and schedule overruns, as well as technical issues, a spokesperson for Space Systems Command told Breaking Defense.
And this one, still the same issue:
THE NEW YORK TIMES
F.A.A. Investigating How Questionable Titanium Got Into Boeing and Airbus Jets
Some recently manufactured Boeing and Airbus jets have components made from titanium that was sold using fake documentation verifying the material’s authenticity, according to a supplier for the plane makers, raising concerns about the structural integrity of those airliners.
The lesson here: this kind of crap is what you get when you let aerospace defense companies be run by people who only care about the money, to the exclusion of all safety, quality, and reliability. How is this any different from the cheap, crummy crap made by slave labor in China, that floods our commercial markets?
GW
Another thing to consider in regard to tank inert mass: is it, or is it not, also the airframe of the vehicle? If it is, what additional loads besides endways gees and internal pressurization, must it carry? If it is not, its mass is in addition to the airframe mass that surrounds it.
I'm only guessing that wall thickness depends upon internal pressure level and material strength as a rough first approximation, and the volume of material to be weighed is the tank surface area times that thickness times some factor a tad above 1, that accounts for local thickness increases, such as at the head-cylinder joints. That takes care of low-density hydrogen, vs the other fuels.
That simple estimate ignores other very real-world mass additions for things like internal surge baffles, and the plumbing that takes propellant from another tank through this one, to the engines. Sometimes that plumbing must be quite heavily insulated.
As for resisting other airframe loads such as bending, and external pressure on one side but not the other (acting to cause collapse laterally), that requires either external or internal stringers, and internal frames, to act similar to a monocoque fuselage.
And I still do not believe in organic matrix polymers that aren't structural junk at 300+F. Composite tanks will inherently require more heat protection that stainless steel, in ANY entry scenario, even the low velocity entries of 1st stages. 300 F is the total temperature of the air at only Mach 2.2 in the stratosphere. That's only 0.64 km/s velocity through the cold stratospheric air.
Aluminum goes slightly hotter than that at about Mach 2.5-2.6.
GW
This kind of thing works for EDL if (and only if) the density of that tenuous atmosphere is known reliably. Period. End of issue.
With one fly-by measurement at only one point in an elongated eccentric orbit, I rather doubt the atmospheric density measurement can be considered reliable.
Let's put it this way: I certainly would bet only $0 on it!
GW
There is no ellipse circularization. The trajectory is straight up to an apogee at no velocity in any direction. The rocket has to take the dead-head payload to 7.91 km/s from 0 km/s, starting from that apogee.
GW
What about the need to rendezvous with whatever you are delivering the payload to?
What about the deorbit burn, necessary to properly dispose of the stage, even if not reusable?
GW
All objects of any shape shed shock waves at any speed faster than sound. At very fast supersonic and especially hypersonic speeds, that shock wave wraps about the object, but not very far from its surface. For pointy objects, the shock wave is attached to the nosetip, while for blunt objects, it is detached as a bow wave just out front. However the lateral sides of the shock wave "wrap" about the object are roughly the same shape, regardless. It cannot be significantly moved by changing the shape of the object. It simply "is".
GW
I checked, Bob is right, SABRE is not a liquid air cycle engine. I honestly thought it was. It does have some sort of translating nozzle design, connecting to two different combustion chambers, one for the airbreather at lower expansion ratio, and one for the rocket at higher expansion ratio. The precooling is how they can do axial-flow compression to very high pressures, and why they can do this past about Mach 3-ish, where the blading in conventional axial-flow compressors on conventional gas turbine engines gets too hot and self destructs under the loads.
That being said, I don't think there is a practical solution for the Skylon airframe (as it is usually depicted) to survive re-entry, not with those engines on its wingtips. I think better thought went into the engine than did the vehicle design. The Wikipedia article about Skylon has the BS about reduced ballistic coefficient reducing entry heating. Use my entry spreadsheet and run the trades yourselves. Yes, lower ballistic coefficient raises deceleration altitude. But only by a little bit here at Earth. So the peak heating is only lowered a little bit. The effect is dramatic at Mars. Not here.
Not with those pointy spikes and nosetip, and the necessarily-low radius wing leading edges! The engine inlet spikes must be sharp, or else they cannot properly provide the external shock wave compression function that they must serve. Blunt them, and the shock detaches as a bow wave out front, which greatly increases the drag of the engine nacelle. But as I wrote in the earlier post, sharp things cannot stay sharp after being through entry. Not and survive the heating.
It also claims a magic skin material which will survive the (unrealistically-low) heating: some sort of ceramic matric carbon fiber composite. To the best of my knowledge, there is as yet no such material, other than as a lab curiosity. Even if it did exist, the shock impingement heating raises the equilibrium leading edge material temperature very close to soaking-out at the effective plasma sheath driving temperature: something in the 7000-8000 K class. Not even a ceramic matric carbon fiber composite (should one actually exist) will be able to withstand that!
Multiple layers of foil between the skin and the load-bearing internal structures and aluminum tanks are supposed to be "insulation" preventing heat from conducting inward from the hot skin. There's better insulations than that, but the whole concept ignores conduction between the skin and the load bearing structures, passing through whatever things physically hold the skin on the airframe at all.
For those who don’t believe in shock impingement heating, the damage it did to the X-15 is documented in NASA Technical Memorandum TM-X-1669, “Flight Experience With Shock Impingement and Interference Heating on the X-15A-2 Research Airplane”, written by Joe D. Watts, dated October 1968. It is available publicly on the internet as a pdf file.
GW
Tom:
Thrusters are never turbo-pump-fed rocket engines. They are always very simple pressure-fed devices. They use helium as the pressurant gas to drive propellant flow, because it will not react with the very reactive NTO or any of the hydrazines.
Any of the hydrazines used with NTO are hypergolic ignition, greatly simplifying the design of thruster systems. They are also hypergolic with nitric acid, usually in the form of inhibited red fuming nitric acid (IRFNA), which has a little bit of NTO and a little bit of water in it. However, it is NTO that gives the higher c* for higher Isp, all else equal.
IRFNA and aniline are another hypergolic combination, but again, NTO-hydrazine gives better c*. IRFNA is often said to be hypergolic with other hydrocarbons like kerosene, but that really isn't reliable enough to be "true" for real engineered systems. Maybe wide-cut, but certainly not plain kerosene. It needs to be a light molecular weight hydrocarbon (like aniline), for hypergolic ignition to be fully reliable.
Hydrogen peroxide is hypergolic with kerosene, but only if really high-test: near 100%, which is very very unstable and short-lived. That's what Andy Beal was using. He was the first liquid rocket operation at McGregor. They did have one peroxide decomposition explosion out there, before Beal decided his rocket initiative was not going to make it as a business. His tall firing test tower was SpaceX's first test stand when they moved in after Beal vacated.
Rob:
Thanks, I wasn't sure which hydrazine Starliner was using. Looks like just about everybody is using MMH today. But, it hasn't been that long since Aerozine-50 was still popular. That one is a 50-50 blend of plain hydrazine and UDMH. I'm not sure, but I think some still use it. I think the Russians use something similar to it.
GW
I just today saw a news story that Starliner's visit to ISS has been extended to June 22. Ostensibly, this is for "more testing". However, the spacecraft is docked with its systems turned off and its helium supply valves closed. Turned on, the helium will leak away to zero in only a few days. That much I have seen stated publicly. This is the helium whose pressure drives the propellants to the thrusters. There are no pumps, the thrusters on both modules are utterly-simple pressure-fed hypergolic-ignition designs. NTO and one or another of the hydrazines.
I suspect that they're really still trying to figure out if the service module propulsion will have enough helium left to power the deorbit burn, and if the attitude thrusters on the capsule will be able to hold attitude through entry. If my suspicions are true, you can be sure no one is going to say anything publicly that might verify them. If it turns out they cannot do the job, expect a rescue using crewed Dragon.
GW
edited 6-15-2024 to insert the word "sure" where it was missing in the 2nd sentence of the 2nd paragraph.
For GW:
If the engine pods were dropped (ejected) from the wings before re-entry, would the problem you've described go away?
Ans.: yes, but it is hard to imagine how to recover them, or how to get them to survive entry tumbling. They are the most valuable portions of the vehicle, even more so than standard rocket or jet engines. That also precludes post-entry propulsion for Skylon, reducing it to a dead-stick glide landing without much divert capability at all.
Also for GW ...
Is the problem you've identified addressable by changing the shape of the engine pod prior to re-entry? I'm thinking of putting a cap on the tips of the engines. Could a cap be designed to cause the shock wave energy to move away from the wings and out away from the aircraft?
Ans.: No. Even a blunt object sheds a detached bow shock, which only a bit more remote from the object (about a cross section dimension away) is just about the same angle and shape as the shock shed by a sharp object.
Finally for GW ...
If using propulsion to reduce velocity prior to entering the atmosphere is the solution, then all we are talking about is the cost of propellant.
Can the Skylon carry enough rocket propellant to drop it's orbital velocity low enough to survive re-entry?
Ans.: No, not in anything resembling its current form. Wig = 275 tons, Wp = 220 tons, and Wpay = 12 tons, per the article. The "allowance" for inert + payload is ignition minus propellant = 55 tons. Subtracting 12 tons of payload means their design concept's inert mass is 33 tons (edit correction 43 tons), an inert fraction some 12% (edit correction 15.6%) of ignition. The thing as it is is 80% propellant. Bear in mind the dV to deorbit from LEO is only about 0.1 km/s. The dV to "survive" entry without serious aerobrake heating is 6-7 km/s.
They think (per the article) that they can use carbon composite for exposed airframe structure and meet that inert percentage with little in the way of heat shield, based on a BS argument about "lower ballistic coefficient" (which in turn tells me no one on that team knew anything substantive about entry heating). But that carbon composite material is limited to failure of the epoxy matrix at material soak-out of only 300 F. It requires protection, even on the lee-side, just as SpaceX found out with their "Starship", which is one of two good reasons they went to stainless steel.
I see little evidence they had anybody on the team that came up with this concept who actually understands reentry. The only way I see to "save" the design is to fold the wings into the wake zone, which cannot be done with a round body. The cross section will have to be almost triangular to do that, with a rather flat belly in that view.
That required cross-section shape will play merry hell with pressurizable round tank shapes that must fit within an airframe instead of actually being the airframe. Between that and adding both a heat shield PLUS INSULATION BETWEEN IT AND THE COMPOSITE AIRFRAME, I would expect a vehicle inert mass fraction nearer 25-35% than 12% (edit correction 15.6%), leaving only 65-75% propellant even at zero payload. There went your dV capability!
One other thing: there are 3 sharp stagnation points on this design: the nose tip, and the two spike cones for the engine inlets. The spikes must be sharp to function properly as supersonic inlets. The nose tip does not have to be sharp, but rounding it does increase drag coefficient. The problem with "sharp" in entry is high heating rate. The old first-order heating correlation equation that still works today says stagnation Q/A = constant * (atm density/nose radius)^0.5 * velocity ^ 3. I use that, and so did Justus and Braun in their seminal paper about entry, descent, and landing on multiple planets. It is from the 1953-vintage warhead entry work of H. Julian Allen. It worked pretty good then, still works pretty good today.
Note the dependence of Q/A upon 1/square-root-of-nose-radius. When that radius goes to zero, that factor in the heating correlation goes infinite. For entry from LEO, when the nose radius is about the same as the cross section dimension of the object, peak Q/A at about 7-8 km/s and 50 km altitude is in the hundreds of Watts/sq.cm (cm not m!!! so these heat flux rates are 4 orders of magnitude larger than most Earthly heat transfer applications). At 10% of that dimension, it's in the low thousands. At 1%, it's in the high thousands. And THAT is why nothing sharply pointed is still sharply pointed after an Earth reentry! (And that assumes no shock impingement phenomena multiplying those heating rates further.)
Do you see a problem with wing leading edges when that "nose radius" ratio is nearer 1% than 10%? Guess why space shuttle leading edges were carbon-carbon slow ablators and not the ceramic tiles. The two-piece Tufroc tiles on the X-37B can survive that kind of stagnation heating, but only just barely (!!!), and even then only on a rather blunt nose tip.
GW
edit corrections done at 3:26 PM CDT 6-14-2024
The ionization thing happens any time you try to combust fuel with air hot enough to be ionized. Down in the stratosphere, that starts happening about Mach 7 or 8 with inlets that try to slow the air subsonic. Scramjets have a different post-capture inlet shape that leaves the air supersonic, just slower supersonic than the oncoming speed. That delays the ionization problem to the vicinity of Mach 10. The static (thermodynamic) temperature of a subsonic stream is very near its total (stagnation) temperature. The static temperature of a supersonic stream is very much lower than the total temperature. In both cases, the temperature ratio total/static = 1 + constant*Mach number-squared, where the constant is computed from the gas specific heat ratio "g" as const = 0.5*(g - 1). This math fails once ionization becomes significant, because the equation of state PV = nRT is no longer valid. The chemical identity of the gas is no longer what it was because of the ionization. Chemistry IS the electron shell, and that is different after ionization begins.
Skylon is intended to leave the sensible atmosphere at speeds not much above Mach 5, where the magnified heating from the impinging spike shocks upon the leading edges, is still tolerable. The faster you go, the more intolerable this becomes, exponentially, because the basic heating is so much larger at higher speeds. You cannot take advantage of "leaving the sensible atmosphere" during reentry descent, BY DEFINITION. You WILL see about orbital speeds about halfway down in altitude from the entry interface, which for Earth from LEO is about Mach 12-13 at about 50-60-some km. The heating there is truly intense, and shock impingement magnifies that problem by a factor on the order of 10.
My point: there is a very good reason why no entry-capable spacecraft has ever flown with any sort of parallel-mounted nacelles or any other structures that could shed shock waves toward the main body. It is the shock impingement overheating risk during reentry. This puts heating into the affected material at a rate that is an order of magnitude (or more) faster than it can be removed by any conceivable means. That affected material then almost-instantaneously soaks out in the local affected zone to the effective driving temperature of the plasma sheath, which at 8 km/s is about 8000 K, at 7 km/s is about 7000 K, etc.
You tell me what materials might still be solid at such temperatures, not to mention structurally strong enough to resist wind loads. See why I talk about the impinging shocks cutting the wings off in mere seconds?
Re-entry vehicles that are lifting bodies or have wings enter the atmosphere at a high angle of attack, but not dead broadside! The space shuttle came back at a very precise 20-40 degrees AOA (40 at peak heating), higher or lower would cause hot plasma impact directly on the windscreen, causing immediate loss of the vehicle and crew. Even "Starship" comes back at 60 degrees, not dead broadside. I do not know the entry AOA for Skylon, but it would be in that same class.
Higher AOA will rip the wings or flaps off. Remember, peak entry gees are usually in the 3-7 gee class (which is why entry angles are always shallow, steep entries are hundreds of gees). Like in airplanes, you can only pull so many gees without ripping the wings off. Making it strong enough to resist more gees usually makes it too heavy to fly. 6+ gees is a lot of gees for any wings to resist. You don't do that at spacecraft-type vehicle inert mass fractions. Skylon CANNOT come back dead broadside, where the shocks don't impinge on the wings. The wind loads would rip the wings right off.
The main known way around these heating and structural integrity conundrums during reentry would be to slow way down before hitting the atmosphere, from about 8+ km/s to something nearer only 1 or 2 km/s (Mach 3 to 6, instead of 25). But that's a big engine dV requirement, which defeats the whole purpose of wanting to aerobrake instead of burning propellant.
There is a second: fold the wings into the wake, and come back dead broadside. But peak gees will be much higher that way. It might still be too much, if there are masses hanging on those wing tips.
GW
My understanding of Skylon and its SABRE (liquid air cycle) engines is airbreathing flight to only about Mach 5 at only about 100,000 feet, during which atmospheric oxygen gets stored in oxygen tanks that are empty at takeoff. At the M5/100k point, those engines shift to LOX-LH2 rocket propulsion, with which the vehicle pulls up steeply to get onto a non-lifting thrusted gravity turn trajectory. It can pull up, and it can accelerate upward along the steep path, PRECISELY because the propulsion is rocket. Rocket takes it to orbit. Single stage is possible because it tanks the oxygen during the early ascent, not before takeoff.
The SABRE engines are liquid air cycle engines. The air is specifically cooled to liquify it, and separate the oxygen from the nitrogen by way of the different boiling points. (This only works with super-cold liquid hydrogen as fuel.) In rocket mode, you dump the nitrogen, and burn the oxygen in the engines with hydrogen fuel from the tanks. In airbreather mode, you dilute the oxygen with at least some of the nitrogen. The combustor and nozzle are more akin to a liquid rocket engine than any gas turbine engine ever built.
This was an old concept dating to the 1950's, but has never been made to work until now. SABRE's makers seemingly have solved the problem of getting really fast heat transfer rates, and they have seemingly solved the problem of how to deal with the moisture in the air which condenses first and freezes all over everything. But, this thing has yet to fly, which is the real proof of the solutions.
I have no problems with any of that, but I do think they have not thought through the whole airframe layout problem, or else its designers had zero understanding of the realities of entry heating. They do NOT have a survivable heat protection scheme, because of shock impingement heating that WILL happen during entry! I have talked about this issue with that design before, here on these forums. It is a FATAL design flaw!
Those tip mounted engine pods have inlet compression spikes whose conical shock waves WILL impinge upon the wing leading edges (there is NO way around that with the parallel-nacelle shape). Those impinging shocks will cut the wings off in a matter of seconds during entry, no matter how thick a piece of carbon ablator they put on those leading edges. This was shown by the well-documented damage to the X-15A-2 when it reached (only !!!) Mach 6.7 at (only !!!)100 kft back in 1968, with an experimental nacelle mounted in parallel to the fuselage, on the ventral fin stub under its tail. Skylon will hit entry heating at Mach 25. Crudely, the expected intensity of stagnation heating varies as velocity-cubed during entry. That will be ~50+ times the heating rate the X-15A-2 dealt with.
Shock impingement heating does not increase gas (or plasma) temperatures, but it does increase the heat flow rate per unit area, right in the zone at the impingement, by a factor crudely near 10. Because of the effect of 4th-power re-radiation rates, the equilibrium temperature of the affected surface will be above all known material melting or ablative destruction points. We've already seen it!
GW
X-30, aka NASP, was to use scramjet to extreme altitudes and very high speeds. Scramjet was absolutely unready to apply in the 1980's. Plus, see what I said about ANY airbreather at extreme altitudes in post #2 above, my point 1 in that posting. To that I would add that the max potential Isp of even a hydrogen-fueled scramjet is pretty much equal to the Isp of a LOX-LH2 rocket ay only about Mach 10.
That very low airbreather Isp at extreme speeds is not just low inlet pressure recovery (which is also inherent), it is also caused by the inability to actually burn fuel with ionized plasma from the inlet, instead of real oxygen-bearing air. Oxidation has to do with the outer electron shell. Once ionized, that shell is damaged or absent. The chemistry "quits". Ionization/recombination energies cannot be converted to thrust by a nozzle, only enthalpy can. Once ionized, nearly all the KE of the oncoming stream goes into ionization/recombination, not enthalpy. That has been well-established for decades, now.
In short: the X-30 as-conceived could never have worked. That was a giant corporate-welfare program.
There has been talk for at least 4 decades now about this-or-that replacement for the SR-71, often now termed "SR-72", long ago often termed "Aurora". Yes, there have been experimental planes aimed at that goal. No, none of them yielded usable results, until possibly recently. This thing Lockheed has been hyping of late they term "SR-72" was to have a combined-cycle gas turbine/ramjet/scramjet engine.
They might have succeeded after a couple of decades trying, if they just tried combined cycle gas turbine/ramjet and separated the scramjet portion almost entirely. If they did succeed, they'll try to keep it secretly flying for a while, like they did the SR-71. It'll be a few years before we know.
I know about the old X-7 and the stovepipe ramjets that pushed it. Those have a flight speed limitation of about Mach 3-ish, due to melting of the flameholders and perforated air-cooled combustor liners and nozzle hardware. Above about Mach 3-3.5, there is simply no such thing as cooling air. That has been known since the 1950's. We've known how to get around those limits since the early 1970's, using sudden dump flameholders, and ablatives for our combustor liners and nozzle hardware. Nothing else actually works.
EXACTLY those technologies were in ASALM-PTV, a ramjet cruise missile prototype, flight tested successfully in 1980. This thing was boosted to ramjet takeover (NOT scramjet!!!) at Mach 2.5, then flew in ramjet power to cruise at Mach 4 and 80,000 feet. It would then dive onto its target at an average Mach 5. In one flight test, we had a throttle runaway failure, and it reached an unintended Mach 6 at only about 20,000 feet! It would have melted and broken up, had this been more than a several-seconds-long transient event.
That hypersonic airbreathing speed record stood until 2004, when NASA broke it with its hydrogen scramjet X-43A, in 3 second burn at Mach 7, boosted all the way to test speed by a rocket, and conducted at just over 100,000 feet (where the scramjet had insufficient thrust to accelerate) in order to limit aeroheating to survivable values. ASALM would cruise for several minutes at Mach 4, 80,000 feet, using a synthetic kerosene as its fuel, and demonstrating steady-state heat protection. We tested the combustor ablative liner as good for 15 minutes, steady state, by retaining the charred-through char as insulation. I'm not going to tell you exactly how we did that, but it WAS a breakthrough!
I have yet to see the limiting concerns properly addressed for airplane launch of an orbital rocket. I detailed those in point 4 of post #2 above. The easiest solution is a low-supersonic airplane with mixed turbine and rocket propulsion, carrying the orbital rocket on its belly. It must pull-up sharply to a path angle near 45 degrees, while maintaining speed (only the mixed propulsion can do that), and drop its rocket directly onto the gravity-turn ballistic path that lets the rocket function without wings for pull-up. If done at Mach 2, that's about 0.6 km/s speed at rocket ignition on a gravity-turn trajectory, leaving roughly 8.5-8.7 km/s delta-vee to be had from the rocket. You could do that single stage with LOX-LH2, or maybe even LOX-LCH4. Might have to be 2-stage if using LOX-RP1. Probably little or no reusability potential there.
Mixed jet/rocket propulsion was a highly-successful technology used for the high-speed research planes from the late 1940's into the early 1970's. The most striking example was the NF-104.
GW
Nice find, Spacenut. That explains what all the talk about a "planned engine out" really was. They had one that would not light, and decided to fly without it. That's where the advantage of so many engines shows up: 1 out of 33 is only a 3% loss of thrust. That one was in the outer ring of 20.
The boostback burn used the inner ring of 10 plus the 3 center engines that gimbal. I saw all 13 burning there.
It was the landing burn that initially used the same 13, transitioning to only the center 3 for touchdown. Only 9 of the inner ring of 10 were lit for that burn (which was the second engine-out), and one could see subsonic flame out to one side blown back up along the side of the stage as it landed. Looked rather like a methane-air fire, most likely a result of the engine that was out. Probably tried to light and had a failure of some kind, that damaged plumbing somewhere causing a methane leak. Had oxygen also spilled, the fire would have been instead an explosion.
GW
Point 1 -- hypersonics is being oversold in the press releases. It has been possible to push payloads hypersonic for many decades now, using rocket propulsion; it's just that the powered range is short. The now-retired AIM-54 Phoenix missile used by the Navy peaked at Mach 5 speeds way back about 1980. It had almost 100 mile range with a peak altitude of near 100,000 feet, on a long arcing trajectory that came back down on its target. The other decades-old hypersonics technology is the "hypersonic boost-glider"; in the 1970's we called them "MaRV" for maneuvering re-entry vehicle. These are simply aerodynamically-maneuverable warheads on ICBM's. There is not the time available for true maneuvering, this technology really only makes the targeting more accurate.
Most of the "hypersonic" things being touted today are either rocket-propelled tactical or strategic cruise missiles, or else maneuverable warheads on ICBMs. My, how things stay the same when they change!
The airbreathing hypersonic things are lagging behind, precisely because scramjet, while sort-of operational, is still technologically difficult and still very "iffy" in terms of reliability. Plus, they are still trying to combine it with turbine, when the inlet duct and thrust nozzle geometries are, quite frankly, utterly incompatible. Fool's errand, that!
Point 2 -- It takes big heavy wings to lift big heavy weights in any sort of airplane, no matter how it is powered. Why carry those big heavy wings to orbit and back, when you do not need them above the stratosphere? That kills any rocket mass ratio you could have. I'll give you one guess as to why horizontal takeoff vehicles to orbit have never materialized. And why it is unlikely they ever will.
Point 3 -- All airbreathing propulsion has an altitude limit called "service ceiling", no matter what kind of airbreather. That includes piston, gas turbine, ramjet, and scramjet, or anything else you want to dream up. This is because the thrust it produces is proportional to the combustion pressure, in turn a fairly-constant ratio to the local atmospheric pressure, at ANY altitude. So, thrust decreases as air pressure decreases. Period! End of issue!
Weight does not decrease with decreasing air pressure. So there is an altitude at which you can no longer climb, nor accelerate. Period! End of issue! And with airbreathing jet propulsion of just about any kind imaginable, it is nearer 80-100,000 feet than anything much higher. Riding only airbreathing propulsion to the edge of space at near-orbital speeds is technically-ignorant marketing bullshit trying to land a pointless technology development program from the government! Always has been, always will be!
Point 4 -- it has been established over the last few decades that what things are the most important for getting things to space from an airplane are: (1) speed at staging off the airplane, (2) flight path angle at staging high enough to go ballistic without lift, and (3) altitude at staging, in THAT ORDER of effectiveness. Not reversed, that order!
Speed is obvious: higher is better. But you must pay for it!
Path angle is NOT obvious, until you think about what it takes to pull up steeply at high speed in very thin air. It takes big heavy wings to generate enough lift to pull up at all, in thin air, and it takes a long time, since the path radius of curvature and resulting path length are large. All that time you are inherently taking on very large amounts of drag-due-to-lift, which can far exceed the zero-lift drag of any vehicle configuration! And THAT adds greatly to your drag loss you have to cover! The faster supersonic you are, the worse this gets, exponentially! This is EXACTLY why the air-launched Pegasus was no more successful than it was, even with subsonic staging from the carrier plane, under 40,000 feet.
Point 5 -- If you want to launch a rocket stage from an airplane, the configuration should indeed resemble the B-58, where that pod underneath fills the same role, but is your rocket-to-orbit instead of a dropped bomb. You can stage supersonic, but you need to do it at relatively low altitudes down nearer 20-30,000 feet where the plane can reach climb path angles on the order of 45 degrees. The B-58 climbing that steeply might only reach about Mach 1.5, which is only about 0.45 km/s velocity. Point is, the airplane has to make the pull-up, so that the dropped rocket needs no wings.
Point 6 -- The other ~9 km/s has to come from the rocket stage. You won't do that single-stage with anything less than LOX-LH2 at any sort of believable inert mass fraction. If you use LOX-RP1, or likely even LOX-LCH4, your dropped stage needs to be a 2-stage rocket. Reusability will be in serious question, in either case. (Pegasus was multi-stage. It was a solid.)
GW
From the ‘Daily Launch” for 6-12-2024, following the link to a “Space News” article dated 6-11-2024, a quotation of the entire Space News article”
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WASHINGTON — NASA confirmed that Boeing’s CST-100 Starliner spacecraft has suffered a fifth, although minor, helium leak in its propulsion system as engineers work to prepare the vehicle for its return to Earth next week.
In a June 10 statement, NASA mentioned that spacecraft teams were examining “what impacts, if any, five small leaks in the service module helium manifolds would have on the remainder of the mission.” That was the first reference to there being five leaks in the spacecraft; NASA had mentioned there were four in a briefing hours after the spacecraft’s June 6 docking with the International Space Station.
In a June 11 statement to SpaceNews, NASA spokesperson Josh Finch said the fifth leak was detected around the time of that post-docking briefing. “The leak is considerably smaller than the others and has been recorded at 1.7 psi [pounds per square inch] per minute,” he said.
NASA was aware of one leak at the time of Starliner’s June 5 launch, having been detected shortly after a scrubbed launch attempt May 6. At the time of launch, NASA and Boeing officials considered that a one-off problem, likely caused by a defect in a seal. However, hours after launch controllers said they had detected two more leaks, one of which was relatively large at 395 psi per minute, said Steve Stich, NASA commercial crew program manager, at the briefing.
A fourth leak was found after docking, although it was much smaller at 7.5 psi per minute. “What we need to do over the next few days is take a look at the leak rate there and figure out what we go do relative to the rest of the mission,” Stich said at the briefing.
NASA closed the helium manifolds in the propulsion system after docking to stop the leaks, although they will have to be opened to use the spacecraft’s thrusters for undocking and deorbit maneuvers. NASA said June 10 that engineers estimate that Starliner has enough helium to support 70 hours of flight operations, while only seven hours is needed for Starliner to return to Earth.
In addition to the helium leaks, engineers are studying one reaction control system (RCS) thruster that shut down during the spacecraft’s flight to the ISS. Four other thrusters were turned off by flight software but later reenabled. An RCS oxidizer isolation valve in Starliner’s service module is also not properly closed.
“We have the commercial crew program, Boeing, ISS teams all integrated, working very well together in order to come up with a forward plan for getting us in the best posture for that undock and reentry,” Dina Contella, NASA ISS deputy program manager, said at a June 11 briefing about a series of upcoming spacewalks at the ISS. “The teams are still working through what are the best ways to go about testing and preparing for undock and reentry.”
Those teams have some time to complete that work. NASA had initially scheduled a June 14 undocking for Starliner, but NASA said June 9 it was delaying the undocking to no earlier than June 18. That delay was to avoid a conflict with a June 13 ISS spacewalk, or EVA, by NASA astronauts Tracy Dyson and Matt Dominick.
“To have it back to back, were we had an EVA followed by undock, was not the most convenient,” Contella said. There are undocking opportunities every few days, governed by the orbital mechanics that set up a landing in the southwestern United States.
The two NASA astronauts who flew Starliner to the ISS, Butch Wilmore and Suni Williams, have been busy both conducting tests of Starliner while at the station while also performing other work, such as science experiments. “Butch and Suni are an extra set of hands,” she said, particularly as other ISS crewmembers prepare for the upcoming spacewalks. “Having Butch and Suni available to perform some key critical science has been outstanding.”
Wilmore and Williams have publicly praised the performance of the spacecraft. “The spacecraft was precise, more so than I would have expected. We could stop on a dime, so to speak,” said Wilmore during a June 10 call with NASA leadership, discussing how the spacecraft maneuvered.
“Our experienced test pilots have been overwhelmingly positive of their flight on Starliner, and we can’t wait to learn more from them and the flight data to continue improving the vehicle,” Mark Nappi, Boeing vice president and commercial crew program manager, said of the astronauts in a June 11 statement.
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My own take on this:
Given the problems related to decades-old plumbing technologies (helium leaks, RCS oxidizer valve, and RCS thrusters), I see gross inattention to detail by Boeing on this design! As an overall spacecraft, it functions well (“stop on a dime” comment). So also did the 737 MAX function very well. But I see the same sort of inappropriate management culture with Starliner as what killed two planeloads of 737 MAX passengers.
In the case of the airliner, a piece of automated flight control intended as a stall prevention (the MCAS) was fed data from ONLY ONE of two angle-of-attack (AOA) sensors, despite those sensors being well-known to have a high failure rate. How to deal with this MCAS when it seriously misbehaved upon AOA sensor failure was LEFT OUT of the pilot’s operating handbook, along with any mention that there was an MCAS installed at all (in violation of the FAR’s), just to enable a lower-cost overall product (this lack “justified” less pilot transition training). It was precisely the lack of pilot training for how to deal with a misbehaving MCAS that killed the two planeloads of people!
And here is Starliner, flying great (now that the originally-defective flight control software has been fixed), but afflicted with plumbing leaks that absolutely should NOT be happening! It is NOT the same design team within Boeing, but it IS the SAME top management culture! And it glaringly shows!
Why do you think the flight control software was originally defective, causing the public spectacle for the repeat unmanned flight test to ISS? Could that management culture have shorted technical excellence to increase short-term profit or decrease short-term costs?
That is EXACTLY what I think has been causing all the problems that have shown up in all recent Boeing products, meaning things designed since about the mid 1990’s.
And NASA’s management culture has never really been fixed even after losing 2 shuttle crews! If it had been, they would have seen these problems with Starliner coming, BEFORE they cropped up in flight. But, no, they can only RESPOND AFTER THE FACT (which is EXACTLY what the second uncrewed flight test to ISS was, a response to an unforeseen problem). Schedule and money are still trumping flight safety at NASA, as well as at Boeing.
In my opinion, Starliner should NOT be certified “ready” until these plumbing problems and balky thruster problems are fixed. If they don’t do that, then if it flies crewed enough times, the odds will inevitably line up to kill a crew.
GW
Void:
I used both links, and they both worked correctly this time! I have no clue what happened the first time I used the first link you posted. All I did was click on the link. It took me to some gal's ad for an unproven weight loss technique that had something to do with coffee. Followed immediately by another ad that I simply did not watch.
GW
I followed the link and all I saw was a series of commercial ads, not anything to do with Starship.
Void, I think that may be the wrong link. Check it out and correct it, if it is.
GW
The unload ramps are cylindrically-curved panels, 4 of some 8 up each aside. These are conceived as metal shells exposed to the environment, covered internally with a layer of something like mineral wool insulation, and a thin layer of plastic sheet to cover up and hold the insulation in place. There might be a couple of stringers that add extra bending strength, just to construct the things.
The curved shape offers quite a bit of extra stiffness in bending compared to a flat panel. However, since this is a one-shot lander, it does not matter if the panels get bent a bit during unload operations. Their real value after unload is their salvage value. Both the metal and the insulation have reuse value. The plastic sheet, not so much. Maybe as dust door mats or something.
GW