JP2004137956A - Solid motor for hybrid rocket engine and its combustion method - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a solid motor for a hybrid rocket engine and its combustion method capable of supplying a liquid/gas propellant produced by gasifying a part of the liquid propellant to an entire solid propellant at a nearly optimum flow, and thereby approximately uniformizing combustion temperature in a combustion chamber, increasing flame temperature, and increasing specific exhaust speed of a nozzle and specific thrust of an engine. <P>SOLUTION: The solid motor for the hybrid rocket engine comprises hollow cylindrical a solid motor case 4 having the nozzle 3, the solid propellant 1 comprising a plurality of segments 1a stored in an axial direction in the solid motor case with spaces, an end surface injector 6 for injecting the liquid/gas propellant 2a toward an end surface in an opposite side of the nozzle of the solid propellant, and a gap injector 7 for injecting the liquid/gas propellant into the gap among the segments. The liquid/gas propellant is reacted with the solid propellant. <P>COPYRIGHT: (C)2004,JPO

Description

[0001]
TECHNICAL FIELD OF THE INVENTION
The present invention relates to a hybrid rocket engine using a liquid propellant and a solid propellant, and more particularly, to a solid motor of a hybrid rocket engine and a combustion method thereof.
[0002]
[Prior art]
Hybrid rocket engines using liquid propellants and solid propellants include those using liquid fuel as liquid propellant and solid fuel as solid propellant, and those using solid fuel and liquid oxidant on the contrary. These hybrid rocket engines are highly safe due to the separation of fuel and oxidant, and have the characteristics (stability and controllability) of solid rockets. Is being eagerly pursued.
[0003]
Hybrid rocket engines can be broadly classified into a gas pushing type and a pump type depending on the liquid propellant supply method. Gas-push type hybrid rocket engines are disclosed in, for example, Patent Documents 1 to 5 and the like. Patent Literatures 6 and 7 disclose pump-type hybrid rocket engines. Further, Patent Document 8 related to the present invention is disclosed.
[0004]
[Patent Document 1]
Japanese Patent No. 2666883 [Patent Document 2]
Japanese Patent No. 2673663 [Patent Document 3]
JP-A-7-19120 [Patent Document 4]
JP-A-7-310594 [Patent Document 5]
JP-A-7-310595 [Patent Document 6]
JP-A-8-93557 [Patent Document 7]
JP-A-8-93558 [Patent Document 8]
JP-A-8-61150
[Problems to be solved by the invention]
In both the gas-push type and the pump type, as illustrated in FIG. 5A, the solid propellant 1 is stored in a solid motor case 4 having a nozzle 3, and from the uppermost stream of the solid propellant 1. A part of the liquid propellant 2 has been jetted at the entire flow rate of the gasified liquid / gas propellant 2a. This figure shows a solid-state motor of the collision combustion type, and 5 is an injector.
[0006]
In such a conventional hybrid rocket engine, the flow rate of the liquid / gas propellant 2a is set so that the ratio with the solid propellant amount 1 becomes a necessary value as a whole. Therefore, even if the mixing ratio is optimal as a whole, the mixing ratio differs between the upstream portion and the downstream portion of the solid propellant, and as shown schematically in FIG. In the temperature distribution, especially in the upstream portion, the liquid / gas propellant 2a becomes excessive (excess oxidant or excess fuel), the flame temperature becomes low, and the overall average combustion temperature does not rise, which is a factor that lowers the performance. .
[0007]
Ideally, the "injector for hybrid rockets" is intended to distribute the oxidant evenly over the entire combustion surface in all phases of combustion and to approximate the combustion process of solid systems [ Patent Document 8] is disclosed.
In this injection device for a hybrid rocket, a liquid component flows through a number of narrow tubes, across an unburned portion of a solid component, and flows toward a combustion surface of an internal combustion port. It originates on the outer surface of the solid fuel block and extends radially into and into a centrally located port.
[0008]
However, since the injection device of this structure, that is, the solid motor, needs to form a large number of tubes in the solid propellant, the shape of the solid propellant 1 is significantly different from that of the conventional one, and its molding is difficult. In addition, there is a problem that can be applied only to an internal combustion type solid motor.
[0009]
The present invention has been made to solve such a problem. That is, an object of the present invention is to provide a liquid / gas propellant 2a in which a part of the liquid propellant 2 is gasified to be supplied to the entire solid propellant 1 at a substantially optimum flow rate, thereby achieving a combustion temperature in the combustion chamber. It is an object of the present invention to provide a solid motor of a hybrid rocket engine and a combustion method thereof, which can make the temperature substantially uniform, increase the flame temperature, and increase the characteristic exhaust speed of the nozzle and the specific thrust of the engine.
[0010]
[Means for Solving the Problems]
According to the present invention, a solid propulsion comprising a hollow cylindrical solid motor case (4) having a nozzle (3) and a plurality of segments (1a) stored in the solid motor case with a gap in the axial direction. A drug (1), an end injector (6) for injecting a liquid / gas propellant (2a) toward an end face opposite to a nozzle of the solid propellant, and a liquid / gas propellant in a gap between the segments. And a gap injector (7) for injecting the liquid and gas propellant and reacting the liquid and gas propellant with the solid propellant.
[0011]
Further, according to the present invention, a solid cylindrical motor case (4) having a nozzle (3) and a plurality of segments (1a) stored in the solid motor case with an axial gap therebetween. Comprising a solid propellant (1) and an inner injector (9) for injecting a liquid / gas propellant (2a) toward the inner surface of each segment, and reacting the liquid / gas propellant with the solid propellant. A solid-state motor for a hybrid rocket engine is provided.
[0012]
Further, according to the present invention, a solid propellant (1) composed of a plurality of segments (1a) is stored in a hollow cylindrical solid motor case (4) having a nozzle (3) with a gap in the axial direction. Distributing the liquid / gas propellant (2a) into the gap between each segment and the end face opposite to the nozzle of the solid propellant and injecting it, and reacting each segment with the liquid / gas propellant at an almost optimum mixing ratio; A method for burning a hybrid rocket engine is provided.
[0013]
According to the apparatus and method of the present invention, the solid propellant (1) is divided into several segments (1a), and a liquid / gas propellant (2a) having an optimal flow rate is injected from the upstream for each segment. Therefore, the combustion temperature in the combustion chamber can be made substantially uniform, and the flame temperature can be kept high. Therefore, as the flame temperature increases, the characteristic exhaust speed (C ^* （T ₀ ^0.5 ) of the nozzle increases, and the specific thrust (I _sp ∝C ^* ) of the engine increases.
As a result, the performance (specific thrust) is improved, so that the amount of the propellant can be reduced, or the payload of the engine can be increased.
Further, when the gap between the segments (1a) of the solid propellant (1) is widened, this portion functions as a pre-combustion chamber, so that the mixing of the liquid / gas propellant and the solid propellant is performed smoothly.
[0014]
According to a preferred embodiment of the present invention, the plurality of solid propellants (1) have alternately axially penetrating gas passages (1b). Further, the plurality of solid propellants (1) may have a gas passage (1b) penetrating in the axial direction at the same position.
Further, the end face injector (6) and the gap injector (7) are provided with a chamber (6a, 7a) to which a liquid / gas propellant (2a) is supplied, and a partition (6b, 7b) for separating the chamber from the combustion chamber. 7b), and the partition is provided with a plurality of injection holes (8) or meshes.
The inner injector (9) includes a chamber (9a) to which a liquid / gas propellant (2a) is supplied, and a cylindrical tube (9b) extending from the chamber to the inside of the solid propellant (1) in the axial direction. The cylindrical pipe has a plurality of injection holes (8).
Furthermore, the ignition energy of each solid propellant (1) uses the combustion gas of the solid propellant further upstream. Further, it is preferable to shift the injection timing of the liquid / gas propellant (2a) to the downstream solid propellant (1).
[0015]
With these structures and methods, the combustion gas on the upstream side can be used to ignite the segments on the downstream side, and the mixing and reaction of the liquid / gas propellant (2a) and the solid propellant (1) can be promoted.
[0016]
According to a preferred embodiment of the present invention, a plurality of solid propellants (1) are stored in a hollow cylindrical solid motor case (4) having a nozzle (3) with a gap therebetween in an axial direction, and a liquid / gas is stored. The propellant (2a) is distributed and injected to the inner surface of each solid propellant, and each solid propellant reacts with the liquid / gas propellant at a substantially optimum mixing ratio.
[0017]
By this method, the flame temperature can be increased, the characteristic exhaust speed of the nozzle can be increased, and the specific thrust of the engine can be increased.
[0018]
In addition, the flow rate of the liquid / gas propellant (2a) to be injected upstream of each solid propellant (1) is expressed as “total liquid propellant flow rate / total solid propellant weight = liquid / gas propulsion for each solid propellant. It is preferable to dispense such that "drug flow rate / weight of each solid propellant".
By this method, the flame temperature, the characteristic exhaust speed of the nozzle, and the specific thrust of the engine can be optimized.
[0019]
BEST MODE FOR CARRYING OUT THE INVENTION
Hereinafter, preferred embodiments of the present invention will be described with reference to the drawings. In addition, the same reference numerals are given to the common parts in the respective drawings, and the duplicate description will be omitted.
[0020]
FIG. 1A is a diagram showing a first embodiment of a solid motor of a hybrid rocket engine according to the present invention, and FIG. 1B is a diagram showing a temperature distribution thereof.
As shown in FIG. 1A, the solid motor 10 of the present invention includes a solid motor case 4, a solid propellant 1, an end face injector 6, and a gap injector 7.
[0021]
The solid motor case 4 is a hollow cylindrical airtight container having the nozzle 3. The solid propellant 1 is composed of a plurality (three in this example) of segments 1a. Each segment 1a is stored in the solid motor case 4 with a gap in the axial direction. The gap between the segments 1a is set to an interval suitable for the collision jet type in this example.
[0022]
2 is a partially enlarged view of FIG. 1, in which (A) is an enlarged view of the end face injector 6 and (B) is an enlarged view of the gap injector 7. As shown in FIG. 2A, the end face injector 6 includes a chamber 6a to which the liquid / gas propellant 2a is supplied, and a partition wall 6b that partitions the chamber 6a from the combustion chamber. A plurality of injection holes 8 are provided in the partition 6b.
With this configuration, the supply / injection amount of the liquid / gas propellant 2a to be injected toward the end face of the solid propellant 1 opposite to the nozzle can be adjusted by the diameter of the injection hole 8 and the overall opening ratio.
[0023]
As shown in FIG. 2B, the gap injector 7 includes a chamber 7a to which the liquid / gas propellant 2a is supplied, and a partition wall 7b that separates the chamber 7a from the combustion chamber. A plurality of injection holes 8 are provided.
With this configuration, the supply and injection amount of the liquid / gas propellant 2a to each gap between the segments 1a can be adjusted by the diameter of the injection hole 8 and the entire opening ratio.
[0024]
FIG. 3 is a partially enlarged view showing another embodiment of FIG. In this figure, a mesh is provided on the partition wall 7b, and the supply and injection amount of the liquid / gas propellant 2a is adjusted by the opening ratio and the thickness of the mesh. This configuration can also be applied to the end face injector 6.
[0025]
FIG. 4 is a diagram showing another embodiment of the present invention. In this figure, in (A), a plurality of segments 1a have gas passages 1b penetrating in the axial direction at the same position. In (B), a plurality of segments 1a have gas passages 1b which penetrate alternately or at the same position in the axial direction. With these structures, the combustion gas on the upstream side can be used to ignite the segments on the downstream side, and the mixing and reaction of the liquid / gas propellant 2a and the solid propellant 1 can be promoted.
Further, with this configuration, the injection timing of the liquid / gas propellant 2a to the downstream segment can be shifted, and the ignition energy of each segment can use the combustion gas of the more upstream segment.
[0026]
In the embodiment shown in FIG. 4C, an inner injector 9 for injecting the liquid / gas propellant 2a toward the inner surface of each segment 1a is provided. The inner injector 9 includes a chamber 9a to which the liquid / gas propellant 2a is supplied, and a cylindrical tube 9b extending from the chamber 9a to the inside of each segment 1a in the axial direction. Further, a plurality of injection holes 8 are provided in the cylindrical tube.
Also with this configuration, the supply and injection amount of the liquid / gas propellant 2a to each gap between the segments 1a can be adjusted by the diameter of the injection hole 8 and the entire opening ratio.
[0027]
Using the solid motor 10 of FIGS. 1 to 3 and FIGS. 4 (A) and 4 (B) described above, in the combustion method of the hybrid rocket engine of the present invention, an axial direction is provided in a hollow cylindrical solid motor case 4 having a nozzle 3. The solid propellant 1 composed of a plurality of segments 1a is stored with a gap therebetween, and the liquid / gas propellant 2a is distributed and jetted to the gap between each segment and the end face opposite to the nozzle of the solid propellant, The segments react with the liquid / gas propellant at an almost optimal mixing ratio.
[0028]
The combustion system in this case is a collision jet system, in which the reaction gas on the end faces of the respective segments 1a is injected to the outside from the nozzle 3 through the gas passage 1b while colliding with each other, thereby obtaining thrust.
[0029]
In this combustion method, the flow rate of the liquid / gas propellant 2a to be injected to the upstream side of each segment is defined as “total liquid propellant flow rate / total solid propellant weight = liquid / gas propellant flow rate per segment / 1 segment”. Dispense to "solid propellant weight". That is, in the example of FIG. 1, when the solid propellant 1 is divided into three equal parts to form three segments 1a, the liquid / gas propellant 2a is also divided into three equal parts and supplied to the respective upstream sides.
[0030]
According to the above-described apparatus and method of the present invention, the solid propellant 1 is divided into several segments 1a, and the liquid / gas propellant 2a having the optimum flow rate is injected from the upstream for each segment. The combustion temperature can be made substantially uniform, and the flame temperature can be kept high, as shown in FIG. Therefore, as the flame temperature increases, the characteristic exhaust speed (C ^* （T ₀ ^0.5 ) of the nozzle increases, and the specific thrust (I _sp ∝C ^* ) of the engine increases. In addition, since the performance (specific thrust) is improved, the amount of propellant can be reduced, or the payload of the engine can be increased.
[0031]
Further, using the solid motor 10 shown in FIG. 4C, in the combustion method of the present invention, a solid propulsion comprising a plurality of segments 1a separated by an axial gap in a hollow cylindrical solid motor case 4 having a nozzle 3 is provided. The medicine 1 is stored, and the liquid / gas propellant 2a is distributed and injected to the inner surface of each segment, and each segment and the liquid / gas propellant are caused to react at a substantially optimum mixing ratio. Other methods are the same as those in FIGS. 1 to 3 and FIGS. 4A and 4B.
[0032]
The combustion method in this case is an inner surface combustion type, in which the reaction gas on the inner surface of each segment 1a is injected from the nozzle 3 to the outside through the gas passage 1b to obtain a thrust. In addition, according to this method, the liquid / gas propellant 2a having the optimum flow rate is injected from the upstream for each segment, so that the combustion temperature in the combustion chamber can be made substantially uniform, and the flame can be increased as in FIG. High temperature can be maintained.
[0033]
It should be noted that the present invention is not limited to the above-described embodiment, and it goes without saying that various changes can be made without departing from the spirit of the present invention.
[0034]
【The invention's effect】
As described above, the present invention has the following features.
1. By dividing the solid propellant into segments and supplying an appropriate liquid / gas propellant for each segment, the combustion temperature can be increased, and the performance (specific impulse) is improved.
2. The space of the liquid / gas injection part functions as a pre-combustion chamber, and the mixing of the liquid / gas propellant and the solid propellant is performed smoothly.
3. Since the performance (specific thrust) is improved, the amount of propellant can be reduced. Or you can increase the payload from this engine.
[0035]
Therefore, the solid motor of the hybrid rocket engine and the combustion method thereof according to the present invention can supply a liquid / gas propellant in which a part of the liquid propellant is gasified to almost the entire solid propellant at an almost optimum flow rate. Thus, the combustion temperature in the combustion chamber can be made substantially uniform, the flame temperature can be increased, and the characteristic exhaust speed of the nozzle and the specific thrust of the engine can be increased.
[Brief description of the drawings]
FIG. 1 is a diagram showing a first embodiment of a solid motor of a hybrid rocket engine according to the present invention.
FIG. 2 is a partially enlarged view of FIG.
FIG. 3 is a partially enlarged view showing another embodiment of FIG. 2;
FIG. 4 is a diagram showing another embodiment of the present invention.
FIG. 5 is a schematic view of a conventional solid motor.
[Explanation of symbols]
1 solid propellant, 1a segment, 1b gas passage,
2 liquid propellant, 2a liquid / gas propellant, 3 nozzles,
4 solid motor case, 5 injectors, 6 end face injectors,
7 gap injectors, 8 injection holes,
6a, 7a chamber,
6b, 7b partition,
9 internal injector, 9a chamber, 9b cylindrical tube,

Claims (11)

A hollow cylindrical solid motor case (4) having a nozzle (3), and a solid propellant (1) comprising a plurality of segments (1a) stored in the solid motor case with a gap in the axial direction; An end injector (6) for injecting a liquid / gas propellant (2a) toward an end face opposite to a nozzle of the solid propellant, and a gap injector for injecting a liquid / gas propellant into a gap between the segments. (7) A solid motor for a hybrid rocket engine, wherein a liquid / gas propellant is reacted with a solid propellant. The solid motor according to claim 1, wherein the plurality of segments (1a) have gas passages (1b) alternately penetrating in the axial direction. The solid motor according to claim 1, wherein the plurality of segments (1a) have gas passages (1b) penetrating in the axial direction at the same position. The end face injector (6) and the gap injector (7) are composed of a chamber (6a, 7a) to which a liquid / gas propellant (2a) is supplied, and a partition (6b, 7b) separating the chamber and the combustion chamber. The solid motor according to claim 1, wherein the partition wall is provided with a plurality of injection holes (8) or a mesh. A hollow cylindrical solid motor case (4) having a nozzle (3), and a solid propellant (1) comprising a plurality of segments (1a) stored in the solid motor case with a gap in the axial direction; An inner injector (9) for injecting a liquid / gas propellant (2a) toward the inner surface of each segment, wherein the liquid / gas propellant and the solid propellant are reacted with each other; Solid motor. The inner injector (9) comprises a chamber (9a) to which a liquid / gas propellant (2a) is supplied, and a cylindrical tube (9b) extending from the chamber to the inside of each segment in the axial direction. 6. The solid motor according to claim 5, wherein the tube is provided with a plurality of injection holes (8). A solid propellant (1) composed of a plurality of segments (1a) is stored in a hollow cylindrical solid motor case (4) having a nozzle (3) with a gap in the axial direction, and a liquid / gas propellant (2a) is stored. ) Is distributed to the gap between each segment and the end face opposite to the nozzle of the solid propellant and injected, and each segment reacts with the liquid / gas propellant at an almost optimum mixing ratio. Burning method. A solid propellant (1) composed of a plurality of segments (1a) is stored in a hollow cylindrical solid motor case (4) having a nozzle (3) with a gap in the axial direction, and a liquid / gas propellant (2a) is stored. ) Is distributed to the inner surface of each segment and injected, and each segment reacts with the liquid / gas propellant at a substantially optimum mixing ratio. The flow rate of the liquid / gas propellant (2a) injected to the upstream side of each segment is calculated by the following equation: total liquid propellant flow rate / total solid propellant weight = liquid / gas propellant flow rate per segment / solid propellant weight per segment. The method according to claim 7, wherein the fuel is distributed so as to be as follows. 10. The combustion method according to claim 9, wherein the ignition energy of each segment uses combustion gas of a more upstream segment. The method according to claim 10, characterized in that the injection timing of the liquid / gas propellant (2a) to the downstream segment is shifted.

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