共查询到20条相似文献,搜索用时 93 毫秒
1.
2.
3.
《固体火箭技术》2021,44(2)
在马赫数2.0,总压0.98 MPa和总温920 K的超声速来流条件下,针对现有常见的凹腔组合式燃料喷注方案出现的燃烧不稳定和火焰吹熄现象,通过改变凹腔上游壁面双路燃料喷注的位置,设计了两种优化的凹腔组合式喷注方案,并对不同燃料喷注方案下的火焰稳定过程进行研究。通过高速摄影和CH*基自发辐射成像技术,详细观测了后缘突扩凹腔燃烧室中乙烯火焰传播过程。研究表明,原始的喷注方案容易发生火焰振荡,并伴随着火焰回传现象以燃烧模式的转换;当量比超过0.3时,就难以实现稳定燃烧,并出现火焰吹熄现象。两种改进的喷注方案均能增强燃料射流与凹腔的相互作用,可在更宽燃料喷注当量比范围内维持火焰不被吹熄。相比于增加上游喷注与凹腔前缘距离的喷注方案而言,增加双路燃料喷注之间距离的喷注方案的稳焰效果更好,燃烧反应区也更加靠近凹腔前缘,燃烧释热也更强。这种喷注方案可为超燃冲压发动机燃烧室中凹腔燃料喷注方案的优化设计提供参考。 相似文献
4.
针对高超声速飞行器巨大的激波阻力,采用数值方法研究了由钝头体、气动杆和侧向喷流构成的组合模型的减阻性能。侧向喷流将弓形激波推离气动杆,组合模型的再附激波明显弱于传统气动杆模型,其阻力系数比气动杆模型低了33.52%,从而验证了本文组合模型优异的减阻效率。进行了组合模型的影响因素分析,随侧向喷流总压比和气动杆的长度的增加,再附激波强度减弱,减阻效率升高,但减阻效率的变化速率逐渐减小。随喷口位置向下游移动,再附激波逐渐增强,减阻效率降低,且减阻效率的变化速率逐渐增加。此外本文还研究了以上参数对流场结构及钝头体压力峰值位置的影响。 相似文献
5.
6.
7.
针对发动机燃气喷流对底部流动的影响开展研究。建立冷喷与热喷计算方法,与经典的高压空气尾喷管喷流试验数据进行了对比,验证了本文建立的三维喷流方法的可靠性。对本文选用的飞行器外形采用冷喷与热喷方法开展了对比计算并与飞行试验值进行比较,分析了两种方法结果的差异。采用热喷方法对来流马赫数 2.5 ,不同飞行高度及喷管进口总压开展计算,研究飞行高度及喷管进口总压对发动机喷流及底部流场的影响。结果表明,保持飞行高度、来流马赫数不变,喷管进口总压增加,底部压力系数逐渐提高。燃气质量浓度最大值位于底部空腔的壁面处,且保持一个恒定值。保持喷管进口总压、来流马赫数不变,飞行高度增加,喷流高速区向后移动且中心区最大马赫数增加。在一定飞行高度下,底部压力系数由负转正,即飞行器底部会出现正推力,这对飞行器的射程会产生重要影响,需要提前评估。 相似文献
8.
9.
对带长深比为10的凹腔结构的燃烧室二维氢燃烧流场进行数值模拟,燃料喷注方式采用凹腔上游喷注加辅加凹腔前壁、底壁、后壁喷注。采用三阶MUSCL格式求解二维含组分守恒N-S方程组,湍流模型采用剪切修正的RNGk-ε湍流模型,对喷氢燃烧工况进行了计算研究,并分别分析了凹腔中不同燃料喷注方式对燃烧特性的影响。结果表明:凹腔是火焰驻留的主要区域;凹腔上游喷注氢,可以使燃料在凹腔中混合燃烧,辅加凹腔中喷氢的三种方式对燃烧状况产生一定的影响。在凹腔前壁、底面辅加喷氢,没有增强凹腔的稳焰特性,对整个燃烧状态影响不大;在凹腔后壁喷氢,能够增加凹腔中的燃料含量,加强了回流效果,对燃烧状态影响较大。三种喷注方式都没有从根本上改变凹腔燃烧流场的特性。 相似文献
10.
水环境下喷管流动分离数值研究 总被引:1,自引:0,他引:1
《固体火箭技术》2020,(1)
为了研究水环境下发动机喷管流动分离现象以及影响因素和规律,基于VOF多相流模型和SST k-ω湍流模型,建立了水环境下固体火箭发动机喷流流场数值仿真模型,并进行了不同喷管扩张比和NPR(燃烧室总压与环境压强之比)下的喷流流场数值模拟。通过数值仿真分析获得了水环境下喷管内发生流动分离时推力、压力特征和流场非定常变化特征,水环境下喷管内流动分离具有强烈的非定常振荡特征,分离激波会在分离点与发动机喷管出口之间呈现推进-返回-推进周期性振荡的流动特征。同时,获得了喷管扩张比和NPR对流动分离特征的影响规律,相同水深环境下不同扩张比喷管对流动分离点位置影响较小; NPR越小,流动分离点的位置处喷管扩张比越小。 相似文献
11.
为研究SMC模式下火箭混合比对RBCC发动机性能的影响规律,完成了氢/氧火箭推力室中心布局、二元定几何结构模型发动机飞行马赫数Ma0=4、高度H=17 km弹道点流场仿真,获得了不同火箭混合比(MR=2、3、4、5、6、8)及燃烧室长度的推力、比冲性能。研究表明:在火箭燃气富燃条件下(MR<8),产生了正的火箭推力增益,且随着混合比的减小,火箭推力增益增加;二次燃烧过程受火箭射流与冲压主流剪切层掺混主导,在给定的基准燃烧室长度下,燃烧效率随着混合比的提高而增加,且火箭射流与冲压主流的超/超射流剪切层燃烧过程一直持续到喷管出口;通过增加燃烧室长度,火箭富燃燃气获得更为充分的燃烧,发动机性能显著提升,但在具体发动机设计中,燃烧室长度的选取需在燃烧效率与结构惩罚之间进行权衡。 相似文献
12.
13.
以飞行马赫数为4.5Ma的RBCC发动机典型工作状态为研究背景,采用大涡模拟研究了支板火箭射流和空气来流形成的超声速反应混合层的掺混燃烧过程,获得了燃烧室内详细的流场结构和流动特征,分析了强射流条件下超声速反应混合层的特性。结果表明由于速度梯度的存在,火箭射流进入燃烧室后与空气来流形成环形剪切层,剪切层内丰富的旋涡结构主导火箭射流和空气来流的掺混燃烧,随着湍流能量的串级输运,化学反应过程中释放的能量将被转化成细观尺度的湍流动能,大尺度旋涡将能量传递给小尺度旋涡并最终耗散,细小尺度的旋涡一方面能够促进燃烧反应物的掺混并强化燃烧过程,另一方面会给化学反应过程带来强烈的脉动,使得局部火焰淬灭,火焰结构表现出明显的非定常性。 相似文献
14.
Cumulative jet formation was regarded aimed at the formation of hypervelocity fragments up to 14 km/s for the investigation of space debris effects on shielding screens. The basic physical problems of jet formation process are analyzed. It is shown that in process of realization of hyper-cumulation conditions for jet formation without complete stagnation point involving formation of the inner zone of constant pressure (dead zone), the flow mass is always greater than slug mass. That opens wide the possibilities for increasing fragment mass in laboratory tests. 相似文献
15.
The mixing process between the injectant and the supersonic crossflow is one of the important issues for the design of the scramjet engine, and the efficiency mixing has a great impact on the improvement of the combustion efficiency. A hovering vortex is formed between the separation region and the barrel shock wave, and this may be induced by the large negative density gradient. The separation region provides a good mixing area for the injectant and the subsonic boundary layer. In the current study, the transverse injection flow field with a freestream Mach number of 3.5 has been optimized by the non-dominated sorting genetic algorithm (NSGA II) coupled with the Kriging surrogate model; and the variance analysis method and the extreme difference analysis method have been employed to evaluate the values of the objective functions. The obtained results show that the jet-to-crossflow pressure ratio is the most important design variable for the transverse injection flow field, and the injectant molecular weight and the slot width should be considered for the mixing process between the injectant and the supersonic crossflow. There exists an optimal penetration height for the mixing efficiency, and its value is about 14.3 mm in the range considered in the current study. The larger penetration height provides a larger total pressure loss, and there must be a tradeoff between these two objection functions. In addition, this study demonstrates that the multi-objective design optimization method with the data mining technique can be used efficiently to explore the relationship between the design variables and the objective functions. 相似文献
16.
17.
合成射流激励器增强同向燃气-氧气掺混数值模拟及机理研究 总被引:2,自引:0,他引:2
建立了将合成射流激励器腔体、出口喉道及外部受控流场作为单连域计算处理的全流场计算模型(X L模型)。基于此计算模型,对合成射流激励器增强同向燃气 氧气掺混的流场进行了数值仿真和机理研究。研究表明,应用合成射流激励器可以显著增强同向燃气/氧气的掺混,其主要控制机理是合成射流激励器对同向燃气/氧气流起到流动方向控制作用,使两侧两股氧气平行射流向内发生偏转,从而大大缩短了每股射流的核心区长度;同时,激励器工作改变和加强了射流出口附近的涡结构,通过涡结构的强对流作用极大地增强了燃气/氧气平行射流在出口附近的混合。 相似文献
18.
《Acta Astronautica》2014,93(1):298-310
Numerical simulations were employed to analyze the flowfield of a scramjet with three-dimensional (3D) sidewall compression inlet, and the effect of inlet distortion on the mixing and combustion process was examined. The numerical approach solved the compressible Reynolds Averaged Navier–Stokes (RANS) equations supplemented with a finite rate chemical reacting model for the combustion of hydrogen fuel and air. Turbulence closure was achieved using Menter shear-stress transport (SST) model. To verify the accuracy of the simulation, the computed wall pressure was compared with the experimental data of the direct-connect combustor test. The metrics employed in the simulations included qualitative assessments related to flow structure as well as quantitative values of fuel mixing efficiency, combustion efficiency and static pressure distribution. Intake sidewalls were found to strongly affect the inlet flow structure, which became more complex in the nonuniform flowfield. The shock train system affected the combustion region located upstream of the injection and led to pairs of asymmetric separation bubbles. Nevertheless, the shock train system dissipated due to the reactions, the combustion patterns of each fuel jets in downstream region were nearly identical, and the degree of improvement of mixing and combustion efficiency near the downstream injectors was less than that near the upstream injectors. 相似文献
19.
Transverse slot injection scheme is very important for the mixing process between the air and the fuel in supersonic flows. The effect of the turbulence model and slot width on the transverse slot injection flow field has been investigated numerically based on the grid independency analysis, and the predicted results have been compared with the experimental data available in the open literature. The obtained results show that the grid scale makes only a slight difference to the wall pressure profiles for all jet-to-crossflow pressure ratios employed in this study, and the wall pressure profile with low jet-to-crossflow pressure ratio is predicted accurately by the RNG k–ε turbulence model, the SST k–ω turbulence model for the flow field with high jet-to-crossflow pressure ratio. High jet-to-crossflow pressure ratio can increase the jet penetration depth in supersonic flows, and the gradient of the length of the upstream separation region is larger than that of the height of the Mach surface. At the same time, when the jet-to-crossflow pressure ratio is maintained constant, the jet penetration depth increases with the increase of the slot width. 相似文献