Air-breathing Engines Research Articles

Study of ways to influence the combustion process in an ejector pulsating air-breathing engine in order to improve its specific performance

The problem of studying the operating cycle of a pulsating air-breathing engine is important in the context of the occurrence of vibration combustion, as well as the transition of combustion to detonation. The characteristics of a detonation-combustion propjet engine should be significantly higher than the characteristics of air-breathing engines. The article presents the results of a search for ways to increase the specific performance of a valveless ejector pulsating air-breathing engine. The results of fire tests of engines with partial conversion of the original fuel and generation during the operating cycle of peroxides and active centers obtained during cold-flame reactions are presented. As a result of the experiments, a displacement of the combustion zones in the engine downstream into the region of the recirculation zone was obtained. The numerical modeling performed showed an agreement with experiment acceptable for engineering calculations. An increase in engine jet thrust was obtained due to partial fuel conversion and a shift in the ignition zone. The transition of combustion to deformation leads to a significant reduction in engine life. The results obtained and a new method of influencing the work process based on a shift in the combustion zone are discussed. It is shown that an increase in the combustion rate of the air-fuel mixture leads to a change in the acoustic feedback mechanism.During the work, research was carried out to find a reactor design combined with a combustion chamber to produce a small proportion of C2H2 and H2. To obtain these products, a production method based on the oxidative pyrolysis of methane was chosen. A technical solution was found; it consisted of installing a finned channel cooled by air on the rear wall of the combustion chamber of the reactor. Cooled plates were placed in the front of the combustion chamber and in the second mixer. In these devices, partial cooling of the emitted products of incomplete combustion occurred and the process of methane decomposition was stopped, which made it possible to increase the concentration of C2H2 and H2 during the combustion chamber purge stroke and thereby increase jet thrust by 25–30%.

Progress of Experimental Studies on Oblique Detonation Waves Induced by Hyper-Velocity Projectiles

Oblique detonation waves (ODWs) are hypersonic combustion phenomena induced by oblique shock waves. When applied to air-breathing engines, ODWs offer high thermal cycle efficiency, adaptability to a wide range of flight Mach numbers, and the advantage of a short combustion chamber, making them highly promising for hypersonic propulsion applications. Despite numerous numerical studies on the heat release and multi-wave flow mechanisms of ODWs, practical applications of oblique detonation engines (ODEs) remain limited due to several technical challenges. These challenges include generating the required high-velocity test environments, achieving effective fuel and oxidant mixing, and measuring the flow field structure in hyper-velocity and high-temperature flows. These limitations hinder the development of ODEs, underscoring the importance of experimental research, particularly for understanding the initiation and propagation mechanisms of ODWs. One of the primary experimental techniques involves inducing oblique detonation using high-velocity models. This method is extensively used to study the initiation process, shock structure, initiation criteria, and ODW propagation. It is advantageous because the state of the experimental mixture is controllable, and the model state can be precisely measured. This paper reviews studies on oblique detonation induced by hyper-velocity projectiles, presenting advances in experimental methods, detonation wave structures, unsteady processes, and initiation characteristics. Additionally, we discuss the deficiencies in existing studies, noting that the current measurement methods fall short of the requirements for observing the ODW initiation process, propagation process, and fine structure. The application of advanced combustion diagnostic techniques and the exploration of the relationship between initiation processes and criteria are crucial for advancing our understanding of ODW initiation and stabilization mechanisms. Finally, we summarize the current state of experimental facilities and measurement techniques, providing suggestions for future research on the measurement of shock waves and chemical reaction zones.

Modelling and experimental analysis of Aero-Engine performance and exhaust emission characteristics Fueled with green fuel blends

This study aims to assess the impact of using green fuel in place of biodiesel on the performance and exhaust emissions of air-breathing engines. GasTurb 13 is utilized to forecast the engine’s performance (kingTech 180 k turbojet engine). Catalytic deoxygenation of vegetable oils produces green diesel, offering an alternative to biodiesel. Physiochemical properties of GD blends (PME30GD20, PME20GD30, PME10GD10) were analyzed, and GasTurb-13 software predicted engine performance and emissions, validated with experimental data. The results show that GD10PME10 exhibited higher density (767.6 g/m3) than pure green fuel, while viscosity dropped by 53.85 % compared to PME. GD outperformed Jet-A1 by 0.74 % in heating value, with GD10PME10 having the highest at 42.63 MJ/kg. Engine performance measures included thrust, mass fuel flow, thrust-specific fuel consumption, and exhaust gas temperature. GD10PME10 showed a 12.5 % thrust increase and the lowest TSFC (95 g/kN.s) at 80,000 RPM compared to Jet-A1. GD20PME30 achieved the lowest EGT (550 °C) at the same RPM. Regarding emissions, GD20PME30 emitted the lowest CO (100,000 RPM, 150 ppm) and 0.5 % less CO2 (90,000 RPM) than Jet-A1. GD10PME10 produced the lowest NOx (12.5 ppm) at maximum speed, while Jet-A1 emitted the least NOx (7.5 ppm) at 120 k RPM. Overall, the data suggested that green fuel might increase the physiochemical properties of biodiesel blends, hence improving the fuel’s capacity to burn more effectively in aero-engines.

Effect of jet splitting using passive strut on the performance and thermoacoustic characteristics of a scramjet combustor

A scramjet engine offers a potential route to achieve supersonic speeds using airbreathing engines. Achieving proper mixing and combustion poses a challenge due to the supersonic inflow of air. Researchers have explored multi-strut configurations to tackle this issue. However, multiple struts supplying fuel inefficiently can lead to fuel loss and reduced efficiency. Alternatively, utilizing a multi-strut setup passively could enhance combustion and mixing efficiency. In this study, two types of jet splitting passive strut configurations were investigated computationally with the improved delayed detached-eddy simulation turbulence model. Implementation of passive strut altered vortical structures, influencing mixing and combustion performance. The splitting of the jet introduces large-scale vortices downstream. Strategically placing the passive strut in the wake of the combustion zone was found to improve both mixing and combustion efficiency. Acoustic loading was seen to increase with the introduction of passive strut. It was observed that the diamond-shaped passive strut has the highest combustion efficiency; however, it suffers from higher acoustic loading. The dynamic mode decomposition analysis revealed the coupling frequency of fluctuating pressure and heat release rate, which causes thermoacoustic loading. Overall, passive strut placement significantly influenced combustion, mixing, and thermoacoustic properties, highlighting the importance of considering passive strut configurations in design optimization for scramjet engines.

Improve the thermal performance of precooler by enhancing the utilization of chemical heat sink: Refined design method and characteristic analysis

The chemical heat sink of the coolant plays a crucial role in enhancing the performance of the precooler, which affects the overall performance of the air-breathing precooled engines. The conventional thermal design method for precoolers cannot accurately evaluate the chemical heat sink of coolants. To address this crucial issue, this paper proposes a refined model that considers the chemical reaction processes of n-Decane, based on a two-dimensional discretization unit model and the plug flow reactor model incorporating molecular-level chemical reaction mechanisms. The validation of this model is conducted by comparing it with publicly accessible data and computational fluid dynamics simulations, thereby demonstrating its excellent accuracy. Subsequently, an analysis is conducted on the factors influencing the chemical heat sink release of the coolant within the precooler, and enhanced design criteria for optimizing chemical heat sink utilization are proposed. Building upon these considerations, the performance of a precooler with a design point of Mach number 5 under off-design conditions was analyzed. The findings indicate that employing rational design methods can enhance the chemical heat sink efficiency of endothermic fuels by 8.0%, along with a concurrent 10.3% increase in the total heat sink. This underscores the significance of considering the impact of the chemical heat sink of endothermic fuels in the design of precoolers. In the case presented in this paper, the contribution of the chemical heat sink to the total heat sink within the n-Decane precooler could reach 35.3%.

Influence of aerodynamic valve structural parameters on detonation initiation and back pressure characteristics of air-breathing pulse detonation engine

In order to enhance the stability of detonation initiation in air-breathing pulse detonation engines (APDE), while simultaneously mitigating the non-stationary back pressure disturbance due to the huge pressure gain in the APDE, a central blunt body aerodynamic valve (aero-valve) was developed and tested with a single-tube pulse detonation combustor (PDC). The detonation initiation and back pressure characteristics of PDC were then discussed. In order to reveal the influence of the aero-valve's structural parameters on the cold-state filling, detonation initiation, and back pressure characteristics of PDC, numerical simulations were carried out based on the experimental setup. The numerical results indicated that the cold-state flow characteristics within the aero-valve was predominantly influenced by the inner diameter of the aero-valve shell (Das), while the cold-state flow inside the combustor was primarily dominated by the axial distance from the thrust wall to the aero-valve shell (Lx) and the inner diameter of the aero-valve outlet (dout). And dout has a larger impact on the cold-state flow inside the combustor. During the hot state, reducing Lx or dout is beneficial for both the generation of detonation and the suppression of back pressure, with the most significant gain effect coming from decreasing dout. The optimal combination scheme for the aero-valve is identified as Das/R = 4.5, Lx/R = 0.25 and dout/R = 0.7, where R is the detonation combustor radius. The PDC equipped with this optimized aero-valve not only has better detonation initiation characteristics, but also a smaller pressure oscillation ratio.

Optimization strategies for the design of pre-cooler based on the synergistic air-breathing rocket engine

A foundation for the engineering of pre-coolers is provided to understand the flow and heat transfer characteristics of air pre-coolers. Using the pre-cooler in the combustor of the Synergistic Air-Breathing Rocket Engine (SABRE) as the research object and selecting the smallest periodic unit of the model, numerical methods were adopted to study the flow field, outlet temperature, total pressure loss, heat transfer coefficient, and heat transfer power under the different number of tube rows, pitch, and helium/air heat capacity ratios. The number of tube rows increased from 7 to 15, the tube pitch increased from 1.5 mm to 3.5 mm, and the helium/air capacity ratio increased from 1.06 to 1.64. The results show that increasing the tube pitch reduces the outlet velocity and decreases the heat exchange efficiency, but the total pressure loss is reduced. Increasing the number of tube rows results in a linear increase in total pressure loss, but the heat exchange efficiency also increases. The increasing trend slows down when the number of tube rows exceeds 11. An increase in the heat capacity ratio can reduce the total pressure loss, increase the heat exchange power, and increase the amount of helium required. It can provide a reliable technical foundation for the practical engineering design of similar pre-coolers.

Optimal design of supercritical He–H2 PCHE in SABER system by multi-objective genetic algorithm

A double-circle straight-channel printed circuit heat exchanger (PCHE) was first employed in a synergistic air-breathing rocket engine (SABER) system for supercritical He and H2 heat transfer. The heat transfer mechanism of the supercritical flow in the PCHE channel was numerically analyzed. For considered the pressure drop, heat transfer efficiency, and ratio of heat flux to weight, simultaneously, the design of the PCHE is multi-objective genetic optimized. The results shown that the supercritical H2 flow is chaotic near the pseudo-critical point, which is a coupled effect of buoyancy and gravity. Chaotic flow leads to an asymmetrical temperature distribution, which deteriorates the heat transfer performance. For 27 numerical experimental cases designed using the center composite surface method, the determine factors of the regression models of the three objectives for both cold and hot sides were all above 92 %. The Pareto optimal solutions for the supercritical He - H2 PCHE design and performance were obtained based on the nondominated sorting genetic algorithm II. From a comprehensive view of the three targets, the optimal designs were the A-4 and B-4 solutions for the cold and hot sides, respectively.

Insights into thermodynamic performance of a hypersonic precooled air-breathing engine with a complicated multi-branch closed cycle

An advanced precooled airbreathing engine with a closed Brayton cycle is a promising solution for high-speed propulsion, of which the Synergetic Air Breathing Rocket Engine (SABRE) is a representative configuration. The performance of the latest SABRE-4 cycle was analyzed in this paper. Firstly, a relatively complete engine performance model that considers the characteristics of turbomachinery and heat exchangers was developed. Then, Sobol’ global sensitivity analysis of key performance parameters was carried out to identify the most influential design variables. Optimal specific impulses under different target specific thrusts were obtained by particle swarm optimization, of which the thermodynamic parameters corresponding to a specific thrust of 1.12 kN·s·kg−1 and a specific impulse of 3163 s were chosen as the design values. Four different control laws were analyzed in contrast, and the charge control method had the strongest ability of thrust regulation as well as maintaining a favorable specific impulse performance. Finally, working characteristics under the charge control and over a typical flight envelope were calculated, in which the average value of the maximum specific impulse was as high as 5315 s. This study would help to deepen the understanding of SABRE-4 thermodynamic characteristics and other precooled airbreathing engine cycles with similar layouts.

Selected Papers from 4th National Aerospace Propulsion Conference NAPC 2022

The National Aerospace Propulsion Conference (NAPC) is a national conference dedicated to aerospace propulsion technologies. Two erstwhile popular conference series, the National Propulsion Conference (NPC) and the National Conference on Air-Breathing Engines (NCABE) were merged to result in NAPC. This biennial conference is intended to bring together the entire propulsion community spanning the industry, academia and the research labs nationwide. The aerospace propulsion-centred conference is considered an ideal opportunity to showcase one’s research activities with peers and foster future collaborations through networking. The 4th edition of the conference, NAPC-2022, comprised a keynote lecture, several topical invited lectures and presentations of contributed papers in parallel sessions. This special issue contains eight manuscripts selected based on peer-reviews of selected papers presented during the conference. The manuscripts broadly cover several aspects pertaining to aerospace propulsion that includes airbreathing engine components such as intakes, compressors, combustors and nozzles. From the rocket propulsion (non-airbreathing), there are papers that investigate the rocket motor grain port alignment and another paper on the heat transfer characteristics of a supersonic rocket nozzle.

Influence of passive strut on the mixing and combustion performance enhancement

The scramjet engine is one of the possible means of reaching supersonic speed using airbreathing engines. As the inflow of air is at supersonic speed, proper mixing and combustion is a challenge. Researchers have used multi-strut arrangements to address the problem. However, the additional supply of fuel from multiple struts with inefficient mixing and combustion leads to losses in terms of fuel and efficiency. Alternatively, the multi-strut configuration could be used to improve the combustion and mixing efficiency passively. In the present investigation, two strut configurations, a fuel injection strut, and a passive strut, have been used. The computational study was done by solving steady Reynolds averaged Navier-Stokes (RANS) equations along with the Shear Stress Transport (SST) k-ω turbulence model, species, and ideal gas equation. Eddy dissipation model was used to solve reacting flow. The introduction of a passive strut in the wake of the fuel injection strut leads to vortices upstream and downstream of the passive strut. These vortices were observed to be responsible for enhancing the mixing and combustion. It was found that judicious placement of the passive strut in the wake of the combustion zone can enhance both the mixing as well as combustion efficiency with minimal loss in total pressure.

Multi-objective optimization of a hypersonic airbreathing vehicle

Multi-objective optimization of a hypersonic airbreathing engine (scramjet technology) was carried out with the aim of maximizing thrust and minimizing drag while satisfying a series of design constraints, such as avoiding unstart (blockage of supersonic flow within the combustion chamber) by ensuring that the pressure ratio across the shock waves remains below the adverse pressure gradient given by the Korkegi limit, geometry correction to achieve shock on-lip condition, and temperature and pressure requirements at the inlet exit. Using the relations presented in the literature, pressure and viscous drag are estimated analytically. The analytical approach is verified against computational fluid dynamics data from Ansys Fluent to solve two-dimensional compressible Reynolds-averaged Navier–Stokes flow equations, with transition shear stress transport as the turbulence closure model. Comparing the total drag and the flow properties at the combustion chamber entrance shows the model's feasibility for the optimization approach. Three different approaches were conducted to formulate the multi-objective function to determine the one that can find the highest number of geometries satisfying the Korkegi limit with the highest net thrust. The best approach was the multi-objective function formulated with the uninstalled thrust, total pressure recovery, and pressure drag, concentrating the search in the region with greater uninstalled thrust and lower drag and nearly doubling the value of net thrust compared to the first formulation, which uses the uninstalled thrust, pressure drag, and viscous drag.

Numerical investigation of conjugated heat transfer mechanisms of supercritical hydrogen-helium in PCHE channels

In this study, straight and z-shaped channel PCHEs are employed as hydrogen-helium heat exchangers in Synergistic Air-breathing Rocket Engine (SABER). Conjugated heat transfer mechanisms and characteristics of supercritical hydrogen-helium are studied numerically. The results of the straight channel indicate that when the mass flow rate of the cold (H2)/hot (He) side is 38/280 kg/(m2·s), a greater comprehensive performance of flow and heat transfer can be obtained. The influence of gravity on the hot side is significantly less than that on the cold side. The increase in plate thickness enhances heat transfer, while the effect of rib thickness is the opposite. As the bend angle decreases, the heat transfer coefficient of the z-shaped channel becomes higher. The heat transfer characteristics of the cold side vary more significantly with the bend angle than that of the hot side.

Design space investigations of scramjet engines using reduced-order modeling

Conceptual design studies performed with reduced-order approaches allow feasibility and sizing considerations of air-breathing engines to be configured at an affordable computational cost. The present study is devoted to exploring the design space of a dual-mode ramjet engine operating in scramjet mode by means of reduced-order analysis to assess the effects of propulsive system design configurations on component level and overall performance characteristics. The approach proposed in this work combines axisymmetric flow configuration used for the design of supersonic/hypersonic intakes and solutions of one-dimensional flow governing equations coupled with finite-rate chemistry and thermophysical properties tables in the numerical domains of the combustor and nozzle components. The scramjet design space is generated by varying parameters which are flight Mach number and altitude, intake truncation angle, intake exit Mach number and equivalence ratio. Performance outputs of total pressure recovery factor, compression ratio, captured air mass flow rate, intake startability index, thrust, specific impulse, fuel consumption and overall efficiency are computed for each design scenario. The generated database is visualized via performance maps and analyzed in terms of propulsive characteristics. A feature importance study is also conducted to quantify the effects of design parameters on the propulsive performance.

Influence of velocity on the transient auto-ignition of ethylene jet in a hot coflow

The difficulty of ignition is one of the main problems in the combustion chamber of the air-breathing combined engine. Based on this, the transient auto-ignition in the turbulent non-premixed flames of ethylene in a high-speed jet in hot coflow was studied using a newly developed burner. High-speed hydroxyl (OH) planar laser-induced fluorescence (up to 10 kHz) and large eddy simulations (LES) were performed to investigate the auto-ignition process. The experimental conditions included injection velocities of 334–491 m/s, and the effect of jet velocity on the mixing process, ignition process, and flame stabilization mechanisms were investigated. The results showed that the numerical simulation results were consistent with the experimental data. The velocity could affect the auto-ignition position, whereas it had no effect on the auto-ignition time. The increase in fuel jet velocity is beneficial to the mixing of fuel and high-temperature gas and helps to increase the premixed combustion zone during the formation of fire kernels. Moreover, as the velocity increased, the flame status changed from a transitional flame to a lifted flame, the growth rates of the flame volume and total heat release increased, and the flame stabilization mechanism changed from flame propagation to auto-ignition.

Numerical investigation on the heat transfer of air/helium precooler for air-breathing pre-cooled engine

Installing a precooler behind the intake is an effective approach for hypersonic air-breathing pre-cooled engine to cool the hot incoming air. Synergetic air-breathing rocket engine is a revolutionary hypersonic air-breathing pre-cooled engine with complex thermodynamic cycle i.e. air cycle, helium cycle. Air/helium precooler is a key component and its configuration and operating condition have great effect on the performance characteristics of air-breathing pre-cooled engine. Thus, the minimum periodic flow and heat transfer model of the precooler are established. The effects of key parameters on the heat transfer performance of precooler are numerically studied. The results indicate that: when the tube row number increases from 7 to 15, the average heat transfer coefficient of air side decreases by 57%, the heat exchange rate increases by 19%, and effectiveness increases by 18.4%. The tube transverse pitch can enhance the heat transfer coefficient of air and helium side, while the heat exchange rate decreases by 33 % when the tube transverse pitch increases from 1.5 to 3.5. The helium inlet velocity can improve the heat transfer performance of precooler and reduce the flow resistance of air side.

Supercritical pyrolysis of in-house developed endothermic fuel and estimation of coke and endothermicity

The present investigation includes the development of potential endothermic fuels for regenerative cooling in air-breathing supersonic engines. The endothermic characteristics of two multi-component fuels, namely AF-1 (a boiling range of157–234 °C) and AF-2 (a boiling range of155–207 °C), were studied by performing pyrolysis experiments using a flow reactor under supercritical conditions, and the results are compared with a single component n-dodecane hydrocarbon Various techniques, such as ASTM D86 distillation, aniline temperature, calorific value, GC, GC-MS, thermogravimetric, and non-dispersive infrared radiation spectroscopy, were used to characterize the feed and product properties. At 700 °C and 55 bar pressure, the cracking conversion of AF-1, AF-2, and n-dodecane feed were 13.2, 15.6, and 25.1 %, respectively. The coke deposition rate was in the order of AF-2 > AF-1 > n-dodecane. In the work, a composition-based enthalpy estimation approach was adopted in estimating the endothermicity of the multi-component fuels. The endothermic heat sink capacity of the in-house developed AF-2 fuel (475 kJ/kg) was marginally higher than the AF-1 (456 kJ/kg) and about 15 % lower than n-dodecane (560 kJ/kg) at 700 °C.

Dynamic Forced Performance of an O-Rings Sealed Squeeze Film Damper Lubricated With a Low Supply Pressure and a Simple Method to Quantify Air Ingestion

Abstract Contemporary squeeze film dampers (SFDs) in air-breathing engines are short in length to limit weight and part count and lubricated with a low feed pressure to reduce oil storage and pumping power. In SFDs, O-rings (ORs) restrict side leakage and increase the viscous damping while a adding a modest centering stiffness. Continuing a long-term project characterizing SFDs for aircraft engines and extending the original work (San Andrés and Rodríguez, 2021, “On the Experimental Dynamic Force Performance of a Squeeze Film Damper Supplied Through a Check Valve and Sealed With O-Rings,” ASME J. Eng. Gas Turb. Power, 143(11), p. 111011) the paper details measurements of the forced performance of an OR sealed damper (OR-SFD) with diameter D = 127 mm, land length L = 0.2 D, and radial clearance c = 0.0022 D. Lubricant ISO VG 2 supplied at 0.69 bar(g) fills an upstream oil plenum and flows into the middle of the land through a single orifice configured with a check valve. Measurements of applied single-frequency dynamic loads, along with the ensuing damper displacements and accelerations serve to identify the parameters of the test structure, ORs, and SFD. The tests encompass centered whirl motions with amplitude r = 0.05–0.45c, and a range of whirl frequencies, ω = 10–130 Hz. Note the squeeze velocity vs = rω reaches 102 mm/s. The ORs force coefficients are nearly invariant with frequency but do depend on the orbit amplitude. The ORs' stiffness (KOR) decreases by 75% as the motion amplitude increases, r → 0.45c, likely due to the large elastic deformations and slow recovery of the ORs material. For small amplitude motions (r/c = 0.05 and 0.10), the ORs damping coefficient (COR) is ∼10% of the overall coefficient for the lubricated system (CL), while for r/c > 0.25, COR ∼ 0.03CL. For small amplitudes of whirl (r ≤ 0.25c), the SFD experimental viscous damping (CSFD) and added mass (MSFD) coefficients, identified over a shorter frequency range (vs<30 mm/s) equal theoretical magnitudes for a fully sealed damper. As the orbit size grows to r = 0.45c MSFD drops nearly by 75% and CSFD decreases by ∼40% due to the onset of both lubricant cavitation and air ingestion occurring for squeeze velocities vs ≥ 24.5 mm/s, as also seen in the recorded dynamic pressures and video recordings of a bubbly mixture leaving the damper. A comprehensive flow model predicts CSFD and MSFD about 8% and 12% larger than the experimentally identified coefficients. A novel approach enables the estimation of the gas volume fraction (GVF) generated in the damper which rapidly increases as vs grows. The simple procedure draws into a deflated balloon the material contents in the squeeze film, weighs the sample, and identifies its volume to produce an estimate of the GVF. The procedure to quantify the GFV will assist in the validation of predictive tools.

Numerical study of a scramjet isolator performance under different sidewall compression angles

Scramjet engines, also known as supersonic combustion ramjet engines, are frequently regarded as a compelling alternative for launching payloads into Earth's orbit. These air-breathing engines have streamlined designs with minimal movement of components. However, the successful design of scramjet engines necessitates overcoming various challenges such as managing the high heat fluxes and pressure loads exerted on the engine walls. Additionally, addressing issues such as shockwave-boundary-layer interactions and the potential occurrence of choked flow within the isolator channel are critical considerations during the scramjet design process. Therefore, this study aims to evaluate sidewall compression in the isolator region to deal with the high heat fluxes and pressure loads inside the scramjet isolator. In addition, this work also investigates how the variation in the angle of attack influences the mass flow rate of the intake and at which range of the angle of attack the intake becomes choked. The CFD analyses include contour images of properties such as Mach number, total pressure, heat flux, and pressure distribution on the walls, and the calculation of performance parameters, including the analysis of the second law of thermodynamics. The study involved varying the compression angle within the range of 4° to 10°. The results of this study demonstrate that implementing sidewall compression in the isolator region allows for the effective management of the position of the heat flux and pressure peaks on the upper wall of the isolator. Regarding the pressure distribution along the upper wall of the isolator, the 10°case presented a pressure peak of approximately 130000 Pa while the 4°case presented 155000 Pa. In addition to this significant decrease in the pressure peak value, its location also changed, with an increase of approximately 8 mm downstream of the isolator by decreasing the compression angle from 10° to 4°. This engineering approach presents a viable solution for mitigating the challenges posed by high heat flux and pressure loads in the intake section. The cost of applying such a solution is to decrease the intake performance – a decrease of approximately 30 % in the isentropic efficiency when comparing a case with no sidewall compression with the sidewall compression cases. In the choked flow study, angles of attack ranging from 4 to 30°were considered. The analysis shows that the choked-flow condition gradually occurs as the angle of attack increases beyond 4°, owing to the shock-on-lip condition. The results at approximately 20° indicate that the isolator becomes completely choked once the mass flow rate abruptly decreases – from around 0.30 to 0.15 Kg/s when comparing the 20°-of-AoA case with the 30° one. This work aims to contribute to the early phase of engine design by avoiding critical failures in the scramjet structure owing to aerodynamic load, thermal stress, and engine unstart.

A review of studies on air-breathing turbine engines from an energy–exergy viewpoint

A review of studies on air-breathing turbine engines from an energy–exergy viewpoint