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  thermal protection
Heat and mass transfer associated with low-velocity vapor flow in the plane channel of a thermal protection system of the radiat
      
Characteristics of the heat and mass transfer in evaporative thermal protection with variation of the external pressure
      
A thermal protection system with evaporation into an inner cavity is considered.
      
Modeling the unsteady heat and mass transfer in the vapor outlet channel of a radiative-evaporative thermal protection system
      
Efficiency of a radiative-evaporative aircraft thermal protection system
      
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  heat shield
Nonstationary gas filtration in a two-layered heat shield
      
A gas mist produced by supplying a cooling gas through a separate slot or a system of slots along the surface is a widespread method of obtaining a mass-transfer heat shield.
      
According to estimates in [1], the heat shield mass for entry of a probe into the atmospheres of the outer planets can make up 20-50% of its total mass; here the radiative component predominates in the aerodynamic heating.
      
The results are compared with the data of a space shuttle flight experiment on the effect of a discontinuity of heat shield catalyticity on the "overequilibrium" surface temperature jump.
      
One of the important problems has been to create reusable heat shield materials.
      
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  thermal protection
Heat and mass transfer associated with low-velocity vapor flow in the plane channel of a thermal protection system of the radiat
      
Characteristics of the heat and mass transfer in evaporative thermal protection with variation of the external pressure
      
A thermal protection system with evaporation into an inner cavity is considered.
      
Modeling the unsteady heat and mass transfer in the vapor outlet channel of a radiative-evaporative thermal protection system
      
Efficiency of a radiative-evaporative aircraft thermal protection system
      
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The determination of heat conduction with mass transfer and chemical reaction is an important problem encountered in design of heat shied for reentry command module. An integral method for solving this problem is provided.Starting with energy balance for definite element, a set of ordinary differential equations has been obtained for temperature at discrete points in present paper. It has been shown the method presented in this paper is more effective in obtaining temperature history in the regions of pyrolysis...

The determination of heat conduction with mass transfer and chemical reaction is an important problem encountered in design of heat shied for reentry command module. An integral method for solving this problem is provided.Starting with energy balance for definite element, a set of ordinary differential equations has been obtained for temperature at discrete points in present paper. It has been shown the method presented in this paper is more effective in obtaining temperature history in the regions of pyrolysis and interface. Comparison with analytic results of one-dimensional heat conduction shows a good agreement. The calculation results have also been consistent with the ground ablation experiments.

有质量引射和化学反应的内部热传导的确定是卫星再入舱热防护设计中遇到的一个重要问题。本文给出解决这类问题的一个积分方法。本文从有限元的能量平衡出发,导出了一组确定离散点温度的常微分方程。计算表明,本文给出的方法对确定热解区和分解面的温度分布更为有效。本文提供的方法对一元热传导的计算结果与解析解是一致的。对地面烧蚀实验状态的计算结果与测量结果也是一致的。

The 300mm diameter solid Propellant rocket motor employs a composite propellant end burning grain. The effective burning time was 42 seconds and the thrust was 720kg. For the motor operating reliability, the motor utilizes a large number of nonmetal materials for thermal protecion of the case and the nozzle, for the lining of the grain, and for the motor assembly seal. The motor repeated firing tests proved that the performance of the materials was good, the design of the structures was reasonable, the technological...

The 300mm diameter solid Propellant rocket motor employs a composite propellant end burning grain. The effective burning time was 42 seconds and the thrust was 720kg. For the motor operating reliability, the motor utilizes a large number of nonmetal materials for thermal protecion of the case and the nozzle, for the lining of the grain, and for the motor assembly seal. The motor repeated firing tests proved that the performance of the materials was good, the design of the structures was reasonable, the technological method was reliable, and all the design criteriaw ere met. This paper discusses the design rationale, the structural features, theperformance, the results of the motor firing test, and the major technical problems of these nonmetal material components.

直径300毫米的固体推进剂火箭发动机采用了复合推进剂端燃药柱。有效燃烧时间42秒,推力720公斤。为保证发动机可靠地工作,发动机采用了大量的非金属材料用于壳体及喷管的热防护、药柱的包覆和发动机装配的密封。多次发动机点火试验证明这些料材的性能良好,结构设计合理,工艺措施可靠,符合设计要求。本文论述了这些非金属材料部件的设计思想,结构特点、性能、点火试验结果以及主要技术问题。

The requirements for the internal insulation arc presented based on its functions in a solid rocket motor. The formulas to compute the charred thickness are derived based on its ablation mechanism and the methods to determine its design thickness arc given based on thermal protection for the chamber case.

根据内绝热层在固体火箭发动机中的作用,对其提出了若干要求;根据内绝热层的烧蚀机理,推导出内绝热层炭优厚度计算公式;根据燃烧室壳体对热防护的要求,给出了确定内绝热层设计厚度的方法。

 
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