Simulation of unsteady combustion in a solid-propellant rocket motor

Various regimes of combustion in end-burning-grain solid-propellant rocket motors were examined within the framework of the phenomenological theory of unsteady combustion. A system of equations capable of describing the interaction between the process of burning and acoustic waves was derived. A specific feature of the problem is that its formulation involves two characteristic times: the acoustic time and oscillation amplitude variation time. These characteristic times differ by about three orders of magnitude, a circumstance that requires a high accuracy of calculations. Based on the quadratic approximation in oscillation amplitude, a simpler method for solving the problem was proposed, according to which only the effects associated with the oscillation amplitude variation time are taken into account. Numerical results were obtained for the simplest model of propellant burning, which contains the minimum number of parameters and disregards entropy waves in the combustion products. The steady and unsteady regimes of burning were identified. In the latter case, nonlinear effects may generate shock waves in the combustion chamber.     

Acoustic characteristics of a rocket combustion chamber: Radial baffle effects

This paper describes methods used for determining the characteristic acoustic modes and frequencies of a liquid-propellant rocket-motor combustion chamber and effects of radial baffles on the chamber’s acoustic field. A multi-point sensing experimental setup, including stationary and moving sensors, was used to measure characteristic frequencies and mode shapes of a combustion chamber. A new technique based on the comparison of signal phase angles from stationary sensors to that of a moving sensor was used to map complex characteristic mode shapes of a combustor. A three-dimensional Helmholtz acoustic solver was also developed using an efficient finite volume approach for complex geometries to simulate the acoustic field inside a combustor. Using this approach the effects of the convergent section of the nozzle and the number of radial baffles on the chamber’s dominant acoustic modes with no mean flow were investigated. We have shown that the classical reduction of characteristic frequency of tangential modes caused by radial baffles is due to longitudinalization of tangential modes and is a function of the blade length and weakly dependent on the number of blades. Also, conjugate spinning modes are decoupled and do not spin in any baffled combustor, independent of the number of blades. On the other hand the converging nozzle section of a combustion chamber modifies pure longitudinal modes in the radial direction and pure tangential modes in the longitudinal direction. Existence of some mixed tangential-longitudinal modes in a combustor is dependent on the ratio of the nozzle throat diameter to the combustor head plate diameter.     

Calculation of the unsteady burning velocity according to a known law of pressure variation during transient processes in a solid-propellant rocket engine

The burning velocity during pressure decay in a model solid-propellant engine caused by the opening of an additional nozzle concurrently with the operation of the main nozzle is calculated; i.e., the inverse problem of internal ballistics is solved. The calculations are performed using experimental pressure versus time dependences and the laws of conservation of mass and energy in the thermodynamic and isothermal approximations. The problem is solved in two formulations: using the basic parameters of the combustion chamber before depressurization and in a formulation based on using initial conditions for pressure decay. In the latter case, the unsteady burning velocities are smaller than the corresponding steady-state velocities at the same pressures and close to the calculated velocities obtained for the same pressure decays by solving the direct problem.     

Experimental demonstration of the operation process of a pulse-detonation liquid rocket engine

A low-frequency demonstrator of the operation process of a pulse-detonation liquid-fuel rocket engine, intended for shaping the future design of a new type of rocket engines for spacecraft control, is for the first time developed and tested. As a result, the conditions are determined under which the demonstrator provides reliable deflagration-to-detonation transition (DDT) in a single-pulse or repetition mode. Because DDT occurs within a very short distance (less than 10 bores of the detonation tube), the operating frequency of the demonstrator can be substantially increased.     

Flow, spray and combustion analysis by laser techniques in the combustion chamber of a direct-injection diesel engine

The purpose of this paper is to show how the analysis of in -cylinder flow, fuel injection, and combustion by means of state-of-the-art optical techniques, as laser light-sheet, laser doppler anemometry and laser shadowgraphy, can help to support the understanding of the interaction of swirl flow development, spray formation, auto-ignition and combustion in near production-line direct-injection diesel engines and thus advances the development of engines with lower fuel consumption and emissions.     

液体火箭发动机中声腔抑制不稳定燃烧的声学分析

发展了声腔的分析和数值模型,对液体火箭发动机不稳定燃烧的抑制作用进行了评定,通过迭代计算研究了二维流动和温度分布变化对声腔调 谐和稳定性能的影响,对不同的声腔几何尺寸和温度梯度的稳定性计算结果表明,燃烧带有较大开口面积的声腔会更大程度地改变振荡的空间分布,这种改变而且影响了驱动和抑制燃烧的机理,讨论了在声腔设计安排中正确选择声腔的几何尺寸,且比较了不同长度和不同直径声腔的阻尼特性,通过考察声吸收系统的方法来最优化系统的阻尼,得到了可供设计参考的结论。     

超声在固体火箭发动机无损检测中的应用

在固体火箭发动机的研制中，无损检测技术非常重要。本文介绍了一些主要的超声检测方法及其应用。脉冲多次反射法用于检测固体火箭发动机壳体与衬层之间的粘结质量。特性参数和扫频超声法以多层介质理论为基础，用于检测固体火箭发动机的深层粘结质量。     

Numerical investigation of the influence of crystallization of ultrafine particles of aluminum oxide on energy characteristics of solid-propellant rocket engine

The results of numerical investigation of a multiphase flow considering coagulation, crushing and crystallization of the particles of polydispersed condensate in the nozzles of solid-propellant rocket engine are presented. The influence of particles crystallization on the energy characteristics of the engine is shown.     

Aluminum combustion in a solid rocket motor environment

A model for prediction of particle radius and oxide cap size/shape versus time for an aluminum particle tracking a stream-tube through a solid rocket motor port has been developed. Following preliminary calculations leading to a postulated flame structure, a quasi-steady model to predict instantaneous consumption of aluminum and generation of condensed oxide (both as a cap on the aluminum particle and as smoke) as a function of instantaneous particle size, ambient conditions, and cumulative amount of oxide in the cap was developed. Finally, this model was imbedded into a framework tracking evolution of particle size, oxide cap size, and ambient conditions, which change as the particle travels along a stream-tube consuming oxidizer and releasing heat. Qualitative agreement was found between predictions and limited observations.     

Numerical analysis of self-excited tangential combustion instability for an MMH/NTO rocket combustor

This study presented the numerical simulation of the tangential combustion instability in a hypergolic liquid bipropellant rocket thrust chamber, which applied fuel liquid film cooling method and unlike impinging injectors. The liquid spray was modeled using Lagrangian approach, while the gas was regarded as Euler phase. Stress-blended eddy simulation and finite rate/eddy–dissipation model were adopted to simulate the turbulent combustion process. Consistent with the experiment results, this work successfully simulated the transformation of tangential combustion instability from standing mode to spinning mode. The mean pressure, amplitude and frequency of limit cycle oscillation were in good agreement with the experiment. There was a detailed analysis about the flow field, Rayleigh index, and driving mechanism of the combustion instability. It was found that the oscillation began with hot spots of heat release rate due to the interaction between the spray of impinging injectors and cooling fuel jet. More than that, cooling fuel jet also contributed to drive the oscillation. In the standing mode, injectors in the inner and outer rings drive the oscillation together, while the spinning mode is mainly driven by injectors in the outer ring. The pressure wave is subsonic and its Mach number is close to 1. It was shown that the pressure wave contained a complex structure divided into three parts. This led to the in-phase of the pressure along the axial direction and the double-peak characteristic of the downstream pressure signal. Besides, a positive feedback closed-loop system associated with periodic oxidizer/fuel ratio was believed to sustain the combustion instability. The oscillation can be maintained when pressure, heat release and oxidizer/fuel ratio are coupled together. The analysis results indicate that rotating detonation is an implication to tangential combustion instability.     

Jet engine electrification as a phenomenon reflecting the evolution of charge instability in an outflowing combustion product mixture

It is hypothesized that charge instabilities developing in a mixture of outflowing combustion products make some contribution to the mechanism of engine electrification in jet engines. Investigations are carried out which support this hypothesis. Zh. Tekh. Fiz. 67, 21–26 (August 1997)     

Performance enhancement of a pulse detonation rocket engine

Utilizing liquid kerosene as the fuel, oxygen as oxidizer and nitrogen as purge gas, a series of multi-cycle detonation experiments was conducted to improve the performance of pulse detonation rocket engine (PDRE). In order to improve the performance of the engine, it is crucial to develop an effective DDT enhancement device with less flow loss and higher survival in hostile detonation tube; therefore, three spiraling internal grooves were tested. The three spiraling internal grooves were semicircle, square and inversed-triangle grooves, respectively. The results showed that the spiraling internal groove can effectively enhance DDT and prolong the operation time of PDRE. The effect of groove shape on thrust enhancement of PDRE and the optimum length of spiraling groove were then investigated. To improve the detonability of liquid kerosene and prolong the durability of PDRE, experiments on the kerosene preheating based on active cooling were conducted. The results demonstrated that with the aid of fuel preheating, the detonation initiation time for liquid kerosene was noticeably reduced and a fully-developed detonation wave was achieved in the position away from igniter 4.67 times the diameter of the detonation tube. By adding the additive to liquid kerosene, the detonation initiation time from 0.75 ms decreased to 0.34 ms and the detonability of fuel was dramatically improved. Finally, experiments were conducted to investigate the effects of the operating frequency on the detonation parameters, the fill fraction and PDRE performance. The results indicated that detonation pressure and temperature vary with the operating frequency of PDRE, and the fill fraction has a significant influence on the specific impulse of PDRE. With the strategy of partial filling in detonation tube, the specific impulse can be remarkably enhanced.     

Features of heat transfer in a pre-nozzle volume of a solid-propellant rocket motor with charges of complex shapes

Local features of thermophysical processes in the channels and pre-nozzle volumes of solid-propellant rocket engines with case-bonded charges of different cross-sectional shapes are considered. The influence of the charge shape on the heat exchange in the nozzle bottom is investigated. It is shown that the value of the Nusselt number at a critical point on the multi-nozzle bottom is determined both by the charge channel form and by the geometry of the pre-nozzle volume. By processing the numerical experimental results the criterial dependences for determining the Nusselt number in the areas of local increase of heat exchange intensity are obtained. The obtained dependences are compared with the known empirical formulas [1–4]. It is found that the use of empirical relationships to estimate the Nusselt number leads to incorrect determination of the parameters of heat transfer on the armored surfaces of the charge, the nozzle covers, and the input parts of the submerged rotating nozzle.     

Effect of combustion chamber insulation on the performance of a low heat rejection diesel engine with exhaust heat recovery

A computer simulation of the turbocharged turbocompound diesel engine system is used to study the effect of combustion chamber insulation on the performance of low heat rejection system configurations with exhaust heat recovery. The analysis is carried out for zirconia coatings of various thickness applied on the cylinder head and piston. It is found that an intercooled turbocompound engine derives a modest thermal efficiency benefit from insulation, e.g. 4.3% improvement at a 60% reduction in heat loss. The addition of Rankine compounding can improve the thermal efficiency of the turbocompound engine by 10–14%, depending on the level of insulation and the system configuration. Furthermore, Rankine compounding can make the otherwise inferior performance of a non-intercooled engine match the performance of an intercooled engine. Finally, use of an insulating material of low conductivity and low heat capacity can increase the thermal efficiency benefits, but at the expense of increased component thermal loading.     

Fishbone-like instability in a looped-tube thermoacoustic engine

Quasi-periodic bursts of acoustic oscillations were observed during the start-up process in a looped-tube thermoacoustic engine. The acoustic oscillations have a constant frequency of 111 Hz, while the bursts have "quasi-periods" in the order of 14-25 s. The quasi-periodic bursts show a new mode of amplitude growth in this thermoacoustic engine. The envelope of the acoustic oscillations has a fishbone-like shape. The nature of the observed fishbone-like instabilities suggests a strong interaction between the acoustic and temperature field.     

Numerical simulation of a methane-oxygen rotating detonation rocket engine

The rotating detonation engine (RDE) is an important realization of pressure gain combustion for rocket applications. The RDE system is characterized by a highly unsteady flow field, with multiple reflected pressure waves following detonation and an entrainment of partially-burnt gases in the post-detonation region. While experimental efforts have provided macroscopic properties of RDE operation, limited accessibility for optical and flow-field diagnostic equipment constrain the understanding of mechanisms that lend to wave stability, controllability, and sustainability. To this end, high-fidelity numerical simulations of a methane-oxygen rotating detonation rocket engine (RDRE) with an impinging discrete injection scheme are performed to provide detailed insight into the detonation and mixing physics and anomalous behavior within the system. Two primary detonation waves reside at a standoff distance from the base of the channel, with peak detonation heat release at approximately 10 mm from the injection plane. The high plenum pressures and micro-nozzle injector geometry contribute to fairly stiff injectors that are minimally affected by the passing detonation wave. There is no large scale circulation observed in the reactant mixing region, and the fuel distribution is asymmetric with a rich mixture attached to the inner wall of the annulus. The detonation waves’ strengths spatially fluctuate, with large variations in local wave speed and flow compression. The flow field is characterized by parasitic combustion of the fresh reactant mixture as well as post-detonation deflagration of residual gases. By the exit plane of the RDRE, approximately 95.7% of the fuel has been consumed. In this work, a detailed statistical analysis of the interaction between mixing and detonation is presented. The results highlight the merit of high-fidelity numerical studies in investigating an RDRE system and the outcomes may be used to improve its performance.     

Analysis of combustion chamber insulation effects on the performance and exhaust emissions of a DI diesel engine using a multi-zone model

A comprehensive two-dimensional multi-zone model of a diesel engine cycle is presented in this study, in order to examine the influence of insulating the combustion chamber on the performance and exhaust pollutants emissions of a naturally-aspirated, direct injection (DI), four-stroke, water-cooled diesel engine. The heat insulation is taken into account by the corresponding rise of wall temperature, since this is the final result of insulation useful for the study. It is found that there is no remarkable improvement of engine efficiency, since the decrease of volumetric efficiency has a greater influence on it than the decrease of heat loss to the coolant, which is converted mainly to exhaust gas enthalpy (significant rise of the exhaust gas temperature). As far as the concentration of exhaust pollutant emissions is concerned, it is found that the rising heat insulation leads to a significant increase of the exhaust nitric oxide (NO) and to a moderate increase of the exhaust soot concentration. Plots of temperature, equivalence ratio, NO and soot distributions at various instants of time inside the combustion chamber, emanating from the application of the multi-zone model, aid the correct interpretation of the insulation effects gaining insight into the underlying mechanisms involved.