Rotating Detonation Engine for Rocket Propulsion

Description:

TECHNOLOGY AREA(S): Space Platforms 

OBJECTIVE: Increase performance of liquid rockets while reducing system size and complexity 

DESCRIPTION: Continuous wave detonation based combustion is an alternative to conventional deflagration based combustion that increases thermodynamic efficiency via a reduction in generated entropy. Due to the high reaction rates, detonative combustion is a nearly constant volume process and is therefore capable of producing substantial increases in pressure (~10x) in a relatively compact volume. As a result, detonative combustion is of interest to the rocket propulsion community where higher efficiency and reduced size/weight are advantageous. The rotating detonation engine (RDE) is a quasi-steady detonative combustion device in which one, or more, transverse detonation waves travel normal to the axial injection of propellant in an annulus. The propellant mix is combusted in the transverse detonation wave and expands supersonically from the open end of the annulus. The propellant is constantly injected and number of transverse detonation waves and their speed is such that the fuel-oxidizer mix is denotatively combusted prior to exiting the annulus. In this way, the detonation, once established, is continuous. Although a number of laboratory model RDEs have been constructed, most research has focused on air breathing terrestrial applications and rocket propulsion has not been extensively or systematically studied. Demonstrations of the rocket cycle and performance verification using fuels and oxidizers of interest is needed prior to adoption of RDE technology by the flight community. In addition, system architectures for propellant flow control, ignition, and thrust control require definition. This solicitation is seeking to construct scaled rocket propulsion gas generator, pre-burner, or thruster based on RDE combustion and demonstrate its performance. Where performance is defined as coupling to turbo-machinery for gas generators and pre-burners, and as specific impulse for thrusters. Gas generator and pre-burner concepts are assumed to applicable for launch applications and relevant fuel/oxidizer combinations (e.g. RP-2 and oxygen, methane and oxygen). Main combustion chamber concepts may either be related to scaled launch applications or to spacecraft related storable propellants, including various monopropellants. Heavy-weight hardware is acceptable for system validation so long as a path to flight weight hardware development is defined. Other critical portions of the system architecture must be sufficiently well defined to ensure future technical demonstration. Recognized uncertainties in the development of RDEs for rocket propulsion applications include performance, heat transfer rates, and scaling to high pressures and fuel-oxidizer ratios. Direct measurement of performance and the energy balance has yet to be realized in these devices. At this time, modeling and simulation is in advance of experimental characterization of pressure gain devices. This solicitation is seeking to develop design methodologies, quantification of performance, and identification of loss mechanisms. Proposed efforts will need to include analysis of injection and mixing dynamics, scaling for geometry and pressure, thermal effects, as well as material compatibility. Laboratory testing will measure performance and identify relevant loss mechanisms. 

PHASE I: Phase I will design a RDE rocket component designed to function with conventional rocket fuels and oxidizers. Device operability should be demonstrated in a laboratory environment 

PHASE II: Phase II builds upon the analysis and testing performed in Phase I to further refine the RDE device with performance measurements. The device will be performance optimized and demonstrated at the system level. Identification of transition partner for a potential technology demonstration follow-on program will complete this task. 

PHASE III: Phase III further matures the technology developed in Phase II. All hardware as well as all supporting subsystems will be fully identified and developed. A demonstration system is readied for a Class D flight demonstration appropriate to the sizing selected by the funding agency. 

REFERENCES: 

1: W.A. Hargus, Jr. and E.J. Paulson, "Rotating detonation rocket technology development at the Air Force Research Laboratory," Joint Meeting of the 11th Modeling and Simulation, and 9th Liquid Propulsion JANNAF Subcommittees, 5-8 Dec. 2016, Phoenix, AZ.

2:  R.D. Smith and S.S. Stanley, "Experimental Investigation of Continuous Detonation Rocket Engines for In-Space Propulsion," Propulsion and Energy Forum, 52nd AIAA/SAE/ASEE Joint Propulsion Conference, July 25-27, 2016, Salt Lake City, UT.

3:  M.L. Fotia, F. Schauer, T. Kaemming, and J. Hoke, "Experimental study of the performance of a rotating detonation engine with nozzle," Journal of Propulsion and Power, Vol. 32, No. 3, pp. 674–681, May-Jun 201

4:  P. Wolanski, "Detonation Engines," Journal of KONES Powertrain and Transport, Vol. 18, No. 3, pp. 515-521, 201

KEYWORDS: Rocket, Detonation, Rotating Detonation Engine, Bipropellant, Monopropellant, Gas Generator, Pre-Burner, Spacecraft Propulsion 

CONTACT(S): 

William Hargus (AFRL/ RQRC) 

(661) 275-6799 

william.hargus@us.af.mil 

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