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In order to realize stable propagation of the ethylene-air rotating detonation wave with less velocity deficit in the annular combustion chamber, a series of experiments have been carried out while changing the combustor width (i.e. 10 mm, 15 mm, 25 mm, and a hollow combustor), mass flow rate of air, and equivalence ratio. The air and ethylene are supplied into the combustor through a slot-orifice impinging injection scheme, where air is injected through a slot with a height of 0.5 mm and ethylene is supplied by 160 orifices with a diameter of 0.35 mm. Experimental results indicate that four propagation modes have been obtained, i.e. the single wave mode, the sawtooth wave mode, the periodic oscillation mode, and the deflagration mode. The sawtooth wave mode and the periodic oscillation mode are considered unstable modes. In the single wave mode, the average propagation velocity of rotating detonations in the annular combustion chamber with a combustor width of 15 mm is between 1121 and 1215 m/s. The minimum velocity deficit is 21.9%, and the relative standard deviation of the propagation velocity is less than 5%, which is much lower than the experimental results of other studies utilizing ethylene and air in the annular combustors. In addition, the reasons may lead to the failure of producing rotating detonations have been analyzed. The effects of the chemical reactivity on the generation of rotating detonation waves have been demonstrated by installing a convergent nozzle at the exit of the combustor. [ABSTRACT FROM AUTHOR]

The numerical and experimental research of rotating detonation engines is usually conducted in ground environmental conditions. Although many breakthroughs have been made, there are still deficiencies in understanding engines operating in real flight conditions. In order to reveal the effect of engine size and initial kerosene droplet size on the rotating detonation engines in flight conditions, the Eulerian-Lagrangian model is adopted, and a series of two-phase kerosene/air rotating detonation cases are simulated in the conditions of Mach 5 and 24 km altitude. When 5 μm and 10 μm droplets are adopted, the RDE behaves like gaseous rotating detonations with clear cellular structures. When the 20 μm and 30 μm droplets are adopted, the rotating detonation waves tend to be divided into two layers and form the λ-shaped shock structure. The results indicate that droplet size and engine size influence detonation wave propagation and flow field mainly through droplet heat absorption and evaporation height. Increasing the engine sizes can promote the two-phase rotating detonation and broaden the initial diameter range capable of obtaining rotating detonation waves. However, as droplet size increases, the fresh mixture layer becomes stratified, with an upper layer predominantly comprising fuel vapor and a lower layer of fuel droplets, which results in an λ-shaped shock structure at the detonation front. The flight operation condition brings in differences in features such as detonation wave height and velocity deficit. The detonation height in flight conditions is higher than that in ground conditions. The most velocity deficit in flight conditions observed in our study is below 15 %, which is lower than the results (22.5 %) in ground operating conditions. These results indicate that the optimal design of the rotating detonation engine must consider the real operation conditions and the effect of engine sizes and droplet sizes. • The propagation characteristics of two-phase rotating detonation waves in flight operation conditions are revealed. • The effects of engine sizes and droplet sizes are analyzed. • Detonation height and propagation velocity are increased in flight operation conditions, compared with the ground test. • The rotating detonation engine design should be based on the real operation conditions. [ABSTRACT FROM AUTHOR]

Through the method of three-dimensional numerical simulation, the rotating detonation combustion characteristics of kerosene-fueled scramjets in the range of Ma＝3~7 were studied. In the flight condition of Ma=3, due to the poor effect of fuel atomization and evaporation, the detonation combustion of kerosene fuel cannot be realized. In the flight conditions of Ma＝4, 5, 6, with the increase of Mach number, the number of wave heads increases gradually, which are single-wave, three-wave, and five-wave modes, respectively. However, the propagation speed gradually decreases. In the scramjet mode, the liquid fuel has a good effect of atomization and evaporation. Nevertheless, kerosene vapor remains in the flow field to varying degrees and is discharged from the combustion chamber without participating in the reaction. In the flight condition of Ma＝7, the flow field will burn in a stationary detonation mode because the incoming flow is close to CJ velocity.

Experimental studies were conducted on the initiation characteristics of rotating detonation waves (RDW) using auto-initiation and pre-detonation methods with high-temperature hydrogen gas. The effects of the two initiation methods on the formation and propagation characteristics of rotating detonation waves were compared. The experimental results showed that the formation process of the RDW was different when using auto-initiation and pre-detonation initiation. However, the parameters of the stable propagation stage of the RDW were similar. The time required for the formation of the RDW by pre-detonation was relatively short. As the equivalence ratio increased, the rotating detonation chamber (RDC) exhibited five propagation modes: two-co rotating detonation waves (TRDW), two-counter rotating detonation waves (TCRDW), a mixed mode, intermittent single rotating detonation wave (ISRDW), and single rotating detonation wave (SRDW). It is worth noting that both initiation methods were capable of achieving rotating detonation within the equivalence ratio range of 0.68–1.96. The pre-detonation initiation can expand the range of stable SRDW modes and has some advantages in terms of RDW propagation stability. • The characteristics of RDW exhibited similarities between two initiation methods. • The time for the formation of the detonation wave by pre-detonation was short. • Detonation be achieved by both the initiation ways with the ER of 0.68–1.96. • Pre-detonation initiation can expand the range of stable single-wave modes. [ABSTRACT FROM AUTHOR]

In this study, a numerical investigation based on the Eulerian-Lagrangian model is conducted to explore a rotating detonation engine (RDE) fueled by liquid ethanol. The focus is on examining the characteristic phenomena of the two-phase rotating detonation wave (RDW) caused by droplet evaporation and varying inlet conditions. To enhance the evaporation of liquid fuel, pre-heated air is used, and both liquid and pre-vaporized ethanol are simultaneously injected. The distribution of ethanol droplets reveals an initial concentration near the injection surface and accumulation in the fuel-refill zone. Here, liquid droplets gradually evaporate after absorbing latent heat from the surrounding gas. The subsequent interactions between the evaporating droplets and the RDW vary with the droplet size. For droplets with diameters of d 0 = 5–15 μ m , after being swept by the RDW, a secondary evaporation process occurs, leading to an enlargement of the width of the reaction zone. However, the chemical reactions still predominantly take place in close proximity to the detonation front. As d 0 further increases, droplet evaporation persists in the post-detonation expansion zone over a long distance until the remaining droplets are fully evaporated and eventually burned in this region. The study also analyzes the extinction of rotating detonations and the emergence of new detonation waves resulting from local explosions and consequent shock collisions. It is demonstrated that variations in the diameter of injected droplets and inlet temperature can lead to different operating modes with varying numbers of RDWs. • Rotating detonation engines (RDEs) fueled by liquid ethanol are investigated by a numerical study. • A multi-stage interaction process between the evaporating droplets and the rotating detonation wave (RDW) is revealed. • The effects of inlet conditions on the phenomenon of wave multiplicity are analyzed. [ABSTRACT FROM AUTHOR]

Rotating detonation experiments with 220 mm outer diameter of combustor are carried out. The experiment used aviation kerosene (RP-3) as fuel and hot air as oxidant. Under the condition that the air flow at the inlet of the combustor is kept constant, when the blockage ratio is 50%, 65%, 75% and 90% respectively, the detonation is successfully initiated and works stably in the experiment. Temperature rise method was used to calculate the combustion efficiency of the rotating detonation combustor outlet. Through sampling, the gas at the rotating detonation combustor outlet was analyzed, and the complete combustion rate based on carbon atoms (CCR) was obtained. There is an optimal blockage ratio and equivalent ratio so that the combustion efficiency and CCR of the rotating detonation combustor reach the maximum value during the experiment. The mode of detonation wave is mainly double-wave collision mode. If the blockage ratio is too large, the propagation mode of detonation wave will change from two wave collision mode to axial pulse mode. At this time, the combustion efficiency and CCR will be greatly reduced. When the blockage ratio is constant, increasing the air flow rate is helpful to improve the combustion efficiency and CCR, and broad the working range of the rotating detonation combustor. The change of air flow rate at combustor inlet has little influence on the wave velocity of detonation wave, and the propagation mode of detonation wave is still two-wave collision mode. [ABSTRACT FROM AUTHOR]

In this study, the operating modes of the rotating detonation engine (RDE) under various inlet conditions were investigated numerically. It was shown that after the sudden reduction of the inlet total pressure, the resulting decrease of the mass flow rate would require the rotating detonation wave (RDW) to adjust its front height and propagation velocity so that it could maintain self-sustained propagation. However, if the inlet total pressure dropped significantly, the number of RDWs in the flow field would decline to match the new inlet conditions. This would lead to an adaptive mode switching process. The approximate matching relation between the characteristic parameters of the RDWs (wave number, propagation velocity, and front height), the geometry of the combustor, and the inlet conditions (mass flow rate) was verified to explain the operating mode switching process and the underlying mechanisms.

In this study, numerical simulations based on an Eulerian-Lagrangian framework are conducted to investigate the liquid-fueled rotating detonation engine (RDE) with heterogeneous kerosene-hydrogen-air mixtures. The hydrogen addition is implemented for combustion enhancement of the liquid kerosene and helps to ignite and achieve a self-sustained two-phase rotating detonation wave (RDW). The effects of droplet evaporation at various initial droplet sizes and kerosene mass flow rates on the structure and propagation of the two-phase RDW are analyzed. Results suggest that with smaller droplet sizes, the structure of the RDW is analogous to that of a gaseous RDW, and a comparison with experimental data suggests that the estimated detonation speed and thrust performance (fuel-based specific impulse) are within the reasonable ranges. However, as the droplet size or the mass flow rate of kerosene increases, the two-phase RDW exhibits characteristic features such as the dual-front laminated structure, micro-explosions and secondary transverse waves. • Rotating detonation engines (RDEs) with a heterogenous kerosene-hydrogen-air mixture are investigated. • Numerical simulations based on an Eulerian-Lagrangian two-way coupled framework are conducted. • A dual-front rotating detonation wave (RDW) caused by droplet evaporation is explained. • Micro-explosions and secondary RDWs are captured at higher kerosene mass flow rates. [ABSTRACT FROM AUTHOR]

In this paper, the effects of the Tesla valve intake structure with or without bypass manifold on suppressing pressure feedback and combustion product backflow in rotating detonation engine fueled by kerosene and enriched air with 29% oxygen content are investigated, with the configuration of convergent–divergent intake without bypass manifold being used for comparison. The working range, detonation wave propagation characteristics, and propulsion performance for the three configurations are evaluated. It is found that the Tesla valve intake structure can effectively improve the suppression of pressure feedback and combustion product backflow through the diversion function of the Tesla valve bypass channel and that the addition of the bypass manifold further improves the suppression of pressure feedback. The Tesla valve intake with bypass manifold configuration has the best ability to suppress pressure feedback, with a minimum value of pressure feedback percentage of about 6.5% and a minimum value of air plenum pressure fluctuation percentage of about 6.7%. By comparison, the minimum value of pressure feedback percentage of the convergent–divergent intake without bypass manifold configuration is about 29.8%, and its minimum value of pressure feedback percentage is about 7.8%. The Tesla valve intake without bypass manifold configuration is most effective in suppressing combustion product backflow, with the effective inlet area ratio of about 95.7%. Under the same conditions, the minimum effective inlet area ratio of the convergent–divergent intake without bypass manifold configuration is about 90.2%. Ultimately, optimizing the engine structure leads to a broader working range and higher propulsion performance. The Tesla valve intake structure broadens the working range by suppressing the generation of the longitudinal pulsed detonation mode. When the bypass manifold is incorporated, some combustible reactants directly enter the rear of the combustion chamber from the head of the chamber, and the working range is reduced. Compared with the convergent–divergent intake without bypass manifold configuration, the working range of the Tesla valve intake with bypass manifold configuration is widened by 383.3%, and that of the without bypass manifold configuration by 700%. Both the Tesla valve intake and bypass manifold increase the strength of detonation waves and improve propulsion performance. The Tesla valve intake with bypass manifold configuration has the highest specific thrust, at about 90.1 N s3/(kg·m). This is increased by an average of 10% compared with the Tesla valve intake without bypass manifold configuration, and by an average of 30.1% compared with the convergent–divergent intake configuration. • Tesla valve inlet and bypass manifold are used to suppress feedback. • Two quantitative evaluation parameters of pressure feedback are established. • The degree of combustion product backflow is evaluated. • The optimized structure can effectively improve the propulsion performance. [ABSTRACT FROM AUTHOR]

Utilization of liquid fuels is crucial to enabling commercialization of rotating detonation engines (RDE) in the near future. In this study, Eulerian-Lagrangian simulations are conducted for rotating detonative combustion with dilute n -heptane sprays and preheated air. Two-dimensional flattened configuration is used and a skeletal chemical mechanism with 44 species and 112 elementary reactions for n -heptane combustion is adopted. The flow structure, droplet distribution, and thermochemical parameters in the refill zone are first analyzed. It is shown that the mixture in the refill zone is heterogeneous, including evaporating droplets, vapor, and air. When the total temperature is below 950 K, the average equivalence ratio increases with the total temperature. When it is higher than 950 K, the average equivalence ratio is almost constant. Subsequently, the chemical explosive mode analysis is applied to identify the controlling reactions and dominant combustion modes in the fuel refill zone and reaction fronts. Results demonstrate that the initiation reaction (R104: n -C 7 H 16 + O 2 → 2-C 7 H 15 + HO 2) and low-temperature reaction (R107: RO 2 → R'O 2 H) are dominant in the upstream and downstream of the refill zone, respectively. The intermediate species from low-temperature chemistry, R'O 2 H, is found to be important for the chemical explosive mode in the undetonated mixture. The influence of species diffusion and dispersed droplets is further analyzed. Results show that vapor autoignition facilitated by droplet evaporation occurs in the refill zone. Finally, the effects of the air total temperature on the detonation propagation speed and RDE propulsion performance are investigated. It is found that the detonation propagation speed and specific impulse increase with air total temperature. The total pressure ratio first increases and then decreases as the air total temperature increases. Moreover, when the total temperature of the preheated air is above 1,300 K, the effects of low-temperature chemistry are negligible. [ABSTRACT FROM AUTHOR]

A three-dimensional upwind conservation element and solution element method (CESE) in cylindrical coordinates is first developed to effectively solve the unsteady reactive Euler equations governing a hydrogen–air rotating detonation engine (RDE) with coaxial structures. The effects of the annular width on the structure of the detonation front and the relationship between the thrust and mass flow rate are then investigated. Additionally, RDEs with various injection conditions are systematically analyzed regarding flow patterns and propulsion performance. The results reveal a positive correlation between the specific impulse and the area ratio of the injection slot to the head-end wall. Nevertheless, the specific impulse shows minimal dependence on the injector slot's location when the area ratio is constant. Ultimately, it is concluded that the area ratio between the injector and the head-end wall is critical in determining the loss of specific impulse. [ABSTRACT FROM AUTHOR]

In order to investigate the effects of gas mixture components on the combustion characteristics of rotating detonation wave, two-dimensional simulation is presented to simulate the propagation process of rotating detonation wave with different methane conversions. The results indicate that there are five propagation modes of rotating detonation wave with different components: single-wave mode, single wave with counter-rotating components mode, double-waves mode, triple-waves mode and quadruple-waves mode. The detonation wave propagates along the forward direction in all five modes. With the increase of methane conversion, multi-wave mode appears in the combustion chamber. The fuel component has a great influence on the heat release ratio of detonation combustion. The velocity of detonation wave decreases with the increase of methane conversion. With the increase of methane conversion, the chemical reaction rate gradually increases, which leads to the intensification of chemical reaction on the deflagration surface. The reaction on the deflagration surface develops to the unburned fuel zone, which eventually leads to the formation of compression waves and shock waves in the fuel refill zone. When the shock wave sweeps through the fresh premixed gas, the reactant is compressed to form a detonation point and then ignite the fuel. A new detonation wave is finally formed. The total pressure ratio decreases with the increasing methane conversion, and the uniformity of the total pressure of outlet decreases with increasing methane conversion. • Component change of multicomponent mixture dictate propagation mode. • The influence of component change on detonation wave velocity and pressure is revealed. • The effect of deflagration surface reaction on the characteristics of detonation wave is analyzed. • Pressure ratio of combustion chamber with mixture of different components is compared. [ABSTRACT FROM AUTHOR]

Large-scale rotating detonation experiments with a 500 mm outer diameter combustor are carried out in this study. The experiment used aviation kerosene (RP-3) as fuel and hot air as oxidant. Detonation is successfully initiated between the equivalent ratio of 0.85 and 1.20. The engine works steadily for 2 s and 3 s. The detonation wave modes are two-wave collision. The average speed of detonation wave is between 920 m/s and 980 m/s. The detonation wave velocity increases with the increase in equivalence ratio, and when the equivalence ratio exceeds 1.0, the increasing trend of detonation wave slows down. Through sampling and analysis of detonation gas, it is found that the uneven distribution of fuel supply pressure in large-scale rotating detonation experiment results in inconsistent local equivalent ratio and uneven combustion. The complete combustion rate based on carbon atoms is used to evaluate the combustion completeness of detonation. With the increase in equivalent ratio, the combustion changes from deflagration to detonation, and the detonation combustion completeness gets better. The combustion completeness of detonation wave is the best near the appropriate chemical equivalent ratio. Once the equivalent ratio exceeds 1.0, then oil-rich combustion occurs and the strength of the detonation wave becomes weaker. [ABSTRACT FROM AUTHOR]

An experimental approach for estimating the detonation cell size for liquid hypergolic propellants is presented and applied for monomethylhydrazine (MMH) and a MON-3 variant of nitrogen tetroxide (NTO) in a non-premixed combustor. The method utilizes a correlation between cell size and reactant fill height in an annular combustor geometry. Reactant fill height is inferred using a control-volume analysis to relate the geometry of the reactant fill zone to the propellant flow rate, detonation wave-speeds and number, and annular gap size. High-speed videography is used to measure the number of waves and their speed over a range of quasi-steady continuous detonation conditions. The detonation criteria is also proven valid in the transient conditions as the decrease of the fill height below a critical value identified in this work matches modal transitions with decreasing number of waves. A power law relating the cell width with induction length further supports the validity of the present technique down to cell sizes of 1–10 µm, a scale not practically resolvable with conventional methods. [ABSTRACT FROM AUTHOR]

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, H 2 –O 2 , H 2 -air, CH 4 -air, C 2 H 6 -air, and C 2 H 2 -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. • Presenting a new algorithm to compute the Chapman-Jouguet variables. • Each detonation wave can achieve a different Chapman-Jouguet pressure. • CFD test cases fuels used: H 2 –O 2 , H 2 -air, CH 4 -air, C 2 H 6 -air, and C 2 H 2 -air. • The RDE detonation reactants mixture height varies with fuel type. • Vorticity is highest at the slip line and shrinks downstream the slip line. [ABSTRACT FROM AUTHOR]

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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. • Numerical simulations are conducted to investigate the operating modes of rotating detonation engines (RDEs). • The RDE can adapt to the changes of inlet conditions and switch its operating mode when necessary. • A quasi-steady analysis of the rotating detention waves is provided to explain the adaptive mode switching process. [ABSTRACT FROM AUTHOR]

An ongoing rotating detonation rocket engine program is investigating the influence of combustor annulus radii on RDRE operating characteristics with flat-faced impinging injectors. To facilitate the isolation of all but the radius of curvature effects in the experiments, the annular gap was kept constant at 5 mm in combustors having either 25-mm or 51-mm outer diameter. The mixing processes were kept similar by utilizing injectors with the same net injector-to-annular gap area ratio (AR = 0.11), same radial separation distance of the orifices, and same center-of-gap impingement distance from the front-end wall. The wave dynamics, plenum pressure, and axial pressure profiles in these RDREs were compared over the mass flux and equivalence ratio ranges of 80 - 400 kg/s/m 2 and 0.26 - 2.6, respectively, with gaseous methane–oxygen propellant. Experiments showed that stable one-wave operation would occur in the 25-mm RDRE at most mass fluxes where stable two-wave operation was established in the 51-mm RDRE. Stable one-wave operation with a single counter-rotating wave was maintained in the 51-mm RDRE at mass fluxes of 240 kg/s/m 2 and below. Under these fueling conditions in the 25-mm RDRE, a counter-rotating wave also appeared while it operated with a single dominant wave. The wave spin speeds were typically 20–40% less than the Chapman–Jouguet detonation speed of the propellant and depended only on mass flux and wave number rather than the annulus diameter. [ABSTRACT FROM AUTHOR]

Rotating detonation propulsion technologies have the potential to create highly efficient engines in a small form factor. However, the detonation dynamics and complex flowfields inside the combustion chamber are greatly dependent on geometry; in particular, the downstream nozzle design affects dynamics inside the combustion chamber. In this work, three-dimensional large eddy simulations of a gaseous methane–oxygen rotating detonation rocket engine are presented for two geometries. The geometries match experimental tests previously conducted at the Air Force Research Laboratory and are chosen to compare engine operation with and without a converging–diverging nozzle. It is shown that flow in the unconstricted chamber exceeds Mach 1 behind the generated oblique shock structure, but that the addition of a 4.4 ∘ converging section results in supersonic flow existing only in the diverging section of the nozzle. The formation enthalpy of the flow is calculated inside the chamber and demonstrates that the difference in pressures and detonation structures associated with the chamber area constriction do not result in a significant change in energy released through combustion. [ABSTRACT FROM AUTHOR]

For high-Mach-number incoming flow circumstances, a rotating detonation ramjet engine configuration is proposed in this research. By installing supporting blocks at the rear of the combustor, this configuration achieves continuous rotating detonation operation. Based on the Comparison of the flow structures obtained from the engine configuration with and without the supporting block before and after detonation ignition respectively, we obtain the intrinsic mechanism of detonation wave's propagation and re-initiation under the action of the supporting block. The supporting block creates a deflagration wave that is almost stationary before detonation ignition. In the detonation-ignited state, the deflagration wave is continually formed and traveling upstream under the influence of the supporting block, which is analogous to the periodical before detonation ignition of a transverse wave structure. The dynamic deflagration wave will cause the incomplete reactants behind the detonation wave to burn as the intensity of the detonation wave decreases. As a result, the incident shock wave is transformed into a Mach stem to achieve the re-initiation of the detonation wave. • The continuous spin detonation with kerosene as fuel is realized at flight Mach number M = 6. • Flow characteristics in the before and after detonation ignition states with and without supporting blocks are compared. • The supporting block may effectively contribute to maintaining the continuous propagation of the detonation wave. • The detonation wave's mechanism of propagation and re-initiation is examined under the influence of the supporting block. [ABSTRACT FROM AUTHOR]

Propulsion systems based on the constant-pressure combustion process have reached maturity in terms of performance, which is close to its theoretical limit. Technological breakthroughs are needed to develop more efficient transportation systems that meet today’s demands for reduced environmental impact and increased performance. The Rotating Detonation Engine (RDE), a specific implementation of the detonation process, appears today as a promising candidate due to its high thermal efficiency, wide operating Mach range, short combustion time and, thus, high compactness. Following the first proofs of concept presented in the 1960s, the last decade has seen a significant increase in laboratory demonstrators with different fuels, injection techniques, operating conditions, dimensions and geometric configurations. Recently, two flight tests of rocket-type RDEs have been reported in Japan and Poland, supervized by Professors Kasahara (Nagoya University) and Wolanski (Warsaw University), respectively. Engineering approaches are now required to design industrial systems whose missions impose efficiency and reliability constraints. The latter may render ineffective the simplified solutions and configurations developed under laboratory conditions. This requires understanding the fundamentals of detonation dynamics relevant to the RDE and the interrelated optimizations of the device components. This article summarizes some of the authors’ experimental and numerical work on fundamental and applied issues now considered to affect, individually or in combination, the efficiency and reliability of the RDE. These are the structure of the detonation reaction zone, the detonation dynamics for rotating regimes, the injection configurations, the chamber geometry, and the integration constraints.

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.

Rotating detonation using ammonia as fuel may be a potential carbon free combustion technology for gas turbine. The detonation wave structure and flow field of a rotating detonation annular combustor are investigated by three-dimensional simulation with detailed chemistry of ammonia/hydrogen-air. The detonation properties, propagation mode, combustor performance and emission characteristics are studied by varying the equivalence ratios and hydrogen concentrations. Both the increases of the combustor pressure and the hydrogen concentration promote the chemical reaction rate of the ammonia burn and the detonation wave velocity gradually increases with increasing hydrogen proportion based on one-dimensional simulation. A stable single-rotating waves resulting in ammonia/hydrogen combustor are observed for a wide range of equivalent ratios only when the hydrogen concentration is at least 0.2. The steady run of the single rotating detonation had an optimal cycle efficiency when the hydrogen concentration is increased to a critical value of 0.3. NOx emissions are more dependent on equivalent ratios than hydrogen concentration in equivalence ratios ranging from 0.70 to 1.40. • Three-dimensional simulation of rotating detonations with detailed chemistry of NH4/H2-air. • The detonations propagate stably in premixed ammonia-oxygen-enriched air mixture with at least 20% hydrogen. • RDE has an optimal cycle efficiency when hydrogen concentration is at a critical value of 0.3. • NOx emissions are more dependent on equivalent ratios than hydrogen concentration. [ABSTRACT FROM AUTHOR]

The thermodynamic cycles of the detonation combustion could potentially deliver a performance increase of 20% beyond the conventional deflagration combustor such as gas turbines and ramjets, thus the detonative combustor is a prospective technology for high-efficiency distributed energy or power systems in the future. Therefore, the core process of rotating detonation combustion using hydrogen as a clean fuel was studied in this paper. A rotating combustor integrated with a supersonic inlet, an isolator, and a nozzle was targeted for the overall performance evaluation using a loose coupling method.To explore the overall performance of a rotating detonation engine (RDE) combustion model, the study proceeds via three main steps. (1) An axisymmetric semi-isentropic inlet is proposed for the RDE combustor model. The inlet consists of a leading-edge cone and an isentropic outward-turning compression surface. The method for quick design of the external-compression surface was developed and has been verified by the comparison of MOC analysis results with those of computational flow dynamics (CFD) simulations. (2) The combustible mixture gas, produced by the subsonic exiting flow of the isolator mixed with the injected hydrogen, is fed to the annular combustor around which a rotating detonation wave propagates. (3) The rotating detonation simulation for an ‘unwrapped’ annular combustor was performed and a stable moving detonation wave was achieved with the typical RDE flow-field pattern. The high efficiency of the RDE combustion mode has been demonstrated by the estimated fuel-based specific impulse of 5268 s. The loose coupling method for RDE’s performance prediction has been verified by the comparison of the calculated specific impulse with theoretical results published in open literature.

The rotating detonation gas turbine boasts several advantageous characteristics, including self-pressurization, minimal entropy increase, high thermal efficiency, and a high thrust-to-weight ratio. These features make it a promising candidate for the next generation of aerospace power devices. This study investigates the flow characteristics and performance attributes of the supersonic turbine stage coupled with a rotating detonation chamber based on different rotational speeds, blade ratio of rotor and stator, and hub radius, and uses DMD (Dynamic Mode Decomposition) to analyze the main modes of the supersonic turbine. The research identifies distinct wave modes in both the stator and rotor blades of the supersonic turbine. In the stator blades, the modes include waveless contact modes, opposite side λ -wave oblique shock wave modes, and derived modes from opposite side λ -wave oblique shock waves. In the rotor blades, the modes encompass non-separation modes (waveless modes, asymmetric dual λ -wave modes, oblique cut wave modes, and single-side λ -wave modes on the pressure surface), leading-edge stator excitation single separation bubble modes, saddle-type separation modes, and dual separation bubble modes. The efficiency and power of the supersonic turbine stage increase with the rotational speed. The turbine efficiency is highest when the blade ratio is 10:7. Under multi-wave mode conditions, detonation waves demonstrate adaptive self-regulation abilities, ultimately achieving equilibrium in speed, detonation wave height, and spacing. The dominant modes in the stator blades and combustion chamber are governed by detonation waves and slip lines, while in rotor blades, the primary mode is the flow-directional vortex mode governed by detonation waves and slip lines at a frequency lower than their propagation frequency. This research holds certain reference significance for understanding the internal flow characteristics of rotating detonation gas turbines, as well as discussing the patterns of efficiency, power, and load characteristics, thereby offering valuable insights for the practical application of rotating detonation gas turbines. • Supersonic turbine stage with different parameters coupled with RDC. • DMD mode characteristics of supersonic turbine coupled with RDC. • Unsteady modes of wave system of interaction between TGVs and combustion chamber. • Unsteady modes of shock wave system of turbine rotor blade. • Performance characteristics of supersonic turbine stage under different parameter conditions. [ABSTRACT FROM AUTHOR]

• RDC passage area has a significant impact on operating range of rotating detonation gas turbine. • The increase of RDC passage area makes the operating range of gas turbine move to the higher power. • Equivalence ratio of RDC replaces turbine efficiency as the greatest influence factor on cycle efficiency increment with different combustor passage areas. • Qualitative analysis based on sensitivity analysis of component parameters is further optimized to announced the comprehensive reason for change of cycle efficiency increment. Application of rotating detonation combustion can significantly enhance gas turbine performance, but the limited flow capacity caused by normally aspirated characteristic of rotating detonation combustor (RDC) severely constrains the operating range of gas turbine. In this paper, the detailed investigation was carried out for the influence of RDC passage area on cycle characteristic parameters for a 25MW single shaft gas turbine. The results demonstrated that changes of RDC passage area had a significant impact on operating range of rotating detonation gas turbine. As RDC passage area increased, operating range of gas turbine moved to the higher power, with the slight increases for cycle efficiency, cycle efficiency increment and net power increment compared to traditional gas turbine at identical net power. For the process when RDC passage area decreased from 0.0272 m2 to 0.0240 m2, the lower working limit of rotating detonation gas turbine decreased to 53.3 % of rated power, in which equivalence ratio of RDC replaced turbine efficiency as the greatest influence factor on cycle efficiency increment. [ABSTRACT FROM AUTHOR]

• The gas generator is combined with the rota detonation engine. • Detonation waves are generated spontaneously without pre-detonator. • The gas generator is combined with the rotating detonation engine. • Low gas temperature is not conducive to the deflagration-to-detonation transformation. • High gas temperature will produce secondary detonation wave. • Rich gas at higher temperatures contains more proportion of highly active ingredients. In this study, the propagation characteristics of a rotating detonation wave (RDW) between ethylene-rich gas and ambient air were investigated experimentally. The traditional pre-detonator initiation method was abandoned in the experiment, but the spontaneous combustion of high-temperature ethylene-rich gas and air was adopted. The RDW auto-initiated through the deflagration-to-detonation transformation (DDT) process. With an increase in the ethylene-rich gas temperature, the modes of RDW propagation appeared as delayed initiation, dual-wave collision, and fluctuating dual-wave collision modes. When the gas temperature was too high, a secondary detonation wave was produced near the head of the rotating detonation chamber (RDC), which affected the propagation efficiency and stability of the RDW. The increase in gas temperature expanded the equivalence ratio range of auto-initiation, which was due to the higher gas temperature, and more highly active components were precipitated. [ABSTRACT FROM AUTHOR]

A three-dimensional upwind conservation element and solution element method (CESE) in cylindrical coordinates is first developed to effectively solve the unsteady reactive Euler equations governing a hydrogen–air rotating detonation engine (RDE) with coaxial structures. The effects of the annular width on the structure of the detonation front and the relationship between the thrust and mass flow rate are then investigated. Additionally, RDEs with various injection conditions are systematically analyzed regarding flow patterns and propulsion performance. The results reveal a positive correlation between the specific impulse and the area ratio of the injection slot to the head-end wall. Nevertheless, the specific impulse shows minimal dependence on the injector slot’s location when the area ratio is constant. Ultimately, it is concluded that the area ratio between the injector and the head-end wall is critical in determining the loss of specific impulse.

To improve the propulsive performance, exhaust nozzles are desired for rotating detonation engines in practical applications. Therefore, their effects on the upstream behaviors inside the combustion chamber are required to be clarified. This study focuses on the characteristics of rotating detonation waves influenced by different outlet configurations. Three types of aerospike nozzle configurations are developed with different throat areas. Ethylene and oxygen-enriched air with an oxygen volume fraction of 50% are utilized in this study. When nozzles are employed, five wave propagation modes (i.e., the single-wave mode, the counter-rotating wave mode, the self-adjustive mode, the periodic oscillation mode, and the deflagration mode) are observed with the equivalence ratio ranging from 0.33 to 1.39, while only single-wave mode appears in the nozzle-less case under the same operating conditions. When increasing the equivalence ratio or reducing the nozzle throat area, those unstable propagation modes are more likely to happen and three boundaries of the equivalence ratio for the wave propagation transformation are clarified. On the contrary, the standard deviations of propagation velocities in the stable propagation process are smaller when employing nozzle configurations compared to the nozzle-less cases, which indicates a more stable state of the wave propagation although the stable propagation range is narrowed. Besides, the pressure rise and specific thrust are analyzed under different outlet configurations. • Five types of propagation modes were observed with different outlet configurations. • Propagation mode evolutions along with the equivalence ratio have been analyzed. • Nozzles enhance the propagation stability but narrowed stable mode operating range. • The propulsive performances of the rotating detonation engine were investigated. [ABSTRACT FROM AUTHOR]

This study aimed to evaluate the effects of the ratio of air injection pressure to hydrogen injection pressure (pR = pA/pH) on the propagation characteristics of rotating detonation waves (RDWs) and the working range of continuous rotating detonation engines (CRDEs). To this end, experiments were conducted under different injection pressure ratios and air mass flow rates with hydrogen as the fuel and air as the oxidant. The results reveal that RDWs can propagate in a stable manner when the injection pressure ratio is 0.67–1.7; however, RDW propagation is unstable when the injection pressure ratio is 0.51–0.67 and 1.7–2.03. The formation of unstable detonation waves may be attributed to the long response times at low pressures and the poor mixing effect of hydrogen and air. Furthermore, under an injection pressure ratio of 2.08–3.09, CRDE initiation failed. As the air mass flow rate increased, the range of the injection pressure ratio increased and the CRDE operated in a stable manner. In addition, the RDW velocity first increased and then decreased as the injection pressure ratio increased. Notably, under an air injection pressure of 3.6 bar, the rotation direction of the RDW changed during propagation. [ABSTRACT FROM AUTHOR]

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen or high-concentration hydrogen peroxide oxidizers with kerosene, hydrogen, and methane fuels. Performance results based on a lumped parameter treatment show that theoretical specific impulse gains of 3–14% are achievable with the highest benefit coming from hydrogen-fueled systems. Assessment of thrust chamber designs for notional space missions shows that both thrust chamber length and diameter benefits are achievable given the tiny annular chamber volume associated with the rotating detonation combustion. While the passing detonation front drastically increases local heat fluxes, global energy balances can be achieved if operating pressures are limited to be comparable to existing or prior space engines. [ABSTRACT FROM AUTHOR]

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The performance of rotating detonation engines (RDEs) is theoretically better than that of traditional aero engines because of self-pressurization. A type of swirl injection scheme is introduced in this paper for two-phase detonation. On the one hand, experiments are performed on continuous rotating detonation of ternary "kerosene, hydrogen and oxygen-enriched air" mixture in an annular combustor. It is found that increasing the mass fraction of hydrogen can boost the wave speed and the stability of detonation waves' propagation. One the other hand, characteristics of kerosene–hot air RDE is investigated for engineering application. Some unstable phenomena are recorded, such as changes of the number of detonation waves, low-frequency oscillations, and sporadic detonation. [ABSTRACT FROM AUTHOR]

Rotating detonation as a kind of pressure gain combustion is expected to greatly improve efficiency when applied to gas turbine engines. In this paper, the operation of rotating detonation combustor and turbine rotor blade was studied. Firstly, the analysis of the interaction between detonation wave and turbine blade shows that the compression of gas by detonation wave and reflected wave will lead to a sharp increase in the temperature at the wall of blade. When the detonation wave propagates, the oscillation amplitudes of pressure and temperature at the turbine inlet are 70% and 75% respectively, and the detonation oblique shock will change the flow trajectory of the air flow, resulting in the flow direction deviating from the incident angle. Then the comparison between detonation and deflagration shows that the total pressure of detonation is higher and will have greater work potential. The torque generated by the blades under detonation has the characteristics of high-frequency oscillation, which may be detrimental to the operation of the engine. [Display omitted] • Coupled operation of rotating detonation and turbine rotor blades. • Characteristics of interaction between detonation wave and turbine. • Variation of pressure and temperature in combustor under deflagration and detonation. • Torque value of turbine rotor under the influence of rotating detonation wave. [ABSTRACT FROM AUTHOR]

In order to investigate the structure and propagation characteristics of hydrogen-assisted kerosene (liquid phase)-air rotating detonation, a modified solver based on OpenFOAMⓇ was proposed and utilized to simulate detonation in the mixture of hydrogen, liquid kerosene and air. The obtained results show that evaporation waves exist in the discrete liquid mist detonation, and the coupled-waves structure formed by the evaporation wave (EW) and the incident shock wave (ISW) propagates upstream. By analyzing the overall and instantaneous propagation characteristics, it is found that the propagation speed of the detonation wave increases with the increase of the equivalence ratio at the fuel rich state. Increasing the droplet size will cause the evaporation wave to be further downstream relative to the ISW, resulting in lower evaporating pressure. The effect of EW on ISW is accomplished by the blast wave traveling upstream which is generated by droplet reaction, while EW is influenced by the ISW by modulating the droplet enthalpy and the pressure attenuation following ISW. Part of the droplet reaction energy is held in the gap between the EW and ISW, consequently the gap acts as an energy storage piston, regulating the ISW and EW propagation speeds. The combustion of droplets with smaller sizes imposes more positive feedback on their evaporation, and increasing the total temperature of the premixed gas promotes the interaction between the EW and the ISW, resulting in a more steady propagating overall structure. • The evaporation wave (EW)'s existence in two-phase rotating detonation is verified. • EW and incident shock wave (ISW) combine and propagate as a general structure. • Mechanism of two-phase detonation wave's propagation is extracted. • Influences of multiple factors on the propagation of EW and ISW are analyzed. [ABSTRACT FROM AUTHOR]

The flows in a combustion chamber shaped as an annular gap between plates with multiheaded rotating detonation of the propane–air mixture entering the chamber in the direction of the symmetry axis from the reservoir with the stagnation parameters specified are investigated numerically. The conditions for the formation of a preset number of waves in the multiheaded detonation wave are determined; they are associated with the dimensions of the combustion chamber and the parameters of the initiators. The values of the highest number of waves for the specified dimensions of the combustion chamber are obtained. It is found that the existence of the highest critical value of the number of waves in the multiheaded detonation is associated with blocking the feed of the combustible mixture. It is found that the mutual angles between the waves that make up the multiheaded detonation are gradually aligned with an uneven arrangement of initiators. For the calculations carried out on the Lomonosov supercomputer of Moscow State University, an original computer complex based on the modified Godunov method and one-stage reaction kinetics is used. [ABSTRACT FROM AUTHOR]

Rotating detonation combustors are prone to a wide range of off-design behaviors during both atmospheric and pressurized operation. The current manuscript deals with the transient mean pressure change seen in back-pressurized combustors after ignition. Using experimental high-speed pressure data acquired during this transience, we report two distinct, but overlapping, mechanisms that are well-known in deflagrative rocket combustors, but are not currently considered in rotating detonation combustors: DC shift and chugging. The former is characterized by sudden and seemingly random alterations of the time-averaged chamber pressure during operation, not unlike the same behavior observed in conventional rockets during instability onset. The latter is a low frequency instability that is inherently coupled to the reactants feed, causing significant oscillations of the mean chamber pressure. We show that these phenomena are strongly linked, where onset of chugging oscillations alter the frequency and pressure of the operating mode, which in turn produces a DC shift of the mean chamber pressure. To discern this complex interaction, we use multivariate polynomial curve-fitting, considering all the appropriate parameters to arrive at an empirical formulation that is seen to reasonably match the observed pressure transience after ignition. • The transient pressure increase after RDC ignition is studied. • The low frequency oscillations accompanying this transience is also investigated. • These two behaviors in RDCs are linked to DC shift and chugging instability in rocket engines. • These two phenomena are also interlinked between themselves. • An empirical correlation is proposed to discern the pressure evolution after ignition. [ABSTRACT FROM AUTHOR]

Research on the application of liquid fuels to continuously rotating detonation was conducted. A new method of mixture preparation was proposed. A special system of liquid fuel injection was designed and tested which is based on injecting into the detonation chamber a preheated liquid fuel partially mixed with hot air at conditions higher than the rich flammability limit. The specially selected conditions allow all liquid fuel to evaporate in the supply system but prevent it from ignition before entering the detonation chamber. Experiments were conducted for two different liquid fuels, extraction gasoline and Jet-A fuel. Research was carried out for different equivalence ratios, and in all tested conditions detonation was achieved. The new tested method of liquid fuel preparation and injection into a cylindrical detonation chamber opens a way of application of liquid fuels to engines which utilize continuously rotating detonation and thus prepares the way for practical application of detonative combustion to turbine engines and jet propulsion systems. [ABSTRACT FROM AUTHOR]

A rotating detonation combustor (RDC) is a novel approach to achieving pressure gain combustion. Due to the steady propagation of the detonation wave around the perimeter of the annular combustion chamber, the RDC dynamic behavior is well suited to analysis with reduced-order techniques. For flow fields with such coherent aspects, the dynamic mode decomposition (DMD) has been shown to capture well the dominant oscillatory features corresponding to stable limit-cycle or quasi-periodic behavior within its dynamic modes. Details regarding the application of the technique to RDC—such as the number of frames, the effect of subtracting the temporal mean from the processed dataset, the resulting dynamic mode shapes, and the reconstruction of the dynamics from a reduced set of dynamic modes—are analyzed and interpreted in this study. The DMD analysis is applied to two commonly observed operating conditions of rotating detonation combustion, viz., (1) a single spinning wave with weak counter-rotating waves and (2) a clapping operating mode with two counter-propagating waves at equal speed and strength. We show that care must be taken when applying DMD to RDC datasets due to the presence of standing waves (expressed as either counter-propagating azimuthal waves or longitudinal pulsations). Without accounting for these effects, the reduced-order reconstruction fails using the standard DMD approach. However, successful application of the DMD allows for the reconstruction and separation of specific wave modes, from which models of the stabilization and propagation of the primary and counter-rotating waves can be derived. [ABSTRACT FROM AUTHOR]

We consider a three-dimensional unsteady flow with one, two, or four rotating detonation waves arising in an annular gap of an axially symmetric device between two parallel planes perpendicular to its symmetry axis. The corresponding problem is formulated and studied. It is assumed that there is a reservoir with quiescent homogeneous propane–air combustible mixture with given stagnation parameters; the mixture flows from the reservoir into the annular gap through its external cylindrical surface toward the symmetry axis, and the parameters of the mixture are determined by the pressure in the reservoir and the static pressure in the gap. The detonation products flow out from the gap into a space bounded on one side by an impermeable wall that is an extension of a side of the gap. Through a hole on the other side of the gap and through a conical output section with a half-opening angle of 45 ∘ , the gas flows out from the engine into the external space. We formulate a model of detonation initiation by energy supply in which the direction of rotation of the detonation wave is defined by the position of the energy-release zone of the initiator with respect to the solid wall situated in a plane passing through the symmetry axis. After a while, this solid wall disappears (burns out). We obtain and analyze unsteady shock-wave structures that arise during the formation of a steady rotating detonation. We measure and compare thrust characteristics for different numbers of detonation waves. It was shown that the thrust weakly depends on the number of waves while the specific impulse increases with increasing number of waves. The study of rotating detonation at various values of the stagnation pressure was conducted. The calculation results show that at a stagnation pressure less than the critical value, the realization of rotating detonation is impossible. It is established that, depending on the stagnation pressure, qualitatively different shock-wave structures can be observed. The analysis is carried out within single-stage combustion kinetics by the numerical method based on the Godunov scheme with the use of an original software system developed for multi-parameter calculations and visualization of flows. The calculations were carried out on the Lomonosov supercomputer at Moscow State University. [ABSTRACT FROM AUTHOR]

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen or high-concentration hydrogen peroxide oxidizers with kerosene, hydrogen, and methane fuels. Performance results based on a lumped parameter treatment show that theoretical specific impulse gains of 3–14% are achievable with the highest benefit coming from hydrogen-fueled systems. Assessment of thrust chamber designs for notional space missions shows that both thrust chamber length and diameter benefits are achievable given the tiny annular chamber volume associated with the rotating detonation combustion. While the passing detonation front drastically increases local heat fluxes, global energy balances can be achieved if operating pressures are limited to be comparable to existing or prior space engines.

In this study, the effects of channel widths on the characteristics of the rotating detonation wave (RDW) were investigated. Pre-combustion cracked kerosene and 50% oxygen-enriched air were taken as the propellant. Keeping the outer diameter (D = 150mm) constant, the channel widths (W) of the combustor range from 15 mm to 50 mm in the experiments. The results indicate that the time for the formation of a stable RDW is longer under the wider channel, while the velocity of the RDW increases significantly with a wider channel. Increasing the ER has a positive effect on the wave velocity and the flow rate has little effect on wave velocity. The wave pressure increases under the higher ER and flow rate. Under the same flow rate and ER, the RDW pressure tends to reach the maximum value when the channel width is 25 mm, and the pressure range is 2 bar to 6 bar. Five kinds of the RDW modes were observed in the experiments, namely the failure "pop-out", single-wave mode, two-counter rotating waves mode, and two-co rotating waves mode. The two-counter rotating waves mode seems to be an intermediate mode of single-wave mode and two-co rotating waves mode in the conducted experiments, and the multi-wave mode is more likely to occur under the narrower channel and the higher oxygen content. [ABSTRACT FROM AUTHOR]

In order to investigate the relationship between high frequency tangential instability and continuous rotating detonation, series of H 2 /Air rotating detonations are experimentally achieved in the hollow chamber with double injection sections. In the center part, gaseous H 2 and air injected by co-axial injector. Near the outer wall, the same propellants are injected in the form of slit-orifice collision. By keeping the total air mass flow rate approximately constant, varying the mixture of the inner and outer injection, series experiments are conducted in the test model with or without Laval nozzle. The results verify the possibility of rotating detonation in the hollow chamber with co-axial injector. To clarify the relationship between continuous rotating detonation and high frequency tangential combustion instability, the intrinsic frequencies of the test model are captured to be compared with propagation frequencies of detonation waves. The results show that they are close to each other when enough propellant assembled near the outer wall. In the combustor, the flame direction in constant pressure mode can change itself into rotating direction spontaneously. The results indicate that rotating detonation is an implication to high frequency tangential instability. • The results verify the possibility of rotating detonation in the hollow chamber with co-axial injector. • The key factor to forming detonation wave with co-axial injector has been discussed. • The results indicate that rotating detonation is an implication to high frequency tangential instability. [ABSTRACT FROM AUTHOR]

Previous experimental studies show that gaseous detonation propagation in an annular tube can be divided into three categories, i.e., stable mode, critical mode and unstable mode. However, the mechanism of detonation propagation with the critical mode is not well understood due to insufficient experimental data. In this paper, detonation propagation with the critical mode in the annular tube was numerically studied for the methane–air mixture based on reactive Navier–Stokes equations, in which the convective term was integrated by the fifth-order weighted essentially non-oscillatory. The numerical results show that the trajectories of the triple points of the shock wave front were drawn to obtain cell structure with "petal" pattern, which is agree with the experimental results. The wall curvature plays an important role for this phenomenon. The outer wall is concave, which can continuously compress the detonation wave and make it stronger. The inner wall is convex, which can produce an isentropic expansion and decouples the reaction front from the shock wave. The transverse shock wave emanating from the triple point moving from the outer wall reignites detonation near the inner wall, which implies that compression of the outer wall plays a key role in the rotating detonation propagation. As the transverse shock wave moves away from the inner wall, the detonation wave near it gradually weakens. Because of the back and forth reflection of the transverse shock wave between the inner and outer walls, the rotating detonation wave can be maintained and its cellular structure presents "petal" pattern. Meanwhile, the averaged detonation velocity at the outer wall is higher than that at the inner wall. [ABSTRACT FROM AUTHOR]

The combustion characteristics in annular rotating detonation combustors are experimentally investigated using 40% oxygen-enriched air and liquid aviation kerosene (RP-3) as reactants. Two types of wall cavity structures, embedded in the outer and the inner wall of the annular combustors, respectively, are employed to study their effects on detonation stability. The characteristics of typical combustion modes observed in the annular combustors are firstly analyzed. The regime of typical combustion modes is then examined in the phase diagram of mass flow rate and equivalence ratio. Results show that the single-wave mode regime can be affected by wall cavities. In comparison to the combustor without a wall cavity, the regime is expanded for the cavity in the inner wall but significantly reduced for the cavity in the outer wall. When the single-wave mode forms, the cavities embedded in the inner wall result in a significant increase of the apparent detonation wave velocity but a decrease in the velocity fluctuation rate, and the effects are changed with the axial location of the cavities. The axial location of the cavity can also influence the stable propagation location of the detonation wave. Additionally, the analysis of the high-frequency pressure signals indicates that the cavities do not affect the initiation process of the formation of the combustion modes. However, the variation of oxidizer mass flow rate and equivalence ratio can evidently influence the evolution time of different combustion modes after ignition. The study indicates that the wall cavity structure has a significant influence on the propagation behavior of rotating detonation waves in the combustors, and it can be utilized to enhance the stability of kerosene two-phase rotating detonation combustion. [ABSTRACT FROM AUTHOR]

The performance of rotating detonation engines (RDEs) is theoretically better than that of traditional aero engines because of self-pressurization. A type of swirl injection scheme is introduced in this paper for two-phase detonation. On the one hand, experiments are performed on continuous rotating detonation of ternary “kerosene, hydrogen and oxygen-enriched air” mixture in an annular combustor. It is found that increasing the mass fraction of hydrogen can boost the wave speed and the stability of detonation waves’ propagation. One the other hand, characteristics of kerosene–hot air RDE is investigated for engineering application. Some unstable phenomena are recorded, such as changes of the number of detonation waves, low-frequency oscillations, and sporadic detonation.

Considering the pressure gain characteristic of continuously rotating detonation (CRD) combustion, a rotating detonation gas turbine cycle was established using methane as fuel. The variation of pressure gain characteristic and its influence on cycle performance at different compressor pressure ratios and air distribution proportions were investigated combined with a series of two-dimensional numerical simulations. The results were demonstrated that compared with the conventional approximate iso-pressure combustion, rotating detonation combustion could significantly improve the thermodynamic performance of gas turbine in various conditions. Due to the difference of entropy generation between rotating detonation and iso-pressure combustion, the pressure gain characteristic of the CRD combustor was provided a higher cycle efficiency. The efficiency gain was higher at lower compressor pressure ratio and air proportion entering CRD combustor. In the calculated cases, the pressurization ratio of the CRD combustor was higher than 1.37, cycle efficiencies of rotating detonation gas turbine were reached 0.395, cycle efficiency increasing rates were more than 22.5%, and power generation enhancements were greater than 1500 kW.