kbd512 Postings

tahanson43206 · 2025-06-06 08:00:50

For SpaceNut re lubrication of bearings ....

A possible solution to the very real problem you have identified is to allow for the rotating habitat batons to turn with only magnets keeping the walls of the rotating components away from each other. This could be done during flight when there are no forces acting to move the fuselage. When it becomes necessary to apply force for navigation, the rotating components could be brought back together so they can accept acceleration loads.

I've never heard of such a system, and there may be practical reasons why it would not work.

I'm pretty sure magnetic bearings do exist, but the machines involved might be small.

(th)

kbd512 · 2025-06-06 12:30:07

tahanson43206,

Using rings of permanent magnets in a Halbach array could lock all rotating and "stationary" parts of the vehicle in place without physical contact. Permanent magnets are typically heavier than electromagnets for the field strength they provide, but they do not require any input electrical power. For the strength of the field required for this application, stronger fields are likely irrelevant. Beyond that, we really don't want the magnets to be "turned off" in this application. The vehicle will function best if the magnets are "on" at all times, and the various parts "float" perhaps a centimeter away from each other.

The label "air" would be replaced with "vacuum" in space:

A planar permanent magnet array:
H08b512fb570a43e481230419bf791198H.png_720x720q50.png

Axial flux electric motor:
EVR%20Chart%20%20(4).jpg

Halbach arrays can also be incorporated into a "curved" inner and outer lip type of geometry, so that the magnetically suspended habitation modules cannot readily "fly apart", even if the magnets themselves fail.

kbd512 · 2025-06-06 13:20:38

tahanson43206,

A much cheaper option would be using full ceramic bearings with minimal Silicone-based lubricant. At 6rpm, the duty cycle is 3,153,600 cycles per year.

A drastically cheaper option would be using standard Babbitt metal bearings with lubricants. Much hay is made over this, but the battery operated tools used by the astronauts to repair the Hubble Space Telescope, such as the bit driver they used to unscrew and re-torque bolts, did in fact use bearing lubricants. The photovoltaic arrays on the ISS do in fact use sealed metal bearings with a vacuum-compatible grease.

That means there are 3 ways to solve this centering and bearing problem:
1. Contactless permanent magnets or electromagnets.
2. Dry or minimally lubricated ceramic bearings.
3. Lubricated metal bearings.

Option 1 is scientifically interesting because it will theoretically never fail from direct contact friction.
Option 2 can work, but is also expensive at the scale required and will still require periodic replacement.
Option 3 will work for many years before replacement and is exhaustively well proven on Earth and in orbit aboard ISS.

tahanson43206 · 2025-06-13 08:51:08

For kdb512 re #278...

There is an additional option that might be worth adding to your list.

A hybrid system would use physical bearings when the habitats are undergoing acceleration by the fuselage, and magnetic bearings during free flight.

This post is to note your contributions to the RDE topic, and to ask what I hope will be a useful question for a future post.

My understanding (so far) of RDE is that it is more efficient in mixing gases for propulsion, and I've picked up the impression that RDE delivers better performance than alternatives.

My impression from reading your posts and other sources, is that a significant benefit of RDE is reduced mass, possibly due to elimination of turbo pumps but I'm not clear on that point.

Aside from those issues, I'm wondering if RDE provides better ISP for a given propellant mixture?

If if does, what is the mechanism that provides that improvement?

(th)

GW Johnson · 2025-06-13 09:02:47

The shock wave moves circumferentially around and around. Flow ahead of it is supersonic, but flow behind it is subsonic, because it is a normal shock. It has to be. So this technology is inapplicable for anything claimed to be "supersonic combustion". The confusion factor with terminology here is with "supersonic" ahead of the wave. That is NOT due to the flow speed coming into this thing from the inlet, it is due to the speed of the circulating wave itself, which must be inherently supersonic in its circumferential speed.

Shock wave-induced combustion is a pressure-gain kind of combustion that we call "detonation", as opposed to merely explosively-fast "deflagration" that occurs with a slight pressure loss. Those are the only two kinds, and achieving stable and controlled detonation combustion has been a very difficult thing to do. For the same pressure gain from the air inlet hardware, the detonation-wave combustion will result in a higher-pressure subsonic chamber condition behind the circulating wave, than could possibly be produced by deflagration in a more conventional combustor scheme. That higher pressure results in a different nozzle throat size, too.

Higher chamber pressure is simply higher thrust potential, which for the same fueling rate is higher specific impulse (if a rocket). That definition of thrust and specific impulse gets a lot more complicated if you do this as an airbreather. But the nozzle thrust term in airbreather thrust really is bigger.

GW

Last edited by GW Johnson (2025-06-13 09:06:06)

kbd512 · 2025-06-14 15:35:10

tahanson43206,

With respect to your question about RDE Isp, I think GW already explained how detonation vs deflagration delivers more thrust for the same propellant mass flow rate.

Expander cycle conventional deflagration-based chemical rocket engines use waste heat from the engine nozzle and combustion chamber to drive the pumps, up to output levels of about 300kN or so, which implies 600kN output levels for a RDE. For RDEs with thrust levels much beyond 600kN, my assumption is that you still require turbopumps, but those pumps will only need to deliver half of the work output which would otherwise be required to feed propellant into conventional deflagration engines.

For a "main engine" using rotating detonation, my assumption is that the engine must still be pump-fed, but the shaft power consumed by the oxidizer and fuel pumps is approximately half that of a thrust-equivalent deflagration-based engine. Let us assert that a particular deflagration-based rocket engine requires 10MW of shaft power. The equivalent RDE should require about 5MW. One would therefore expect that the pump mass for the RDE is about half that of the deflagration engine. This assumption may not be 100%-true-to-reality accurate, but it will be a reasonably close approximation.

As far as air-breathing RDEs are concerned, I think Dresser-Rand's "rampressor" (supersonic ram-air compressor), which was originally designed to compress CO2 exhaust gas from coal-fired power plants for storage, can deliver compressed air to the engine with a supersonic inlet velocity. Unlike a conventional gas turbine engine, there's no need to slow the incoming airstream to subsonic velocities before feeding it into the compressor stages of the gas turbine. The rampressor uses a special inlet ramp geometry to permit shockwave-assisted gas compression, rather than massive wave drag, which is what you get when the tips of prototypical "fan blades" used by conventional gas turbine engine compressor stages approach or exceed the speed of sound. Conventional blade-based compressor stages could deliver 2:1 CO2 compression ratio per stage. The rampressor designed by Dresser-Rand delivered 10:1 compression per stage, with the ability to recuperate some of the waste heat to reduce the power required to drive the compressor.

The thrust that a rampressor based turbojet engine can deliver is roughly double that of a conventional turbojet. Another way of saying that is a gas turbine engine using rampressor technology to feed air into the engine has double the thrust-to-weight ratio of a conventional blade-based subsonic air compression gas turbine engine, because it needs fewer compressor stages and is far more compact than a conventional turbojet engine as a result.

A "RamGen" Engine with Dual Supersonic Inlet and Dual Supersonic Expander Stages:

The air going into the engine inlet is supersonic, combustion happens at supersonic velocities, and gas expansion happens at supersonic velocities. This would destroy or severely damage a conventional turbojet engine. However, the "rampressors" are deliberately designed to compress and expand supersonic air. The supercomputer-designed inlet ramp geometry uses the "trapped" shockwave to actually assist with compressing the incoming air stream, which is how the test models achieved 10:1 to 12:1 compression per stage, vs the 2:1 compression achievable with "normal" fan blades. A conventional turbojet or turbofan engine would use complex inlet geometry to trap the shockwaves from the supersonic flow and also slow the incoming flow to subsonic speeds, loosing kinetic energy and generating heating in the process.

The most advanced models of GE's GEn-9X turbofan (an advanced GE-90 derivative engine design) that power jets like the Boeing 787 use, IIRC, 14 stages of compression, and achieve an overall pressure ratio (OPR) of about 58:1 during certain phases of flight, with 42:1 at takeoff. OPR is a figure of merit that describes the overall efficiency of a jet engine. A 2 stage Rampressor achieves a 100:1 OPR.

In gas turbine engines, the overall pressure ratio (OPR) is a crucial parameter that significantly impacts engine performance, efficiency, and thrust output. It represents the ratio of the total pressure at the exit of the compressor to the total pressure at the inlet. A higher OPR generally leads to increased thermal efficiency, more thrust, and improved fuel consumption.

A Review of Shock Wave Compression Rotary Engine Projects, Investigations and Prospects

British Aerospace HOTOL:

North American-Rockwell's STS orbiter delivered 12,500kg to a 400km (ISS-like) circular orbit.
BAE's HOTOL was intended to deliver 7,000kg to 8,000kg to a 300km circular orbit.

HOTOL would carry about 205,000kg of LOX/LH2 propellant, with the fuselage containing 48,000kg (677.488m^3) of LH2 and 157,000kg (137.599m^3) of LOX. Isp was said to be about 700s overall, and around 2,000s at sea level when consuming atmospheric Oxygen to power the engines. Total burn time to achieve orbit was to be about 730s. HOTOL would transition to using pure rocket propulsion between Mach 5 and Mach 6, which correlates well with prototypical booster burnout velocity for a TSTO.

I envision a SSTO somewhat similar to BAE's HOTOL concept, meaning a HTHL vehicle that takes off and lands on a conventional runway, but powered by rampressors to do away with those large and heavy variable geometry inlets, ultimately force-feeding atmospheric Oxygen into RDEs. Rather than including all the machinery to convert atmospheric air into a liquid, we instead use hot high pressure gas from the rampressors, and use some of the waste heat to convert RP1 into a vapor to feed into the RDEs. RP1 was selected over LH2 to both reduce the volume of the vehicle for a given payload mass, and to avoid the complexities of handling LH2 on the ground. The vehicle will be heavier as a result, but we can get more payload to orbit using higher thrust from RP1 and greatly reduced turbopump size / mass / volume.

This vehicle would use advanced materials for structure and thermal protection that simply did not exist when the HOTOL concept was being developed in the 1980s, along with and an advanced combined cycle propulsion system that uses RamGen and RDE engine tech that also did not exist.

The purpose of this vehicle, once again, is purely passenger transport to orbit to either a space station or awaiting interplanetary transport vehicle. We already have Starship, New Glenn, and perhaps Vulcan to do the heavy lifting. If we apply pure RDE power, meaning no air-breathing, and the same materials tech to a VTVL TSTO, it's still likely to deliver more payload for less money. Even our miraculously powerful new engines and advanced materials cannot compensate for discarding 2/3rds of a vehicle's dry mass on its way to orbit.

Something similar to HOTOL will deliver 1t of payload per 31.25t of vehicle and propellant mass. I'm sure we can improve on that using modern materials, but only to a point. HOTOL would've used bleeding edge engine and materials tech when it was developed. The only real night-and-day performance improvement is related to the engines.

Starship 3 will deliver 1t per 25t of vehicle and propellant mass with full reusability, or 1t per 12.5t to 16.7t fully expended. No SSTO is ever going to compete with an expendable TSTO on payload performance, and it's pointless to try. We are never going to beat basic flight physics, even if we manage to devise a SSTO that is only marginally worse than TSTO.

The only real "game changer" that might be possible in another 30 years or so is using fusion to superheat a monopropellant like LH2 or LCH4. That kind of SSTO may make using TSTOs for heavy cargo delivery a thing of the past, but only because fusion rather than chemical energy supplies the heat input, and it will probably need to be aneutronic fusion to minimize shielding requirements, which implies He3, which implies lunar mining. That makes fusion-powered SSTOs a far future technology that requires greenfield development work, and lots of it.

Until then, we really need to have mass-competitive air-breathing SSTOs for minimal risk passenger delivery to orbit, to allow airline transport services to take shape. We'll continue to use advanced TSTOs for heavy cargo until something like aneutronic fusion heat engines (~7,000s of Isp with H2) makes chemically powered SSTOs and TSTOs superfluous. On top of that, we really must have a propellantless launcher, either a rotary slingshot type machine or electromagnetic gun, for cost-effective delivery of bulk delivery of metals and liquids.

kbd512 · Yesterday 02:50:05

For clarity, in the first illustration above you will see a different variant of the HOTOL vehicle with 995m^3 of tank volume allocated to LH2 (70,496kg at most), plus 114m^3 (130,074kg) of LOX tank capacity, and possibly up to 140m^3 of LOX, although the actual capacity of the aft LOX tank near the engine cluster is not provided. LH2 fuel mass for HOTOL is listed as 48t, but there were multiple variants of the vehicle and I don't have definitive data to know which propellant capacities applied to which specific variant. What ought to be readily apparent is how much internal vehicle volume LH2 fuel requires.

Engine Type Specific Impulse with Hydrogen vs Hydrocarbon Fuels:

University of Texas at Arlington - Aerodynamics Research Center - Pulsed Detonation Engines

PDE = Pulsed Detonation Engine (RDE is a type of PDE)

Presumably, what the chart actually means is that a RDE being fed atmospheric Oxygen can achieve that kind of Isp, because pure rocket propulsion Isp is shown at the bottom of the chart, and ranges between about 360s (RP1) and 470s (H2). Detonation engines don't achieve substantially greater Isp (thrust generating efficiency) than conventional deflagration-based chemical rocket engines when all of the oxidizer and fuel must be carried onboard the vehicle. I think the term applicable here is "Effective Isp". When the atmosphere can be used as reaction mass and to provide the oxidizer consumed to generate thrust, the Effective Isp for turbojets, turbofans, and pulsed engines greatly exceeds that of pure rocket engines, to include PDEs / RDEs.

Anyway...

In my proposed SSTO vehicle design using vertical launch and pure rocket propulsion, propellant mass was 2,185,289kg to deliver about 100t to orbit.

SSTO Engine Technology Post #29

1,875.496m^3 of LOX/RP1 propellant was required.

SSTO Propellant Masses and Volumes for a STS Orbiter Equivalent Payload to Orbit:
LOX: 1,597,846kg (1,192.422m^3)
RP1: 587,443kg (683.074m^3)
Total: 2,185,289kg (1,875.496m^3)

Saturn V S-IC Stage Propellant Masses and Volumes:
LOX: 1,444,506kg (1,266m^3)
RP1: 647,800kg (790m^3)
Total: 2,092,306kg (2,056m^3)

This was intended to deliver the same total dry mass to orbit as the historical Space Shuttle, inclusive of its nominal max payload. Effective Isp over the entire flight regime (sea level to orbit) was 90% of the RD-180's Vacuum Isp. Engine mass was derived from Raptor 3's 200:1 engine TWR and from the requirement for a 1.5:1 vehicle TWR at ignition, in accordance with GW's admonishment to ensure that liftoff thrust is sufficient to get the vehicle "smartly moving downrange". This was presumed to be feasible on the basis of Raptor 3's 200:1 TWR while burning LCH4 fuel, which is only about half as dense as RP1, and consequently requires larger and heavier turbopumps than a RP1 burning Raptor 3 thrust-equivalent. Essentially, if a 200:1 engine TWR is achievable with LCH4, then it's also achievable with RP1. If anything, the engine should get lighter and provide a higher TWR because the fuel pump requires less mass / volume / input shaft power to deliver the same flow volume to the combustion chamber. There's nothing too outlandish about these assumptions.

When we talk about a pure rocket powered SSTO designed to deliver 500 passengers to orbit, we're discussing a vehicle with roughly double the Space Shuttle orbiter's total internal volume, on-par with the S-IC stage of the Saturn V, but using much lighter and stronger structural materials, modern thermal protection materials, and thrust equivalent to the 5X F-1 engines attached to the S-IC stage, albeit using much lighter / more powerful engines. If we're able to substitute the onboard oxidizer mass and volume by using air-breathing engines up to normal booster burnout velocity, then our vehicle's total internal volume becomes more similar to that of the Space Shuttle orbiter or HOTOL design, and we can say with a fair degree of confidence that such a vehicle will be able to attain orbit and return to a runway on Earth, just as the Space Shuttle did.

Now back to the "big idea" behind using RDEs:

Thrust-equivalent RDE should require half as much input shaft power, thus much smaller / lighter pumps, and consequently, a bump in engine TWR from 200:1 to 400:1 ought to be possible using the same materials (high grade stainless steels and Nickel-Copper alloys. RCC is only 2g/cm^3 vs 8g/cm^3 for stainless and Nickel-Copper alloys, so a 800:1 to 1,600:1 engine TWR becomes feasible. Since NASA has already tested scale RDEs and RCC engine components (the combustion chamber and nozzle), I think the material and engine tech are "ready to apply", even though more development is still required. RCC turbopump components were fabricated and tested for Project Timberwind, many moons ago. It's not outlandish to think that all the major components can be fabricated from RCC and that regenerative cooling is largely optional. We may still choose to extract heat from the combustion chamber and nozzle to reduce pumping power, not because the engine would melt if we did not. At 800:1 TWR or higher, even when Saturn V thrust levels are demanded, engine mass becomes a minor fraction of total vehicle mass, so propellant tank mass and TPS mass will become the focus of development work.

All together now, we're using supersonic air compressors to extract atmospheric Oxygen and atmospheric Nitrogen to use as reaction mass, possibly driven by sCO2 gas turbines that transfer heat from the compressed air into the fuel, to force feed gobs of air and vaporized fuel into rotating detonation wave engines, so that our sea level to orbit propulsion system can generate the thrust required for a mass that a SSTO vehicle can tolerate.

kbd512 · Yesterday 23:14:30

Buran, Space Shuttle, HOTOL Size Comparison:

Gross Liftoff Mass; Orbiter Vehicle Mass; Max Payload Capability
Buran: 2,524,000kg; 75,000kg; 30,000kg (presumably 251km circular orbit)
Space Shuttle: 2,029,633kg; 78,000kg; 24,400kg (204km circular orbit) or 12,500kg (407km circular orbit)
HOTOL: 250,000kg; 50,000kg; 8,000kg (300km circular orbit)

HOTOL LH2 vs RP1 Mass and Volume Comparison for Equal Stored Energy
1. HOTOL had sufficient internal volume to store 995m^3 (70,495.75kg; 10,080,892.25MJ) of LH2 fuel, at most, ignoring the fact that some void space must be present to vent boil-off.
2. 10,080,892.25MJ / 43MJ/kg = 234,439.35kg (285.902m^3) of RP1
3. HOTOL also had up to 140m^3 (159,740kg) of LOX onboard, at most, which implies that 26,623.33kg (3,807,136.67MJ) of LH2 was consumed using pure rocket propulsion (presuming the typical 6:1 O/F ratio). 88,538.06kg of RP1 provides equivalent energy, which implies that 230,198.96kg (201.752m^3) of LOX (presuming the typical 2.6:1 O/F ratio for RP1 fueled engines) must be carried aboard an RP1 powered HOTOL variant for pure rocket propulsion.
4. That means total onboard propellant volume for a RP1 powered HOTOL equivalent is 487.654m^3, so 230,198.96kg LOX + 234,439.35kg RP1 = 464,638.31kg of total propellant mass.
5. The LOX/RP1 propellant mass is slightly more than double that of HOTOL's stated LOX/LH2 propellant load, but the total propellant volume in the vehicle is less than half of the LH2 volume alone. The TWR of RP1 fueled engines is more than double that of LH2, so engine mass would be significantly reduced, which directly translates into greater payload mass since HOTOL is a SSTO.

The most obvious "better option" associated with choosing RP1 power is the ability to carry significantly more useful payload mass to orbit in a vehicle with the same internal volume as an equivalently-sized LH2 powered vehicle.

kbd512 · Today 08:38:53

I watched a YouTube video last night on how the RB545 engine actually works:
HOTOL - Anatomy of a spaceplane engine

Apparently, HOTOL's RB545 engine would dump 1/2 to 2/3 of its entire LH2 load directly overboard (never combusted in any engine) after using it as a heat sink to help "almost liquefy" the incoming airstream to reduce the compression work required to supply atmospheric air to the engine for the first phase of the flight from takeoff to Mach 5 or so. That aptly explains how a vehicle with an average Isp of over 700s needed so much LH2 onboard.

Using a rampressor, sCO2 gas turbine, and a heat sink, we can dramatically improve upon that without dumping any fuel overboard.

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