Hypersonic Propulsion Research Articles

Effect of geometry parameters on the hydrocarbon fuel flow rate distribution in pyrolysis zone of SCRamjet cooling channels

Hypersonic air-breathing propulsion (>Mach 5) based on SCRamjet promotes the revolutions of both military and civilian applications. However, the engine thermal protection is seriously challenged due to high Mach number. One of the key problems is mal-distribution of hydrocarbon fuel in regenerative cooling channels under the complex thermal boundaries and channel geometry, which may cause structural failure. To improve the cooling channel design and achieve better fuel flow distribution, the effect of channel geometry parameters on flow distribution in pyrolysis zone is numerically studied. Both geometry induced and heat flux induced flow mal-distribution are concerned. The study indicates that the flow and heat transfer features in pyrolysis zone of parallel channels differs with those in non-pyrolysis zone. The flow distribution deviations in pyrolysis zone are lower. When the channel aspect ratio varies, thermal stratification dominates flow distribution more than the variation of heat flux on heated surface in pyrolysis zone, which is opposite of that in non-pyrolysis zone. Besides, thinner rib and smaller total flow area achieve better cooling performance while consuming less fuel chemical heat sink. The results help to conclude possible references to the design of cooling channels with fuel pyrolysis considered.

Study on the effects of geometry on the initiation characteristics of the oblique detonation wave for hydrogen-air mixture

The Oblique Detonation Wave Engine (ODWE) may act as a hypersonic propulsion system operating at high Mach numbers, which is an important member in the family of Scramjet. Hydrogen is a promising fuel for Scramjet, which provides wider Mach number range and is environmentally friendly. The geometry of the engine greatly affects the performance of the ODWE using hydrogen fuel. This investigation focuses on a novel wedge proposed recently, which may be utilized in scramjet engines. The wedge consists of two sub-wedges and a step. This research focuses on how the geometry of the wedge affects the initiation characteristics of the oblique detonation. Simulations are conducted on basis of Euler equations and a 9-species and 19-reactions mechanism. It is found that a larger leading wedge angle leads to a shorter initiation length. A larger step angle induces a longer initiation length. Few effects are observed on the initiation characteristics for the current range of depth. The streamline surface at the rear of the step weakens the rear shock wave and induces a longer initiation length. The streamline surface at the tip of the step begins to take effect when the initiation position is away from the step. This research provides basis for understanding the performance of the oblique detonation wave under different geometries and provides theoretical basis for scramjet engine design.

Characteristics of hydrogen jet combustion in a high-enthalpy supersonic crossflow

The facilitation of a stable combustion process is of utmost importance for the realizability and performance of hypersonic propulsion systems. To elucidate the turbulent combustion characteristics, wall-modeled large eddy simulations of a transverse jet injection into a heated supersonic flow are conducted employing a detailed reaction mechanism. The computation framework utilizes an adaptive central-upwind weighted essentially nonoscillatory (WENO-CU) scheme to achieve the sixth-order accuracy in smooth flowfields, while keeping a good shock-capturing ability. The reacting zones agree well with experimental measurements in terms of the instantaneous distribution of OH radicals. And the flame penetration height has been predicted with an error of less than 17%. It is found that the turbulent reacting flow is dominated by nonpremixed combustion mainly taking place in the near-wall region and jet windward shear-layer. Moreover, the autoignition process, which plays a critical role in stabilizing supersonic combustion, shows to favor a fuel-lean or not very fuel-rich environment of a high enthalpy. Local scalar dissipation induced by turbulence gives rise to a rapid fuel mixing with the surrounding air. However, this effect may also lead to the decrease in local temperature.

Active control method for restart performances of hypersonic inlets based on energy addition

The restart performance of the inlet is crucial to a hypersonic air-breathing propulsion vehicle. An active flow control method based on energy addition for the restart of hypersonic inlets is brought forward in this paper. Using Reynolds Averaged Navier–Stokes equations method, the effects of energy addition parameters on the restart performances of hypersonic inlets are investigated and the restart process is presented. This study verifies that the present method is feasible to improve the restarting capability of hypersonic inlets. The restart performances rely heavily on the energy addition parameters. Using energy addition properly, it usually yields a better inlet performance and more importantly enables the inlet restart again; otherwise, the large separation bubble still exists. Investigations on energy addition parameters further reveal that energy addition always increases the mass flow rate, and its center far from the cowl facilitates restarting the inlet and decreasing the stagnation pressure losses. It also appears advantageous to work with energy addition at a moderate power consumption and effective radius. Besides, the restart process indicates that the parabolic shock induced by heated region plays an important role on the restart performance. It alters the separation bubble, broadens the flow passage and enables more deflected air into the inlet. The adverse pressure gradients in the entrance are greatly changed as well. Finally, results also suggest that an appropriate effective time for energy addition is sufficient to restart the inlet. The variations of hysteresis loops show that energy addition prompts the separated bubble to disappear and the inlet to restart in advance. As a result, the restarting Mach number declines and the operation range of hypersonic inlets prominently enlarges.

Effects of magnetohydrodynamics on temperature and shock standoff distance in a supersonic flow over a blunt body

Magnetohydrodynamics has been one of the promising ways of increasing efficiency of supersonic as well as hypersonic propulsion systems, mostly by applying magnetic fields over shockwaves. This paper presents the effects of magnetic fields with different strengths both along and transverse to an electrically conducting flow at Mach 2.94 over a blunt body using a density-based solver implemented with MHD equations in OpenFOAM. The solver utilises Riemann method with an AUSM+ flux splitting technique along with limited linear interpolation. The effects of imposed magnetic field on temperature and shock standoff distance are observed. Movement of the shockwave due to the application of magnetic field on the geometry, along with a consequent change in shock standoff distance is presented in the paper. The influence of a magnetic field's direction on the Mach number of the flow is also shown. Likewise, the stagnation temperature of the blunt body is demonstrated to be independent to the direction of applied magnetic field.

Supercritical-Pressure Heat Transfer, Pyrolytic Reactions, and Surface Coking of n-Decane in Helical Tubes

Supercritical-pressure heat transfer of a hydrocarbon fuel has practical importance in fuel precooling technology in hypersonic air-breathing propulsion systems. Heat transfer of n-decane in helica...

Research on control-oriented coupling modeling for air-breathing hypersonic propulsion systems

A control-oriented coupling hypersonic propulsion system model is proposed to enable the rapid development of a propulsion model for air-breathing hypersonic flight vehicles (ABHV) in early stages of the design process and facilitate design control and analysis. An airframe/propulsion coupling hypersonic inlet model was established based on oblique shock wave theory. An isolator model in which the effects of the back pressure of the combustion chamber were adjusted by the static pressure ratio of the isolator was established. A hypersonic combustion model was also established, taking into account the fuel flow, cross-sectional area, wall friction, combustion efficiency, and exothermic reactions based on quasi-1D flow theories. Nozzle/afterbody modeling was established based on identification of the free boundary (i.e., the location of the shear layer) by Newton collision theory, and flow parameters were determined according to the influence coefficient method. The mass flow rates of air in the design state and two typical non-design states were determined geometrically based on the application of oblique shock wave theory. A propulsion coupling model that reflects the coupling of propulsion system and aerodynamics, as well as the physical mechanisms of the propulsion mechanism, was then established based on air flow rates obtained and the momentum theorem. Simulation results of airframe/propulsion integrated module air-breathing hypersonic flight vehicles (ABHVs) by the proposed model were compared to results achieved by numerical 3D Computational Fluid Dynamics (CFD) models. Results indicated that the efficacy and accuracy of the proposed models met the established requirements of control-oriented modeling, thus facilitating dynamic modeling and control in the early stage of the design process.

Synthesis of aluminum nanoparticles as additive to enhance ignition and combustion of high energy density fuels

High energy density fuels are critical for hypersonic aerospace propulsion but suffer from difficulties of ignition delay and incomplete combustion. This research reports aluminum nanoparticles (Al NPs) assisted ignition and combustion of high energy density JP-10 fuel. Al NPs with a size of 16 nm were fabricated through a mild and simple method by decomposing AlH3·Et2O with the addition of a surfactant ligand. The uniform size distribution, nanoscaled size and surface ligand make Al NPs stably suspend in JP-10, with 80% NPs being dispersed in the liquid fuel after six months. A shock tube test shows that the presence of 1 wt-% Al NPs can significantly shorten ignition delay time at temperature of 1500 to 1750 K, promote the combustion, and enhance energy release of JP-10. This work demonstrates the potential of Al NPs as ignition and combustion additive for high energy density fuel in hypersonic applications.

Uncertainty and sensitivity analysis of flow parameters on aerodynamics of a hypersonic inlet

The performance of the inlet is crucial to the cruise flight of a hypersonic air-breathing propulsion vehicle. The objective of this work is to investigate the uncertainty and sensitivity of pressure field and the performance parameters for a hypersonic inlet due to the uncertainty of five flow parameters, including freestream Mach number, Reynolds number, angle of attack, temperature and wall temperature. The steady Reynolds Averaged Navier-Stokes equations are solved to predict the inlet start and unstart flows within the hysteresis loop. Then, a point-collocation non-intrusive polynomial chaos method (NIPC) is utilized to quantify the uncertainty and sensitivity in the output quantities of interest. The uncertainty analysis in pressure field shows that Mach number and angle of attack of freestream make dominant contributions to the total uncertainty, and the Mach number has remarkable impacts in the isolator. In the start flow, the angle of attack exerts its prominent influence in the post-shock regions, while Mach number mainly dominates these regions ahead of and around the shocks. The reason may be interpreted as the much greater pressure derivatives with respect to angle of attack in the post-shock regions. Significant discrepancies are presented for the unstart flow. The reflected shock waves in the unstart flow are less sensitive to the variations of flow parameters. The external flow field, separation bubble and reflected shocks are significantly affected by angle of attack. Besides, the uncertainties of the performance parameters in the start flow are about twice those in the unstart flow. The sensitivity analysis further reveals that Mach number is the major contributor to the total uncertainty of performance parameters. The correlation coefficients via linear regression method clearly illustrate the relationships between the five input parameters and the performance parameters.

Parametric study on the hydrocarbon fuel flow rate distribution and cooling effect in non-uniformly heated parallel cooling channels

Advanced aero-engines with higher flying Mach number, like SCRamjets, offer good solution to the hypersonic propulsion. However, the thermal boundary condition of the engine cooling channels is normally non-uniform, which may cause fuel mal-distribution. Serious fuel cooling capacity waste and even over-temperature may occur. To ensure the reasonable flow distribution and wall temperature, a parametric study of the cooling channel geometries is conducted under non-uniform thermal boundary condition. A 3D numerical model considering the flow and heat transfer under supercritical pressure in parallel cooling channels is developed and validated. The results indicate that the flow distribution and cooling effect both improve with the reasonable design of the channel geometry. Higher aspect ratio (AR) improves the flow distribution and cooling effect. The variations of rib thickness and total channel flow area show greater impact on the heat transfer performance than on the flow distribution. At last, the results of channel geometry optimization are validated under different heat flux distributions. It can be concluded that the channel geometry optimization in this work improves the flow distribution and lowers the wall temperature, which provides reference for the design of cooling channels in aero-engines like SCRamjets.

Transition Between Different Initiation Structures of Wedge-Induced Oblique Detonations

Oblique detonation waves (ODWs) have been widely studied due to their application potential for airbreathing hypersonic propulsion. Moreover, various formation structures of wedge-induced oblique detonation waves have been revealed in recent numerical investigations. Given the inflow conditions, the wave configuration is dependent on the wedge angle. Hence, any wedge-angle change will induce a transient ODW evolution to transition from one configuration to another. In this study, the transient development created by instantaneously changing the wedge angle is investigated numerically, based on the unsteady two-dimensional Euler equations and one-step irreversible Arrhenius chemical kinetics. The evolution caused by the abrupt wedge-angle change from one smooth initiation structure to another, both with a curved oblique shock/detonation surface at high-Mach-number regime, is investigated. Two processes are analyzed; the first consists of the downstream transition of the ODW initiation region the by decreasing the angle, and the second is the upstream transition by increasing the angle. In the downstream transition, the overall structure moves globally and readjusts continuously, generating an intermediate kinklike initiation structure. In the upstream transition, a localized reaction region forms and induces a more complex process, mainly derived from the different responding speeds of the oblique shock and detonation waves. To avoid the generation of the new localized explosion region, which causes an abrupt change in the initiation position and potentially affects the ODWE’s stability and performance, it is suggested to vary the wedge angle in incremental steps within a certain time interval.

Mathematical modeling and characteristic analysis for over-under turbine based combined cycle engine

The turbine based combined cycle engine has become the most promising hypersonic airbreathing propulsion system for its superiority of ground self-starting, wide flight envelop and reusability. The simulation model of the turbine based combined cycle engine plays an important role in the research of performance analysis and control system design. In this paper, a turbine based combined cycle engine mathematical model is built on the Simulink platform, including a dual-channel air intake system, a turbojet engine and a ramjet. It should be noted that the model of the air intake system is built based on computational fluid dynamics calculation, which provides valuable raw data for modeling of the turbine based combined cycle engine. The aerodynamic characteristics of turbine based combined cycle engine in turbojet mode, ramjet mode and mode transition process are studied by the mathematical model, and the influence of dominant variables on performance and safety of the turbine based combined cycle engine is analyzed. According to the stability requirement of thrust output and the safety in the working process of turbine based combined cycle engine, a control law is proposed that could guarantee the steady output of thrust by controlling the control variables of the turbine based combined cycle engine in the whole working process.

Experimental and theoretical investigation of power generation scheme driven by thermal cracked gaseous hydrocarbon fuel for hypersonic vehicle

Hypersonic vehicle has become one of the research focuses because of its broad application prospects. The power supply issue in the long range/endurance of hypersonic propulsion system turns out to be one of the most important targets. Because the conventional power supply devices cannot meet the demands of a long range hypersonic air-breathing flight. In this work, a new scheme based on the expansion of high temperature cracked hydrocarbon fuel is proposed to solve the power generation problem. An analytical model for the expansion characteristics of the high temperature cracked hydrocarbon fuel is developed and validated by experiment data. The experimental results show that positive output work can be obtained with fuel working temperature even as low as 800 K and increases with fuel temperature. It also shows that the expansion work doing capability of cracked hydrocarbon fuel is much larger than the un-cracked fuel and greatly influenced by the fuel temperature, especially at the cracking temperature range. Especially, both experimental and calculation results show that the pressure also affects the high temperature thermal cracked hydrocarbon fuel expansion capability through influencing pyrolysis pathway of fuel and corresponding cracked components. The sensitivity analysis shows that the expansion work doing capability of fuel is more sensitive to the fuel temperature, especially at the hydrocarbon fuel pyrolysis starting and rapid increasing temperature region which is from 785 K to 855 K.

Supersonic Combustion in Air-Breathing Propulsion Systems for Hypersonic Flight

Great efforts have been dedicated during the last decades to the research and development of hypersonic aircrafts that can fly at several times the speed of sound. These aerospace vehicles have revolutionary applications in national security as advanced hypersonic weapons, in space exploration as reusable stages for access to low Earth orbit, and in commercial aviation as fast long-range methods for air transportation of passengers around the globe. This review addresses the topic of supersonic combustion, which represents the central physical process that enables scramjet hypersonic propulsion systems to accelerate aircrafts to ultra-high speeds. The description focuses on recent experimental flights and ground-based research programs and highlights associated fundamental flow physics, subgrid-scale model development, and full-system numerical simulations.

Mixing enhancement and penetration improvement induced by pulsed gaseous jet and a vortex generator in supersonic flows

A good injection strategy with high penetration and mixing efficiency relates to the overall performance of the airbreathing hypersonic propulsion system. In the current study, the transverse injection flow field with a vortex generator placed in front of jet has been investigated, and the jet mixing enhancement and penetration improvement induced by the pulsed jet has been evaluated as well. The obtained results predicted by the three-dimensional Reynolds-average Navier – Stokes (RANS) equations coupled with the two equation k-ω shear stress transport (SST) turbulence model show that the jet-to-crossflow pressure ratio has an great impact on penetration depth and mixing efficiency in the steady jet flow field. In the case of lower jet-to-crossflow pressure ratio, higher mixing efficiency but lower penetration depth are shown. On the other hand, compared with the steady jet flow field, the pulsed jet flow field with the periodic variation of jet-to-crossflow pressure ratio of the fuel injection has better comprehensive performance, and the penetration depth and mixing efficiency are improved simultaneously.

Experimental and Numerical Investigation of a Fluidically Variable Hypersonic Inlet

A fluidically variable hypersonic inlet with fixed geometry is investigated in this paper by both experimental and numerical methods. The results show that the inlet maintains the shock-on-lip condition both at Mach 4.92 and 5.74 successfully. At Mach 4.92, the shock-on-lip condition is naturally obtained without any need of secondary flow injection. When the inlet operates at Mach 5.74, the fluidic control technique works to prevent the external ramp shocks from entering the duct, and the secondary flow ratio used to control the first and second ramp shocks is 1.03 and 0.67%, respectively. Compared with the fixed-geometry inlet under the same design constraints, the mass flow ratio and the total pressure recovery of the fluidically controlled inlet are improved by 23.8 and 3.95% at Mach 4.92, respectively. While operating at Mach 5.74, the total pressure recovery of the fluidically controlled inlet is slightly lower than the fixed-geometry one because the curved ramp shocks lead to additional total pressure loss. As a result, the fluidically controlled inlet has excellent performance at low operating Mach numbers, which is beneficial to the performance of hypersonic propulsion system at the acceleration stage.

Mixing augmentation induced by a vortex generator located upstream of the transverse gaseous jet in supersonic flows

The mixing process plays a very important role in the engineering realization of the scramjet engine, and sufficient mixing between the incoming supersonic air and the fuel relates to improve the overall performance of the airbreathing hypersonic propulsion system. In the current study, the delta wing is placed in front of the injector to promote the mixing of fuel and supersonic crossflow, and the effects of the delta wing height and the jet-to-crossflow pressure ratio have been investigated numerically based on grid independency analysis and code validation. The obtained results predicted by the three-dimensional Reynolds-average Navier–Stokes (RANS) equations coupled with the two equation k–ω shear stress transport (SST) turbulence model show that the delta wing has a highly remarkable improvement on mixing characteristics such as mixing efficiency and fuel penetration depth. However, the delta wing also shows additional losses of stagnation pressure. In the case of higher values of delta wing height and jet-to-crossflow pressure ratio, higher penetration and more losses of stagnation pressure are shown. At the same time, the mixing efficiency decreases with the increase of the jet-to-crossflow pressure ratio irrespective of the height of the delta wing, and there is an optimum height of the delta wing for each jet-to-crossflow pressure ratio to achieve the maximization of rapid fuel–air mixing. In addition, the hydrogen content in the recirculation region between the orifice and the delta wing is a result of both the jet-to-crossflow pressure ratio and the height of the delta wing. In conclusion, the design of the flow field with the delta wing to realize objective comprehensive optimal performance is a multi-objective problem, and it should be solved by the multi-objective design optimization approach.

Experimental Investigation of Brazilian 14-X B Hypersonic Scramjet Aerospace Vehicle

The Brazilian hypersonic scramjet aerospace vehicle 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) to be tested in flight into the Earth’s atmosphere at an altitude of 30 km and Mach number 7. The 14-X B has been designed at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Institute for Advanced Studies (IEAv), Brazil. The IEAv T3 Hypersonic Shock Tunnel is a ground-test facility able to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of the 14-X B into the Earth’s atmosphere at hypersonic flight speeds. A 1 m long stainless steel 14-X B model was experimentally investigated at T3 Hypersonic Shock Tunnel, for freestream Mach numbers ranging from 7 to 8. Static pressure measurements along the lower surface of the 14-X B, as well as high-speed Schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp, provided experimental data. Experimental data was compared to the analytical theoretical solutions and the computational fluid dynamics (CFD) simulations, showing good qualitative agreement and in consequence demonstrating the importance of these methods in the project of the 14-X B hypersonic scramjet aerospace vehicle.

Numerical study of nonequilibrium plasma assisted detonation initiation in detonation tube

Nonequilibrium plasma has shown great merits in ignition and combustion nowadays, which should be especially useful for hypersonic propulsion. A coaxial electrodes configuration was established to investigate the effect of alternating current (AC) dielectric barrier discharge nonequilibrium plasma on the detonation initiation process in a hydrogen-oxygen mixture. A discharge simulation-combustion simulation loosely coupled method was used to simulate plasma assisted detonation initiation. First, the dielectric barrier discharge in the hydrogen-oxygen mixture driven by an AC voltage was simulated, which takes 17 kinds of particles (including positively charged particles, negatively charged particles, and neutral particles) and 47 reactions into account. The temporal and spatial characteristics of the discharge products were obtained. Then, the discharge products were incorporated into the combustion model of a detonation combustor as the initial conditions for the later detonation initiation simulation. Results showed that the number density distributions of plasma species are different in space and time, and develop highly nonuniformly from high voltage electrode to grounded electrode at certain times. All the active species reach their highest concentration at approximately 0.6T (T denotes a discharge cycle). Compared with the no plasma case, the differences of flowfield shape mainly appear in the early stage of the deflagration to detonation transition process. None of the sub-processes (including the very slow combustion, deflagration, over-driven detonation, detonation decay, and propagation of a self-sustained stable detonation wave) have been removed by the plasma. After the formation of a C-J detonation wave, the whole flowfield remains unchanged. With the help of plasma, the deflagration to detonation transition (DDT) time and distance are reduced by about 11.6% and 12.9%, respectively, which should be attributed to the active particles effect of nonequilibrium plasma and the local turbulent enhancing effect by the spatial characteristics of discharge. In addition, as the duration of forming a shock wave in the combustor is shortened by approximately 8.1%, it can be inferred that the plasma accelerates the DDT process more significantly before the flow becomes supersonic.

고온 공기 가열기 개발 현황 조사 및 고찰

극초음속 공기 흡입 추진기관의 지상 시험 시 추진기관의 고속 비행 조건을 모의하기 위해 고온 공기를 공급할 수 있는 고온 공기 가열기가 필요하다. 본 논문에서는 다양한 고온 가열기들을 조사하여 유형별로 정리하고, 장단점을 비교하였다. 가열기는 크게 유동장 내 연소 가열기, 아크 가열기, 축열식 가열기, 열교환식 가열기 4종류로 분류할 수 있었으며, 각각의 장단점이 다양하므로 목적에 맞게 가열기를 선정하여야 한다. A high temperature heater is required to supply high temperature air to the hypersonic propulsion system in order to simulate high velocity flying condition during the ground test. Various high temperature heaters were reviewed, categorized, and analyzed in this paper. Heaters were categorized in 4 groups; in-stream combustion heater, electric arc heater, storage heater, and heat exchange heater. Each group has its own advantages and disadvantages, so the heater should be selected considering its purpose.