Numerical study of gas-solid reactive flow in a cylinder

A theoretical model is described for the numerical simulation of unsteady gas-solid reactive flow in packed beds of granular propellants in a tubular geometry. The approach couples continuity, momentum and energy equations in each phase along with porosity and granular stresses. Propellant bed combustion process is neglected, but the mass flow rate due to the combustion is derived using a pressure-based burning rate correlation. As the burning of solid propellant begins, pressure, temperature, density and other gas parameters begin to change rapidly. To catch these changes, a CFD approach with explicit McCormack method is used to solve the coupled system of equations. A moving mesh is used to consider the moving boundary. Pressure history, moving boundary velocity and other parameters of the mixture are obtained and compared with other numerical data.     

A study on the characteristics of combustion variations of compression ignition engine

This paper describes the results of a study of the variation of combustion characteristics in a precombusion chamber type water-cooled diesel engine. Statiscal analysis on cycle-by-cycle variation of combustion characteristics such as rate of pressure rise, heat release rate, and mass burning rate from combustion pressure-crank angle data of one thousand cycles were made under several operating conditions. The influence of engine speed and coolant temperature upon maximum frequency of combustion characteristics are discussed also.     

淹没水射流冲蚀性能试验研究

试验研究了靶距、射流压力，冲蚀时间及淹没水深对射流性能的影响关系，对淹没水射流冲蚀性能作了全面分析，并将试验结果进一步应用于清洗工程中，取得了较好效果。     

Flow instability due to cryogenic cavitation in the downstream of orifice

Flow instability in LRE (liquid rocket engine) occurs due to various reasons such as flow interactions with valve, orifice and venturi, etc. The inception of cavitation, especially in the propellant feeding system, is the primary cause of mass and pressure oscillations because of the cyclic formation and depletion of cavitation. Meanwhile, the main propellant in a liquid rocket engine is the cryogenic fluid, which properties are very sensitive to temperature variation. And the change of propellant properties to temperature variation by thermodynamic effect needs to be properly taken into account in the flow analysis in order to understand basic mechanisms for cryogenic cavitation. The present study focuses on the formation of cryogenic cavitation by using the IDM model suggested by Shyy and coworkers. The flow instability was also numerically investigated in the downstream of orifice with a developed numerical code. Calculation results show that cryogenic cavitation can be a primary source of flow instability, leading to mass fluctuations accompanied by pressure oscillations. The prediction of cavitation in cryogenic fluid is of vital importance in designing a feeding system of an LRE. This paper was recommended for publication in revised form by Associate Editor Jun Sang Park Changjin Lee received his B.S. and M.S. degrees in Aeronautical Engineering from Seoul National University in 1983 and 1985. He then went on to receive his Ph.D. degree from University of Illinois at Urbana- Champaign in 1992. Dr. Lee is currently a Professor at the department of Aerospace Engineering at Konkuk University in SEOUL, Korea. His research interests are in the area of combustion instabilities of hybrid, liquid rocket and jet propulsions. Tae-Seong Roh received his B.S. and M.S. degrees in Aeronautical Engineering from Seoul National University in 1984 and 1986. He then went on to receive his Ph.D. degree from Pennsylvania State University in 1995. Dr. Roh is currently a Professor at the department of Aerospace Engineering at Inha University in Incheon, Korea. His research interests are in the area of combustion instabilities, rocket and jet propulsions, interior ballistics, and gas turbine engine defect diagnostics.     

隔板应用于发动机燃烧室不稳定燃烧抑制中的二维数值计算分析

针对某水下航行器的燃烧室,用有限元的方法对隔板抑制燃烧不稳定性的声学机理进行了初步研究。利用ANSYS软件数值计算了隔板长度变化以及隔板数目变化时,对应的燃烧室声模态,并比较了它们的变化对燃烧室声模态造成的影响。计算表明:隔板抑制燃烧不稳定主要是通过对燃烧室腔体的声学共振频率进行平移,避开了与其相耦合的频率,从而达到抑制燃烧不稳定的目的。同时随着隔板长度的增加,靠近隔板处的声腔部分的切向声模态轴向化会逐步加剧,这样就可以得到更好的抑制作用。     

A steady-state combustion modelling of composite solid propellants

By depicting the transfer of heat and combustion reaction to take place within thin gas layers close to the propellant surface burning in a steady-state fashion, a mathematical equation has been deduced to describe the burning rate of solid propellant as a function of initial grain temperature and chamber pressure. It has been also assumed that chemical reaction could take place in premixing-diffusing zone but were carried out mainly in the reaction-flame zone. All these phenomena taken place in each zone of combustion have been assumed to be steady-state. In the present investigation, the equation, γ=k·(1/R(T _i +C))ⁿ. exp (-E _a/R(T _i +C)(P/z) is being presented and it is compared with experimental data. The proposed model has been tested and evaluated vis-a-vis strand burner data for three different propellants based on CTPB, and it has been found that the deviation of the computed burning rates from the measured rates ranged up to 2%.     

含ACP的无烟改性双基推进剂燃烧特性研究

快燃物ACP作为一种有效含能助剂用来提高推进剂的燃速具有一定的效果。通过DSC和TG-DTG研究了ACP对无烟改性双基推进剂(CMDB)热分解特性的影响,采用靶线法研究了不同含量的ACP对无烟改性双基推进剂燃速和燃速压力指数的影响,用燃烧火焰单幅照相技术和微热电偶测温获得了含ACP无烟改性双基推进剂在稳态燃烧条件下的火焰结构和燃烧波温度分布,分析了该推进剂中主要组分对燃烧性能的影响。结果表明:含ACP推进剂的燃速随ACP含量的增加而增大,推进剂燃烧火焰结构随ACP含量增加而更加明亮。     

Transient analysis of hybrid rocket combustion by the zeldovich-novozhilov method

Hybrid rocket combustion has a manifestation of stable response to the perturbations compared to solid propellant combustion. Recently, it has revealed that the low frequency combustion instability about 10 Hz was occurred mainly due to thermal inertia of solid fuel. In this paper. the combustion response function was theoretically derived by use of ZN (Zeldovich-Novozhilov) method. The result with HTPB/LOX combination showed a quite good agreement in response function with previous works and could predict the low frequency oscillations with a peak around 10 Hz which was observed experimentally. Also, it was found that the amplification region in the frequency domain is independent of the regression rate exponentn but showed the dependence of activation energy. Moreover, the response function has shown that the hybrid combustion system was stable due to negative heat release of solid fuel for vaporization, even though the addition of energetic ingredients such as AP and Al could lead to increase heat release at the fuel surface.     

固体火箭推进剂燃烧微粒场的激光全息诊断

大幅度提高金属燃料的含量是实现高能高效推进剂配方的一个发展方向。现针对含铝量2／%,粒度及含铝量6%,粒度15±2μm的两种推进剂展开了一系列的研究工作。应用离轴4F全息技术,进行了固体火箭推进剂燃烧流场的粒子分布研究。在常压、2MPa和4MPa实验工况下,获得了清晰的全息图。     

Response of radiation driven transient burning of AP and HMX using flame modeling

The radiation driven response function (R _q ) for AP and HMX propellant was obtained and compared with experimental results by using a simple αβγ flame model rather than with detailed chemistry. For an AP propellant, the profile of heat release was assumed by the experimental data. The calculatedR _q shows a frequency shift of the peak amplitude to the higher frequency and a decrease in the maximum amplitude as radiation increases. In addition, it was found the increase in the total flux could enhance the mean burning rate[`(r)]_b\bar r_b while the phase differences between the radiation and resulting conduction could consequently reduce the fluctuation amplitude Δy _b . Fortunately, this is the qualitative duplication of the behavior recently observed in the experiments of RDX propellants. For HMX, the response functionR _q has been calculated and showed a quite good agreement with the experimental data. Even though the fairly good agreement ofR _q with experimental ones, the unsteady behavior of HMX was not reproduced as the radiation input increased. This is due to lack of the material properties of HMX or the physical understanding of HMX burning at high pressure.     

Investigation of self-excited combustion instabilities in two different combustion systems

The objective of this paper is to characterize dynamic pressure traces measured at self-excited combustion instabilities occurring in two combustion systems of different hardware. One system is a model lean premixed gas turbine combustor and the other a fullscale bipropellant liquid rocket thrust chamber. It is commonly observed in both systems that low frequency waves at around 300Hz are first excited at the onset of combustion instabilities and after a short duration, the instability mode becomes coupled to the resonant acoustic modes of the combustion chamber, the first longitudinal mode for the lean premixed combustor and the first tangential mode for the rocket thrust chamber. Low frequency waves seem to get excited at first since flame shows the higher heat release response on the lower frequency perturbations with the smaller phase differences between heat release and pressure fluctuations. Nonlinear time series analysis of pressure traces reveals that even stable combustion might have chaotic behavior with the positive maximum Lyapunov exponent. Also, pressure fluctuations under combustion instabilities reach a limit cycle or quasi-periodic oscillations at the very similar run conditions, which manifest that a self-excited high frequency instability has strong nonlinear characteristics.     

Controlling plant dynamics change characteristics in thrust response of hybrid rocket engine

The hybrid rocket engine (HRE) has attracted much attention as one of the promising engines for the propulsive subsystem of a VTVL. A recent study reported that the thrust response in HRE exhibits different delay curves depending on the direction of thrust modulation, which is a typical example of system nonlinearity called as plant dynamics change (PDC). A series of step response tests was conducted to investigate the PDC occurrence and its nonlinear characteristics in HRE combustion by measuring response delays of various components. Test results confirmed that thrust response delay varied depending on the thrust direction and thrust range, and thus PDC of different characteristics occurred. In addition, results showed that the hysteresis of the chamber response is the critical factor related to the development of PDC. To improve the thrust control accuracy of the VTVL system simply by reducing the PDC in HRE thrust control, advanced PID control method with gain scheduling was used. In the result, the size of the integrated absolute error (IAE) was successfully reduced. In addition, the results confirmed that PID with gain scheduling can be an effective solution to improve the thrust control accuracy of VTVL using a hybrid rocket engine.

     

A study on the unsteady temperature field of combustion chamber wall in a turbocharged gasoline engine

To design and develop a turbocharged engine, it is necessary that a lot of studies should be done to find the chrracteristics of engine performance and thermal flow. To accomplish this purpose, turbocharger equipped to a naturally aspirated gasoline engine was utilized. A thin-film type temperature probe was made and installed onto the combustion chamber wall to measure unsteady temperature. The unsteady heat flux at combustion chamber wall was evaluated by one dimensional unsteady conduction equation.     

Pulsed illumination, closed circuit television system for real-time viewing of unsteady (> 1 micros) events

A pulsed illumination closed circuit television system is described whereby fast (times <33 ms), unsteady events can be observed in real time. A low-power helium-neon laser beam is modulated to send a short duration light pulse through the unsteady test medium. The light is refracted according to the instantaneous optical properties of the medium. The refracted light travels to a solid state television camera, known as a charge injection device (CID), in which the sensor array is charged within microseconds. The scanning of the charged array then follows, requiring the standard 33 ms for information transfer to video tape and a TV monitor. The image is thus formed during the laser pulse duration (which presently is 10 to 100 micros, but shorter duration pulses are possible with more powerful lasers), but no more than one image every 33 ms can be observed and recorded. Thus this method is particularly suited for the investigation of high frequency periodic events in which one can observe both a single image, or an ensemble average of as many as 100 images, occurring at corresponding times in different cycles. The reported applications include the recording of steady and transient propane torch flames, of the transient fuel injection process in a motored internal combustion engine, and of the propagation of a flame under firing conditions in the engine. In the shadowgraph and Schlieren modes the method is particularly suited for application to periodic combustion events such as those occurring in internal combustion engines. The method then presents the following advantages over high-speed filming (> 3000 pictures/s); real-time observation and recording of chamber events at any crankangle; real-time observation and recording of the effects of changes in the engine variables (speed, load, spark timing, injection pressure and duration, chamber swirl, etc.) on the combustion events; real-time observation and recording of ensemble averages and cycle-to-cycle variations. The technique also eliminates the delays and unknowns of film processing. Finally, the cost of this system is similar to that of a high-speed camera.     

固体火箭发动机抗烧蚀防热涂层的研究

直径及开口都较大的固体火箭发动机燃烧室大都采用橡胶基绝热层,但是,对于长细比大,或者是开口较小的客体,采用橡胶基绝热层在工艺上难以实现.为了寻找同时具备隔热效果良好,且针对小口径壳体工艺可行性高的内防热材料,我们开展了以环氧树脂、橡胶为基体,云母粉等耐高温无机填料组成的防热材料.通过试件测试及产品验证,证明该防热涂层也是固发燃烧室一种较为适宜的烧蚀防热材料,它不受被保护产品的几何形状限制,烧蚀率较小.     

Experimental study of the role of gas-liquid scheme injector as an acoustic resonator in a combustion chamber

In a liquid rocket engine, the role of gas-liquid scheme injector as an acoustic resonator or absorber is studied experimentally for combustion stability by adopting linear acoustic test. The acoustic-pressure signals or responses from the chamber are monitored by acoustic amplitude. Acoustic behavior in a rocket combustor with a single injector is investigated and the acoustic-damping effect of the injector is evaluated for cold condition by the quantitative parameter of damping factor as a function of injector length. From the experimental data, it is found that the injector can play a significant role in acoustic damping when it is tuned finely. The optimum tuning-length of the injector to maximize the damping capacity is near half of a full wavelength of the first longitudinal overtone mode traveling in the injector with the acoustic frequency intended for damping in the chamber. When the injector has large diameter, the phenomenon of the mode split is observed near the optimum injector length and thereby, the acoustic-damping effect of the tuned injectors can be degraded.     

Simulation of combustion in a porous-medium diesel engine

The future internal-combustion (IC) engines should have minimum emissions level under lowest feasible fuel consumption. This aim can be achievable with a homogeneous combustion process in diesel engines. We used a porous medium (PM) to homogenize the combustion process. This research studies simulation of a direct-injection diesel engine, equipped with a chemically inert hemispherical PM. Methane is injected into a hot PM, assuming mounted up the cylinder in head. Combustion with lean mixture occurs inside PM. A numerical model of PM engine was carried out using a modified version of the KIVA-3V code. PM results were evaluated with experimental data of unsteady combustion-wave of methane in a porous tube. The results show the mass fraction of methane, CO, NO and temperature in solid and gas phases of the PM and in-cylinder fluid. Also presented are the effects of injection timing and compression ratio on combustion.

     

Performance and emission characteristics of a low heat rejection spark ignited engine fuelled with E20

In internal combustion engines, the concept of low heat rejection (LHR) using thermal barrier coating on the surface of combustion chamber is gaining attention. Thermal barrier coating reduces the heat transfer to the cooling system, protects engine components from peak heat flux and fluctuating temperature produced during combustion and improves the performance of the engine. Information in the literature is plentiful for LHR diesel engine and only few studies exist on LHR spark ignited engine. The application of thermal barrier coating in spark ignited engine is limited by pre-ignition and knocking due to elevated combustion chamber temperature. A spark ignited engine with moderate insulation on the combustion chamber and higher octane fuel can overcome this difficulty. The objective of the present experimental study is to quantify the changes in performance and emission characteristics brought by partial thermal insulation on the combustion chamber of a four stroke spark ignited engine fueled with E20 blend. Partial thermal insulation was created by coating 0.3 mm thick Alumina (Al₂O₃) on the cylinder head, inlet and exhaust valves. The changes are quantified with respect to unmodified engine fueled with gasoline. The combustion parameters such as flame development and rapid burn duration are also estimated and compared. The results indicate that partially insulated SI engine when fueled with E20 improves performance and reduces emission. A maximum of 48% reduction in THC and 50% reduction in CO emission at part load was achieved.

     

Self-Ignition of Fuel Consisting of Hydrogen Peroxide,Kerosene, and Pyrophoric Additives

Abstract—The possibility of creating an environmentally safe self-igniting fuel mixture for liquid-propellant rocket engines is investigated. The mixture is based on kerosene with pyrophoric additives and an oxidant: highly concentrated hydrogen-peroxide solution. The self-ignition of the components is studied experimentally. Tests are also conducted in the combustion chamber of a low-thrust liquid-propellant rocket engine.     

Hot-fire injector test for determination of combustion stability boundaries using model chamber

This study realizes the conceptual method to predict combustion instability in actual full-scale combustion chamber of rocket engines by experimental tests with model (sub-scale) chamber. The model chamber was designed based on the methodologies proposed in the previous work regarding geometrical dimensions and operating conditions, and hot-fire test procedures were followed to obtain stability boundaries. From the experimental tests, two instability regions are presented by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio, which are customary for combustor designers. It is found that instability characteristics in the chamber with the adopted jet injectors can be explained by the correlation between the characteristic burning or mixing time and the characteristic acoustic time. In each instability region, dynamic behaviors of flames are investigated to verify the hydrodynamically-derived characteristic lengths of the jet injectors. Large-amplitude pressure oscillation observed in upper instability region is found to be generated by lifted-off flames.