An Ex Rocket Man's Take On It: launch

Showing posts with label launch. Show all posts

Monday, September 1, 2025

On Pressure Vessels

One can build an unpressurized liquid tank to any desired shape, unless it is so large that the depth of liquid inside exerts a significant pressure all by itself. The same is not true of pressurized vessels, even small ones, beyond a “single handful of psi” gage pressure inside it. There are very good reasons these tanks are only made in spherical shapes, or as right circular cylinders with hemispherical or spherical-segment ends.

This document presumes the reader knows what “gage pressure” means, and what mechanical stresses and strains are.

Figure 1 shows a cylindrical pressure vessel holding gage pressure P inside. This could be a tank, or even a pipe or a tube. The “hoop stress” is given by the formula shown top right, which is sometimes referred to as Barlow’s pipe stress formula, which is a measure of the stress trying to split the cylinder open longitudinally. Technically, you use the cylinder inside diameter for this formula (a requirement working with pipe and tubing), but as long as the material thickness is small compared to the diameter, something true of rocket cases and propellant tanks, there is little-or-no perceptible difference between inside, outside, and average diameters.

Bottom left, the cross-section of the cylinder, or any section through a sphere, is shown. For a cylinder, the “axial stress” is that which resists the pressure trying to part one end of the vessel from the other. For a sphere, it’s just the “membrane stress” that resists splitting the sphere apart, no matter the section orientation. That stress works out to be just half the hoop stress, as shown.

Figure 1 – Pressure Vessel Stresses In the Vessel Material

For a cylindrical pressure vessel fitted with hemispherical ends, Figure 2 takes this notion a bit further, showing as it does the stress distributions on a tiny patch of material, located on either side of the joint. The axial stresses match, while the hoop direction stresses differ by a factor of 2. There are (at least theoretically) no stresses perpendicular to the material itself (the radial direction from the cylinder axis, or from the hemisphere’s center).

This hoop stress mismatch at the joint corresponds to a strain mismatch in the circumferential direction, in turn corresponding to a radial displacement mismatch between the cylinder material, and the end (or “head”) material. The cylinder swells radially under pressure twice as much as the end or “head” swells radially, as measured at the joint. This distorts both the cylinder and the head locally at the joint, as the materials bend locally, in order to try to stay joined. This induces large bending stresses locally, which add in certain ways to the hoop and axial (or membrane) stresses already described. That makes the joint quite vulnerable to local overstress failure.

The ”fix” for this is to locally thicken the cylinder and head materials at the joint. In effect, the extra material “sops up” the extra imposed stresses. For boilers, there are very specific guidelines for how much local thickening is needed, and how far “local” extends away from the joint. Those rules are the ASME boiler code, which is legally mandatory everywhere in the country, for designing and building boilers. Every provision represents a life lost learning that lesson. This is serious business!

Figure 2 – There Are Distortions With Extra Stresses At the Joints

There are choices allowed for how to implement those localized thickenings at the joints. Those are depicted in Figure 3, and also apply to solid rocket motor case designs. You can increase the thickness toward the inside, or toward the outside, or even some of both, just as long as enough extra material is supplied. Finite-element stress-strain analysis can refine this further.

The same figure also shows a variation on the spheroidal end, where only a sphere segment is used as the end membrane. This requires a connection ring that resists radial swelling at about the same rate as the end resists radial swelling at its attachment joint. That way, no thickening of the membrane is required, the ring supplies that for the membrane. You still need a local thickening of the cylinder material at its attachment to the ring, because of the radial swelling mismatch.

This spherical segment and ring approach lets one enclose more volume within a given length, without making the assembly any heavier than a full hemispherical end. This is how most solid rocket motor case closures are designed. It is a well-proven solution. The flatter the membrane, the heavier the ring gets, though. It’s a trade-off.

Figure 3 – Full Hemispherical Ends Vs. Spherical Segment Ends With Rings

There are many possible reasons for wanting to use other shapes for one or both cylindrical pressure vessel end heads. As long as membrane stress is insignificant (meaning very low pressure indeed!), you can do that! But as soon as the pressure (and the stresses it induces) become significant, those other shapes rapidly become infeasible. This is shown in Figure 4, where the two spheroidal options are depicted, along with elliptical and conical ends. A flat end fails even worse than the elliptical, from similar stresses that are just higher, plus a sharp corner effect that locally greatly magnifies the local stresses even further, right at the corner.

There is one positive benefit to a conical end, if an axially-directed load must be carried from the cone tip into the cylinder. This is an efficient load path for such a load. But it is still a lousy pressure vessel choice! It will require a lot of internal stiffeners to keep it from trying to “go round”. Those are going to add significant weight, there is no way around that problem! You must trade off the axial load path advantage against the big weight gain incurred to make a conical shape a pressure vessel.

Figure 4 – Which End Shapes Work and Which Do Not, and Why

It is a common belief that an elliptical shape is as good as the ring and spherical segment. This is not true from a pressure vessel design standpoint, as the figure shows. The elliptical head will try to “go round”, inducing severe bending stresses. It is volumetrically efficient, which is why many unpressurized railroad tank cars use elliptical heads on cylindrical bodies. But these tank cars are not pressurized! Their evident abundance is deceptive regarding the pressure vessel issues.

That same effect is why you want to use circular cylinders as your basic tank body for a pressure vessel, not some elliptical (or other) cross section shape. Those other shapes will try to “go round” upon pressurization, leading to enormous bending stresses and very rapid failure. If you must put pressurized storage within some oddly-shaped volume, you must fill it with multiple small circular cylinders! The non-circular cross section is not, and will never be, a successful pressure vessel design! This is why air mattress floats are made the way that they are, for example: multiple cylinders, connected at the ends so as to fill together all at once.

Addendum: Exact Analysis

The exact analysis for right circular cylinders and spheres is given in Figure 5, along with the geometries that allow these formulations to be made from very simple measurements.

Figure 5 – Exact Formulations

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Search code 01092025

Search keywords launch, ramjet, space program

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Friday, August 1, 2025

Air Launch to Low Earth Orbit

There is a conundrum associated with launching to low Earth orbit from an airplane. The illustration tries to sum up the various parts of it. That is not to say that it cannot be done, because it already has. But, it may, or may not, be an attractive way to do the mission.

The first part of this conundrum is the low speed of the launch aircraft (which for the Pegasus system is a wide-body subsonic airliner). That forces the dropped rocket vehicle to be at least two-stage, despite the advantage of the low stratospheric launch altitude. As it says in the illustration, speed at drop is the biggest influence on the rocket vehicle design, and altitude the least, although both are beneficial. Mach 0.85 at 45,000 feet is but 822 feet/sec (0.25 km/s). The drag loss of the rocket vehicle is (at least theoretically) less, because it starts in thinner air up high.

The second part of this conundrum is not so obvious: the level path angle of the carrier airplane at the drop point. A low-loss non-lifting ballistic trajectory begun at stratospheric altitude would need a path angle at ignition on the order of 45 degrees, maybe even a little more. So, either the carrier airplane, or the rocket vehicle, has to pull up rather sharply, to reach that path angle from level flight. One or the other must do this!

The usual airplane flying high in the stratosphere is at or near its “service ceiling”, where there is barely enough wing lift being produced at an efficient angle of attack, to hold up the weight, and essentially all the thrust the airbreathing engines can make is just overcoming drag at the flight speed! The airplane can neither accelerate path-wise, nor can it climb! That is the definition of “service ceiling”, and for most planes, it falls in the 45,000-55,000 foot altitude range, at high subsonic speeds. There have been exceptions: the U-2 variants and the SR-71 variants could fly higher, being very specialized designs.

Left unaddressed in the airplane, the service ceiling problem puts the sharp pull-up task squarely upon the rocket vehicle to be dropped. There are only two choices: put wings on the rocket vehicle, or fly it at very large angles of attack, so that the cross-path vector component of its thrust is effectively a large lift force.

Pegasus used large wings, on the first stage of a two-stage rocket vehicle. Those add both weight and drag, especially drag-due-to-lift at the large lift coefficient needed to pull up sharply. That pretty-well eats up the advantage gained by launching the rocket at elevated altitude in the thin air. The wings are bigger than you would want, precisely because of that thin air! And that problem is why there have just not been that many Pegasus launches.

Leaving the wings off of the rocket vehicle forces you to pitch it up to very large angles of attack, in the 45-75-degree range, to get enough of a cross-path vector component of the rocket thrust, to serve as the necessary lift force for a sharp pull-up maneuver. That reduces the path-wise vector component of thrust, while at the same time greatly increasing vehicle drag. So, you accelerate slowly( if at all) in rocket thrust during the pull-up maneuver, using up a great deal of rocket propellant that adds nothing to your speed. That also eats up any advantage of launching in the thin air, way up high!

The only other feasible alternative is to add another large source of thrust to the launch airplane, so that it can execute the pull-up maneuver into a zoom climb, without stalling and falling out of the sky, out-of-control. Generally speaking, you would add a source of thrust immune to the service ceiling effect, and that is rocket thrust! Your launch airplane would have to be modified for mixed (parallel-burn) rocket and gas turbine propulsion, somewhat similar to the NF-104 and some of the early high-speed X-planes.

So far, no air-launch carrier plane has had this design approach, but it certainly would be possible! And it would take care of the high path angle requirement that is second only to speed at launch in importance, while keeping the wings on the airplane where they belong, and not on the rocket vehicle!

That leaves speed at launch, the most important variable affecting the rocket vehicle design. There are (or have been) very few supersonic aircraft designs that are also large enough to serve as a drop aircraft for a rocket vehicle of any significant size. Those would include the B-58 Hustler (long-retired, and none are left), the SR-71 (also retired, but very expensive to operate indeed), and the B-1B bomber (currently in service as a military strategic bomber).

The modifications to include rocket propulsion to the SR-71 likely would not fit within its very-critical shape. The M-21 variant that launched the D-21 drone was limited in payload size, to the size of that drone (not very large). A rocket might be added in the tail cone of a B-1B, but its payload would be limited to that which would fit in its bomb bay. That B-1B option would reach a low supersonic launch speed at the high path angle needed, with a rather-dangerous zoom climb and recovery after drop.

That brings up the danger of supersonic store separation. There is a very good reason that most military aircraft, even those capable of supersonic flight, are limited to high-subsonic weapon release speeds. That is because the inherent wobbles of a released store will include pitch-up, thus developing lift. At high enough speeds, that lift generated by the wobbling store will exceed its weight, and it can easily fly up and collide with the drop aircraft, before the store’s drag can pull it behind.

It cost a destroyed airplane and the life of one of the two crew, to learn this lesson with the M-21 trying to launch a D-21 drone (without a booster) at just about Mach 3. That is why the drone was re-fitted with a big booster, and launched subsonically from B-52’s instead. It’s not that supersonic store separation cannot be done (because that booster separated at Mach 3 from the D-21). But successful supersonic store separation is very difficult to achieve, and the risks of doing it are inherently very high.

So how fast a drop speed can be obtained? That depends upon the gas turbine engines powering the launch aircraft. Those are seriously limited by the high air temperatures associated with capturing supersonic air. Most are limited to about Mach 2.5. There are a very few that went faster: those powering the XB-70 at Mach 3, those powering the SR-71 variants at Mach 3.2, and the 500 hour short-life, replace-don’t-overhaul engines in the Mig-25 at Mach 3.5. So, to have a wide range of possible engines available for new designs, it looks like Mach 2.5 at drop is “about it” with gas turbine. Maybe Mach 3.

So, the answer would seem to be a mixed-propulsion airplane with gas turbine propulsion, augmented by parallel-burn rocket propulsion, added to enable the zoom-climb by a sharp pull-up maneuver. This would be at high altitude near 45,000 feet, for the drop of the rocket vehicle. To do this successfully, the very difficult supersonic store separation problem must be very carefully addressed! Both aircraft and crews are at serious risk.

Mach 2.5 at that altitude would be 2419 feet/second (0.737 km/s), less than 10% of low circular orbit speed, so one is still looking at a two-stage rocket vehicle to reach orbit. Deliverable payload would be limited in size by the size of the drop aircraft, since that in turn limits the size of the rocket vehicle it can drop.

In a word, this has already been done with subsonic carrier aircraft, although it has proven no more attractive than vertical rocket launch, at best. The supersonic release has yet to be tried, and will prove both difficult and dangerous, although the improvement in attractiveness may be worth that effort and risk. No one yet knows.

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Search code 01082025;

search keywords: aerothermo, airplanes, launch, space program

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Update 8-2-2025: Please do not misunderstand, air launch to LEO is possible and in fact has been done more than once! It's just not easy, because many of the problems associated with it are hard. They are hard enough that the attractiveness of this approach is still in question, relative to the tried-and-true vertical rocket launch.

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Update 8-4-2025: For an air launch-to-orbit carrier aircraft, the gas turbine speed limitation could be gotten around by instead using ramjet propulsion, which for a true high speed design might reach speeds between Mach 3 and Mach 4 in the stratosphere, limited mainly by atmospheric drag of something inherently not a “clean” missile shape.

One would still need the rocket component of a mixed-propulsion parallel-burn scheme to achieve the necessary climb angle at launch of the rocket payload, and one would still need to solve the dangerous supersonic store separation problem. But this would get the highest possible speed at launch, at the right launch angle, and at an altitude high enough to be beneficial.

The downside is that ramjet has no static thrust! You will need some sort of booster to reach ramjet takeover speed, and the necessary high-speed ramjet design is going to have a takeover speed in the Mach 1.8 to 2.5 range. Given that rocket is needed to reach the high climb angle at launch, that same rocket is likely the propulsion needed to reach takeover speed.

Speeds will be limited by the percentage of frontal blockage area occupied by each of the two propulsion systems. The airbreather is fundamentally lower in frontal thrust density than is the rocket, so it needs to occupy the larger fraction of the total frontal blockage area.

Being a lower percentage of vehicle frontal blockage area than the ~100% of a “clean” missile design, the max possible speed capability of a ramjet (near Mach 6) cannot be reached with this kind of a vehicle. But the ramjet weighs far less than any possible turbojet propulsion! That makes a smallish rocket system feasible for getting off the ground with wings, and reaching Mach 1.8 to 2.5 takeover speed at relatively low altitude.

From there, you climb in ramjet to high altitude at speeds near Mach 2.5, and pull over level to accelerate to top speed in the thin air. Fire up the rocket to climb steeply for the supersonic store separation, then shut down the rocket and throttle-back the ramjet to execute a zoom climb and descent back into air dense enough to support controlled flight. Cruise back in ramjet, then glide to a landing with the rocket in reserve for go-around capability.

The real trade-off here, yet to be evaluated, is whether to integrate the two propulsion systems into some sort of combined-cycle rocket-ramjet, or leave them as separate systems to be operated entirely separately. Combined-cycle usually seriously compromises the performance of both components, while parallel-burn does not, instead running into the fraction-of-frontal area problem.

And there is also the problem of there being “no such thing as cooling air” above about Mach 3 to 3.5 in the stratosphere. Vehicle designs flying faster than that will need one-shot ablatives for their ramjet combustor and nozzle heat protection. Which means you must swap-out the entire combustor and nozzle unit after every flight! Given that eventuality, you could do a solid propellant integral booster in the combustor and nozzle unit, like a great big JATO motor, for the initial takeoff. That reduces the volume (and cross-sectional area) of the on-board propellants for the liquid rockets.

None of these issues have been resolved for an air launch-to-orbit application.

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Monday, June 23, 2025

Starship Explosion

Update 7-5-2025: search code 23062025.

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From AIAA’s “Daily Launch” email newsletter for Monday, 6-23-2025. This ship was intended for Flight Test 10. It blew up before they ever ignited the engines. Quote:

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SpaceX traces Starship test-stand explosion to failure of pressurized nitrogen tank

By Mike Wall published 3 days ago (on Space.com)

"Initial analysis indicates the potential failure of a pressurized tank known as a COPV."

SpaceX thinks it knows why its newest Starship spacecraft went boom this week.

The 171-foot-tall (52-meter-tall) vehicle exploded on a test stand at SpaceX's Starbase site late Wednesday night (June 18) as the company was preparing to ignite its six Raptor engines in a "static fire" trial.

A day later, SpaceX narrowed in on a likely cause.

"Initial analysis indicates the potential failure of a pressurized tank known as a COPV, or composite overwrapped pressure vessel, containing gaseous nitrogen in Starship's nosecone area, but the full data review is ongoing," the company wrote in an update on Thursday (June 19).

"There is no commonality between the COPVs used on Starship and SpaceX's Falcon rockets," the company added. So, launches of the workhorse Falcon 9, which has already flown 75 times in 2025, should not be affected.

The Starship explosion did not cause any reported injuries; all SpaceX personnel at Starbase are safe, according to the update. People living around the site, which is near the border city of Brownsville, shouldn't be worried about contamination from the incident, SpaceX said.

"Previous independent tests conducted on materials inside Starship, including toxicity analyses, confirm they pose no chemical, biological, or toxicological risks," the company wrote. "SpaceX is coordinating with local, state, and federal agencies, as appropriate, on matters concerning environmental and safety impacts."

That said, the explosion did damage the area around the test stand, which is at Starbase's Massey site (not the orbital launch mount area, from which Starship lifts off).

"The explosion ignited several fires at the test site which remains clear of personnel and will be assessed once it has been determined to be safe to approach," SpaceX wrote in the update. "Individuals should not attempt to approach the area while safing operations continue."

Wednesday night's explosion occurred during preparations for Starship's 10th flight test, which SpaceX had hoped to launch by the end of the month. (Static fires are common prelaunch tests, performed to ensure that engines are ready to fly.) That timeline will now shift to the right, though it's not clear at the moment by how much.

The incident was the latest in a series of setbacks for Starship upper stages. SpaceX lost the vehicle — also known as Ship — on the last three Starship flight tests, which launched in January, March and May of this year.

Starship's first stage, called Super Heavy, has a better track record of late. For example, on Flight 7 and Flight 8, the huge booster successfully returned to Starbase, where it was caught by the launch tower's "chopstick" arms as planned.

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My take: if the description “in the nosecone” for the location of the COPV is correct, then it is located very close to the oxygen header tank (as the version 1 with 1200 metric tons propellant capacity was laid out), which is also in the nose of the vehicle, ahead of the “cargo bay” area. Such a COPV explosion would easily rupture that oxygen header tank. Compressed gases drive great explosive violence (with shrapnel) when such vessels burst. See Figure 1.

Figure 1 – Inboard Profile of Starship Version 1

There would seem to be an oxygen header tank transfer piping line down the windward “belly” of the cargo bay section, based on descriptions I have read. In the explosion slow-motion video, the cargo bay splits open through its heat shield, right where that transfer line supposedly is, with gush of something white (not fire) bursting through, followed immediately by an explosion engulfing about the top half of the vehicle, and a second or so later by a second explosion seemingly centered lower down.

The main propellant tanks below the cargo bay would be the main methane tank forward, with the methane header tank located inside, at the base of that tank, and finally the main oxygen tank, just ahead of the engine bay. The upgraded version 2 has a bigger propellant capacity, but should be laid out similarly.

I would hazard the guess that the COPV explosion and bursting oxygen header tank somehow put a large force on the transfer line, which split open the belly at the cargo bay, allowing liquid (and vapor) oxygen out through that split, as well as releasing a few tons of liquid oxygen to fall down on top of the main methane tank.

My guess is that spilled header oxygen and vented methane vapors are much of the first explosion. Bear in mind that the impact of a few tons of liquid oxygen on the top of the main methane tank would rupture it as well, adding some fuel to that first explosion pulse. That first explosion pulse would massively rupture the main methane tank, and also likely the main oxygen tank below it. That’s the second pulse of the explosion, which was larger and longer, reflecting the larger mass of reactants.

All of that scenario is just an educated guess on my part.

As for the nitrogen tank, said to be a “COPV”, or “composite overwrapped pressure vessel”, maybe that is not the right choice this early in the flight test program. Such a design is a metal shell that is simply too thin to hold the pressure, overwrapped by a yarn or fabric-reinforced composite material, to bring it up to strength at a lighter weight.

Here’s the problem: no composite material has a large plastic (post-yield) strain capability. If the COPV over-pressures for any reason whatsoever, failure will be sudden, without any warning! Maybe a heavier all-metal nitrogen tank, one with much more plastic strain capability, would be a better choice until the other bugs all get worked out. At least you could see it stretch before it explodes. You do not want to fly even experimentally, with too many possible failure modes! See Figure 2.

Figure 2 – Stress-Strain Curves for Low and High Plastic Strain Capability

Lots of things look good “on paper”, but there are a lot of other things to worry about, many of which cannot be put on that paper. This is where the “older hands”, with many years of school-of-hard-knocks experiences, can be effectively very much wiser than youngsters fresh out of school. SpaceX has no “old hands” on its staff: they hire no one over about age 40 or 45. There’s no gray heads visible anywhere in that organization.

PS: note in Figure 1 the "typical" inert mass of the Starship Version 1 upper stage, as 120 metric tons. I do not have a figure for the Version 2 inert mass, but it simply cannot be very much different from Version 1! They are all built the same way. The hearsay bandied about on the internet, about 80 tons, or even less, is simply BS!

Sunday, June 1, 2025

SpaceX “Starship” Flight Test 9

Update 6-3-2025: A close version of this article ran in today's Waco "Tribune-Herald" as a column on the opinion page.

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The recent Flight 9 test resulted in the loss of both the upper stage spaceship and the lower stage booster. A lot of stories class this as just another failure in a series of three, but it is a little more complicated than that.

The “Superheavy” booster portion of this test was actually a lot more successful than many believe. They never intended to recover this one, instead choosing to stress it possibly to destruction, on a return flight profile that was likely fatal. And that is exactly what happened.

On the plus side, this was the very first re-flight of the reusable “Superheavy” booster. This very same stage launched the Flight 7 “Starship” test, with most of those 33 engines reused on this Flight 9 test. What was “new” on Flight 9 was twofold: (1) the deliberate overstress return, and (2) the demonstration of a “quick flip” during hot staging.

The “quick flip” was entirely successful! Until this flight, after staging, the booster would flip slowly using only vectored thrust from a few of its engines, which also settled the propellants in the tanks so that the engine pumps could maintain a good suction.

This time around, the hot stage ring that protects the forward end of “Superheavy” from the hot blast of “Starship’s” engines during staging, was built to an asymmetric shape, so that “Starship’s” exhaust was deflected out one side only. The force of that asymmetric exhaust deflection flipped “Superheavy” around much faster, saving propellant for the landing. That was completely successful!

The stressful return involved flying the booster at an angle to the oncoming wind, instead of streamlining straight in, tail first. That being off-angle increases the aerodynamic drag, so that less braking propellant might be used. But, there are structural limits to how much off-angle is survivable.

They found that limit experimentally, essentially breaking up that “Superheavy” with too much wind load coming from the side, where it is nowhere near as strong. That is something very important to know “for sure”. It is just like pulling too many gees that can rip the wings off an airplane! You must know your limits!

The Flight 9 “Starship” upper stage spacecraft did not experience engine failures during the ascent burn from staging to orbital insertion, unlike what happened on Flights 7 and 8. It appears that they may have fixed at least a part of what was going wrong on those two flights.

However, there was a slow loss of propellant tank pressure, due a leak (or leaks) somewhere. Their attitude thrusters are cold gas thrusters powered by the propellant tank pressure. And there must be at least some pressure to successfully ignite an engine.

Whether this loss of propellant tank pressure had anything to do with opening the payload door is unknown to me, but not long after that door failed to open, is when the vehicle went into an uncontrolled spin for lack of attitude thrusters. And they called off the engine relight test because of it, too.

Not having attitude control is why the ship broke up upon reentry over the Indian Ocean. The heat shielding is only on the belly and around the nose, so that if you are pointed in the wrong direction, you are unprotected from the extreme heating of orbital reentry. The vehicle was spinning completely out of control as it entered.

SpaceX is going to have to figure out what caused the leak and loss of tank pressure. That won’t be easy. But it has to be done, and I hope they obtained the necessary data to figure out that problem, from this test.

Getting a reentry survival is crucial toward verifying their heat shield design. I would suggest that SpaceX add some independent attitude thrusters to these flight tests, so that loss of tank pressure does not mean loss of attitude control. The Draco attitude thrusters on their Dragon capsules would be a good choice, with lots of real experience behind them.

I do not know for sure, but had Flight 9 been equipped with independent attitude thrusters, that particular “Starship” might have made it back through reentry successfully. That would have been a huge “plus” for this mission! They will still have to find and fix the propellant tank pressure leak, in order to make the final landing burn, however.

Photo courtesy UPI

Thursday, May 1, 2025

Vehicle Assembly and Refueling Facility in LEO

Described herein is a concept (only) for a facility in low Earth orbit for the assembly and fueling of interplanetary vehicles requiring hyperbolic departure (and arrival), in particular those associated with space-tug assisted departures and arrivals. Such a facility need not be a 1-to-2-decade long international project to build, if it is docked together out of modules that fit within the payload spaces of the current launcher fleet! That should be easily achievable for a “clean sheet of paper” design like this!

For lunar missions, the departure from LEO is not hyperbolic, although it is elliptic at very-near-escape perigee speeds! Depending upon the choice of the extended departure (and arrival) ellipse, the LEO departure velocity requirement for a lunar mission can be reduced to near-zero, with the space tug assuming most or all of that velocity requirement just getting the craft onto the ellipse.

No calculations have been made, these results are concept only, as is perfectly reasonable at this early stage! The basic design concept has two core sections, one made of pressurized modules docked together, and the other a truss core to which a multitude of propellant tanks are attached.

Attached to one end is the Power and Propulsion Section, where solar electricity is made, stored, and distributed. This section includes propulsion sufficient to address the needs for countering orbital decay, conducting debris avoidance, and performing end-of-life safe disposal.

The pressurized-core section is the Vehicle Assembly and Refueling Section, where interplanetary vehicles are assembled from modules, mated to space tug vehicles as appropriate, and fueled-up from the propellant depot section for the relevant missions. Many remote-operated arms similar to those used at the International Space Station (ISS), and previously on the old Space Shuttle, are installed to make vehicle assembly and handling operations as safe and easy, as is possible. This is a manned microgravity facility, probably manned by rotating crews, as is the ISS.

The truss core section has multiple propellant tanks attached to it, with the propellant feed lines and electric power lines housed inside the truss. This is the Propellant Depot Section, presumed to be kept supplied by unspecified tanker flights up from Earth. It would have both cryogenics and storables, to supply a variety of on-orbit needs. Its capacity is also easily expandable.

The concept for the Vehicle Assembly and Fueling Section is sketched in Figure 1. The concept for the Power and Propulsion Section is sketched in Figure 2. The concept for the Propellant Depot Section is sketched in Figure 3. All figures are at the end of this article.

This kind of a facility would be easiest to keep supplied, if located in a low-inclination eastward Earth orbit. That presumes vehicle modules, propellants, and supplies are shipped up from the surface. It would clearly be useful in any event, but it is an essential enabling item for making use of reusable space tugs for elliptic departure and arrivals, as described elsewhere in Reference 1.

Vehicle Assembly and Fueling Section

As the sketches in the first figure indicate, this facility is built up from many cylindrical modules docked together, and each is to be small enough to fit in the payload spaces of the existing launcher fleet. Some of these are oriented along the section axis, and the others are perpendicular to it, but all are in one plane. These could be either hard-shell modules, or inflatables with hard structural cores, or a mix of both types. That choice remains unspecified, at this time.

The modules along the core axis provide much crew living space, lots of storage space for life support and other supplies, airlocks for space-walk activities, plus any equipment for Earth observations (potentially replacing those functions after the ISS is decommissioned). These modules would be equipped with external cradle mounts, to help hold the vehicles being assembled, thus freeing up the arms for other tasks that are part of the assembly process. Some of the hatches should be closed, when the modules are not in use by the crew.

The modules perpendicular to the core provide the spaces for the arm operators to work. The arms are affixed to the module ends. These modules need large windows, by which the arm operators can see their workpieces in order to work. Assembly work areas are disposed along this section, on two opposite sides. It should be able to handle a busy traffic load, if arranged in this way.

Per Reference 2, I am suggesting that this section’s internal atmosphere follow the “Rule of 43”, that being a two-gas oxygen-nitrogen system, at 43 volume percent oxygen, and 43% of a standard atmosphere total pressure. This is very close to the best atmosphere that I found (which was 43% oxygen, 43.5% of an atmosphere total pressure), and it is easier to remember! It has the same oxygen concentration (as mass per unit volume) as sea level Earthly air at 70 F, so the “predicted fire burn rate danger” from the usual Arrhenius overall-chemical rate equation, is no worse than that down here on Earth, at sea level on a warm day.

Further, the “pre-breathe criterion” allows no pre-breathe requirement be imposed for donning pure-oxygen space suits, of helmet pressures down to as low as only 3.002 psia (155.2 mm Hg)! That criterion says the ratio of nitrogen partial pressure to suit oxygen helmet pressure, may not exceed 1.2, in order to avoid the nitrogen blow-off time otherwise required. (The absolute minimum tolerable suit pressure for functional cognitive capability is 2.675 psia (138.3 mm Hg), before applying a 10% leak-down factor. The cognition margin is very slightly negative once leaked down.)

As a further bonus, the proposed oxygen partial pressure is the same as that at about 2500 m altitude, so there should be no long-term hypoxia risks, or even any reproductive health risks for female crew, based on centuries of human experiences living up to that altitude, but not above it.

Power and Propulsion Section

Most likely, the “best” propulsion choice for this application is a hypergolic storable bi-propellant system, pressure-fed for the greatest engine simplicity and reliability. Tanks would be bladdered, with inert gas (helium) expulsion at effectively the feed pressure to the engines. Propellants would likely be nitrogen tetroxide (NTO) oxidizer and monomethyl hydrazine (MMH) fuel, although the other hydrazines could also serve, which include plain hydrazine, unsymmetrical dimethyl hydrazine (UDMH), and Aerozine-50 (a 50-50 blend of plain hydrazine and UDMH). These tanks would need a thin layer of insulation topped with a very reflective aluminum foil, plus electric tank heaters to prevent freezing while shadowed.

There is a core module to this section that connects to the rest of the station on one end, and the engines and their propellant tanks on the other. It would have multiple “fins” mounted to the sides, some being waste heat radiators, the others being solar photovoltaic panels. There would be controls, batteries, and distribution switching equipment inside, plus a docking module for capsules bringing crew and supplies. This core module is pressurized for easy access, but the hatch into it should be closed, when crew are not working in there.

Propellant Depot Section

This section has a modular truss core containing the multiple types of propellant feed lines, and the necessary power lines. The propellant tanks are mounted to its periphery, as sketched in the third figure. There are basically two types of propellant tanks, those equipped to store and deliver cryogenics, and those equipped to store and deliver storables. Each propellant species must have its own line fittings, not interchangeable, so as to prevent incorrect hookups (which would most likely be disastrous). There are no pressurized modules in this section.

For the storables (which includes rocket-grade kerosene RP-1), the bladder in the tank provides the means to transfer propellant, driven by inert gas pressure that everts the bladder, as indicated in the third figure. These tanks will also need some insulation topped by reflective foil, and some in-tank heaters, much like the tanks on the Power and Propulsion Section. The difference is that the inert gas pressure can be lower, just enough to drive the transfer, and not at all far above the level to prevent “hot room temperature” boiloff.

The cryogenics are different, in that there are no feasible bladder materials that could survive eversion at cryogenic temperatures. These have to be metal tanks with no bladders, although they do need a layer of insulation topped by reflective aluminum foil, plus cryocooler equipment.

In zero gee, the propellant will initially be free-floating globules, eventually settling into a thick film coating the entire inner surface of the tank with a vapor void up the core, but with no pressure other than enough inert gas pressure to stop boiloff. The slightest touch causes the thick film to break up into free-floating globules again. Up to now, the only way to control this behavior into a stable pool from which a pump can draw suction, was to use thrusters to accelerate the vehicle. You can’t do that with tanks on a space station whose orbit you do not want to change.

I had previously come up with the spinning tank concept to fling the propellant to the outer wall by centrifugal force. From there, a pump could draw suction from openings along the tank sides instead of one end. This was conceptualized as the vehicle docking with the tank, in turn undocked from the station. The docked vehicle and tank would move away to a safe distance and then spin-up in “rifle-bullet” mode, to fling tank contents to the outer walls. Then pumped transfer could take place, followed by de-spin, then redocking the tank with the station, and finally undocking the vehicle from the station’s tank. While this would work, it does involve multiple docking operations, and spin-up/de-spin of some massive objects. But it was better than trying to store spinning tanks on the station. This concept was described in Reference 3.

I have since revised the concept to just spinning the propellant inside the stationary tank, by means of moving vanes inside the tank. The suction pickups remain on the outer periphery. If you use a pair of counter-rotating vane sets inside the tank, separated by a perforated baffle, you can avoid gyroscopic forces being applied to the station. This concept is shown in the third figure.

There are no dock/undock operations associated with a propellant transfer by this means. The vanes can be spun by a coaxial counter-rotating shaft assembly, entering one end of the tank through a gland seal, with the drive motor left accessible for repairs and replacements. You just spin up for the transfer. Otherwise, nothing moves. This is also the least mass to spin up, reducing the energy requirements for spin-up/de-spin, and eliminating any and all thruster operations.

I put the oxygen (LOX) tanks closest to the Vehicle Assembly and Fueling Section, because that is the largest species volume being used, and that shorter length minimizes the power line losses for the motors powering the liquid spin. As the third figure shows, I put the cryogenic fuels hydrogen (LH2) and methane (LCH4) adjacent to the oxygen, because their volumes are also large, to reduce transmission line losses a bit further. I separated the NTO from the MMH with the storable RP1, in order to minimize the possibility of spilled hypergolics coming into contact, even in vacuum. That is a crucial safety consideration!

The truss can be extended further, with additional tanks installed, either for other propellant species, or for additional capacity, or both. This is because there is no other section to the station beyond the Propellant Depot Section.

I think this “spin the propellant inside stationary tanks” concept may be easier to develop and implement than the alternative “each tank is its own syringe” concept, because (1) the rotating-shaft gland seal technology already exists, even for cryogenics, (2) the required piston seal concepts and associated leakage recovery concepts for the “syringe” do not yet exist for cryogenic fluids, and (3) the tank-and-equipment masses and dimensions would be lower: vanes and motor vs a syringe piston and its driving equipment. The hardware has to ride up to LEO inside existing payload spaces, after all!

Conclusions

#1. A combined vehicle assembly facility and propellant depot in LEO could enable all sorts of interplanetary missions very effectively, and even missions to lunar orbit, plus replace the Earth-observation functions that will likely cease for a while when the ISS is decommissioned.

#2. This type of LEO facility is an enabling item to put an effective space tug operation into effect, that uses elliptic departures and elliptic arrivals, to reduce the velocity requirements of interplanetary (and lunar) vehicles.

#3. This kind of fueling operation could use “spin-the-fluid-in-a-stationary-tank” to reduce the overall energy requirements of propellant transfer, eliminate any need for the use of ullage thrusters, and also eliminate many dock/undock operations.

#4. All the other technologies required to build this thing already exist.

References (use date and title in the archive tool on the left, to access quickly):

#1. G. W. Johnson, “Tug-Assisted Arrivals and Departures”, posted to “exrocketman” 1 December 2024.

#2. G. W. Johnson, “Refining Proposed Suit and Habitat Atmospheres”, posted to “exrocketman” 2 January 2022.

#3. G. W. Johnson, “A Concept for an On-Orbit Propellant Depot”, posted to “exrocketman” 1 February 2022.

Figures:

Figure 1 – Concept Sketches For the Vehicle Assembly and Fueling Section

Figure 2 – Concept Sketches For the Power and Propulsion Section

Figure 3 – Concept Sketches For the Propellant Depot Section

Update 5-4-2025:

Conversations with a friend led me to understand that what I have in mind for the vane-equipped tank may not be readily apparent to the reader. Please see the sketch in Figure A below, as you read the following more detailed descriptions.

There are just two sets of vanes inside the tank, mounted on shafting that causes them to counter-rotate. Their tips spin circumferentially, but in opposite directions (which avoids applying gyroscopic forces to the depot station). There is a perforated baffle between the two sets of vanes so that the two volumes of fluid which are affected by the vanes, also rotate circumferentially, independently in each section.

Yet the baffle is perforated, so that the radially-measured levels of the fluid, flung out to the tank wall, are equal in the two sections. It is one shaft, with a gland seal at one tank head, and an internally-mounted bearing at the other. There is a gear box near the middle that makes the shafting turn in opposite directions in the two sections. There are probably 4+ vanes in each section, mounted to the shafts. If you forgo “instant” response, these vane and shaft assemblies can be fairly lightweight construction.

The propellant pickup is along one side of the tank, not one or the other end head, since the centrifugal forces will fling the propellant to the outer cylindrical wall, forming a big hollow cylindrical "form" in each of the two sections, as wetted to the local outer wall. These propellants are moving in opposite directions circumferentially in the two sections, induced to do so by the spinning vanes. But that circumferential motion does not really affect the suction drains along the tank cylindrical wall!

You spin the vanes to withdraw propellants, but you need no spin to pump propellant into the tank. I put the drive motor outside the propellant tank for its safety (remembering the in-tank stirring-fan device that caused the explosion on Apollo-13), and for easy maintenance and repair. Cryogenic gland seal technology already exists, in rocket engine turbopumps. The vane shaft motor and the propellant withdrawal line are on the end that attaches to the core structure of the orbital propellant depot space station. All the power lines and fluid delivery piping is inside that core.

This is a heavier solution than ullage thrusters, so this is definitely only for a propellant depot in orbit (where you do not want to disturb the orbit with ullage thrust), not the vehicles it is supposed to supply with propellants on orbit.

However, per the not-to-scale concept sketch in Figure B below, it might be "just the thing" for the "payload" propellant tanks of a dedicated tanker vehicle sent up to supply this depot station. Those will be rather small compared to the rest of the upper stage delivery craft, in turn small compared to its launch booster. The sketch is not to scale, in order to provide clarity about which tanks are vane-equipped, and which are not. Ultimately, this tanker needs to be a fully-recoverable vehicle.

Figure A – Concept Details for Cryogenic Vane-Ullage Propellant Tank

Figure B – Concept Sketch for Dedicated Tanker Vehicle

Update 5-27-2025:

I extended my investigations into some preliminary design analysis to determine some “typical” design characteristics for spinning-vane tanks of the counter-rotating type. I ran my numbers in metric, since that is what most space industry people are now using.

It is customary to use rotation speed measured in revolutions per minute, rpm. The conversion to or from radians per second looks like this:

ω, rad/s = (N, rpm)*(2*pi rad/rev)/(60 sec/min)

or N, rpm = (ω, rad/s)*(60 sec/min)/(2*pi rad/rev)

The radial acceleration component “a” felt at the radius tip for a rotation speed ω is:

a, m/s² = (R, m)*(ω, rad/s)² or ω, rad/s = [(R, m)/(a, m/s²)]^0.5

You divide that “a” by the standard value of g = 9.80667 m/s² to express that acceleration in “standard gees”, as n = a/g, which is often also customary, and dimensionless.

The distance s that something falls through, under constant acceleration a, depends upon the square of the fall time. Radial acceleration is not constant, but we will ignore that:

s = 0.5*a*t²

The floating spherical globules are everywhere throughout the tank cross section, so the average distance through which a globule must fall is R/2. Substituting that into the previous equation and solving for a fall time, we have:

t = (R/a)^0.5

which is a sort of time constant t_const for the process of all the globules accelerating outward along the radius. (Note that for axial “ullage” acceleration, you would use the tank length L instead of its radius R for a measure of the average fall length s = L/2 in the same equation for time constant.)

The process is quite stochastic, but the usual “rule of thumb” is to wait 3 time constants, and then the process is pretty much completed. Thus:

t_settle ~ 3*t_const

Figure C shows N, rpm versus n, gees, parametric upon values of R from 1 to 5 meters. There is clearly a wide range of possible values. The corresponding settling time estimates versus n, gees, parametric upon R are given in Figure D just below.

Figure C – Results for N, rpm, Versus n, gees, Parametric Upon R, m

Figure D – Results for t_settle, sec, Versus n, gees, Parametric Upon R, m

Amazingly enough, those t_settle values actually formed a tight band of solutions across a fairly wide range of R from 1 to 5 meters! If one selects n = 0.1 gees, the settling times all fall 3 < t_settle < 7 sec! So, the design problem is rather conveniently bounded at about n = 0.1 applied rotational gees, with any conceivable settling times under about 10 sec!

Applying that revelation, I plotted the necessary rotation rates N, rpm, versus the tank radii R, m, for only 0.1 radial gee. Those data are given in Figure E, and indicate a bound on the rotation rate as 4 < N < 10 rpm, for all radii from 1 to 5 meters! That rather conveniently bounds the rotation rates to rather low values. Tip speeds are also closely bounded as about 1 to just over 2 m/s, also low numbers.

Bear in mind that if you wait about 10 sec, you have all the settling times covered! That is a rather short settling time! The smaller value of 3 sec occurs at the smaller R = 1 m, so yet smaller tanks are no problem at all. At this time in history, it would be difficult to imagine successfully transporting to orbit a tank much larger than 10 m in diameter, so we have that end of things fairly well covered, also! And the highest tip speed is less than twice the max flow speed recommended for liquids in pipes. We should be okay there!

Figure E – Required Rotation Rates vs Radius at 0.1 Gee

This Figure E and a nominal settling time of 10 seconds pretty well has most conceivable tank sizes “covered” for designing-in counter-rotating vanes. Such would require no dock and undock operations to transfer the propellants, and would apply no unwanted forces due to spin reactions or gyroscopic resistance forces to whatever structure or vehicle such a tank is mounted upon (precisely because there are counter-rotating sets of vanes). The pump suction points are on the tank lateral wall instead of the ends. So what? The tank itself does NOT spin! So, there is no balance problem, other than the vanes themselves, when you build one of these tanks. You just spin up the vanes, wait about 10 seconds, then start your transfer pump. How simple is that?

Again, such would enable cryogenic propellant storage and transfers from an orbiting propellant depot station, to any vehicle departing from there. Such could also be the “delivered payload” tanks of a dedicated supply tanker vehicle, arriving at that same station. The implications for making deep space missions cheaper with reusability, and cheaper still by means of space-tug-assisted elliptic departure, are simply staggering!

Tuesday, April 1, 2025

About Nuclear Pulse Propulsion

Depending upon the detail method chosen, this kind of revolutionary propulsion could be in flight test within 5 years, and flying in its initial form in 10 years. That would be the 1950’s fission technology. The other versions might perform better, but lack the materials, and the necessary detonation or containment technologies, that are required even to build test devices. They still might not be ready to fly in 50 to 100 years, if fusion is involved, according to some experts.

This article is based upon a NASA paper and two Wikipedia writeups about pulse propulsion, plus George Dyson’s book “Project Orion” about his father’s work on the Orion project at General Atomics in San Diego, CA, in the 1950’s and early 1960’s. That last begins with company R&D work leading up to their first government contract.

Between those four sources, a pretty good picture of the propulsion is available, particularly the technologically-ready fission charge version originally pursued in the 1950’s and early 1960’s. For readers wishing to pursue this further, those references are:

AIAA paper 2000-3856 “Nuclear Pulse Propulsion - Orion and Beyond”, by G. R. Schmidt, J. A. Bonometti, and P. J. Morton, then from the NASA Marshall Space Flight Center, Huntsville, AL.

Wikipedia article “Nuclear Pulse Propulsion”, as retrieved 2-26-2025, and last edited January 2025.

Wikipedia article “Project Orion (nuclear propulsion), as retrieved 2-26-2025, and last edited February 2025.

George Dyson, “Project Orion – the True Story of the Atomic Spaceship”, published 2002 by Henry Holt and Company, New York City. (A second edition is forthcoming very soon, if not already available.)

The basic external detonation concept is depicted in the Figure 1 sketch below. All figures are at the end of this article.

The idea is to explode a fission bomb at a safe distance behind the vehicle, which vaporizes a reaction mass, and blows that reaction mass into the pusher plate at the rear of the vehicle. This requires a sort of “shaped charge” technology for the fission device, in order to increase the amount of reaction mass intercepted by the pusher plate. There are shock absorbers between the pusher plate and the rest of the vehicle, to smooth-out the high-gee “hits” into a nearly-continuous and almost-steady acceleration.

Figure 2 gives some indication of the effects of vehicle size on this process. Bigger mass is inherently a larger pusher plate dimension, which then intercepts a larger fraction of the plasma blast created by vaporizing the reaction mass. The bigger mass also reduces the high-gee “hit” from the explosion, making the shock absorber design easier.

At the time this technique was pursued in the 1950’s and early 1960’s, it was thought that the main environmental concern would be the radiation fallout in the atmosphere from a surface launch. Being fractional-kiloton (KT) devices, leading to low-KT devices once out of the atmosphere, there is less fallout than one might otherwise suspect, more or less comparable to one atmospheric test of a low-range megaton (MT) thermonuclear device. See Figure 3.

The risk of electromagnetic pulse (EMP) effects was not really recognized until after the “Starfish Prime” megaton-range nuclear test, that was conducted in space in 1962. Some data about that are given in Figure 4. This really restricts where and how one might surface-launch such a vehicle. These vary inversely with the square of the distance, and directly with yield. Bear in mind that we simply do not have the necessary technological capabilities yet, to build such vehicles out in space, so surface launch is still the only means available in the short term.

Based on the data in the four references, a rough approximation to the performance values one might expect are given in Figure 5. These are quite remarkably high, sufficient for sending crews pretty much anywhere within the solar system, on relatively short trips. Some of the fusion concepts, if they can really be made to work, might even be suitable for interstellar missions.

It is this author’s opinion that we need to get over our fears about “nuclear things in near-Earth space”, modify the Space Treaty to allow this kind of propulsion, and simply “get on with the war”. However, finding massive new fissionable material resources is inherent to support the quantities of fissionable material that will be necessary. If not on Earth, then “out there” somewhere.

As a conceptual example, the one-way velocity requirement to go from low Earth orbit to rendezvous with asteroid 2024YR4 is about 5.9 km/s as depicted in Figure 6. Back to low Earth orbit, about 4 years later, the same velocity requirement can decelerate one back to low orbit, excluding all rendezvous and course correction requirements. If one wanted to send a large expedition to explore mining this asteroid (to include bringing significant mass home), using a pulse propulsion vehicle already based in low orbit, the rough sizing of such a vehicle’s weight statement might look like the numbers given in Figure 7. The contrast with a chemically-powered vehicle is quite stark. But with pulse propulsion, the basic message is to build it big!