Rotating Detonation Engine

Void · 2023-02-04 08:50:31

Marcus House had a video out that mentioned this engine. Possibly as much as 25% more efficient, it is said.

https://en.wikipedia.org/wiki/Rotating_ … ion_engine

https://www.youtube.com/watch?v=UutHG8Y2UuQ

Done

Last edited by Void (2023-02-04 08:54:36)

tahanson43206 · 2023-02-04 10:39:17

For Void re New Topic

Index to posts that may be contributed by NewMars members over time:

Index:
kbd512: https://newmars.com/forums/viewtopic.ph … 55#p232155
Explanation of rocket engine operation and review of RDE advantages and possible uses

kbd512: https://newmars.com/forums/viewtopic.ph … 96#p232196
Comparison of existing SpaceX Super Heavy to possible RDE configuration with composite fuel tanks

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Original contents:

Thanks for providing the Wikipedia link, and the YouTube video....

The YouTube video includes quite a bit of space coverage before the Rotating Detonation rocket engine coverage, but ** that ** coverage is worth the wait!

There is animation ** and ** (what appears to be) a slow motion camera recording of the rotation.

If anyone can find out more about the mechanical systems needed to allow this (apparently more efficient) combustion process to work, I would be interested.

By way of speculation/question .... does this ignition/combusion method make any difference in the perennial quest for SSTO?

A bug-a-boo of SSTO efforts until now has (I gather) been the need to slow gases to be burned to subsonic speeds.

(th)

RobertDyck · 2023-02-04 12:16:27

tahanson43206 wrote:

A bug-a-boo of SSTO efforts until now has (I gather) been the need to slow gases to be burned to subsonic speeds.

Any rocket engine works by mixing fuel and oxidizer in a combustion chamber. The throat of the engine is the transition from combustion chamber to nozzle. The de Laval nozzle was originally developed in the 19th century by Gustaf de Laval for use in steam turbines. The throat of such a nozzle causes gas flow to transition from subsonic to supersonic. Supersonic flow behaves quite differently than subsonic, so the transition is very important. The transition ensures high pressure in the combustion chamber while gas flows continuously out the nozzle. Rocket engines produce thrust using Newton's law of motion: accelerating a mass aft cause the engine to accelerate forward; momentum of reaction mass accelerated aft is exactly equal to momentum of engine propelled forward. And the rest of the rocket is attached to the engine. So if you accelerate reaction mass 'm' from zero to 'v' velocity, then momentum is m*v. So if a rocket engine expels 'm' reaction mass per time period from zero to 'v', then the engine will produce m*v thrust per the same time period. The rocket including fuel tank, oxidizer tank, plumbing, engine, fuel and oxidizer in those tanks, and payload will total much more mass than reaction mass expelled by the engine. So change in velocity of the rocket is less than exhaust velocity. Also note total mass of the rocket will decrease as fuel and oxidizer are consumed, making the math more complicated.

Example: a Falcon 9 rocket carries 17,400 kg of payload. The upper stage has a dry mass of 4,000 kg, and carries 75,200 kg of LOX and 32,300 kg of RP1. Total mass when the engine ignites is 128,900 kg. Note the fairing is discarded before upper stage engine ignition. That stage has one engine, thrust 934 kN. One Newton of force is 1 kg*m/s² (one kilogram metre per second squared). 1 kN = 1,000 Newtons so at engine ignition, acceleration is 934,000 / 128,900 = 7.246 m/s². If you want to know how thrust in Newtons is calculated, it's reaction mass expelled by the engine per second multiplied by exhaust velocity in metres per second.

(In the United States "meter" is spelled that way. In the rest of the English speaking world it's spelled "metre".)

Again, as fuel and oxidizer is consumed, mass of the upper stage will decrease. Thrust by the engine will remain constant, so acceleration will increase as propellant (reaction mass, fuel + oxidizer) is consumed.

For a liquid fuel chemical rocket, it's most efficient if fuel and oxidizer fully combust. You don't want unburned fuel expelled out the rocket nozzle. Unburned fuel is still mass out the nozzle, but it doesn't produce energy. Burning fuel with oxidizer undergoes a transition from liquid to gas, which results in gas expansion. More importantly, burning fuel releases energy. That energy is in the form of heat, and gas expands when it's hot. That gas expansion produces pressure in the combustion chamber. More pressure results in greater exhaust velocity. Greater exhaust velocity produces more momentum from the same exhaust mass (reaction mass), so more efficient engine. This means you want complete combustion. All rocket engines have subsonic gas flow in the combustion chamber, but the closer the gas is to the throat, the faster it moves. Some fuel could be expelled unburned. Rotating detonation is an attempt to ensure complete combustion.

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Slowing gas to subsonic speed is something modern jet engines do. A RAM jet has a shaped inlet that slows air flow, causing compression. Fuel is injected into the combustion chamber and ignited. Transition from liquid to gas causes combustion products to expand. Heat causes air and combustion products to expand much more. Then the RAM jet has a nozzle similar to a rocket engine nozzle which causes transition from subsonic to supersonic. This works best when inlet air is supersonic before it's slowed, so the transition from supersonic to subsonic prevents pressure in the combustion chamber from simply escaping out the front. A turbojet engine has a compressor in front to achieve gas compression, and prevent pressure from escaping out the front. A turbine in the exhaust nozzle captures some of the energy of the pressure to drive the compressor. Gas flow out the compressor in the back have less resistance to than forward through the compressor. So both RAM jet and turbojet engines use subsonic combustion.

There is a new type of jet engine called Supersonic Combustion RAM jet, or SCRAM jet. This does not slow air flow to subsonic speed, instead air flows through the combustion chamber at supersonic speed. It's still slower than intake air, but not slowed below the speed of sound. This allows an aircraft to fly at very high speed, expected to be faster than 5 times the speed of sound (mach 5). Faster than mach 5 is called hypersonic. Theoretically a RAM jet could fly up to mach 6, but faster than that is difficult because slowing intake air to subsonic speed creates so much drag. And exhaust velocity must be faster than the aircraft for the engine to provide any thrust at all. So a SCRAM jet engine is an attempt to fly faster than mach 5 or mach 6. There are a couple problems with a SCRAM jet. First how to you ensure gas flows the correct way without subsonic/supersonic transition? Second and most importantly, how do you achieve complete combustion? With supersonic combustion, air flowing through the combustion chamber of a SCRAM jet does not remain in the combustion chamber vary long. And after combustion, heat must flow from exhaust gas to air, allowing that heat to cause air expansion. That expansion creates pressure, which provides thrust. The air/fuel/exhaust mixture in a SCRAM jet engine doesn't remain in the engine very long. One solution is to make the engine physically longer. That increases engine mass.

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My point is slowing gasses to be burned to subsonic speed is not strictly necessary. It's... complicated.

kbd512 · 2023-02-04 15:05:57

We now have supersonic CO2 compressors that accept an inlet stream of gas, like a turbojet, at supersonic speeds, up to Mach 3 or so, and then a single stage, using the shock waves to assist with compression of the gas rather than generating massive wave drag, achieving 10:1 compression of the gas in a single compression stage. The resultant compressor stage of a supersonic compressor more closely resembles a drill bit's cross-section than a traditional row of fan blades we would expect to see inside a subsonic gas turbine compressor stage / section, which is basically all of them, to include the engines inside the SR-71 and XB-70. Figuring out the "ramp geometry" to make this work required a lot of supercomputer time, which didn't exist before the early 2000s. The end result of that development program was each stage becoming much more compact and performant than traditional blade-based compressors, so fewer compression stages were required, thus jet engines can roughly double their thrust-to-weight ratio while being much shorter in overall length for a given compression ratio.

This new type of turbojet engine was known as "RamGen":
Ramgen Technologies

It's essentially a "ramjet" packaged within a gas turbine engine.

They produced a 2-stage 100:1 compressor, which included supersonic "rampressor" (supersonic compression stage) and "ramexpander" (supersonic expansion stage).

Ramgen Power Systems - Workshop on Future Large CO2 Compression Systems

Since I know few of you actual read through the links posted, read this from Page #18 the link immediately above:

100:1 CO2 compressor --> 2-casings/2-stages/Intercooled
No aero Mach# limit
10+:1 pressure ratio; 400°F temperature rise
1400 fps tip speeds; Shrouded rotor design

I would think that conventional jet engines, capable of operating up to Mach 6, with double the thrust-to-weight of existing conventional designs, would allow us to design a SSTO with the "booster stage" being LCH4 rampressor coolant / fuel, plus O2 from the atmosphere using the rampressor or United Kingdom's SABRE O2 liquefaction if Hydrogen is the only fuel used, followed by a full-flow staged combustion rocket engine optimized for high-altitude / vacuum operation. The Hydrogen-fueled concept is scientifically interesting, but a "Phase I" design should combine RamGen turbojet engines with a few Raptor engines.

This would enable routine takeoffs from conventional airport runways, high speed dash / ascent to 1st stage booster burnout velocities, followed by rocket-powered ascent to orbit using engines optimized for the pressures at high altitudes / vacuum of space. Starship could then be assembled on-orbit and optimized for use exclusively as a lander for lunar / Mars operations.

kbd512 · 2023-02-04 15:53:14

Here's how I see development of a rational SSTO "spaceline transport system" playing out:

1. Develop the RamGen engine technology to the point where we have engines capable of providing the thrust required to depart an airport runway with full fuel to achieve both "booster stage burnout velocity / altitude" using ramjet / supersonic compressor / expander gas turbine engines. This includes sharp leading edge ultra-high temperature ceramics and stainless or super alloy airframes using minimal surface insulation applied for reentry TUFROC / TUFFI durable ceramic insulation using mechanical fasteners to connect all heat shielding elements to the vehicle's airframe. Fuel will be LCH4, with O2 from the atmosphere. The "upper stage thrust" will be provided by Raptor engines with nozzles optimized for exo-atmospheric operation.

2. Continue to develop rotating detonation wave rocket engines (RDEs) to the point that they're well-understood and reliable. Replace the Raptors with the new RDEs to evaluate performance. Re-qualify the vehicle after the RDEs have proven themselves reliable and durable.

3. Evaluate the operational costs of the new "Spaceliner" over several years of operation to determine which components are costing the most money, therefore which direction to take development efforts. At the same time, LH2-fueled SABRE development, with RamGen improvements, should be the focus of development funding.

4. Requalify several airframes to use LH2 fuel and rampressor interstage coolant. This shouldn't be too difficult, because using LH2 as a coolant was one of the RamGen development items that was completed about 10 years ago.

5. Fly the LCH4-fueled and LH2-fueled airframes side-by-side for several years to determine which combination of fuel / coolant provides the most payload performance for the least cost. LH2 will probably win the payload performance argument decisively, but operational cost remains an open question. The 1st gen (Saturn V) and 2nd gen (Space Shuttle) engines and fuels programs produced robust payload performance improvements, but cost went up dramatically. Can we "do better" in the cost department using modern tech? That's an open question that can only be properly evaluated using side-by-side comparison of substantially similar airframes, both optimized to use a given fuel technology, but using the same basic airframe design.

Anyway, that's how I believe we should proceed with a practical SSTO development program that "puts the final nail in the coffin" of the fuel and engines arguments about what is the most performant / most practical / most cost-effective way to send people into orbit, without subjecting them to the dangers and stresses of a high g-load multi-stage rocket launch.

Rocket landings on the moon and Mars appear to be the only feasible way to reliably land there since no runways exist to enable "rolling touchdowns". Even on the moon, which has no atmosphere to speak of, you could still achieve a rolling touchdown if you had a runway, and then stop the aircraft / spacecraft on the ground where normal brakes or arresting gear works (even on the moon). On Mars, you do have some atmosphere to work with to cushion landings, even though landing speeds will be pretty high relative to Earth. Concorde would land at 187mph and takeoff at 250mph, so this is perfectly doable as part of routine passenger transport operations. A 747 landing without flaps would also touch down at speeds near 250mph.

RGClark · 2023-02-17 08:25:26

kbd512 wrote:

…
I would think that conventional jet engines, capable of operating up to Mach 6, with double the thrust-to-weight of existing conventional designs, would allow us to design a SSTO with the "booster stage" being LCH4 rampressor coolant / fuel, plus O2 from the atmosphere using the rampressor or United Kingdom's SABRE O2 liquefaction if Hydrogen is the only fuel used, followed by a full-flow staged combustion rocket engine optimized for high-altitude / vacuum operation.
…

You had me at SSTO.

Having an integrated jet-engine with the rocket engine has long been proposed to make a SSTO. The problem has been the jet engine part has been too heavy. This proposal doubles the the T/W of jet engines. That would put it at about 20 to 1. I would like the T/W doubled again to ca. 40 to 1 to be confident it can make a feasible SSTO with high payload.

Robert Clark

Last edited by RGClark (2023-02-17 08:26:45)

tahanson43206 · 2023-12-22 21:01:05

It's been a while since we've had an update on the Detonation Engine ... here is a high level (30,000 foot) glimpse....

https://www.yahoo.com/news/nasas-detona … 00213.html

Gizmodo
NASA's Detonation Engine Revs Up for 4 Minutes in Breakthrough Test
Passant Rabie
Fri, December 22, 2023 at 9:30 AM EST·1 min read
22
The Rotating Detonation Rocket Engine combustor during a 251-second hot fire test in fall 2023.
NASA just put its new propulsion system to the test, powering a 3D-printed rotating detonation rocket engine for a sustained burn that lasted three times as long as the first test.
The Rotating Detonation Rocket Engine, or RDRE, produced more than 5,800 pounds of thrust for a total of 251 seconds (a little longer than four minutes) during a recent test at NASA’s Marshall Space Flight Center in Huntsville, Alabama, the space agency announced this week.
RDRE was tested for the first time in 2022, producing more than 4,000 pounds of thrust for nearly a minute. The latest test was designed to “better understand how to scale the combustor to different thrust classes, supporting engine systems of all types and maximizing the variety of missions it could serve, from landers to upper stage engines to supersonic retropropulsion, a deceleration technique that could land larger payloads – or even humans – on the surface of Mars,” NASA wrote.
NASA engineers are currently working to figure out how to scale the technology for higher performance, hoping to develop a fully reusable 10,000-pound (4,500-kilogram) RDRE.
“The RDRE enables a huge leap in design efficiency,” Teasley said. “It demonstrates we are closer to making lightweight propulsion systems that will allow us to send more mass and payload further into deep space, a critical component to NASA’s Moon to Mars vision.”
For more spaceflight in your life, follow us on X (formerly Twitter) and bookmark Gizmodo’s dedicated Spaceflight page.

The report is dated December but it appears to be covering a test in fall of 2023.

(th)

tahanson43206 · 2025-05-16 17:15:08

Our last update was in late 2023. Here is a report on what appears to have been a successful live test:

https://www.yahoo.com/tech/science/arti … 00224.html

I'll try to capture some of the text

Extreme Tech
Rotating Detonation Rocket Engine Survives First Subsonic Flight Test
Jon Martindale
Thu, May 15, 2025 at 12:16 PM EDT
RDE rocket launched from a rail system.
(Credit: Venus Aerospace)

A US-based propulsion company has successfully launched and flown a new rocket powered by a unique rotating detonation engine. Although relatively small by rocket standards, the test could pave the way for much more dramatic test flights, in turn leading to hypersonic rockets, jets, and even more efficient spacecraft in the not-too-distant future.
Rotating detonation engines (RDE) have been theorized and partially tested for decades and considered a potential future solution to inefficient chemical rocket space travel. It works by continually detonating explosions in a circular channel, allowing the igniting flame to travel at supersonic speeds. That makes for faster combustion and more efficient fuel usage. Some theoretical RDE designs could end up more than 25% more efficient than traditional combustion within rocket engines.
That's all in the future, but for now, a successful test of the 2,000-pound rotating detonation engine is a sizeable milestone for the technology and the company behind it. Venus is hailing this as the first successful test of this kind of rocket motor in the US. Indeed, this is ahead of NASA's own rotation detonation tests in 2023, which were merely static fires of an admittedly more powerful engine. The Japan Aerospace Exploration Agency, however, did successfully test its own rotation detonation engine in space in 2021, with further experiments in 2024.
That doesn't perturb Venus, though. Following the half-minute flight of this test rocket, it plans to continue developing the motor further.
"This milestone is a testament to what's possible when engineering rigor meets entrepreneurial urgency," said Dr. Rodney Bowersox, associate dean for research and professor of aerospace Engineering at Texas A&M University. "Rotating detonation rocket engines have been a scientific curiosity for decades. Venus is showing the world that they aren't just academically interesting—they're buildable, testable, and operational under real-world conditions. This is how aerospace innovation should look."

Engineered to work with Venus' VDR2 air-breathing detonation ramjet, it claims that this combination could enable aircraft to take off from a runway and reach speeds up to Mach 6, achieving and maintaining hypersonic speeds without requiring rocket boosters.
All of this is in service of developing the company's Stargazer M4, a passenger jet that could reach Mach 4 and enable hyper-fast transport for up to 12 passengers around the world, travelling almost anywhere in just a few hours.
"This milestone proves our engine works outside the lab, under real flight conditions," said Andrew Duggleby, co-founder and chief technology officer at Venus. "Rotating detonation has been a long-sought gain in performance. Venus' RDRE solved the last but critical steps to harness the theoretical benefits of pressure gain combustion. We've built an engine that not only runs, but runs reliably and efficiently—and that's what makes it scalable. This is the foundation we need that, combined with a ramjet, completes the system from take-off to sustained hypersonic flight."
Long term, this kind of technology could revolutionize travel and transport in Earth's atmosphere and beyond, paving the way for more efficient flight and space travel. That's still a long way off, but these early signs are promising that the technology could see some commercial use within the next decade.

(th)

tahanson43206 · 2025-05-24 07:55:54

For all... please keep a watch for articles on Rotating Detonation engine progress. In separate correspondence GW Johnson explained the likely cause of the low frequency oscillations/pulsations he has been observing since he lives close enough to the SpaceX McGregor test facility to have his house shaken when the new engines are tested. he has observed low frequency (sub audible) movement of the windows in his home.

GW's explanation is quite interesting, and I am hoping he will agree to post it in the forum.

In an attempt to translate, I get the impression that in traditional rocket engines, the physical design of the nozzles and the nature of the mixture is such that vortexes form as the gases attempt to combine. Apparently under some circumstances, a feedback loop can occur as the gases interact. I had inquired about causes of oscillation, and learned that the kind of oscillation (quarter wave) of organ pipes is quite different from the oscillation observed in rocket engines.

My question is whether the new Rotating Detonation engine design would be immune to the oscillation to which traditional rocket engines are subject.

(th)

tahanson43206 · 2025-06-08 08:08:45

The Wikipedia article on Rotating Detonation Engines appears to be updated as recently as May of 2025

https://en.wikipedia.org/wiki/Rotating_ … ion_engine

This page was last edited on 26 May 2025, at 00:20 (UTC).

One interesting hint I picked up from the article is that there may be a torque associated with the technique.

That seems reasonable to me, since force is being generated in a circle.

Airplanes already deal with torque, so an airplane application might not be a problem.

I wonder if the amount of torque would be great enough to be a problem for a rocket.

Google has a great number of citations for Rotating Detonation Engine.

There are multiple videos available.

(th)

Void · 2025-06-08 20:41:17

A thing that has confused me about chemical rocket engines, is do they have to do their burns in the shortest time possible to get the most desired effect?

Of course, when rising though the Earth's atmosphere maximum tolerable thrust is desirable as you want to reduce gravity losses. For example, the wrong thing to do would be to hover above a point until the propellant are all consumed.
But the machines being lifted can tolerate only a certain level of force towards the maximum, which would be an explosion, I suspect.

So, I suspect it would be some time before the rotating detonation engine could be used to make a machine rise from the surface of the Earth, though the Earth's atmosphere.

But once in LEO, and practical vacuum in orbit, could a small engine be used to do a long, small burn to propel a ship such as the Moon or Mars? The presumed oscillations which apparently may be part of the device might be managed if it were a small engine, relative to the mass of the machine it was mounted on, a spacecraft.

That is my guess, that it might be the lowest hanging fruit for the new engine described.

Ending Pending

Last edited by Void (2025-06-08 20:49:10)

kbd512 · 2025-06-11 02:23:05

Void,

Brief impulsive burns are the preferred way to operate all rocket engines to minimize gravity losses, but some engines don't generate enough thrust to do that. The insufficient thrust problem primarily applies to the various electric propulsion schemes, but sometimes applies to low power chemical propulsion for interplanetary spacecraft and satellites as well.

Gravity loss reduction is critical when attempting to achieve orbital velocity from a launch pad. Some user here, maybe PhotonBytes, has proposed using an air-breathing mixed propulsion scheme whereby a spaceplane repeatedly dips into the upper reaches of the atmosphere to obtain Oxygen for combustion, following a sinusoidal trajectory around the Earth until sufficient velocity is achieved that very little pure rocket propulsion (onboard oxidizer plus fuel), at greatly reduced Isp, is required to attain orbital velocity. Gravity losses will be higher in this scheme simply due to total thrusting time required to attain orbit, but the Isp is so much higher from not having to carry most of the oxidizer onboard the vehicle that so long as said spaceplane can tolerate the extreme aerodynamic heating it will be subjected to, the end result is a significant propellant consumption reduction. How well that works in real life remains to be seen. I place this scheme in the "easier said than done" category. I think it's probably attainable, but at what cost?

Gravity loss is highest (near or at 100%) when thrusting "straight up" to clear the dense lower atmosphere, as is typical of a TSTO design that uses most or all of the propellant in the booster stage to rapidly ascend through the lower atmosphere before the gravity turn and firing of the second stage, which provides the majority of the delta-V necessary to achieve orbit. If your rocket was powerful enough and you only intended to attain escape velocity, you could theoretically thrust "straight up" or at least "straight to your intended interplanetary target", and leave Earth's gravity well that way.

I believe the Nova booster concept, which was to be equipped with 8 F-1 engines in its first stage, was designed for a "direct ascent trajectory". For various reasons, mostly related to energy and monetary economics, direct ascent was ultimately rejected as the most practical means to achieve the lunar landings during the Apollo Program. Direct ascent requires a much more powerful rocket. The Saturn V, with only 5 F-1 engines in its first stage, was the "economy model" moon rocket. We had to master orbital rendezvous and docking to use a comparatively less powerful Saturn V, but those on-orbit maneuvers were subsequently proven to work quite well, and are used today every time a spacecraft visits ISS. Saturn V could deliver 140t to LEO. Saturn C8 (Nova) could deliver 210t to LEO, and the SpaceX Starship Super Heavy V3 aims for about 200t with full reusability or up to 300t when fully expended. If we were to expend the upper stage of Starship and include a Hydrogen powered third stage, then we'd have a very powerful direct ascent moon rocket.

After you're already in orbit, you can perform multiple burns, firing the engine at certain points along a highly elliptical orbit that are most favorable to increasing velocity (gravity assist) along your intended trajectory. This typically applies to escape maneuvers, but is also used by satellites going to higher orbits to reduce propellant consumption. GW will correct me if I'm wrong, but I believe his space tug concept also capitalizes on gravity assist thrusting to take the payload to just shy of escape velocity so the upper stage can be refueled and reused. Any engine capable of multiple restarts makes this possible. RL-10 upper stage engines could restart 15 and later up to 25 times. Sometimes restart limitations are imposed due to ignition, especially when using TEA/TEB or other pyrophoric compounds to achieve ignition, and sometimes stress on various engine components or cumulative damage. This will vary from engine to engine. RS-25 engines use the equivalent of a spark plug, so if there is enough electrical power then the number of possible restarts are not limited by the count of onboard pyro cartridges.

To the best of my knowledge, RDEs can achieve thrust-to-weight ratios (TWRs) better than conventional chemical rocket engines, so engine thrust would not be a limiting factor. Gravity and atmospheric drag, however, are always going to be limiting factors. After you're in orbit, as you noted, a RL-10 thrust level engine, provided that all components can handle the stress of firing, could perform long duration burns or multiple burns. Examples of small engines performing extended duration burns include engines powered by storable chemical propellants. Russian interplanetary space probes have upper stages with small engines that fire for quite some time, for example. Some engines can fire for cumulative times measured in terms of multiple hours, although most have single burn duration limits because their nozzles are made from uncooled refractory metals that can only withstand burns of a given duration. If the engine combustion chamber and nozzle were regeneratively cooled or made from Carbon-Carbon that can readily withstand the temperatures generated by even the hottest burning propellants, so long as the turbopumps / valves / seals / injectors / combustion chamber / nozzle and any other attached machinery can withstand the forces and temperatures involved, such an engine could theoretically fire for hours on end to achieve the desired Delta-V, if need be.

At least in theory, since RDE fuel injection systems operate in a manner more similar to a piston-driven internal combustion engine with electronic fuel injectors, they don't use turbopumps to feed the propellants. RDEs have very few components and fewer moving components than typical turbopump-fed rocket engines. They do require electronic propellant feed control with exceptionally precise injection timing. When sequenced correctly, it's as if the detonation of the propellant forms a sort of "standing detonation wave" that endlessly "swirls" around the combustion chamber. After initial ignition, in steady state operation no igniter is required to maintain combustion. Just as the first propellant injection and detonation sequence completes the second detonation sequence begins, and in so doing generates continuous thrust using a very rapidly pulsing rather than continuous fuel injection. Heat from the combustion chamber and nozzle can also be used to convert cryogenic liquids into gasses. You have more heated surface area to work with, as compared to a conventional rocket engine design with a bell nozzle, so the engine can generate more thrust than a typical pump-fed expander cycle engine like the RL-10, before a conventional expander cycle engine runs into heat transfer rate limitations from the limited surface area available to provide input power.

NASA has experimented with Hydrogen, Methane, and RP1 fueled RDEs with regenerative cooling. NASA has also experimented with Reinforced Carbon-Carbon (RCC) nozzles that don't absolutely require regenerative cooling. RCC is 1/4th the mass of the commonly used Nickel-Copper alloys (2g/cm^3 vs 8g/cm^3) and won't melt if left uncooled. The first goal behind using RCC was engine mass reduction / TWR improvement for upper stage engines. The RDE scale engine models used thus far are about half as heavy as conventional rocket engines for the same thrust output and dramatically more compact. That means a RDE fabricated from RCC could be about 1/8th the mass of a typical RL-10 for the same thrust output, so putting 8 RDEs on the back of an upper stage would provide about 880kN of thrust for the mass of a single RL-10. Add a few more engines and you achieve J-2X / Merlin thrust levels from something that's not much more mass than the RL-10. At that point, even with a fairly substantial payload, your burn is less than an hour in duration.

The current / active RL-10 versions are 170kg to 300kg, so a thrust-equivalent RDE would weigh about 85kg to 150kg. If advanced materials such as RCC are used throughout the engine, then the engine might weigh as little as 21.25kg to 37.5kg. That's a trivial amount of weight for a RL-10 thrust-class Hydrogen burning engine, mostly because it doesn't require an enormous Hydrogen turbopump. At some point traditional turbopumps are likely required to feed enough propellant, but the 10kN to 44kN RDEs that NASA has experimented with thus far don't require them. It's easy to see what the real advantages are. There's a modest Isp bump from using detonation, but engine size, mass, complexity, and fabrication time / cost matters at least as much. As we've already noted, as long as the engine can withstand the stress of firing, multiple hours of firing time are feasible and increased total burn time can be favorably traded-off against other engineering considerations such as the force that an upper stage engine will subject the entire vehicle to as the propellant mass is expended. Maybe you don't want an excessive amount of acceleration applied to a large interplanetary vehicle equipped with fully deployed solar and radiator arrays. A single large pump-fed engine can only throttle to a certain point. Multiple smaller RDEs would give you more throttling range to work with.

Air-breathing RDE variants have been simulated to function up to ~5.8km/s, which is well beyond burnout velocity for all booster stages I'm aware of. After that, you only need to add a couple more km/s using onboard oxidizer to attain orbital velocity. You also get a badly-needed TWR improvement with Hydrogen burning RDEs for a SSTO vehicle, although there's still not much you can do to reduce the additional Hydrogen tank mass, nor TPS mass if you intend to make the vehicle fully reusable. It's an interesting pathway towards a practical SSTO, but one that requires a lot more work. Mixed propulsion with air-breathing Methane burning RDEs is probably the way forward for SSTOs. You get to offload a lot of the oxidizer mass you'd need to carry in the vehicle, an Isp increase over RP1 without resorting to absurdly large LH2 tanks, and sky-high TWRs. Raptor 3's TWR is 200:1. Even if the best we can do is 400:1, the mass of the engine hardware becomes trivial, but 800:1 is at least possible. We could implement almost arbitrarily large SSTOs (300t to 500t to LEO) with payload mass fractions comparable to today's TSTOs.

Void · 2025-06-11 04:30:42

Thanks, I think that is helpful. A small Rotating Detonation Engine might have value in a mission once your mission is above the atmosphere, is some part of what you posted, I believe.

Ending Pending

kbd512 · 2025-06-11 09:02:48

Void,

Correct. After you're in orbit, overcoming the gravity losses associated with low TWR becomes far less important than other factors such as engine mass / complexity / physical size, because you would generally take advantage of the Oberth effect during thrusting periods to minimize gravity losses. Aerodynamic drag losses are almost nonexistent in LEO, although there is still some atmosphere which will ultimately reduce your vehicle's velocity through collision, over a period of weeks to months.

In a low orbit, traveling at roughly 7.8km/s, your vehicle already has quite a bit of momentum, which ought to be in your desired direction of travel, previously established by your launch trajectory. To go to Mars on a 6 month free return trajectory from an ISS-like orbit, you need to add another 3.9km/s. Whether said velocity increment is added over 15 minutes or 5 hours should be impacted relatively little by gravity losses.

Void · 2025-06-11 20:21:08

As it happens, I stumbled upon this today: https://www.msn.com/en-us/news/technolo … r-AA1Gs1WK Quote:

Texas startup making history with revolutionary rocket engine
Story by Brian Spencer • 1d •
4 min read

Quote:

"Detonation engines have been theorized for many, many decades," Duggleby said. "To really frame this in the right perspective - the SpaceX rocket engine is only about 2% better than the Apollo engines that took astronauts to the moon. And that 2% was fought over for decades.
"This technology is a 10 to 30% jump. So it's just, just massive."

Video: https://www.bing.com/videos/riverview/r … ORM=VRDGAR Quote:

Texas startup making history with revolutionary rocket engine
YouTube
Straight Arrow News
1 day ago

Seems like it could change things in space as well.

Ending Pending

Last edited by Void (2025-06-11 20:27:10)

kbd512 · 2025-06-13 06:56:20

Void,

33X 200:1 TWR Raptor-3 engines weigh 49,500kg.
33X 800:1 TWR RDEs would weigh about 12,375kg.

The propellant tanks weigh 80,000kg when made from 304L stainless, and the interstage weighs about 20,000kg. If the propellant tanks were made from CFRP, then mass is around 20,000kg. Those two changes can save around 97,125kg, cutting the dry mass of the Super Heavy Booster in half. That is a massively consequential upgrade to both engine performance and propellant tank dry mass.

Between those two night-and-day performance enhancements to our chemical rocket technology, SSTOs become not merely technically feasible but reasonably practical. Flight physics dictates that a TSTO will always soundly beat a SSTO on payload performance unless dry mass approaches zero. We really can't do much about that, and any SSTO mass reduction will likely carry over to a TSTO design with better payload performance. However, TSTO does not beat SSTO in overall simplicity of operation. By its very nature, any vehicle that deliberately breaks apart into multiple pieces during its flight, with separate recovery locations for the different pieces of the vehicle, is going to be more complex, time consuming, and expensive to operate. When a mixed propulsion SSTO can take off and land on a conventional runway like other jet aircraft, VTVL TSTOs will be relegated to specialist heavy cargo delivery roles because we won't be able to justify the operating expenses or dangers associated with operating them for routine commercial orbital passenger service.

So long as it's not too extreme, justification for the additional propellant consumption of a SSTO vs TSTO for passenger service is little different than justifying the additional fuel consumption of much faster jet aircraft capable of carrying more passengers over shorter duration flights vs slower piston engine aircraft that carry fewer passengers. Reduced propellant consumption is always desirable, but by mass most of the propellant is oxidizer rather than fuel. That is why a mixed propulsion scheme that significantly offsets the onboard oxidizer mass by consuming atmospheric Oxygen is the most likely path to a commercially successful SSTO that is at least GLOW-competitive with TSTOs.

I view RDEs, mixed propulsion consuming atmospheric Oxygen, advanced high strength composites, and advanced ceramic composites for engine components and thermal protection as the technological underpinnings required for the success of the National AeroSpace Plane (NASP) Program. NASP began in earnest in the 1980s, although conceptual studies began in the early 1970s. Here in 2025, it's starting to look like we finally have the technology to make a practical NASP vehicle, which was intended to be a SSTO with mixed propulsion.

It took about 50 years to go from Wright Flyer to Boeing 747, 50 years from Saturn V to a rapidly reusable super heavy lift launch vehicle (Starship Super Heavy), and about 50 years yet again, to go from a 1.5 stage-to-orbit semi-reusable space planes (Space Shuttles) to rapidly reusable SSTO space planes (NASP), so time-wise I would say we're on-track.

RAND Report on NASP:
The National Aerospace Plane (NASP): Development Issues for the Follow-on Vehicle - Executive Summary

Void · 2025-06-13 09:31:40

So, that implies to me that over the next 50 years, the method of Starship will eventually be considered outdated. But that does not mean that it is inappropriate now.

I imagine that it will be similar to how we may look at technology made 50 years ago.

But for now Stainless Steel and Raptors are likely to be competitive. But as the other competition emerges, they will all have to try to improve their edge little by little.

With robotic labor hardware costs may go down, in which case building with lighter materials may become the sensible thing to do.

I am glad that there is room for better engines and lighter materials over time.

Ending Pending

Last edited by Void (2025-06-13 09:35:21)

kbd512 · 2025-06-13 12:03:50

Void,

Absent exceptionally powerful new propulsion technologies, likely involving fusion, TSTOs like Starship are always going to have their place in the expanding menu of orbital launch vehicle options.

The trends are clear:
1. full reusability with minimal repair and refurbishment
2. streamlining of operational costs by reducing repair costs and time
3. substantial improvements to engine performance as we push towards the theoretically achievable Isp values and combustion efficiency limits for conventional chemical engines
4. exploration of non-rocket launch methods for lower value cargo such as propellants and metals
5. new materials for thermal protection

That said, no amount of materials science is going to overcome basic orbital flight physics. A TSTO will have a higher payload mass fraction than a SSTO, period. Thus far, no miraculous new strong but light material nor more efficient engine technology can compete with the ability to discard 2/3rds of your vehicle's inert / dry mass at booster burnout, because pretty much any tech you can apply to a SSTO can also be applied to a TSTO, to even greater effect, especially if full reusability is important. The only area where TSTO cannot compete with SSTO is the simple fact that a SSTO nominally stays in one pretty piece, from launch to landing, and doesn't require multiple separate recovery, transport, and stacking operations. Starship Super Heavy partially reduces that problem by returning the booster to the pad it launched from. That's a great first step, but the upper stage stilll has to land somewhere and get stacked on top of a booster again. Whatever savings there are in terms of propellant cost reduction, it probably cannot offset all that added operational complexity.

SpaceX wants to launch Starship several times per day. With a HTHL SSTO, after the vehicle lands, it's towed back to the gate, inspected, gassed up, towed back to the runway, and then it's ready for its next flight. That might feasibly be done in the span of a few hours. Even a VTHL SSTO can be pretty straightforward to relaunch. A VTVL TSTO requires multiple pads with supporting infrastructure, if only because the upper stage cannot readily land on top of its own booster, multiple cranes are needed for stacking, inspecting two vehicles vs one must be done (only a major time imposition because it probably needs to be done in two or more different places for a large fleet operation), etc. All that adds up to more time, more touch labor, and more cost.

I want SpaceX to succeed, or any body else to succeed for that matter, but launching this way 1,000+ times per year (at least 3,500 per year to send 1,000 Starships to Mars every 26 months, or about 10 per day) is going to be challenging without a standing army of support personnel and sprawling infrastructure investment. Doable? Yes. Very cheap and practical with limited head count? Almost certainly not, at the scale required- too many simultaneously moving pieces. Maybe AI and those Tesla automatons can coordinate the touch labor operations to the degree required, plus they can stay in a shelter on the pad, immediately ready to inspect the booster or upper stage the moment it's captured on the pad. We will never run recovery and vehicle inspection operations that way with humans for obvious reasons. Ground crew are not allowed anywhere near the vehicle during the launch and recovery.

All I know for sure about vertical launch and vertical landing is that it's very tricky to do correctly and routinely. It's not impossible, obviously, but far from simple and easy. Automated landing helps mightily, but what about 24/7/365 airline transport style operations? We don't do that with rocket launches. If a hair looks out of place, we stop the clock, reassess, and everyone involved weighs in before we proceed. We run launch operations that way to stave off disaster. The primary precious cargo we're transporting to Mars is lots and lots of people. Losing even one vehicle per year is unacceptable. Starship is clearly not an optimal solution for an airline transport style operation, even though it's acceptable for exploration missions. Neither a SSTO nor the upper stage of a TSTO alone is the correct solution for colonization. Refueling up to a half dozen times in LEO is impractical at the scale they want to implement. We need some kind of fusion-based in-space propulsion system for practical large scale colonization operations. A solid core NTR is woefully inadequate for that task.

tahanson43206 · 2025-06-13 12:42:28

Reminder:

Void wrote:

Marcus House had a video out that mentioned this engine. Possibly as much as 25% more efficient, it is said.
https://en.wikipedia.org/wiki/Rotating_ … ion_engine
https://www.youtube.com/watch?v=UutHG8Y2UuQ
Done

Void, this is your topic, and you can depart from the title if that is your preference.

However, NewMars readers might hope the topic might have something to do with the title.

It would be disappointing to a serious reader to find the topic drifting.

(th)

Void · 2025-06-13 17:41:49

A reasonable caution (th). Still, kdb512's broad picture provides a canvas onto which we might hope to apply a Rotating Detonation Engine in time.

Both the engine, and the lightweight materials rebel against the stipulation that the Earth is almost too massive for chemical rockets to get to orbit. We tend to dismiss the Earth as a source of cost appropriate materials for activities in space.

No blame to you (th) to express such a concern, it is appropriate, but I am trying to find a loophole in reality to get just a bit better in the game.

Now, if I bring this material into this topic: https://newmars.com/forums/viewtopic.ph … 54#p232154
Quote: OK I will try to go a bit further with this:

So, possibly SpaceX could make a propellent transporter like this.

Blue Origin might try to make a "Jarvis" like this.

Relatively Space who has abandoned a reusable 2nd stage, might eventually try something like this.

If something like the Starship system were cut into three stages:
1) The Superheavy would be very close to what is being developed.
2) The Starship would be cut into a base part that would be a cylinder and have almost all of everything except for:
3) The Main Tanks, and perhaps a Rotating Detonation Engine.

Perhaps someday RDE's would replace the Raptors on Starship, but for now a small engine would be a good early method to explore the new technology.

The scheme is a bit different than the standard reuse or expend options for Starship. Section #3 would be mass produced at a low cost, and put into orbit, the liquid propellants could go to a depot above LEO, and the materials of the tank may be reused in orbit to build structures or to provide propellants. I had thought that the #3 main tanks might be made of an Aluminum/Lithium Alloy, but kdb512's option might have merit.

There have been imposed a form of logic that main tanks are of low value. The Space Shuttle abandoned its main tank to destruction although far out ideas existed to bring it to orbit as an option and use it for something.

Vulcan has a similar logic where the 1st stage main tanks are to be discarded and the engines and avionics are to be saved, (Eventually in future versions), by an expandable one-time heat shield.

I want to put is backwards. The main tanks for the 2nd stage still contain some of the liquid propellants, and in my opinion the tank walls if for instance of Aluminum/Lithium, may be useful to reform into orbital structure and/or to be made into propellants for Magdrive or Neumann Drive. Having been lifted to such an altitude with so much cost and effort, the sensible thing is to bring them fully to orbit, perhaps with a small very efficient engine.

Perhaps with a "Rotary Detonation Engine".

In doing this we turn the wet and dry mass of the Main 2nd Stage tanks into value in and above LEO.

We also get rid of the task of bringing the main tanks back though reentry with heat shielding.

And I should explain my intentions for the portion of the 2nd Stage which would not be the Main Tanks or the Rotatating Detonation Engine.

The Cylinder Section of the 2nd Stage as shown in the drawing would need to come down though the air braking, using flaps just like a Starship but not having a pointed nose. The Starship doe not so much glide as to do a sky-dive, so this may not be impossible.

I anticipate that the heat shield method could be active cooling with the Methane Tank having extra capacity to provide the coolant needed.

While this is a significant cost, even so we avoided having to use active cooling for the Main 2nd Stage Tanks, because we left them in orbit.

This version would be created specifically as it might efficiently provide both liquid propellants such as Methane, Oxygen, Argon, Hydrogen, but it's Aluminum/Lithium Structure may be relatively easy to cut into pieces and even melted and used in things like 3D print processes.

So, yes much more than the Rotating Detonation Engine are discussed, but it would be an important part of the whole process.

Ending Pending

Last edited by Void (2025-06-13 18:12:21)

tahanson43206 · 2025-06-15 06:16:24

In the kbd512 Postings topic, kbd512 provides on overview of engine types and capabilities, and concludes with description of an RDE application using atmosphere for oxygen and nitrogen throw mass.

https://newmars.com/forums/viewtopic.ph … 27#p232227

A highlight of the post is a chart showing propulsion system performance.

The chart compares hydrogen and hydrocarbon fuels.

(th)

kbd512 · 2025-06-15 17:25:07

Theoretical total quantity of energy output available from 1m^3 of LH2 and RP1 fuel:
LH2 at 70.85kg/m^3 and 143MJ/kg: 10,131.55MJ/m^3
RP1 at 820kg/m^3 and 43MJ/kg: 35,260MJ/m^3

That means 235.632kg of RP1 (when combusted with 612.643kg of LOX) stores about as much energy as 70.85kg of LH2 (when combusted with 425.1kg of LOX), understanding that neither of the Oxidizer / Fuel ratios for LOX/LH2 (6:1) or LOX/RP1 (2.6:1) are precisely stoichiometric, even though both seem to produce the best combination Isp (propellant economy) and thrust (raw power to accelerate the vehicle) for each type of rocket engine. It may also be the case that precisely stoichiometric O/F ratios start melting engine components due to excessive heating. It's probably a combination of economy, power, reaction kinetics in chemical combustion, and material limits.

That also means 495.95kg (1.373m^3) of LOX/LH2 is about equal to 848.275kg (0.824m^3) of LOX/RP1. This appears to be a very clear point of advantage in favor of LH2, until we recall that we also get approximately half as much thrust per unit of engine mass by using LH2 and require significantly more total propellant tank volume for equal Total Impulse, at which point it looks more like a great way to increase the vehicle's volume and thus dry mass.

To make Total Impulse / energy output completely equal, we need 3.48X greater fuel tank volume for the LH2, so 246.558kg of LH2 (35,260MJ / 10,131.55MJ) = ~3.48m^3 and 3.48 * 70.85kg of LH2 = 246.558kg. That means 246.558 * 6 (O/F Ratio for LH2) = 1,479.348kg. The LOX/RP1 equivalent is 2,952kg, so 820kg * 2.6 (O/F ratio for RP1) = 2,132kg. Therefore, 2.869m^3 of LOX/RP1 stores the same energy as 4.251m^3 of LOX/LH2. If getting to orbit using a single stage was merely a matter of propellant mass, then case closed. Hydrogen wins the argument handily. If only it was that simple...

For any SSTO propellant tank, which accounts for the great majority of the vehicle's internal volume, the limiting factor on propellant tank mass and stiffness, in order to resist progressive deformation to structural failure, appears to be the propellant tank's internal pressurization load required to feed the propellant, rather than structural failure associated with the mass of the propellant itself when applied acceleration loads are factored in. The very low density of LH2 requires significantly higher internal pressurization loads to force-feed very low density Hydrogen into the turbopump inlets. The internal pressurization load is so great for LH2 that for vehicle acceleration loads up to 3g, propellant tank mass to support the propellant mass need not increase at all, even when it contains very dense LOX (1,141kg/m^3) vs LH2 (70.85kg/m^3) or RP1 (820kg/m^3). Essentially, any propellant tank built to withstand LH2 internal pressurization loads is de-facto sufficiently strong and stiff to withstand acceleration loads applied while carrying LOX or any lighter propellant, such as RP1.

The real measurable difference is that any vehicle requiring 48% less internal volume will be both stronger and stiffer because it's made from materials with fixed / unchanging tensile strength and stiffness values, regardless of size. Those two material properties don't change one iota as the vehicle's physical size increases, so the vehicle either has to become stronger and stiffer, which increases its dry mass fraction, or else it becomes weaker and less able to withstand repeated stresses, which is not tolerable past a certain point. On top of that structural issue, a vehicle with 48% greater internal volume will also have more surface area, which must be covered with additional thermal protection materials if the vehicle is to be reusable.

Something has to give here. If you insist on using LH2 because it provides a significant wet mass and Isp advantage, then you're designing a vehicle with meaningfully reduced structural margins by default. This is what ultimately doomed Lockheed-Martin's VentureStar SSTO which was being developed to replace the Space Shuttle. VentureStar's structural mass margins were so razor-thin that the LH2 tanks burst, cracked, or otherwise failed pressurization tests. Reducing its propellant tank volume by 48% and doubling the engine thrust surely would've helped its viability as a fully and rapidly reusable SSTO Space Shuttle analog vehicle, but everyone was fixated on using LH2, so the project failed because Lockheed-Martin was unwilling to pay for fabricating additional copies of stronger yet heavier propellant tanks that would've blown their mass budget and resulted in a vehicle that failed to meet payload performance requirements.

A vehicle operator pays for the propellant and the vehicle, but the vehicle manufacturer pays for the materials and labor required to construct the vehicle. If you can make the vehicle 48% smaller while providing the same payload performance, most of the time that's going to work in the favor of both the manufacturer and the operator. To this day, despite all our technological advances related to handling cryogens, there are precisely zero airline transport services using LH2 to fuel their fleet of aircraft. Many experiments with both LH2 and LCH4 fuels have been conducted by governments and airline transport services, but the value proposition was never there, and there's zero infrastructure to use these new fuels. To this day, Hydrogen with energy-equivalence to a given mass of kerosene still costs more money than kerosene. Until we can achieve cost parity, we may as well stick with cheaper and easier to use fuels.

For a SSTO, we should be fixating on vehicle dry mass, rather than vehicle wet mass, because that's dictating the vehicle's useful payload. What' I'm hoping for is that air-breathing RDEs substantially reduce the onboard LOX mass fraction. The energy density advantage of kerosene, and the smaller / lighter vehicle it allows for, will make up for its gravimetric energy density deficiency, when compared to LH2.

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Engineered to work with Venus' VDR2 air-breathing detonation ramjet, it claims that this combination could enable aircraft to take off from a runway and reach speeds up to Mach 6, achieving and maintaining hypersonic speeds without requiring rocket boosters.

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