Rotating Detonation Combustor Research Articles

Numerical study on the integration of supersonic turbine guide vanes and three-dimensional hydrogen/air rotating detonation combustor

Hydrogen/air rotating detonation turbine engine is expected to become a new generation of aerospace power plant because of its compact structure, high cycle thermal efficiency, and superior thrust performance. It can also reduce fuel consumption, save energy, and reduce carbon emissions. However, the highly unsteady oscillation characteristics of the outlet flow of the rotating detonation combustor make it difficult to integrate the supersonic turbine with the rotating detonation combustor. In this paper, the supersonic turbine guide vanes are designed by the method of characteristics and Bessel parameterization and are integrated with three-dimensional hydrogen/air rotating detonation combustors for numerical studies. The effects of aligned mode and misaligned mode on the coupling of supersonic turbine guide vanes and rotating detonation combustor are discussed carefully. The results show that the supersonic turbine guide vanes can make the rotating detonation wave change from a single-wave mode to a double-wave alternating strength and weak propagation mode. It can effectively suppress the oscillation of the combustion chamber outlet airflow. In the aligned mode, the peak pressure at the outlet of the supersonic turbine is about 70% lower than that at the cascade inlet, the pressure oscillation amplitude is reduced by 93.33%, and the temperature amplitude is reduced by 23.81%; the average total pressure loss coefficient of the cascade is 11.63%. In the misaligned mode, compared with the cascade inlet, the peak value of the pressure signal at the cascade outlet decreases by about 50%, while the pressure oscillation amplitude decreases by about 33.33%, and the temperature oscillation amplitude decreases by 11.11%; the average total pressure loss coefficient of the cascade is 4.83%. The supersonic turbine guide vanes have a better suppression effect on the oscillation signal in the aligned mode, but the relative total pressure loss is relatively large. This is because that the oblique shock wave, channel shock wave, and supersonic turbine guide vanes interact to generate more complex wave system and secondary flow in the aligned mode. These features provide important reference information for the coupling of supersonic turbines and rotating detonation combustors.

Coupling study of supersonic turbine stage and two-dimensional hydrogen/air rotating detonation combustor

The hydrogen/air rotating detonation turbine engine has the advantages of self-supercharging, small entropy increase, high thermal efficiency, high thrust-to-weight ratio, low fuel consumption, and low carbon emissions. However, the high-frequency and high-speed oscillation characteristics of the outflow of the rotating detonation combustor pose challenges for the turbine to extract work efficiently. In this study, a supersonic turbine stage designed using the Python code of the method of characteristics coupled with a two-dimensional rotating detonation combustor is numerically investigated. The propagation characteristics of the detonation wave in the aligned mode and misaligned mode and interaction with the supersonic turbine stage are carefully discussed. The results show that the coupling of the supersonic turbine stage and the rotating detonation combustor will cause the detonation wave to change from a single-wave mode to a three-wave co-propagating mode. The Kelvin–Helmholtz instability of the slip line increases after passing through the induced oblique shock. In the multi-wave mode, the detonation wave is self-adaptive, and multiple detonation waves interact to automatically adjust the propagation velocity, intensity, and distance of each detonation wave, and finally achieve a dynamic balance. The supersonic turbine stage has good operating performance under the condition of rotating detonation flow, its power level can reach 110 kW, and the maximum stagnation adiabatic efficiency of the supersonic turbine stage can reach 86%. The supersonic turbine guide vanes can greatly reduce the oscillation amplitude of the incoming flow. In the aligned mode, the supersonic turbine guide vanes has a more obvious effect of suppressing the amplitude of the incoming flow. The total pressure loss of the supersonic turbine stage is smaller, and the supersonic turbine rotor can extract work more efficiently in the aligned mode. These findings provide a valuable reference for further research on the hydrogen/air rotating detonation turbine engine, ultimately leading to the practical application of an energy-saving, high-efficiency, and low-emission hydrogen/air rotating detonation turbine engine.

Rotating detonation combustion of liquid kerosene under near-ramjet limit conditions

The feasibility of using preheated kerosene and pre-evaporation to assist in liquid kerosene detonation ignition and self-sustaining combustion under near-ramjet limit conditions was investigated. Additionally, the detonation decoupling and self-sustaining mechanisms were analyzed in depth. Owing to the forced heat transfer effect of supersonic flow on kerosene droplets, self-sustaining detonation combustion could not be achieved using preheated kerosene alone. The detonation wave was still decoupled due to the forced heat exchange effect of the cold kerosene droplets on the detonation wave, even if the distribution of C12H23 reached the detonation conditions through the pre-evaporation of fuel.

Experimental study on a premixed rotating detonation combustor using Tesla inlet configuration fueled by kerosene

The working performance of a rotating detonation combustor (RDC) can be enhanced by improving the mixing effect. Presently, most non-premixed injection schemes are adopted for safety purposes. However, the propellant in the premixed injection is completely mixed and thus can improve the working performance of RDC. In this study, kerosene and 28.5% oxygen-enriched air (at room temperature) were taken as the propellant, and the Tesla valve inlet structure was installed upstream of the combustion chamber to successfully realize a stable and self-sustaining propagation of the rotating detonation wave (RDW) under a premixed injection scheme. The propagation characteristics of the RDW and the performance of RDC under premixed and non-premixed injection conditions were then compared. The lean flammability boundary of RDC was effectively widened, and the intensity of the RDW was improved under premixed injection due to the better mixing effect of propellant and higher activity of combustible mixture. However, the high propellant activity led to premature combustion of the fuel, thus reducing the equivalence ratio boundary of RDC by about 50%.The RDW intensity was significantly higher under premixed injection than under non-premixed injection when the airflow rate was increased. The pressure feedback was quantitatively analyzed using the percentage of the pressure fluctuating amplitude in the air plenum to the pressure amplitude of the detonation wave (PFdw). The feedback intensity was closely associated with the intensity of the RDW (minimum PFdw; about 22%). The specific thrust was increased by about 48.3% (maximum specific thrust; about 743 N·s·kg−1), and the specific fuel consumption was reduced by 36.8%(minimum specific fuel consumption; about 0.24 kg·(N·h)-1) under the premixed condition. Meanwhile, the flow field of the inlet throat of RDC under premixed injection was subsonic, which greatly reduced the inlet pressure loss and further improved RDC performance.

Coupling characteristic analysis and propagation direction control in hydrogen–air rotating detonation combustor with turbine

Two-dimensional numerical simulations are performed to investigate the interactions between a rotating detonation combustor (RDC) and turbine. The flowfield structure and coupling characteristics are analyzed under different equivalent ratios. Furthermore, the formation mechanism of the detonation wave direction is illustrated. Results show that although RDCs increase the average turbine work, they enhance the flowfield oscillation and maintain the quasi-periodic fluctuation in the turbine work. Compared to counter-clockwise propagation, the pressure oscillation attenuation and total pressure loss through the turbine guide vane increase when the detonation wave propagates clockwise. When the detonation wave height is low, RDCs are more prone to quenching due to the impact of reflected waves. After re-ignition, the detonation wave direction is unrelated to the initial ignition direction and is random. Introduction of deflecting wedges into RDCs enables the automatic control of the detonation wave direction but causes extra total pressure loss of at least 3.6%.

Impact of mixture mass flux on hydrodynamic blockage ratio and Mach number of rotating detonation combustor

To analyze non-ideal phenomena, such as burned gas backflow and non-detonation combustion, which affect the rotating detonation wave Mach number, simultaneous self-luminous visualization, time-averaged static pressure, fluctuating pressure, and thrust measurements with gaseous ethylene and oxygen were performed. Consequently, by doubling the number density of the fuel injectors, the hydrodynamic blockage ratio at the oxidizer inlet increased approximately 1.7-fold under the same oxidizer inlet area conditions. This may be attributed to the increase in the detonation propagation Mach number owing to the enhanced mixing of fuel and oxidizer. The relationship between the parasitic combustion fraction in front of the rotating detonation wave and the Mach number was also investigated by using a distributed heat release model. Consequently, it was suggested that experimental Mach number decreased from approximately 4.1 to 2.8 with increase in a mixture mass flux, and the theoretical detonation wave propagation Mach number was 7.3.

Numerical study of the mechanisms of the longitudinal pulsed detonation in two-dimensional rotating detonation combustors

A numerical study of the longitudinal pulsed detonation (LPD) is conducted in the present paper. The occurrence mechanism of the LPD, called shock wave amplification by coherent energy release, is verified preliminarily in this study. To be specific, upstream propagating shock waves, which originate from the outlet, induce a specific gradient of reactant distribution, and then detonation waves are ignited and evolve along the gradient in close succession. It is worth noting that the occurrence of LPD does not mean that the LPD will necessarily be sustained. The low injection pressure ratio PR (i.e., the ratio of inlet pressure to outlet pressure) = 1.3 is found to be conducive to the sustenance of the LPD instability in the baseline model. A lower PR (PR ≤ 1.2) or a slightly higher PR (1.4 ≤ PR ≤ 1.8) shall lead to an unstable detonation or quenching of detonations, while a much higher PR (PR > 1.8) contributes to the formation of stable canonical rotating detonation waves. In addition, the combustion regimes of five combustors of different heights at different PR are explored. As the combustion chamber height increases, the PR of the sustainable LPD is nearly linearly increasing, and its operating frequency decreases gradually. The calculation formula between the sustainable LPD propagating frequency and the natural acoustic resonance frequency of the combustor is employed and discussed, but in consideration of its imperfection, further investigation is required.

Evaluation of a computational strategy to model transitory injection in rotating detonation combustors

The efficiency of a Rotating Detonation Combustor (RDC) strongly depends on the transitory injection process of fresh reactants in the combustion chamber: poor propellant mixing induces losses of combustion efficiency and consequently low detonation speed and unstable detonation propagation. Moreover, dilution of fresh reactants with burnt gases during injection increases the deflagration losses and decreases the pressure gain provided by the detonation. Numerical simulation can help design an efficient injector to reduce these losses. In this study, the modeling strategy previously proposed by ONERA to simulate the transitory injection process is applied to two existing experimental RDC (from Nagoya University and TU Berlin) and one in-development RDC from ONERA. The computational domain represents only one injection element, convenient for a parametric study at low computational cost. A custom initial condition is used to model the expansion process of burnt gases past a detonation wave. The initial condition parameters are discussed and a method is proposed to correctly set them. The TU Berlin RDC is studied in more detail: mixing efficiency up to 70% is obtained, and 5% of deflagration losses are estimated according to the assumptions of the simulation. Based on the numerical results, detonation speed was evaluated at various distances from the injection plane taking into account the heterogeneities of the fresh mixture. The measured speed lies within the predicted range.

Integration of a transonic high-pressure turbine with a rotating detonation combustor and a diffuser

Abstract In this paper, a diffuser is used to integrate a transonic high-pressure turbine with a rotating detonation combustor (RDC). The paper focuses on the required design modifications to the turbine endwalls (EW) to enable high efficiency, while preserving the airfoil blade-to-blade geometry. The main challenge is the stator passage unstarting, due to the high inlet Mach number. First of all, steady Reynolds Averaged Navier Stokes simulations were performed to compare the efficiency of turbines with constant-radius EWs to turbines with axisymmetric EWs. A modified EW design prevented the unstarting of the stator passage, enabling a significant gain in performance. Afterward, the influence on the turbine efficiency and damping due to the unsteadiness from the diffuser-like fluctuations of the RDC was evaluated with unsteady Reynolds Averaged Navier Stokes simulations with a mixing plane approach (MPA). Full unsteady simulations were carried out on selected inlet conditions and compared to the mixing plane results. This parametric study provides turbine designers with recommended diffusion rates along the vane EWs. Additionally, we provide guidance on the upstream diffuser design, specifically the required damping and outlet Mach number.

Effects of a straight guide vane on the operating characteristics of rotating detonation combustor

Effects of a straight guide vane on the operating characteristics of rotating detonation combustor

Profiling cross-sectional area of a radial rotating detonation combustor to increase pressure gain

The detonation wave(s) in a rotating detonation combustor (RDC) produce highly unsteady flow with varying supersonic and subsonic segments. The non-homogeneous RDC flow must be conditioned to improve its uniformity, which can increase the performance and operation of the downstream hardware for propulsion and power generation applications. In this study, the RDC channel and converging nozzle are combined into a single flow conditioning device, and Gradual to rapid (GTR), Linear, and Rapid to Gradual (RTG) strategies to profile the cross-sectional area of the combined channel are introduced. An inward flowing disk-RDC (D-RDC) operated on gaseous methane-oxygen is numerically investigated using a quasi-two-dimensional computational fluid dynamic (CFD) simulation with a curvilinear coordinate system. The study focuses on the combined channel to isolate the effects of area profiling and neglects radial to axial turning of the flow necessary to integrate it, for example, with an aerospike nozzle located downstream. Simulations show that the RTG area profile yields a higher performance compared to GTR and Linear cases. The RTG profile is also shown to improve pressure and temperature uniformity of the RDC flow at the throat. Results are analyzed to show that a rapid decrease in the cross-sectional area closer to the detonation for the RTG profile improves the stability of the detonation process and constrains the expansion of product gases behind detonation to orient the flow radially inwards. Therefore, profiling the cross-sectional area of an RDC can be vital in improving the operability and effective integration of the downstream hardware.

Visualization and Performance Evaluation of a Liquid-Ethanol Cylindrical Rotating Detonation Combustor

Rotating detonation combustors (RDCs) are among the combustors that use supersonic combustion waves known as detonation waves, and are expected to simplify engine systems and improve thermal efficiency due to their supersonic combustion and compression performance using shock waves. Research is also being actively conducted worldwide on a cylindrical RDC; a RDC without an inner cylinder, which is expected to simplify and downsize the combustor. However, most of the research was performed using gas propellants, and liquid propellants were rarely used. Since liquid propellants are used in many combustors, it is important to evaluate the performance of RDCs with liquid propellants. In this study, a cylindrical RDC with a liquid ethanol–gas oxygen mixture was constructed and tested at a flow rate of 31.5 ± 5.0 g/s, an equivalence ratio of 0.46–1.39, and a back pressure of 14.5 ± 2.5 kPa. The thrust was shown to depend strongly on the combustor bottom pressure history. In addition, the start-up process of the cylindrical RDC with liquid fuel was clarified by self-luminous and CH* radical visualizations. It was found that the detonation wavefront propagated at a distance of 2–3 mm from the combustor bottom, and the main combustion region was 10–15 mm in height.

On the wave propagation modes and operation range in rotating detonation combustor with varied injection and outlet throat

The rotating detonation combustor (RDC) has received remarkable attention in the aerospace community. In this work, an experimental RDC model supplied by liquid kerosene and oxygen-enriched air is established. A parametric survey is performed with different injection throats, outlet restrictions, and equivalence ratios to analyze the rotating detonation wave propagation modes comprehensively. Dynamic pressure transducers and a high-speed camera are both employed to identify wave modes synchronously. Overall, the propagation modes are found to be highly dependent on the injection throat and combustor outlet restriction. With a large injection to annulus area ratio of 0.3, a single-wave mode is characterized when no restriction is added at the combustor outlet. Reducing the outlet area leads to a decrease in the wave frequency and a narrower steady rotating detonation propagation regime. The propagation stability of the rotating detonation is strengthened when the injection to annulus area ratio decreases to 0.2. A dual-wave collision mode and a four-wave collision mode are observed, depending on the outlet restriction. A preliminary stable RDC operation range correlated with outlet to injection throat area ratio and equivalence ratio is achieved. Furthermore, an interval value of the outlet to injection throat area ratio is proposed to reach the potential positive total pressure gain. These findings should serve as the reference for RDC configuration design in air-breathing and gas-turbine propulsion systems.

Liquid fuel refill dynamics in a rotating detonation combustor using megahertz planar laser-induced fluorescence

The highly dynamic response of a continuously injected liquid fuel jet to rotating detonation waves is a critical parameter in the design, performance optimization, and modeling of a rotating detonation engine (RDE). In this work, this response is spatio-temporally resolved from the fuel injection point to the detonation channel using planar laser-induced fluorescence (PLIF) at imaging rates up to 1 MHz. A rotating detonation combustor (RDC) is operated on hydrogen and air to sustain stable detonation waves that interact in a one-way coupled manner with a single liquid fuel jet that propagates into the combustion chamber with cycle periods of ∼250 µs. Diesel is utilized as a realistic fuel surrogate with higher aromatic compounds to enable fluorescence excitation using the 355 nm third-harmonic output of a burst-mode Nd:YAG laser. By optimizing the technique to accommodate orders of magnitude variations in the fuel density throughout the injection process, the PLIF data enable measurements of (i) the overall refill dynamics after the arrival of the detonation wave, (ii) changes in the liquid spray trajectory with microsecond temporal resolution, and (iii) the time required to reestablish the quasi-steady axial refilling process as a function of peak chamber pressure relative to the air- and liquid-injector pressure drops. As the passage of the detonation wave imparts significant changes in the momentum flux ratio, the qualitative liquid break-up process and spatial distribution of the spray also vary significantly in time. Only as the injection system recovers late in the cycle does the fuel spray eventually return to a quasi-steady position and allow comparisons with theoretical jet trajectories. These data, enabled by ultra-high-speed PLIF imaging, represent some of the first detailed measurements for quantifying the dynamic response and recovery of liquid jets exposed to periodic detonations in an operating RDC.

Investigation of rotating detonation gas turbine cycle under design and off-design conditions

As cycle efficiency of gas turbine can be significantly improved due to the pressure gain characteristic of rotating detonation combustion, a cycle calculation model of rotating detonation gas turbine based on component characteristic curves of a single shaft gas turbine of 25 MW and limited maximum flow capacity of rotating detonation combustor was established to investigate the variations of cycle performance and component characteristics under design and off-design conditions. Thereafter, sensitivity analysis of component parameters was conducted to explore the comprehensive reason for change of cycle efficiency increment. The results demonstrated that compared with those of traditional gas turbine, higher cycle performance and lower compressor pressure ratio were gained in rotating detonation gas turbine. Increment of cycle efficiency and cycle net power increment significantly increased with the drop of load, which reached to 0.0248 and 3.261 MW when turbine inlet total temperature was 1450 K and increased to 0.0539 and 4.008 MW when turbine inlet total temperature was 1279 K. Cycle performance was significantly affected by the changes of compressor pressure ratio, compressor efficiency, equivalence ratio of RDC and turbine efficiency, among which turbine efficiency rather than compressor outlet total temperature had the most significant effect on cycle efficiency increment under the calculated conditions.

Numerical investigation on the interactions between rotating detonation wave complex and planar turbine cascade

AbstractDue to pressure gain characteristics, the rotating detonation combustor (RDC) can be integrated into gas turbines by further improving the system's performance. The matching of unsteady flows in the RDC and turbine is a key challenge. Here, interactions in an RDC and planar turbine cascade are investigated using the unsteady Reynolds time‐averaged Navier‐Stokes method. The effects of the blade profiles on interactions between the oblique shock wave (OSW) and turbine cascade are analyzed, and the suppression effect of the turbine cascade on high‐frequency pressure oscillations is evaluated. The results show that the OSW interacts with the leading edge, pressure side, suction side, and trailing edge of the turbine stator and rotor blades, which generates complex wave systems in the cascade passage. The opposite propagation direction causes the OSW to be aligned or misaligned with the stator blade. The position of the OSW that operates on the turbine cascade changes, and two different wave system structures appear. The various pitches of the turbine cascade at different blade heights cause differing intensities and propagation directions in the reflected wave as generated by the action of the OSW and turbine cascade and the variable flow fields at different blade heights. The turbine cascade can significantly suppress the high‐frequency pressure oscillations, while the pressure amplitude attenuation rate can reach over 80% when the OSW propagates through the turbine cascade. These findings expound on the interaction mechanism between the rotating detonation wave complex and the turbine stator and rotor blades.

The influence of component parameters on cycle characteristic in rotating detonation gas turbine

A cycle calculation model of methane-air directly mixing rotating detonation gas turbine was built combining two-dimensional numerical simulations of rotating detonation combustor (RDC) with convergent inlet structure. Controlled variable method was used to investigate the cycle performance improvement for different component parameters compared with traditional gas turbine, and sensitivity analysis to gain the influence weight of component parameters was carried out. The results demonstrated that cycle efficiency increasing rates were all more than 11.5% in simulated conditions. With the decrease of compressor ratio, compressor efficiency, turbine efficiency and equivalence ratio of RDC, cycle net power increments were all in a downward trend, while cycle efficiency increment showed different variation trends especially when equivalence ratio of RDC and compressor pressure ratio decreased. Turbine efficiency was the most significant factor for cycle performance improvement and the influence of compressor efficiency had been stable. Compressor pressure ratio had different influence weights on cycle efficiency increment when above and below 8.00. Equivalence ratio of RDC had an opposite effect on cycle efficiency increment and different influence weights on net power increment when above and below 0.75.

Numerical Study of the Influence of Inlet Mass Flow Rate on Rotating Detonation Flow Field Characteristics and Pressure Gain Performance

ABSTRACT A three-dimensional numerical study is performed to investigate the effect of inlet mass flow rate on flow field characteristics and pressure gain of rotating detonation combustor. The results show that the mass flow rate has an important effect on the behavior of the detonation wave. When the inlet mass flow rate is increased, the flow Mach number of inlet throat increases and flow develops to a close critical state. After ignition, two counter-rotating detonation waves will collide in the combustor and the number of collisions decreases with the increase of inlet mass flow rate. The strong wave eventually engulfs the weak wave and develops into a single-wave mode. When the inlet mass flow rate is reduced, the detonation wave velocity and refill height of premixed gas increase. Deflagration area of flow field reduces and the stability is improved. In addition, when the inlet mass flow rate is reduced, the total pressure loss caused by expansion wave decreases. The pressure gain and fuel-based specific impulse are increased. However, increasing the total pressure gain also improves the static pressure fluctuation ratio of inlet throat. The phenomenon of high-pressure feedback under small flow rate condition is also more prominent.

Rotating Detonations in Hollow and Flow-Through Combustors

This paper focuses on rotating detonation combustor (RDC) experiments involving two configurations, a hollow and flow-through RDC, using hydrogen–air and ethylene–air mixtures. Two air injection areas with gaps of 1.778 and 3.81 mm were used at different air mass flow rates of 0.2, 0.3, 0.4, and 0.5 kg/s and equivalence ratios from 0.6 to 2 in 0.2 increments. Results show repeatable detonations for hydrogen–air mixtures in both configurations; however, the detonative range is wider, stronger, and faster in the hollow configuration. Speeds exceeding the Chapman–Jouguet velocity and pressures matching the ideal model were observed. Sustained detonations for ethylene–air mixtures were not attained with the flow-through configuration. Installing a back plate to convert the configuration into a hollow combustor promoted rotating detonations with the ethylene–air mixture at select operating points. The results of this study showcase the strong promise of producing and sustaining rotating detonation in nonannular RDC configurations.

Influences of thermal physical property parameters on operating characteristics of simulated rotating detonation ramjet fueled by C12H23

Two-phase rotating detonation ramjets are considered to be suitable for aerospace applications due to their high thermodynamic cycle efficiency. These engines have an extremely complex internal flow field, in which the liquid fuel undergoes physical and chemical processes such as fragmentation, evaporation, mixing, and combustion; these processes also interact with detonation waves that have significant gradients. This makes it difficult to simulate a three-dimensional (3D) full-process rotating detonation combustion chamber. Here, based on the Euler–Lagrangian simulation method, a 3D numerical combustion chamber was simulated using kinetic theory and the constant thermal physical property parameter (TPPP) calculation method. The accuracy of these methods was then compared with the existing experimental results and theoretical values. Calculating the TPPPs using kinetic theory brought about a relatively high-pressure peak and detonation wave temperature; the detonation wave profile was also finer and more precise. The detonation wave propagation velocity of the two-phase detonation is estimated to be about 60% of the theoretical gas-phase CJ velocity. The calculation method of physical parameters has relatively little influence on the engine’s operating frequency and the detonation wave's propagation velocity but has a more significant influence on the peak pressure. Constant TPPPs can be used when the Kelvin–Helmholtz–Rayleigh–Taylor model with insufficient precision is used to consider the breakup of droplets and leads to the acceleration of the propagation speed of two-phase detonation waves.