WO2015111253A1

WO2015111253A1 - Thruster

Info

Publication number: WO2015111253A1
Application number: PCT/JP2014/077030
Authority: WO
Inventors: 康智田中; 中村　武志; 雅人石崎
Original assignee: 株式会社Ihi
Priority date: 2014-01-22
Filing date: 2014-10-09
Publication date: 2015-07-30

Abstract

[Problem] A thruster having a main body using a ceramic base composite material, that suppresses contamination caused by leaked gas leaking from micro-cracks formed in the main body. [Solution] Provided is a thruster comprising: a main body using a ceramic base composite material; and a cover material having a covering section that covers the outer surface of a micro-crack generation area in the main body. The covering section of the cover material is provided at a distance from the outer surface of the micro-crack generation area in the main body. The inner surface of the covering section of the cover material and the outer surface of the main body form a gap that rectifies, in the thrust direction, leaked gas that leaks from the micro-cracks. A leaked gas outlet is provided positioned at an end section on the rear side of the gap in the thrust direction and causes leaked gas to flow out. The micro-crack generation area in the main body is at least one area among a chamber, a throat, and a nozzle skirt. A main component of the cover material is either titanium, nickel, or an alloy of these.

Description

Thruster

The present invention relates to a thruster used for an artificial satellite, a spacecraft, and a flying object.

The thruster is a member used for attitude control of artificial satellites, spacecrafts, and flying objects, and is attached to a predetermined position on the outer surface of the aircraft such as artificial satellites. The basic structure of the thruster is composed of an injector assembly, a chamber, a throat, and a nozzle skirt.

Fuel for propelling satellites and the like is introduced from the fuel tank into the chamber via the injector assembly. The introduced fuel is mixed and burned in the chamber. The combustion gas is compressed by the throat and then injected from the nozzle skirt, and the attitude of the artificial satellite or the like is controlled by the thrust.

Conventionally, Nb alloy with a special coating can be used as a thruster material. However, since Nb alloys are heavy and expensive, a lightweight and inexpensive thruster material is required. Instead of the Nb alloy, there are a cooling type inexpensive thruster and a thruster using light monolithic ceramics. However, the above-described thruster requires a cooling mechanism, so that the problem of weight reduction remains. Moreover, since a brittle material is used, the reliability of strength is low.

There is a composite material as a material that is lightweight and does not break brittlely. Examples of the composite material used for the thruster include a carbon / carbon (C / C) composite material. However, a thruster manufactured from a C / C composite material tends to have a reduced thrust due to repeated use. One reason is ablation by combustion gas.

That is, during fuel combustion, ablation occurs in the C / C thruster due to the heat of the combustion gas in a high temperature / oxidizing atmosphere. When ablation occurs, the C / C thruster deforms, and thinning is particularly noticeable on the throat inner wall where heat tends to collect. In the thruster deformed from the initial shape, the flow path of the combustion gas changes and the thrust is reduced. Therefore, the C / C thruster is not suitable for repeated use and is not used for attitude control of spacecrafts and satellites. The main application of C / C thrusters is the first stage rocket nozzle and auxiliary rockets for disposable launches.

As a composite material that can be used repeatedly, a ceramic matrix composite material (CMC) reinforced with ceramic fibers is proposed. An example of CMC is a SiC / SiC composite material. Patent Document 1 discloses a SiC / SiC thruster having high airtightness and high thermal shock resistance. The SiC / SiC thruster is suitable for repeated use because it does not generate ablation due to combustion gas.

However, in the conventional SiC / SiC thruster, after the first mission, very fine cracks (microcracks) are generated due to thermal shock. When microcracks occur, combustion gas inside the thruster may leak from the microcracks. Since the leaked leak gas is very small, it does not affect the thrust of the thruster. However, if a component derived from leak gas adheres to a fuselage such as an artificial satellite equipped with a thruster, a camera, a sensor, or the like installed in the artificial satellite, smooth operation and accurate observation may be hindered.

A CMC thruster whose surface is coated is proposed as a leak gas leakage prevention means. However, the coating material adhered to the outer surface of the CMC thruster is affected by the thermal stress generated by the CMC thruster and breaks or peels off. This phenomenon is particularly remarkable in the region where the combustion heat is concentrated in the CMC thruster. Further, the combustion heat is transmitted to the coating material, the temperature difference between the inside and the outside (coating material surface) of the CMC thruster is enlarged, and the thermal stress is increased, which may promote the generation of microcracks.

Japanese Patent Application No. 11-19416

However, a thruster that can prevent contamination due to leak gas-derived components is desired. An object of the present invention is to prevent contamination due to leak gas leaking from a microcrack of a thruster.

The present invention includes a main body using a ceramic matrix composite material and a cover material having a covering portion that covers an outer surface of a microcrack generation region of the main body, and the covering portion of the cover material is a microcrack generation region of the main body. The gap between the inner surface of the covering portion of the cover material and the outer surface of the main body forms a gap for rectifying the leak gas leaking from the microcracks in the thrust direction, and the gap is located on the rear side in the thrust direction. It is a thruster provided with the leak gas outflow port which is located in the edge part of this and flows out leak gas. The microcrack generation region of the main body is at least one of a chamber, a throat, and a nozzle skirt. The main component of the cover material is preferably titanium, nickel, or an alloy thereof.

The present invention can prevent and reduce contamination by leak gas leaking from the microcracks of the main body.

It is a perspective view which illustrates the thruster of this invention. It is the cross-sectional schematic which illustrates the thruster of this invention.

The thruster of the present invention includes a main body using a ceramic matrix composite material and a cover material having a covering portion that covers a microcrack generation region of the main body. The covering portion of the cover material is provided to be separated from the outer surface of the main body. FIG. 1 is a perspective view illustrating a thruster of the present invention. FIG. 2 is a schematic cross-sectional view illustrating the thruster of the present invention.

1 and 2, 100 is a thruster, 110 is a main body, and 120 is a cover material. The main body 110 is a rotating body in which an axis X is a rotation axis and an outer radius j of a cross section perpendicular to the axis X is appropriately changed along the axis X. The rotating body has a hollow and includes areas of an injector assembly 111, a chamber 112, a throat 113, and a nozzle skirt 114. The chamber 112 communicates with a fuel tank (not shown) via an injector assembly 111. As a result, fuel is supplied from the fuel tank to the chamber 112. The combustion gas burned in the chamber 112 is compressed by the throat 113 and then injected from the nozzle skirt 114. In the present invention, the thrust direction is a direction in which an artificial satellite or the like to which the thruster 100 is attached is propelled. The thrust direction is determined along an array including the injector assembly 111, the chamber 112, the throat 113, and the nozzle skirt 114. The area close to the fuel tank is the front side in the thrust direction, and the area away from the fuel tank is the rear side in the thrust direction. When specifically shown in FIG. 2, the thrust direction is substantially parallel to the axis X. The injector assembly 111 side is the front side in the thrust direction, and the nozzle skirt side is the rear side in the thrust direction.

The cover material 120 is a rotating body having a rotation axis that is coaxial with the rotation axis of the main body 110, and includes a mounting portion 121 that attaches the cover material 120 to the main body 110 and a covering portion 122 that covers the outer surface of the main body. The inner dimension of the covering portion 122 is larger than the outer dimension of the corresponding region of the main body 110. Therefore, when the cover material 120 is attached to the main body 110, the cover portion 122 of the cover material 120 is separated from the opposing outer surface of the main body, and the gap between the inner surface of the cover portion 122 of the cover material 120 and the outer surface of the main body 110 is increased. 130 is formed.

2, an annular leak gas outlet 140 is formed at the end of the main body 110 on the nozzle skirt side by a gap between the main body 110 and the cover member 120, as shown in an enlarged sectional view A in the vicinity of the nozzle skirt. . The leak gas outlet 140 allows leak gas to flow out. The end of the main body 110 on the nozzle skirt side and the cover member 120 may or may not be coupled as long as the leak gas outlet 140 is formed. Reference numeral 150 denotes a microcrack generated by a thermal shock given by heating the thruster of the present invention. The arrow shown in the gap 130 is the flow of leak gas. The leak gas is rectified in the gap 130 and escaped to the outside from the leak gas outlet 140 located at the rear end of the gap 130 in the thrust direction. The arrows shown in FIG. 2 represent the rectification direction of the leak gas. As shown in FIG. 2, the leak gas is rectified along the outer surface of the main body and the inner surface of the cover portion of the cover material, and flows to the leak gas outlet 140.

The main body of the thruster of the present invention is manufactured using a ceramic matrix composite material. By using the ceramic matrix composite material, the main body can be made lightweight and tough. Examples of ceramic-based composite materials include SiC / SiC composite materials with SiC fibers and SiC matrix, carbon / SiC (C / SiC) composite materials, and carbon / carbon with an oxidation resistant coating (C / C) Composite materials.

The main body uses a ceramic fiber fabric to form a megaphone-type (partial cone-shaped) fabric molded body including an injector assembly, chamber, throat, and nozzle skirt illustrated in FIGS. 1 and 2, and the fabric molded body It can be produced by impregnating a ceramic matrix.

Furthermore, a ceramic matrix composite material in which another matrix layer having a low binding force of ceramic fibers may be formed in the voids of the ceramic matrix. Thereby, the main body of a thruster can be made excellent in lightness, airtightness, and impact resistance. Examples of the material used for the other matrix layer include a combination of an organic silicon polymer such as polycarbosilane and a carbon-based organic polymer such as phenol and Si (metal).

The main body having the above structure is manufactured by a hybrid process in which a ceramic matrix layer is formed by CVI treatment and another ceramic matrix layer is formed by PIP treatment after forming a fiber molded body by net shape molding of ceramic fibers. be able to. Other manufacturing methods include SPI treatment in which carbon powder and Si powder are mixed and fired as exemplified in Japanese Patent No. 5138901, and MI treatment in which phenol containing carbon powder is fired, carbonized, and impregnated with molten Si. .

When the fuel mixed in the chamber of the main body burns, microcracks are generated in the main body due to the thermal shock. In the present invention, the “microcrack generation region” means a region where a microcrack is generated by thermal shock, and includes a region where the generation of the microcrack is expected based on a microcrack generation condition described later. .

The microcrack generation condition is that combustion heat gathers and the temperature becomes high. For example, a region where the combustion heat has passed through the body and the temperature has been raised to 1000 ° C. or more is likely to generate microcracks due to the progress of sintering and crystallization.

Another condition is that the heated region is thick. That is, when the thick region of the main body is heated with combustion heat, the temperature difference between the inside and the outside of the region increases. Therefore, a large thermal stress is applied to the region, and microcracks are easily generated. When the temperature condition is 1000 ° C or higher, a SiC / SiC composite material with a thickness of 3 mm or more is liable to generate microcracks.

Other microcrack generation conditions include stress caused by differences in the thermal expansion coefficients of different materials such as C / SiC composite materials and reinforcing fibers. When the absolute value of the difference in thermal expansion coefficient between the reinforcing fiber and the matrix is 3.0 × 10 ⁻⁶ (/ K) or more, microcracks are likely to occur. The absolute value of the difference in coefficient of thermal expansion is 0 to 0.5 × 10 ^-6 (/ K) in the case of SiC / SiC composite material, which is unlikely to cause microcracks. In the case of C / SiC composite material, it is 3.0 to 4.5 × 10 ^-6 (/ K), which is the main cause of micro cracks. In the case of C / C composite material with oxidation resistant coating, it is 0 to 1.0 × 10 ^-6 (/ K), but the absolute value of the difference in coefficient of thermal expansion from the oxidation resistant coating is 3.0 × 10 ^-6 (/ K) Many of the above materials are the cause of microcracks. A microcrack is generated by satisfying one of the above microcrack generation conditions in a protruding manner or by satisfying a plurality of conditions in a composite manner.

For example, the main body includes the chamber 112, the throat 113, and the nozzle skirt 114 shown in FIG. 2, and the thickness of each region is 1 to 2.5 mm for the chamber, 1 to 2.5 mm for the throat, and 1 to 2.5 mm for the nozzle skirt. It is assumed that a material manufactured by using a SiC / SiC composite material is used. The throat is easy to collect combustion heat due to its structure and tends to become high temperature. Therefore, from the viewpoint of reducing the influence of thermal stress, the thickness of the main body of the above example is smaller than the thickness of any region of the chamber or the nozzle skirt in the region where the thickness of the throat is the smallest.

However, when the temperature gradient inside the main body is large and an excessive thermal shock is given to the throat, micro cracks are generated even if the thickness is reduced. Therefore, when combustion gas of 2000 ° C. or higher passes through the inside of the main body in the above example, microcracks are most likely to occur at the throat, and a small number of microcracks are also generated at the chamber and nozzle skirt.

Specific examples of the microcrack generation region having the microcrack generation conditions exemplified above include a throat, a nozzle skirt, and a chamber. The covering portion of the cover material of the present invention covers one or more regions of these specific examples. The covering portion of the cover material, which will be described later, is not limited to a mode of covering the outer surfaces of the chamber, the throat, and the nozzle skirt of the main body as illustrated in FIGS. 1 and 2, and covers the outer surfaces of the chamber and the throat. And a mode of covering the outer surface of the throat.

When microcracks occur, a very small amount of combustion gas and soot in the main body leaks out of the main body as leak gas through the microcracks. The present invention includes a cover material that covers a microcrack generation region. As illustrated in FIGS. 1 and 2, in the present invention, a gap 130 is formed between the cover portion 122 of the cover material and the main body. Therefore, the combustion heat in the main body 110 is hardly transmitted to the cover material 120, and the temperature rise of the cover material 120 is suppressed. Further, the cover material 120 is not easily affected by the thermal stress generated in the main body 110. Therefore, breakage of the cover material 120 can be prevented. Thus, the present invention improves the leakage gas leakage prevention effect in the microcrack generation region.

Further, the leak gas leaked from the micro crack 150 of the main body 110 is rectified in the thrust direction in the gap 130 and is released to the outside from the leak gas outlet 140 formed at the end of the nozzle skirt. Therefore, the leak gas does not stay in the gap 130. The present invention suppresses the diffusion of leak gas in the vicinity of a fuselage such as an artificial satellite to which a thruster is attached. As a result, contamination due to leak gas-derived components can be avoided.

That is, according to the present invention, it is possible to suppress the adhesion of leak gas-derived components to an airframe such as an artificial satellite, maintain the thermal design at the initial design, and obtain a desired heat insulation effect. In addition, it prevents the leak gas-derived components from adhering to the cameras, sensors, etc. equipped in the aircraft, contributing to accurate observation and smooth operation.

As illustrated in FIG. 2, the cover material used in the present invention is a rotating body having a rotation axis coaxial with the rotation axis X of the main body 110 and has a hollow. Comparing the inner radius h of the covering portion 122 of the cover material 120 with the outer radius j of the corresponding region of the main body 110 in the cross section in the direction perpendicular to the rotation axis X, the inner radius h of the covering portion 122 is Greater than 110 outer radius j. As a result, when the cover material is attached to the main body, the present invention covers the microcrack generation region in a state where the cover portion of the cover material is separated from the outer surface of the main body, and the gap 130 is formed by the inner surface of the cover portion and the outer surface of the main body. Can be formed.

Therefore, the shape of the cover material is determined according to the outer shape of the main body. In the direction orthogonal to the rotation axis X, the separation distance d between the inner surface of the covering portion and the outer surface of the main body facing the cover portion may be changed along the rotation axis X. The megaphone-type cover material shown in FIG. 2 has a shape in which the covering portion 122 gradually increases the inner radius h of the covering portion along the rotation axis X direction from the attachment portion 121 side. When the megaphone-type cover material is used, the separation distance d becomes the largest at the throat 113 and becomes smaller from the throat 113 toward the chamber 112 or the nozzle skirt 114. The throat 113 is an area where the temperature tends to increase and microcracks are most likely to occur. Therefore, the amount of leaked gas is relatively large. By increasing the separation distance d of such a region, heat transfer to the cover material 120 can be suppressed, and the leakage gas can be kept in the gap 130 to a minimum. The distance d between the inner surface of the covering portion and the outer surface of the main body relative thereto is the difference between the inner radius h of the covering portion of the cover material and the outer radius j of the corresponding region of the main body in the cross section perpendicular to the rotation axis X In other words, (d = h−j).

From the viewpoint of suppressing the retention of leak gas in the gap 130 and promoting rectification, the preferable separation distance is 15 to 30 mm, more preferably 20 mm in a region where the amount of microcracks generated is large and the leak amount of leak gas is large. is there. The separation distance in the region where the amount of microcracks generated is small and the amount of leaked gas is small may be about 10 mm. The separation distance at the leak gas outlet 140 is preferably 0.5 to 2 mm, more preferably 1 mm. By providing the separation distance within the above preferable range, the leak gas is appropriately rectified in the thrust direction in the gap 130 and flows out from the leak gas outlet 140 toward the rear side in the thrust direction. In addition, since the cover material 120 and the leak gas do not come into local contact, the thruster of the present invention avoids a partial increase in the temperature of the cover material and is less deteriorated.

The material of the cover material used in the present invention is preferably a lightweight material having high heat resistance and high airtightness corresponding to the environment in which the thruster is used and leak gas. For example, when the present invention is attached to an artificial satellite, the heat resistant temperature of the cover material is preferably 600 ° C or higher, and more preferably 800 ° C or higher. Specifically, titanium, nickel, and alloys thereof are preferably selected. The cover material contains 50% by mass or more, preferably 96% by mass or more, based on the total mass of the cover material, with the preferred material as a main component.

The attachment part of the cover material is not particularly limited as long as the covering part can be disposed at a position corresponding to the microcrack generation region of the main body. A preferable mounting position of the cover material is a region at least in front of the throat in the thrust direction on the outer surface of the main body, and more preferably a region of the injector assembly of the main body. Since the injector assembly region is kept at a low temperature, the thermal stress is relatively low. Therefore, even if the cover material is brought into close contact with the injector assembly, there is no risk of damage.

The shape of the attachment portion is a shape that can be attached to an airframe such as an artificial satellite to which the present invention is attached, and can be fitted to the main body of the present invention. Specific examples include a ring shape. Examples of methods for attaching the cover material to the main body include the following methods. The above-described attachment region of the main body is fitted to the attachment portion, and the outer surface of the main body and the inner surface of the attachment portion are joined. Thereafter, a cover material covering portion is joined to the attachment portion by means such as welding. As another method, the attachment region of the main body may be fitted to a cover material in which the attachment portion and the covering portion having the above shape are integrated. Alternatively, there is a method in which a flange is joined to the main body and a cover material is bolted to the flange. In any case, the length in the thrust direction of the joining region between the main body and the cover material is preferably at least 8 to 15 mm, and more preferably 8 mm.

The attachment part of the main body and the cover material can be joined using a known high heat-resistant brazing material. Examples of such a brazing material include a brazing material containing any one of titanium, zirconium, and hafnium. Specifically, active silver wax or the like. When joining using brazing material, the thruster of the present invention can provide a clearance of about 0.1 to 0.2 mm between the inner surface of the mounting portion and the outer surface of the mounting region of the main body at the melting temperature of the brazing material. Designed to.

The more specific joining process includes at least a fitting process, an infiltration process, and a cooling process. In the fitting step, the above-mentioned clearance is formed by fitting the attachment region of the main body to the attachment member of the cover material. The brazing material is poured into the clearance in the permeation process, and the brazing material is infiltrated into the cover material and the main body, and then held at a temperature exceeding the melting temperature of the brazing material for 10 minutes or more. Thereafter, in the cooling process, the cover can be attached to the main body by cooling at a cooling rate of about 1 ° C / min. A covering portion can be joined to the attachment portion to obtain the thruster of the present invention.

The cover material of the present invention contributes to preventing contamination due to diffusion of leak gas, extending the life of the thruster, and has the effect of protecting the main body. The above effects are remarkably exhibited when applied to a main body using a ceramic matrix composite material, but can also be applied to a main body manufactured from other materials. The above cover material can obtain the above-described effects even when attached to a main body in which a C / C composite material is used.

The fuel supplied to the chamber of the present invention may be a liquid fuel or a solid fuel, and specific examples include nitric oxide (NTO) / nitrogen (N ₂ ), NTO / monomethylhydrazine (MMH), and the like.

Examples of the present invention will be described below. However, the present invention is not limited to the following examples.

[Example]
A main body made of SiC / SiC composite material formed into the shape shown in FIG. 1 was prepared. The shape of the main body is a shape including an injector assembly, a chamber, a throat, and a nozzle skirt as shown in FIG. Regarding the length in the thrust direction, the injector assembly of the main body was 1 cm, the chamber was 1.3 cm, the throat was 1.4 cm, and the nozzle skirt was 3.4 cm. The outer diameter of the cross section in the direction perpendicular to the axial center was 3.0 cm in the injector assembly, and gradually decreased toward the throat. The minimum outer diameter of the throat cross section was 1.6 cm. The outer diameter of the cross section increased from the throat to the nozzle skirt, and the outer diameter of the cross section at the end of the nozzle skirt was 4.4 cm.

The attachment of the cover material comprising the attachment portion and the covering portion of the present invention to the main body was performed by first fixing the attachment portion to the main body and then welding the covering portion to the attachment portion. A titanium alloy ring having an inner diameter of 3.3 cm was prepared as an attachment part for the cover material. The injector assembly (outer diameter: 3.0 cm) of the main body was heated, and 8 mm in length in the thrust direction was fitted to the ring. Activated silver wax was poured as a brazing material into the clearance between the injector assembly and the ring of the main body, and further heat treatment was performed. The heat treatment temperature was 750 ° C., and the heat treatment holding time was 5 minutes. Thereafter, the ring was cooled at a cooling rate of 1 ° C./min for 720 minutes, and the ring was fixed to the outer surface of the injector assembly.

A covering part made of the same titanium alloy as the ring was prepared. The shape of the covering portion was a megaphone-type partial conical shape shown in FIG. 1, and the inner diameter gradually increased from the end portion on the side to be welded to the ring toward the opposite end portion. As for the dimensions, the length in the thrust direction was 7.1 cm, the inner diameter of the end on the welding side with the ring was 3.3 cm, and the inner diameter of the opposite end was 3.7 cm. The covering portion was welded to a ring fixed to the main body to cover the main body chamber, throat, and nozzle skirt, thereby obtaining the thrusters of the examples.

The separation distance between the outer surface of the main body and the inner surface of the cover material was calculated from the size of each region of the main body and the size of the covering portion of the cover material. The maximum separation distance was 1.2 cm of the gap formed on the throat outer peripheral surface. The width of the annular leak gas outlet formed on the outer peripheral surface of the nozzle skirt at the end of the main body was 2 mm.

The above embodiment has the effects of the present invention described above. When actually attaching to an artificial satellite, a spacecraft, a flying object, etc., since this invention is manufactured with the size corresponding to an attachment object, it is larger than the size of said Example. However, the dimensional ratios of the main body and the cover material, and the gaps formed by these are attached with those manufactured at the same dimensional ratio as in the above embodiment. The present invention is lightweight and can suppress contamination due to leak gas-derived components of an airframe such as an artificial satellite to which a thruster is attached. Moreover, leak gas can be made to flow out in the thrust direction.

100 thrusters
110 body
111 Injector assembly
112 chambers
113 Throat
114 nozzle skirt
120 Cover material
121 Mounting part
122 Cover
130 gap
140 Leak gas outlet
150 micro crack
X axis of rotation (axis)
d Distance between the outer surface of the body and the inner surface of the cover material
h Cover radius
j Outer radius of the body

Claims

A main body using a ceramic matrix composite material and a cover material having a covering portion that covers the outer surface of the microcrack generation region of the main body, the covering portion of the cover material being separated from the outer surface of the microcrack generation region of the main body The gap between the inner surface of the covering portion of the cover material and the outer surface of the main body forms a gap that rectifies the leak gas leaked from the microcracks in the thrust direction, and is located at the end of the gap on the rear side in the thrust direction. And a thruster provided with a leak gas outlet for allowing leak gas to flow out.
The thruster according to claim 1, wherein the microcrack generation region of the main body is at least one of a chamber, a throat, and a nozzle skirt.
The thruster according to claim 1 or 2, wherein a main component of the cover material is titanium, nickel, or an alloy thereof.