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Normal shock trains are a flow phenomenon of significance to ramjet engines, but it remains unclear what its structure is decided by and how it evolves with the incoming Mach number. To seek a theoretical explanation, the minimum entropy production principle is generalized to the quasi-steady behavior of normal shock trains in two-dimensional straight channels with uniform incoming flow. Numerical simulations are also performed to validate the model together with the data collected from public literature. The analysis suggests that the flow parameters of a normal shock train depend on the inviscid shock-shock interactions rather than the local boundary-layer separations, though the angles of two incident shocks should still be equal as similar to the case that complies with the free-interaction theory. The shock feet’s positions, meanwhile, are allowed to be coincident or not, free from the entropy restriction. This freedom of position explains why both symmetric and partially asymmetric normal shock trains could be found previously. Further theoretical calculations reveal the inclinations of two incident shocks increase first and then decrease with the incoming Mach number, peaking at 48.570 degrees when the Mach number reaches 1.753. It is also indicated that the Mach number range allowing for a normal shock train is 1.652 to 2.254, giving evidence for past observations. [ABSTRACT FROM AUTHOR]

In recent years, many turbofan engines that lack flight certification have been repurposed for other applications, such as power generation or providing air mass flow. This article investigates the feasibility and characteristics of using a standard turbofan combustion chamber in a ramjet application. One can of the combustion chamber from the D30K engine, which is of a can-annular type, was selected as the basis for this study. A redesign process was undertaken to determine the geometry of the combustion chamber based on existing scientific and technical literature. The geometric ratios of the D30K engine were used for validation, demonstrating satisfactory adaptation.The Gas Turb software was employed to determine the flow entrance conditions for a combustion chamber applied in a ramjet engine. Subsequently, the Fluent software was used to perform simulations of combustion using a non-premixed liquid phase method. The results indicated that, in this configuration, the induced vortices were not strong enough to stabilize the flame, making using a flame holder essential.Additionally, a parametric study was conducted to investigate the effects of the flame holder's size, number, position, and distance on parameters such as pressure loss, maximum temperature, Mach number, and flame stability. The results demonstrate that a multi-flame holder setup positively impacts performance, while more oversized flame holders are not recommended due to undesirable effects. [ABSTRACT FROM AUTHOR]

In recent years, many turbofan engines that lack flight certification have been repurposed for other applications, such as power generation or providing air mass flow. This article investigates the feasibility and characteristics of using a standard turbofan combustion chamber in a ramjet application. One can of the combustion chamber from the D30K engine, which is of a can-annular type, was selected as the basis for this study. A redesign process was undertaken to determine the geometry of the combustion chamber based on existing scientific and technical literature. The geometric ratios of the D30K engine were used for validation, demonstrating satisfactory adaptation.The Gas Turb software was employed to determine the flow entrance conditions for a combustion chamber applied in a ramjet engine. Subsequently, the Fluent software was used to perform simulations of combustion using a non-premixed liquid phase method. The results indicated that, in this configuration, the induced vortices were not strong enough to stabilize the flame, making using a flame holder essential.Additionally, a parametric study was conducted to investigate the effects of the flame holder's size, number, position, and distance on parameters such as pressure loss, maximum temperature, Mach number, and flame stability. The results demonstrate that a multi-flame holder setup positively impacts performance, while more oversized flame holders are not recommended due to undesirable effects.

For high-Mach-number incoming flow circumstances, a rotating detonation ramjet engine configuration is proposed in this research. By installing supporting blocks at the rear of the combustor, this configuration achieves continuous rotating detonation operation. Based on the Comparison of the flow structures obtained from the engine configuration with and without the supporting block before and after detonation ignition respectively, we obtain the intrinsic mechanism of detonation wave's propagation and re-initiation under the action of the supporting block. The supporting block creates a deflagration wave that is almost stationary before detonation ignition. In the detonation-ignited state, the deflagration wave is continually formed and traveling upstream under the influence of the supporting block, which is analogous to the periodical before detonation ignition of a transverse wave structure. The dynamic deflagration wave will cause the incomplete reactants behind the detonation wave to burn as the intensity of the detonation wave decreases. As a result, the incident shock wave is transformed into a Mach stem to achieve the re-initiation of the detonation wave. • The continuous spin detonation with kerosene as fuel is realized at flight Mach number M = 6. • Flow characteristics in the before and after detonation ignition states with and without supporting blocks are compared. • The supporting block may effectively contribute to maintaining the continuous propagation of the detonation wave. • The detonation wave's mechanism of propagation and re-initiation is examined under the influence of the supporting block. [ABSTRACT FROM AUTHOR]

This paper presents an analysis of approaches to ensuring the performance of a ramjet combustor using an unconventional method of measuring the thrust in testing an experimental model of a liquid-fuel combustor under flight conditions simulated by a fired heater. The thrust and drag of the combustor model were measured in firing tests with an attached air duct using a thrust meter. Thrust is a parameter that characterizes the energy potential of any engine. The conventional method of determining the thrust and economic characteristics of a heat engine involves lengthy multi-parameter studies and calculations of the operation performance. The new method allows one to detect the effect of any parameter or element of the structure on the physical process in a combustor in units of the developed thrust and evaluate the change in any parameter by the thrust level directly in firing tests. The proposed thrust measurement system makes it possible to refine the strategy of implementing the combustor operation at an early stage of testing, thus reducing the amount of intermediate calculations and saving a huge amount of time and money. This, combined with static pressure measurements along the combustor length, allows a guaranteed prediction of the direction of search for improving the operation performance. Ways are shown to improve the ramjet combustor performance using a thrust meter to evaluate the overall engine performance under conditions of limited technical and methodical possibilities in firing tests. [ABSTRACT FROM AUTHOR]

Experiments were conducted in this study to reveal the flow characteristics of a variable cross-section S-shaped isolator, when applying the steady-state back pressure at the isolator outlet. The self-excited oscillation characteristics of the shock train generated under the influence of steady-state back pressure at an incoming flow speed of Mach 2, were also studied. The findings suggest that pressure oscillation within the area affected by the shock train's flow field was significantly more potent than outside the affected area. Moreover, the forward movement velocity of the shock train in the variable cross-section S-shaped isolator was not uniform. The forward movement speed was slower when encountering sharp turns and faster during gentle turns. In the shock train flow field, high-frequency pressure oscillations, which mainly stemmed from the oscillations of the separated shock legs, propagated more readily within the flow field than low-frequency pressure oscillations. The significant separation of the shock train flow field will switch between the top and bottom walls, and the frequency of pressure oscillation in the large separation region is low. On another note, the closer the distance between two points is in the shock flow field, the stronger the coherence of pressure oscillations will be. In the distance upstream of the shock train flow field, the turbulent boundary layer (TBL) determines pressure oscillations instead of the shock train flow field, so the coherence was very high. [ABSTRACT FROM AUTHOR]

Compression by shock waves is a specific way of compressing gases. It has been practically applied for many years in supersonic flying objects. The idea of using this method in rotary engines is extremely appealing because one disk can replace several or a dozen disks of an axial compressor, significantly reducing the weight and production costs of the engine and lowering the fuel consumption due to possible increased compression ratio. This paper presents a review of existing technical solutions and the results of published research devoted to the construction of shock wave compression rotary engines: patents, scientific publications describing various research methods, numerical calculations, and the experimental results of unusual technical solutions. The characteristic solutions and problems that arose during the implementation of these methods are presented and described. Judging from the presented overview, these have wide application possibilities, and an enormous intellectual and financial effort has been put into the construction of such engines. Conversely, there is a rather hermetic group of scientists involved in this activity. [ABSTRACT FROM AUTHOR]

The screening method has been used to investigate the effects of magnesium fluoride, polytetrafluoroethylene, and ammonium hexafluorotitanate additives in a composite propellant on the basis of boron or aluminum dodecaboride as a fuel (37%), ammonium perchlorate as an oxidizer and an active binder on the parameters of propellant combustion. At pressures of 1.2 and 2.4 MPa, propellant combustion rates were measured, samples of condensed combustion products were taken, and burning residue in the form of a carcass was investigated. Using the cerimetric method of chemical analysis of the selected combustion products, incompleteness of propellant combustion was indentified and the efficiency of its energy release was evaluated. It is shown that the addition of small quantities (1%) of the said substances to the propellant makes it possible to control the velocity of its combustion and, in some cases, to raise the efficiency of energy release. [ABSTRACT FROM AUTHOR]

Oscillating flow is one challenge for wide-Mach-number-range flight with supersonic/hype-rsonic vehicles. Aiming to mitigate the large-amplitude forced shock-wave oscillation, a 2D wavy surface has been implemented onto the flat-bottomed wall of the Sajben diffuser under downstream pressure disturbance. The oscillating SBLI in the diffuser is captured using the finite volume method with the second-order implicit dual-time-stepping method. Impacts of the wavy surface on the forced shock-wave oscillation are numerically investigated. It is found that increasing the wavy surface's depth benefits mitigation of the shock-wave oscillation amplitudes on the walls under the given conditions, but that decreasing the wavy surface's length may increase or decrease the oscillation amplitudes, depending on the specific value. The mitigating mechanism is interpreted from two viewpoints, i.e., the shock-wave stability and the work performed by a moving shock-wave. The transient second shock-wave temporally appears in the flow field and can be explained by the post-shock expansion. [ABSTRACT FROM AUTHOR]

The energy efficiency of a supersonic ramjet engine with detonation combustion of a hydrogen-air mixture in an axisymmetric Laval nozzle is considered. First of all, within the framework of the one-dimensional (1D) theory, it is shown that extreme combustion corresponding to the Chapman-Jouguet (C.-J.) deflagration, like constant pressure combustion, surpasses detonation in energy efficiency. However, this advantage becomes insignificant with an increasing the free-flow Mach number. Based on the results of the 1D theory, an axisymmetric Laval nozzle is constructed, which provides steady-state detonation combustion of a stoichiometric hydrogen-air mixture at an altitude of about 40 km and with freestream Mach numbers 9 and 12. The efficiency of detonation combustion in this case is about 35% instead of the maximum 65% according to the 1D theory. • In terms of efficiency, detonation reaches deflagration at a high flight Mach number. • In 1D-theory a stationary combustion in Laval nozzle has a maximum flight Mach number. • 1D theory gives overestimated values of the efficiency and expansion of the diffuser. • The detonation slope expands the limits of its applicability in terms of speed and height. • The steady-state detonation in the Laval nozzle can provide an efficiency of about 35%. [ABSTRACT FROM AUTHOR]

The mathematical model and numerical method proposed in this study make it possible to study the regularities of gasification of solid porous fuel when hot gases are filtered through it and estimate the maximum operating time of the propulsion system. The features of gasification are studied for the case when polymethyl methacrylate is used as a solid fuel. It is shown that by changing the characteristics of the system and the process conditions, it is possible to change the flow rate of gaseous products from the gas generator, which will affect both the operating conditions of the cooling system for the materials of the combustion chamber structure and the value of the engine thrust. It is also shown that the operating time of a jet engine with a low-temperature gas generator increases with decreasing temperature and consumption of gaseous solid fuel combustion products, the interphase heat transfer coefficient, and the characteristic pore size of the fuel. [ABSTRACT FROM AUTHOR]

Experiments were conducted in this study to reveal the flow characteristics of a variable cross-section S-shaped isolator, when applying the steady-state back pressure at the isolator outlet. The self-excited oscillation characteristics of the shock train generated under the influence of steady-state back pressure at an incoming flow speed of Mach 2, were also studied. The findings suggest that pressure oscillation within the area affected by the shock train’s flow field was significantly more potent than outside the affected area. Moreover, the forward movement velocity of the shock train in the variable cross-section S-shaped isolator was not uniform. The forward movement speed was slower when encountering sharp turns and faster during gentle turns. In the shock train flow field, high-frequency pressure oscillations, which mainly stemmed from the oscillations of the separated shock legs, propagated more readily within the flow field than low-frequency pressure oscillations. The significant separation of the shock train flow field will switch between the top and bottom walls, and the frequency of pressure oscillation in the large separation region is low. On another note, the closer the distance between two points is in the shock flow field, the stronger the coherence of pressure oscillations will be. In the distance upstream of the shock train flow field, the turbulent boundary layer (TBL) determines pressure oscillations instead of the shock train flow field, so the coherence was very high.

Oscillating flow is one challenge for wide-Mach-number-range flight with supersonic/hype-rsonic vehicles. Aiming to mitigate the large-amplitude forced shock-wave oscillation, a 2D wavy surface has been implemented onto the flat-bottomed wall of the Sajben diffuser under downstream pressure disturbance. The oscillating SBLI in the diffuser is captured using the finite volume method with the second-order implicit dual-time-stepping method. Impacts of the wavy surface on the forced shock-wave oscillation are numerically investigated. It is found that increasing the wavy surface’s depth benefits mitigation of the shock-wave oscillation amplitudes on the walls under the given conditions, but that decreasing the wavy surface’s length may increase or decrease the oscillation amplitudes, depending on the specific value. The mitigating mechanism is interpreted from two viewpoints, i.e., the shock-wave stability and the work performed by a moving shock-wave. The transient second shock-wave temporally appears in the flow field and can be explained by the post-shock expansion.

Design and manufacture of a hypersonic pre-cooled turbojet engine have been developed so far. Wind tunnel tests of the engine have been conducted using the Ramjet Engine Test Facility (RJTF) at JAXA Kakuda Space Center, which simulates Mach 4 flight conditions. The shock oscillation was found and mass capture ratio (MCR) was around 40.1%, which is lower than the design target. This paper shows CFD analysis to clarify details of the experiment results and to find the way to improve the intake performance. In this paper, a virtual nozzle is used to simulate experiment results. From CFD analysis, the intake buzz occurs. Shock patterns from CFD analysis agree with the experiment results quantitatively. When back pressure decreases, the intake buzz doesn’t occur and MCR increases to 79.1% and max TPR is 43.6%. [ABSTRACT FROM AUTHOR]

Results of a complex experiment-calculated investigation of the structure of a nonstationary gas flow in the working path of a hypersonic ramjet engine are presented. Methods of experimental determining the pressure and temperature distributions over the wall of an axisymmetric channel of complex geometry with a sudden widening are proposed and described. The problem on the gas flow over a ramjet engine positioned in a wind tunnel was solved. [ABSTRACT FROM AUTHOR]

A new and intuitive method for evaluating the efficiency of the compression process in a supersonic air intake is proposed. The method is based on a comparison of the geometric compression ratio of the flow tube in the air intake under consideration with the maximum possible isentropic compression. It is shown that a higher value of the geometric compression ratio in the air intake allows obtaining higher thrust characteristics, with other conditions being identical. [ABSTRACT FROM AUTHOR]

The coexistence of multiple stable states is indicative of self-organising processes occurring in the course of the combustor-inlet interactions in a ramjet engine and give rise to the appearance of various nonlinear phenomena. This paper provides a dynamic model that can describe the multiple stable states and the corresponding nonlinear effects to further investigate the dynamic interactions between combustor and inlet in a ramjet engine. Our study shows the whole engine can display distinct dynamic behaviours ranging from irreversibility to hysteresis and to various mode transitions, depending on different physical parameters. With the model, we also illustrate the role of the instability of the normal shock wave in impacting the whole engine's nonlinear dynamics. Additionally, we extend the previous studies of the classification of combustor-inlet interactions from a static framework to a dynamic framework, which helps to clarify the transient processes of the nonlinear interactions. This work offers a quantitative illustration of the combustor-inlet interactions in a ramjet engine by revealing its nonlinear dynamics and associated characteristics, therefore advancing our understanding of the nonlinear phenomena that exhibit in ramjet engines. [ABSTRACT FROM AUTHOR]

This study experimentally confirms the intensification of kerosene combustion in the case where a mixture simulating steam reformation products in a model combustion chamber with a supersonic velocity of the flow at the inlet. It is shown that the used mixture has a higher chemical activity than ethylene. The use of steam reformation products of hydrocarbon or synthetic propellants in schemes with pulse-periodic combustion control increases combustion efficiency without the use of special design solutions for organizing initiation and stable combustion. [ABSTRACT FROM AUTHOR]

The creation of Russian supersonic passenger planes is considered. The development of promising airliners for no more than 20 passengers is discussed. A hybrid power unit consisting of turbojet and supersonic ramjet engines is proposed. The proposed hybrid system performs well both in ascent and descent and in cruising. This approach has undeniable economic and engineering benefits. [ABSTRACT FROM AUTHOR]

The article considers several solid ethinyl derivatives (actually existing diethinylbenzene and a number of compounds not yet synthesized) as potential dispersants of fuels for direct-flow rocket engines. The density and enthalpy of the formation of the considered hypothetical ethinyl derivatives are estimated. Some of important parameters are studied characterizing their effectiveness as dispersants. The calculation and parametric studies of the effectiveness of the compounds as fuel components for direct-flow rocket engines of specific atmospheric aircraft of the Meteor-type are carried out, the results of which confirm their efficiency. [ABSTRACT FROM AUTHOR]

The standard enthalpy of formation of 1,4-diethynylbenzene experimentally to be 500.6 ± 6.7 kJ/mol. Calculations show that 1,4-diethynylbenzene has a relatively high adiabatic combustion temperature (about 1970 K at a pressure of 5 atm) and very high heat of combustion in oxygen (42 MJ/kg); therefore, 1,4-diethynylbenzene can be used as the basis to develop an effective dispersing agent of solid fuel that provides an adiabatic combustion temperature of up to 2500 K at a heat of combustion far exceeding the values provided by HMX and the other azides of N-heterocycles previously proposed for this purpose. [ABSTRACT FROM AUTHOR]

The identification of the chain origin of gas combustion at atmospheric and elevated pressure has opened up new aspects of the theory and wide-ranging possibilities for effectively managing the processes of combustion, explosion, and detonation by controlling the branching and loss rates of active intermediate particles, such as atoms and radicals, by using inhibitors and promoters. In this paper, we briefly describe methods for preventing fires and explosions of methane-air mixtures, including those in coal mines, ignition and explosion of hydrogen-air mixtures, and the transition of combustion to detonation in the current ramjet engine model. The results of interdepartmental tests and experimental data illustrating the destruction of a stationary detonation wave by the small impurities of the lower hydrocarbons are given. [ABSTRACT FROM AUTHOR]

Results of a complex experiment-calculated investigation of the structure of a nonstationary gas flow in the working path of a hypersonic ramjet engine are presented. Methods of experimental determining the pressure and temperature distributions over the wall of an axisymmetric channel of complex geometry with a sudden widening are proposed and described. The problem on the gas flow over a ramjet engine positioned in a wind tunnel was solved. [ABSTRACT FROM AUTHOR]

It is of great significance to study the hysteresis phenomenon in the unstart/restart process of supersonic inlets for restart control in the practical operation of ramjet engines. In this paper, the unsteady numerical simulations are conducted for the unstart/restart processes of a ramjet engine with a two-dimensional twin-duct inlet configuration under asymmetric incoming-flow conditions (M a 0 = 2.6,β = 2°), and the unstart/restart characteristics of the inlet caused by the plug throttling and the thermal choking are compared and analyzed. The results show that from the perspective of source incentives, the hysteresis loop of the restart process of the leeward inlet caused by the thermal choking is larger than that of the restart process of the leeward inlet caused by the plug throttling, which indicates that it is difficult to verify the restart performance of the ramjet with the two-dimensional twin-duct inlet configuration in the actual operating process by the inlet's plug-throttling experiment in the conventional wind tunnel. In addition, in the practical operations, the increase of fuel gas causes the transition of the leeward inlet to the unstart state, and part of the fuel gas reacts and combusts stably on the external compression surface of the leeward inlet, which easily leads to the burning of the inlet if there is without any special thermal protection. [ABSTRACT FROM AUTHOR]

In this study, a rocket-aided ramjet engine (RARE) configuration is proposed that is aimed at improving the thrust gap and enhancing the engine thrust stability in the ramjet-to-scramjet mode transition of dual-mode scramjet (DMSJ) engines. The combustion mode and relevant heat release characteristics in the kerosene-fueled combustor during the mode transition are analyzed in detail through numerical simulations and ground direct-connect tests at the simulated flight condition of Mach 6. In the ramjet mode, the combustion mode distribution is dominated by diffusion and supersonic combustion, whereas the heat release comes mainly from diffusion and subsonic combustion. The subsonic combustion mode is that which has the highest heat release efficiency, accounting for 28.6% of the total combustion volume, and contributing 67.9% of the total heat release. The combustion heat release level in the rocket-aided scramjet mode is higher than the simple sum of the pure rocket and scramjet modes. The RARE combustor effectively maintains the heat release level and stabilizes the distribution of several different combustion modes from the ramjet to the scramjet mode using the embedded rocket. Furthermore, a robust method for a smooth ramjet-to-scramjet mode transition is provided for ramjet engines. Compared with the engine operation without embedded rocket, the engine thrust variation is reduced from 50.77% to 1.33% with the aid of the embedded rocket, which enhances the operational stability of the engine considerably during the mode transition. [ABSTRACT FROM AUTHOR]

Boron is one of the most attractive additives of fuel used for ramjet engines, due to its high volumetric and gravimetric heating value. Studies focusing on the ignition and combustion model of a single boron particle have been taken place for a long time, the most outstanding model is proposed by the group of Kuo, which is simple and can be well used in CFD application. However, some drawbacks were found in this model, and a new model was proposed. In the new extended model, some physical processes of boron ignition and combustion were considered. At the ignition stage, the balance of (BO) n on the liquid oxide layer surface was considered. It was diffused from the boron-boron oxide interface, and then consumed by both the heterogeneous reactions and the evaporation of diboron dioxide on the surface. The evaporation process of boron oxide and reaction between water vapor and boron oxide were considered. At the combustion stage, the evaporation and boiling processes of boron were considered due to the high combustion temperature, some chemical reaction rates were adjusted. The influence of forced convection on the particle was considered in both ignition and combustion stage. A careful validation shows that the extended model is suitable for a single boron particle, especially at the combustion stage. Effects of gas temperature, gas pressure, oxygen mole fraction on boron ignition and combustion in ramjet engine were studied and could be useful for ramjet engine design. [ABSTRACT FROM AUTHOR]

The advantages of the bicalibre configuration of the volumetric packaging in comparison with the single-calibre one are shown within the framework of the concept of ballistic efficiency (BE) enhancement of the aerial vehicles (AV) with prevailing coast phase. The rational range of booster and coaster calibre ratios of bicalibre aerial vehicles for both solid propellant jet engines and combined propulsion units are defined. [ABSTRACT FROM AUTHOR]

This paper presents a numerical study of the combustion of a hydrogen-air mixture in a model ramjet combustor with separate hydrogen and air supply during activation of O molecules by resonant laser radiation at a wavelength of 762.3 nm and 193.3 nm. The calculation is made using the parabolized Navier-Stokes equations taking into account chemical reactions, laser irradiation, and the nonuniformity of air parameters at the combustor inlet due to the complex gas-dynamic structure of the flow in the air intake. It is shown that the combustion completeness at the combustor outlet can be increased by a factor of 2.8 by redistributing the hydrogen supply through the system of fuel tank pylons. Further increase in the combustion completeness can be obtained by exposure of a narrow flow region to resonant laser radiation, more effectively at a wavelength of 193.3 nm. The combination of laser exposure with hydrogen supply redistribution increases the combustion efficiency by a factor of more than 4.7 compared to the base case. In this case, this provides a 95% increase the longitudinal force component in the portion of the internal engine duct that provides a positive contribution to the thrust. Estimation of the energy efficiency of using laser radiation shows that the laser energy input required to achieve this effect is 40-80 times (depending on the fuel supply method) less than the increase in the chemical energy (compared to the case of no laser exposure) released due to fuel combustion. [ABSTRACT FROM AUTHOR]

Although ramjet has the advantages of high-speed flying and higher specific impulse, the performance parameters will decline seriously with the increase of flight Mach number and flight height. Therefore, the investigation on the thermodynamic performance of ramjet is very crucial for broadening the working range. In the current study, a typical ramjet model has been employed to investigate the performance characteristics at wide working conditions. First of all, the compression characteristic analysis is carried out based on the Brayton cycle. The obtained results show that the specific cross-section area (A 2 and A 5 ) and the air-fuel ratio ( f ) have a great influence on the ramjet performance indexes. Secondly, the thermodynamic calculation process of ramjet is given from the view of the pneumatic thermal analysis. Then, the variable trends of the ramjet performance indexes with the flow conditions, the air-fuel ratio ( f ), the specific cross-sectional area (A 2 and A 5 ) under the fixed operating condition, equipotential dynamic pressure condition and variable dynamic pressure condition have been discussed. Finally, the optimum value of the specific cross-sectional area (A 5 ) and the air-fuel ratio ( f ) of the ramjet model at a fixed work condition (Ma=3.5, H=12 km) are obtained. [ABSTRACT FROM AUTHOR]

Multi-variant three-dimensional numerical simulations demonstrate the feasibility of the continuous- detonation process in an annular combustor of a ramjet power plant operating on hydrogen as fuel and air as oxidant in conditions of flight at a Mach number of M = 5.0 and an altitude of 20 km. Conceptual schemes of an axisymmetric power plant, 400 mm in external diameter and 1.3 to 1.5 m in length, with a supersonic intake, divergent annular combustor, and outlet nozzle with a frusto-conical central body are proposed. Calculations of the characteristics of the internal and external flows, with consideration given to the finite rate of turbulent-molecular mixing of the fuel mixture components with each other and with the combustion products, as well as the finite rate of chemical reactions and the viscous interaction of the flow with the bounding surfaces, have shown that, in these flight conditions, the engine of such a power plant has the following performance characteristics: the thrust, 10.7 kN; specific thrust, 0.89 (kN s)/kg; specific impulse, 1210 s; and specific fuel⋅consumption 0.303 kg/(N h). In this case, the combustor can operate with one detonation wave traveling in the annular channel at an average velocity of 1695 m/s, which corresponds to a detonation wave rotation frequency of 1350 Hz. It is shown that, an operating combustor has regions with subsonic flow of detonation products, but the flow is supersonic throughout its outlet section. [ABSTRACT FROM AUTHOR]

We discuss the method of comparing pulse detonation engines (PDE) and engines with combustion in subsonic flow (ramjet) by means of their specific impulse used by the "Center of Pulse-Detonation Combustion" (CPDC). We demonstrate that the method used by CPDC to calculate the performance of PDE overstates the value of specific impulse relative to its actual value by a factor of at least two. In contrast, the values of specific impulse for ramjet are understated. As a result, the specific impulse of PDE significantly exceeds that of ramjet or is close to it. We investigate these misleading conclusions, and demonstrate their complete failure. [ABSTRACT FROM AUTHOR]

The flow and combustion dynamics in the ramjet sustainer of an integrated rocket-ramjet (IRR) engine are investigated systematically. The physical model includes the entire engine flowpath, from the freestream in front of the inlet to the exit of the exhaust nozzle. The flowfield obtained previously is used as the initial condition for the present analysis, so that the complete operation history of the engine can be obtained. Various physiochemical processes dictating the engine internal flow development and flame behavior are examined. Emphasis is placed on the interactions between the inlet and combustion dynamics, as well as the overall engine performance. The key mechanisms for driving and sustaining the flow oscillations are explored. In addition, the acoustic wave structures in the engine flowpath are identified. [ABSTRACT FROM AUTHOR]

performance study of a water ramjet engine is described. The engine is powered by the reaction of a magnesium-based propellant and ingested water. In this study, a solid propellant, which consisted of a large percentage of magnesium, a binder and a small amount of oxidant, was used as a hydro reactive fuel. Cold water was injected into the combustion chamber as a main oxidant. A scaled-down experimental engine was tested in a direct-connect ground testing system to characterize the factors influencing the engine performance. The results show that the increasing of total water/fuel ratio, an addition of secondary water intake along the combustion chamber, a larger magnesium content in the solid propellant, a smaller primary water injection angle towards the coming main flow, and a higher primary injection pressure were all able to promote the engine performance. The maximum engine performance was obtained in test 08, and with all tests, an appropriate set of parameters and conditions for the optimum engine performance were determined [ABSTRACT FROM AUTHOR]

This paper discusses an approach to implementing the operation and designing the flow duct of a ramjet combustor. Methods of air compression, ignition, flame stabilization, and fuel combustion in the flow are considered for the purpose of implementing an effective process on a short length with moderate total pressure losses. Advantages of narrow-mode ramjet engines are noted. Comparative results of experiments on the burning of hydrocarbon fuels in the ramjet combustors are given. [ABSTRACT FROM AUTHOR]

This paper gives a brief overview of investigations and developments of combustion chambers of high-speed ramjet engines performed at Moscow Aviation Institute (MAI). Shadow photographs are given which show changes in the structure of regular shocks versus heat input to the flow during hydrogen combustion. The impact of oblique shock waves on combustion is studied by visualizing OH radical luminescence. Because of the significant dependence of the combustion effectiveness on the flow parameters and structure, it is relevant to study ramjet engine designs with a variable air flow duct of the combustion chamber, which allows adjustment of the working section to obtain the required parameters and characteristics over a wide range of flight conditions. A large series of tests of such combustion chambers have been performed under a joint French—Russian program. Measurement results corresponding to flight Mach numbers 2–7 are given. Mixtures of chemically pure methane and hydrogen was used as fuel. Characteristics of a three-stage heater of model flow developed at MAI for testing wide-range combustion chambers are presented. [ABSTRACT FROM AUTHOR]

A preliminary study on the operating performance of a certain kind of water ramjet engine was implemented theoretically. With regard to this proposed powerplant operating in an available water environment, inspiration drawn from a supercavitation phenomenon coupled with a hydroreactive characteristic of several metals was the inducement for its attractive development. In terms of the requirement to provide the highest possible performance, a multiple water-injection mechanism was employed; i.e. the introduction of sufficient ambient water was suggested in order to maintain a successful combustion, while in the meantime not being too large to avoid a potential condensation phenomenon at the nozzle exit. Then against a ramjet carrying a metal fuel grain with a 50 per cent magnesium mass fraction, a performance analysis proceeded on the basis of a proposed thermodynamic calculation theory. The upper limit of the primary water-fuel ratio was predicted as 2.4 when considering the conditional temperature of the main magnesium-water reaction. Peak values of both specific impulse and thrust exist on the order of 4695.86 N s/kg and 2.57 kN respectively, while the water-fuel ratio approached 4.2. Furthermore, rules governing the characteristic velocity along with the water-fuel ratio were also examined. Then, altering physical conditions such as metal fuel formation and chamber pressure, a conclusion was drawn that a higher performance was related to a relatively higher chamber pressure and a higher magnesium mass fraction as long as normal working conditions were guaranteed. [ABSTRACT FROM AUTHOR]

The fuel injection system is the main process in supersonic vehicles. In this article, injection of fuel jet through a trapezoidal strut in the coaxial supersonic air stream is investigated. The numerical method is applied for the simulation of three-dimensional trapezoidal lobe strut in supersonic flow. Hydrogen gas is injected from the end of the trapezoidal strut injector to augment fuel mixing and diffusion downstream by intensifying vortices inside the domain. The power of the vortices varies by the injector velocity and free stream condition. To simulate our model, RANS equations with the SST turbulence model are solved. Effects of free-stream velocity and fuel jet total pressure on fuel distribution and mixing are also investigated. Our results indicate that fuel mixing enhances about 18% when a rectangular strut is replaced by a trapezoidal one. An achieved finding disclosed that the strength of vortices is amplified by the presence of a trapezoidal strut injector. [ABSTRACT FROM AUTHOR]