An interstage turbine burner (ITB) integrates a combustor into the transition between the high and low pressure turbines. It is a promising technology for enhancing gas turbine engine efficiency and reducing emissions. The ITB with trapped vortex cavity stabilization (TVC-ITB) presents outstanding combustion stability and thus has attracted extensive attention. Due to the particularity of its position, the downstream low pressure turbine guide vane will have a significant impact on the performance of the ITB. However, this critical coupling effect has been ignored in previous studies. This study investigates TVC-ITB configurations with and without a turbine vane. It combines experimental measurements, numerical simulations, and time-scale analysis to elucidate the vaneinduced ignition and lean blowout mechanisms. Results demonstrate that turbine vanes enhance TVC-ITB flame stability. The ignition equivalence ratios are reduced by 13.7%-40% and the lean blowout equivalence ratios by 4.8%-22.8% relative to the vaneless configuration. Numerical simulations reveal that vane blockage effects and pressure-suction surface interactions increase local static pressure, thereby altering the combustor flow field. The primary recirculation zone is expanded by 66.7%, resulting in increased residence time for chemical reactions. Time-scale analysis reveals a 70% increase in mixing residence time in the combustor with vanes. The critical stability criterion (Damko & uml;hler number >= 1) is ensured to be met at low fuel equivalence ratio conditions.

The thermal oxidation coking of endothermic hydrocarbon fuels in realistic fuel systems is a key factor limiting heat sink utilization and the reliability of aviation propulsion and thermal management systems. This study investigates the thermal oxidation coking deposition of aviation kerosene RP-3 in tubular flow, with a specific focus on how flow regime, characterized by the Reynolds number, governs the diffusion and deposition of coking precursors. To quantify the effect of hydrodynamics, the Reynolds number was systematically varied from 1500 to 16,000 by adjusting inlet velocity, mass flow rate, and tube diameter. The results show that the dominant mechanism of coke formation alternates between temperature control and diffusion control depending on the flow regime. In the laminar flow regime (Re < 2300), temperature is the primary factor governing coke formation, whereas in the turbulent regime (Re > 2300), the effect of physical diffusion on coking deposition strongly depends on tube diameter, bulk fuel temperatureand wall temperature. Based on the experimental trends observed over this wide Reynolds number range, the study infers how increasing Re modifies the relative roles of thermal conditions and mass transfer in driving coking, thereby clarifying the dependence of thermal oxidation behavior on flow regime. The insights obtained provide a practical basis for selecting appropriate Reynolds number ranges and for optimizing fuel system geometry and operating conditions to mitigate coking in advanced aviation thermal management applications.

Aero-engine emissions have attracted more interest in environmental pollution. The staged combustion in the developing combustors has been recognized as a potentially effective approach to achieving low emissions. The swirling-flow single trapped vortex combustor (SSTVC), integrates the superior mixing capability of swirl combustion, and the broad operational range of trapped vortex combustion. Experiments on the SSTVC demonstrate the role of fuel staging under various operating conditions. A three-dome rig is employed to obtain systematic data on flame structures, outlet temperature profiles, combustion efficiency, and emission index of carbon monoxide, unburned hydrocarbons, and nitrogen oxides. It is found that the combustion efficiency improves progressively with the cavity equivalence ratio under the Only Cavity Mode, reaching a maximum of 96.8% as the cavity equivalence ratio approaches 1.4. In the Two Stage Mode, efficiency further improves to 99.9%, while the nitrogen oxides emission index remains as low as 7.2 g/kg even at the highest fuel air ratio. Fuel staging strongly influences flame propagation and temperature uniformity, with the cavity equivalence ratio approaches 1.4 identified as the optimal value for stable combustion and low emissions. Finally, the empirical correlations for nitrogen oxides emissions in both modes are constructed. It can be used as a predictive tool for designing low-emission aero-engine combustors.

To investigate the effect of flow field, particularly specifically the large-scale axial vortex, on the combustion performance of a triple-swirler combustor, two triple-swirlers (referred to as S1 and S2) were developed to establish two distinct flow fields in a gas turbine combustor. The flow field of S1 is characterized by a larger recirculation zone at the center spanwise cross-section and a chaotic divergent flow at the axial cross-sections. In contrast, the flow field of S2 features a smaller recirculation zone at the center spanwise cross-section and a more uniform flow pattern with a prominent large-scale axial vortex at the axial cross-sections. Experimental investigations were conducted at atmospheric pressure and a temperature of 473 K to assess the combustion performance of the combustor under these differing flow fields. The evaluation of combustion performance included lean ignition performance, lean blowout limit, combustion efficiency, and pattern factor. The findings indicate that the combustion behavior of the combustor is strongly influenced by the flow field. S2 outperforms S1 in terms of lean ignition performance, lean blowout limits, and pattern factor under most operational conditions. However, when the fuel-air ratio (FAR) is below 0.02 or above 0.4, S1 exhibits superior combustion efficiency compared to S2. Conversely, for the remaining FAR conditions, S2 achieves higher combustion efficiency than S1.

In this paper, ignition and lean blowout (LBO) performance of an annular mixed-flow trapped vortex combustor (MTVC) were investigated under sub-atmospheric pressure. Experimental investigations were conducted to evaluate the ignition and lean blowout (LBO) performance under various sub-atmospheric pressure conditions. The findings indicate that MTVC has excellent ignition performance, the ignition pressure of the combustor can reach 35.3 kPa under ambient temperature, and the corresponding altitude is close to 8,000 m. As the pressure drops, the range of ignitable velocities diminishes, leading to increase in ignition fuel to air ratio (FAR) and LBO FAR. These shifts can be attributed to the more challenging combustion conditions and heightened inlet velocity resulting from reduced pressure. Notably, the ignition and LBO performance see substantial enhancements with rising temperatures. However, the positive impact of elevated temperature hardly compensated for the detrimental impact of lower pressure on ignition. There is minimal noticeable impact of the inlet Mach number on the ignition and LBO performance. Numerical simulations are carried out for both unreactive and reactive flow to verify the experiment results.

The tandem Turbine-Based Combined-Cycle (TBCC) engine is critical for enabling wide-range flight vehicles. Its combined combustor must accommodate rapidly varying inlet conditions across three modes, including afterburner, transitional and ramjet, each characterized by significantly different flow fields and combustion performance. Meanwhile, the flow field characteristics within the trapped-vortex cavity flameholder (TVCF) during TBCC mode transition remain unclear. In this study, the flow field characteristics downstream of the TVCF were investigated using particle image velocimetry (PIV) under four representative Rear Variable-Area Bypass Injector (RVABI) angles α = 0°, 5°, 14°, and 23°.,The results indicate that, a stable trapped vortex structure is maintained within the cavity, a large-scale recirculation zone forms behind the struct. The fundamental flow field morphology remains consistent across all tested angles. However, at α = 5°, the primary cavity vortex exhibits flattening and displacement. When α ≥ 14°, the deflection angle of the high-velocity mainstream becomes developed, enhancing the recirculation zone behind the struct: the recirculation zone behind the struct enlarges significantly, and the low-velocity region expands. The study reveals the physical mechanism by which the inflow affects the flow field structure: the momentum of the cavity fore inlet flow determines the radial position of the recirculation zone behind the struct. Simultaneously, the high-speed deflected inflow reduces the static pressure in the flameholder wake region, thereby weakening the squeezing effect on the recirculation zone and driving its expansion under a lower adverse pressure gradient.

Subsonic-supersonic mixing inflows in combined-cycle engines pose unique flame stabilization challenges compared to uniform inflows. The flame evolution characteristics and velocity vector distribution downstream of an evaporative flameholder was investigated using a high-speed camera and particle image velocity system. Comparative analysis of subsonic uniform inflow and subsonic-supersonic mixing inflow reveals three key findings: (1) Both conditions exhibit the same spark ignition process: kernel generation, flame growth, and stable combustion, with flames anchored in the recirculation zone; (2) Asymmetric vortices induced by subsonic-supersonic shear layer reduce the stable flame projected area by 81.14 % compared to symmetric vortices in uniform flows, and increase the ignition delay time by 18.97 %; (3) Distinct extinction modes emerge: sequential attached-detached flame extinction under subsonic uniform inflows, and direct attached flame extinction under subsonic-supersonic mixing inflows. These results provide critical criteria for flameholder design in combined-cycle engines with extreme flow gradients.

Conventional bluff-body flameholders in afterburners suffer from limited flame propagation from the recirculation zone into the mainstream, restricting combustion space utilization and efficiency. This paper proposes a shear-layer transport-enhanced flameholder (STF). Unlike lobed mixers installed upstream to enhance inlet airflow mixing in afterburners, the STF integrates small-scale lobed devices directly onto the sidewalls of the flameholder. Streamwise vortices are generated at the trailing edge, specifically enhancing mass and heat exchange between the recirculation zone and the mainstream flow. The flow field, fuel spray, and combustion characteristics of the STF are experimentally investigated. Additionally, an aerodynamic wake width evaporating flameholder (AWEF) and an AWEF with plain plates (AWEF_P) are set as contrast flameholders for comparison. Results reveal that the STF introduces spatially alternating flow patterns. The crest section promotes outward transport from the recirculation zone to the mainstream, while the trough section facilitates inward transport from the mainstream to the recirculation zone. This mechanism is fundamentally distinct from conventional lobed mixers, which target bulk stream mixing without direct interaction with a recirculation zone. Under the tested conditions, the STF crest section increases fuel penetration depth by 21.9% and 56.6%, and flame width by 20.6% and 50.8%, relative to the AWEF and AWEF_P, respectively. Flame projection area is enlarged by 27.1% and 48.1%, and combustion efficiency is improved by up to 10.4% at an equivalence ratio of 0.62. The present study is conducted as a proof-of-concept investigation. A new design method is proposed to enhance afterburner performance through localized aerodynamic modifications rather than global scaling or auxiliary systems.

Unmanned aerial vehicles (UAVs) have been widely deployed in both military and civilian domains. As future UAV platforms aim for higher flight speeds and extended endurance, their propulsion systems must maintain stable performance over a broad operational envelope, particularly under high throughflow inlet conditions. In this study, a novel small gas turbine combustor is proposed. A prototype combustor was designed and experimentally evaluated for its combustion characteristics. The results demonstrate that this combustor not only preserves structural compactness but also delivers favorable performance under high throughflow conditions, including low total pressure loss, broad ignition and blowout limits, uniform temperature distribution, and high combustion efficiency. Specifically, at an inlet Mach number of 0.30, the total pressure loss was measured at 4.85%. The ignition and blowout limits were recorded at fuel-air ratios of 0.0052 and 0.0044, respectively, under an inlet Mach number of 0.35. At atmospheric pressure and an inlet Mach number of 0.35, the temperature distribution factor at the exit was below 0.14, with peak combustion efficiency reaching 97.45%. Overall, under the present experimental conditions, the proposed combustor shows promising potential for compact high-throughflow combustor.

The combustion process in rocket-assisted subsonic ramjet engines represents a key advancement in integrated aerospace propulsion, particularly for embedded rocket-based systems. These engines offer the potential to improve combustion performance at altitudes of 25-35 km. However, the significant temperature and velocity differentials between the rocket jet and the subsonic ramjet flow restrict heat and mass transfer. Investigating the relationship between combustion performance and inlet parameters under subsonic-supersonic mixing conditions offers a promising approach to enhancing thrust performance. This study introduces subsonic and supersonic airflow mixing via a flat-plate shear layer in a rectangular channel, with an evaporative flameholder placed centrally to assess combustion. Results reveal that combustion efficiency decreases as the equivalence ratio exceeds 0.2, while the static temperature ratio has minimal impact on efficiency but strongly influences the maximum flame stabilization limit. As the temperature ratio increases from 1.30 to 1.80, the flame limit narrows from 1.656 to 0.237. Higher pressure ratios initially enhance combustion efficiency and flame coverage but eventually cause a decrease. The flame limit broadens from 0.900 to 1.626 as the pressure ratio increases from 1.12 to 1.50. While Mach number changes have little effect on efficiency, the flame limit exhibits an initial rise followed by a drop. Novel findings include an asymmetrical flame pattern and a "Z" shaped outlet temperature distribution, contributing to optimized combustion strategies for combined-cycle engines.

There is a high risk for bluff-body stabilized flame to suffer blowing off under some practical operation conditions, and the triggering mechanism of the lean blow-off (LBO) is still ambiguous. The blow-off is a transition process from flame to global extinction, inevitably accompanied by flame front changes. This study investigated two potential changes to the flame front, including the merging and the local extinction of the flame fronts. The LBO behavior in a two-dimensional (2D) bluff-body stabilized premixed flame was investigated by 5 kHz OH*-chemiluminescence (CL). Flame necking and a following pinch-off leading to permanent downstream extinction and reaction within the recirculation zone (RZ) were observed. In addition, the 10 Hz particle image velocimetry (PIV) and planar laser-induced fluorescence (PLIF) of OH radical were also applied simultaneously, and the flame characteristics of the stable and near-the-LBO conditions were recorded and compared. Statistical analysis regarding strain rates, OH-PLIF boundary gradients, velocity directions close to the local extinction sites, and the distribution of turbulent velocity fluctuations relative to mean progress variable contours are provided. The results demonstrate that the flame pinch-off predominantly occurs in the stagnation point region (SPR), and the flame front gradients in this region under near-LBO conditions are similar to those in stable states, indicating that local extinction is not the primary cause of pinch-off. Local extinction mainly occurs in the upstream shear layer, where unburnt gas briefly enters the burnt region but is rapidly reignited by high-temperature gas. Moreover, the overlap between the flame front and high-velocity fluctuation zones in the SPR promotes the merging of flame branches, leading to pinch-off, which was confirmed based on the critical images before pinch-off. This study can help optimize the active control strategy of the combustion stability in 2D bluff-body stabilized premixed flames.

Understanding of flow characteristics of combustor under sub-atmospheric conditions is critical for plateau ignition and high-altitude relight of aero-engine. In this study, the impacts of sub-atmospheric operating pressure p(op) on flow characteristics of a twin-stage swirler combustor have been experimentally and numerically investigated. Specifically, two reference velocities V-ref, i.e., 3.0 and 5.5 m/s, are considered. The sub-atmospheric p(op) ranges from 0.08 to 0.02 MPa, while the p(op) of 0.11 MPa represents an atmospheric condition. Results show that as the p(op) decreases from the atmospheric pressure to sub-atmospheric ones, there exist two flow patterns, i.e., Y- and V-type. The critical pressure of their transition is associated with the V-ref. A higher V-ref has a lower critical pressure, indicating its higher capacity to withstand the sub-atmospheric pressure effect. A non-linear map between the radial pressure difference across the second stage swirler airflow and the p(op) is observed, except in the region near the swirler exit. The flow pattern transforming from Y-type to V-type leads to an expansion of the central toroidal recirculation zone and an increase in the reflux ratio. Further, the simplified force balance analysis by reductio for swirl flow shows that both the centrifugal and pressure differential forces should reduce linearly as decreasing the p(op), but the pressure differential force actually decreases non-linearly and more rapidly than the centrifugal force when the p(op) approaches some certain value. Consequently, the swirl flow expands radially outward for force balances, accounting for the transition of the flow pattern from Y-type to V-type.

This paper presents a novel combined cycle propulsion scheme based on parallel combustion. The parallel combustion cycle engine (PCCE) comprises turbine engine components that form an internal bypass flow path, an outer combustor of parallel combustion set up in the external bypass, and an intake adjustment device set up in the intake to regulate the air input of the internal and external bypasses. Furthermore, a shunt adjustment body is installed at the outlet of the internal bypass compressor, which allows a portion of the high-pressure airflow compressed by the compressor to be diverted into the outer combustor. This can be utilized in the propulsion system of Ma0–7 wide-speed range flight vehicle. The paper constructs the principle structure of PCCE, describes the four operating modes and its working principle, and establishes its ideal thermodynamic cycle process. The results of the thermodynamic cycle analysis demonstrate that the PCCE in dual-combustion mode (Ma0.7–2) exhibits a broader speed range and higher thermal efficiency. In the transition mode (Ma2–4), the shunt adjustment body facilitates the utilization of compressed air to enhance specific thrust at Ma2.5 by 2.1

The diffuser plays a critical role in enabling efficient combustion in an aeroengine afterburner by decelerating and pressurizing airflow. However, deteriorating diffuser inlet conditions and structural size limitations have made flow separation more likely to occur within the diffuser. This study explores an active control method using boundary layer blowing to suppress flow separation in the diffuser. An afterburner blown diffuser is designed and experiments are conducted to determine the effects of divergence angles (12°, 15°, 18°, and 24°) and blowing flow ratios (0%–5.48%) on the internal flow characteristics. The results indicate that the effect of the blowing flow ratio are not always consistent across different divergence angles. For relatively smaller divergence angles (e.g., 12°, 15°, and 18°), a moderate blowing flow ratio can effectively suppress the onset of flow separation, but excessive blowing flow ratio induces flow separation on the opposite wall. At larger divergence angle (e.g., 24°), the diffuser exhibits four distinct flow patterns as the blowing flow ratio increases, which differ completely from those observed at the other three divergence angles. At this divergence angle, the size of the separation vortex within the diffuser decreases as the blowing flow ratio increases, and its position shifts significantly. Based on the study of the flow characteristics of the diffuser at various divergence angles, six flow patterns within the afterburner blown diffuser are summarized, and the changes in flow patterns are analyzed. Despite multiple flow patterns, total pressure loss decreases with increasing blowing flow ratio across all divergence angles.

Within the context of exploiting efficient cooling methods for advanced gas turbine combustors, understanding the fundamental physics for impinging-film cooling under various operational conditions is of significance. In this paper, impacts of different interaction modes between coolant and hot mainstream on the impinging-film cooling are quantitatively evaluated via active subspace (AS) method. Three interaction modes are considered, i.e., a transitional flow (TF), a turbulent boundary layer (TBL) and a wall jet (WJ). Sensitivities and uncertainties of cooling effectiveness eta with respect to coolant mass flow rate Mcand model parameters (C mu, C epsilon 1, C epsilon 2, Prtw) are estimated. Results show that 1-D active subspaces are sufficient to map eta in TF and TBL modes while high-dimensional active subspaces are warranted for WJ mode, indicating its more complicated interaction between cold and hot flows. eta is the most sensitive to and dominated by Mcespecially in the slot and far fields, and turbulent effects are more significant in the near field than other places. Specifically, an increase in Mcor a decrease in turbulent level monotonously improves eta in TF and TBL modes while initially increases eta then reduces it for WJ mode. Further analysis of flow characteristics of WJ mode demonstrates that the reduction in eta results from the strengthened impingement-induced streamwise vortexes and thereby, enhanced mixing between the coolant and mainstream. The propagation of the input uncertainty to eta is space-and operational condition-dependent, consistent with the evolution of active subspaces.

This paper proposes a combined flameholder comprising a pilot flameholder and a bluff body flameholder, interconnected by a radial V-gutter. The configuration of this combined flameholder is widely utilized in current afterburners and ram combustors, and its outstanding performance in future advanced combined cycle engines has been proved. To avoid the issue of uniform airflow at the inlet observed in most previous studies, this research installed a diffuser to simulate relatively realistic inlet airflow conditions of combustor. Particle Image Velocimetry (PIV) and high-speed camera are employed to measure the flow field structure of the combined flameholder, as well as the ignition flame evolution process under inlet Mach number of 0.3 - 0.5, temperature of 500 - 700 K, and pressure of 0.05 - 0.1 MPa. This paper reveals the combustion organization method of the combined flameholder through combining flow field characteristics and the ignition flame evolution process. Furthermore, the research findings indicate that flame evolution processes, flame projection area, and ignition delay time are significantly influenced by the inlet conditions. The variations in these flame evolution characteristics with different inlet conditions are summarized and the contributing factors are analyzed. The results of the paper are expected to serve as a reference for the optimization and design of flame stabilization systems in current and future afterburners and ram combustors.

Subsonic uniform and supersonic mixing flows are widely present in the combustion chamber of a multi-mode combined scramjet engine, where the significant pressure, velocity, and temperature gradients of the two impose severe limitations on flame propagation. The vortex structures downstream of an evaporative flameholder were experimentally measured under subsonic-supersonic mixing inflow and subsonic uniform inflow. Results indicate that, under subsonic-supersonic mixing inflow conditions, an asymmetric vortex structure in the recirculation zone is formed due to the presence of a subsonic-supersonic shear layer, which exerts compressive or stretching effects on the recirculation zone downstream of the flameholder, in contrast to the symmetric dual-vortex structure observed under subsonic uniform inflow conditions. The asymmetric recirculation zone displays fluid being entrained from one vortex to another, with the degree and direction of entrainment dependent on the supersonic expansion state. Additionally, a simplified flow field structure is proposed for comparing the differences between the subsonic-supersonic mixing inflow and subsonic uniform inflow, along with the introduction of five regions and two stages to describe the flow field structure downstream of the flameholder under subsonic-supersonic mixing inflow. Analysis of the flow characteristics downstream of the flameholder under subsonic-supersonic mixing inflow and subsonic uniform inflow conditions can offer design insights for the combustion scheme of aerospace-integrated propulsion systems.

The effect of the inlet Mach number ratio and the static temperature ratio between the supersonic and the subsonic flows on the distortion characteristics of the recirculation zone downstream of the flameholder was experimentally investigated. The fluid structure was visualized and measured by a calibrated particle image velocimetry system under the subsonic–supersonic mixing flow, which was constructed by a flat plate to divide the rectangular flow channel. The present results reveal that the fluid structure downstream of the flameholder remains stable over a wide range of the inlet Mach number ratio and exhibits insensitivity to changes in the inlet static temperature ratio. Increasing the inlet Mach number ratio enhances the distribution area and the mixing effect of the subsonic–supersonic shear layer, resulting in an amplified asymmetry of the dual-vortex and a larger vortex size near the supersonic mainstream. The static temperature ratio influences the ejection effect of the supersonic flow on the subsonic flow, causing the recirculation zone to shift toward the supersonic mainstream and facilitating interaction with the subsonic–supersonic mixing layer. The findings contribute to a better understanding of the interaction between the subsonic–supersonic shear layer and the recirculation zone, offering guidance for improved subsonic–supersonic mixing flow designs in combustion systems.

The swirling-flow single trapped vortex combustor (SSTVC) provides a new design approach for aviation engine combustors, showing promising potential for achieving high-temperature rises or low emissions. However, current research on it is limited to demonstrating its feasibility and lacks further investigation into its combustion organization, combustion, and emission characteristics. In this study, experiments are conducted to investigate the flow field, combustion, and emission characteristics of the SSTVC under 5 inlet reference velocities and 3 inlet temperatures. As the inlet reference velocities increase from 9.96 m/s to 14.95 m/s, noticeable changes are observed in the backflow velocity of the cavity, increasing from 33.8 m/s to 48.3 m/s. By combining the flow field structure and flame features, the combustion organization of the SSTVC is summarized. With an increase in the inlet reference velocities, the combustion efficiency remains almost unchanged initially at 95.75% and then increases to 96.55%. The outlet temperature distribution improves, with the pattern factor value decreasing from 0.23 to 0.14. The carbon monoxide emissions initially increased from 94.91 to 99.89 and then decreased to 88.80, the total unburned hydrocarbons decreased from 23.22 to 16.03, and the nitrogen oxide emissions showed no significant change. With an increase in the inlet temperature, all parameters improved, with the combustion efficiency increasing from 93.71% to 96.61%, the pattern factor decreasing from 0.26 to 0.12, and all emission parameters decreasing as well.