Air-breathing Vehicle Research Articles

Numerical study on flow and combustion properties of oblique detonation engine in a wide speed range

Ensuring safe flight is a fundamental prerequisite for developing hypersonic propulsion systems. A comprehensive investigation of the steady boundary associated with oblique detonation wave in a wide speed range was conducted, with the aim of exploring the feasibility of oblique detonation engine across a diverse array of flight conditions. In this study, the wedge angle applicable in a wide-speed range was acquired via the analysis of oblique detonation wave polar curve. The configuration of the internal injection oblique detonation engine was subsequently designed and established, considering the effect of fuel-air inhomogeneity and complex wave system interactions within a confined combustor. The compressible Euler equations coupled with a 9-species and 19-step chemical reaction mechanism are employed to simulate the oblique detonation process. Ultimately, the safe flight envelope of an air-breathing vehicle equipped with the internal injection oblique detonation engine is mapped across a broad range of Mach numbers, demonstrating the engine’s capability to operate within the Mach 8 to 12 range. Furthermore, the findings reveal that decreasing either the flight Mach number or altitude results in unsteady oblique detonation wave within the internal injection oblique detonation engine combustor, however, reducing the equivalence ratio can stabilize the oblique detonation wave once again. This study provides valuable guidance for the design and wide-speed-range operation of an internal injection oblique detonation engine.

An Investigation of Internal Drag Correction Methodology for Wind Tunnel Test Data in a High-Speed Air-Breathing Vehicle Using Numerical Analysis

AbstractAccurate understanding of aerodynamic characteristics is critical for air vehicle design. This study applies the impulse-momentum theorem to estimate internal drag of an asymmetric air vehicle. Both computational fluid dynamics (CFD) and wind tunnel tests were employed. Numerical analysis using STAR-CCM + was conducted across various Mach numbers, while wind tunnel tests were conducted under only specific Mach number conditions. The measured data from wind tunnel tests, which is considered representative, were compared with CFD plane-averaged values, indicating fairly good agreement. However, discrepancies arose when comparing CFD cell-averaged values with the plane-averaged values. To address this, a correction function was developed to describe the differences. It was observed that the ratio of internal drag coefficient from the cell-averaged values to that of the plane-averaged values followed an exponential function of Mach number. This approach allows wind tunnel data to be transformed into a form akin to cell-averaged values, enhancing internal drag estimation accuracy. Future research will explore additional methodologies to further refine and validate the proposed approach.

Refined mode classification and performance analysis method of dual-mode scramjet vehicles for flight trajectory optimization

Hypersonic air-breathing vehicles, utilizing ramjet and scramjet propulsion systems, are emerging as a promising option for future hypersonic transportation due to high specific impulse. The performance of hypersonic aircraft is significantly influenced by their flight trajectory, as combustion efficiency and aerodynamic forces vary markedly with altitude. Although numerous studies have focused on optimizing these trajectories, they have not adequately considered the potential for operational failures in the propulsion system, such as unstart and blowout. This study introduces a trajectory optimization approach for hypersonic aircraft that proactively addresses and mitigates these operational failures. This is achieved by establishing an operational classification process for a dual-mode scramjet engine and proposing a failure mode avoidance strategy using a Gaussian process classifier. Optimized ascent trajectories are achieved through the development of a two-stage robust sequential convex programming approach. The results of trajectory optimization indicate that the failure mode avoidance strategy proposed is crucial in obtaining minimal time trajectories, and that the optimal trajectories include drop motion in the initial stage flight in order to increase combustion efficiency.

An experimental study on the combustion characteristics of the cavity-stabilized supersonic premixed flame

With the overwhelming trend of fast and economic aviation, it is becoming increasingly difficult for fuel to burn stably and efficiently in a constrained space and time period in the combustor of a hypersonic air-breathing vehicle. Conventional supersonic diffusion flame runs the risk of blowing out in the incoming flow at such high velocity. Upstream fuel injection of the flame holder is gaining popularity due to its prominent advantage in improving mixing efficiency, and this fuel injection method contributes to the organization of supersonic premixed combustion. In order to study the combustion characteristics of the supersonic premixed flame, corresponding experiments are carried out on a direct-connected scramjet platform with the inflow total temperature of 1700 K and the inflow Mach Number of 2.0. Gaseous ethylene with an equivalence ratio of 0.15 was injected transversely 670 mm upstream of a cavity. Optical measurements including laser-induced fluorescence on the hydroxyl radical (OH-LIF), CH*chemiluminescence photography, and high-speed photography were performed to investigate the steady-state and transient combustion characteristics of the cavity-stabilized supersonic premixed flame. From steady-state OH-LIF and CH* results, the flame could be divided into reactant zone, preheating zone, intense reacting zone, and product zone. It could be inferred that the preheating zone induced by the hot product in the cavity acts as a radical source and heat source for combustion stabilization in the cavity. In addition, the transient flame morphology analysis was carried out from the high-speed photography results. The flame base position and flame spreading angle were abstracted from the binarized image. The results showed that there existed a subtle flame flashback phenomenon for the flame base, whereas the flame spreading angle maintained a relatively constant value with a small deviation. The frequency-domain analysis revealed that both the flame base position and flame spreading angle reached the peak amplitude in the direct current component (0 Hz), exhibiting no explicit high-amplitude resonance component within 1500 Hz compared to those in the diffusion combustion mode. In all, the supersonic premixed flame presented a stabilized flame morphology.

Conceptual design methodology and performance evaluation of turbine-based combined cycle inward-turning inlet with twin-design points

Turbine-based combined cycle (TBCC) engines attract intensive attention due to adequate high working efficiency within the full speed range of air-breathing vehicles. Challenges of TBCC engines lie mostly in common-used components in particular the inlet system, which needs to satisfy the mass flow distribution of different sub-engines and to reconcile the performance requirement in the full speed range. In the present work, a new conceptual design methodology of the three-dimensional inward-turning TBCC inlet with twin-design points is proposed according to the basic flow field with double incident shock waves. The shock structures and flow field parameters at the high-speed design point Ma=4 and mass flow distribution at the low-speed design point Ma=3 both can be designed by the basic flow field. On that basis, a typical inlet model is designed and the corresponding aerodynamic performance is evaluated at two design points. Simulation results prove that the shock structure, the mass flow distribution, and the averaged throat parameters of the TBCC inlet are all consistent with the basic flow field with sufficient accuracy at twin-design points, where the maximum relative error is only 6.3%. Meanwhile, the aerodynamic performance of the inlet shows that the inlet has stable and efficient operation characteristics, which fully affirms the feasibility of the design method.

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.

Preliminary design of a hypersonic airbreathing vehicle based on four steps: An academic approach to learning

This article proposes a systematic approach for the preliminary design of a hypersonic airbreathing (scramjet) vehicle within an academic context. The goal is to serve as an educational tool, focusing on teaching students the fundamental physical concepts behind scramjets. The article provides essential equations for generic scramjet designs while explaining the simplifications made from an academic perspective. It demonstrates the use of spreadsheets as a user-friendly tool to modify parameters and observe the behavior of shock waves. The article outlines a four-step procedure. It includes estimating atmospheric properties, designing the supersonic inlet, calculating combustion chamber parameters, and determining postcombustor properties. An academic exercise illustrates these steps for a scramjet flying at a specific velocity and altitude. The results show that this procedure, combined with spreadsheets, enables a rapid prediction of thermodynamic properties, enhancing the understanding of the fundamental physics behind hypersonic airbreathing vehicles.

A novel ray method for the coking of n-decane in circular tubes with flexible cross-section area

Coking of endothermic hydrocarbon fuels significantly deteriorates the efficiency of regenerative cooling for hypersonic airbreathing vehicles. A novel ray method based on dynamic mesh techniques is proposed to describe the coking behavior in a circular tube with a flexible cross-section area. This method can be used to identify metal and deposit coke efficiently, which promotes the applicability of the dynamic mesh method in flexible structures. The cooling performance of n-decane before and after coking is numerically simulated and compared. The deposition brings higher thermal risks by elevating the solid temperature. However, the deposition reduces the channel’s cross-sectional area and increases the flow speed, leading to decreased flow residence time. Because the suppressing effect from lower flow residence time outweighs the promotion effect from increased temperature, the n-decane conversion decreased after coking.

Thermodynamic analysis of a high Mach number scramjet engine with secondary combustion for thrust enhancement

Intending to augment the thrust of air-breathing vehicles operating at high flight Mach numbers, especially under extreme flight conditions, this study introduced a pioneering methodology predicated on secondary combustion. This innovative scheme uses the product of combustion of hydrocarbon fuel and air to further react with metal fuel to obtain thrust gain. This study was mainly based on theoretical calculation. A preliminary parametric design and performance analysis of a Mach 8 scramjet engine with an afterburner was carried out based on the thermodynamic cycle analysis method, considering the gas component changes. The engine model design with an inward two-dimensional inlet was adopted, along with an oblique injection of hydrocarbon fuel and two combustors in series. In addition, the expansion of gas in a supersonic nozzle was modeled with quasi-one-dimensional flow equations. By maintaining an equivalence ratio of 1.0 for the hydrocarbon fuel in the primary combustor, meticulous exploration of the powder-to-gas ratio in the afterburner, ranging from 0.04 to 0.2, was conducted. Similarly, an investigation into the temperature ratio at the outlet of the inlet, spanning the range of 3.8 to 7, was performed. Through a judicious optimization process involving both the inlet compression temperature ratio and the magnesium powder's powder-to-gas ratio, the engine designed can obtain a maximum thrust gain of 57.3%. The analysis underscores the advantageous correlation between heightened powder-to-gas ratios and augmented thrust production. Setting the upper limit of the powder-gas ratio to 0.16 is recommended to maintain optimal performance. This research also provided a new idea for expanding the application of hydrocarbon fuels in higher flight Mach number scramjet engines.

Similarity Analysis for Two-Dimensional Hypersonic Inlets Restart

The restart ability of inlets is significant to the robustness of hypersonic airbreathing vehicles. Classic starting theories based on the one-dimensional and inviscid flow assumption cannot predict the restart boundary accurately. To overcome the classic theories, the restart characteristics of two-dimensional contraction ducts in viscous flow are numerically studied. The restart boundaries for different configurations and flow conditions are determined, and the restart similarity laws are discovered. With the proposed dimensionless quantities (the dimensionless cowl length and the effective throat-entrance area ratio), the influences of multiple factors on restart are described simultaneously in a concise self-similar formulation. The restart characteristics can be described with any two variables in cowl angle, dimensionless cowl length, and effective area ratio. The restarts of two-dimensional inlets are classified into three types of the short-cowl, transitional, and long-cowl restarts according to the features of restart boundary lines and the flow structure. The three restart types are dominated by the competition between the cowl shock system and cowl tail expansion waves, the cowl shock system alone, and the complementation between the cowl shock system and throat choke, respectively. The wall pressure criteria for restart are correlated with the dimensionless cowl length and the characteristic separation scale at the restart critical state. It is observed that Kantrowitz limit is not generalizable for restart boundaries.

Thermo-Structural Response Prediction of UHTCC Control Surface for Hypersonic Cruise Vehicle

One of the most complex issues that govern the design of a hypersonic vehicle is aerodynamic heating which is much more severe for an air-breathing vehicle as it operates within the atmosphere for longer durations. Due to the detrimental impact of aerodynamic heating on the vehicle structure, it is imperative to develop a robust thermal protection system that can protect the vehicle from high temperatures. However, the development of sophisticated TPS for slender components is impractical and, therefore, requires a hot structure system that can operate at substantial aerothermal loads. The aim of this study is to predict the impact of aero-heating on a UHTCC-based hot structure system of a hypersonic cruise vehicle through uncoupled thermo-structural analysis of the control surface. Validation of numerical methods implemented is done through literature comparison and grid independency tests.

Performance Improvement of Kerosene Fuelled Scramjet Combustor Through Modified Fuel Injection - A Cfd Study

Full scale kerosene fuelled scramjet combustor of a hypersonic air-breathing vehicle is numerically explored for two different fuel injection systems. A commercial software CFXTASCflow has been used for the reacting flow computation within the combustor. Three dimensional Navier-Stokes equations are solved along with k-ε turbulence model to analyse the flow field. Eddy Dissipation Model (EDM) with fast rate chemical kinetics is used for combustion modeling. Lagrangian Particle Tracking Method (LPTM) has been used to simulate the trajectory of the kerosene droplets. Kerosene is injected from the struts. Reacting flow simulations have been carried out for the fuel equivalence ratio of 1.0. The overall flow characteristics in the combustor are presented through the distribution of various thermochemical parameters at different cross sections. Kerosene which is injected in liquid phase completely vaporizes within the combustor. Considerable improvement in the performance in terms of thrust and combustion efficiency is observed with the modified fuel injection system.

Thermochemical non-equilibrium flow characteristics of high Mach number inlet in a wide operation range

The high-temperature non-equilibrium effect is a novel and significant issue in the flows over a high Mach number (above Mach 8) air-breathing vehicle. Thus, this study attempts to investigate the high-temperature non-equilibrium flows of a curved compression two-dimensional scramjet inlet at Mach 8 to 12 utilizing the two-dimensional non-equilibrium RANS calculations. Notably, the thermochemical non-equilibrium gas model can predict the actual high-temperature flows, and the numerical results of the other four thermochemical gas models are only used for comparative analysis. Firstly, the thermochemical non-equilibrium flow fields and work performance of the inlet at Mach 8 to 12 are analyzed. Then, the influences of high-temperature non-equilibrium effects on the starting characteristics of the inlet are investigated. The results reveal that a large separation bubble caused by the cowl shock/lower wall boundary layer interaction appears upstream of the shoulder, at Mach 8. The separation zone size is smaller, and its location is closer to the downstream area while the thermal process changes from frozen to non-equilibrium and then to equilibrium. With the increase of inflow Mach number, the thermochemical non-equilibrium effects in the whole inlet flow field gradually strengthen, so their influences on the overall work performance of the high Mach number inlet are more obvious. The vibrational relaxation or thermal non-equilibrium effects can yield more visible influences on the inlet performance than the chemical non-equilibrium reactions. The inlet in the thermochemical non-equilibrium flow can restart more easily than that in the thermochemical frozen flow. This work should provide a basis for the design and starting ability prediction of the high Mach number inlet in the wide operation range.

Dynamic Transition of Unsteady Supersonic Flow From Mach to Regular Reflection Over a Moving Wedge

Abstract The design of supersonic and hypersonic air-breathing vehicles is influenced by the transition between the Mach reflection (MR) and regular reflection (RR) phenomena. The purpose of this study is to investigate the dynamic transition of unsteady supersonic flow from MR to RR over a two-dimensional wedge numerically. The trailing edge of the wedge moves downstream along the x-direction with a velocity, V(t) at a freestream Mach number of 3. An unsteady compressible inviscid flow solver is used to simulate the phenomenon. The Arbitrary Lagrangian–Eulerian (ALE) technique is applied to deform the mesh during the wedge motion. The dynamic transition from MR to RR is defined by two criteria, the sonic and the von Neumann. Moreover, the lag in the dynamic transition from the steady-state condition is studied using various nondimensional angular velocities, κ, in the range of [0.1-2]. The lag effect in the shock system is remarkable at the high values of κ greater than 1.5. Furthermore, the dynamic transition from MR to RR happens below the dual solution domain (DSD) because the shock is upstream curved during the wedge motion.

The analysis on the flow characteristics of supersonic inlet considering structural vibration

Inlet is the key part which can influence the performance of airframe-propulsion integration for the air-breathing vehicle. In supersonic or hypersonic condition, the local vibration of the structure will change the integrated performance greatly. The work is to study the influence of the structural vibration of the compression surface on the inlet’s performance, and the vibration factors such as amplitude, wavelength and frequency are evaluated. The influence of the vibration factors on the inlet’s performance is studied. The research shows that the structure of the flow field for the inlet is changed. The influence is more obvious when the vibration occurs to be more closed to the lip. Besides, the wavelength has more influence than the amplitude, and the influence of the frequency is smallest.

Air-breathing hypersonic vehicle trajectory optimization with uncertain no-fly zones

The air-breathing hypersonic planes are considered to be the future of commercial airlines, which can fly from NYC to London in under an hour. For air-breathing vehicles, 3D trajectory planning will become extremely important due to its significant impact on flight performance. Past research on this issue was not comprehensive enough. Thus, we proposed a hybrid method for generating the optimal trajectories efficiently. No-fly zones are specified for geopolitical restrictions in air-breathing hypersonic vehicle missions. However, previous studies focused on no-fly zone constraints with fixed locations and boundaries. For robust execution, we must take into account no-fly zones’ uncertainties, which arise due to uncertain localization, measurement errors, and no-fly zones’ movement. A chance-constrained approach is presented here to deal with such uncertainties. The key idea of this method is controlling risk when the flight path is close to the uncertain no-fly zones. First, the trajectory planning of air-breathing hypersonic vehicles is modeled as a chance-constrained optimization (CCOPT) problem. With the help of the convexification and linearization techniques, we can approximate the non 3.2.2 CCOPT problem as a convex second-order cone programing under the Gaussian distribution assumption. It can be solved to global optimality by the interior point method. Finally, we give numerical simulation results to prove the effectiveness of this method.

Hierarchical variable fidelity evolutionary optimization methods and their applications in aerodynamic shape design

This paper proposes two hierarchical evolutionary optimization methods based on variable fidelity analysis and search space contraction for aerodynamic shape design, i.e. hierarchical evolutionary Pareto and Nash games. One of these techniques is used in the optimization process, namely the advantages of high and low fidelity flow simulation. The high-fidelity model provides solution accuracy while the low-fidelity model reduces the computational cost. Especially, the search space contraction and the population size reduction are introduced in the process of transition from the optimization on low fidelity simulation to the optimization on high fidelity simulation, so that the optimization based on high fidelity simulation can get the high-precision optimal solution with a relatively less Central Processing Unit(CPU) cost. They are applied to the single objective natural laminar wing shape design at transonic flow and the multidisciplinary shape optimization of a hypersonic air-breathing vehicle respectively. The optimization results show that regardless of a single objective or multi-objective/multidisciplinary optimization problem, the new hierarchical optimization methods proposed in this paper can improve the optimization efficiency by 5-10 times.

Enhancing Mechanical and Combustion Performance of Boron/Polymer Composites via Boron Particle Functionalization.

High-speed air-breathing propulsion systems, such as solid fuel ramjets (SFRJ), are important for space exploration and national security. The development of SFRJ requires high-performance solid fuels with excellent mechanical and combustion properties. One of the current solid fuel candidates is composed of high-energy particles (e.g., boron (B)) and polymeric binder (e.g., hydroxyl-terminated polybutadiene (HTPB)). However, the opposite polarities of the boron surface and HTPB lead to poor B particle dispersion and distribution within HTPB. Herein, we demonstrate that the surface functionalization of B particles with nonpolar oleoyl chloride greatly improves the dispersion and distribution of B particles within HTPB. The improved particle dispersion is quantitatively visualized through X-ray computed tomography imaging, and the particle/matrix interaction is evaluated by dynamic mechanical analysis. The surface-functionalized B particles can be uniformly dispersed up to 40 wt % in HTPB, the highest mass loading reported to date. The surface-functionalized B (40 wt %)/HTPB composite exhibits a 63.3% higher Young's modulus, 87.5% higher tensile strength, 16.2% higher toughness, and 16.8% higher heat of combustion than pristine B (40 wt %)/HTPB. The surface functionalization of B particles provides an effective strategy for improving the efficacy and safety of B/HTPB solid fuels for future high-speed air-breathing vehicles.

The Effect of Ramp Location in a Strut Based Scramjet Combustor under non reacting flow field

The scramjet engine is considered to be the high-speed propulsive system for hypersonic air-breathing vehicles. Researchers focus on various geometries of strut injectors to enhance the mixing and combustion efficiencies of the scramjet combustor. This paper reveals the non-reacting flow characteristics of various ramp locations in a scramjet combustor with strut injection. The ramps are symmetrically placed at three different locations upstream of the strut within the combustor. Air enters into the combustor inlet with Mach number 2, and the fuel is injected at the sonic velocity from the strut. The 2D supersonic flow characteristics are numerically investigated using ANSYS 18.0 software. The steady-state flow simulations are performed in this study. The flow field is modelled using RANS equations, and the k-ω SST turbulence model and default constants are chosen to investigate the flow characteristics within the combustor. The flow characteristics of the ramp based DLR scramjet combustors are compared with the baseline DLR scramjet model. The computational results are validated with the experimental data. The shock wave interaction from the ramps enhances the distribution of hydrogen in the lateral direction of the flow than the baseline strut configuration. The ramps create recirculation regions that could enhance fuel-air mixing. The ramp induced strut based DLR scramjet engine provides increased total pressure loss of the baseline DLR scramjet.

Alleviation of lateral spillage of two-dimensional hypersonic inlet using waverider-configuration chines

Hypersonic inlet is an important part of the propulsion system of hypersonic air-breathing vehicles. However, the performance of the two-dimensional hypersonic inlet, a major type of hypersonic inlets, is considerably deteriorated for lateral spillage. In this study, waverider-configuration chines mounted on the lateral sides of a two-dimensional three-staged external-compression hypersonic inlet for a Mach number of 6.0 are investigated to determine their ability to alleviate the lateral spillage. The chines are built by using a waverider design method. The numerical results suggest that a severe flow spillage induced by three-dimensional effect shows up near the lateral edge of the inlet without chines, which degrades the mass-flow ratio and flow uniformity. In contrast, the waverider-configuration chines effectively alleviate the lateral spillage. Consequently, the mass-flow ratio and the flow uniformity are both improved significantly.