JP2670253B2 - Rocket engine injectors - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

The present invention relates to an injector used in a combustion chamber of a rocket engine.

[0002]

2. Description of the Related Art FIG. 3 shows a conventional rocket engine injector. Reference numeral 3 denotes an outer wall of a metal injector having a circular cross section. Inside the outer wall 3, a partition wall 4 made of a metal having a circular section is arranged concentrically with the outer wall 3, and an arrow is drawn in the partition wall 4. As shown, the liquid oxidizer 2 is supplied, and the gas fuel 1 that has passed through the combustion chamber of the rocket engine and the cooling passage of the rocket nozzle is supplied between the partition wall 4 and the outer wall 3 as shown by the arrow. . The ends of the outer wall 3 and the partition wall 4 are open to the combustion chamber of the rocket engine, and the outer wall 3 forms an injection hole, and the gas fuel 1 and the liquid oxidizer 2 are injected into the combustion chamber through this injection hole. . In addition, 5 is a wall of the injection chamber to which the outer wall 3 is attached.

[0003]

In the conventional rocket engine injector shown in FIG. 3, the outer wall 3 of the injector is used.
The size of the injection hole formed in is fixed. When changing the thrust of the rocket engine, there are a method of changing the combustion pressure and a method of changing the mixing ratio of the liquid oxidant and the gas fuel (oxidant mass flow rate / fuel mass flow rate).

When increasing the combustion pressure, the flow rates of the liquid oxidizer and the gas fuel injected into the combustion chamber are increased with the mixing ratio of the liquid oxidizer and the gas fuel kept constant. When increasing the mixing ratio of the liquid oxidant and the gas fuel, the flow rate of the liquid oxidant injected into the combustion chamber is increased and the flow rate of the gas fuel is decreased.

In the conventional injector shown in FIG. 3 in which the size of the injection hole is fixed, when the combustion pressure and the mixing ratio deviate greatly from the design point, the reason is as follows.
Significant performance degradation occurs.

When the combustion pressure is changed, if it is lowered too much, the injection flow velocity of the gas fuel and the liquid oxidizer becomes slow, and combustion oscillation occurs.

On the other hand, if the combustion pressure is too high, the following problem occurs. That is, as shown in FIG. 4A, when the combustion pressure P _c is increased, the gas fuel density ρ _g increases in proportion to the combustion pressure P _c , but the liquid oxidant density ρ _l increases. It takes a constant value regardless of P _c . In order to increase the combustion pressure P _c , it is necessary to increase the gas fuel flow rate and the liquid fuel flow rate, and at this time, the following equation 1 is established.

[0008]

(Equation 1)

In the case of a fixed injection hole, since A is constant, Equation 2 holds. This relationship is shown in FIG.

[0010]

(Equation 2)

Therefore, the injection flow rate ratio (V _g / V _l ) of gas fuel and liquid oxidant (hereinafter referred to as the injection flow rate ratio) is shown in FIG.
As shown in (b), the combustion pressure P _c increases (1 → 2)
And tends to be low. In a rocket engine injector, the characteristic exhaust velocity efficiency (η) is shown in FIG. 4 because the mixing becomes slower as the injection velocity ratio (V _g / V _l ) becomes lower.
It becomes lower as shown in (c).

To summarize the above, as the combustion pressure (P _c ) rises, the density of the gas fuel (ρ _g ) increases, and the density ρ _l of the liquid oxidant remains constant, so that the combustion pressure (P _c ) Increase → Injection velocity ratio ( _Vg / _Vl ) decrease → Characteristic pumping speed efficiency η
The relationship of decline is established.

In the case of changing the mixing ratio, if the injection flow velocity ratio of the gas fuel and the liquid oxidizer is largely changed from the design point, the mixing becomes slow, and the characteristic exhaust velocity efficiency also becomes low.

The above problems are caused by changes in the injection flow velocities of the liquid oxidizer and the gas fuel when changing the thrust of the rocket engine.

The present invention is intended to provide a rocket engine injector which can solve the above problems.

[0016]

[Means for Solving the Problems]

(1) The present invention provides an injector for a rocket engine, which is configured to inject into the combustion chamber the gas fuel that has flowed around the liquid oxidizer injected into the combustion chamber of the rocket engine through the cooling passages of the combustion chamber and the rocket nozzle. The portion of the outer wall of the injector where the injection hole is formed is made of a shape memory alloy that deforms so as to reduce the size of the injection hole during high pressure combustion.

(2) Further, the present invention is a rocket engine in which the gas fuel flowing through the cooling passage of the combustion chamber and the rocket nozzle from around the liquid oxidizer injected into the combustion chamber of the rocket engine is injected into the combustion chamber. In the injector of
The portion of the outer wall of the injector that forms the injection hole was made of a shape memory alloy that deforms so as to reduce the injection hole when the liquid oxidizer and the gas fuel are highly mixed.

In the present inventions (1) and (2),
Since the gaseous fuel is injected into the combustion chamber from around the liquid oxidant, the fuel flows along the outer wall in the portion of the outer wall of the injector forming the injection hole.

In the present invention (1), when the combustion pressure in the combustion chamber is changed and raised, the liquid oxidation is injected into the combustion chamber with the mixing ratio of the liquid oxidizer and the gas fuel kept constant. Increase the flow rates of agent and gas fuel. Since the gas fuel injected into the combustion chamber passes through the combustion chamber of the rocket engine and the cooling passages of the rocket nozzle to cool these parts before that, the gas fuel passing through the injection holes of the injector increases as the flow rate increases. The temperature of the fuel decreases, and the temperature of the shape memory alloy decreases accordingly. The shape memory alloy is deformed by the decrease in the temperature at the time of high pressure combustion as a trigger to reduce the injection hole. As a result, the injection flow velocity of the gas fuel is increased, and the proper injection flow velocity ratio is maintained.

In the present invention (2), when the mixing ratio (liquid oxidant mass flow rate / gas fuel mass flow rate) of the liquid oxidizer and the gas fuel injected into the combustion chamber is changed and increased, the combustion is performed. The flow rate of the liquid oxidant injected into the chamber is increased and the flow rate of the gas fuel is reduced. The gas fuel injected into the combustion chamber cools these parts through the cooling chamber of the rocket engine and the cooling passage of the rocket nozzle before that, so when the flow rate decreases, the gas fuel that passes through the injection holes of the injector The temperature of the fuel rises, and with this, the shape memory metal is deformed by the temperature rise at the time of a high mixing ratio of the liquid oxidizer and the gas fuel as a trigger to reduce the injection hole. Thereby, the injection flow rate of the gas fuel and the liquid oxidant is increased, and an appropriate injection flow rate ratio is maintained.

The shape memory alloys in the present inventions (1) and (2) are nickel-based or copper-based alloys (TiNi alloys, Cu-Zn-A) used as combustor materials.
1 alloy, Cu-Al-Ni alloy, etc.) can be used.

[0022]

DESCRIPTION OF THE PREFERRED EMBODIMENTS One embodiment of the present invention will be described with reference to FIGS. In this embodiment, the injector of the conventional rocket engine shown in FIG. 3 is improved as follows, and the parts that are not changed are the same as in FIG.
The same reference numerals as in the above denote the same parts, and a description thereof will be omitted.

That is, in this embodiment, the outer wall 3 of the injector is
Part 3a forming an injection hole opening at least in the combustion chamber
Was composed of a nickel-based, copper-based, etc. shape memory alloy. In the case of changing the combustion pressure, when the temperature is low, this portion 3a has a portion having a "f" -shaped cross section and a narrowed inner diameter as shown in FIG. 1, and constitutes this portion 3a. When the temperature rises to a constant value, the shape memory alloy deforms outward so as to enlarge the injection hole.

On the other hand, in the case of changing the mixing ratio, when the temperature is high, as shown in FIG. 1, there is a section with a "concave" cross section and the inner diameter is reduced, and the temperature drops to a constant value. Then, the injection hole is deformed outward so as to enlarge.

In the present embodiment, when the combustion pressure in the combustion chambers is increased in order to increase the thrust of the rocket engine, the mixing ratio of the liquid oxidizer 2 and the gas fuel 1 is kept constant and the fuel is injected into those combustion chambers. Increase the flow rate.
Before the gas fuel 1 is cooled through the combustion chamber of the rocket engine and the cooling passage of the rocket nozzle, the flow rate of the gas fuel 1 injected from the injection hole of the injector into the combustion chamber is increased. Then, the temperature also decreases, and the temperature of the shape memory alloy of the portion 3a forming the injection hole also decreases accordingly. The shape memory alloy is deformed by the temperature change as a trigger to reduce the size of the injection hole.

As a result, as described in detail in the [Problems to be solved by the invention] and [Means for solving the problems], even if the flow rates of the gas fuel 1 and the liquid oxidizer 2 are changed, the injection is performed. It is possible to maintain the speed ratio at an appropriate value, prevent combustion oscillation, improve the characteristic exhaust speed efficiency, and improve the performance of the rocket engine.

In the above-described embodiment of the present invention, the combustion pressure of the rocket engine is changed. However, the mixing ratio (oxidant mass flow rate / fuel mass flow rate) of the liquid oxidant and the gas fuel of the rocket engine is changed. When changing, the following shape memory alloys are adopted.

That is, the portion 3a of the outer wall 3 forming the injection hole
When the temperature is high, the shape-memory alloy that constitutes is has a portion with a narrowed inner diameter with a cross-section as shown in FIG. 1, and when the temperature drops to a certain value, the injection hole expands. It is designed to be deformed outward as if it were.

When increasing the mixing ratio of the liquid oxidizer 2 and the gas fuel 1 of the rocket engine, the flow rate of the liquid oxidant is increased and the flow rate of the gas fuel is decreased. The temperature of the gas fuel that has cooled the combustion chamber of the rocket engine and the rocket nozzle rises as the flow rate decreases, and the temperature of the shape memory alloy also rises accordingly. The shape memory alloy is deformed by the temperature rise at the time of the high mixing ratio as a trigger to reduce the injection hole. As a result, even if the flow rate of the gas fuel 1 is reduced, the proper injection flow velocity can be maintained, combustion oscillation can be prevented, characteristic exhaust velocity efficiency can be improved, and rocket engine performance can be improved. .

FIG. 2 is a diagram comparing the combustion performance of the rocket engine equipped with the injector shown in FIG. 3 with this embodiment. In the fixed type conventional injector shown in FIG. 3, the characteristic exhaust velocity efficiency decreases as the combustion pressure and the mixing ratio increase as shown by the line 7, but in the present embodiment, the line 8
As shown in 6 and 6, the characteristic exhaust velocity efficiency is once improved at the deformation point of the shape memory alloy (point indicated by the line 8), so that the overall performance can be improved.

[0031]

As described above, according to the present invention, even if the combustion pressure or the mixing ratio is changed to change the thrust of the rocket engine, the injection flow velocity ratio of the gas fuel and the liquid oxidizer is maintained at an appropriate value. In addition, it is possible to prevent the occurrence of combustion vibration and the reduction of the characteristic exhaust speed efficiency, and to improve the performance of the rocket engine.

[Brief description of the drawings]

FIG. 1 is a sectional view of an embodiment of the present invention.

FIG. 2 is a diagram showing the combustion performance of a rocket engine equipped with an embodiment of the present invention and a conventional injector.

FIG. 3 is a sectional view showing an injector of a conventional rocket engine.

FIG. 4 (a) is a combustion pressure (P _c ), a gas fuel density (ρ _g ), and a flow velocity (V) in the rocket engine injector.
_g ), a graph showing the relationship between the density (ρ _l ) of the liquid oxidizer and the flow velocity (V _l ), and FIG. 4 (b) shows the combustion pressure (P _c ) in the rocket engine injector and the injection flow velocity ratio (V _g / V _l) graph showing the relationship between, FIG. 4 (c) is a graph showing the relationship between the injection flow rate ratio in the rocket engine injector (V _g / V _l) with the characteristic exhaust velocity efficiency (eta).

[Explanation of symbols]

 1 Gas Fuel 2 Liquid Oxidizer 3 Outer Wall 3a Portion of Outer Wall Constituting Injection Hole 4 Partition Wall 5 Combustion Chamber Wall

 ─────────────────────────────────────────────────── ─── Continuation of the front page (72) Inventor Hideyuki Taguchi 1200, Higashitanaka, Komaki City, Aichi Prefecture Mitsubishi Heavy Industries, Ltd. Nagoya Induction Propulsion System Factory

Claims (2)

(57) [Claims] 1. An injector of a rocket engine configured to inject into the combustion chamber gas fuel that has flowed from around a liquid oxidizer injected into the combustion chamber of the rocket engine into the combustion chamber and the cooling passage of the rocket nozzle. An injector for a rocket engine, characterized in that a portion of the outer wall of the vessel forming the injection hole is made of a shape memory alloy that deforms so as to reduce the injection hole during high pressure combustion. 2. An injector of a rocket engine, wherein the gas fuel flowing through the combustion chamber and the cooling passage of the rocket nozzle from around the liquid oxidant injected into the combustion chamber of the rocket engine is injected into the combustion chamber. An injector for a rocket engine, characterized in that a portion of the outer wall of the vessel forming the injection hole is made of a shape memory alloy that is deformed so as to reduce the injection hole at a high mixing ratio of liquid oxidizer and gas fuel.