不足热负荷由管道补燃提供。
最后提出补燃系统的控制结构方案原理图。
The control scheme for this auxiliary combustion system is presented.
该研究从理论上找到了提高补燃效率的途径。
The way improving afterburning efficiency was theoretically found.
建立了全流量补燃循环发动机的静态特性的数学模型。
The mathematical models of the static characteristics of the full-flow staged combustion cycle rocket engine were presented.
随着补燃室长度的增大，发动机推力、比冲和燃烧效率均提高。
As the length of secondary combustion chamber increases, the thrust, specific impulse and combustion efficiency increase respectively.
通过模块化编程，建立了新一代高性能补燃发动机启动过程的部件模型。
The dynamic models of engine parts in a new high-performance staged combustion cycle liquid propellant rocket through the module codified were built.
通过数值计算，得到了冲压发动机补燃室内绝热层温度分布以及烧蚀率。
The temperature contour in the insulation and the ablation rate can be obtained by the numerical simulation.
气液同轴直流离心式喷嘴对补燃发动机燃烧室高频燃烧稳定性有显著影响。
The bipropellant gas-liquid coaxial orifice-swirl injector has great effects on the high-frequency combustion instability of the LOX/kerosene staged combustion cycled engine.
补燃室内复杂的温度分布和空气与燃气的掺混、燃烧及流动状态有密切关系。
In the afterburner, the complicated temperature distribution has close relation with the dilution, combustion and flow of the air and gas.
补燃室内空气同燃气的掺混、燃烧过程较为复杂，影响补燃效率的因素很多。
Mixing and combustion process are very complex, and there are many factors affect combustion efficiency.
喷口中心线与补燃室轴线的角度增加对碰撞有催进作用，但影响效果相对较小；
The increase of the angle between centerline of nozzle orifice and axis of secondary combustor can promote collision number, but the effect is small.
首先建立研究静止增程固体火箭冲压发动机补燃室燃烧特性的物理和数学模型。
Physical and mathematic model of investigating combustion characteristics of secondary combustion chamber for static extended solid rocket ramjet was established.
所得的结论对全流量补燃循环发动机气-气喷注器的优化设计有一定的指导意义。
The results presented in this paper may provide valuable reference for the optimization design of the gas-gas injector for full flow staged combustion cycle engine.
以全流量补燃循环氢氧发动机系统为研究对象，对其的动态响应特性进行了研究。
The full flow staged combustion cycle and its superior performance compared to the traditional cycle are discussed.
因此，开展固体冲压发动机补燃室内绝热层传热烧蚀特性研究，就显得非常有意义。
Therefore, it is very important and necessary to specially research the heat transfer and ablation properties of insulation in ramjet chamber.
以采用强迫起动方式的补燃发动机涡轮泵联试系统为研究对象，建立了系统的动态模型。
The dynamic model is established for a turbo-pump test system of a staged combustion cycle engine which employs forced start.
基于闭式补燃循环液体火箭发动机流量大的特点，欲设计效率水平高的航天反力式涡轮。
Attributing to the large flow rate, high efficiency level reaction turbine is adoptive in staged combustion cycle liquid rocket engine.
应用概率密度函数非预混模型，对固体火箭冲压发动机补燃室内的湍流燃烧进行了数值模拟。
The numerical simulation of turbulent combustion in the afterburner of solid rocket ramjet was performed by applying the probability density function non -premixed mode.
研究表明，补燃室压强的变化对燃烧效率的影响是有限的，喷管损失主要与喷管结构尺寸有关；
The results show that the effect of combustion efficiency on chamber pressure is limited, nozzle loss is mainly caused by its structure and size.
基于双轴涡扇发动机热力学模型，采用高低压涡轮间补燃的方法，实现了对发动机性能的改善。
On the basis of the thermodynamics model of two-spool turbofan engine, performance of two-spool turbofan engine was improved by after-burning between high and low turbines.
提高空燃比，可增强补燃室中燃气的回流和轴向涡流强度，加大掺混力度，从而提高燃烧效率。
The intensity of the gas backflow and axial swirl ond the dilution in the afterburner are enhanced by increasing the air-to-fuel ratio, and then the combustion efficiency is improved.
本文从实验和数值模拟两方面，对静止和旋转增程固体火箭发动机补燃室燃烧流场进行了研究。
The combustion flow field in secondary combustion chamber for static and spinning extended solid rocket ramjet were studied experimentally and numerically in detail.
超声速进气道的性能优劣决定了补燃室流场内的压强和空气流量，进而影响到冲压发动机的工作性能。
The performance of supersonic inlet determines the pressure and the air flux in combustor, then affect the work of ramjet.
混合发动机燃烧效率比较低，使用预燃室和补燃室、合理设计喷注器以及采用扰流装置，可以提高燃烧效率；
The combustion efficiency of hybrid rocket motor is very low. In order to increase combustion efficiency, we can use vaporization chamber and mixing chamber.
硼颗粒在通过固体火箭冲压发动机的燃气发生器喷管进入补燃室的过程中，颗粒通常都会受到高速气流的作用。
The effects of various parameters on boron particle ignition of ducted rocket secondary chamber were studied with the model developed by the King.
考虑了进气道尺寸、进气夹角、头部距离、一次喷嘴扩张比和喷嘴数目等结构因素对补燃室掺混与燃烧性能的影响。
The effect of inlet size, inlet Angle, dome height, primary nozzle expansion ratio and nozzle Numbers on combustion efficiency were taken into account.
研究还发现，硼颗粒的点火燃烧主要取决于硼颗粒在补燃室内的滞留时间和所处的环境温度，而与补燃室的长度没有直接关系。
The study also shows that the ignition and combustion of boron particle lie on the resorting time and the environment temperature, and have no direct relationship with after-burning chamber's length.
为了分析发动机设计参数对反应流场的影响，用不同的条件进行计算，并由此分析了补燃室几何结构和液体燃料初始颗粒直径对燃烧效率的影响。
Calculations were carried out under different chamber configurations and initial droplet diameters, from which the effects of these conditions on the combustion efficiency were analyzed.
为了分析发动机设计参数对反应流场的影响，用不同的条件进行计算，并由此分析了补燃室几何结构和液体燃料初始颗粒直径对燃烧效率的影响。
Calculations were carried out under different chamber configurations and initial droplet diameters, from which the effects of these conditions on the combustion efficiency were analyzed.
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