爆轰波对喷嘴中过度膨胀气流的影响

S. S. Vasyliv, O.O. Kirichenko
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摘要

本文提出利用爆炸过程来解决高机动性飞行器的控制问题。这项工作的目标是研究一种利用爆炸冲击波对喷嘴中气体流动的影响来控制火箭发动机推力矢量的新方法。众所周知,通过向火箭发动机的超音速喷嘴区域喷射气体来控制推力矢量的方法具有比动量损失最小、响应速度快的特点,并能产生很大的侧向力。然而,就目前的火箭技术水平而言,有必要改进这些特性。引爆被认为是加强影响主气流和产生侧向力的过程的一种方法。它的特点使得开发脉冲弹道修正系统成为可能。为了研究这种系统的特点,对超音速喷嘴流的爆轰波效应进行了实验研究。开发并制作了一个喷嘴模型和一个气体发生器,用于引发与主超音速气流相互作用的爆炸波。在实验过程中,发现了喷嘴在超膨胀模式下工作时,主气流与喷嘴壁被爆轰波分离的效应。这种效应的持续时间远远长于爆炸产物对喷嘴内主气流的影响,因此可用于开发一种新的推力矢量控制方法。为了更好地理解实验结果,对试验条件下超音速气流中的爆轰波传播进行了数值模拟。模拟是使用 Solid Works 软件包以非稳态二维形式进行的。模拟的目的是估算流体结构,以及在向超音速喷嘴区域注入爆燃产物时,流体结构变化所产生的相对侧向力值。获得了压力场的时间演变。确定了向喷嘴中的超音速气流注入起爆产物所产生的相对侧向力。所介绍的喷气发动机推力矢量控制特征和方法可用于在各种速度范围内运行的无人驾驶系统。
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Effect of a detonation wave on an overexpanded gas flow in a nozzle
This paper proposes to use the detonation process to solve the problem of controlling highly maneuverable flying vehicles. The goal of the work is to study a new way of the thrust vector control of a rocket engine using the effect of a detonation shock wave on the gas flow in the nozzle. It is known that the method of thrust vector control by gas injection into the supersonic nozzle area of a rocket engine features one of the lowest losses of specific momentum and a high response speed and produces a significant lateral force. However, for the current level of rocket technology, there is a need to improve these characteristics. Detonation is considered as a method to intensify processes that affect the main gas flow and produce a lateral force. Its features make it possible to develop a system for pulse trajectory correction. In order to study the features of such a system, experimental studies of the detonation wave effect on a supersonic nozzle flow were conducted. A nozzle model and a gas generator for initiating a detonation wave interacting with the main supersonic air flow were developed and made. In the course of experiments, the effect of separation of the main flow from the nozzle wall by the detonation wave during the nozzle operation in the overexpansion mode was revealed. The duration of this effect was much longer than that of the detonation product effect on the main air flow in the nozzle, thus allowing it to be used in the development of a new thrust vector control method. To better understand the experimental results, a numerical simulation of the detonation wave propagation in a supersonic flow was carried out for the test conditions. The simulation was carried out in a non-stationary 2D formulation using the Solid Works software package. The goal of the simulation was to estimate the flow structure and the value of the relative lateral force produced by the change of this structure during detonation product injection into the supersonic nozzle area. The time evolution of the pressure field was obtained. The relative lateral force produced by detonation product injection into the supersonic air flow in the nozzle was determined. The presented features and method of jet engine thrust vector control may be used in unmanned systems operating in a wide range of speeds.
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