Rotating Detonation Engine Research Articles

Experimental study on the rotating detonation engine based on a gas mixture

A direct connection test of a rotating detonation engine was conducted. The outer and inner diameters of the annular combustors were 206 and 166 mm, respectively. High enthalpy air was used as an oxidizer, and a mixture of hydrogen, carbon monoxide, and methane was used as fuel with a volume fraction of 5/4/1. The mixture was injected through small holes, and air was injected through annular slots. The effects of combustor length, width of annular slots, and the equivalent ratio on formation, development, propagation, and flameout of rotating detonation waves were analyzed, and several modes of rotating detonation were observed. It was found that when the width of the air annular slot was within the range of 3–5 mm, the pressure of the detonation wave was augmented with an increased slot width. As width increased, detonation waves became unstable. In study test conditions, an annular slot width of 6 mm was the critical condition for the formation of stable detonation. When the slot width was 4 mm and combustor length 160 mm, the phenomena of the conversion between single and double waves, double-wave collision, and conversion of the propagation direction occurred at different equivalent ratios. When the equivalent ratio was 1.2/1 and the slot width was within the range of 3–6 mm, the slot width was inversely related to the detonation wave velocity. When the slot width was 4 mm and the equivalent ratio was 1.0/1–1.2/1, the slot width was positively correlated with the detonation wave velocity. When the combustor length was shortened to 80 mm, the propagation mode of the detonation wave was changed to a single wave first and then to a double wave in the same direction, and the velocity reduced from 1130.69 to 1024.16 m/s. The injector used in the test inhibited the propagation of back pressure from the combustor.

Effects of detonation instability and boundary layer on flame propagation behavior in millimeter-scale smooth tubes

Small pre-detonators have potential applications in practical pulse detonation engines or rotating detonation engines. The effects of detonation stability parameter and area divergence on flame propagation in capillary tubes are experimentally analyzed using high-speed cinematography. The four mixtures (namely, stoichiometric methane/oxygen, ethylene/oxygen, propane/oxygen and stoichiometric acetylene/oxygen/50% Ar mixtures) used have different values of stability parameter. The inner diameters of the tubes (namely, 0.5, 1.0, 2.0, and 4.0 mm) investigated provide a variation in area divergence at initial pressures p0 ranging from 5 to 100 kPa. The results show that stability parameter and area divergence play positive and inhibitory roles, respectively, in the process of flame propagation. Five propagation modes are observed (namely, steady detonation, stuttering detonation, galloping detonation, deflagration, and no flame), among which the stuttering detonations have three different forms, depending on the product of stability parameter and area divergence. For the most-unstable mixture (stoichiometric methane/oxygen, with the largest values of stability parameter), the range of the product of stability parameter and area divergence in which steady detonation can occur is the narrowest. The results also indicate that the stoichiometric ethylene/oxygen and stoichiometric propane/oxygen, with similar values of the product of stability parameter and area divergence, have the same propagation modes. Further analysis suggests that the Chapman–Jouguet deflagration-to-detonation transition (DDT) distance L decreases with increasing of stability parameter and area divergence for all four mixtures. Compared with the other three mixtures, stoichiometric methane/oxygen, which has the smallest values of L because it is the most unstable mixture with the greatest propensity for hotspot formation, is most likely to undergo DDT under the same initial and boundary conditions. These conclusions agree well with the various characteristics of L obtained for larger tubes.

On “new methodologies of processing” of results calculated for rotating detonation engines and of results of force measurements for ramjet models tested in supersonic test benches

On “new methodologies of processing” of results calculated for rotating detonation engines and of results of force measurements for ramjet models tested in supersonic test benches

Numerical investigations of the nozzle performance for a rocket-based rotating detonation engine with film cooling

This study numerically investigates the nozzle performance of a rotating detonation engine with film cooling. The premixed stoichiometric hydrogen/air mixture serves as the reactant, and the pure air is pumped into the film cooling holes. The diverging section of the nozzle is designed based on the characteristic method and the maximum thrust theory. The spike truncation ratio is optimized to ΔLspike = 40% Lcowl, reducing the weight and maintaining a high axial thrust coefficient of 0.9456. In contrast with the steady condition, the flow loss of the practically transient condition mainly comes from three parts. The first is the pressure loss of the continuously sweeping shock waves, the second is the viscous effect of the flow passage, and the third is the energy dissipation of the recirculation zone in the base region. When the film cooling strategies are introduced on the inner and outer walls, the hole type a/b < 1.0/1.0 is preferred thanks to a more considerable coolant coverage induced by more violent interactions between the radial jet and the mainstream. The cooling effects and the aerodynamic performance can be further balanced within a moderate injection velocity. The final optimized nozzle configuration with ΔLspike = 40% Lcowl, a/b = 1.0/2.0, and Vsec = 100 m/s corresponds to an axial thrust coefficient of 0.9552, a discharge coefficient of 0.9344, a pressure gain of -16.48%, and a root mean square deviation of the flow deflection angle of 14.42°.

Characteristics of rotating detonation wave fueled by liquid kerosene with increasing equivalence ratios

To investigate the characteristics of rotating detonation waves fueled by liquid kerosene with increasing equivalence ratios, the experiments were performed through increasing the supply of liquid kerosene under three different mass flow rates of oxidizer. Various mode transitions were realized, including transition from unstable detonation to stable single-wave mode, transition from unstable detonation to stable double-wave mode and transition from double-wave mode to single-wave & double-wave hybrid mode. The results demonstrate that the increase of equivalence ratios leads to the improvement of mixture reactivity under lean-fuel limit, which enhances the propagation stability and pressure peak of rotating detonation waves, and then results in the transition from unstable detonation to a stable propagation mode. In addition, when the mass flow rate of oxidizer is approximately 1100 g/s, the transition from double-wave mode to single-wave & double-wave hybrid mode is observed by increasing the equivalence ratio from 0.6 to 0.93. It is concluded that the mutation of dynamic balance between the RDW and injection plenum caused by the increasing of equivalence ratio is one of the reasons inducing the mode transition. This work illustrates the feasibility of modulating the propagation mode in rotating detonation waves fueled by liquid kerosene, and promotes the application of rotating detonation engine.

Thrust Performance of Converging Rotating Detonation Engine Compared with Steady Rocket Engine

Rotating detonation engines (RDEs) have been actively researched around the world for application to next-generation aerospace propulsion systems because detonation combustion has theoretically higher thermal efficiency than conventional combustion. Moreover, because cylindrical RDEs have simpler combustors, further miniaturization of conventional combustors is expected. Therefore, in this study, with the aim of applying RDEs to space propulsion systems, a cylindrical RDE with a converging–diverging nozzle was manufactured; the combustor length was changed to 0, 10, 30, 50, and 200 mm; and the thrust performance and combustion mode with the different combustor lengths were compared. As a result, four combustion modes were confirmed. Detonation combustion occurred with a combustor length of : that is, a converging rotating detonation engine. The thrust performance of this engine was 94 to 100% of the theoretical rocket thrust performance, which is equivalent to the thrust performance of conventional rocket combustion generated at . This study shows that detonation combustion can significantly reduce engine weight while maintaining thrust performance.

Applying a finite difference scheme to the flow field analysis of rotating detonation engines using multiple fuels

The objective of this study is to apply the finite difference method (FDM) scheme to generate profiles of velocity, vorticity, pressure, temperature, and density of the flow field inside the chamber of a rotating detonation engine. In addition, this study discusses the Kuo's algorithm used in the computation of Chapman-Jouguet (CJ) variables. A new computational algorithm is introduced to compute the CJ variables, the shock, and the post-shock conditions. Five different fuels are tested, H2–O2, H2-air, CH4-air, C2H6-air, and C2H2-air. The results show good performance for the FDM scheme, and the new algorithm is adopted in this study to perform all CJ calculations. The new algorithm's results are compared with the experimental data available for the tested fuels and show close results. Another important result is that the highest vorticity forms along the slip line of the combustion chamber, then it gradually shrinks along the path of the slip line.

Experimental research on solid fuel pre-combustion rotating detonation engine

The solid-fuel rotating detonation engine with gas generator is experimentally studied in this paper. This study solved the problems of ignition and stable detonation of pre-combustion gasification injection. The feasibility of solid pre-combustion rotating detonation is verified. A set of gas generators was designed and ignition tests were carried out. The detonation wave velocity is about 2625 m/s in the stable stage, which is about 20% higher than that generated by hydrogen/oxygen detonation. A maximum thrust of 69 N is obtained. Stable detonation pressure signal was obtained in our experiments. It is found that the detonation wave has experienced four stages in the formation process: deflagration, deflagration to detonation transition (DDT) process, galloping detonation, and stable detonation. In the stable detonation stage, coexistence of detonation waves with different strength is identified by high-speed photography images.

Supersonic combustion heat flux in a rotating detonation engine

Supersonic combustion heat flux in a rotating detonation engine

Deflagration-to-Detonation Transition in a Semi-Confined Slit Combustor Filled with Nitrogen Diluted Ethylene-Oxygen Mixture

The conditions for the mild initiation of the detonation of homogeneous stoichiometric ethylene-oxygen mixtures diluted with nitrogen up to ~40%vol. in a planar semi-confined slit-type combustor with a slit 5.0 ± 0.4 mm wide, simulating the annular combustor of a Rotating Detonation Engine (RDE), are determined experimentally using self-luminous high-speed video recording and pressure measurements. To ensure the mild detonation initiation, the fuel mixture in the RDE combustor must be ignited upon reaching a certain limiting (minimal) fill with the mixture and the arising flame must be transformed to a detonation via deflagration-to-detonation transition (DDT). Thus, for mild detonation initiation in a C2H4 + 3O2 mixture filling the slit, the height of the mixture layer must exceed the slit width by approximately 10 times (~50 mm), and for the C2H4 + 3(O2 + 2/5 N2) mixture, by approximately 60 times. The limiting height of the mixture layer required for DDT exhibits a sharp increase at a nitrogen-to-oxygen mole ratio above 0.25. Compared to the height of the detonation waves continuously rotating in the RDE combustor in the steady-state operation mode, for a mild start of the RDE, the fill of the combustor with the explosive mixture to a height of at least four times more is required.

Development of a Liquid-Propellant Rocket Powered by a Rotating Detonation Engine

This paper presents work on a demonstrator rocket powered by a rotating detonation engine (RDE) fed by liquid propellants. Key to the success of the project was the development of a suitably efficient and lightweight RDE. After selecting the propellants: nitrous oxide () and propane (), laboratory tests were carried out on RDEs with different combustion chamber geometries. The conical geometry showed the best performance, and it was selected for further development; on this basis, the target RDE was designed and built in two versions: uncooled and regeneratively cooled. These were tested in laboratory conditions, but only the regeneratively cooled engine met the requirements. After successful laboratory testing of the engine, a demonstration rocket was designed and built. Static tests made in a vertical position confirmed that the rocket had sufficient operational performance (net thrust and total impulse). Ultimately, the liquid-propellant RDE-powered rocket made a successful flight. To the authors’ knowledge, it was the first liquid-fuel rocket in the world to take off and fly exclusively by means of RDE propulsion.

Adaptive operating mode switching process in rotating detonation engines

In this study, the adaptive mode switching process of rotating detonation waves (RDWs) in response to the change of inlet conditions is numerically explored and analytically explained. Once a quasi-steady rotating detonation flow field is obtained, the inlet total pressure of the reservoir is decreased and maintained at different levels for a certain period of time in order to observe the feedback on the RDWs. It shows that instead of being quenched immediately, the RDWs can adapt to the change of inlet conditions by a mode switching adjustment; it decreases the number of detonation waves and begins to run in a new operating mode, a process that is found to occur swiftly and is corroborated by experimental observations; moreover, an analytical analysis indicates that for a stable RDW flow field, the characteristic parameters such as the number and height of detonation waves are related to the combustor geometry and inlet conditions, which determine whether the RDWs are required to switch into a new operating mode as a means to ensure stability. Additionally, the results indicate that the rotating detonation engine (RDE) can work in different operating modes at the same mass flow rate, or conversely, the RDE can remain in the same operating mode even if the mass flow rate varies.

Experimental research on the performance of hollow and annular rotating detonation engines with nozzles

• The performance of hollow and annular RDEs with nozzles are compared. • The hollow RDE can reduce the injection pressure loss. • The hollow chamber brings a higher performance loss. Rotating detonation engines (RDEs), which are hoped to achieve significant performance improvements in propulsion systems, have received much attention in recent years. The annular RDE and hollow RDE are two of the major configurations of which the structure is very similar to each other. However, there has been a lack of experimental research discussing the differences in their performance. In the present study, experiments on hollow and annular RDEs are conducted. Four configurations in total are tested at ambient conditions: hollow and annular combustion chambers with two exit throat areas. The gaseous methane and gaseous oxygen are used as propellants. The results reveal that the c*-efficiency of the two hollow combustion chamber configurations is on average 0.7 and 0.79, which is lower than the 0.75 and 0.84 of annular chambers. However, it is exhibited that the hollow RDE can ensure high detonation efficiency at a lower injection pressure ratio. From this perspective, the hollow RDE can be considered a more promising configuration for practical application because it provides a new approach to reducing injection pressure loss. In addition, the analysis of nozzle indicates that the Laval nozzles overperform the aerospike nozzles, which goes against previous studies. The mechanism that accounts for this phenomenon needs further investigation.

A flame technique to isolate the detonation/product interface relevant to rotating detonation engines

A novel experimental technique is proposed to study the detonation propagation in a layer of non-reacted gas weakly confined by combustion products. This problem is relevant to rotating detonation engines, where transverse detonations are confined by products of a previous rotation cycle, and other applications such as industrial safety. The experimental technique utilizes a flame ignited along the top wall in a long channel. The preferential growth of the flame along the long direction of the channel creates a finger flame and permits to create a narrow layer of unburned gas. A detonation ignited outside of this layer then propagates through the layer. This permits to conduct accurate observations of the detonation interaction with the inert gas and determine the boundary condition of the interaction. The present paper provides a proof-of-concept demonstration of the technique in a 3.4 m by 0.2 m channel, in which long finger flames were observed in ethylene-oxygen mixtures. The flame is visualized by high-speed direct luminosity over its entire travel, coupled with pressure measurements. A direct simulation of the flame growth served to supplement the experiments and evaluate the role of the induced flow by the flame growth, which gives rise to a non-uniform velocity distribution along the channel length. Detonation experiments were also performed at various layer heights in order to establish the details of the interaction. The structure was visualized using high speed Schlieren video. It was found that an inert shock always runs ahead of the detonation wave, which gives rise to a unique double shock reflection interaction.

Research of Rotating Detonation Engine at Air Systems Research Center, Ministry of Defense

Research of Rotating Detonation Engine at Air Systems Research Center, Ministry of Defense

Numerical Simulation of Flow-induced Vibration of a Rotating Detonation Engine

Numerical Simulation of Flow-induced Vibration of a Rotating Detonation Engine

Experimental Proof of Concept of a Noncircular Rotating Detonation Engine (RDE) for Propulsion Applications

A noncircular engine cross-section could provide great flexibility in the integration of propulsion into the airframe. In this work, a tri-arc RDE was constructed and tested as an example of noncircular cross-sectioned RDE. The operational characteristics of detonation wave propagation and thrust performance were investigated and compared with an equivalent circular RDE under the same operating conditions. High-speed camera images, short-time Fourier transform (STFT), and fast Fourier transform (FFT) were used for the investigation. The tri-arc RDE showed very similar characteristics to the circular RDE but exhibited slightly better stability and propulsion performance than the circular RDE. We consider that repeated curvature changes positively affect the stability of detonation wave propagation. The experimental data show contradicting results from the numerical analysis with a homogeneous mixture assumption in which the detonation pressures at the convex corner were greater than those at the concave corner. It is reasoned that the tri-arc injector design provides a non-uniform mixture composition, resulting in a strong detonation at the convex corner. Overall, the noncircular RDE of a tri-arc shaped cross-section is demonstrated, one which performs slightly better than an ordinary circular-shaped RDE both in detonation stability and performance.

Variably Premixed Rotating Detonation Engine for Evaluation of Detonation Cycle Dynamics

A variably premixed rotating detonation engine using gaseous hydrogen and air reactants is introduced to enable investigation of key cycle processes while varying the homogeneity of the reactant inlet conditions. Two chamber configurations are investigated, the first with reactants filling the entire span of a straight annular channel and the second with a slightly larger channel and a backward-facing step. The first configuration permits both premixed and non-premixed fuel injection, enabling mixing quality modulation. The second configuration is operated only at fully premixed conditions. Operating modes and detonation wave speeds are characterized using exhaust-plume imaging, while the chamber heat release field is captured by transverse imaging through a transparent outer body. Tests using the first configuration were characterized by unstable counterpropagating modes with low detonation wave speeds regardless of the state of premixing, while the second configuration rendered single-wave behavior with wave speeds up to 86% of the Chapman–Jouguet velocity. Comparisons with a simple computational fluid dynamics model of the second configuration indicate that reactant preheating significantly influences the detonation wave topology, highlighting the potential utility of the test platform for isolating key physics associated with the effects of reactant premixing, preheating, and chamber geometry on rotating detonation engine operation.

Effects of inlet and secondary flow conditions on the flow field of rotating detonation engines with film cooling

A three-dimensional simulation of the rotating detonation engine (RDE) with film cooling is conducted. The aim of this study is to analyze the fluid dynamics and heat transfer of the detonation flow field under the influence of cooling flow from the film holes. Results suggest that when the rotating detonation wave sweeps the film holes, the shape of the wave structure will deform, and the detonation products will invade and block the outflow from the film holes; however, this only occurs temporarily. The structure of the detonation wave will quickly restore to its stable form and, meanwhile, the cooling flow also recovers rapidly and provides adequate protected area on the wall surface and effective thermal protection time in a full propagation cycle of the detonation wave. A parametric analysis indicates that the effective outflow time improves with the increase of the mass flow rate of the cooling flow; on the other hand, the cooling efficiency is more significant downstream from the inlet of the combustor to the outlet. In addition, the thrust and specific impulse of the RDE are also examined under the influence of film cooling.

Large Eddy Simulation of Combustion for High-Speed Airbreathing Engines

Large Eddy Simulation (LES) has rapidly developed into a powerful computational methodology for fluid dynamic studies, between Reynolds-Averaged Navier–Stokes (RANS) and Direct Numerical Simulation (DNS) in both accuracy and cost. High-speed combustion applications, such as ramjets, scramjets, dual-mode ramjets, and rotating detonation engines, are promising propulsion systems, but also challenging to analyze and develop. In this paper, the building blocks needed to perform LES of high-speed combustion are reviewed. Modelling of the unresolved, subgrid terms in the filtered LES equations is highlighted. The main families of combustion models are presented, focusing on finite-rate chemistry models. The density-based finite volume method and the reaction mechanisms commonly employed in LES of high-speed H2-air combustion are briefly reviewed. Three high-speed combustor applications are presented: an experiment of supersonic flame stabilization behind a bluff body, a direct connect facility experiment as a transition case from ramjet to scramjet operation mode, and the STRATOFLY MR3 Small-Scale Flight Experiment. Several combinations of turbulence and combustion models are compared. Comparisons with experiments are also provided when available. Overall, the results show good agreement with experimental data (e.g., shock train, mixing, wall heat flux, transition from ramjet to scramjet operation mode).