UA98431C2 - Liquid-propellant engine with afterburning of exhaust turbine gas in supersonic part of the nozzle and with controlled thrust vector - Google Patents

Abstract

The invention relates to rocket-space engineering and can be used as sustainer liquid rocket engine of the higher stages of carrier rockets and acceleration blocks of spacecrafts. A liquid rocket engine with afterburning of exhaust turbine generator gas in supersonic part of the nozzle and with controlled thrust vector comprises a combustion chamber with supersonic nozzle, turbo-pump system for supply of fuel components with gas generator and with turbine of active type exhaust collector of which is connected by gas conduit to circular collector of the system for injection of exhaust turbine gas to the nozzle. For afterburning of generator gas and control of thrust vector of chamber of the engine in the wall in supersonic part of the nozzle jets are provided for injection of oxidation component of fuel connected with one circular collector divided with partitions into eight sections that are connected by a system of pipelines to high-pressure main of the engine. For control of thrust vector of the engine each two opposite in diameter sections that are placed in the planes for pitch and yaw control of thrust vector, connected to each other by a pipeline with distributor valve. The inlet branch pipe of each distributor valve is connected by a pipeline to distributor valve for supply of oxidation component to the system for thrust vector control and to intermediate sections of injection of oxidizer for afterburning of generator gas. Advantage of the invention is in increase of functionality of the system for the engine thrust vector control due to use of new pattern of distribution of losses of oxidation fuel component and increase of thrust vector control efficiency.

Description

The invention belongs to rocket and space technology and can be used as a cruise liquid rocket engine (RRD) of the upper stages of carrier rockets and booster units of space vehicles.

RRDs of the specified purpose are known with a turbopump system for supplying fuel components and afterburning the gas spent on the turbine in the supersonic part of the nozzle (1).

The disadvantage of the engine according to the patent (1) (with afterburning of the turbine exhaust gas in the supersonic part of the nozzle of the main chamber) is that it does not regulate the thrust vector (VT), in connection with this, the functional capabilities and ballistic characteristics of the engine installation are sharply reduced in as a whole

Mechanical and thermogasodynamic methods and devices for adjusting the thrust vector of a rocket engine are known from many sources, in particular (2, 3, 4). But the use in engines with the afterburning of exhaust turbine gas in the supersonic part of the nozzle (according to the above-mentioned patent 1) of known devices for regulating VT is impractical or impossible. Some known solutions for the VT regulation system are not compatible with solutions for exhaust gas afterburning or require the use of additional devices, units, and systems in the engine, while the mass of the engine structure increases, the volume, term, and costs of the engine as a whole increase.

The closest analogue of the engine (prototype), where new solutions are applied, is a liquid rocket engine with afterburning of spent generator gas on the turbine and with an adjustable thrust vector according to the Ukrainian patent for the invention Mo 94359.

Common essential features of the prototype engine and the claimed engine include the fact that the engine is mounted stationary and tightly assembled. It contains a combustion chamber with a supersonic nozzle, which are cooled regeneratively by the main fuel components, a turbopump system for supplying two fuel components of approximately the same pressure to the engine combustion chamber, and a gas generator that generates working gas for the turbine with a large excess of fuel. The outlet manifold of the turbine is connected by a gas pipeline to the annular manifold made in the middle part of the supersonic nozzle and connected by an annular gap to the internal space of the nozzle; on the wall of the nozzle in the zone of the gap of injection into the nozzle of the generator gas is made in one plane of the cross-section of the nozzle section of the annular collector with

With nozzles for injecting the oxidizing component of the fuel into the nozzle, which is connected to a pipeline that takes the oxidizing component from the high-pressure main line of the engine.

The axis of symmetry of each section is located in the flight control planes of the rocket stage with respect to pitch and heading, every two diametrically opposed injection sections are connected by a pipeline to a hydraulic distributor, which is controlled by an actuator under commands from the aircraft flight control system, and the inlet pipe of each hydraulic distributor of the pitch and of the course is connected to pipelines supplying the oxidizing component for afterburning in the nozzle of the exhaust generator gas, which can be connected directly, or using an additional hydraulic distributor with a high-pressure engine line before or after the chamber path.

The disadvantage of the prototype engine is that it has an unfriendly injection system of the oxidizer component in the nozzle due to the fact that the injection sections have a large angle of the injection sector

Vvy-907, at the same time, we have large losses of lateral forces due to the influence of the curvature of the nozzle surface, and we also have a relatively low efficiency of interaction of the main flow of the nozzle with the controlled flow of the oxidizing component of the injected fuel.

The basis of the invention is the task of improving the engine through the use of new schematic and design solutions regarding the injection of the oxidizing component of the fuel into the nozzle in order to adjust the thrust vector of the prototype engine, taking into account the characteristics of the afterburning of the spent generator gas on the turbine.

The problem is solved by the fact that not the entire annular injection manifold is used to adjust the VT, but only its four parts. This is achieved by using the following variants of the engine design: - eight partitions are made in the annular collector with injection nozzles of the fuel oxidizing component in such a way that the collector becomes divided into eight sections (discrete). The axis of symmetry of the four injection sections is located in the flight control planes of the rocket stage with respect to pitch and course; every two diametrically opposite oxidizer injection sections (pitch or heading sections) are connected by a pipeline to a hydraulic distributor, which is driven by a drive on commands from the flight control system, the inlet pipe of each hydraulic distributor, as well as the other four sections are connected by pipelines to the engine high-pressure line , which supplies the oxidizer to the nozzle for 60 afterburning the generator gas;

- in order to determine the optimal modes of afterburning the exhaust turbine gas in combination with the optimal modes of control of the engine thrust vector, the angle of the sector of the VVP injection annular collector is 30";

The essence of the invention is explained by the drawing in fig. 1, which shows the diagram of the engine with the afterburning of exhaust turbine gas in the supersonic part of the nozzle and the system of gas-dynamic adjustment of the thrust vector.

The rocket engine includes a combustion chamber 1 with a supersonic nozzle 2, a turbopump system for supplying fuel components to the chamber, which includes a gas generator 3, a turbine 4, an oxidizer pump 5, a fuel pump 6. Through the oxidizer 7 and fuel 8 pipelines, the fuel components after cooling the chamber walls enter the combustion chamber 1. At the same time, through pipelines 9 and 10, the fuel components enter the gas generator 3, where generator gas with a large excess of fuel is created, which enters the turbine 4 and, having given part of the energy to the turbine, enters the exhaust manifold 11. The exhaust manifold 11 of the turbine connected to the annular collector 12 of the injection into the nozzle of the spent generator gas on the turbine 4, which is located in the supersonic part of the nozzle and is connected by a gap 13 to the flow (inner) part of the nozzle. To ensure the uniformity and stability of the geometric characteristics of the gap 13, the walls of the gap are connected by stiffening ribs 14, which are evenly located along the perimeter of the gap 13. The input fuel collectors of the oxidizer 15 and fuel 16 with the cooling channels of the chamber end with the output collectors 17 and 18, from which the fuel components come into the nozzle head 19 of the combustion chamber 1. For the afterburning of the generator gas spent on the turbine 4, on the supersonic part of the nozzle, nozzles for the injection of the oxidizer 20 are made, which enters the nozzles from the annular collector 21, made on the wall of the nozzle and connected by a branched system of pipelines 22 and 25 3 output collector 17 of the oxidizing component. To adjust the thrust vector of the engine, the injection manifold 21 of the oxidizing fuel component in the nozzle 2 is divided by partitions 23 into eight sections. Four sections 21.1, 21.2, 21.3, 21.4 (section A-A, Fig. 2) are located in the VT control planes along the pitch and heading channels. Each two opposite sections 21.1, 21.2 and 21.3, 21.4 are interconnected by pipelines 25.1 and 25.2 with hydraulic distributors 24.1 and 24.2, which are controlled by actuators 26.1 and 26.2, which are connected by the control system (SU) of the thrust vector along the planes

From pitch and course. In order to determine the optimal modes of afterburning the exhaust turbine gas in combination with the optimal modes of control of the engine thrust vector, the angle of the Vvp injection annular manifold sector is 30". The other four sections 21.5, 21.6, 21.7, 21.8 are interconnected by a pipeline 28 with a hydraulic distributor 27, which driven by actuator 29.

The hydraulic distributor 27 with the drive 29 is installed in the pipeline 22 to expand the range of adjustment of the thrust vector of the engine in one of the pitch or heading planes without increasing the consumption of oxidizer for injection into the nozzle. At the same time, the hydraulic distributor 27 redistributes the oxidizer consumption between the traction vector control system (through the branched system of pipelines 22) and the afterburning system of the turbine exhaust gas (through the pipeline 28).

The device works like this.

The turbopump system 3, 4, 5, 6 supplies fuel components through pipelines 7 and 8 to the inlet manifolds 15 and 16 with the cooling paths of the combustion chamber 1. After cooling the combustion chamber, the heated fuel components enter the outlet manifolds 17 and 18, from which they enter the injector head 19, burn at the optimal ratio of second consumption, combustion products flow out of the nozzle and create jet thrust. Through pipelines 9 and 10, the fuel components, which are taken from the main fuel pipelines 7 and 8, enter the gas generator 3. The generator gas from the generator Z enters the turbine 4, from the exhaust manifold 11 it enters the ring manifold 12. From the manifold 12, the generator gas is blown into the supersonic flow of the nozzle 2 through the annular gap 13. For afterburning the gas blown into the nozzle, the oxidizing component is taken from the collector 17 and fed through the main line 22 into the distribution hydraulic system for adjusting the thrust vector to the hydraulic distributors 24 and further to the collector 21 of the injection nozzles 20. In the supersonic part of the nozzle 2 the oxidizer interacts with the injected generator gas, which is thus afterburned, increasing the thrust of the engine. If there is no need to adjust the thrust vector by the hydraulic distributors 24, the oxidizing component is sent evenly to all sections of the collector 21. To adjust the thrust vector along the pitch and heading channels, the hydraulic distributors 24, according to control commands of the control system to the drives 26, redistribute the oxidant consumption unevenly between the opposite sections depending on the required lateral control forces. At different oxidizer consumptions, unbalanced lateral forces are created in opposite sections and thus the thrust vector of the engine changes. To 60 provide maximum lateral forces, for example, in the pitch plane, all the oxidizer passing through the hydraulic distributor 24.1 is sent to one of the sections 21.1 or 21.2, in the other, only the leak can remain. Similarly, the thrust vector in the course plane is regulated. To increase the lateral forces in the engine thrust vector control system, the hydraulic distributor 27 reduces the oxidizer consumption in the exhaust gas afterburning section of the turbine, which is not involved in the engine thrust vector control system.

REFERENCES

1. Patent for the invention 86966, Ukraine, IPC G23НО/00. The method of afterburning the spent generator gas of the turbine of the turbopump unit of the liquid rocket engine and the device for its application / Kovalenko M. D., Strelnikov G. O., Kovalenko G. M., Khomenko O. V.,

Kovalenko T.O., Syrotkina N.P.; the applicant and patent holder Institute of Technical Mechanics of the National Academy of Sciences and

NKAU. - a200610599; statement 06.10.2006; published 10.06.2009, Bull. 11. - 16 p. 2. Missiles and space vehicles of the "Southern" design bureau. Under the general editorship

S.M. Konyukhova. - Dnipropetrovsk: OSO "ColorGraf". - JSC "Tandem - U", 2001. - 240 p.

Z. Kovalenko N.D. The rocket engine as the executive body of the systems! control of rocket flight / N.D. Kovalenko. - Dnipropetrovsk: ITM of NASU and NKAU, 2004. - 412 p. 4. Patent for the invention 86958, Ukraine, IPC RO2KUO/00, gO2KO/42. Liquid rocket engine with adjustable thrust vector / M.D. Kovalenko, G.O. Strelnikov, H.M. Kovalenko, O.V. Khomenko,

Kovalenko T.O., Syrotkina N.P.; the applicant and patent holder Institute of Technical Mechanics of the National Academy of Sciences and

NKAU. - 2006 07625; statement 07.07.2006; published 10.06.2009, Bull. Mo. 11. - 10 p.m. 5. Patent for the invention 94359, Ukraine, IPC RO2KU/00, RO2KOU/42. A liquid rocket engine with the afterburning of spent generator gas on the turbine and with an adjustable thrust vector /

Kovalenko M.D., Strelnikov G.O., Kovalenko H.M., Ignatiev O.D., Kovalenko T.O., Syrotkina

N.P.; the applicant and patent holder is the Institute of Technical Mechanics of the National Academy of Sciences and the National Academy of Sciences. - a201005331; statement 04/30/2010; published 04/26/2011, Bul. Mo 8. - 10 p.

Claims (1)

FORMULA OF THE INVENTION A liquid rocket engine with afterburning in the supersonic part of the nozzle of the generator gas spent on the turbine and with an adjustable thrust vector, which includes a camera Combustion with a supersonic nozzle, which has cooling paths with fuel inlet and outlet manifolds, a turbopump system for supplying fuel components with a gas generator operating on the main fuel components with a large excess of fuel, with an active turbine type, the exhaust manifold of which is connected by a gas pipeline with an annular a gas collector made in the middle part of a supersonic nozzle with an annular slot for the injection of spent generator gas on the turbine, which connects the cavity of the injection collector with the cavity of the nozzle, as well as nozzles for injecting the oxidizing component of the fuel into the nozzle, made in the nozzle wall with a uniform location along the perimeter of the transverse cross-section of the nozzle and are united by one fuel collector divided by partitions into sections, connected by a pipeline system to the high-pressure main line of the oxidizing fuel component, which is distinguished by the fact that eight partitions are made in the manifold of the injectors for injecting the oxidizing component into the nozzle of the fuel oxidizing component, which separate it into eight sections, and the axis of symmetry of the four injection sections is located in the flight control planes of the rocket stage with respect to pitch and course, every two diametrically opposed oxidizer injection sections (pitch or course sections) are interconnected by pipelines with hydraulic distributors that are controlled by actuators under commands from flight control system, while the inlet nozzle of each hydraulic distributor, as well as the other four sections, are connected by pipelines to the high-pressure main of the engine, which supplies the oxidizer to the nozzle for afterburning the generator gas, and the four sections for injecting the oxidizer component of the fuel into the nozzle, which are located in the pitch planes and the course, made with the optimal angle of the VVP injection sector for adjusting the thrust vector - 30".