RU2705982C1 - Low-thrust single-component liquid-propellant rocket engine - Google Patents

Low-thrust single-component liquid-propellant rocket engine Download PDF

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Publication number
RU2705982C1
RU2705982C1 RU2019108547A RU2019108547A RU2705982C1 RU 2705982 C1 RU2705982 C1 RU 2705982C1 RU 2019108547 A RU2019108547 A RU 2019108547A RU 2019108547 A RU2019108547 A RU 2019108547A RU 2705982 C1 RU2705982 C1 RU 2705982C1
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RU
Russia
Prior art keywords
decomposition chamber
nozzle
fuel supply
support
set screws
Prior art date
Application number
RU2019108547A
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Russian (ru)
Inventor
Николай Михайлович Вертаков
Георгий Иванович Казаков
Original Assignee
Федеральное государственное унитарное предприятие "Опытное конструкторское бюро "Факел" ФГУП "ОКБ "Факел"
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Priority to RU2019108547A priority Critical patent/RU2705982C1/en
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Publication of RU2705982C1 publication Critical patent/RU2705982C1/en

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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60Constructional parts; Details
    • F02K9/68Decomposition chambers

Abstract

FIELD: astronautics.
SUBSTANCE: invention relates to space engineering, particularly, to single-component liquid-propellant rocket engines included in low-thrust propulsion systems of satellites for orbital maneuvering. Single-component low-thrust liquid-propellant engine comprises fuel decomposition chamber 1 with gas-dynamic nozzle 2, fuel supply tube 3 from control valve 4 fixed on circuit board 5 to inlet part of decomposition chamber 1. Decomposition chamber 1 with nozzle 2 and the fuel supply pipe 3 are coaxially located inside power casing 6 rigidly fixed on circuit board 5. Inlet end face of decomposition chamber 1 is permanently connected to support bushing of three-petal shape 7, having tail cylindrical part 8, at that fuel supply pipe 3 is coaxially located inside support sleeve with clearance. Elements of the decomposition chamber fixation elements located at the level of the fuel supply pipe inlet 3 into decomposition chamber 1 and at the level of the outlet part of nozzle 2 are made in the form of two sets of locating screws 10 and 11, three in each group. Heads of adjusting screws 10, 11 enter into threaded joints with power casing 6 and radially directed to engine axis. Ends of adjusting screws 10 pass through guide holes in intermediate ring 12, touching outer surface of tail cylindrical part 8. Intermediate ring 12 with a gap is fixed with mounting screws 10 in a suspended state above cylindrical tail part 8. Ends of adjusting screws 11 thrust against groove 13, made on outer surface of support ring 14. Support ring 14 is arranged on outer cylindrical surface of nozzle 2 in its outlet part. On inner surface of support ring 14 there is collar 15 providing contact of minimum area of support ring 14 with outer surface of nozzle 2 at all stages of engine operation. In power housing 6 there are openings 16, at the places of arrangement of threaded holes for adjusting screws 10 lugs 17 are made.
EFFECT: invention provides mechanical strength of fuel supply pipe and decomposition chamber as a whole with simultaneous maximum reduction of heat flow from decomposition chamber to fuel supply tube and power housing.
6 cl, 4 dwg

Description

The invention relates to space technology and can be used to create single-component liquid rocket engines that are part of the satellite thrust propulsion systems for solving orbital maneuvering problems.

A known engine for a spacecraft operating on hydrazine, comprising a control valve, a fuel supply pipe including a support tube with a capillary tube coaxially placed therein, and a decomposition chamber with a nozzle. The inlet end of the feed tube is fixed to the outlet of the control valve, and the outlet end is in the inlet of the decomposition chamber (GB Patent GB No. 1470664, IPC F02K 9/02, 1973).

A disadvantage of the known engine is the cantilever mounting of the decomposition chamber on the fuel supply pipe relative to the control valve. To eliminate the risk of destruction of the supply tube from mechanical influences when the spacecraft is launched, it becomes necessary to make the support tube strong enough to hold a relatively large mass in the form of a decomposition chamber at the output end of the supply tube. However, an increase in the wall thickness of the support tube leads to a substantial increase in the heat flux along the support tube from the decomposition chamber to the control valve. During the operation of the engine, during the decomposition of fuel, a significant amount of heat is released. After the engine is turned off, most of the heat from the decomposition chamber is conductively transferred to the support tube; as a result of heat “pumping” from the support tube to the capillary tube, the risks of ensuring engine operability increase, especially in pulsed operation modes and with relatively low traction.

A one-component electrothermal thruster is known, comprising a control valve, an elongated curved fuel supply tube, a decomposition chamber, and a nozzle. To reduce the heat flux from the decomposition chamber, a rigid perforated cylinder is placed between the end of the decomposition chamber and the mounting flange, which simultaneously acts as a stiffening element to ensure a stable position of the decomposition chamber relative to the mounting flange with holes for mounting the motor (US Patent US No. 7665292, IPC F03H 1 / 00, 2003)

The known engine has the following disadvantages:

- when the engine is turned on, a significant amount of heat is released in the decomposition chamber. After the engine is turned off, most of the heat from the decomposition chamber is conducted to the fuel supply pipe, which is rigidly fixed by the ends in the decomposition chamber and the mounting flange. When the engine is turned on, relatively cold (liquid) fuel enters the feed pipe. As a result, in the presence of a perforated cylinder rigidly fixed between the decomposition chamber and the mounting flange, the supply tube experiences large cyclic thermal loads and corresponding deformations, as a result of which the risks of fatigue defects in the fuel supply tube increase, especially in pulsed operation modes;

- removal of a significant heat flux from the decomposition chamber along a perforated cylinder with a relatively low thermal resistance leads to inefficient use of the heat of the chamber released from the decomposition of the fuel, and, accordingly, to the deterioration of the characteristics of the engine as a whole.

Known electrothermal gas traction unit, adopted for the prototype, containing a control valve, a fuel supply pipe, a decomposition chamber and a nozzle. An end plate is fixed to the control valve, onto which a power housing is fixed in the form of a cylinder made of titanium or stainless steel. At the levels of entry into the decomposition chamber and at the exit from the nozzle, elements for fixing the position of the decomposition chamber are made, made in the form of diaphragms with spokes that support the decomposition chamber body within the power housing, and the diaphragm spokes rigidly connect the fuel supply tube and the edge of the nozzle to the power housing ( GB Patent GB No. 2095336, IPC F02K 9/42, 1981).

The known engine has the following disadvantages:

- due to the fact that the diaphragm spokes rigidly connect the output part of the fuel supply tube and the edge of the nozzle to the power housing, cyclic stresses from the decomposition chamber cause cyclic stresses in the materials of both the feed tube in the axial direction and in the spokes of the diaphragms each time and turning off the engine. This reduces the mechanical strength of both the feed tube and the diaphragms with spokes, and, as a result, the mechanical strength of the engine during its operation;

- the removal of a significant heat flux from the decomposition chamber to the power housing through the diaphragms with spokes leads to the inefficient use of the heat of the chamber released from the decomposition of the fuel, and, accordingly, to the deterioration of the engine as a whole.

When creating the invention, the problem was solved of eliminating the deformation of the fuel supply pipe from the effects of mechanical loads during the launch of the spacecraft and due to temperature effects from the decomposition chamber during engine on and off and, accordingly, ensuring the mechanical strength of the fuel supply pipe and the decomposition chamber as a whole while maximum reduction of heat flow from the decomposition chamber to the fuel supply tube and power housing.

The problem is solved due to the fact that in the well-known single-component liquid propellant small thrust rocket engine containing a fuel supply pipe, a decomposition chamber with a nozzle coaxially placed in a power housing rigidly mounted on a circuit board, and elements for fixing the position of the decomposition chamber in the power housing installed according to the invention, the input end surface of the decomposition chamber is inseparably connected to the support sleeve of a three-petal shape, having at the entrance to the decomposition chamber and at the exit level of the nozzle a tail cylindrical part, and on the petals, a landing surface is made with the smallest possible contact area with the end surface of the decomposition chamber, and the fuel supply pipe is located with a gap in the channel of the support sleeve, an intermediate ring with guide holes is placed with a gap on the end of the tail cylindrical part of the support sleeve the outer surface of the nozzle with a gap placed a support ring with a groove on its outer surface, and the elements for fixing the position of the decomposition chamber are made in the form set screws, the heads of which are threaded to the power housing and radially directed to the axis of the engine, the ends of the set screws located at the entrance to the decomposition chamber inserted into the guide holes of the intermediate ring and come into contact with a minimum contact area with the outer surface of the tail the cylindrical part of the support sleeve, and the ends of the set screws located at the level of exit from the nozzle are inserted into the groove of the support ring until touching with a minimum area contacts.

Also, the problem was solved due to the fact that at least 3 set screws were installed at the level of entry into the decomposition chamber and at the level of exit from the nozzle.

Also, the problem is solved due to the fact that a support shoulder is made in the center of the inner surface of the support ring.

Also, the problem was solved due to the fact that the power housing, the support sleeve, the intermediate ring, the support ring and the set screws are made of a material with high thermal resistance.

The fastening on the input end surface of the decomposition chamber of a three-leaf support sleeve having a tail cylindrical part on which the set screws are supported with a minimum contact area allows, on the one hand, to eliminate the radial displacement of the input part of the decomposition chamber, and, on the other, to reduce the thermal flow from the decomposition chamber to the power housing due to the minimum contact area of the decomposition chamber with the petals of the support sleeve and set screws with the tail cylindrical part th sleeve, and also due to the gap between the tail cylindrical part of the support sleeve and the intermediate ring.

Placing on the outer surface of the nozzle with a clearance of the support ring with a groove for the set screws on its outer surface allows the mechanical strength of the decomposition chamber with the nozzle and at the same time reduce the heat flux from the nozzle to the set screws and further to the power housing.

The implementation of the support flange in the center of the inner surface of the support ring allows to minimize the contact of the nozzle with the support ring and, accordingly, to reduce the heat flux from the nozzle to the power housing.

The implementation of the power housing, the support sleeve, the intermediate ring, the support ring and the set screws of a material with high thermal resistance can reduce heat flow from the decomposition chamber and nozzle.

Thus, the maximum heat retention from the decomposition of fuel in the decomposition chamber ensures maximum engine performance.

The implementation of the elements for fixing the position of the decomposition chamber in the form of set screws allows you to additionally make the required adjustment of the axis of the nozzle and the decomposition chamber of the engine at the stage of its manufacture.

The invention is illustrated by drawings, where in FIG. 1 shows a general view of a single-component liquid propulsion thruster; in FIG. 2 - section aa; in FIG. 3 - remote element B; in FIG. 4 - remote element B.

The engine comprises a fuel decomposition chamber 1 with a gas-dynamic nozzle 2, a fuel supply pipe 3 from a control valve 4, mounted on a circuit board 5 with a flange for connection to a spacecraft (not shown in the drawing), to the input part of the decomposition chamber 1. Decomposition chamber 1 s the nozzle 2 and the fuel supply pipe 3 are coaxially located inside the power housing 6, made of a material with high mechanical strength and relatively high thermal resistance, for example, of a titanium alloy, and rigidly fixed to my tazhnoy plate 5. The input end surface of the decomposition chamber 1 is one-piece, for example, by welding, connected to the supporting sleeve of a three-leaf form 7 having a cylindrical tail portion 8, and on the petals there is a landing surface 9 with the smallest possible contact area with the end surface of the decomposition chamber 1, and the fuel supply pipe 3 is coaxially located inside the support sleeve with a gap. The supporting sleeve 7 with the tail of the cylindrical part 8 is made of a material with high thermal resistance. Elements for fixing the position of the decomposition chamber located at the level of the input of the fuel supply pipe 3 to the decomposition chamber 1 and at the level of the outlet part of the nozzle 2 are made in the form of two groups of setscrews 10 and 11, three in each group, the material of which has the highest thermal resistance, for example, titanium alloy. The heads of the set screws 10, 11 are threaded to the power housing 6 and radially directed to the axis of the engine. The ends of the set screws 10 pass through the guide holes in the intermediate ring 12 and are in contact with the outer surface of the tail cylindrical part 8 with a minimum contact area. The intermediate ring 12 is fixed with a gap by the set screws 10 in suspension above the cylindrical tail part 8 due to the minimum clearance between the set screws 10 and guide holes in the intermediate ring 12, as well as by selecting a specific inner diameter of the intermediate ring. The ends of the set screws 11 abut against a groove 13 made on the outer surface of the support ring 14, the width of which is larger than the diameter of the screws. The support ring 14 with the minimum necessary clearance to counter radial thermal deformations of the nozzle and the support ring is placed on the outer cylindrical surface of the nozzle 2 in its output part and is made of a material with high thermal resistance. A bead 15 is made in the center of the inner surface of the support ring 14, which ensures that the minimum area of the support ring 14 is in contact with the outer surface of the nozzle 2 at all stages of the engine operation. In the power housing 6, in order to facilitate the design and reduce the heat flux along the set screws to the places of their fastening, as well as along the power case in the direction of the mounting plate 5, windows 16 are made, while tides 17 are made in the locations of the threaded holes for the set screws 10 to provide of sufficient length for the strength of threaded joints.

The assembly and operation of the engine is as follows.

At the preliminary stage of engine assembly, to ensure the required orientation of the geometric axis of the nozzle of decomposition chamber 1, the engine is aligned within the elastic deformation of the fuel supply pipe 3, i.e. using the adjusting screws 10, 11, the angular and linear position of the axis of decomposition chamber 1 with the nozzle 2 relative to the base seating surface and the holes in the mounting plate 5 is adjusted. During the final adjustment at the stage of final assembly of the engine, to prevent self-unscrewing at all stages of its application, fixing is performed threaded joints of set screws 10, 11 in the power housing 6, for example, with high-temperature glue.

In the process of launching the spacecraft, when mechanical loads act on the engine, the set screws 10, 11 eliminate the radial displacement of decomposition chamber 1, and during the fire operation of the engine, they provide in the axial direction free deformations of decomposition chamber 1 and fuel supply tube 3 from thermal effects during turning on and off. When turning the engine on and off, the heat flux from the decomposition chamber to the remaining engine elements is reduced to a minimum due to the minimum contact areas of the set screws 10 with the tail portion 8 of the support sleeve 7 and the set screws 11 with the support ring 14, and also due to the presence of a gap between the support ring 14 and nozzle 2.

The inventive one-component liquid propellant rocket engine of low thrust passed the cycles of ground-based experimental tests and showed both high mechanical strength under the influence of various mechanical loads and high efficiency, and operational efficiency, including during long-term life tests both in continuous and in pulsed modes work.

Claims (6)

1. A one-component liquid propellant small thrust rocket engine containing a fuel supply pipe, a decomposition chamber with a nozzle coaxially placed in a power housing rigidly mounted on a circuit board, and elements for fixing the position of the decomposition chamber in a power housing mounted at the entrance to the decomposition chamber and at the level of exit from the nozzle, characterized in that the input end surface of the decomposition chamber is inseparably connected to the supporting sleeve of a three-petal shape having a tail cylindrical part, and on the petals in the landing surface is filled with the smallest possible contact area with the end surface of the decomposition chamber, and the fuel supply pipe is located with a gap in the channel of the support sleeve, an intermediate ring with guide holes is placed with a gap on the end of the tail cylindrical part of the support sleeve, and a support ring is placed on the outer surface of the nozzle with a gap a ring with a groove on its outer surface, and elements for fixing the position of the decomposition chamber are made in the form of set screws, the heads of which are threaded connections with the power housing and are radially directed to the axis of the engine, the ends of the set screws located at the level of the entrance to the decomposition chamber inserted into the guide holes of the intermediate ring and come into contact with the minimum contact area with the outer surface of the tail cylindrical part of the support sleeve, and the ends of the installation screws located at the exit level of the nozzle are inserted into the groove of the support ring with a minimum contact area until they touch.
2. A one-component liquid propellant thruster according to claim 1, characterized in that at least three set screws are installed at the level of the entrance to the decomposition chamber.
3. A one-component liquid propulsion thruster according to claim 1, characterized in that at least three set screws are installed at the exit level from the nozzle.
4. A one-component liquid propellant small thrust engine according to claim 1, characterized in that a support flange is made in the center of the inner surface of the support ring.
5. A one-component liquid propellant thruster according to claim 1, characterized in that tides are made in the power housing at the places of threaded connections with set screws.
6. A one-component liquid propellant small thrust rocket engine according to claim 1, characterized in that the power housing, support sleeve, intermediate ring, support ring and set screws are made of a material with high thermal resistance.
RU2019108547A 2019-03-25 2019-03-25 Low-thrust single-component liquid-propellant rocket engine RU2705982C1 (en)

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Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2095336A (en) * 1981-03-19 1982-09-29 Secr Defence Electrothermal gas thrust unit
RU2096647C1 (en) * 1993-03-24 1997-11-20 Опытное конструкторское бюро "Факел" Monopropellant liquid rocket thruster and method of starting it
US5941062A (en) * 1995-05-11 1999-08-24 Societe Europeenne De Propulsion Pulse rocket engine
RU2154748C2 (en) * 1996-09-09 2000-08-20 Опытное конструкторское бюро "Факел" Monopropellant thruster

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2095336A (en) * 1981-03-19 1982-09-29 Secr Defence Electrothermal gas thrust unit
RU2096647C1 (en) * 1993-03-24 1997-11-20 Опытное конструкторское бюро "Факел" Monopropellant liquid rocket thruster and method of starting it
US5941062A (en) * 1995-05-11 1999-08-24 Societe Europeenne De Propulsion Pulse rocket engine
RU2154748C2 (en) * 1996-09-09 2000-08-20 Опытное конструкторское бюро "Факел" Monopropellant thruster

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