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燃烧效率
相关语句
  combustion efficiency
    THE EXPERIMENTAL INVESTIGATION OF COMBUSTION EFFICIENCY OF INTEGRAL ROCKET/RAMJET
    固体火箭—冲压组合发动机燃烧效率的实验研究
短句来源
    THE EFFECT OF DIRECT HEATING ON COMBUSTION EFFICIENCY OF THE CHAMBER OF JET ENGINE IN THE SIMULATION TEST ON THE GROUND
    直接加热对空气喷气发动机燃烧室地面模拟试验燃烧效率的影响研究
短句来源
    A THEORY OF COMBUSTION EFFICIENCY COMPUTATION FOR LIQUID ROCKET ENGINE COMBUSTOR
    液体火箭发动机燃烧室燃烧效率计算理论
短句来源
    A STUDY OF COMBUSTION EFFICIENCY FOR LIQUID ROCKET ENGINE WITH VARIABLE THRUST
    变推力液体火箭发动机燃烧效率的研究
短句来源
    The Flow Field and Combustion Efficiency Characteristics of Afterburner with Numerical Method
    数值分析加力燃烧室的燃烧效率和燃烧流场
短句来源
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  burning efficiency
    The analysis of combustion residues shows that the burning efficiency is obviously improved after coated by PBT,and the amount of substance proportion between B 2O 3 and B increases from 122.7 to 18.51.
    燃烧残渣分析结果表明,经包覆后的B的燃烧效率明显提高,残渣中B2O3和B的摩尔比由包覆前的122.7增加到18.51。
短句来源
    The burning efficiency of B in propellant was obviously improved after coated by LiF, and the amount of substance proportion between B and B_2O_3 decreases from 37.5∶1 to 3.1∶1.
    包覆后推进剂中硼的燃烧效率明显提高,推进剂燃烧残渣中硼与B2O3的摩尔比由包覆前的37.5∶1变为包覆后的3.1∶1。
短句来源
    It has advantages in intensifying burning, increasing burning efficiency and reducing pollution.
    在声能作用下的煤炭燃烧特性不仅燃烧效率和燃烧强度高,而且燃烧烟气中污染物含量低,有利于大气环境保护.
短句来源
  “燃烧效率”译为未确定词的双语例句
    Analysis of Combustion stability and efficiency of FWS-9 Under the Condition of High Altitude and Low Velocity
    某型发动机高空、低速条件下燃烧稳定性及燃烧效率分析
短句来源
    The paper also discusses size scale-up studies of hybrid rocket motors and the numerical simulation of two-phase hybrid combustion.
    本文还基于颗粒轨道模型建立了两相混合燃烧模型,并通过数值模拟研究了液氧颗粒尺寸和蒸发过程对两相混合燃烧效率的影响。
短句来源
    By applying these sets,this combustor can easily excite the pulsating combustion. The frequency and the most amplitude of pulsating pressure were 100 Hz ̄110 Hz,3 8 kPa.
    试验表明,采取以上措施,改型瑞克管可以激发脉动燃烧,压力脉动频率在100Hz~110Hz,最大压力脉动的幅值约为3.8kPa,燃烧效率接近100%。
短句来源
    The operating process of liquid apogee engine (N2O4/MMH) thrust chamber is studied. By applying the SIMPLE algorithm with staggered grid, the numerical model is established. The atomization mechanism and decomposition combustion of hypergolic rocket propellant are considered in the model.
    研究了远地点发动机(N2O4/MMH)的推力室工作过程,考虑了自燃推进剂的雾化、蒸发以及化学反应流动过程,采用交错网格系统的SIMPLE算法,得到了不同边区冷却流量对推力室的内流场和燃烧效率的影响结果,数值计算的结果与理论分析相符合,为推力室工作过程的稳定性分析提供了重要参考。
短句来源
    The combustion efficiencies were 0.82, 0.67 and 0.60, respectively.
    燃烧室出口的燃烧效率分别为0·82,0·67和0·60。
短句来源
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  combustion efficiency
The theoretical and experimental results show that film combustion has a higher combustion efficiency, a lower pollutant emission and a better working performance.
      
The model is developed on the basis of a "burn-out curve", that is, the dependence of the combustion efficiency on the longitudinal coordinate and the design features of the chamber.
      
Nitrogen addition also enhances the diffusion rate and combustion efficiency.
      
The temperatures of the combustor are stable and the char combustion efficiency is about 98%.
      
The concentration of elemental components in diesel soot, generally, varied with operating conditions, which affected fuel and oil consumption, combustion efficiency (soot production), and mechanical strain.
      
更多          
  burning efficiency
Plasma fuel systems that increase the coal burning efficiency are discussed.
      
We perform a numerical study of the burning efficiency in a closed vessel.
      
Length scales provide some understanding of the injection of cryogenic propellants in rocket chambers on mixing efficiency, which translates to burning efficiency and performance.
      
In comparison, about 9-17 Pg of above-ground dry matter (4-8 Pg C) is exposed to fires, indicating a worldwide average burning efficiency of about 50%.
      
As said before, burning efficiency is related to several factors such as vegetation water content and fine fuel moisture content.
      
更多          
  efficiency of combustion
In so doing, a glowing flame is formed, whose geometric dimensions define the energy efficiency of combustion of fuel.
      
The efficiency of combustion in a supersonic high-temperature flow at spread and local hydrogen feed is considered.
      
The influence of the amount of sample transported into the oxyhydrogen flame on the efficiency of combustion is presented.
      
Effect of oxidizer particle size distribution on efficiency of combustion catalysts
      
Efficiency of combustion of hydrogen-kerosene fuel in a straight duct
      
更多          


A simplified model for the disign of hypergolic liquid propellant rocket engine combustors is presented.In contrast with Priem's basic model,the effects of decomposition reaction of fuel,secondary atomization of droplets and non-uniformity of mixture ratio and flow flux in the cross-section of the chamber on the combustion efficiency have been considered.All equations required for the calculation of combustion efficiency are derived and the computer program is given.As an example of the application of this method,the...

A simplified model for the disign of hypergolic liquid propellant rocket engine combustors is presented.In contrast with Priem's basic model,the effects of decomposition reaction of fuel,secondary atomization of droplets and non-uniformity of mixture ratio and flow flux in the cross-section of the chamber on the combustion efficiency have been considered.All equations required for the calculation of combustion efficiency are derived and the computer program is given.As an example of the application of this method,the combustion efficiency of “C” engine is calculated.Comparison of calculated Combustion efficiencies with experimental results shows satisfactory agreement with errors within 1-5%.Curves of total vaporization rate of fuel and vaporization rate per unit length as functions of the axial distance of combustor for given rocket engine are also presented.These curves show that the effects of decomposition reaction,secondary atomization of droplets and drop size distribution on vaporization rate are very substantial.Finally, the conditions,under which this model can be used,are discussed

本文提出了一个自燃推进剂液体火箭发动机燃烧室设计简化模型。与 Priem 基本模型不同,本文考虑了燃料分解反应,液滴二次雾化以及燃烧室截面上混合比和流强的不均匀性对燃烧效率的影响。文中导出了计算所需的全部公式,提出了计算程序。应用本方法计算了 C 发动机的燃烧效率。计算效率与试验结果相当符合,其误差在1~5%以内。文中还给出了燃料总蒸发速率和单位长度上的蒸发速率随该发动机燃烧室轴向距离变化的曲线。曲线表明,分解反应,液滴二次雾化和液滴尺寸分布对蒸发速率的影响很大。最后讨论了本模型使用的条件。

The Integral Rocket Ramjet (IRR) which usually uses fuel-rich solid propellant is a type of modern power plant for new generation tactical missiles. In this paper the combustion process of fuel-rich aluminized solid propellant in IRR was investigated. For the purpose of acheiving better mixing flow field in the ramjet chamber, some various types of rocket nozzles were tried.The static pressure profile of non-parallel four-port secondary flow, the static and total pressure profile of primary flow, and the velocity...

The Integral Rocket Ramjet (IRR) which usually uses fuel-rich solid propellant is a type of modern power plant for new generation tactical missiles. In this paper the combustion process of fuel-rich aluminized solid propellant in IRR was investigated. For the purpose of acheiving better mixing flow field in the ramjet chamber, some various types of rocket nozzles were tried.The static pressure profile of non-parallel four-port secondary flow, the static and total pressure profile of primary flow, and the velocity profile of mixing flow have been measured.Three conditions under which the aluminum particles could fully burn up were also discussed. On the analytical basis of the effect of rocket nozzle types on the mixing flow field, multiple subsonic rocket nozzles were finally selected. The secondary combustion efficiency with this type of nozzles was increased by about 30% than that with single sonic (or supersonic) rocket nozzles.In the combustion efficiency experiments, the weight of the solid propellant grain is 5.5-8.9 kg, operating time is 14 to 23 seconds.

本文研究了固冲发动机含铝推进剂的燃烧过程,为了合理组织主、次流的掺混流场,选用不同型式的火箭喷管进行试验。测量了四孔非平行进气的次流静压分布;单独主流的总、静压分布;音速单喷管等六种喷管的主次流掺混流场的速度场。 本文分析了铝颗粒完全燃烧的三条件,并结合火箭喷管型式对冲压室内掺混流场的分析,选用了具有4×φ12—15°斜喷口的多孔分流式亚音速喷管为火箭喷管,进行了燃烧效率试验。实验结果表明,冲压室燃烧效率比采用音速(或超音速)单喷管时提高30%左右。 燃烧效率试验用的装药为含铝贫氧推进剂,重量5.5—8.9公斤,工作时间14—23秒。

The evaporation-decomposition process of a liquid hypergolic bipropellant droplet at high-pressure and high-temperature conditions is studied thoroughly, and the model of high pressure equilibrium evaporation for this liguid droplet is presented, in this model there are considerations for droplet surface regression, effects of non-ideal gases, changes of the properties, decomposition and dissociation of the species. Using this model, the equilibrium vapovization constants of UDMH and N_2O_4 droplets at different...

The evaporation-decomposition process of a liquid hypergolic bipropellant droplet at high-pressure and high-temperature conditions is studied thoroughly, and the model of high pressure equilibrium evaporation for this liguid droplet is presented, in this model there are considerations for droplet surface regression, effects of non-ideal gases, changes of the properties, decomposition and dissociation of the species. Using this model, the equilibrium vapovization constants of UDMH and N_2O_4 droplets at different environment pressures. temperatures and convective intensities are calculated. The results indicate that there exists a critical environment pressure at definite temperature for given species above which there is no equilibrium evaporation. For UDMH, when T_∞=3200°K, Le=1, the pressure p?? at which there appears supercritical evaporation equal 54 atm, and for N_2 O_4 P_∞= 120 atm.The results also indicate that the vaporization rate for UDMH is greater than that for N_2O_4. So we can draw a conclusion that under ordinary working conditions of liquid propellant rocket engine UDMH behaves supercritical vaporization and N_2O_4. behaves subcritical vaporization and combustion efficiency is controled by the vaporization rate of N_2O_4. These have been confirmed by testing of liquid propellant rocket engine.

本文详细分析了自燃推进剂组元液滴在高温高压燃烧室环境下的蒸发——分解燃烧过程,提出了该种液滴的高压平衡蒸发计算模型.模型考虑了液滴界面移动、非理想气体效应、流体物性的变化以及组元的分解和离解效应.应用本模型计算了UDMH和N2O4液滴在不同介质压力、温度和对流强度下的平衡蒸发常数.计算表明,存在一个介质界限压力,超过这一压力就达到超临界蒸发.对于UDMH,当T_∞=3200°K,Le=1.0时,界限压力P_∞=54大气压,而对于N2O4,P_∞=120大气压.计算还表明,UDMH的蒸发速度远大于N_2O_4的蒸发速度.因而可以得出结论,在一般液体火箭发动机的工作条件下,UDMH为超临界蒸发,而N2O_4为亚临界蒸发,而且发动机的燃烧效率主要受N_2O_4的蒸发速度所控制.这一结论已为发动机试车所证实.

 
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