JP2007239473A - Rocket engine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a rocket engine of which a fuel is easy to be handled at low cost, not requiring a special cooling structure for a combustion chamber and reducing cost. <P>SOLUTION: The rocket engine 1 is supplied with a propellant and an oxidant into a propelling chamber 10. Liquefied natural gas 41 is used as the propellant for the rocket engine 1. An ablator 31 is included on the inner surface side of a combustion chamber 12 of the propelling chamber 10. The combustion chamber 12 is used as an ablator combustion chamber. The ablator 31 forms a decomposed gas layer getting endothermically decomposed at a high temperature and having a heat insulating property and cools a high temperature part facing combustion gas 35 produced in the combustion chamber 12. Liquefied natural gas 41 creates film cooling 33 in the combustion chamber 12 to cool the combustion chamber 12. <P>COPYRIGHT: (C)2007,JPO&INPIT

Description

  The present invention relates to a rocket engine, and more particularly to a rocket engine that uses liquefied natural gas as a propellant and that does not require a special cooling structure in a combustion chamber.

As an engine used as the upper stage of a rocket, since high reliability is required, a regenerative cooling rocket engine using liquid hydrogen as a propellant has become mainstream. (For example, refer nonpatent literature 1).
FIG. 4 shows a conventional regenerative cooling rocket engine, and FIG. 5 shows a conventional regenerative cooling combustion chamber. FIG. 6 shows a cross-sectional structure taken along line AA in FIG.

  As shown in FIG. 4, in the regenerative cooling rocket engine, liquid hydrogen and oxidant are supplied to the combustion chamber 53 of the thrust chamber 52 through pipes 55 and 56 by driving of pumps 50 and 51, respectively. When liquid hydrogen is supplied to the combustion chamber 53, the liquid hydrogen is gasified by receiving the radiant heat of the thrust chamber 52 by passing a pipe 55 installed around the outside of the thrust chamber 52. Is sent into the combustion chamber 53.

In the combustion chamber 53 of the thrust chamber 52 shown in FIG. 5, a coolant channel 54 is provided in the wall portion of the combustion chamber 53 as shown in FIG. The cooling flow path 54 functions as the piping 55 in FIG. 4, and the liquid hydrogen sent from here cools the combustion chamber 53 as a cooling propellant.
Japan Aerospace Society, “Second edition Aerospace Engineering Handbook” Maruzen Co., Ltd., September 30, 1992, p. 925

However, in the regenerative cooling type engine adopting such a structure, liquid hydrogen used is a cryogenic propellant of about −253.16 ° C. (20K) although performance is high. Is very difficult to handle and the production cost of liquid hydrogen is high.
Further, since it is necessary to form the complicated coolant flow path 54 by machining as described above, the combustion chamber does not have to be a complicated structure having a special cooling structure. In other words, the development and manufacturing costs of the thrust chamber, including the combustion chamber, were high.

  Accordingly, the present invention has been made paying attention to the above-mentioned conventional problems, and provides a rocket engine that is easy to handle and inexpensive, and that does not require a special cooling structure in the combustion chamber and can reduce costs. It is an object.

  In the rocket engine of the present invention, in the rocket engine in which propellant and oxidant are supplied into the thrust chamber, liquefied natural gas is used as the propellant, and the combustion chamber of the thrust chamber has an ablator. It is characterized by. In this case, liquefied natural gas is easier to handle and cheaper than conventional liquid hydrogen, and a special cooling structure is not required in the combustion chamber, so that the cost can be reduced.

  The rocket engine according to the present invention is preferably characterized in that the ablator is formed inside a metal outer cylinder of the combustion chamber. In this case, the ablator is endothermically decomposed at a high temperature by solid sublimation (vaporization) to form a decomposition gas layer having thermal insulation, and a high temperature facing the liquefied natural gas combustion gas formed in the combustion chamber. By cooling the portion, the combustion chamber can be reliably cooled.

  The rocket engine of the present invention preferably has a pump for sending the liquefied natural gas to the combustion chamber of the thrust chamber through a pipe, and the pipe does not bypass the outer peripheral portion of the thrust chamber. It is connected to the combustion chamber. In this case, since the liquefied natural gas can be sent directly to the combustion chamber of the thrust chamber without passing through the bypassed pipe, the structure of the liquefied natural gas supply system can be simplified and the rocket engine can be downsized.

  The rocket engine of the present invention is preferably characterized in that a part of the liquefied natural gas is supplied along the inner surface of the combustion chamber to cool the combustion chamber with a film. In this case, the combustion chamber of the thrust chamber is cooled by the ablator, and the combustion chamber of the thrust chamber can be reliably cooled using liquefied natural gas.

  According to the rocket engine of the present invention, the fuel is easy to handle and inexpensive, and a special cooling structure is not required in the combustion chamber, so that the cost can be reduced.

EXAMPLES Hereinafter, although an Example demonstrates this invention still in detail, this invention is not limited to a following example.
FIG. 1 shows a preferred embodiment of the rocket engine of the present invention. A rocket engine 1 in FIG. 1 is an ablator-cooled liquefied natural gas engine, for example, used as an upper stage engine of a rocket.

  As shown in FIG. 1, the rocket engine 1 is a liquid rocket engine that uses liquefied natural gas (LNG) 41 as fuel (propellant). The rocket engine 1 supplies the liquefied natural gas 41 and the oxidant 42 to the combustion chamber 12 of the thrust chamber 10 that is a thrust chamber, burns at high pressure, and jets the combustion gas from the nozzle 13 to obtain thrust.

  The rocket engine 1 is an engine that is easy to handle fuel and inexpensive, and does not require a special cooling structure in the combustion chamber 12 and can reduce costs. The rocket engine 1 in FIG. 1 has a fuel tank 2, an oxidant tank 3, a supply system 40, and a thrust chamber (thrust chamber) 10.

  The fuel tank 2 in FIG. 1 stores liquefied natural gas 41 as a propellant. The liquefied natural gas 41 is obtained by cooling and compressing natural gas mainly composed of methane, which is also called liquefied methane gas (LMG), and has a temperature of −150 ° C. to −160 ° C.

  Although liquid hydrogen used as a normal fuel has high performance, it is a cryogenic propellant that remains in a liquid state only at an extremely low temperature state, so that it is very difficult to handle liquid hydrogen. In order to gasify liquid hydrogen, a configuration in which liquid hydrogen piping is installed along the outer shape portion of the thrust chamber, and radiant heat from the thrust chamber is applied to the liquid hydrogen to gasify the liquid hydrogen. It is. In storage of liquid hydrogen, it is necessary to pay attention to thermal insulation of the tank to reduce evaporation loss. And the production cost of liquid hydrogen is high.

  In contrast, the liquefied natural gas used in the embodiments of the present invention has a lower handling performance than liquid hydrogen, but the handling temperature is high. Therefore, the liquefied natural gas can be supplied directly into the combustion chamber through the piping. easy. Moreover, liquefied natural gas is inexpensive because it is generally used in city gas and the like.

  The oxidant tank 3 in FIG. 1 stores an oxidant 42. As the oxidant 42, for example, liquid oxygen (LOX) can be employed. The liquefied natural gas 41 in the fuel tank 2 and the oxidant 42 in the oxidant tank 3 are supplied to the combustion chamber 12 side of the thrust chamber 10 by a fuel supply system 40 described below.

The fuel supply system 40 shown in FIG. 1 will be described.
The fuel supply system 40 is disposed between the fuel tank 2, the oxidant tank 3, and the thrust chamber 10.
1 includes a fuel supply pump 4, an oxidant supply pump 5, turbines 6, 7, a gas generator 8, pipes 20, 21, 22, 23, 24, 25, 26, 27, and a turbine exhaust. It has a tube 28.

  The fuel supply pump 4 and the turbine 6 are connected by a shaft 6S, and the oxidant supply pump 5 and the turbine 7 are connected by a shaft 7S. The fuel supply pump 4, the turbine 6 and the shaft 6S constitute a fuel turbo pump (FTP) which is a propellant supply pump, and the oxidant supply pump 5, the turbine 7 and the shaft 7S are oxidation which is an oxidant supply pump. An agent turbo pump (OTP) is configured.

In FIG. 1, the pipe 20 connects the fuel supply pump 4 and the gas generator 8, and the pipe 21 connects the oxidant supply pump 5 and the gas generator 8.
The pipe 22 connects the fuel supply pump 4 and the injector 11 of the thrust chamber 10 directly, preferably so as to be as short as possible. Unlike the conventional rocket engine piping, the piping 22 for supplying the liquefied natural gas 41 does not need to be arranged along the outer portion of the thrust chamber 10. That is, one end 22B of the pipe 22 is connected to the fuel supply pump, one end 22C of the pipe 22 is connected to the injector 11, and a midway portion 22D of the pipe 22 needs to be arranged along the outer shape of the thrust chamber 10. There is no.

Similarly, the pipe 23 connects the oxidant supply pump 5 and the injector 11 of the thrust chamber 10 directly, preferably with a length as short as possible.
Since the piping 22 can be shortened compared to the conventional case, the structure of the supply system 40 can be simplified, and the rocket engine 1 can be downsized.
The pipe 24 is disposed between the gas generator 8 and the turbine 6, and the pipes 26 and 27 are disposed between the turbines 6 and 7 and the turbine exhaust pipe 28. The end of the turbine exhaust pipe 28 is inserted into the nozzle 13.

The liquefied natural gas 41 in the fuel tank 2 in FIG. 1 is supplied to the gas generator 8 through the fuel supply pump 4 and the pipe 20, and the oxidant 42 in the oxidant tank 3 is supplied through the oxidant supply pump 5 and the pipe 21. The gas generator 8 is supplied.
The gas generator 8 shown in FIG. 1 generates a liquefied natural gas 41 and an oxidant 42 injection gas to the turbines 6 and 7 in order to rotationally drive the turbines 6 and 7.
When the gas generator 8 injects the injection gas to the turbines 6 and 7 via the pipes 20 and 21, the turbines 6 and 7 rotate along the directions of the arrows, so that the fuel supply pump 4 and the oxidant supply pump 5 is activated.

When the fuel supply pump 4 of FIG. 1 is operated, the liquefied natural gas 41 in the fuel tank 2 is directly supplied to the injector 11 side through the pipe 22, and when the oxidant supply pump 5 is operated, The oxidant 42 is directly supplied to the injector 11 side through the pipe 23.
The injection gas supplied from the gas generator 8 to the turbines 6 and 7 via the pipes 20 and 21 is supplied into the nozzle 13 of the thrust chamber 10 via the pipes 26 and 27 and the turbine exhaust pipe 28. The injection gas 44 is effectively used as fuel.

Next, a structural example of the thrust chamber 10 shown in FIG. 1 will be described.
The thrust chamber (thrust chamber) 10 is a so-called ablator thrust chamber having an ablator inside, and the thrust chamber 10 has a bell shape in the illustrated example. The thrust chamber 10 has an injector 11, a combustion chamber 12, and a nozzle 13.
1 shows the appearance of the injector 11, the combustion chamber 12, and the nozzle 13, while FIG. 2 shows an example of a cross-sectional structure at a portion X of the combustion chamber 12 and the nozzle 13 in FIG. FIG. 3 shows an example of a cross-sectional structure along the axial direction of the combustion chamber 12 and the nozzle 13.

As shown in FIGS. 1 to 3, the combustion chamber 12 and the trumpet type nozzle 13 are formed continuously, and a throat portion 46 is formed in the middle of the thrust chamber 10.
As shown in FIG. 1, in this embodiment, as an example, a range related to the axial direction of the combustion chamber 12 is indicated by a symbol H, and a range related to the axial direction of the nozzle 13 is indicated by a symbol M. The injector 11 is disposed at the end of the combustion chamber 12.

  The thrust chamber 10 shown in FIGS. 2 and 3 has a metal outer cylinder 30. An ablator 31 is formed on the inner surface side of the metal outer cylinder 30 of the combustion chamber 12. In the example shown in FIGS. 2 and 3, the ablator 31 is preferably formed inside the combustion chamber 12. The combustion chamber 12 can also be referred to as an ablator combustion chamber, and a combustion gas 35 of liquefied natural gas 41 is formed in the combustion chamber 12 as shown in FIG.

  As a material of the metal outer cylinder 30, a material having a large strength / specific gravity value, for example, stainless steel, a cobalt alloy, a nickel alloy, or the like can be adopted in order to reduce the weight.

  As the ablator 31 formed so as to cover the inner surface of the metal outer cylinder 30, an ablation material that forms a decomposition gas layer having a thermal insulation property by endothermic decomposition at a high temperature by solid sublimation (vaporization), for example, graphite, Silica, resin, GFRP (glass fiber reinforced plastic) and the like can be employed.

The injector 11 shown in FIG. 3 injects the liquefied natural gas 41 and the oxidant 42 that are pressurized and sent in the supply system 40 shown in FIG. 1 into the combustion chamber 12. In this case, the liquefied natural gas 41 is preferably injected as film cooling 33 into the combustion chamber 12 using the injector 11, as shown in FIG.
The film cooling 33 actively cools the surrounding portion of the combustion gas 35 in the combustion chamber 12 using the liquefied natural gas 41 as a cooling medium. Thereby, since it is not necessary to provide a complicated special cooling structure part for flowing the cooling fluid in the combustion chamber 12, the manufacturing cost of the combustion chamber 12 can be reduced and the weight can be reduced, or the thrust including the combustion chamber 12 can be reduced. The manufacturing cost of the chamber 10 can be reduced and the weight can be reduced.

Next, an operation example of the rocket engine 1 configured as described above will be described.
The liquefied natural gas 41 in the fuel tank 2 in FIG. 1 is supplied to the gas generator 8 through the fuel supply pump 4 and the pipe 20, and the oxidant 42 in the oxidant tank 3 is supplied through the oxidant supply pump 5 and the pipe 21. The gas generator 8 is supplied.
The gas generator 8 of FIG. 1 supplies the liquefied natural gas 41 and the injection gas of the oxidant 42 to the turbines 6, 7. The agent supply pump 5 operates.

When the fuel supply pump 4 of FIG. 1 is operated, the liquefied natural gas 41 in the fuel tank 2 is directly supplied to the injector 11 side through the pipe 22, and when the oxidant supply pump 5 is operated, the oxidant in the oxidant tank 3 is operated. 42 is directly supplied to the injector 11 side through the pipe 23.
The injection gas supplied to the turbines 6 and 7 by the gas generator 8 through the pipes 20 and 21 is supplied into the nozzle 13 of the thrust chamber 10 through the pipes 26 and 27 and the turbine exhaust pipe 28. By being supplied, the injection gas 44 is effectively used as fuel.

The injector 11 shown in FIG. 3 injects the liquefied natural gas 41 and the oxidant 42 that are pressurized in the supply system 40 shown in FIG. 1 into the combustion chamber 12. The liquefied natural gas and the oxidant are injected into the combustion chamber 12 in a liquid state, but the liquefied natural gas and the oxidant are mixed after evaporation to cause a combustion reaction in the gas phase.
In this case, a part of the liquefied natural gas 41 is preferably injected as film cooling 33 into the combustion chamber 12 as shown in FIG. The film cooling 33 can positively cool the surrounding portion of the combustion gas 35 in the combustion chamber 12 by using the liquefied natural gas 41 and the oxidant 42 as a cooling medium.

  As illustrated in FIG. 2, the ablator 31 formed so as to cover the inner surface of the metal outer cylinder 30 is subjected to endothermic decomposition at a high temperature by solid sublimation (vaporization) to form a decomposition gas layer having thermal insulation properties. Thus, the high-temperature portion facing the combustion gas 35 formed in the combustion chamber 12 can be cooled. FIG. 2 shows a state in which the ablator 31 releases the decomposition gas by solid sublimation by a plurality of arrows 39. The thickness of the ablator 31 is reduced to the extent that it does not affect the propulsion performance.

  As described above, in the rocket engine of the preferred embodiment of the present invention, liquefied natural gas is used as the fuel, and the temperature when used is higher than that of conventional liquid hydrogen, and the fuel is easy to handle and inexpensive. It is. Further, in the preferred embodiment of the rocket engine of the present invention, the combustion chamber 12 has an ablator 31, and the combustion chamber 12 does not require a special cooling structure that is conventionally required. The cost for manufacturing the combustion chamber 12 can be reduced.

  In the preferred embodiment of the rocket engine of the present invention, the ablator 31 is preferably formed inside the metal outer cylinder 30 of the combustion chamber 12. In this case, the ablator 31 is endothermically decomposed at a high temperature by solid sublimation (vaporization) to form a decomposition gas layer having thermal insulation, and a high temperature facing the combustion gas 35 formed in the combustion chamber 12. By cooling the portion, the inside of the combustion chamber 12 can be reliably cooled.

  The rocket engine of the present invention preferably has a fuel supply pump 4 for sending liquefied natural gas 41 to the combustion chamber 12 of the thrust chamber 10 through a pipe, and the pipe 22 bypasses the outer periphery of the thrust chamber 10. Without being connected to the combustion chamber 12. In this case, since the temperature of the liquefied natural gas 41 is higher than that of the liquid hydrogen, the liquefied natural gas 41 is easily gasified, and the liquefied natural gas 41 can be sent directly to the combustion chamber 12 of the thrust chamber 10 without passing through the bypassed pipe. The structure of the supply system 40 of the gas 41 can be simplified, and the rocket engine 1 can be downsized.

  The rocket engine of the present invention is preferably cooled by supplying a part of the liquefied natural gas 41 along the inner wall surface of the combustion chamber 12 of the thrust chamber 10 and flowing the inside of the combustion chamber 12 as the film cooling 33. . In this case, the inner wall surface of the combustion chamber 12 is cooled by the ablator 31 and the liquefied natural gas 41 is used to reliably cool the vicinity of the inner wall surface of the combustion chamber 12. Thereby, the cooling effect of the combustion chamber 12 can be further enhanced and the cooling performance of the rocket engine 1 can be improved.

By the way, the present invention is not limited to the above-described embodiments, and various modifications can be considered without departing from the scope of the claims.
For example, in the example of FIG. 1, the supply system 40 has the gas generator 8, and the fuel pumps 4, 5 are liquefied by rotating the turbines 6, 7 with the gas generated by the gas generator 8. Natural gas 41 and oxidant 42 are supplied to the combustion chamber 12.

However, the present invention is not limited thereto, and the gas generator 8 is not used, and the fuel pumps 4 and 5 are liquefied natural by rotating the turbines 6 and 7 by returning the vaporized fuel to the turbines 6 and 7 from the injector 11 side. The gas 41 and the oxidant 42 may be supplied to the combustion chamber 12.
Further, without using the gas generator 8, by returning a part of the combustion gas in the combustion chamber from the combustion chamber 12 side to the turbines 6 and 7 and rotating the turbines 6 and 7, the fuel pumps 4 and 5 are The liquefied natural gas 41 and the oxidant 42 may be supplied to the combustion chamber 12.

  In the example of FIGS. 1 to 3, the ablator 31 is formed in the metal outer cylinder 30 of the combustion chamber 12. However, the ablator 31 may be formed in the metal outer cylinder 30 of the combustion chamber 12 and in the metal outer cylinder 30 of the nozzle 13 continuing to the combustion chamber 12. In any case, the ablator 31 may be formed at least in a portion corresponding to the periphery of the combustion gas 35 shown in FIG.

Further, the ablator 31 is provided only in the metal outer cylinder 30 of the combustion chamber 12, in the metal outer cylinder 30 of the combustion chamber 12 and a part of the metal outer cylinder 30 of the nozzle 13, and in the metal outer cylinder 30 of the combustion chamber 12. And the nozzle 13 may be formed over the entire metal outer cylinder 30.
In FIG. 1, as an example, the range of the combustion chamber 12 is indicated by the symbol H, and the range of the nozzle 13 is indicated by the symbol M. However, the range of the combustion chamber and the nozzle range are not limited to the illustrated example. Can be arbitrarily selected. The rocket engine of the embodiment of the present invention is not limited to the upper stage of the rocket, but may be used in the middle stage or the lower stage.

It is a figure which shows the structure of the preferable Example of the rocket engine of this invention. It is sectional drawing which shows the structural example of the part X of the thrust chamber of the rocket engine of FIG. It is a figure which shows the combustion chamber and nozzle of the thrust chamber of the rocket engine of FIG. 1, and the combustion gas in a combustion chamber. It is a figure which shows the conventional regenerative cooling type rocket engine. It is a figure which shows the conventional regenerative cooling combustion chamber. It is a figure which shows the cross-section of a combustion chamber in the AA line of FIG.

Explanation of symbols

1 Rocket Engine 2 Fuel Tank 3 Oxidant Tank 4 Fuel Supply Pump 5 Oxidant Supply Pump 6 Turbine 7 Turbine 8 Gas Generator 10 Thrust Chamber (Thrust Chamber)
11 Injector 12 Combustion chamber (ablator combustion chamber)
13 Nozzle 20 Piping 21 Piping 22 Piping 23 Piping 24 Piping 26 Piping 27 Piping 28 Turbine exhaust pipe 30 Metal outer cylinder 31 Ablator 33 Film cooling 35 Combustion gas 40 Supply system 41 Liquefied natural gas (fuel)
42 Oxidant 44 Injection gas 46 Throat portion 50 Pump 51 Pump 52 Thrust chamber 53 Combustion chamber 54 Coolant flow path 55 Piping 56 Piping

Claims (4)

In rocket engines where propellant and oxidizer are supplied into the thrust chamber,
A rocket engine using liquefied natural gas as the propellant and having an ablator in the combustion chamber of the thrust chamber.   The rocket engine according to claim 1, wherein the ablator is formed inside a metal outer cylinder of the combustion chamber.   A pump for sending the liquefied natural gas to the combustion chamber of the thrust chamber through a pipe, and the pipe is connected to the combustion chamber without bypassing the outer peripheral portion of the thrust chamber; The rocket engine according to claim 1 or 2.   4. The rocket engine according to claim 3, wherein a part of the liquefied natural gas is supplied along the inner surface of the combustion chamber to cool the combustion chamber with a film.

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