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首页> 外文会议>Fifth International Conference on Electronic Measurement amp; Instruments (ICEMI'2001) Vol.1 Nov 18-21, 2001 Guilin, China >Design of a Temperature-measurement and Control System for Solid Propellant Rocket Motor
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Design of a Temperature-measurement and Control System for Solid Propellant Rocket Motor

机译：固体火箭发动机的测温与控制系统设计

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摘要

The paper states a way of incorporating a new type Flash single-chip, signal acquisition circuits, temperature-control circuits into a temperature-control system for a solid propellant rocket motor, to fulfil the settings on the initial temperature of its charge. Beginning with a few words on why it is important to design such a system, this article mainly contains two parts: the hardware design and the software design. Solid propellant rocket motors are rocket motors that use solid propellant, which is formed into charges of certain shape and then crammed into the combustor. In the combustor, the charge is ignited and burns away, producing gases at high temperature and pressure. The gases vent through the nozzle at a high velocity, and behave as a jet of propulsive medium, which produces reaction propulsion. Compared with the liquid counterparts, the solid propellant rocket motor has simpler construction and higher reliability, and is easier to be en-charged, transported and maintained, so it is widely used in many types of missiles. But, the combustion rate of the charge, as shown in fig.1, is greatly effected by its initial temperature. For a given charge, if the initial temperature is increased, the combustion rate will be increased, and the combustion pressure and hence the resulting thrust force will also be increased, whereas the duration of the combustion will be shortened. This dependence works in the opposite way if the initial temperature is decreased. So the magnitude of the initial temperature effects the thrust-time curve of the rocket motor, causing great degradation to its performance. In a strictly required missile, there must be a temperature-control system to ensure that the initial temperature of the charge meets the design requirement for the missile flight.
机译：该论文提出了一种将新型Flash单片机，信号采集电路，温度控制电路并入用于固体推进剂火箭发动机的温度控制系统的方法，以实现其装药初始温度的设定。首先说一下为什么设计这样的系统很重要，本文主要包含两个部分：硬件设计和软件设计。固体推进剂火箭发动机是使用固体推进剂的火箭发动机，该固体推进剂形成一定形状的炸药，然后塞入燃烧室。在燃烧室中，装料被点燃并燃烧掉，从而在高温和高压下产生气体。气体以高速度通过喷嘴排出，并表现为推进介质的喷射，从而产生反应推进。与液体对等物相比，固体推进剂火箭发动机结构更简单，可靠性更高，更易于装填，运输和维护，因此广泛用于多种类型的导弹。但是，如图1所示，装料的燃烧速率受其初始温度的影响很大。对于给定的装料，如果增加初始温度，则燃烧速率将增加，并且燃烧压力以及因此产生的推力也将增加，而燃烧的持续时间将缩短。如果降低初始温度，则这种依赖性以相反的方式起作用。因此，初始温度的大小会影响火箭发动机的推力时间曲线，从而导致其性能大大下降。在严格要求的导弹中，必须有一个温度控制系统，以确保装药的初始温度满足导弹飞行的设计要求。

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