JPS59128955A - Rocket engine - Google Patents

Abstract

PURPOSE:To obtain optimum thrust force X time, by arranging in the axial direction a combustion control member, which is installed partially in the circumferential direction of grains loaded in a rocket chamber and which has a portion closed toward the periphery. CONSTITUTION:A combustion control member 3, which is installed in the circumferential direction of grains 1 loaded in a rocket chamber 4 and which has a portion closed toward the periphery, is arranged in the axial direction. Thereby the web according to the inner-surface combustion system will become a one having substantially a diameter 2-3 times as large as the grains' 1, to cause substantial increase of the combustion progressive web. The combustion pattern thus obtained will transfer from the combustion initial face 1a to the combustion progressive face A one after another, wherein a comparatively large thrust force is obtained in the initial stage, while a comparatively small, approx. constant and stable thrust force is obtained continuously after passage of a certain period of time. This provides effectiveness for launching an aeronautical body in vertical position to a high altitude, as well as the reach can be stretched even in launching aslant.

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a rocket engine suitable for use as a high-altitude flying vehicle.

As is well known, in order to effectively launch a flying object to a high altitude, there is an optimal combination of the thrust of the rocket engine and the duration of thrust action (usually referred to as operating time 5). (Patent Application Publication No. 1976-762, Patent Application Publication No. 1973
1-6762).

However, most conventional rockets have been of the internal combustion type, which does not provide the optimum combination of thrust and time. Additionally, the end-burning type lacks thrust and is inefficient. There is also a special end-combustion type in which a high-burning-rate propellant, etc. is embedded on the shaft, as disclosed in some of the above-mentioned patent applications, but it is not widely used. In this regard, various attempts have been made to obtain a rocket engine with optimal thrust x time, but none have been satisfactory.

The purpose of the present invention is to obtain a rocket engine with an optimal thrust x time using only grains with a burning rate that can be easily obtained at present. This objective is achieved by a structure in which a combustion restricting material is disposed along the axial direction of the rocket chamber, and a part of the combustion restricting material is closed to the outer periphery. Furthermore, by arranging the combustion limiter partially along the axial direction of the exhaust chamber, it is attempted to obtain even better thrust.

Hereinafter, one embodiment of the present invention will be described based on the drawings.

The cross-sectional shape of an embodiment of the present invention is shown in FIGS. 1 to 3. Normally, the internal combustion type web is 25% to 30% of the grain diameter, but in the present invention, the combustion restriction material 3 partially disposed in the circumferential direction of the grain 1 makes it substantially normal. A web with a diameter of 2 to 3 times the diameter of grain 1 is obtained.

Among these, the effect of substantially increasing the combustion progressing web in the embodiment shown in FIG. 5 will be explained as shown in FIG. 4. In this example, it has increased by about 2.5 times. In addition, 1a in FIG. 4 shows the initial stage of grain combustion. The combustion pattern according to this example is shown in FIG. That is, in the initial stage of combustion, a relatively large thrust F is obtained, and after a certain period of time, a relatively small, almost constant, stable thrust is continuously obtained. Regarding such a combination of thrust force and time, various design changes can be made by changing the position and shape of the combustion restricting material 6. If a rocket engine is designed using the present invention, the initial thrust can be made sufficiently large, and the desired thrust can be obtained thereafter. It is not necessary to use booster grains in conjunction with this. Depending on requirements, the initial thrust can be further increased by disposing the combustion restricting material 5 not over the entire length along the axial direction of the rocket chamber, but only partially as shown in FIG. 6, for example. In addition, 5 in FIG. 6 indicates a nozzle. Furthermore, the combustion limiting material disposed within grain 1 may be made of a soft material, or may have a two-layer or multi-layer structure to alleviate stress caused by the temperature environment, handling, etc. Therefore, destruction of grain 1 or inherent stress distortion can be prevented. An example thereof is shown in FIG.

Since the present invention is configured as described above, it is possible to obtain a sufficient initial thrust similar to that when using the internal combustion type, and after that, it is possible to continuously obtain a substantially constant thrust. It becomes easy to design a combustion pattern with a good relationship between thrust and time. Furthermore, by making the combustion restricting material soft or having a two-layered or multi-layered structure, the grain 1 will not be destroyed or distorted.

The present invention allows the flying object to l'! ! Not only is it effective for launching almost vertically to high altitudes, but it also makes it possible to greatly extend the range when flying diagonally at an angle to the ground compared to normal internal combustion rocket engines. It is something to do.

[Brief explanation of the drawing]

1 to 3 and 7 are cross-sectional views showing one embodiment of the present invention, FIG. 4 is an enlarged view of main parts showing the progress of combustion, and FIG. Thrust (F), time (f
) A graph of the curve, and Figure 6 is a vertical cross-sectional view of an embodiment of a rocket engine when applied to an actual rocket. ... Grain combustion initial surface 2 ... Rocket engine 3 ... Combustion restriction material 4 ... Rocket chamber 5 ... Nozzle exit angle Kaoru Furuya

Claims (1)

[Claims] 1. A combustion restricting material that is partially disposed in the circumferential direction of the grain loaded in the rocket chamber and partially closed to the outer periphery is disposed along the axial direction of the rocket chamber. A rocket engine characterized by having a structure in which: 2. The rocket engine according to claim 1, having a structure in which the combustion restricting material is partially disposed along the axial direction of the rocket chamber. 3. The rocket engine according to claim 1 or 2, in which the combustion-side limiting material is made of a soft material, or has a two-layer or multi-layer structure.