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

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.

------
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.

------
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)

New Mars Forums

Announcement

#1 2023-02-04 08:50:31

Rotating Detonation Engine

#2 2023-02-04 10:39:17

Re: Rotating Detonation Engine

#3 2023-02-04 12:16:27

Re: Rotating Detonation Engine

#4 2023-02-04 15:05:57

Re: Rotating Detonation Engine

#5 2023-02-04 15:53:14

Re: Rotating Detonation Engine

#6 2023-02-17 08:25:26

Re: Rotating Detonation Engine

#7 2023-12-22 21:01:05

Re: Rotating Detonation Engine

Board footer