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Enhancement of Detonation Wave Dynamics in Rotating Detonation Combustors (RDC)

Description:

RT&L FOCUS AREA(S): Hypersonics; General Warfighting Requirements

TECHNOLOGY AREA(S): Air Platforms; Weapons

OBJECTIVE: Develop a Rotating Detonation Combustor (RDC) with passive acoustic control technology such as lining the augmentor duct, capable of suppressing unwanted secondary waves that subtract energy from the main precessing detonation wave, resulting in increased combustion efficiency.

DESCRIPTION: Currently, military gas turbines use subsonic deflagration (not detonation) thrust augmentors to reheat the core flow prior to expansion through the engine nozzle. This requires the use of bluff body flame stabilization methods that are achieved by the insertion of mechanical structures into the core flow near the exit of the turbine. These flame stabilization methods anchor the flame near the turbine exit and establish an environment for constant pressure combustion or “pressure loss” combustion. These features cause significant flow disruption and an unnecessary total pressure loss when the augmentor is not in operation.

Alongside the performance issues of conventional thrust augmentors there exists a prevalent combustion instability issue which manifests itself in the form of transverse and longitudinal waves inside the augmentor duct. Combustion instabilities are caused by heat release fluctuations that excite natural acoustic modes in resonating chambers. These fluctuations give rise to powerful pressure oscillations that lower the combustion efficiency of the device and ultimately may damage engine hardware [Ref 4]. One acoustic mode of interest is the transverse (or radial duct) mode typically occurring in the range of 1-10 kHz. In some cases, transverse mode pressure oscillations can exceed 9% of the base pressure, equivalent to pressure amplitudes of 0.6-0.8 bar (peak-to-peak) in a 7 bar combustor.

An RDC will also show signs of transverse instability wave modes however at a lower magnitude. To attenuate combustion instability modes generated by the RDC, a passive suppression technology, such as acoustic absorption coatings will be applied to the inner walls of the RDC duct. These absorption coatings can be comprised of porous high-temperature composites engineered to absorb a given frequency range.

Addressing the core issue of conventional augmentors, an RDC-based augmentor does not require a mechanical structure for flame stabilization and would eliminate this performance and durability loss. By replacing the typical flame stabilization geometry with an RDC-based augmentor the mechanical complexity is greatly reduced and “pressure gain” combustion can be exploited by harnessing detonations which are shock-coupled, supersonic combustion waves trapped in a continuous ‘spin mode’ about an annulus [Refs 1-3]. During pressure gain combustion, there is an effective rise in pressure across the RDC which allows for greater thermal power efficiencies to be achieved compared to constant pressure combustors while using the same amount of fuel.

By utilizing detonation waves, higher rates of reactant mass consumption are achieved due to wave propagation speeds that are typically three orders of magnitude greater than deflagrations of the same reactant mixture. This indicates an overall enhancement in combustion power density. The benefit of enhanced power density can be appreciated by engine size and mass reductions. Typical deflagration-based thrust augmentors require exceedingly long augmentor ducts due to the reduced levels of oxidizer, which attenuate chemical kinetic rates. This in turn leads to low efficiency combustion and an exceptionally low combustion power density.

By implementing RDC-based thrust augmentors, naval aircraft engines will showcase increased power thermal density and increased thermal efficiency as much as 7% of a standard Brayton cycle while providing a safer mode of operation [Ref 5].

PHASE I: Develop a passive acoustic suppression technique, such as a coupon-sized acoustically absorptive lining material tuned to the frequency range of interest and verify its performance via an appropriate demonstration, such as the design and fabrication of a real-ultrasonic, pressurized impedance tube. Both radial and axial modes spuriously triggered in an RDC environment should be targeted, as well as various base pressures. A broadband characterization of the acoustic performance of such material is expected. The technology solution should survive in the RDC environment experiencing pressures from 1 to 20 atmospheres and temperatures in excess of 2000o K [Ref 6]. Develop a Phase II plan.

PHASE II: Develop an RDC benchtest demonstrator with an acoustic suppression technique, such as a coating, proving enhancement of detonation wave dynamics and reduced structural loading of the augmentor duct. Testing will involve a range of fuel-to-oxidizer ratios and base pressures. High speed diagnostics will be used to assess the RDC response. Upon successful completion of benchtest experiments, the developed RDC technology should be integrated in a combustor test facility to demonstrate operations at sea level conditions.

PHASE III DUAL USE APPLICATIONS: Integrate the RDC into an engine for ground-based demo. Develop plans for flight worthy hardware. RDC technology can also be developed for commercial gas turbine applications for aviation, marine, and land-based power generation.

REFERENCES:

  1. Bykovskii, F.A., Zhdan, S.A., Vedernikov, E.F. and Samsonov, A.N. “Scaling Factor in Continuous Spin Detonation of Syngas - Air Mixtures.” IOP Conf. Series: Journal of Physics: Conf. Series 899 (2017) 042001. https://iopscience.iop.org/article/10.1088/1742-6596/899/4/042001/pdf
  2. Bykovskii, F.A., Zhdan, S.A. and Vedernikov, E.F. “Continuous Spin Detonations." Journal of Propulsion and Power, vol. 22, no. 6, pp. 1204-1216, 2006. https://arc.aiaa.org/doi/abs/10.2514/1.17656?journalCode=jpp
  3. Bykovskii, F.A., Zhdan, S.A. & Vedernikov, E.F. Continuous Spin Detonation in Annular Combustors. Combust Explos Shock Waves 41, 449–459 (2005). https://doi.org/10.1007/s10573-005-0055-6
  4. Lieuwen, Tim C. “Unsteady Combustor Physics.” Cambridge University Press, 2012. https://www.cambridge.org/core/books/unsteady-combustor-physics/77A89B0BB731551B6FB2BF7F632F9A8B
  5. Sousa, J., Paniagua, G., and Collado Morata, E. “Thermodynamic analysis of a gas turbine engine with a rotating detonation combustor." Applied Energy, vol. 195, 2017, pp. 247-256. https://www.sciencedirect.com/science/article/abs/pii/S0306261917302684
  6. Schwer, D.A. and Kailasanath, K., “Feedback into Mixture Plenums in Rotating Detonation Engines,” AIAA Paper 2012-0617, 50th AIAA Aerospace Sciences Meeting, Nashville, TN, 9-12 Jan 2012. https://doi.org/10.2514/6.2012-617
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