Hypersonic Airbreathing Engines Research Articles

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.

Effects of Fuel Concentration Gradient on Stabilization of Oblique Detonation Waves in Kerosene–Air Mixtures

The transition structures and wave stabilization features of the wedge-induced oblique detonation wave (ODW) affect the combustion of the hypersonic air-breathing detonation engine, which is associated with the inhomogeneity of the inflow mixture. This study numerically investigates the influence of inhomogeneous kerosene–air mixtures on the stabilization of ODW for the first time, considering the inter-phase heat and mass transfers and focusing on the flow structure. The multiphase reacting flows are solved by the two-way coupling Eulerian–Lagrangian method. The inhomogeneous degree of fuel–air premixing is represented by the gradient of the liquid fuel equivalence ratio. A new pattern of transition wave structure from the shock-induced deflagration to oblique detonation is found. Under the fuel-rich condition before the shock-induced deflagration wave, a diamond-shaped wave structure is generated due to the large fuel concentration gradients. This flow structure is formed on the wedge without oscillations and is expected for a well-stabilized ODW. The initiation length of ODW is used to value the combustion performance. Its dependence on the inhomogeneous premixing degree displays a W-shaped curve. The chemical heat release influences the initiation length more obviously than the evaporative cooling in the fuel-lean conditions before the shock-induced deflagration. The ODW stabilization is enhanced by the heat released from the fuel-rich chemical reaction. Generally, the two-phase oblique detonation is determined by the competitiveness between the evaporative heat loss and chemical heat release. A uniform fuel–air mixture may not be optimal for detonation initiation based on the results of the present study.

Topology optimization of regenerative cooling channel in non-uniform thermal environment of hypersonic engine

Regenerative cooling technology is widely used in the active thermal protection system of hypersonic airbreathing engine. However, the overall efficiency of active cooling with limited total coolant quantity needs to be improved due to flow non-uniformity, hydrocarbon fuel pyrolysis and other factors. This paper presents a fluid–structure coupled topology optimization design of the regenerative cooling channel to improve heat transfer efficiency. To further verify the advantages of topology-optimized cooling channels under real operating conditions, detailed three-dimensional numerical simulations were conducted to investigate the supercritical flow and heat transfer processes of hydrocarbon fuel in the topology-optimized channels. Results reveal that compared to the traditional straight cooling channel, the topology-optimized channel has improved the heat transfer efficiency and flow distribution significantly. Many micro-channels perpendicular to the mainstream direction which can be termed as “capillary-like” zone in the optimized channel distribute the inlet fuel to the adiabatic walls, thus significantly improving the overall heat transfer. Remarkable cooling capacity robustness of the topology-optimized cooling channels has been found under different non-uniform heat flux conditions. Comparing to the straight channels, the averaged temperature of the heated wall in the optimized channels can be decreased up to 5.8% and the pressure drop can be decreased up to 20.6%. The uniformity of the fuel temperature distribution is about 2 times better than that of the straight channel. However, although overall heat transfer efficiency is high, complex flow and heat transfer phenomenon occurs in the topology-optimized cooling channel, such as the reverse flow phenomenon induced by adverse pressure gradient near the entrance and the stagnated flow in the “capillary-like” zone near the exit, which impair the local heat transfer and cause the existence of high temperature regions.

Experimental investigation of hypersonic flight-duplicated shock tunnel characteristics

Hypersonic air-breathing propulsion is one of the key techniques for future aviation and the ground aerodynamic testing for full scale test models with sufficient test time at flight conditions is of fundamental importance for verifying hypersonic air-breathing engines. Based on the backward detonation-driven concept, the hypersonic flight-duplicated shock tunnel (or JF-12 shock tunnel) has been successfully constructed and calibrated. This facility is capable of reproducing airflow for Mach numbers ranging from 5 to 9 at an altitude of 25–50 km, with a test duration of more than 100 ms. To quantify the performance of the shock tunnel, experiments were conducted to investigate the aerodynamic characteristics of the test flows and the effects of several critical techniques that play important roles in the operation of the shock tunnel. The stagnation pressure was constant within ±5 $$\%$$ and the average stagnation pressure varied by less than 0.048 $$\%/$$ ms. The variation of the stagnation pressure in repeated experiments is less than 2.0 $$\%$$ , indicating the good repeatability of the wind tunnel. The non-uniformity of the Mach number in the core flow field at the nozzle exit was within ±2.5 $$\%$$ . Additional, a uniform flow field is established upstream of the nozzle exit. The axial gradients of the flow field are small since the Mach number varies less than 1.7 $$\%/$$ m. Findings regarding the ignition technology, diaphragm ruptures, detonation driver capacity, incident shock-wave decay, and tunnel operation mode are also presented. The findings of this study are not only helpful for operating the shock tunnel, but can also assist the future development of hypersonic wind tunnels.

Development of C/PyC/SiC CMCs by CVI Process for Hypersonic Vehicle Application

Hot structures are essentially required for the hypersonic vehicle applications owing to their operation in the dense flight environments. This paper aims at the development and testing of ceramic matrix composite components for air breathing hypersonic vehicle application. In this regard, 2.5D C/PyC/SiC composites were prepared by isothermal Chemical vapour infiltration (CVI) process. X-ray CT analysis was employed to understand the extent of CVI densification in the ceramic matrix composites laminates, and it was under stood that that the intra-bundle infiltration precedes the inter-bundle infiltration with an edge (high) to center (low) gradient. The composite specimens showed an average tensile strength of 206 ± 21 MPa and an average flexural strength of 405 ± 25 MPa, respectively, at room temperature. Thermal conductivity of composite specimens shows an increase in trend with temperature. The composite laminates were successfully tested for use in the air intake ramps of hypersonic air breathing engines.

Minimization of entropy generation of a closed Brayton cycle based precooling-compression system for advanced hypersonic airbreathing engine

Fuel indirect precooled engines have the potential to radically reform the paradigm of propulsion for next generation in- and trans-atmospheric vehicles. The engine is characterized by a sophisticated thermodynamic cycle with a series of heat exchangers. Reducing irreversibility of the precooling-compression system (PCS) is the key to ensure high engine performance, for which a conjugated optimization coupled with heat exchanger design was performed to evaluate the performance level that can be achieved under the state of technologies. The results indicate that irreversibility of the PCS can be reduced but at the penalty of heavier and larger size of heat exchangers. Moreover, it shows that precooler and regenerator together contributes 75–84% of the total entropy generation rate within the PCS, while hydrogen pump occupies the second largest irreversibility source. Smaller precooling temperature can help to improve the extent of irreversibility of the PCS and increase the maximum achievable air pressure ratio and engine specific thrust, nevertheless the increased fuel consumption and air side pressure drop and the decreased engine specific impulse shouldn’t be ignored. Larger regenerator effectiveness is preferred from all aspects of the PCS and engine level figures of merit as long as the total weight and length of the heat exchangers are not increased aggressively.

Quasi-0D-Model Mapping of Reflected Shock Tunnel Operating Conditions

Reflected shock tunnels have been widely used for ground testing of hypersonic airbreathing engines. These facilities are able to replicate the total enthalpy and pressures experienced along hypersonic flight trajectories. Designing the relevant operating conditions is complex, and extensive guidance about operation condition design in reflected shock tunnels is scarce. To facilitate the operation of a reflected shock tunnel, it may be advantageous to develop a database of attainable conditions and the operation parameters required to achieve them. For this, a model to calculate global thermodynamic states and shock speeds also accounting for piston location and tube length (quasi-0D-model) is proposed. This model uses the NASA Chemical Equilibrium with Applications code to solve real gas and high-temperature effects. A number of conditions are calculated to produce maps of operating conditions. As test cases, the reflected shock tunnel under development at Tokai University, and T4 at the University of Queensland are used. Maps obtained sweeping rupture pressure or compression ratio, and initial test gas pressure for different driver compositions are presented. This methodology can be adapted to similar facilities. Moreover, observation of the presented maps provides a quick reference of the influence of each parameter when modifying existing conditions or fine-tuning tunnel operation.

Configuration optimization of the tandem cooling-compression system for a novel precooled hypersonic airbreathing engine

The tandem cooling-compression (TCC) system possesses the potential to reduce the compression work and cooling fuel consumption in fuel indirect precooled engines. Selecting of system branch number and allocating the working fluid flow are key factors to achieve the optimum system performance. On account of this, configuration optimization for the system is carried out from the perspective of compression work minimization, with the optimum system design strategies are proposed, and the necessity of using the multi-segment precooler for efficient air precooling is clarified as well. The strategies reveal that only when the branch number and the total heat capacity ratio of the system are equal, the usual uniform flow distribution (UFD) strategy is meanwhile an optimum design choice. Based on the numerical models developed, the correctness of the proposed strategies is verified by Differential-Evolution algorithm, with the potential advantages of which are evaluated also at both the component and system levels. Compared with the UFD strategy, the results indicate that generally, the compression work and cooling fuel consumption can be further reduced, and also the operational range can be remarkably extended for the system designed with the proposed optimum strategies. Moreover, assisted with the multi-segment precooling scheme, the schemes can bring about even larger extent of fuel savings and air precooling depth, higher working fluid and air side pressure ratios, and ultimately considerable gains in engine performance.

Thermodynamic analysis of the influential mechanism of fuel properties on the performance of an indirect precooled hypersonic airbreathing engine and vehicle

Fuel indirect precooled engines are promising power to realize the vision of greener and faster air transportation. Efficient operation of the engines relies critically on sophisticated energy management schemes, within which the fuel plays a decisive role. Research here intends to make clear the mechanism of how the energy conversion process of this innovative cycle is driven by properties of the fuel. A unified engine model was developed for analysis, through which it indicates that besides the properties commonly known, the triple point temperature and stoichiometric heat capacity (SHC) of the fuel are also crucial to efficient engine operation. Numerical engine method and the Concorde airplane were proposed to evaluate the synthetic influence of the multiple properties on engine and vehicle performance, with four seemingly promising fuels, i.e., the methane, n-decane, methanol and hydrogen were selected for further assessment. Experiments were performed to evaluate the contribution of chemical endothermic effect (CEE) to SHC. On account of the results, the impacts of the CEE and the quantitative significance of fuel properties on engine and vehicle performance are revealed. Also, it indicates that the system fueled with hydrogen shows superior performances in engine specific impulse and specific thrust, in spite that the corresponding flight range of the vehicle is extremely poor. On the contrary, the systems using hydrocarbon fuels (i.e., methane, n-decane or methanol) exhibit better performance in vehicle flight range, but their related engine-level performances such as the specific impulse/thrust are usually lower than that using hydrogen. Considering this complementary feature, it implies that combined use of several fuels (for instance, hydrogen + kerosene) may provide more balanced and better overall performance than that of the single fuel scheme.

Numerical analysis and design optimization of supersonic after-burning with strut fuel injectors for scramjet engines

Scramjets are a class of hypersonic airbreathing engine that offer promise for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen. An after-burning scheme can be used to effectively utilize the remaining hydrogen by supplying additional oxygen into the nozzle, aiming to augment the thrust. This paper presents the results of a single-objective design optimization for a strut fuel injection scheme considering four design variables with the objective of maximizing thrust augmentation. Thrust is found to be augmented significantly owing to a combination of contributions from aerodynamic and combustion effects. Further understanding and physical insights have been gained by performing variance-based global sensitivity analysis, scrutinizing the nozzle flowfields, analyzing the distributions and contributions of the forces acting on the nozzle wall, and examining the combustion efficiency.

Development of a Premixed Combustion Capability for Dual-Mode Scramjet Experiments

Hypersonic airbreathing engines rely on scramjet combustion processes, which involve high-speed, compressible, and highly turbulent reacting flows. The combustion environment and the turbulent flames at the heart of these engines are difficult to simulate and study in the laboratory under well-controlled conditions. Typically, wind-tunnel testing is performed that more closely approximates engine development rather than a careful investigation of the underlying physics that drives the combustion process. The experiments described in this paper, along with companion datasets, aim to isolate the chemical kinetic effects and turbulence–chemistry interaction from the fuel–air mixing process in a dual-mode scramjet combustion environment. A unique fuel-injection approach is adopted that produces a uniform fuel–air mixture at the entrance to the combustor and results in premixed combustion. This approach relies on the mixing enhancement of a precombustion shock train upstream of the dual-mode scramjet’s combustor. For the first time, a stable flame, anchored on a cavity flameholder, is reported for a scramjet combustor operating in premixed fuel–air mode. The new experimental capability has enabled numerous companion studies involving advanced diagnostics such as coherent anti-Stokes Raman scattering, particle image velocimetry, and planar laser-induced fluorescence.

Tip-to-tail numerical simulation of a hypersonic air-breathing engine with ethylene fuel

End to end CFD simulations of external and internal flow paths of an ethylene fueled hypersonic airbreathing vehicle with including forebody, horizontal fins, vertical fins, intake, combustor, single expansion ramp nozzle are carried out. The performance of the scramjet combustor and vehicle net thrust-drag is calculated for hypersonic cruise condition. Three-dimensional Navier–Stokes equations are solved along with SST-k–ω turbulence model using the commercial CFD software CFX-14. Single step chemical reaction based on fast chemistry assumption is used for combustion of gaseous ethylene fuel. Simulations captured complex shock structures including the shocks generated from the vehicle nose and compression ramps, impingement of cowl-shock on vehicle undersurface and its reflection in the intake and combustor etc. Various thermochemical parameters are analyzed and performance parameters are evaluated for nonreacting and reacting cases. Very good mixing (~98%) of fuel with incoming air stream is observed. Positive thrust–drag margins are obtained for fuel equivalence ratio of 0.6 and computed combustion efficiency is observed to be 94 %. Effect of equivalence ratio on the vehicle performance is studied parametrically. Though the combustion efficiency has come down by 8% for fuel equivalence ratio of 0.8, net vehicle thrust is increased by 44%. Heat flux distribution on the various walls of the whole vehicle including combustor is estimated for the isothermal wall condition of 1000K in reacting flow. Higher local heat flux values are observed at all the leading edges of the vehicle (i.e., nose, wing, fin and cowl leading edges) and strut regions of the combustor.

극초음속 공기흡입식 추진기관 지상 시험설비의 구성 및 특성에 관한 연구

극초음속 공기흡입식 추진기관 지상 시험설비의 구성 및 특성에 관한 연구

Mixing Characteristics of Inclined Fuel Injection via Various Geometries for Upstream-Fuel-Injected Scramjets

Efficient fuel/air mixing plays a crucial role in successful operation of hypersonic airbreathing engines, particularly scramjets, where fuel must be injected into high-speed crossflow and mixed with air at an extremely short timescale. This paper presents the results of a numerical study that investigates the effects of various orifice shapes on fuel mixing characteristics into hypersonic airflow at Mach 5, aiming at the application to scramjet operation with upstream fuel injection at Mach 10. The performance of the injectors at an inclination angle of 45 deg are assessed with respect to various criteria such as the mixing efficiency, streamwise circulation, total pressure recovery, fuel penetration, and spread. Streamwise slot injectors have been found to yield higher mixing efficiency than the other injectors tested (namely, square, circular, diamond, and triangular injectors), owing to the buffering effects. Apparent higher total pressure recovery has been obtained with these rectangular injectors, but their advantages have diminished significantly with the alignment of the trailing-edge position. The highest vertical penetration has been achieved by the square injector, whereas the performances of other injectors with a sharp leading edge have been found to be affected considerably by the axis switch phenomenon due to inclined injection at higher injection pressure. A major influence of the fore and aft shapes of the orifice has been observed on the three-dimensional bow shock formation and wake recirculation, respectively.

Validation of Three-Dimensional Simulation of Flow through Hypersonic Air-breathing Engine

<p>The flow path of a conceptual hypersonic air-breathing scramjet engine integrated with the vehicle (without combustion) has been simulated numerically using ANSYS CFX software with the SST turbulence model. The computations were performed for the free stream Mach number of 6 and angle-of-attack of 5°. A strong separation bubble was observed on the bodyside wall in the internal compression region where the reflected cowl shock impinges on body which in turn increases the static pressure substantially. The external-internal flow field of the hypersonic mixed compression intake, shock-boundary layer interactions, and the shock-shock interactions present in the internal compression region have qualitatively been obtained and analysed. The variation of centreline pressure along the bodyside wall close to the symmetry plane obtained from numerical simulation centreline has been compared with the experimentally measured data. It has been observed that the computed wall pressure matches fairly well with the measured values in the external ramp compression region, internal compression region and in the combustion chamber. The flow patterns and the pressure variations near the middle wall and the fuel injecting strut locations have also been analysed.</p><p><strong>Defence Science Journal, Vol. 65, No. 4, July 2015, pp. 272-278, DOI: http://dx.doi.org/10.14429/dsj.65.6979</strong></p>

Thrust augmentation optimization through supersonic after-burning in scramjet engine nozzles via surrogate-assisted evolutionary algorithms

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen which under ideal conditions, can be utilized to introduce an after-burning scheme. After-burning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a single-objective design optimization considering three design variables with the objective of producing maximum thrust augmentation. It is found that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of variance-based global sensitivity analysis, force contribution breakdowns, analysis of the nozzle flowfields, analysis of the surface pressure and shear stress distributions acting on the nozzle wall and analysis of the combustion efficiency.

Thrust augmentation through after-burning in scramjet nozzles

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable access-to-space and high-speed atmospheric transport. After-burning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a numerical study of three parameters and the effect that they have on thrust augmentation. These parameters include the injection pressure, injection angle and streamwise injection position. It is shown that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of a force contribution breakdown, analysis of the nozzle flowfields and finally the analysis of the surface pressure and shear stress distributions acting upon the nozzle wall.

画像解析法による極低温二相流のボイド率測定に関する研究

The hypersonic air-breathing engine, which is currently under development by Japan Aerospace Exploration Agency (JAXA), uses liquid hydrogen as the fuel. In order to accurately control the fuel flow rate during the start-up, it is essential to measure the heat transfer and pressure drop of the two phase flow. These two characteristics depend on void fraction, flow velocity and flow regime; thus, measurement methods for these values are required to be established. In this study, the void fraction measurement method by the high-speed image analyses has been developed. Two images taken from top and side directions by high-speed cameras of 1000 fps are used for “two-direction semi-automatic analysis”. We also develop “One-direction full automatic analysis” which uses only the side view as a full-automatic and high sampling rate method with little accuracy deterioration. The preliminary verification test using vertical pipe and acrylic ball shows favorable results within 2.2€% error against the theoretical value. A cryogenic experiment using two-phase nitrogen flow was also conducted. Sampling rate of “One-direction full automatic analysis” can be up to 1000 Hz. The difference between the results of two methods was as minor as 5€% when the void fraction was below 30€%.

A computational study of the HyShot II combustor performance

Experimental and flight data for hypersonic air-breathing engines are both difficult and extremely expensive to obtain, motivating the use of computational models to enhance the understanding of the complex physics involved. Here, a comprehensive numerical study has been carried out for the HyShot II scramjet combustor. This study makes use of Reynolds Average Navier Stokes (RANS) based models and Large Eddy Simulation (LES) based models with semi-detailed reaction kinetics. In this investigation we focus on the underlying flow-mixing-combustion physics at different operating conditions tested in the High Enthalpy Shock Tunnel Göttingen (HEG). To account for the complex flow in the HEG facility a zonal approach is employed in which RANS is used to simulate the flow in the HEG nozzle and test-section, providing the necessary inflow boundary conditions for the combustor RANS and LES, being the focus of this analysis. Specifically, we here combine results from RANS and LES computations with data from the HEG experiments and the target HyShot II flight-tests at two different flight-altitudes (28 and 33km). The LES model is observed to capture the experimental wall-pressure and heat-flux data very well for both the 33 and 28km altitude cases, whereas the RANS model is only able to predict the wall-pressure and heat flux data for the 28km altitude case. Based on the LES results, the flow at both altitudes is found to be unsteady, but with unsteady transitional flow features dominating the 33km case. Moreover, these results show that the equivalence ratio is of key importance to the resulting flow, mixing and combustion physics, with richer mixtures being prone to transitional flow features. The LES results are also used to describe the flow physics in detail for both altitudes, and the key processes responsible for the transition between the two combustion modes observed.

How Much Compression Should a Scramjet Inlet Do?

The supersonic combustion ramjet, or scramjet, is the engine cycle most suitable for sustained hypersonic flight in the atmosphere. This paper examines a key decision in the design of the inlet or intake of these hypersonic airbreathing engines, namely, the level of compression to be performed. Too much compression can lead to onerous system level issues including the need for bleed or variable geometry, while too little compression can result in low cycle efficiency and poor combustion of fuel. An analysis of the important factors that affect the choice of scramjet inlet compression ratio has been performed for hydrogen-fueled scramjets at Mach 6, 8, 10, and 12. It was found that contrary to classical thermodynamic analyses, scramjet cycle efficiency reaches an optimum at a relatively low compression ratio between 50 and 100 for all Mach numbers. Practical constraints related to nonequilibrium flow effects, inlet starting, and boundary-layer separation were also shown to prompt a desire for low compression ratio. The lower limit on compression was found to be set by the need to complete the combustion reaction in the available engine length and is therefore dependent on engine scale. On the basis of these factors it is recommended that scramjet inlet compression ratio be set to the minimum that satisfies the robust combustion requirement, with the caveat that it not be below 50 in order to maintain high cycle efficiency. For typical wind-tunnel-scale engines, this results in a requirement for the inlet to compress airflow entering the combustor to a pressure of approximately 1/2 atm, regardless of the flight Mach number.