Rotating Detonation Combustor Research Articles

Shape Optimization of a Diffusive High-Pressure Turbine Vane Using Machine Learning Tools

Machine learning tools represent a key methodology for the shape optimization of complex geometries in the turbomachinery field. One of the current challenges is to redesign High-Pressure Turbine (HPT) stages to couple them with innovative combustion technologies. In fact, recent developments in the gas turbine field have led to the introduction of pioneering solutions such as Rotating Detonation Combustors (RDCs) aimed at improving the overall efficiency of the thermodynamic cycle at low overall pressure ratios. In this study, a HPT vane equipped with diffusive endwalls is optimized to allow for ingesting a high-subsonic flow (Ma=0.6) delivered by a RDC. The main purpose of this paper is to investigate the prediction ability of machine learning tools in case of multiple input parameters and different objective functions. Moreover, the model predictions are used to identify the optimal solutions in terms of vane efficiency and operating conditions. A new solution that combines optimal vane efficiency with target values for both the exit flow angle and the inlet Mach number is also presented. The impact of the newly designed geometrical features on the development of secondary flows is analyzed through numerical simulations. The optimized geometry achieved strong mitigation of the intensity of the secondary flows induced by the main flow separation from the diffusive endwalls. As a consequence, the overall vane aerodynamic efficiency increased with respect to the baseline design.

Investigation on the integrated characteristics of n-decane/air rotating detonation combustor and supersonic turbine

In this study, the two-dimensional numerical simulations are conducted to study the flow field characteristics, turbine performance and loss mechanism of integrated system of rotating detonation combustor and supersonic turbine under two different directions of detonation wave propagation. The results indicate that the propagation direction of RDW affects the incident angle between OSW and guide vanes, resulting in different operating modes for aligned and unaligned modes. OSW in the turbine cascade undergoes the leading-edge shock, blade surface reflection and trailing edge diffraction. The backward-propagating rake-type shock envelope is formed due to leading-edge shock of the rotor. In misaligned mode, the stator has higher damping of both pressure and temperature. The stator significantly improves the circumferential uniformity of both velocity and pressure, particularly when operating in aligned mode. The total pressure loss in aligned mode is less, therefore the turbine achieves higher rim work and efficiency. The viscosity is one of the main sources of loss in the flow field. The reflected shock waves at the leading edge of the rotor and in the stator cascade are the primary factors contributing to the leading-edge vortices on the stator vanes.

Analysis of waves dynamics in a rotating detonation combustor fueled by kerosene

The wave dynamics play a crucial role in the operation characteristics of the rotating detonation engine. We conducted numerical simulations of a rotating detonation combustor (RDC) using multicomponent reactive Navier–Stokes equations coupled with a discrete phases model. The RDC in this research employs a configuration with multiple coaxial injectors supplying oxygen-enriched air and kerosene spray at room temperature. To accurately identify and analyze waves within the RDC, we proposed a three-dimensional transient detonation wave detection method based on the combined parameters of normal Mach number and heat release rate in the flow field. Two typical wave modes, referred to as single-wave mode and counter-waves mode, are identified and then selected to conduct a detailed wave dynamics analysis. The general wave behavior is discussed, and velocity deficit is compared for these two wave modes. For the single-wave mode, intermittent micro-explosions are observed generating retonation waves periodically in the unburnt pockets behind the rotating detonation shock front. For the counter-waves mode, we analyzed the collision process of the two waves and the coupling/decoupling of the shock front with the detonative heat release zone, revealing the reason for significant velocity deficits in this wave mode. This research demonstrates that micro-explosions intermittently occur in the multiphase RDC in both single-wave and counter-waves modes and generate micro explosion shock waves periodically, which influence the complicated wave dynamics behavior.

The impact of film cooling on the heat release within a rotating detonation combustor

Rotating detonation combustors establish a detonation wave that continuously circulates inside a small annulus. The presence of the detonation wave and the downstream oblique shock within the small annulus coupled with high mass flow induces a high heat load to the combustor wall. Preliminary analysis shows that for higher thermal power, internal air cooling alone is not sufficient to remove the heat out of the walls to maintain them below the maximum temperature of the metal. A possible solution is to use film cooling to reduce the heat flux to the combustor walls. One issue, though, is that the introduction of film cooling provides additional air into the system that can influence the combustion process as well as providing a location for secondary combustion.This paper represents the first investigation to study the secondary implications on combustion of using film cooling in a rotating detonation combustor. The TU Berlin RDC architecture was modified with the introduction of 480 film cooling holes placed in the oblique shock region. High fidelity LES investigations were performed for different coolant plenum pressures to show the benefits of using film cooling. However, due to the presence of unburnt fuel in this post-detonation region, the coolant can result in additional combustion leading to an increase in temperature near the wall. One the one hand, these secondary reactions result in an increase of the overall heat release increasing combustion efficiency, however this also results in higher temperatures and reduced film cooling effectiveness. A simulation performed with nitrogen as a coolant enabled the effects of increased mixing caused by the ejection of coolant gases to be separated from the additional heat release. The simulation with nitrogen shows a reduction of 88% in the local heat release in the post detonation region resulting in similar performance as the uncooled case and significantly cooler walls.

Propagation and stabilization properties of rotating detonation waves in a hollow combustion chamber with heated air inflow

The application of rotating detonation to conventional engine poses challenges such as high-temperature inflow conditions. In this study, experiments were conducted to discuss the propagation and stabilization properties of rotating detonation waves within a hollow rotating detonation combustor using pure heated air/ethylene as propellant. The temperature and equivalence ratio are varied to explore the boundary of detonative limits. The results show that the pressure of detonation waves is influenced by temperature fluctuations, whereas changes in air temperature have minimal impact on velocity losses. In addition to detonation, the deflagration zone is observed within the combustor. When the air temperature is increased, the deflagration zone expands gradually. The static temperature of air increases from 304 to 618 K, resulting in an increase in the area of the glowing zone from 17.7% to 60.0%. The expansion of the zone indicates a rise in deflagration percentage, and a stable single-wave mode is achieved mainly under fuel-rich conditions. The detonative equivalence ratio range is narrowed with increasing heated temperature at 484–618 K. Unstable propagation modes, characterized by sawtooth waves, are frequently observed at detonative lower limit, where the equivalence ratio is around 1.

Experimental study on the suppression of inlet blockage in rotating detonation combustor by porous-wall

The rotating detonation combustor (RDC) is renowned for its ability to provide substantial pressure gains. Nonetheless, during the stable operation of the RDC, the high-pressure rotating detonation wave (RDW) at the combustor inlet induces an increase in upstream chamber pressure, ultimately compromising engine stability and performance. To fully harness the performance advantages of the turbine-based continuous rotating detonation engine (TBCRDE) while maintaining engine stability, a porous-wall RDC has been developed to alleviate intake blockage and mitigate upstream chamber pressure rise. The operating modes, pressure rise characteristics, and performance parameters of both the porous-wall RDC and the reference configuration were systematically evaluated across varying air flow rates and nozzle designs. This analysis concentrated on the operational characteristics of the porous-wall RDC and its mechanisms for suppressing upstream chamber pressure rise. The findings reveal that the porous-wall RDC significantly extends the stable operating range and effectively reduces upstream chamber pressure rise by minimizing intake blockage. Specifically, the stable operating range is enhanced by 50 % at an outlet area ratio of 0.33, with stable rotating detonation combustion achieved at an outlet area ratio of 0.25. At an air flow rate of 1 kg/s and an outlet area ratio of 0.33, the chamber pressure rise is optimally suppressed, demonstrating a maximum reduction of approximately 16.4 %. The total pressure recovery coefficient of the combustor was analyzed, taking into account both intake loss and combustor pressure gain capabilities, and the propulsion performance of the two configurations was compared. The porous-wall RDC effectively reduces intake loss while slightly diminishing combustor pressure gain capability, resulting in a marginal increase in the total pressure recovery coefficient. Although this leads to a slight reduction in propulsion performance during chamber pressure rise suppression, the overall engine matching environment benefits from enhanced matching stability. Consequently, other engine components experience a reduced performance decline. Therefore, the implementation of a porous-wall structure is anticipated to improve the overall propulsion performance of the engine.

Influence of fuel inhomogeneity on detonation wave propagation in a rotating detonation combustor

Rotating detonation combustor (RDC) is a form of pressure gain combustion, which is thermodynamically more efficient than the traditional constant-pressure combustors. In most RDCs, the fuel–air mixture is not perfectly premixed and results in inhomogeneous mixing within the domain. Due to discrete fuel injection locations, local pockets of rich and lean mixtures are formed in the refill region. The objective of the present work is to gain an understanding of the effects of reactant mixture inhomogeneity on detonation wave structure, wave velocity, and pressure profile. To study the effect of mixture inhomogeneity, probability density functions of fuel mass fractions are generated with varying standard deviations. These distributions of fuel mass fractions are incorporated in 2D reacting simulations as a spatially/temporally varying inlet boundary condition. Using this methodology, the effect of mixture inhomogeneity is independently investigated to determine the effects on detonation wave propagation and RDC performance. As mixture inhomogeneity is increased, detonation wave speed, detonation efficiency, and potential for pressure gain all decrease, ultimately leading to the separation of the reaction zone from the shock wave.

Analysis of coupling supersonic turbine stage with rotating detonation combustor under different turbine parameters

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.

On the Challenges of Integrating a Rotating Detonation Combustor with an Industrial Gas Turbine and Important Design Considerations for Row-1 Blades

Abstract In an effort to increase the efficiency and performance of gas turbine power cycles, pressure gain combustion (PGC) has gained significant interest. Since Rotating Detonation Combustors (RDC) can provide a quasi-steady mode of operation, research has been triggered to integrate RDC with power-generating gas turbines. However, the presence of subsonic and supersonic flow fields which are generated due to the shock waves that stem from the detonation wave front and the highly non-uniform temperature and velocity profiles may cause a depreciation in the turbine performance. The current study seeks to investigate the challenges of integrating the RDC with nozzle guide vanes (NGV) of an industrial, can-annular gas turbine and attempts to understand the major contributors that impact efficiency and identify the key areas of optimization that need to be considered for maximizing performance. The RDC was integrated with the NGVs through a non-optimized straight duct-type geometry with a diffuser cone. 3-Dimensional Numerical analyses were performed to investigate sources of total pressure loss and to understand the unsteady effects of RDC which contribute towards the deterioration of performance. The entropy generation at different regions of interest was calculated to identify the major irreversibilities in the system. Also, total pressure and temperature distribution along the radial direction at the exit of the transitional duct is presented.

Quantitative Measurements in the Exhaust Flow of a Rotating Detonation Combustor Using Rainbow Schlieren Deflectometry

Abstract This study employs Rainbow Schlieren Deflectometry (RSD) to characterize the unsteady, supersonic/subsonic exhaust plume of a Rotating Detonation Combustor (RDC). Firstly, RSD images are analyzed to quantify the frequency and strength of flow oscillations and their relationship to the detonation wave. Secondly, a 3D tomographic algorithm is used to obtain the local 3D density field across the whole region of interest (ROI). The tomographic analysis relies upon wave rotation to infer projection data of the 3D exhaust plume at multiple view angles using a single RSD/camera system and was previously validated using phantom data from computational fluid dynamics analysis of an RDC. The annular RDC operated on methane and 2/3 O2 - 1/3 N2 oxidizer mixture is equipped with a converging nozzle to pressurize the combustion chamber. The product flow exiting the nozzle throat expands across an unoptimized conical aerospike attached to the center body of the RDC. RSD images provide a temporal resolution of 369 ns and spatial resolution of 100 ?m in a 6.4 high and 25.6 mm wide ROI of the exhaust plume. A rainbow filter is calibrated to convert hue in color schlieren images into deflection angle data. This data is used to characterize the unsteady flow oscillations that show excellent agreement with PCB pressure probe measurements acquired inside the combustion chamber. Tomographic analysis yields a 3D local density field that shows distinct features, consistent with published numerical simulations of the RDC exhaust plume.

Performance Evaluation of Different Rotating Detonation Combustor Designs Using Simulations and Experiments

Abstract Rotating Detonation Engines (RDEs) present a potential avenue for enhancing the efficiency of gas turbine combustors through the utilization of a detonation-driven process. However, integrating an RDE with a downstream turbine poses a significant challenge due to the shock-laden and highly unsteady nature of the RDE exit flow, coupled with a high degree of flow periodicity. In contrast, gas turbines are designed to operate with relatively small velocity and temperature fluctuations at the turbine inlet. The objective of the present study is to develop an understanding of how geometric profiling can improve RDE performance and mitigate exhaust flow field unsteadiness. Three RDE designs are analyzed: an annular combustor with a constant cross-sectional area, an annular combustor with a converging nozzle near the exit, and an annular combustor with a rapid to gradual (RTG) area convergence. Three-dimensional unsteady reacting simulations are conducted for each configuration using the same fuel-oxidizer composition and mass flow rate condition. All simulations are validated against experimental measurements, and RDE performance is assessed based on total pressure gain/loss, unsteadiness at the exit of the RDE, and detonation effectiveness. Results show significant performance improvement for both the convergent nozzle and the RTG profile compared to the constant cross-sectional area configuration, and the RTG design performed better than the convergent nozzle design. Therefore, strategically constricting the flow in an RDE can be used to optimize the performance of an RDE and should be considered in future designs.

Performance Characterization of a Radial Rotating Detonation Combustor with Axial Exhaust

Abstract This experimental study characterizes the performance of a radial rotating detonation combustor (RDC). An aerospike nozzle for rocket propulsion has been integrated into the center of the combustor, although the same combustor could also be coupled with turbomachinery. The radial RDC utilized a rapid to gradual (RTG) area change in the flow direction to effectively confine the detonation region close to the inlet plane and to improve the uniformity of the flow exiting the RDC. Three test cases were analyzed, (a) a baseline case at a total reactant mass flow rate, m˙ = 0.136 kg/s and equivalence ratio, ϕ = 0.6, (b) a higher reactant flow rate, m˙ = 0.318 kg/s and ϕ = 0.6, and (c) a higher ϕ = 0.8 at m˙ = 0.318 kg/s. All tests were conducted using methane and a 67% oxygen and 33% nitrogen (by mole) oxidizer mixture. Measurements were acquired using CTAP probes inside the combustion channel and along the aerospike to characterize the performance, PCB and ion probes near the detonation region to identify wave modes and their variations during the test, and thrust measurements using a six-axis force sensor. Results show highly complex wave modes with multiple co-rotating and/or counter-rotating waves depending upon the reactant flow rate. The pressure and thrust measurements are consistent with the wave mode analysis. In general, a positive (combustor only) pressure gain was inferred when losses associated with the injection system were excluded.

Investigation of rotating detonation gas turbine cycle with different combustor passage areas

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.

Time/frequency domain analysis of detonation wave propagation mechanism in a linear rotating detonation combustor

Controlling the rotating detonation combustor (RDC) is crucial for enhancing propulsion system safety, efficiency, and overall performance. Swift and precise identification of the operational mode of RDC is of great significance for refining the propulsion dynamics of detonation-based technologies. In this study, two-dimensional numerical simulations are performed to investigate time-domain characteristics and frequency-domain characteristics of various operational modes of the RDC, hydrogen–air mixtures are used as fuel, and fast Fourier transformation (FFT) is applied to the analysis of frequency-domain. The alteration of the RDC operational mode is achieved through adjustments to the chemical equivalent ratio (ER) of the fuel, spanning from 0.52 to 3.50. Our findings reveal four distinct operational modes within this equivalent ratio range: ignition failure (ER ≤ 0.52 and ER ≥ 3.50), collision failure (ER = 0.85, 1.60 and 1.65), single wave (0.53 ≤ ER ≤ 0.80 and 1.70 ≤ ER ≤ 3.40) and multiple wave mode (0.90 ≤ ER ≤ 1.50). For the ignition failure mode, the detonation waves in RDC decouple from the flame and fail after ignition. In the collision failure mode, the inability of detonation waves to propagate sustainably within the RDC is evident. Frequency-domain analysis underscores the significance of a characteristic low-frequency (0.3 kHz) in distinguishing this mode from others. For single wave mode, it represents the steady and continuous propagation of a detonation wave in the RDC. Frequency-domain analysis reveals characteristic frequencies approximately equal to integer multiples of the propagation frequency (fD), where fD ∼ 10fD serves as the primary characteristic frequencies. The multiple wave mode can be categorized into two subtypes: the forward and reverse single detonation waves appear alternately or the forward and reverse single detonation waves finally co-direction double detonation waves. In the former subtype, characteristic frequencies are integer multiples of the fD, but the amplitude distribution is different with single wave mode. In contrast, primary characteristic frequencies of the latter subtype are fD ∼ 4fD, accompanied by two adjacent characteristic frequencies near each integer multiples of fD. These findings promise advancements in the controllability of RDC and lay the foundations for the development of RDC control systems.

Comparative Analysis of Total Pressure Measurement Techniques in Rotating Detonation Combustors

Abstract Current total pressure measurement techniques in RDCs are based on different assumptions and therefore show different applicability for specific RDC operating conditions, and few studies have directly compared these techniques. Therefore, this study comprehensively tested three total pressure measurement techniques: the direct Kiel probe method, the Mach-corrected CTAP method, and the equivalent available pressure (EAP) method under different RDC geometries and mass flow rates, and compared them with their corresponding uncertainties considered. The results show that for all tests in this study, the EAP method shows the largest uncertainty range up to 24%, which is mainly contributed by the load cell calibration process, while the direct Kiel probe method has the lowest uncertainty range, which is consistently below 7%. These uncertainties were incorporated into the comparison between the three techniques via Gaussian process regression, showing that the direct Kiel probe method and the Mach-corrected CTAP method can present EAP-like total pressure. In particular, the total pressure of the SWCC and L modes measured by the three techniques is very comparable. This work shows that the comparability of total pressure techniques depends on the specific RDC environment, and provides the possibility to evaluate the RDC performance with the simplest implementation.

Performance and Exergy analysis of TurboJet and TurboFan configurations with Rotating Detonation Combustor

Conventional gas turbine is reaching its saturation level and further improvement in the cycle efficiency is possible through redesigning the cycle. One of the effective methods is to introduce a Pressure Gain Combustor (PGC) in place of the conventional deflagration combustor as PGC contributes to a considerable gain in thermal efficiency. Pulsed Detonation Combustor (PDC) and Rotating Detonation Combustor (RDC) are the two most prevalent combustor designs for pressure gain combustion. The major challenges of integrating PGCs to existing conventional cycles are the highly unsteady exhaust flow from PGC, high exhaust temperatures from PGC, potential back flow from combustor to compressor. As turbines are designed for much steadier flows, this unsteady exhaust flow and high temperature exhaust gas from PGC is challenging for the turbine of the cycle and its design. One of the solutions is to have an ejector between the PGC and turbine in order to reduce the flow fluctuations, flow velocity and flow temperature. In this work, different cycle layouts of Turbojet and Turbofan engines working with RDC are described. Low order RDC and ejector models are used in this paper. As the ejector is being used, the compressed air from the compressor is split for RDC, ejector, turbine blade cooling. The different cases for the compressed air split are also examined. The performance parameters of the engines are evaluated for different compressor pressure ratios, fan pressure ratios, bypass ratios. Finally, an exergy analysis is performed on all components.

Effects of the outlet contraction ratio on the performance of liquid kerosene/air two-phase rotating detonation combustors

Experiments and numerical simulations were conducted across a range of contraction ratio to investigate the detonation combustion characteristics of liquid aviation kerosene/air rotating detonation combustor. The research indicates that extremely low and high contraction ratio adversely affect the stable operation of the two-phase rotating detonation combustors. At a contraction ratio that is too low (19.03 % in the RDC discussed in this paper), the atomization and mixing properties of the two-phase fuel are poor, with a lower reaction pressure in the fresh mixture, hindering the initiation of detonation waves. On the other hand, a contraction ratio (85.7 % in the RDC discussed in this paper) that is too high enhances the energy of the detonation wave; it triggers oblique shockwave feedback, preventing the fresh mixture from entering the combustor and thus leading to the quenching of detonation waves. With an increase in the contraction ratio of outlet configurations, there is a noticeable rise in both the peak pressure and combustion efficiency of detonation waves. In contrast, the wave height gradually decreases, and the wave velocity shows a pattern of initially increasing and then decreasing. Along the axial position inside the two-phase rotating detonation combustor, the upstream detonation wave is caused by the reflected shockwaves increasing the pressure of the upstream fresh mixture. In contrast, the downstream detonation wave results from increased fresh mixture pressure and enhanced fuel atomization and mixing due to the increased contraction ratio of outlet configurations and wave feedback. So, the detonation wave exhibits a strong-weak-strong distribution characteristic.

Research on the Combustion Mode and Thrust Performance of Rotating Detonation Scrarmjet Engines

Decades of research in hypersonic science and engineering have shown that scramjet engines are the preferred choice for powering hypersonic vehicles. In these engines, oblique shock combustion and rotating detonation combustion are two common combustion modes. This article primarily focuses on the numerical simulation and theoretical analysis of these two combustion modes. The research indicates that as the combustible mixture injected into the combustion chamber increases, the combustion mode transitions from deflagration combustion to self-sustaining detonation combustion and then to forced detonation combustion to maintain stable combustion in the chamber. Under self-sustaining detonation combustion mode, as the amount of combustible mixture injected into the chamber increases, the number of detonation waves gradually increases from one to multiple and the angle between the detonation waves and the inflow changes from acute to right angles. In the forced detonation combustion mode, the minimum shock wave intensity required to initiate detonation combustion can be obtained by drawing a tangent from the Rayleigh line. The structure generating oblique shock waves in the combustion chamber needs to match the shock wave intensity. In the detonation combustion mode, both the exit total temperature and specific impulse of the combustion chamber increase with the increase in Mach number at the inlet, while the pressure ratio decreases. This article provides a reference for the design of the combustion chamber in supersonic combustion ramjet engines.

Flow and Performance Characterization of Rotating Detonation Combustor Integrated with Various Convergent Nozzles

In this study, convergent nozzles of various area ratios (ARs) are used downstream of an annular rotating detonation combustor (RDC) to increase the operating pressure and approach sonic conditions at the nozzle throat. Reactant methane and oxygen-enriched air (67% O2 and 33% N2 by volume) are supplied in counterflow arrangement from two separate plenums located at the base of the RDC annulus. Based on experimentation, a total mass flow rate of 0.32 kg/s was chosen to achieve stable, single-wave mode RDC operation for all test cases, allowing for one-to-one comparisons. The internal performance of the RDC was characterized by ion probes and pressure measurements (wall static and oscillating) in supply plenums and across different axial locations of the combustor. Particle image velocimetry (PIV) at 100 kHz was utilized to measure axial and circumferential velocity components within a two-dimensional region of interest located downstream of the converging nozzle exit. Results show higher internal performance of the RDC with increasing AR of the convergent nozzle. PIV measurement illustrated that the flow oscillation amplitudes decrease with an increasing AR of the converging nozzle. The exit flow contained significant nonuniformity and unsteadiness even with a converging nozzle of AR 2.0, indicating incomplete choking of the flow at the nozzle throat.

Numerical simulations and theoretical analysis of the forward shock wave in a non-premixed air-breathing rotating detonation combustor

The forward shock wave induced by the detonation wave propagates in the inlet and restrains the air from being injected into the rotating detonation engine, and the stable continuous propagation of the detonation wave is affected. Numerical simulation and theoretical derivation of air flow velocity and detonation wave height were carried out in a rotating detonation combustor. The influence of the forward shock wave on air flow velocity and air flow velocity on detonation wave height was analyzed. A non-premixed air-breathing rotating detonation combustor was numerically simulated with the two-dimensional Navier–Stokes equations and a kerosene/air reaction model based on the RYrhoCentralFoam solver. Analytical equations were derived and validated to calculate the air flow velocity across the forward shock wave and the height of the detonation wave. The results show that the air flow velocity across the forward shock wave is inversely proportional to the detonation wave velocity, incoming flow temperature, and pressure ratio at the forward shock wave. In addition, the angle of the forward shock wave is inversely proportional to the velocity of the detonation wave and incoming flow and is proportional to the incoming flow temperature and the pressure ratio. The forward shock wave leads to a blocked zone of injection in the combustor, as the inflow velocity decreases or even reverses across the shock wave. As a result, the forward shock wave induced by the detonation wave restrains the recovery of the reactant fill zone, which has an influence on the height of the detonation wave.