Expander bleed cycle rocket engine with aerospike nozzle

Abstract

To provide a rocket engine that maintains optimal performance even in an environment where atmospheric pressure changes by mounting an aerospike nozzle on an expander bleed cycle rocket engine. [Solution] A rocket engine equipped with a regenerative cooling nozzle, a turbine 9 that rotates with fuel vapor vaporized by the regenerative cooling nozzle, a fuel transport turbopump 10 driven by the turbine, an oxidizer transport turbopump 11 driven by the turbine, and a combustion chamber 5 in which the oxidizer and fuel are mixed and burned, wherein the regenerative cooling nozzle is shaped into an aerospike nozzle, and the problem is solved by a rocket engine that uses an expander bleed cycle in which the fuel vapor vaporized by the regenerative cooling nozzle is discharged outside the combustion chamber after use.

Description

[Technical field]

[0001] [Background technology]

[0002] In order to improve the propulsive performance of rocket engines, a component called a nozzle is used to control the speed and pressure of the combustion gas. For example, Patent Document 1 uses a bell nozzle.

[0003] One type of nozzle is an aerospike nozzle, which is arranged inside the combustion gas. For example, Patent Document 2 uses an aerospike nozzle.

[0004] As a thermal cycle for rocket engines, there is an expander cycle in which a turbine is rotated by fuel vapor vaporized by a regenerative cooling nozzle, an oxidizer and fuel are transported to a combustion chamber by a turbo pump and burned, and the fuel vapor after the turbine rotation is exhausted in a separate system. For example, Patent Document 3 is a rocket engine using an expander bleed cycle. [Prior art documents] [Patent documents]

[0005] [Patent Document 1] Patent No. 4885301 Patent Publication (B2) [Patent Document 2] Patent No. 7091386 Patent Publication (B2) [Patent Document 3] JP 2008-202542 Published Patent Application (A) Summary of the Invention [Problem to be solved by the invention]

[0006] The specific impulse, which indicates the performance of a rocket engine, is best when the nozzle exit pressure is equal to the pressure around the rocket engine.

[0007] The bell nozzle described in Patent Document 1 has the advantage that the nozzle is outside the combustion gas, making it easy to dissipate heat, but because the nozzle outlet pressure is determined by the nozzle shape, in an environment where the atmospheric pressure changes from moment to moment with altitude, it is not possible to select the optimal nozzle outlet pressure in accordance with the atmospheric pressure at the appropriate time, and the best specific impulse cannot be maintained.

[0008] The aerospike nozzle described in Patent Document 2 has the characteristic of automatically maintaining an optimal nozzle outlet pressure according to atmospheric pressure, but because the nozzle is located inside the combustion gas, it is difficult to dissipate heat and is susceptible to damage from heat.

[0009] In the expander bleed cycle rocket engine described in Patent Document 3, when the engine is enlarged to increase thrust, the volume of the combustion chamber expands in proportion to the cube of the engine size, but the surface area of ​​the regenerative cooling nozzle expands in proportion to the square of the engine size, so the amount of heat absorption by the regenerative cooling nozzle tends to be insufficient.

[0010] An object of the present disclosure is to provide an expander bleed cycle rocket engine equipped with an aerospike nozzle, which maintains optimal performance in an environment where atmospheric pressure changes. [Means for solving the problem]

[0011] A regenerative cooling nozzle; a turbine rotated by the fuel vapor vaporized by the regenerative cooling nozzle; a fuel transport turbopump driven by the turbine; an oxidant transport turbopump driven by the turbine; A combustion chamber in which an oxidizer and a fuel are mixed and burned; A rocket engine comprising: The regenerative cooling nozzle is shaped into an aerospike nozzle shape, An expander bleed cycle is used in which the fuel vapor vaporized by the regenerative cooling nozzle is discharged outside the combustion chamber after use. The solution is rocket engines. Effect of the Invention

[0012] According to the present disclosure, it is possible to provide a rocket engine that maintains optimal performance even in an environment where atmospheric pressure changes. [Brief description of the drawings]

[0013] [Figure 1] FIG. 1 is a cross-sectional view for explaining the configuration of a rocket engine 1. As shown in FIG. [Explanation of symbols]

[0014] 1. Rocket engine 2 Rotation Axis 3. Oxidizing Agent 4 fuel 5. Combustion chamber 6 Regenerative cooling nozzle end section 7 Regenerative cooling nozzle throat 8 Regenerative cooling nozzle divergent section 9. Turbine 10 Fuel transport turbopump 11. Oxidizer transport turbopump 12 Combustion gas edge line (atmospheric pressure) 13 Outer edge of combustion gas (low air density) 14 Combustion gas outer edge (vacuum) DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0015] Embodiments (examples) of the present disclosure will be described in more detail with reference to the drawings. FIG. 1 is a cross-sectional view of a rocket engine. The specifications of rocket engine 1 are as follows: Rocket engine 1 is an expander bleed cycle engine with a vacuum thrust of 150 tonnes, a total length of 3 m, a diameter of 2 m. The rocket engine 1 has a basic structure that is symmetrical about a rotation axis 2 . The oxidizer 3 is liquid oxygen, but may be dinitrogen tetroxide, hydrogen peroxide, nitrous oxide, or other liquid oxidizers. The fuel 4 is liquid hydrogen, but could also be liquid methane, hydrocarbons, methyl alcohol, or other liquid fuels. The combustion chamber 5 is made of steel, but may be made of other metals or carbon fiber composite materials. The combustion chamber 5 withstands the heat and pressure generated by the combustion of the fuel. The end portion 6 of the regenerative cooling nozzle is made of steel, but may be made of other metals or carbon fiber composites. The regenerative cooling nozzle end section 6 has a shape with a varying cross-sectional area that increases the velocity of the combustion gas generated in the combustion chamber 5 . The regenerative cooling nozzle throat 7 is made of steel, but may be made of other metals or carbon fiber composites. The regenerative cooling nozzle throat 7 has a shape in which the cross-sectional area changes so that the speed of the combustion gas passing through it becomes equal to the speed of sound. The regenerative cooling nozzle divergent portion 8 has a basic structure that is axially symmetrical with respect to the rotation axis 2 . The regenerative cooling nozzle divergent portion 8 is made of steel, but may be made of other metals or carbon fiber composites. The regenerative cooling nozzle divergent section 8 has a cross-sectional area change shape that reduces the pressure and increases the velocity of the combustion gas. The regenerative cooling nozzle divergent section 8 has an aerospike shape in which the nozzle is inside the combustion gas. The regenerative cooling nozzle divergent portion 8 has a fuel pipe joined to its inside and has the function of transferring heat of the combustion gas to the fuel. The turbine 9 rotates by the fluid energy of the fuel vapor that is vaporized by receiving heat from the combustion gas in the regenerative cooling nozzle divergent portion 8 . After rotating the turbine 9, the fuel vapor is discharged from the tip of the regenerative cooling nozzle divergent portion 8. The turbine 9 is disposed within the regeneratively cooled nozzle divergent section 8 . The axis of rotation of the turbine 9 coincides with the axis of symmetry of the rocket engine 1 . The fuel transport turbo pump 10 is driven by the turbine 9 to pressurize the fuel 4 and transport it to the combustion chamber 5 . A fuel delivery turbopump 10 is disposed within the regeneratively cooled nozzle divergent section 8 . The rotation axis of the fuel transport turbopump 10 coincides with the axis of symmetry of the rocket engine 1 . The oxidizer transport turbo pump 11 is driven by the turbine 9 to pressurize the oxidizer 3 and transport it to the combustion chamber 5 . The oxidizer transport turbopump 11 is disposed within the regenerative cooling nozzle divergent section 8 . The rotation axis of the oxidizer transport turbopump 11 coincides with the axis of symmetry of the rocket engine 1. The shape of the combustion gas during operation of the rocket engine 1 is optimized according to the atmospheric pressure to maintain the best specific impulse. The combustion gas outer edge (atmospheric pressure) 12 is the shape of the combustion gas when the pressure around the rocket engine 1 is atmospheric pressure. The combustion gas outer edge (atmospheric pressure) 13 is the shape of the combustion gas when the pressure around the rocket engine 1 is low air density. The combustion gas outer edge (in a vacuum) 14 is the shape of the combustion gas when the pressure around the rocket engine 1 is a vacuum.

[0016] In order to rotate the turbine 9, it is necessary to have part of the fuel absorb the heat of the combustion gas from the regenerative cooling nozzle divergent section 8. However, the shape of the regenerative cooling nozzle divergent section 8 makes it difficult to release the radiant heat of the combustion gas to the surrounding area, which is advantageous in that it is easy to secure the heat.

[0017] The turbine 9, fuel transport turbopump 10, and oxidizer transport turbopump 11 are installed by effectively utilizing the internal space of the regenerative cooling nozzle divergent portion 8, which is advantageous in shortening the overall engine length and reducing the structural weight.

[0018] The rotation axes of the turbine 9, fuel delivery turbopump 10, and oxidizer delivery turbopump 11 coincide with the axis of symmetry of the rocket engine 1, which provides good structural balance and is advantageous for reducing the structural weight.

[0019] Since the pressure near the tip of the regenerative cooling nozzle divergent section 8 is reduced to the same as the ambient pressure of the rocket engine 1, there is an advantage in that the exhaust pressure is small when the fuel vapor after rotating the turbine 9 is discharged from the tip of the regenerative cooling nozzle divergent section 8.

[0020] The uses of the rocket engine according to the present disclosure are not particularly limited, but it can be suitably used as a high-performance rocket engine in satellite launch rockets, rockets for travel in outer space, defense weapons, and the like.

[0021] The embodiments disclosed herein are illustrative in all respects and should not be construed as limiting. The scope of the present invention is defined by the claims, not the above description, and all modifications within the meaning and scope of the claims are intended to be embraced.

Claims

1. A regenerative cooling nozzle; a turbine rotated by the fuel vapor vaporized by the regenerative cooling nozzle; a fuel transport turbopump driven by the turbine; an oxidant transport turbopump driven by the turbine; A combustion chamber in which an oxidizer and a fuel are mixed and burned; A rocket engine comprising: The regenerative cooling nozzle is shaped into an aerospike nozzle shape, An expander bleed cycle is used in which the fuel vapor vaporized by the regenerative cooling nozzle is discharged outside the combustion chamber after use. Rocket engine.

2. 2. The rocket engine according to claim 1, wherein the turbine, the oxidizer transport turbopump, and the fuel transport turbopump are mounted inside the regenerative cooling nozzle when viewed from the side.

3. 2. The rocket engine according to claim 1, wherein the rotation axis of the turbine, the rotation axis of the oxidizer transport turbopump, and the rotation axis of the fuel transport turbopump are aligned with an axis of axisymmetric shape of the regenerative cooling nozzle.

4. 2. A rocket engine according to claim 1, wherein the fuel vapor after rotating the turbine is discharged from a tip of the regenerative cooling nozzle.

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