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This study focuses on the great challenges for combined cooling and power supply on hypersonic aircrafts. To address the issues of low thermal efficiency and high fuel consumption of heat sink by the existing CO2 supercritical Brayton cycle, a novel fuel-based CO2 transcritical cooling and power (FCTCP) system is constructed. A steady-state simulation model is built to investigate the impacts of combustion chamber wall temperatures and fuel mass flow rates on the FCTCP system. Thermal efficiency of the CO2 transcritical cycle reaches 25.2~32.8% under various combustion chamber wall outlet temperatures and endothermic pressures. Compared with the supercritical Brayton cycle, the thermal efficiency of novel system increases by 54.5~80.9%. It is found from deep insights into the thermodynamic results that the average heat transfer temperature difference between CO2 and fuel is effectively reduced from 153.4 K to 16 K by split cooling of the fuel in the FCTCP system, which greatly enhances the matching of CO2–fuel heat exchange temperatures and reduces the heat exchange loss of the system. Thermodynamic results also show that, in comparison to the supercritical Brayton cycle, the cooling capacity and power generation per unit mass flow rate of working fluid in the FCTCP system increased by 75.4~80.8% and 12.9~51.6%, respectively. The FCTCP system exhibits a substantial performance improvement, significantly enhancing the key characteristic index of the combined cooling and power supply system. This study presents a novel approach to solving the challenges of cooling and power supply in hypersonic aircrafts under limited fuel heat sink conditions, laying the groundwork for further exploration of thermal management technologies of hypersonic aircrafts. [ABSTRACT FROM AUTHOR]

Abstract The role of efficient fuel mixing and a stable flame holder is crucial in enhancing the performance and capabilities of scramjet engines for high-speed flight. The present research paper has tried to disclose the fuel mixing efficiency of 3-lobe annular nozzle on the mixing mechanism of the fuel jet behind the strut. In addition, using internal air jet flow for increasing the circulation strength and fuel mixing behind the strut is also examined in this study. Numerical simulation of the flow and fuel jet behind the strut is done to reveal the main physics related to the mechanism of fuel mixing inside the combustor with the proposed injection system. The results of our simulation show that using annular 3-lobe fuel jet improve the fuel mixing via production of the multiple vortex pairs within the combustor behind the strut. The use of internal air jet also enhances the fuel mixing efficiency up to 90% in combustor of scramjet engine.

This study focuses on numerically exploring the effect of coupling factors on enhanced combustion performance in a two-strut scramjet combustor. The impact of five geometric variables, namely d (the vertical distance between the two struts), ur (lr) (the radius of the upper(lower) strut' tip), and ul (ll) (the length of the upper(lower) strut), is investigated through univariate analysis. While increasing ur and lr intensifies shocks at the cost of great total pressure loss, decreasing ul (ll) strengthens the upper(lower)flame but compresses the other. Besides, changing d influences the interaction of shocks and shear layers. To investigate the coupling effect of these variables, the combustion efficiency of the two-strut combustor is optimized through the SEGA algorithm based on a Kriging model built by small sampling points in 5D design space, which increases combustion efficiency by 14% and thrust by 27%. The optimum design features small ur and zero lr , larger d and ll and smaller ul. Further analysis performed on it revealed that the mechanism for enhancing combustion performance involves coupling the five geometry variables to assist combustion of the fuel closer to the upper expanding wall via stronger shock waves. Simultaneously, it weakens the shocks induced by the other strut to balance the total pressure loss. • Study coupling factors for efficient combustion in a two-strut scramjet combustor. • Enhance combustion in an asymmetry two-strut configuration via optimization. • Alternate shocks to enhance ignition and decrease total pressure loss. [ABSTRACT FROM AUTHOR]

This study serves as a research endeavor aiming to explore the behavior of the coupling flow effects of the single expansion ramp nozzle (SERN) in over-expansion conditions during the static start-up process. The open-source program OpenFOAM and its solver "rhoCentralFoam" are employed in the 2D simulation and the two critical geometric variations, the shape of the ramp and the length of the flap beyond the throat, are considered in the geometric variation. The result shows the preferable propulsion performance in the FSS (Free Shockwave Separation) state compared to RSS (Restricted Shockwave Separation). FSS also plays the role of the initial, albeit transient, separation, which originates from the shockwave from the throat and will eventually transform into a stabler RSS state. For the 100% flap length configuration in this study, the axial thrust can achieve a high value of 500 N/m in the FSS state and decrease to around 450 N/m, on average, in the RSS state. The trust angle also shows a preferable performance of around −13° in FSS compared to −30° in RSS. Regarding geometric modifications, both modifications, shorting the flap and bell-shaped ramp adjustments, manifest similar effects. Both conical and bell-shaped short flap configurations demonstrate an axial thrust from around 1750 to 1900 N/m and a thrust angle of around −45°. However, the flap shortening, which may demonstrate an attitude compensation effect, exhibits a more pronounced effect compared to the bell-shaped modification. [ABSTRACT FROM AUTHOR]

Reliable ignition is a fundamental prerequisite for the scramjet engines. In order to improve the reliability of ignition, further enhance the application of spark ignition in engineering, and improve the ignition ability in the field of supersonic combustion, an innovative long electrode distance high-energy spark igniter (LHSI) system was proposed by our research team. To reveal the instantaneous ignition characteristics of the LHSI, a simulation model of the LHSI is established and verified by experiments. Based on this, the formation and development of the initial flame kernel of vaporized kerosene and the ignition characteristics of LHSI in a scramjet combustor were simulated and compared with the conventional high-energy spark igniter (HSI). The analysis revealed that the success of the cavity ignition is greatly influenced by the local equivalent ratio of the initial flow field and the penetration depth of the initial flame kernel. Meanwhile, the fore recirculation zone of the cavity plays a key role in stabilization and development of the initial flame kernel. Compared with the HSI, the LHSI can produce initial flame kernel with higher penetration and greater size, which is the main reason why LHSI has a broader ignition boundary. The simulation results show that the oil-rich and oil-lean ignition boundaries of LHSI are about 17 % and 55 % wider than those of HSI, respectively.

Scramjet engines represent one of the most promising alternatives to power hypersonic vehicles, although the peculiarities of supersonic combustion raise many challenges. Computational Fluid Dynamics (CFD) simulations can be used to model highly turbulent reactive flows in such complex configurations and avoid expensive experimental campaigns. A particularly challenging aspect when dealing with numerical simulations of turbulent combustion is accurately modelling the interactions between chemistry and turbulence. These interactions can play a significant role in supersonic flows and require an appropriate closure since the conditions encountered in supersonic flows differ from those of conventional combustion devices. For this reason, an assessment of the role of turbulence-chemistry interactions in supersonic combustion is needed. This work evaluates the performance of the Partially-Stirred Reactor (PaSR) closure of the mean chemical source term by performing Reynolds-Averaged Navier–Stokes (RANS) simulations of two selected test cases, namely, the Burrows and Kurkov combustor and the German Aerospace Centre (DLR) scramjet. The PaSR approach requires the definitions of two important quantities: mixing and chemical time scales. While the chemical time scale is related to finite-rate chemistry, supersonic flow conditions can affect the most appropriate definition of the mixing time scale. For this reason, several formulations of the mixing time scale are tested in this work. The PaSR model is validated against experimental data and also compared to the finite-rate approach without turbulence-chemistry interactions (FR/No-TCI) and the Eddy Dissipation Concept (EDC). In the Burrows and Kurkov test case, the PaSR model can accurately predict the main quantities of interest (e.g., temperature and chemical compositions), showing better results than the FR/No-TCI and similar to the EDC model. The validation on the DLR case confirms these trends and highlights the importance of accounting for turbulence-chemistry interaction in supersonic combustion modelling. However, the EDC model is not able to predict a stable ignited flame in the DLR combustor. The presence of recirculation zones near the strut injector of the DLR scramjet can explain the better performance of the PaSR model versions employing the integral-based and geometric mean formulations, which estimate comparable values of the mixing time but higher than those of other time scale approaches considered in this work. • Supersonic combustion is characterised by similar chemical and mixing time scales. • The Partially Stirred Reactor model is suitable for RANS simulations of supersonic combustion. • The model performs better than a laminar finite-rate approach and Eddy Dissipation Concept. • The choice of the mixing time scale affects the Partially Stirred Reactor accuracy. • The most appropriate mixing time scale is selected for two scramjet combustion test cases. [ABSTRACT FROM AUTHOR]

The active regenerative cooling technology faces the bottleneck problem of insufficient heat transfer capacity when the scramjet flies at a higher Mach number. It is proposed to strengthen the heat transfer performance of the regenerative cooling channel with airfoil-fins. To verify the enhanced heat transfer effect of the airfoil-fin channel in principle, an experimental test platform for flow and heat transfer of ambient air in NACA0021 symmetrical airfoil-fin channels and NACA4822 asymmetric airfoil-fin channels (with cross-section sizes of 50 mm × 50 mm) was built. The Nusselt number of the heated surface was obtained based on the steady-state liquid crystal technique. The results show that, the heat transfer intensities of NACA0021 symmetrical airfoil-fin channels and NACA4822 asymmetric airfoil-fin channels improve by 0.17%~17.1% and 18.4%~52.1%, respectively. Correspondingly, PECs are 1.04 and 1.24, respectively, with the volume flow of ambient air at 50 m³/h. The NACA4822 asymmetric airfoil-fin channel can enhance the heat transfer performance of the middle heating surface under the condition of a large flow rate. The flow pressure drop in the airfoil-fin channels also increases correspondingly, where the pressure drop in the NACA4822 airfoil-fin channel is the largest. The asymmetry of the airfoil-fin causes the continuous accumulation of flow turbulence intensity, resulting in a significant increase in the downstream pressure drop. The work is helpful for further research on the flow and heat transfer characteristics of supercritical fluids in airfoil-fin channels, and broadens the application temperature range of the active regenerative cooling technology for scramjets. [ABSTRACT FROM AUTHOR]

The forced ignition instability and combustion instability in a scramjet engine were investigated experimentally and numerically using modal analysis. In this study, proper orthogonal decomposition (POD) was applied to OH+ chemiluminescence image data obtained during previous supersonic combustion experiments. In this scramjet combustor, a microrocket torch was installed in the cavity. First, to clarify the combustion instability phenomenon in the cavity and flow paths, a POD analysis was performed using the OH+ chemiluminescence image data. The maximum power spectrum density peak shifted to the high-frequency region in high-order modes. The combustion oscillations between 100 and 500 Hz were attributed to injector flame feedback from the fuel or torch jet gas because the fuel and torch gas were directly injected into the shear layer, where fuel combustion occurred. Next, to clarify the forced ignition instability phenomenon in the cavity, the POD modes near the torch jet orifice were investigated. The low frequencies near the torch jet orifice within the cavity were likely associated with the torch gas injector flame feedback because the frequency spectra of the inner pressure oscillations in the micro-rocket torch matched the frequency spectra peak near the torch jet orifice in the cavity. [ABSTRACT FROM AUTHOR]

This study focuses on the great challenges for combined cooling and power supply on hypersonic aircrafts. To address the issues of low thermal efficiency and high fuel consumption of heat sink by the existing CO2 supercritical Brayton cycle, a novel fuel-based CO2 transcritical cooling and power (FCTCP) system is constructed. A steady-state simulation model is built to investigate the impacts of combustion chamber wall temperatures and fuel mass flow rates on the FCTCP system. Thermal efficiency of the CO2 transcritical cycle reaches 25.2~32.8% under various combustion chamber wall outlet temperatures and endothermic pressures. Compared with the supercritical Brayton cycle, the thermal efficiency of novel system increases by 54.5~80.9%. It is found from deep insights into the thermodynamic results that the average heat transfer temperature difference between CO2 and fuel is effectively reduced from 153.4 K to 16 K by split cooling of the fuel in the FCTCP system, which greatly enhances the matching of CO2–fuel heat exchange temperatures and reduces the heat exchange loss of the system. Thermodynamic results also show that, in comparison to the supercritical Brayton cycle, the cooling capacity and power generation per unit mass flow rate of working fluid in the FCTCP system increased by 75.4~80.8% and 12.9~51.6%, respectively. The FCTCP system exhibits a substantial performance improvement, significantly enhancing the key characteristic index of the combined cooling and power supply system. This study presents a novel approach to solving the challenges of cooling and power supply in hypersonic aircrafts under limited fuel heat sink conditions, laying the groundwork for further exploration of thermal management technologies of hypersonic aircrafts.

In present paper, computational fluid dynamic is applied to evaluate hydrogen mixing impacts of the single extruded lobe injector at supersonic cross flow. The aim of this investigation is to introduce efficient injection system by elongation of the extruded nozzle at high-velocity air flow. The shock waves and temperature contour for the two nozzle heights are compared. The importance of vortex produced inside the domain on the fuel diffusion is also explained. The usage of annular and coaxial air and fuel jet for the improvement of the proposed configurations are also investigated. Obtained results shows that elongation of extruded nozzle form 2 mm to 4 mm improves the fuel mixing about 50%. However, the injection of the inner air jet results in 85% higher fuel mixing in combustion chamber of the scramjet engine. • Influence of shock generator on fuel mixing of single ejected nozzle at supersonic flow is studied. • The impact of induced shock waves on fuel jet interactions of the single jets are studied. • CFD method is used for the modeling of the cross jet injected from side of ejected 2-lobe nozzles. • The development and power of vortex have significant role on the fuel mixing downstream the jet. [ABSTRACT FROM AUTHOR]

This paper investigates an intelligent strut design scheme for achieving high-performance combustion of a combustion chamber under wide-ranging flow conditions. The design process involves optimizing the base strut configuration for the design point and obtaining the Pareto front. The subsequent selection of the optimal configuration is followed by analyzing and interpreting the optimization results using single-factor analysis rules. The findings reveal that the selected optimal configuration enhances the combustion efficiency by 8.22% and reduces the force by 66.12% compared to the traditional DLR combustion chamber. Additionally, a surrogate model is constructed to investigate the influence of the rotation angle of the intelligent strut's rotating blades on the combustion efficiency and force at a Mach number range of 2–3.5. Results indicate that the intelligent strut is capable of improving combustion efficiency via regulation, and the efficiency can quickly reach a desirable level. These findings suggest that the proposed design scheme has potential applications in achieving high-performance combustion in combustion chambers exposed to wide-ranging flow conditions. • This paper investigates an intelligent strut design scheme for achieving high performance combustion. • The design process involves optimizing the base strut configuration for the design point and obtaining the Pareto front. • This design scheme aims to achieve the best comprehensive combustion. Performance based on the design point. • Results indicate that the intelligent strut is capable of improving combustion efficiency. via regulation. [ABSTRACT FROM AUTHOR]

The elliptical hole has the potential to promote jet breakup and improve atomization and mixing quality. The flow behavior in the nozzle and the outflow characteristics have an important influence on the atomization and mixing process of the downstream fuel. Therefore, the inner cavitation flow processes and outflow characteristics for elliptical holes for scramjet engine at injection pressures of 3 MPa and 5 MPa were investigated using the Volume of Fluid (VOF) method and Large Eddy Simulation (LES). The numerical findings demonstrated that the cavitation region and vortex structures of elliptical holes mainly appeared on the long axial plane. The cavitation intensity of elliptical holes was weaker than that of the circular hole. Compared with the circular hole, the gas phase volume fraction of elliptical hole with the aspect ratio of 4 was reduced by about 22.1 and 21.1% respectively under 3MPa and 5MPa injection pressure. The application of elliptical holes inhibited the generation of cavitation and had large mass flow rate and flow coefficient. The mass flow rate of the elliptical hole with the aspect ratio of 4 was the largest, and the mass flow rate increased by about 5.1% compared with that of the circular hole. Elliptical holes showed a smaller velocity coefficient and a larger area coefficient when cavitation existed. The elliptical hole with the aspect ratio of 4 held the smallest velocity coefficient and largest area coefficient, with the difference of 12.1% and 18.9% respectively from the circular hole. Another important point was that elliptical holes enhanced the turbulent disturbance at the hole outlet, which facilitated the breakup of downstream jet. [ABSTRACT FROM AUTHOR]

To acquire the ignition promotion methods for scramjet engine, the auto-ignition and reaction kinetic characteristics of hydrogen/n-dodecane mixture under mid and high temperatures were revealed by detailed chemical kinetic mechanism and experimental data. The results illuminated that the increases of n-dodecane's content and ϕ significantly declined the ignition delay time (IDT) of hydrogen mixture (by two orders of magnitude at most) at mid-temperature high-pressure (T = 950 K& p = 40 bar). The addition of n-dodecane induced the OH radical formation in initial stage and the advance of fuel's H-abstraction reactions (hydrogen: R3, n-dodecane: R2741/R2742/R2743/R2746). Whereas, the corresponding reduction in hydrogen content inhibited the establishment of active radical pool in dual-fuel reaction system before ignition time point. The ignition promoting effect of advanced initial reaction stage overweighted the ignition inhibiting effect of weakened rapid reaction stage in mid-temperature low-reactivity dual-fuel reaction system. Interestingly, at high-temperature high-pressure (T = 1300 K& p = 40 bar), the IDT firstly decreased (by 18.8%), and then continuously increased (by 105.3% at most) with the elevating n-dodecane's content. The minor addition of n-dodecane (1% of mole fraction) at high-temperature had a similar ignition promoting effect as that at mid-temperature. Whereas, the further reduction in hydrogen content dramatically suppressed the formation of active radical pool (H/OH/HO 2 /H 2 O 2), which overweighted the ignition promoting effect of slightly advanced initial reaction stage. In other words, the hydrogen played a more critical role than the n-dodecane in high-temperature reaction system due to the originally high reactivity of hydrogen. The auto-ignition properties of hydrogen/n-dodecane mixture relied on the competition relationship of initial and rapid reaction stages under specific thermodynamic conditions and blending ratios. This investigation not only reveals internal interaction mechanism between fuels and active radicals in hydrogen/n-dodecane reaction system, but also provides fundamental insights for the precisely ignition promition methods of hydrogen/aviation-kerosene dual-fuel scramjet engine. • Auto-ignition and reaction kinetic properties of H 2 /n-C 12 H 26 mixture were revealed. • Blending n-C 12 H 26 dramatically reduced the IDT of H 2 mixture at mid-T. • IDT firstly reduced and then increased with increasing ratio of n-C 12 H 26 at high-T. • Ignition of H 2 /n-C 12 H 26 relied on competition of initial and rapid reaction stages. [ABSTRACT FROM AUTHOR]

This study serves as a research endeavor aiming to explore the behavior of the coupling flow effects of the single expansion ramp nozzle (SERN) in over-expansion conditions during the static start-up process. The open-source program OpenFOAM and its solver “rhoCentralFoam” are employed in the 2D simulation and the two critical geometric variations, the shape of the ramp and the length of the flap beyond the throat, are considered in the geometric variation. The result shows the preferable propulsion performance in the FSS (Free Shockwave Separation) state compared to RSS (Restricted Shockwave Separation). FSS also plays the role of the initial, albeit transient, separation, which originates from the shockwave from the throat and will eventually transform into a stabler RSS state. For the 100% flap length configuration in this study, the axial thrust can achieve a high value of 500 N/m in the FSS state and decrease to around 450 N/m, on average, in the RSS state. The trust angle also shows a preferable performance of around −13° in FSS compared to −30° in RSS. Regarding geometric modifications, both modifications, shorting the flap and bell-shaped ramp adjustments, manifest similar effects. Both conical and bell-shaped short flap configurations demonstrate an axial thrust from around 1750 to 1900 N/m and a thrust angle of around −45°. However, the flap shortening, which may demonstrate an attitude compensation effect, exhibits a more pronounced effect compared to the bell-shaped modification.

CHEMKIN-Pro was used to build a custom-made reduced reaction mechanism for a 3D simulation of C2H4/Air reaction in a supersonic combustion ramjet (scramjet) combustor for the JAXA S-520-RD1 flight test performed on 24 July 2022. It was also attempted to reproduce the effect of unavoidable H2O vitiation in the facility (ground) test on combustion characteristics. USC Mech II (a 111-species mechanism) was applied as the master mechanism. First, 1D simulation was conducted by applying a plug flow reactor model; 34-, 23-, 20-, 18-, and 19-species mechanisms were suggested as candidates for reduced mechanism. Then, 0D ignition delay simulation by applying a perfectly stirred reactor model by CHEMKIN-Pro, 2D reacting flow simulation by CRUNCH CFD, and 3D reacting flow simulation by RANS were performed to select the most suitable reduced mechanism. The 20-species (96-elementary-reactions) mechanism consisted of H2, O2, N2, H2O, CO2, CO, H, O, OH, HO2, CH2, CH2*, CH3, HCO, CH2O, CH3O, C2H2, C2H3, C2H4, and CH2CHO and was selected as the best-reduced reaction mechanism for C2H4-fueled 3D simulation under the flight and facility conditions corresponding to the RD1 flight test. [ABSTRACT FROM AUTHOR]

This study numerically investigated the combustion instability and characteristics of a laboratory-scale gaseous hydrogen-fueled scramjet combustor. For this purpose, a numerical simulation with an improved detached eddy simulation and a detailed hydrogen/oxygen reaction mechanism was performed. The numerical framework used high-resolution schemes with high-order accuracy to ensure high resolution and fidelity. A total of five fuel injection pressures were considered to characterize the combustion instability as a function of the equivalence ratio. A sampling time of up to 100 ms was considered to sufficiently accumulate several cycles of low-frequency combustion instability dynamics with a period in the order of 100 Hz. Numerical results revealed the repetitive formation/dissipation dynamics of the upstream-traveling shock wave, and it acts as a key factor of combustion instability. The period and derived principal frequency of these upstream-traveling shock waves is several ms. The frequency analysis showed that the instability frequency increased in the low-frequency range as the combustion mode transitioned from the cavity shear-layer to the jet-wake type. This characteristic was derived from the transition in combustion mode at the same equivalence ratio. Therefore, it suggests that the instability frequency shifting is governed by the combustion mode rather than the equivalence ratio. These comprehensive numerical results demonstrated not only the effect of the equivalence ratio but also the important role of the combustion mode on the low-frequency combustion instability. [ABSTRACT FROM AUTHOR]

Air heater is a ground test engine for generating and supplying hot vitiated air to scramjet engine combustor at designed parameters. It has different mixing zone modules consisting of an outer flange and an inner shell. Cooling channels are there in between the flange and inner shell. At the aft end and fore end of the mixing zone modules inner cylinder and the flange are joined by EB weld. After the ground test, the air heater was kept in the test bay for a long time. Subsequently during the cooling water flow trial through air heater, water leakage was noticed at the EB weld joint between the inner cylinder and flange. Scale like formation over the flange interface near the EB weld was also noticed. In situ metallographic analysis revealed that the cracks have initiated from the weld region and propagated in transgranular manner into the flange side as well as inner shell parent metal. The nature of cracks responsible for leakage and the chemical composition of the brown colored patches on the flange surface near the EB weld as well as at the middle region of inner shell were examined. Multiple cracks with extensive branching are observed. Cracks have been initiated from the corrosion products formed at the weld region. From the nature of the crack and presence of corrosion, cracking is attributed to Stress Corrosion Cracking. [ABSTRACT FROM AUTHOR]

The scramjet model used in the control system needs to consider both accuracy and real-time requirements. In this paper, a quasi-one dimensional component level model method for scramjet is proposed. Under the premise of meeting the real-time requirement, this modeling method has higher calculation accuracy, and can obtain higher thrust performance without exceeding the limits. In order to improve the calculation accuracy of the scramjet model, the boundary layer correction and the interaction between shock wave and boundary layer are considered in the inlet and isolator modeling process by using reference temperature method. For the scramjet compression components, the inlet and isolator models are simplified based on shock train theory to improve real-time performance. Meanwhile, the combustor model and nozzle model are established and the combustor model is established with the combination of equal area and isothermal based on temperature limitation. Considering the boundary layer correction, the accuracy of the inlet and isolator model is improved by 8.79 and 7.9% respectively. Compared with the equal area combustion mode, the combined combustion mode can get higher thrust. In the C + + simulation environment, the average calculation time of each state point is about 25 ms which meets the real-time requirements. [ABSTRACT FROM AUTHOR]

Abstract Both the fuel transfer control scheme and the knowledge of fuel temperature and pressure values before injection are important to a scramjet engine. A measurement method of the fluid temperature suitable for small channel is presented, and a fluid transfer control scheme of the area of injecting holes is designed using the balance equations of fluid mass, energy and entropy. A convective heat transfer experiment is conducted to evaluate the temperature measurement and the fluid transfer control design method of the heat absorbing supercritical hydrocarbon fuel. The results show that during the heating process the injecting pressure meets the requirement, and the fluid transfer control system works well. Such results suggest that the design method in the present work would be useful in the design of regenerative cooling for scramjet engines.

In this study, experimental methods are used to investigate the effects of different vibration and pressure parameters on heat transfer performance are analyzed. The results show that at a subcritical pressure (0.1 MPa), the external vibration begins to affect the heat transfer when the fuel passes the phase-change point and becomes gaseous. At a near-critical pressure (3 MPa), the external vibration deteriorates the heat transfer of fuel across the critical-temperature zone. At the supercritical pressure (5 MPa), the external vibration intensifies the heat transfer in the hot end of the channel only when the fuel is below the critical temperature and the internal wall is above the critical pressure. Combined with data analysis, it can be seen that the external vibration mainly acts on the temperature boundary layer of the fuel oil to affect the wall temperature and heat transfer coefficient. [ABSTRACT FROM AUTHOR]

The efficient injection of fuel in the combustion chamber is significant for the advancement of the current supersonic vehicles. This study has focused on the role of non-alignment injector configuration on the fuel mixing of the hydrogen jet inside the combustion chamber of the scramjet engine. The fuel distribution and vortex generation behind both annular and coaxial jets are modeled and compared to disclose the mechanism of fuel diffusion in the combustion chamber. A three-dimensional model of two nozzle angles is produced to visualize the complicated flow interaction of four non-aligned circular jets at supersonic cross-flow. Our results indicate that the strength of the vortex pair produced by the core of the fuel jet is extended more when the angle of the nozzles is increased. The flow visualization of the jet also confirms that the flow becomes more complex in non-alignment jet configurations and fuel mixing is improved by the addition of the internal air jet in the suggested injection model. • The role of non-aligned injector system for efficient fuel mixing of hydrogen jets is investigated. • CFD is used to investigate the flow and fuel mixing inside the combustor of scramjet engine. • Role of the vortex pairs on the fuel mixing and penetration is presented. [ABSTRACT FROM AUTHOR]

Traditional deep learning technologies often overlook crucial spatiotemporal information when reconstructing supersonic combustion flow fields. These models can only reconstruct the current flow field, resulting in poor visual quality due to the atomization effect. This research introduces experimental studies on scramjet combustion flows with variable injection pressures at the inlet of a scramjet Mach 2.5, creating image datasets of flame evolution processes over different predicted time spans (5, 10, and 15 ms). It develops a multi-scale attention algorithm with deblurring (DB-MSAN) capabilities to predict the flame image after a certain period using the wall pressure signal from the previous moment. The algorithm employs two multi-head attention channels at different scales to extract timing information, encouraging the model to focus on the anisotropy of the input data and to identify more significant features. The DB module utilizes the adapted atmospheric scattering model to jointly estimate the transmission map and atmospheric light, achieving an exceptional defogging effect with a minimal computational expense. The model's overall performance is assessed regarding efficiency and accuracy, comparing DB-MSAN's performance differences with those of Chen's and the ResNet16 models across multi-span, large-span, various test sets, and sparse wall measurement points. Experimental results indicated that, compared to other methods, DB-MSAN achieves a maximum peak signal-to-noise index improvement of 76.10 % and an SSIM index improvement of 155.26 % when predicting average-quality images. The prediction accuracy remains stable, even with sparse pressure input. In addition, DB-MSAN strikes an optimal balance between the number of parameters and model accuracy; it represents only 16.97 % of the volume compared to Chen's model, which is 585 MB. [ABSTRACT FROM AUTHOR]

This study numerically investigated the combustion instability and characteristics of a laboratory-scale gaseous hydrogen-fueled scramjet combustor. For this purpose, a numerical simulation with an improved detached eddy simulation and a detailed hydrogen/oxygen reaction mechanism was performed. The numerical framework used high-resolution schemes with high-order accuracy to ensure high resolution and fidelity. A total of five fuel injection pressures were considered to characterize the combustion instability as a function of the equivalence ratio. A sampling time of up to 100 ms was considered to sufficiently accumulate several cycles of low-frequency combustion instability dynamics with a period in the order of 100 Hz. Numerical results revealed the repetitive formation/dissipation dynamics of the upstream-traveling shock wave, and it acts as a key factor of combustion instability. The period and derived principal frequency of these upstream-traveling shock waves is several ms. The frequency analysis showed that the instability frequency increased in the low-frequency range as the combustion mode transitioned from the cavity shear-layer to the jet-wake type. This characteristic was derived from the transition in combustion mode at the same equivalence ratio. Therefore, it suggests that the instability frequency shifting is governed by the combustion mode rather than the equivalence ratio. These comprehensive numerical results demonstrated not only the effect of the equivalence ratio but also the important role of the combustion mode on the low-frequency combustion instability.

Efficient mixing and combustion in supersonic flow is critical for developing hypersonic airbreathing propulsion systems such as scramjet engines. The short residence time of a few milliseconds of inflow air in a scramjet combustor make it very difficult to achieve efficient and fast fuel/air mixing. In supersonic flow, the compressibility effect suppresses the development of planar free shear layers, which play an important role in low-speed turbulent mixing. For this reason, a mixing strategy based only on planar shear layers cannot be effective, as it requires the use of impractically long engines that result in weight and heat transfer penalties. One of the efficient methods to enhance and accelerate mixing in supersonic flow is to induce streamwise vortices. Because streamwise vortices swirl perpendicular to the airflow, they are less susceptible to compressibility effects and are advantageous for enhancing the mixing process between the injectant and the airflow. Experimental and computational studies on mixing enhancement by streamwise vortices have been extensively conducted over several decades. Therefore, this paper reviews and summarizes research progress on mixing enhancement using streamwise vortices for four widely studied physical devices: lobed mixers, ramp injectors, pylon injectors, and strut injectors. In addition, the mechanisms for generating streamwise vortices, the flow structures of these physical devices, and the effect of streamwise vortices on mixing enhancement are provided. • The mixing enhancement strategies using streamwise vortices were reviewed and summarized. • The mechanisms generating streamwise vortices by physical devices (lobed mixers, ramps, pylons, and struts) were provided. • The geometrical effects of the physical devices on the fuel/air mixing in supersonic flow were presented. • Remarks were provided for future work on the physical devices for the application in scramjet combustors. [ABSTRACT FROM AUTHOR]

Both the fuel transfer control scheme and the knowledge of fuel temperature and pressure values before injection are important to a scramjet engine. A measurement method of the fluid temperature suitable for small channel is presented, and a fluid transfer control scheme of the area of injecting holes is designed using the balance equations of fluid mass, energy and entropy. A convective heat transfer experiment is conducted to evaluate the temperature measurement and the fluid transfer control design method of the heat absorbing supercritical hydrocarbon fuel. The results show that during the heating process the injecting pressure meets the requirement, and the fluid transfer control system works well. Such results suggest that the design method in the present work would be useful in the design of regenerative cooling for scramjet engines. [ABSTRACT FROM AUTHOR]

The effects of back pressure and cavity L/D ratio on the shock wave structure in the cold flow field of a typical cavity-based scramjet combustor with combined inlet and isolator is investigated numerically in the selected scramjet models. The scramjet with a throat ratio of TR 0.0 and cavity L/D 6.04 was analyzed. To perform such analysis, steady, 2-D RANS was used with SST k-ω. From the analysis, the value of static pressure along the cowl surface, contours of Mach number and pressure were obtained. The scramjet was modeled with different TR 0.1, 0.2, 0.25 and 0.3 with the same cavity L/D 6.04 and different cavity L/D 4.04, 9.04 and 12.04 with the same TR 0.25. All the models were analyzed with the same inlet conditions and the results were obtained. From the analysis, it was observed that the increase in back pressure moves the shock train towards the inlet of the isolator which leads to 'engine unstart' after the throat ratio of TR 0.1. Also, it is observed that there is an optimal L/D ratio of the cavity L/D 9.04 which restricts the propagation of high-pressure waves obtained in the combustor. [ABSTRACT FROM AUTHOR]

It is vital to analyze the flame characteristics and identify the flame states efficiently in scramjet. Flame combustion and oscillation behaviors in a hydrogen-fueled cavity-stabilized scramjet combustor have been investigated in this study. A 500 Hz high-speed PLIF (planar laser-induced fluorescence) technique with a 20-cm-wide view field is introduced to characterize the combustion flow. The spatial distribution of the flame structure under different fuel injection conditions is studied. The results indicate that the position of the flame peak shifted downstream of the cavity when increasing the injection pressure to a high level. The dynamic characteristics of flames under different flow conditions are analyzed, and the correlation between flame features and hydrogen injection conditions is obtained. These flame features include the vertical range, the flow direction position of the peak flame, the flow direction position of the centroid, the vertical position of the centroid, the flame area, and the flame circumference. For the present cases, the flame of 4.0 Mpa is more unstable than that of 1.5 Mpa under any flow conditions. Moreover, an experiment based on feature extraction results shows that the KNN (K-nearest neighbor) classifier could achieve high accuracy for flame state recognition in this scramjet combustor. • High-speed OH-PLIF technique and feature extraction methods were performed. • The combustion characteristics of flame in a scramjet combustor were obtained. • Flame structure features were used to investigate flame oscillation characteristics. • The flame state recognition by the KNN classifier was carried out. [ABSTRACT FROM AUTHOR]

Supercritical CO 2 closed-Brayton-cycle demonstrates considerable application potential on hypersonic vehicles, effectively addressing power supply issues during prolonged flights. However, the CO 2 flow rate is strictly limited by the cycle's cooling source (aviation kerosene), rendering traditional regenerative cooling inadequate for sufficiently cooling the scramjet engine. To address this challenge, this study proposes an integrated cooling scheme that combines regenerative cooling, ceramic thermal-protective layer, and fuel film cooling. A two-dimensional coupled numerical model was developed, simultaneously considering the convective heat transfer of supercritical CO 2 and the shear mixing of supersonic fuel film. The results indicate that the high-enthalpy gas imposes an extreme aerodynamic heat load on the regenerative cooling channel wall, resulting in a notable deterioration in heat transfer for supercritical CO 2. Implementing kerosene film cooling within the combustor can effectively mitigate the abnormal heat transfer behavior of supercritical CO 2 and significantly reduce viscous dissipation heat within the gas boundary layer. By optimizing the fuel film injection layout, this combined cooling scheme reduced wall heat flux by 43 % and temperature by over 400 K compared to traditional regenerative cooling, thus providing effective thermal protection for the scramjet engine. • Coupled numerical simulation between S-CO 2 cooling channel and scramjet is achieved. • Heat transfer deterioration occurred in the S-CO 2 cooling channel under aero-heating. • Even with a ceramic insulation layer, S-CO 2 cooling channel can't cool the scramjet. • Supersonic fuel film mitigates the heat transfer deterioration in the S-CO 2 channel. • Double fuel film spraying combined with S-CO 2 channel is an effective cooling scheme. [ABSTRACT FROM AUTHOR]

The usage of zigzag multi-jet configuration for efficient fuel circulation along the supersonic combustion chamber has been extensively examined in this article. To analyze the flow and fuel jet interactions, computational fluid dynamic is applied as an efficient technique to visualize the flow and fuel jet in the zigzag injection system. Injection via an annular nozzle is compared with coaxial air and fuel jet to find the effective mechanism for well-organized fuel mixing in the combustion chamber of the scramjet engine. The circulation and fuel mixing of the zigzag injection is inspected at a supersonic air stream with Mach = 4. The contour of the Mach value and stream shows that the deflection of the air stream after contact with the core of the upstream jet redirects to the core of the downstream jet and consequently, the fuel mixing is enhanced in the combustion chamber. The usage of an internal air jet expands the performance of fuel mixing of annular jets up to 100 % in the zigzag nozzle arrangement. • The role of zigzag injector for fuel mixing of hydrogen jet in combustion chamber is investigated. • Computational method is applied to model fuel jet in supersonic cross flow. • Fuel mixing and shock wave interactions in zigzag arrangements are comprehensively studied. [ABSTRACT FROM AUTHOR]

In scramjet engines, ignition must take place within a residence time on the order of milliseconds. In this study, secure ignition conditions for specified n-octane pyrolysis fuel components used in autoignition or forced-ignition by plasma jet torch in a high-speed flow were numerically investigated. First, the ignition delay time within the combustor and cavity flame-holder was estimated using chemical reaction analysis. Three fuel components (n-C8H18, all pyrolysis fuel (15 components, decomposition rate under 11%)), and pyrolysis gas fuel (eight gas fuel components, decomposition rate under 11%) could not self-ignite within the combustor and cavity residence time. Secondly, ignition using a plasma jet torch in the cavity was numerically investigated. In the case of forced-ignition by plasma jet torch, all pyrolysis fuel (No. 3) and n-C8H18 could ignite within the cavity residence time with less input energy than pyrolysis gas fuel (No. 3) under three kinds of Mach number flight conditions (M0 = 4, 6, and 8). Moreover, the effect of shortening the ignition delay time by raising the plasma jet torch gas temperature and O radical rate within the cavity was investigated. Ignition of the three kinds of mixture fuel was more greatly affected by the torch injection temperature than the O radical rate in the cavity under all Mach number flight conditions. [ABSTRACT FROM AUTHOR]

Scramjet is an air breathing engine in which the combustion takes in a supersonic air stream by generating high pressure. Wedge shaped strut plays a major role in the scramjet engine to mix and compress air at high speeds, igniting the air fuel mixture and sustaining the flame. But the geometric profile of wedge in which the air gets blocked partially or completely is unanswered. Therefore, the effect of blockage on intake characteristics is investigated in this present study. Numerical investigations are carried out using ANSYS Fluent at Mach 5 to study the changes in static pressure, stagnation pressure, temperature and mass flow rate for different percentage of blockage which is varied from 10% to 45%. Unstart-free air intake for scramjet engine is ensured by keeping the percentage blockage of fuel injection struts between 10% and 35%. The Mach number at supersonic regime until 35% of blockage turns suddenly into subsonic conditions afterwards for further increase in blockage percentage. Moreover, as the blockage is further increased beyond 35%, bow shock is observed ahead of the cowl and the flow through the intake is subsonic, thereby indicating 'unstart' intake. Based on this geometry in air intake, smooth gas dynamic compression or ramming effect is obtained up to 35% of blockage. [ABSTRACT FROM AUTHOR]

The detailed mechanism of supersonic mixing enhancement by a novel device is investigated computationally based on the Navier-Stokes equations. The device is a wall-mounted three-dimensional cavity, several parts of the rear face of which are cut out, and the lower wall of the cut-out part has a slope directed toward the injector from which the injectant is discharged. The flow structures and behaviors observed in the computational results are confirmed in the experimental results. Experimental information about the high-frequency (several tens of kHz) pressure oscillation in the cut-out part is acquired in this study using a pressure-sensitive paint (PSP) responding to rapid change in pressure. The computational results reveal that a hairpin-like vortex is formed inside the cavity. The vortex works on the shear layer across the cavity to deflect it upward. The deflected shear layer shields the injectant from the primary flow, which causes the injectant to penetrate deep into the primary flow. The computational results also clarify that the expansion and compression waves produced in the rectangular part of the cavity are delivered effectively to the injector along the slope of the cut-out part, which is confirmed in the PSP measurement results. The delivered waves actuate the jet issuing from the injector to fluctuate up and down periodically with a large amplitude. • The mechanism of supersonic mixing enhancement by a three-dimensional cavity is investigated. • The hairpin-like vortex formed inside the cavity makes the injectant to penetrate deep into the primary flow. • The pressure waves contributing to mixing enhancement are captured computationally and experimentally. • The pressure waves actuate the jet issuing from the injector to fluctuate up and down. [ABSTRACT FROM AUTHOR]

Liquid fuel-only injection into combustion air is widely used in gas turbine combustors for power generation, particularly in large capacity type. This combustion air goes by the name of ‘crossflow’. For the liquid fuel-only injection, breakup process, trajectory of liquid jet and droplet sizes have been extensively investigated. However, there remain much space to be devised for improvement in atomization. In the present study, twin-fluid injection is introduced instead of liquid-only injection into a gas turbine combustor, where atomization air is supplied through the side port and mixed with liquid in the mixing port at right angle, and two phases mixture is injected into the crossflow. A simple-structured twin-fluid atomizer is designed. An experimental study was conducted to compare the behaviours between liquid-only and twin-fluid injected transversely into a crossflow. High-speed photography technique is used to clarify the atomization characteristics of jet/spray. One feature of twin-fluid injection is that annular liquid film appears at the exit of the mixing port, and the film easily disintegrates by aerodynamical force of the crossflow. The jet/spray of the twin-fluid injection is not blown away easily by crossflow, and flow across the crossflow. Furthermore, the jet/spray of the twin-fluid spread and droplets disperse more widely in the crossflow. In addition, the twin-fluid injection produces numerous fine droplets and few coarse droplets.

In this study, we investigated the heat transfer characteristics of kerosene for various fuel flow rates in the regenerative cooling heat exchanger of scramjet engines. We used hydrogen for combustion and kerosene for the heat exchanger regenerative cooling. The internal combustor wall surface temperature was measured using a phosphor thermometer. The fuel pressure and temperature were also measured. Significant changes were observed when kerosene flow rate changed from 6 to 8 g/s. The fuel boiling point at specific pressures and temperatures was obtained using differential scanning calorimetry; at the kerosene boiling point, they were found to be consistent with the temperature and pressure values between 6 and 8 g/s fuel flows, which confirmed that the fuel changes from liquid to gas under these circumstances, leading to considerable changes in the surface and fuel temperature tendencies. These findings could help in the optimal design of scramjet engines regenerative cooling heat exchangers. [ABSTRACT FROM AUTHOR]

To propose a design guideline for a scramjet combustor at low cost, an ignition model for a scramjet combustor is required. Therefore, the overall aim of this study was to propose an ignition model that could be used to estimate the forced ignition limits within a cavity. In this study, we developed two forced ignition models (FIMs): the forced ignition model in the recirculating zone (FIMR) and the forced ignition model in the shear layer (FIMS). We evaluated the forced ignition limits using a micro-rocket torch within a scramjet combustor. The validated results of the FIMs indicated that the modified FIMR incorporating cavity refresh time (defined as the time required for complete replacement of all gases in the cavity) was the most appropriate method for calculating the forced ignition limit within the recirculation zone. However, in the case of Φ2.0 + 50 % He torch, the modified FIMR could not precisely estimate the forced ignition limit, showing that further modification to the present prediction model was necessary. The forced ignition process in the cavity was investigated using OH* chemiluminescence images, which indicated that the fuel mixture ignited in the shear layer; therefore, it was necessary to investigate ignition not only within the recirculation zone but also in the shear layer. In the case of Φ2.0 + 50 % He torch, the forced ignition limit could not be predicted by the FIMR, but it could be predicted by the FIMS. These results demonstrate that the FIMS provides improved prediction accuracy for the forced ignition limit in the scramjet combustors. [ABSTRACT FROM AUTHOR]

The combustion of the engine in the air vehicle causes the presence of charged particles in the jet stream, and the electric field displayed varies with engine type and fuel characteristics. In order to show the magnitude and distribution of charged particles in combustion products during scramjet operation, we constructed microphysical model and fluid calculation model based on the physical characteristics of supersonic combustion to simulate the variation of charged particles concentration in the scramjet combustion chamber and the charge density distribution outside the combustion chamber. The obtained results show that the charged particles in combustion products are mainly composed of ions, electrons, ion clusters, and charged soot particles. The electron concentration can be almost negligible at the exit of the combustion chamber. The other charged particles in the combustion products are ejected into the atmosphere with the jet. The charge density distribution of the jet has a certain symmetry in a short time due to the influence of the flow field. It is found that the type of fuel and the conductivity of the surface materials in the combustion chamber affect the electrified properties of the combustion products to some extent. This study has a reference value for further analysis of the detectable electric field generated during scramjet operation. [ABSTRACT FROM AUTHOR]

An airbreathing hypersonic vehicle(AHV) generally adopts a scramjet engine as its propulsion, which needs strict conditions of flight dynamic pressure. A new dynamic pressure control system is presented, in which the AHV autopilot directly uses dynamic pressure to track the dynamic pressure command. Firstly, the dynamic pressure model of the AHV is established and the state equations of dynamic pressure control are given. Then, the dynamic pressure control system is implemented using the LQR optimal control theory. The dynamic pressure error is augmented in order to add an integral control to track the dynamic pressure with zero steady error. The structure of the dynamic pressure control system is also obtained. The simulation results show that the dynamic pressure control system has a good performance for guaranteeing the dynamic pressure of the scramjet engine with the disturbance of drag and thrust considered.

Suppose that curved compression surface of inlet consists of segments. Two curved surfaces formed by equal compression angles of the micro-element segments and a slight increase in the compression angle of the micro-element segments are designed respectively. The numerical simulation method is used to compare the performance of two curved surfaces with the reference three-wedge compression surface. Select NASA classic test data, in order to determine the turbulence model and calculation method chosen by the numerical simulation Fluent software. The results show: the configuration of the segment compression angle deeply affects compression efficiency of the curved surface compression system. Pressure gradient distribution on the compression surface with constant compression angles segments is nearly constant along the incoming flow direction, and the curved compression surface easily resist the separation of the boundary layer compared to three-wedges compression surface. The approximate calculation method of the bending shock profile is given.

A key technological requirement for realizing a supersonic combustion ramjet (scramjet) engine is supersonic mixing enhancement. Streamwise vortices, which have larger circulation and break down more quickly into small-scale turbulent eddies, can be effectively used for supersonic mixing enhancement. In the present study, to clarify the condition for mixing transition in supersonic streamwise vortices, we performed visualization by the schlieren method as well as particle image velocimetry and hot-wire anemometry measurements of a supersonic mixing field that introduced streamwise vortices at the freestream Mach number M∞ = 1.8. It was confirmed that the mixing transition state of streamwise vortices was spread over the whole vortex region owing to the shear layer instability at the outer edge and the unstable swirling motion of the streamwise vortices. [ABSTRACT FROM AUTHOR]

Flame dynamics under various backpressure conditions were experimentally investigated using direct flame visualization, high-speed CH* chemiluminescence imaging, and wall pressure measurements. The stagnation pressure and temperature used in the present study were 100 kPa and 2500 K, respectively, with a freestream Mach number of 4.5. Rectangular scramjet models with and without a cavity were used to explore the effects of the cavity on flame dynamics when operating in scramjet mode, ramjet mode, and unstart. The flow rate of the ethylene jet was varied to impose backpressures corresponding to each operation mode. For both models, reverse flame propagation was observed for ramjet mode and unstart. For ramjet mode, flame fluctuation occurred within the isolator due to the coupling of fluid dynamics and combustion. The presence of a cavity enhanced combustion and reduced flame fluctuation in both scram and ramjet mode. The cavity promoted unstart because of the greater heat release from combustion. Further research using spatially resolved diagnostic techniques is needed to identify the flame locations for ramjet mode and unstart. [ABSTRACT FROM AUTHOR]

Due to the wide flammability range and low ignition delay of Hydrogen, it can burn rapidly and generate a large amount of thrust. It is for this reason alone that many researchers have promoted hydrogen as a potential fuel in scramjet engines. As H 2 is also environmentally safe and clean and can be produced from abundant sources, several researchers across the globe support the use of H 2 fuel. However, its lower volumetric energy density and higher flammability range also have an adverse effect in on-board storage system for aircraft propulsion applications. This review gives a brief representation of Hydrogen fueled scramjet engine as well the challenges associated with H 2 fuel. Additionally, the advantage of hydrogen as a fuel as compared to other hydrocarbon fuel is also discussed here thoroughly. • The motivations for selecting H 2 fuel in Scramjet Engines have been discussed. • Recent works in scramjet engines using H 2 fuel have been reviewed as well summarized. • Challenges faced by H 2 fuel is also discussed briefly. [ABSTRACT FROM AUTHOR]