Ramjet Engine Research Articles

Computational study of hydrogen and air co-flow jets mixing in dual-combustor ramjet in presence of ramp shock generator

Flame stabilization distribution and efficient fuel mixing are highly significant in the performance of ramjet engine. In this paper, numerical studies are done to investigate the supersonic combustion of hydrogen and air co-flow in presence of the ramp edge. In this study, flame maintenance and combustion formation in a dual-combustion ramjet engine are fully investigated. The primary emphasis of this work is to disclose the impact of produced shocks due to the existence of ramp edge on the structure of fuel and air-jet immediately downstream of injector. The shock interactions of two different angles (10 deg and 30 deg) of the ramp edge are investigated. Favre-averaged conservation equations are considered for the simulation of compressible flows inside the combustor. Various mechanisms expressing the flame spreading characteristics are also examined. Depending on the angle of the ramp edge, the produced shock waves influence on size and strength of vortices. Our findings also show that high angles of ramp edge augment the fluctuation of the vortices by the splitter and consequently, fuel mixing enhances in the combustion chamber of ramjet engine. Our findings also confirm that the similar jet pressure would be more effective on the vortex production downstream of the splitter. The application of ramp edge with angle of 30 deg would increase average momentum thickness of the mixing layer up to 55%.

Evaluation of novel wall function approach for supersonic flow problems involving separations induced by geometry and shocks

Supersonic combustion ramjet engine involves complex flow physics including shock-turbulent boundary layer interaction (STBLI) and shock-induced separations. The general approach while using an averaged Navier-Stokes equations to capture these flow separation, due to the flow physics and/or geometry accurately, is to use a low-Reynolds turbulence model. This requires the placement of first grid point in the viscous sublayer ( y + ≃ 1 ), which implies a need of finer mesh near the wall leading to an overall increase in number of cells in the computational domain and thus an increase in overall computational time. The wall-function method is a substitute to this approach which allows for the use of a coarser mesh in the region near to the wall. However, the standard wall-function methods are generally not very successful in capturing flow separations. In this work, we have employed a generalized wall function approach which incorporates the effect of both pressure gradient and compressibility in its formulation and have evaluated its performance for complex supersonic flow cases involving shock-turbulence boundary layer interactions (STBLI) and shock induced separations, which have not been done previously. We have shown that this method captures the flow physics accurately even with much coarser meshes near the wall. Further, insensitivity of the method to the placement of first grid point has also been demonstrated.

무격자 기법을 이용한 초음속 예혼합 조건에 대한 화학 반응 수치 해석

The scramjet is the supersonic-combustion ramjet engine operating in the hypersonic regime. Recently, importance of designing regenerative cooling system is highlighted for enhancing combustion performance and overcoming severe thermal environment. To optimize a design of cooling channel, investigation of thermal aerodynamics, thermal structure, cooling channels, combustion heating and heat transfer should be comprehensively performed. Although finite volume method(FVM) is widely used in numerical simulation, the meshless method is efficient for complex geometries. In this study, supersonic premixed combustion is numerically investigated using a meshless method as a fundamental research about combustion analysis based on meshless method. Numerical results are compared with the hydrogen-oxygen combustion results provided by the NPARC alliance validation archive.

Influence of isolator section on the shock augmented mixing in SCRAMJET engine

The paper aims to numerically study the effect of isolator design on shock-augmented mixing in a generic supersonic Ramjet (SCRAMJET) engine and reduce what would be a very complicated thermodynamics+aerodynamic optimization in the design to a simpler aerodynamic one. Changing the length of the isolator section effectively varies the thickness of the boundary layer that is formed in the isolator, which then enters into the SCRAMJET combustor. The study is conducted on a three-dimensional geometry using the Menter's SST model for turbulence on an in-house unstructured grid RANS solver. In this paper, the authors quantify the effect of the incoming boundary layer thickness on the shock-augment air-fuel mixing in the SCRAMJET combustor. It is observed that for the maximum variation in the isolator length, the normalized mixing volume only improves by 7%, which is only a marginally better mixing of air and fuel obtained as with an increase in the isolator length. The authors conclude that the Prandtl-Meyer expansion fan (PMEF) formed at the flame-holder effectively isolates the combustor from the isolator section to a degree that the isolator dimensions can therefore be chosen entirely on considerations of flow homogenization of inlet air before it enters the combustor section of the SCRAMJET engine. Although, the paper analyses a generic SCRAMJET with specific dimensions, the results are more generally applicable than for the specific design actually considered in the paper. Thus this study is also relevant to SCRAMJET engines that may differ radically from our considered design in form and dimensions.

Численное моделирование процессов дросселирования прямоточного воздушно-реактивного двигателя

Aimed at determining the maximum allowable pressure in the combustion chamber causing no flow separation in the air intake device, numerical simulation of the throttling process in a ramjet engine was carried out. Calculations were made in two formulations: with local and global time steps. When comparing results obtained using these approaches, the possibility was found to determine the air intake device throttling characteristics without losing the accuracy at the initial stages of designing the ramjet engines with introduction of the local time step. The non-stationary model created the basis for selecting the optimal integration step to minimize computational resources. The flow pulsations period was evaluated for the event of a surge. These pulsations frequencies were registered in the extremely low frequency range.

Low toxic nitromethane based monopropellant for gas generator cycle air turbo ramjet engine

The present study proposes ethanol-nitromethane mixture as a safe, storable, and low toxic monopropellant for a gas generator cycle air turbo ramjet engine and evaluates its propulsion performance. The present study proposes that nitromethane is mixed with ethanol to adjust gas generator combustion temperature. The author developed the computational code for the present analysis and verified its accuracy. The specific thrust of ethanol-nitromethane is nearly identical to that of ethanol-liquified oxygen and is higher than methanol-hydroxylammonium nitrate aqueous solution and hydrazine. Hydrazine has the highest Isp among the propellants in the present analysis. However, Isp of ethanol-nitromethane is nearly equal to that of ethanol-liquified oxygen and higher than those of methanol-hydroxylammonium nitrate aqueous solution, and hydroxyl terminated polybutadiene-ammonium perchlorate. With the propellant-to-air ratio range in the present study, ethanol-nitromethane has the stoichiometric condition at the ram combustor around the propellant-to-air ratio range from 0.2 to 0.25 and can obtain high ram combustion temperature. This result is favorable for ethanol-nitromethane because it improves the specific thrust and the specific impulse. Therefore, ethanol-nitromethane can be a promising low-toxic liquid monopropellant for the air turbo ramjet engine.

A Mixing Process Influenced by Wall Jet-Induced Shock Waves in Supersonic Flow

With the development of hypersonic air-breathing propulsion systems, such as the supersonic combustion ramjet (Scramjet) and rocket-based combined cycle (RBCC) engines, the mixing process of supersonic airstream with fuel in the engine combustor has been drawing more and more attention. Due to the compressibility effects, the mixing process in a supersonic condition is significantly inhibited. In the present paper, the novel strategy of wall-jet induced shock waves (WJISW) is put forward to realize mixing enhancement. The interaction process between WJISW and the supersonic mixing layer is researched and the enhanced-mixing mechanism is revealed, employing large eddy simulation (LES) methods. The fine vortex structures of the flow field are well captured and presented, utilizing the numerical schlieren technique. Detailed visualization results indicate that WJISW in a low frequency condition can result in the ‘region action mode’ (RAM) never reported before. The drastic dynamic behaviors including growth, deformation, and distortion in the interaction region can undoubtedly promote the mixing of upper and lower streams. The Reynolds stress distributions along the streamwise x-direction suggest that more intense fluctuations can be achieved with a low frequency WJISW. Moreover, a sharp increase in mixing layer thickness can be realized in the interaction region. The dynamic mode decomposition (DMD) analysis results show that the mixing layer evolution process is dominated by the mode induced by WJISW, which leads to the coexistence of both large- and small-scale structures in the flow field. The entrainment process corresponding to large-scale vortices and the nibbling process corresponding to small-scale vortices can obviously promote mixing enhancement. It is suggested that the present proposed strategy is a good candidate for enhanced-mixing with application to Scramjet and RBCC combustors.

Effects of magnesium particle size on combustion characteristic of martian ramjet engine

Experiments were conducted to determine the typical magnesium particle size effects on the Mg–CO2 Martian ramjet combustion characteristic. It was carried out at a mainstream mass flow rate of 110 g/s and a temperature of 800 K. In this study, two typical particle sizes (25.2 and 64.6 μm) of magnesium particles that are more easily obtainable in the industry were selected. The morphology, composition, and combustion efficiency corresponding to different particle sizes and equivalence ratios (ER) were analyzed. Results indicate that the 25.2-μm magnesium powder has a wider equivalence ratio adaptation range (ER = 0.13–0.27) and yielded less condensed phase deposition compared to 64.6-μm Mg powder. Combustion efficiency of 85.6% was achieved under the aforementioned conditions. Experimental data shows that 15% of the fine magnesium powder participated in the heterogeneous reaction and caused the pulsating combustion, demonstrating a low frequency (roughly 1 Hz) pressure oscillation. The amplitude even reached 14.6% of the chamber pressure. Remarkably, the condensed phase product produced by the small particle size is mainly magnesium oxide, resulting in an obvious agglomeration phenomenon. In contrast, unburned magnesium particles were detected in the combustion products of large magnesium powder due to the longer combustion time of coarse particles.

Reduced order design and investigation of intakes for high speed propulsion systems

Ramjet propulsion is commonly preferred to power supersonic and hypersonic vehicles for cruising faster than Mach 3. This is an elegant solution owing to the lean architecture which does not embody any rotating parts. Although the geometry of the engine is simple as compared to turbine based configurations, the flow physics through the engine duct is quite complex and the flow speeds modulate between the supersonic and subsonic regimes multiple times. The design and performance analysis of ramjet engines are vital to ensure that propulsion system can satisfy the flight trajectory requirements. Therefore, this study introduces a reduced order holistic approach for design and assessment of the flow development in high-speed propulsion systems composed of generic elements of ramjet/scramjet engine configurations. Accordingly, the intakes designed based on axisymmetric flow templates are used to provide the necessary freestream flow modulation prior to the isolator through which a normal shock assumption is applied. The resultant flow properties are utilized for the combustion module where the flow expansion within the combustor and nozzle components are computed based on 1D steady inviscid flow equations coupled with detailed chemistry approach and JANAF tables. The module was validated and verified with the experimental and numerical data obtained for a dual-mode ramjet/scramjet combustor. Consequently, the parameters such as thrust, fuel consumption and specific impulse are calculated to quantify the engine performance for each design. Finally, the employment of the low fidelity model is demonstrated over a family of ramjet flow paths where the design space is confined based on the requirements of a high-supersonic cruise vehicle.

Busemann diffuser for supersonic ramjet on detonation combustion of kerosene vapor

The possibility of using a convergent-divergent nozzle with a Busemann diffuser in ramjet engines for aerospace engineering applications is studied. Kerosene vapors are considered as a fuel, the combustion of which is carried out in the detonation mode at a hypersonic speed of the oncoming air flow. It is shown that at the altitude of about 40 km with the flight Mach number of 9, the Busemann diffuser can provide stable detonation combustion of kerosene with the efficiency of over 50%. Modeling was carried out on the basis of two-dimensional unsteady Euler equations of motion for axisymmetric flows of a multicomponent reacting gas using the modification of the Godunov numerical scheme, which provides the second order of the solution approximation in terms of spatial variables. The previously published kinetic scheme of ignition and combustion of kerosene vapors is supplemented by the processes of low-temperature combustion of the initial fuel.

Закономірності зміни оптимальних складу і параметрів силових установок з прямоточними контурами від швидкості надзвукового крейсерського польоту

There are two aircraft concepts for transoceanic flights at supersonic speed. The first is the aircraft with moderate supersonic cruising speed M = 1.7...2.2, and the second is the aircraft with cruising speed M ≥ 3. Propulsion with a ramjet engine (ramjet) is preferable for high supersonic speeds. However, a ramjet has no starting thrust and is uneconomical at subsonic flight speeds. Combined propulsion with a turbojet duct for takeoff and a ramjet duct for supersonic cruise flight are used to overcome this contradiction. There are known propulsion with turboramjet engines, as well as various propulsion schemes with turbofan engines (afterburning and duct-burning ones), in which the outer contour together with the afterburner and nozzle can be considered the ramjet duct. When the turbojet duct is turned off, the operating process of such turbofan engines is practically the same as the ramjet, which allows using the advantages of a ramjet in supersonic cruise flight. Because flight at supersonic cruising speed can be provided by different propulsion compositions, the choosing problem of the composition and parameters of the propulsion for an aircraft with supersonic cruising speed is relevant. A calculation-analytical method was used to select the composition and parameters of the propulsion, which is based on the maximum relative mass of the payload criterion. Using the aircraft mass balance equation this criterion can be represented as a minimum conditions for the relative mass of the fuel and propulsion. Using this method, for an aircraft with a given mass, aerodynamic characteristics and a given flight profile, for the indicated propulsion compositions, the change patterns of the minimum relative mass of fuel and propulsion in the range of M = 2.5...4 are established. The established patterns allow choosing the propulsion composition and parameters depending on the speed of the supersonic cruising flight or choosing the flight speed of the aircraft at a given propulsion composition. The dependance of the propulsion working process optimal parameters for takeoff and cruising mode on the speed of supersonic cruising flight according to the criterion minimum of relative mass of fuel and propulsion were obtained.

Preliminary Design of a Ramjet Engine: An Analytical Approach

This work is based on preliminary design of Ramjet engine. Ramjet engine is simplest type of gas turbine engine used, which consists of non-moving parts for its operation. It is mainly used for power generation at supersonic speeds. This type of engine is mostly used in missiles, with a few applications in aircrafts. The present approach to preliminary design was based on mathematical equations considering ideal conditions. For the design, aero-thermodynamic equations were used starting with intake and followed by diffuser, combustor and nozzle. For the design purpose, initially Mach 2 and thrust of 10 kN was selected as desired condition. Subsequently, it was analyzed for the varying Mach numbers starting from Mach 1.5 to Mach 4 at desired thrust of 10 kN. Also, the design was analyzed for varying thrust from 6 kN to 22 kN at desired Mach number of 2. The results obtained have been reported and effects of variation of parameters have been represented graphically. The graphs were obtained using GNU Octave 6.4.0. The design achieved could be used for further steps of CAD model generation and subsequent analysis for the required purpose. It can serve as a base for detailed design of the engine.

Computational evaluations of emissions indexes released by the STRATOFLY air-breathing combined propulsive system

PurposeThe purpose of this study is to provide the description of a computational methodology to model the combined propulsive systems of hydrogen propelled air-breathing scramjet vehicles and to evaluate the pollutant and climate-changing emissions.Design/methodology/approachEmissions indexes of nitrogen oxide (EINO) and water vapour released by the air turbo rocket (ATR) and dual mode ramjet (DMR) engines of the STRATOFLY air-breathing, hypersonic scramjet vehicle, propelled by hydrogen/air were evaluated. ATR engine operation was assessed for several cruise conditions in both subsonic and supersonic flight regimes in Ecosimpro software, which is an object-oriented thermodynamic design and simulation platform. ATR combustor inlet flow conditions play a key role in the computation of species mass fractions, and these conditions are highly dependent on turbomachinery performance and engine flight regime. A propulsive operational database was created by varying mass flow rates of fuel and flight conditions such as cruise speed and altitude to investigate possible engine operations. The all-inlet conditions in this map are provided to the Cantera-Python chemical/combustion chemistry solver implementing a specially designed and formulated 0D kinetic-thermodynamic methodology successfully used to model and simulate the electric spark ignition required to activate the combustion process of the reacting mixture in the ATR combustion chambers, whereas the coupled aero-thermodynamic/aero-propulsive 0D/1D code i.e. Scramjet PREliminary Aerothermodynamic Design (SPREAD), designed and developed by the Italian Aerospace Research Centre (CIRA) was used for DMR calculations. Results show low emissions of NO according to the optimized design of the ATR; on the other hand, a tuning of operational conditions is needed for DMR, with its complete re-design to be more conclusive. Analogously, the released amount of water vapour is in good agreement with the required combustion efficiency and the expected propulsive performance.FindingsResults show low emissions of NO according to the optimized design of the ATRs; on the other hand, a tuning of operational conditions is needed for DMR, with its complete re-design to be more conclusive. Analogously, the released amount of water vapour is in good agreement with the required combustion efficiency and the expected propulsive performance.Originality/valueApplications of innovative 0D/1D chemical kinetic methodology and in-house codes.

Busemann diffuser for supersonic ramjet engine with detonation combustion of hydrogen-air mixture

The expediency of using a Busemann diffuser in a scramjet engine with detonation combustion of hydrogen-air mixtures is studied. A method is developed for calculating an air intake that provides spontaneous initiation of oblique detonation in a hydrogen-air mixture under conditions of rarefied atmosphere at hypersonic freestream Mach numbers. The geometry of the Laval nozzle (convergent-expanding) with the Busemann diffuser for the free flow Mach number 9 is determined. It is shown that at an altitude of about 40 km, the detonation combustion of the hydrogen-air mixture in this nozzle can provide more than 0.4 tons of thrust and more than 35% efficiency. The mathematical model is based on unsteady two-dimensional Euler equations for axisymmetric multicomponent reacting gas flow. To simulate chemical transformations, the detailed kinetic scheme is used that takes into account 33 non-equilibrium reactions for nine components of the mixture. The heat capacity, enthalpy, and entropy of a mixture are calculated using the reduced Gibbs energy of the gas components. Numerical modeling is carried out according to the modified Godunov scheme of the second order of approximation in spatial variables.

Studies on a swirling heptane spray flame by large-eddy simulation

Swirling spray flames are widely encountered in gas-turbine combustors, rocket and ram-jet engines and internal combustion engines. Large-eddy simulation (LES) with different sub-grid scale (SGS) stress and combustion models is frequently used for studying the combustion mechanism of liquid fuel and design of combustors. In this paper, LES with a statistical-moment combustion model was used for studying a swirling spray flame. The statistical results for velocity, temperature and species concentration were validated by both experimental results and DNS database, showing the reliability of the simulation results. The instantaneous results give the spiral structures of flames formed by swirl, which can intensify combustion.

Numerical Simulation of Dynamic Process of Variable Flow Gas Generator

In this paper, the research on the flow-adjustable gas generator of the solid rocket ramjet engine is carried out, and the internal flow field of the gas generator during the actuation of the flow control valve is analyzed. The steady and unsteady constant value simulations of the gas generator's working process are carried out respectively, and the static and dynamic flow field distributions in the control valve of the solid jet engine are obtained. The negative regulation in the gas flow adjustment process of the solid-jet engine is reproduced, and the reason for this phenomenon is analyzed based on the simulation results.

Effects of the Air Inlet Angle on the Combustion and Ablation Environment of a Hybrid Powder-Solid Ramjet

A hybrid powder-solid ramjet (HPSR) that combined the advantages of a solid rocket ramjet (SRJ) and a powder ramjet engine was investigated in this research. To improve the combustion efficiency and optimize the inner wall thermal protection of the afterburner, the effects of the air inlet angle on the combustion and wall ablation environment were studied. The standard k - ε model, the eddy-dissipation model (EDM), and the boron particle ignition and combustion model were adopted to simulate the two-phase flow in the afterburners with different air inlet angles (45°, 60°, 75°, and 90°). The results showed that the global flow field and the distribution of the vortexes in the afterburner that had a significant influence on the ablation environment of the inner wall and the combustion efficiency were determined by the impact effect and the squeezing effect of the ram air on the primary fuel gas, which was affected by the air inlet angle. As the air inlet angle increased, the total combustion efficiency of the four cases first increased and then decreased, reaching 80.38%, 81.64%, 84.34%, and 83.26% for angles of 45°, 60°, 75°, and 90°, respectively. At the same time, the inner wall ablation became more severe because both the erosion effect of the condensed phase particles and the gas-flow scouring effect were enhanced, and a large temperature gradient was generated on the inner wall. The study results can provide a reference for designing the air inlet angle of an HPSR.

Influence of sinusoidal shock generator on mixing performance of hydrogen and air co-flow jets in dual-combustor ramjet

This article presents a comprehensive study on the impacts of a wavy shock generator on the flow structure and mixing of air and hydrogen jets in dual combustion ramjet engine. Our focus is on flame and flow characteristics directly downstream of air and fuel injector where shock wave and shear layers are presented. Three different sinusoidal profiles are considered to disclose the impacts of shock strength on fuel penetration and diffusion. SST turbulence model is applied to obtain turbulence closure. Our studies show that the existence of the wavy shock producer amplifies the barrel shock near the fuel injector. Besides, the formation of expansion waves by the presence of a wavy shock generator increases the shock interactions and augments the mixing performance. Our findings indicate that the propagation and interaction of the expansion waves with boundary shear layer augment vortices downstream of splitter and expands hydrogen mixing region substantially.

Air-to-Air Heat Exchanger in a Turbofan Ramjet Engine

Air-to-Air Heat Exchanger in a Turbofan Ramjet Engine

Experimental study of high-speed air intake performance by side clearance

Japan Aerospace Exploration Agency (JAXA) and universities plan the “High Mach Integrated Control Experiment” (HIMICO), which is a hypersonic flight test by JAXA's S-520 sounding rocket. The experimental aircraft for HIMICO is powered by a ramjet engine, whose intake is a rectangular mixed compression type with a design Mach number of 5.In this research, the effects of the side clearance (SC) on the intake performance are investigated at Mach 3.4 by supersonic wind tunnel tests. The side clearances are thin slits between the 2nd and 3rd ramps and sidewalls, which is provided for driving the variable geometry intake. Three types of models with different SC widths were compared. As the results, it is found that the 2nd ramp SC, which is located upstream of the throat, causes the performance deterioration and the 3rd ramp SC, which is located downstream of the throat can expand the starting range. Hybrid SC which has only the 3rd ramp SC obtains sufficient Mass Capture Ratio (MCR) and maintains the start condition over a wider nozzle throat height range due to the air bleed from the 3rd ramp SC. However, the performance of Hybrid SC deteriorates when adding the sideslip angle, β=5°, because of a leakage from the bleed slit located between the 2nd and 3rd ramps.