Investigation of flow characteristics in supersonic combustion ramjet combustor toward improvement of combustion efficiency

Supersonic combustion ramjet (scramjet) is a variant of ramjet in which the combustion takes place at supersonic velocity. The flow physics inside the scramjet combustor is quite complex due to the fact that the mixing and completion of the combustion take place in a short time, which is of the order of milliseconds. This study focuses on flow characteristics within the combustion chamber of the scramjet engine that is designed to improve energy efficiency by enhancing combustion efficiency. The effect on combustion performance and thereby the energy efficiency on using strut‐based flame stabilizer is evaluated at different positions. Reynolds averaged Navier‐Stokes equations are solved with the Shear Stress Transport k‐ω turbulence model. Single strut configuration is used to validate with the experimental data. Single strut is then compared with three‐strut configuration. In the three‐strut configuration, the location of the primary strut is kept constant, and the secondary struts are relocated in x and y directions. Combustion performance was evaluated for the cases of flow from primary strut only and through three struts. It was found that the placement of secondary strut in a three‐strut configuration plays a vital role in improving energy efficiency. A maximum of 33.86% improvement in combustion efficiency was observed in comparison to the single strut combustor. A reduction in unburned fuel was observed, making the system more energy efficient. If the struts are not placed optimally, the combustion performance of the combustor was observed to be lower than that of a single‐strut configuration. The shock reflection and expansion fans within the primary combustion zone and the secondary strut region enhance the combustion efficiency. The wall static pressure was observed to increase with the addition of secondary struts. For certain strut configurations, flow separation was seen on the combustor walls. If the secondary strut was placed close to the primary strut, combustion efficiency was found to enhance. It was seen that combustion efficiency was also enhanced for the cases of reacting flow from primary strut only. It could also help to increase fuel efficiency, as additional fuel is not supplied to the secondary strut, making the overall system energy efficient. As the secondary strut is introduced, total pressure loss also increases. It could also be noted that if the combustor length was increased, there could be further increased in combustion efficiency.     

超燃冲压发动机喷焰辐射试验研究

介绍了超燃冲压发动机喷焰辐射特性试验与测量方法,在燃料为航空煤油、来流马赫数为6、当量油气比为0.8的试验状态下,获得了长1 m量级超燃冲压发动机喷焰辐射强度和光谱。喷焰辐射强度在发动机喷口附近随喷射距离缓慢下降;辐射强度显现出的高频震荡、低频稳定特性,反映出燃料燃烧过程的微观脉动性和宏观稳定性。根据喷焰光谱,喷焰辐射能量集中在波长4.364 3 μm附近,主要辐射源为CO₂.     

超燃冲压发动机前体/进气道和隔离段气动设计

采用等激波角设计方法并考虑温度、激波与附面层干扰等的影响，对超音速燃烧冲压发动机的二维混压式高超音速前体／进气道和隔离段的设计进行了探索，给出了前体／进气道和隔离段的几何结构和尺寸。运用二维CFD数值计算手段，对所设计进气道结构进行了修正，并计算了设计状态和非设计状态性能和流场。研究表明，文中所设计的进气道结构简单、附加阻力较小、总压恢复系数较高，所给出的设计方法对于前体／进气道和隔离段的初步设计具有较好的适用性。     

高超声速临近空间飞行器跳跃飞行轨迹优化

以高超声速临近空间飞行器的连续跳跃飞行为研究对象,建立了其纵平面运动模型和弹道优化模型,在考虑攻角值、攻角变化率、法向过载、动压、热流及落地条件等约束下,利用遗传算法对一个飞行周期的跳跃飞行轨迹进行优化,并给出了燃料消耗率最小的飞行轨迹.最后与常值稳态巡航飞行进行比较分析.结果表明,连续跳跃飞行可以大量减少燃料消耗.     

凹腔对含硼固体火箭超燃冲压燃烧特性的影响

凹腔常用来增强超燃冲压发动机中空气与燃气掺混、提升火焰稳定性及燃烧效率,然而超音速燃烧室内的燃气流场特性依赖于凹腔结构及其分布。为优化凹腔结构及其分布,提升固体火箭超燃冲压发动机补燃室内的燃烧性能,采用数值方法计算分析凹腔长深比、后倾斜角对含硼固体火箭超燃冲压发动机燃烧特性的影响。计算结果表明：在凹腔长度不变时,取凹腔长深比分别为5.00、3.75、3.00、2.50、2.18、1.85、1.67,硼颗粒燃烧效率与比冲随着长深比减小先增大、后减小,在长深比为1.85时最大;当凹腔长深比为1.85、凹腔后倾角从90°变化到175°时,随着凹腔后倾角增加,硼颗粒燃烧效率增加,175°时燃烧效率最大,但其总压恢复系数及比冲最小,比冲在165°时最高。     

超音速气流中煤油蒸发与掺混特性研究

数值模拟了超燃冲压发动机模型燃烧室中煤油在空气中的蒸发与掺混特性,分别计算了煤油温度、煤油液滴直径、雾化角等射流参数的影响;结果表明：煤油温度越高,雾化颗粒直径越小,则颗粒蒸发成气相所需的时间越短,相应的浓度就越高,雾化角的变化对蒸发和掺混的效果不明显;模拟结果对超燃冲压发动机煤油雾化和掺混设计具有一定的指导意义.     

超燃冲压发动机进气道内激波/边界层干扰研究

外界存在多种因素引起隔离段反压增加,将直接影响到进气道内波系的结构和强度,进而影响到因激波诱导而分离的边界层情况．在设计状态下,建立二维的进气道模型,运用RSM湍流模型,对单个超燃冲压发动机进气道隔离段内复杂流场进行数值模拟．得到对应不同反压时,超燃冲压发动机进气道的激波与边界层相互作用情况．研究发现,反压为0时,隔离段内分离区较大,出口马赫数较低,尾部波系较弱,为燃烧室稳定燃烧提供有利条件．随着反压的增加,即使总压恢复系数有所增加,但隔离段尾部的气流稳定性降低．因此,要合理调节隔离段反压大小,以便更好地协调总压损失和尾部气流稳定性的矛盾,对于进气道工作最优化的控制起到一定的参考作用．     

Observer‐based fault‐tolerant control of hypersonic scramjet vehicles in the presence of actuator faults and saturation

An adaptive sliding mode observer (SMO)–based fault‐tolerant control method taking into consideration of actuator saturation is proposed for a hypersonic scramjet vehicle (HSV) under a class of time‐varying actuator faults. The SMO is designed to robustly estimate the HSV states and reconstruct the fault signals. The adaptive technique is integrated into the SMO to approximate the unknown bounds of system uncertainties, actuator faults, and estimation errors. The robust SMO synthesis condition, which can be formulated as a set of linear matrix inequalities, is improved by relaxing structure constraints to the Lyapunov matrix. An anti‐windup feedback control law, which utilizes the estimated HSV states and the fault signals, is designed to counteract the negative effects of actuator saturation induced by actuator faults. Simulation results demonstrate that the proposed approach can guarantee stability and maintain performance of the closed‐loop system in the presence of HSV actuator faults and saturation.     

Numerical simulation of the interaction between shock train and combustion in three-dimensional M12-02 scramjet model

The characteristics of combustion flow fields and performance for hypersonic M12-02 scramjet were numerically simulated and analyzed. The compressible two-equation k-w SST turbulence model was employed for the turbulence model and the 9-species, 27-reaction-step hydrogen-air reaction mechanism was used as the reaction kinetics model. The numerical method was verified and a good agreement was obtained between the results of the numerical simulations and the experimental data. The results showed that shock waves from the upper and lower walls respectively crossed with each other near the central axis, forming a ‘diamond’ shape in the high-temperature combustion region. Compared to the conventional scramjet engine, most of the fuel reaction was in pure supersonic combustion mode for this hypersonic scramjet engine. Changes in the distribution of fuel on the upper and lower walls could have an appreciable impact on the combustion flow field. Average fuel distribution between upper and lower walls is benefit for combustion enhancement while the heat transfer in the corner of the side wall is severe and should be avoided during operation. The flame investigation showed that it cannot automatically predict the flame surface temperature in advance only based on the equivalence ratio Φ according to diffusion combustion theory. Compared to Φ = 1.0 condition, the flame surface temperature for Φ = 0.8 condition is higher as the complicated interaction between shock waves and combustion, which makes the local air temperature and mixing extent in flame surface is more appropriate. However, in terms of the overall engine performance, the Φ = 1.0 condition has the better combustion efficiency along the whole flow path.