RU2532454C1 - Method of liquid propellant rocket engine boosting by thrust and liquid propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: invention relates to the field of rocket engine technology, oriented at space transport systems. The method of liquid propellant rocket engine boosting by thrust based on variation of energy parameters of functioning, according to the invention, boosting is done by supplying some gas from a gas tract of at least one of fuel components, or generator gas, or their mixture, at least to one additional turbine, interacting at least with one of the main turbopump sets (TPS), and after leaving it the gas is sent for further use or discharge. After discharge from the additional turbine the gas is sent to the inlet of the discharge nozzle or into the nozzle of the chamber, or to the inlet of the additional turbine, or into a heat exchanger. A liquid propellant rocket engine (LPRE), comprising a chamber, control and regulation units, the TPS with the main and at least one additional turbine, besides, the gas tract of the main turbine is connected to the inlet to the chamber, in which according to the invention the gas tract of at least one of the components or their combustion products is equipped with an additional manifold, connecting it with the inlet of at least one additional turbine and outlet from it, at the same time the manifold is equipped with a local control system installed upstream or downstream the main turbine, and the outlet of the manifold is connected with the system of gas removal and/or the system of its reuse.

EFFECT: invention provides for higher cost-effectiveness of an LPRE at rated mode of operation and further increase (by more than 1,3 times) of thrust during engine boosting.

4 cl, 2 dwg

Description

The present invention relates to the field of rocket propulsion, focused on space transport systems.

One of the requirements for a liquid rocket engine (LRE) is the requirement to provide the ability to control the magnitude of the thrust during the flight of the rocket, including, in the direction of its increase, i.e. forcing. The process of changing the thrust of the liquid propellant rocket engine occurs by changing the fuel consumption through the combustion chamber, which, in turn, is achieved by changing the pressure of the fuel supply. The latter is achieved for liquid propellant engines equipped with a turbopump fuel supply system by changing the rotor speed of the turbopump assembly (TNA) by changing the turbine power.

A known method of changing the power of the turbine TNA when regulating the thrust of the rocket engine by changing the temperature of the gas in front of the turbine and the associated change in the mass flow of gas. By this method, engines are controlled that have a two-component gas generator for generating a turbine working fluid (see the diagram in the book of T.M. Melkumov et al. Rocket Engines, M .: Mashinostroenie, 1968, p. 11, Fig. 1.5).

The disadvantage of this method is the inability to force the engine by more than 5-10% due to a significant increase in the temperature of the generator gas or a decrease in its economy in nominal mode when operating at a low temperature of the generator gas.

A known method of boosting the thrust of a liquid propellant rocket engine containing a gas turbine driven by steam of one of the fuel components formed in the cooling path of the combustion chamber, based on a change in the energy parameters of operation, namely, an increase in the temperature of the gas in front of the turbine, characterized in that the steam is injected with a metered amount of another fuel component before it is fed to the turbine and the resulting fuel mixture is ignited - prototype (see RF patent No. 2451202 of 04/27/2 011, IPC F04D 29/00, F02K 11/00).

The disadvantage of this technical solution is the complication of the engine circuit and the reduction of reliability indicators (new units are introduced: a gas generator with an ignition system, automation and regulation units), as well as the possibility of soot formation in an additional gas generator of an oxygen-hydrocarbon rocket engine.

A liquid-propellant rocket engine is known, comprising a combustion chamber with a cooling path and a nozzle head, a fuel pump, an oxidizer pump and a turbine communicated by an inlet with a cooled path of the combustion chamber, and a device containing a diffuser and a nozzle is mounted in a pipe connecting the chamber cooling path and the turbine.

The disadvantages of this technical solution are the disadvantages described above.

Known liquid rocket engine containing a combustion chamber with a cooling path and a nozzle head, a gas generator, a fuel pump, an oxidizer pump and a turbine communicated with the inlet of the gas generator, and the output with the nozzle head having an additional turbine, the input of which is communicated with the exit of the cooling path, and the output is with a nozzle head (see RF patent No. 2352804, IPC F02K 9/44, dated December 6, 2007 - prototype).

A feature of the liquid propellant rocket engine adopted for the prototype is that the temperature of the vapor of the fuel component evaporated in the chamber jacket formed by the deterministic heat removal value (for a fixed combination of the heat-transfer surface area and the mass flow rate of the fuel component through the cooling path) is low (450-500 K). This temperature is much lower than the permissible level under the condition of ensuring the turbine operability (up to 1200 K) and, significantly, with increasing thrust, the temperature decreases to the level of 150-250 K, thereby reducing the effective power of the turbine.

Since for engines with a thrust of 250 ÷ 400 tf the power of an additional turbine is 10-15% of the power of the main turbine, the main boost in engine thrust (1.2 ... 1.3 times) is due to an increase in the gas temperature in front of the main turbine, this leads moreover, the turbine in the boost mode operates at the maximum possible temperature of 800 ÷ 000 K, and in the nominal mode it operates at a temperature of 550 ÷ 750 K, which reduces the average trajectory efficiency of the LPRE.

The task of the invention is to eliminate the noted disadvantages of the prototypes, namely increasing the efficiency of the liquid propellant rocket engine at the nominal operating mode and further increasing (more than 1.3 times) the thrust during engine boosting.

The problem is solved in that in the known method of boosting the thrust of the rocket engine, based on a change in the energy parameters of functioning, according to the invention, the boosting is carried out by supplying a part of the gas from the gas path of at least one of the fuel components, or the generator gas, or their mixture, at least one additional turbine that interacts with at least one of the main ТНА, and after exiting from it gas is sent for further use or disposal.

In addition, at the exit from the additional turbine, gas is directed to the inlet of the discharge nozzle, or to the chamber nozzle, or to the inlet of the additional turbine, or to the heat exchanger.

The specified method is implemented in a liquid-propellant rocket engine containing at least one chamber, control and regulation units, TNAs with a main and at least one additional turbine, the gas path of the main turbine being connected to the entrance to the chamber, in which according to the invention a gas path of at least one of the components or their products of combustion is equipped with an additional line connecting it to the input of at least one additional turbine and exit from it, while the line is equipped with a local control system, located ennoy before or after an additional turbine, and the output line connected to the gas evacuation system and / or system reuse.

In addition, the gas reuse system after the additional turbine is made in the form of at least one additional turbine or line connecting the output of the additional turbine to the supersonic part of the chamber nozzle, or an additional discharge nozzle, or heat exchanger.

The essence of the proposed rocket engine is illustrated by the schematic diagrams shown in the drawings: FIG. 1 is a diagram of a rocket engine with a gas generator, FIG. 2 is a diagram of a rocket engine without a gas generator, where the following notation is adopted:

1, 2 - lines for supplying fuel components to pumps;

3, 4 - pumps;

5, 6 - lines for the removal of fuel components from pumps;

7 - line supply cooler to the chamber;

8 - gas generator;

9 - camera;

10 - gas supply line to the main turbine;

11 - the main turbine;

12 - regulation system;

13 - additional turbine;

14 - gas supply line to the nozzle manifold of the chamber;

15 - gas discharge nozzle;

16 - gas supply line after cooling the chamber to the main turbine;

17 - line for supplying the fuel component after the pump to the mixing head of the chamber.

The proposed engine (Fig. 1) consists of highways for supplying fuel components 1 and 2, pumps 3 and 4, highways for withdrawing fuel components from pumps 5 and 6, highways for supplying 7 coolers to the chamber, gas generator 8, chambers 9, highways for supplying gas 10 to the main turbine 11, the control system of the engine 12, the additional turbine 13, the gas supply line to the nozzle manifold of the chamber 14, the gas discharge nozzle 15, the gas supply line after cooling the chamber to the main turbine 16, the fuel component supply line after the pump mixing chamber head 17.

The engine operates as follows.

The fuel components come from the launch vehicle (LV) tanks along the lines 1 and 2 to the input of the pumps 3 and 4. From the pumps, the fuel components go through the lines 5, 6 to the gas generator 8, and along the highway 7 to the cooling chamber 9. From the gas generator along the line 10 it enters the main turbine 11, after which the main flow goes to the mixing head of the chamber, and the remaining flow rate, passing through the control system 12, enters the input of the additional turbine 13, after which it can be discharged along the line 14 to the collector on the nozzle of the chamber or to the nozzle discharge 15. bezgazogeneratornyh circuits (2), the gas after cooling chamber through line 16 is input to the main turbine and further moves as previously described schemes for the gasification, and one of the components of the fuel after the pump on line 17 enters the mixing chamber head.

The supply of a part of the gas to the additional turbine after the main turbine allows a significant increase in power (more than 1.3 times) due to the pressure difference on the additional turbine (π _T = Pvh / Pvyh) without increasing the gas temperature on the main turbine. At the same time, accordingly, costs and pressures increase, including in the combustion chamber. In boost modes, engine efficiency decreases by 5-10 kgf · s / kg, however, with an increase in starting overload from 1.2 ÷ 1.5 to 1.5 ÷ 2.5, gravitational losses decrease, which not only compensates for the loss in efficiency, but also provides technical profitability of this offer.

It should be noted that the possibility of forcing the engine allows, in case of failure of some engines, to perform the flight mission.

Claims (4)

1. A method of forcing the thrust of a liquid propellant rocket engine based on a change in the energy parameters of operation, characterized in that the forcing is carried out by supplying part of the gas from the gas path of at least one of the fuel components, or the generator gas, or a mixture thereof, to at least one an additional turbine interacting with at least one of the main turbopump units, and after exiting from it, gas is sent for further use or disposal. 2. The method according to claim 1, characterized in that at the exit from the additional turbine the gas is directed to the inlet of the discharge nozzle, or to the nozzle of the chamber, or to the inlet of the additional turbine, or to the heat exchanger. 3. A liquid rocket engine containing at least one chamber, control and regulation units, a turbopump unit with a main and at least one additional turbine, the gas path of the main turbine being connected to the chamber entrance, characterized in that the gas path is at least one of the components or their products of combustion is equipped with an additional line connecting it to the input of at least one additional turbine and exit from it, while the line is equipped with a local control system, p memory location before or after an additional turbine, and the output line connected to the gas evacuation system and / or system reuse. 4. The liquid rocket engine according to claim 3, characterized in that the gas utilization system after the additional turbine is made in the form of at least one additional turbine or line connecting the output of the additional turbine to the supersonic part of the chamber nozzle, or an additional discharge nozzle, or heat exchanger.

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