Search Results (43)

As hypersonic aircraft applications become increasingly extreme, traditional regenerative cooling channels primarily using high-temperature alloys as wall materials can no longer simultaneously meet the dual requirements of thermal protection and lightweight design. This study, based on hypersonic environments with Mach numbers exceeding 8, selects five materials with significant advantages from metals, ceramics, and C/SiC composite materials to conduct a coupled design of wall materials for the flow and heat transfer characteristics of n-decane under 3 MPa pressure. The results show that the heat transfer ability of different material combination schemes is closely related to the thermal–physical properties of the materials, and the materials with obvious advantages in specific thermal–physical properties are dominant. Under a heat flux of 1.5 MW/m², the GH3128 + Cu composite scheme demonstrates a 17.5% increase in Nusselt number and a 17.6% improvement in comprehensive heat transfer coefficient compared to the traditional high-temperature alloy scheme. When the heat flux triples to 4.5 MW/m², the temperature variation of the GH3128 + Cu composite scheme is only 50% of that of the steel + C/SiC composite scheme. This indicates that multi-material coupling exerts both synergistic effects and inhibitory effects on flow and heat transfer characteristics, highlighting the importance of flexible material selection tailored to different tactical and technical requirements. Full article

This paper presents a comprehensive study on the control law design for the turbine-to-ramjet mode transition in an adaptive-cycle turbine-based combined cycle (TBCC) engine, aiming to mitigate the persistent “thrust gap” challenge. An integrated conceptual configuration of a hypersonic vehicle with a parallel-duct TBCC system, which replaces the conventional turbofan with a three-bypass adaptive cycle engine (ACE), is proposed. High-fidelity performance models for both the ACE and the scramjet are developed, with a Kriging surrogate model employed to accelerate the computationally intensive ACE simulations during the transition. A co-optimization framework is established, defining a comprehensive performance index that balances thrust tracking accuracy and control smoothness under rigorous intake-engine flow matching constraints. Using sequential quadratic programming (SQP), the control schedules for the ACE’s variable geometries are optimized. Comparative analyses reveal that the ACE, with its flexible bypass management and multiple adjustable mechanisms, can actively adapt its airflow demand to match the restricted intake supply. Consequently, the optimized ACE-based TBCC reduces total airflow fluctuation during the Mach 3–3.5 transition from 106% (conventional turbofan baseline) to 42.5%, while maintaining required thrust. This work quantitatively demonstrates the superior flow-handling capability of adaptive cycle technology, providing a viable and effective solution for ensuring stable and efficient mode transition in future hypersonic TBCC propulsion systems. Full article

Hydrocarbon fuels are extensively employed as coolants in the regenerative cooling systems of scramjet engines. However, the pyrolytic coking of hydrocarbon fuels at high temperatures introduces complex adverse effects on the flow and cooling processes. In this study, a numerical model was developed to investigate the coupling processes of fluid flow, heat transfer, pyrolysis and pyrolytic coking in the heated tube, under both a constant outer wall heat flux of 1.8 MW/m² and a constant outer wall temperature of 1150 K. The multi-step pyrolytic reaction mechanism and the kinetic coking model were applied to simulate the pyrolytic coking processes of RP-3. The results reveal that the amounts of catalytic coking and lateral growth exhibit significant differences in magnitude, as well as in their spatial and temporal variations. Under a constant outer wall heat flux, coking evidently increases the outer wall temperature and thermal resistance, leading to a narrowed flow passage and a reduction in the residence time and RP-3 conversion rate. Under a constant outer wall temperature, coking decreases the heat absorption flux, resulting in a lower fluid temperature, which primarily affects the efficiency of the endothermic pyrolytic reaction. The results obtained in this research can provide practical insights for the development of regenerative cooling technology. Full article

Scramjet operation requires a comprehensive understanding of the internal flowfield, encompassing fuel–air mixing and combustion. This study investigates transient flow and flame development within a HIFiRE-2 scramjet engine combustor, which features two opposing cavities and dual sets of fuel injectors—the upstream (primary) and downstream (secondary) injectors. These cavities function as flameholders, creating circulating flows with elevated temperatures and pressures. Shock waves form both ahead of fuel plumes and at the diverging and converging sections of the flowpath. Special attention is given to the interactions among these shock waves and the shear layers along the supersonic core flow as the system progresses towards a quasi-steady state. Driven by increased backpressure, bow shocks and disturbances induced by the normal, secondary fuel injection and the inclined, primary fuel injection move upstream, amplifying the cavity pressure. These shocks generate adverse pressure gradients, causing near-wall flow separation adjacent to both injector sets, which enhances the penetration and dispersion of fuel plumes. Once a quasi-steady state is achieved, a feedback loop is established between dynamic wave motions and combustion processes, resulting in sustained entrainment of reactive mixtures into the cavities. This mechanism facilitates stable combustion in the cavities and near-wall separation zones. Full article

Local overheating in cavities seriously threatens the safe operation of scramjet engines. To investigate the formation mechanism and control methods of local overheating, this paper conducts numerical simulations on the cooling process of cavity sudden expansion (S-E) structures. A three-dimensional numerical model coupled with pyrolysis reactions is established and validated through experiments. The effects of thermal pyrolysis reactions and cooling channel parameters on flow distribution are analyzed, and comparative studies with different channel parameters are performed. The results show that S-E structures are prone to uneven fuel flow distribution, leading to local over-temperature phenomena, and thermal pyrolysis reactions will aggravate this phenomenon to a certain extent. Increasing the aspect ratio of the channel can enhance the pressure drop at the inlet of the S-E structure and improve the uniformity of flow distribution. When the aspect ratio increases from one to eight, the mass flow distribution deviation ϕ_m decreases from 0.954 to 0.181. More uniform flow distribution under a larger aspect ratio avoids local over-temperature in the S-E structure, and reduces the coking risk caused by local excessive pyrolysis. This work reveals the fundamental characteristics of cooling heat transfer in the S-E structure of Scramjet engines and can provide recommendations for the design of cooling channels. Full article

The pylon has been identified as a highly promising method for enhancing mixing efficiency in scramjet combustors. This work systematically assessed the impact of spanwise, streamwise, and oblique multi-pylon combinations in a supersonic cold flow through numerical simulations, employing pylon-aided ethylene fuel injection under low dynamic pressure conditions. The Reynolds-averaged Navier–Stokes (RANS) equations with the SST k-ω turbulence model are applied during the simulation. Numerical results reveal that, in comparison to the streamwise combination, the spanwise combination exhibits superior flow field characteristics in terms of mixing efficiency, penetration depth, and total pressure loss. For a given injection condition, an optimal distance between pylons exists in the spanwise combination, with the angle between two pylons having minimal influence on mixing efficiency. The oblique multi-pylon combination yields poorer mixing enhancement efficiency and fuel penetration but incurs less total pressure loss in the near field when compared to the spanwise combination. Additionally, the oblique multi-pylon combination demonstrates enhanced mixing efficiency further downstream of the injector than the spanwise combination. This investigation into fuel injection schemes based on multi-pylon combinations offers valuable insights for the structural design of scramjet engines. Full article

Two-staged fuel injection configuration for scramjet combustors has been shown to be effective in distributing heat release in the combustor for preventing the unstart transition of the engine by suppressing peak pressure while increasing the pressure thrust. In this study, the effect of fuel species on combustion characteristics in a two-staged fueled scramjet combustor was investigated. Wall pressure measurements in a two-staged fueled scramjet combustor were conducted in a combustion wind tunnel facility with fuels having different reactivity, such as H₂ and CH₄. Reynolds-Averaged Navier–Stokes/Large Eddy Simulation (RANS/LES) hybrid simulations were performed to verify the interaction characteristics between the primary and secondary combustion zones for different fuels. The experimental results confirmed that pressure peaks at injections were clearly separated in the hydrogen case, while these interacted with each other in the methane case with a lower reactivity than H₂. The RANS/LES Hybrid analysis predicted this effect of fuel reactivity on the pressure distribution, namely, the heat release delay of the first stage fuel caused the interaction with the second stage fuel heat release. The results indicate that the need to design the staged fueled combustor, i.e., the injection stage interval accordingly to the reactivity of the fuel. Full article

The burning chamber wall of the ramjet engine is facing an extremely thermal environment during normal conditions. Thermal protection measures must be taken on the wall surface of the combustion chamber. At the same time, the aircraft faces high-power electrical demand problems under high-speed cruising states. To address these issues, a second fluid-closed Brayton cycle system was introduced in this paper. Helium was utilized as the secondary fluid medium, while kerosene was used as the final heat sink. The ramjet engine chamber wall was cooled by the helium cycle system. At the same time, part of the heat absorbed by the helium cycle was transformed into electric power by a generator. This work proposes a new method of thermal management in a closed cycle. Unlike traditional methods, this proposal can regulate the mass flow rate of helium based on the requirement of heat load. A zero-dimensional numerical calculation method was established for thermodynamic analysis. The results show that as the equivalence ratio of 0.8~1.5 for the kerosene flow rate, the system can suffer the thermal load of 200~350 kJ/kg on the combustion chamber wall at the maximum kerosene allowable temperature. To ensure the normal operation of the circulating system, the mass flow ratio between the helium and the air changes from 0.02 to 0.045. Compared with the direct kerosene cooling method, the second fluid circulation method leads to the kerosene equivalent saving ratio by 2% to 14%; at the same time, such a system could generate 160~500 kJ/kg of electrical energy. This new thermal management method can achieve kerosene saving, electric power generating and suffering more thermal loads under the premise of satisfying normal work. Full article

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. Full article

This study investigates the operational characteristics of the Rotating Detonation Engine (RDE), with a focus on fuel injector design. Inspired by the similarity between the fuel injection structure of RDE and the Jet in Supersonic Crossflow (JISC) of a scramjet, experimental research on fuel injectors with jet penetration was conducted. Five injectors were designed, each with a fixed fuel injection area or injection hole diameter. Experiments determined practical injection areas, and an empirical correlation was used to calculate jet penetration heights. Under conditions of a total mass flow rate of 105 ± 5 g/s and an equivalence ratio of 1.05 ± 0.1, combustion modes were analyzed. Initial detonation occurrence was assessed through pressure history, with a detailed analysis via image post-processing. The results indicated that the injector D4N15, with the highest jet penetration height, exhibited deflagration, while D4N23 showed chaotic propagation. The injector D2N60 demonstrated relatively unstable behavior in sustained detonation cases. Thrust comparisons revealed that D4N30, with wider hole spacing and higher jet penetration height, exhibited approximately 12.5% higher specific impulse compared to D1N240. These outcomes confirm the significant impact of jet penetration height and hole spacing on detonation propagation and engine performance. Full article

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. Full article

A micro-pulse detonation engine (μPDE) was designed and installed to a direct-connect scramjet combustor of Pusan National University (PNU-DCSC). The active excitation on the scramjet combustor was experimentally studied using the μPDE operating at frequencies of 10–20 Hz. A vitiation air heater (VAH) was used to supply high-enthalpy vitiated air to the isolator and the scramjet combustor at a Mach number of 2.0, with a total temperature of 1600 K and a total pressure of 1.68 MPa. The exit of μPDE was located at the center of the cavity of the scramjet combustor. Active excitation was performed at equivalence ratios of 0.111 and 0.163, and characteristics were analyzed through Schlieren recording and bottom wall pressure measurement. As a result, when the detonation emitted from the μPDE entered the scramjet combustor, it instantly formed a shock train and moved forward within the scramjet combustor. The flame instantaneously changed from the cavity shear layer flame to the cavity flame. Through bottom wall pressure measurement, it was also observed that active excitation resulted in a significant pressure increase near the cavity compared to when active excitation was not performed. This revealed combustion characteristics, indicating improved combustion efficiency from the pressure increase in the scramjet combustor. Full article

To enhance the heat-transfer performance of scramjet engines, a numerical simulation was conducted on the heat-transfer process of RP-3 aviation kerosene under supercritical pressure within a horizontal micro-fine circular tube. The intrinsic mechanism of the heat-transfer process was analyzed, summarizing the impacts of mass flux, inlet temperature, and gravitational acceleration. Furthermore, four commonly used buoyancy criterion numbers were compared and evaluated. The results indicate that the heat-transfer process can be divided into five phases: heating inlet phase, normal heat-transfer phase, heat-transfer deterioration phase, heat-transfer enhancement phase, and high-temperature normal heat-transfer phase. The heating inlet phase is significantly influenced by the inlet temperature, while the heat-transfer deterioration is affected both by the thermal property variations of the aviation kerosene and the buoyancy effects. Lower mass flux and hypergravity conditions all exacerbate heat-transfer deterioration. Inlet temperature, however, does not affect the heat-transfer pattern. Among the criteria, Gr_q/Gr_th provides the best prediction of buoyancy effects in horizontal circular tubes. Full article

This study focuses on the three-dimensional flow and combustion characteristics of a cavitied scramjet engine with multi-position injection. A single-equation large eddy simulation (LES) turbulence model is employed, with a detailed reaction mechanism for hydrogen combustion, as described by Jachimowski. The combustion characteristics of hydrogen in the scramjet combustion chamber are analyzed. Based on the combustion chamber model, the influence of different equivalence ratios, injection timing, injection positions, and injection pressures on the flame formation and propagation process are compared. The results indicate that within a certain range, an increase in the equivalence ratio enhances the combustion intensity and chamber pressure. In the case of multi-position injection, the order of injection from different nozzles has little effect on the final flame stabilization mode and pressure distribution. The opposite-side distribution of nozzles can effectively improve the fuel efficiency and the internal pressure. Furthermore, when the nozzles are closely placed in the opposite-side distribution, the combustion efficiency increases, although this leads to a higher total pressure loss. In scenarios where the fuel injection duration is short, an increase in the injection pressure at the upstream nozzles of the cavity results in a higher local equivalence ratio, as well as reduced fuel mixing and ignition time. Full article

Scramjet based on solid propellant has become a potential choice for the development of future hypersonic vehicles. In this paper, a boron-containing solid rocket scramjet based on the central strut injection was proposed, and the ground direct-connect experiment with the equivalence ratios of 0.43 to 2.4 under the flight condition of Mach 6, 25 km was carried out. The pressure and flow rate over time were measured in the experiment. The results show that the engine can realize stable supersonic mode or subsonic mode combustion by changing the gas flow rate. The engine can effectively increase the combustor pressure, reduce the unstable combustion time, and advance the strong combustion position by increasing the gas flow rate. The engine achieved high combustion efficiency when the equivalence ratio was about 1, with a maximum of 88.28%. A numerical simulation analysis was also carried out in this paper. Compared to the experimental results, the pressure error obtained by numerical simulation was less than 4%, and the typical position error was less than 3%, suggesting that the simulation model can be used to predict the behavior of scramjet. Full article