Thrust characteristics of an airbreathing pulse detonation engine in supersonic flight at various altitudes

The main thrust characteristics, such as thrust force, specific impulse, specific fuel consumption, and specific thrust, of a pulse detonation engine (PDE) with an air intake and nozzle in conditions of flight at a Mach number of 3 and various altitudes (from 8 to 28 km above sea level) are for the first time calculated with consideration given to the physicochemical characteristics of the oxidation and combustion of hydro-carbon fuel (propane), finite time of turbulent flame acceleration, and deflagration-to-detonation transition (DDT). In addition, a parametric analysis of the influence of the operation mode and design parameters of the PDE on its thrust characteristics in flight at a Mach number of 3 and an altitude of 16 km is performed, and the characteristics of engines with direct initiation of detonation and fast deflagration are compared. It is shown that a PDE of this design greatly exceeds an ideal ramjet engine in specific thrust, whereas regarding the specific impulse and specific fuel consumption, it is not inferior to the ideal ramjet.     

Thrust characteristics of an airbreathing pulse detonation engine in flight at mach numbers of 0.4 to 5.0

Multidimensional calculations are performed to demonstrate that, by its characteristics, the pulse detonation engine (PDE) is a unique type of ramjet propulsion system, which can be used in both subsonic and supersonic aircraft. By a number of examples, it is shown that, in various thrust characteristics, such as the specific impulse, specific fuel consumption, and specific thrust, the PDE substantially exceeds ramjet engines.     

Testing of solid fuel ramjet with measurement of thrust characteristics in aerodynamic facilities

A design of an axisymmetric solid fuel ramjet consisting of a multi wedges nose air intake, solid fuel gas gene-rator, combustion chamber, and a nozzle, was developed. According to this design, a ramjet model for tests in the ground wind-tunnel facilities was fabricated. Experiments with solid fuel combustion were carried out in the Transit-M and T-313 wind tunnels, ITAM SB RAS, at air-flow Mach numbers М = 2.5?5.0. High values of the internal and net excess thrust were obtained.     

Investigations of three-dimensional flow characteristics in a liquid ramjet combustor using the PIV method

Three-dimensional flow characteristics in a liquid fuel ramjet combustor were investigated using the PIV method. The combustor had two rectangular inlets that form a 90-degree angle with each other, with intake angles of 30 degrees. Three guide vanes were installed in each rectangular inlet to improve flow stability. The experiments were performed in a water tunnel test with the same Reynolds number as Mach 0.3 at the inlet. PIV software was developed to measure the characteristics of the flow field in the combustor. Accuracy of the developed PIV program was verified with a rotating disk experiment and standard data. The experimental results showed that the two main streams from the rectangular intakes collided near the plane of symmetry and generated two large longitudinal vortices, which was in agreement with three dimensional computational results. A large and complex threedimensional recirculating flow was measured behind the intakes.     

Estimation of energetic efficiency of heat supply in front of the aircraft at supersonic accelerated flight. Part 1. Mathematical models

Fuel economy at boost trajectory of the aerospace plane was estimated during energy supply to the free stream. Initial and final flight velocities were specified. The model of a gliding flight above cold air in an infinite isobaric thermal wake was used. The fuel consumption rates were compared at optimal trajectory. The calculations were carried out using a combined power plant consisting of ramjet and liquid-propellant engine. An exergy model was built in the first part of the paper to estimate the ramjet thrust and specific impulse. A quadratic dependence on aerodynamic lift was used to estimate the aerodynamic drag of aircraft. The energy for flow heating was obtained at the expense of an equivalent reduction of the exergy of combustion products. The dependencies were obtained for increasing the range coefficient of cruise flight for different Mach numbers. The second part of the paper presents a mathematical model for the boost interval of the aircraft flight trajectory and the computational results for the reduction of fuel consumption at the boost trajectory for a given value of the energy supplied in front of the aircraft.     

Estimation of energetic efficiency of heat supply in front of the aircraft at supersonic accelerated flight. Part II. Mathematical model of the trajectory boost part and computational results

The fuel economy was estimated at boost trajectory of aerospace plane during energy supply to the free stream. Initial and final velocities of the flight were given. A model of planning flight above cold air in infinite isobaric thermal wake was used. The comparison of fuel consumption was done at optimal trajectories. The calculations were done using a combined power plant consisting of ramjet and liquid-propellant engine. An exergy model was constructed in the first part of the paper for estimating the ramjet thrust and specific impulse. To estimate the aerodynamic drag of aircraft a quadratic dependence on aerodynamic lift is used. The energy for flow heating is obtained at the sacrifice of an equivalent decrease of exergy of combustion products. The dependencies are obtained for increasing the range coefficient of cruise flight at different Mach numbers. In the second part of the paper, a mathematical model is presented for the boost part of the flight trajectory of the flying vehicle and computational results for reducing the fuel expenses at the boost trajectory at a given value of the energy supplied in front of the aircraft. Sections, formulas, and figures have a numbering continued from the first part of the paper printed in the journal “Thermophysics and Aeromechanics”, 2008, Vol. 15, No. 4, P. 537–548.     

用于高超音速推进的反循环发动机热力性能分析

符号表fIkg空气对应的燃料流量ISP燃料比冲Ti后燃烧室出口温度马比推力，燃料空气当量比WT涡轮输出功MT飞行马赫数＄前登燃烧室出口温度。T涡轮膨胀比H飞行高度T4＊热交换器出口温度7TC压气机压比1引言八十年代以来，美、俄、德、法、英、日等主要空间大国均提出了各自的高超音速计划，如NASP、Sanger、HOTOL、STAR等。各国的方案在Ma＞6．5均采用火箭发动机或超燃冲压发动机，对于Ma＜6．5采用何种方案则分歧较大，有涡轮一冲压组合发动机方案、ATR方案、LACE方案等。本文对反循环发动机（InverseCgcleEngine）方案作了进…     

Three-dimensional numerical simulation of the characteristics of a ramjet power plant with a continuous-detonation combustor in supersonic flight

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 ₀ = 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.     

Combustion characteristics of a dual-mode scramjet combustor with cavity flameholder

Combustion characteristics of a laboratory dual-mode ramjet/scramjet combustor were studied experimentally. The combustor consists of a sonic fuel jet injected into a supersonic crossflow upstream of a wall cavity pilot flame. These fundamental components are contained in many dual-mode combustor designs. Experiments were performed with an isolator entrance Mach number of 2.2. Air stagnation temperatures were varied from 1040 to 1490 K, which correspond to flight Mach numbers of 4.3–5.4. Both pure hydrogen and a mixture of hydrogen and ethylene fuels were used. High speed imaging of the flame luminosity was performed along with measurements of the isolator and combustor wall pressures. For ramjet mode operation, two distinct combustion stabilization locations were found for fuel injection a sufficient distance upstream of the cavity. At low T₀, the combustion was anchored at the leading edge of the cavity by heat release in the cavity shear layer. At high T₀, the combustion was stabilized a short distance downstream of the fuel injection jet in the jet-wake. For an intermediate range of T₀, the reaction zone oscillated between the jet-wake and cavity stabilization locations. Wall pressure measurements showed that cavity stabilized combustion was the steadiest, followed by jet-wake stabilized, and the oscillatory case. For fuel injection close to the cavity, a hybrid stabilization mode was found in which the reaction zone locations for the two stabilization modes overlapped. For this hybrid stabilization, cavity fueling rate was an important factor in the steadiness of the flow field. Scramjet mode combustion was found to only exist in the cavity stabilized location for the conditions studied.     

Characteristics of pulse detonation engine versus ramjet characteristics

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.     

On the possibility of reprocessing spent nuclear fuel and radioactive waste by plasma methods

The concept of plasma separation of spent nuclear fuel and radioactive waste is presented. An approach that is based on using an accelerating potential to overcome the energy and angular spread of plasma ions at the separation region inlet and utilizing a potential well to separate spatially the ions of different masses is proposed. It is demonstrated that such separation may be performed at distances of about 1 m with electrical potentials of about 1 kV and a magnetic field of about 1 kG. The estimates of energy consumption and performance of the plasma separation method are presented. These estimates illustrate its potential for technological application. The results of development and construction of an experimental setup for testing the method of plasma separation are presented.     

Flame dynamics under various backpressures in a model scramjet with and without a cavity flameholder

Flame dynamics under various backpressure conditions were experimentally investigated using direct flame visualization, high-speed CH* chemiluminescence imaging, and wall pressure measurements. The stagnation pressure and temperature used in the present study were 100 kPa and 2500 K, respectively, with a freestream Mach number of 4.5. Rectangular scramjet models with and without a cavity were used to explore the effects of the cavity on flame dynamics when operating in scramjet mode, ramjet mode, and unstart. The flow rate of the ethylene jet was varied to impose backpressures corresponding to each operation mode. For both models, reverse flame propagation was observed for ramjet mode and unstart. For ramjet mode, flame fluctuation occurred within the isolator due to the coupling of fluid dynamics and combustion. The presence of a cavity enhanced combustion and reduced flame fluctuation in both scram and ramjet mode. The cavity promoted unstart because of the greater heat release from combustion. Further research using spatially resolved diagnostic techniques is needed to identify the flame locations for ramjet mode and unstart.     

Effects of the penetration height of ethylene transverse jets on flame stabilization behavior in a Mach 2 supersonic crossflow

Effects of fuel jet penetration height on supersonic combustion behaviors were investigated experimentally in a supersonic combustion ramjet model combustor at a Mach speed of 2 and at a stagnation temperature of 1900 K. The jet-to-crossflow momentum flux ratio was varied to control the fuel-jet penetration height, using several injectors with different orifice diameters: 2, 3, and 4 mm. First, transverse nitrogen jets were observed to identify a relationship between the fuel jet penetration height and the momentum flux ratio by focusing Schlieren photography. Then, supersonic combustion behaviors of ethylene were investigated through combustion pressure measurements. Simultaneously, time-resolved images of CH* chemiluminescence and shadowgraphs were recorded with high-speed video cameras. Furthermore, a morphology of supersonic combustion modes was investigated for various equivalence ratios and fuel penetration heights in a two-dimensional latent space trained by the shared Gaussian process latent variable models (SGPLVM), considering CH* chemiluminescence images and the shock parameters. The results indicated that the penetration height of nitrogen jets was a function of the jet momentum flux ratio; this function was expressed by a fitting curve. Five typical combustion modes were identified based on time-resolved CH* chemiluminescence images, shadowgraphs, and pressure profiles. Even for a given equivalence ratio, different combustion modes were observed depending on the fuel penetration height. For an injection diameter of 3 and 4 mm, cavity shear-layer and jet-wake stabilized combustions were observed as the scram modes. On the other hand, although the cavity shear-layer and lifted-shear-layer stabilized combustions were observed, no jet-wake stabilized combustion was observed for an orifice diameter of 2 mm. Fuel penetration heights above the cavity aft wall were expected to affect the combustion behavior. Finally, a morphology of the supersonic combustion modes was clearly shown in the two-dimensional latent space of the SGPVLM.     

玉米秸秆生物油-柴油乳化油的燃烧特性

试验用生物油是玉米秸秆快速热解液化的产物,主要成分为含氧有机混合物和水,不宜直接作为燃料使用,但与柴油乳化后可实现其在发动机中应用.在一台未作改动的直喷式柴油机上研究了玉米秸秆生物油质量分数分别为10%(B10)和20%(B20)的生物油-柴油乳化油的燃烧特性.结果表明:与0号柴油相比,乳化油的滞燃期延长,预混燃烧放热...     

涡轮-冲压组合发动机模态过渡段性能模拟和概念探讨

１概述涡轮－冲压组合发动机是可望用于天地往返运输系统和高超声速民航运输的吸气式发动机。在地面起飞和低速飞行阶段以涡轮发动机模态工作,在高空高速阶段以冲压发动机模态工作．涡轮模态和冲压模态的相互转换过程称为模态过渡段。在过渡段中两种发动机共同工作以联合循环方式运行。组合发动机以联合循环方式工作的性能,不仅与组成它的涡轮发动机和冲压发动机本身的型式和特征有关,而且受到两类发动机相互关系以及调节机构的影响。所以,涡轮冲压组合发动机模态过渡段稳态和瞬态过程的研究,是组合发动机性能研究的重要组成部分［１－…     

Characteristics of dual-combustion ramjet

The authors discuss a possibility to use a diverging dual-combustion chamber as applied to high-supersonic boost ramjets operating at flight Mach numbers up to M_f = 8–10. Due to diverging, this chamber allows beginning the ramjet operation from flight Mach numbers M_{f ini} = 2–3. The diverging combustion chamber is characterized by a ratio of its exit cross-sectional area relative to the cross-sectional area of air-intake throat. This expansion area ratio is determined at M_f = M_{f ini}, but it should be the same at all flight Mach numbers M_f ? M_{f ini}, and depends on two factors: the location of a normal shock in the air-intake throat and the condition of reaching the critical velocity at the chamber exit. The dual-combustion chamber provides heat supply in its alone channel first to the subsonic flow and then, along with acceleration of the flying vehicle, to the supersonic flow, which is bound with a decrease in relative heating of working gas. Calculations of characteristics of an exemplified dual-combustion ramjet considered with a twodimensional air-intake were performed in the range of M_f = 3–7.     

Experimental and kinetic modeling study of combustion of JP-8, its surrogates and reference components in laminar nonpremixed flows

Experimental and numerical studies are carried out to construct reliable surrogates that can reproduce aspects of combustion of JP-8 and Jet-A. Surrogate fuels are defined as mixtures of few hydrocarbon compounds with combustion characteristics similar to those of commercial fuels. The combustion characteristics considered here are extinction and autoignition in laminar non premixed flows. The “reference” fuels used as components for the surrogates of jet fuels are n-decane, n-dodecane, methylcyclohexane, toluene, and o-xylene. Three surrogates are constructed by mixing these components in proportions to their chemical types found in jet fuels. Experiments are conducted in the counterflow system. The fuels tested are the components of the surrogates, the surrogates, and the jet fuels. A fuel stream made up of a mixture of fuel vapors and nitrogen is injected into a mixing layer from one duct of a counterflow burner. Air is injected from the other duct into the same mixing layer. The strain rate at extinction is measured as a function of the mass fraction of fuel in the fuel stream. The temperature of the air at autoignition is measured as a function of the strain rate at a fixed value of the mass fraction of fuel in the fuel stream. The measured values of the critical conditions of extinction and autoignition for the surrogates show that they are slightly more reactive than the jet fuels. Numerical calculations are carried out using a semi-detailed chemical-kinetic mechanism. The calculated values of the critical conditions of extinction and autoignition for the reference fuels and for the surrogates are found to agree well with experimental data. Sensitivity analysis is used to highlight key elementary reactions that influence the critical conditions of autoignition of an alkane fuel and an aromatic fuel.     

Exergy analysis of ramjet

An exergy analysis of possible regimes of energy supply to the air flow in the ramjet duct is carried out. A condition for the supply of a given heat amount to supersonic flow and a condition of the passage across the sound velocity are obtained for a duct with variable cross-sectional area. An analysis of the flow in a model ramjet duct at a pulsed-periodic energy supply is carried out. For a clear demonstration of possible schemes of heat supply in such a duct, a diagram in the temperature-exergy coordinates is proposed. A duct configuration in which the heat supply to supersonic flow is realized with regard for the limitation of the gas static temperature is proposed.     

Influence of resonators on the acoustic and propulsion performance characteristics of a ramjet ejector chamber under conditions with vibration hydrogen combustion

The influence of acoustic resonators on the acoustic and propulsion performance characteristics of a ramjet ejector chamber under conditions with vibration hydrogen combustion was experimentally examined. In the study, resonators having identical throats and different cavity diameters were used. For fixed-volume resonators the best propulsion performance characteristics were achieved in the case in which the cavity diameter differed little from the resonator throat diameter.     

Investigation of the efficiency of regenerative cooling of the ramjet combustor by gasification products of energy-intensive material

The results of mathematical modeling of the thermal state of combustion chambers with regenerative cooling for ramjet engines of promising flying vehicles are presented. The cooling of combustion chambers by the gasification products of a combined charge of the energy-intensive material is considered, where the polyethylene is used as a stuff, and the HMX-based compounds are used as the active substance. The flow rates of the cooling eneregy-intensive material are determined, which provide acceptable levels of temperatures of combustion chambers at various modes of engines operation are determined.