7
$\begingroup$

Nasa just announced successful testing of their new rotating detonation engine:

https://www.nasa.gov/centers/marshall/feature/nasa-validates-revolutionary-propulsion-design-for-deep-space-missions

I was curious if they had a published paper on this on the measured or theoretical ISP for such an engine?

$\endgroup$
1
  • 3
    $\begingroup$ News: Verizon gets fined for allowing links to NASA's 'wheel of explosions' video... $\endgroup$ Jan 26, 2023 at 22:35

1 Answer 1

4
+100
$\begingroup$

Have they published a paper on this?

Yes. Some of the researchers working on the RDRE and associated test campaign at the Marshall Space Flight Center very recently published a paper in the AIAA SciTech Forum. Unfortunately I don't think the full document is accessible with out subscription. They do not cite values for Isp or C*

..on the measured or theoretical ISP for such an engine?

Not by these NASA folks, but by others. Both experimental and theoretical values for Isp were published by some industry researchers in a 150lbf RDRE using GOX paired with Methane, Ethane, and Ethylene in JPP. Again, this may be hidden behind a paywall. They find Isp's in the range of ~100-200s.

Surprise. RDE's are not supremely game changing and are subject to the same propellant energy limitations that other rocket engines face. There is only so much power to be extracted from a given propellant combination.

The primary benefit RDRE's might provide is reduced weight. They may also achieve higher Isp's than traditional rocket engines under the same propellant and head pressures thanks to the pressure gain combustion, but we are talking on the order of 10's of seconds improvement. The weight reduction, however, is significant and impactful. They offer more complete combustion in shorter chambers, and, as I alluded to, potentially smaller pumps for the same 'Chamber Pressure'; the nozzle length may also be significantly reduced if supersonic combustor exhaust can be maintained (a big challenge in the current space).

I should also mention that there are now almost too many RDE's to count. They are popping up in universities all over the world. Very trendy. My lab campus currently has at least 4 RDE test stands. The reason for this is that the 10's of seconds of Isp gain I mentioned. That may not make space enthusiasts salivate, but it certainly will make aircraft manufacturers and operators start full on drooling. The vast majority of experimental RDE's right now are motivated by air-breathing jet combustor applications. You will easily find Isp data published in this sea of research, but most of it is cited as thrust per fuel weight flow.. The Isp for an air-breathing engine. These will be an order of magnitude higher than Rocket Isp's

All that said, I think we can estimate the Isp of the NASA RDRE

They cite the following information:

Thrust = 4,171 lbf

Chamber Pressure = 622 PSI

Inner Body Diameter = 5.59 inches

Annulus Gap Width = 0.33 inches

Propellant: GOX/CH4

This is enough information to make a rough Isp estimate. It is rough because I did not find combustion temperatures cited, so the ratio of specific heats (and mass flow) can not be precisely known; additionally I will be assuming chocked flow with in the annulus, which is not precisely accurate for an RDE.

But here we go:

I ran NASA's CEA with GOX/CH4 and found took the gamma at the optimal O/F ratio (found this to be ~3.5). The ratio of specific heats is then ~1.15 and the chamber temperature is ~3,600 K.

These assumptions should give a very optimistic value for Isp.

I use the choked mass flow equation with the above equation to get:

Mass Flow = 16.7 kg/s (36.8 lbm/s)

Then the Isp based on their peak thrust conditions is 113 seconds

This is similar to the values achieved in the smaller RDE's but I suspect it is actually an underestimate. But we will need to wait for more information to be published on this specific experiment.

Using the cited expansion ratio of 5, CEA estimates an ideal vacuum Isp of 327s

I should ALSO mention that an RDC installed in a jet engine gets to use TURBINES to extract energy and condition the exhaust, benefits that a rocket engine does not have. The exhaust flow is quite chaotic and with majorly circumferential velocity components. These will create cosine losses in the thrust and reduce the Isp, but can actually be used to the benefit of turbine efficiency.

$\endgroup$
1
  • 2
    $\begingroup$ Again, I learned something! $\endgroup$ May 18, 2023 at 17:24

Your Answer

By clicking “Post Your Answer”, you agree to our terms of service and acknowledge you have read our privacy policy.

Not the answer you're looking for? Browse other questions tagged or ask your own question.