RU2801019C1 - Method of operation of closed cycle liquid rocket engine with afterburning of oxidizing and reducing generator gases without complete gasification and liquid rocket engine - Google Patents

Abstract

FIELD: rocket and space technology.

SUBSTANCE: proposed group of inventions relates to the design of liquid rocket engines. The invention discloses the use of a scheme with two gas generators and afterburning of reducing and oxidizing generator gases without complete gasification. The non-gasified residue is fed into heat exchangers and used to pressurize the fuel tanks, which increases the payload mass by reducing the mass of the tank pressurization systems and reduces the overall cost of the installation.

EFFECT: invention provides increased efficiency of a high thrust rocket engine and high payload.

2 cl, 1 dwg, 2 ex

Description

The proposed group of inventions relates to the field of rocket and space technology, namely to the design of liquid rocket engines. The claimed technical solutions are intended for propulsion systems used in the first stages of heavy and super-heavy launch vehicles in order to launch a payload weighing up to 100 tons into low earth orbit.

The prior art method of compensating for differences in the physical properties of fuels in a universal generatorless liquid-propellant rocket engine and a liquid-propellant rocket engine (hereinafter referred to as LRE) (Patent RU 2358142, Publ. 10.06.2009, Bull. No. 16). In a method for compensating for differences in the physical properties of fuel components, based on matching the operating modes of the universal rocket engine supply units, according to the invention for a generatorless engine with separate HP, when it is transferred from hydrogen to LNG (methane), the fuel consumption (LNG, methane) is first increased to the required value for ensuring reliable cooling of the chamber, after cooling before supplying fuel to the TNAG turbine, its total consumption is divided into two parts, one of which is fed to the TNAG turbine, and the other is dumped, and after passing through the TNAG, the process of fuel division is repeated, while one part of it is sent for combustion in the combustion chamber, and the other is discarded or sent for further use. The discharged parts of the fuel flow can be used as a working medium, for example, for steering nozzles, for a turbine for driving an engine oscillation system, for pressurizing tanks, repeatedly as a working medium for a fuel pump and/or chamber fuel. EFFECT: invention ensures operation of the engine both on fuel components "oxygen+hydrogen" and on fuel "oxygen+liquefied natural gas" (methane). The disadvantage of this technical solution is the impossibility of obtaining high pressures in the chamber and, as a result, high specific impulses, since the proposed scheme of the propulsion system does not have a gas generator(s).

A known method of operating a liquid-propellant rocket engine with a turbopump supply of oxygen-methane fuel (patent RU 2166661, publ. 10.05.2001 Bull. No. 13). 20-50% of the methane fuel entering the engine is spent on regenerative cooling of the chamber, after which it is burned in it, and the oxygen oxidizer is fed partly directly into the chamber and partly into the reduction gas generator, where excess fuel entering the gas generator at a pressure above the initial pressure is burned in the oxidizer refrigerant; the resulting reducing gas is used in the turbine, after which it is burnt out in the chamber. The disadvantage of this technical solution is the high consumption of reducing gas required to ensure the operation of the pumps, the low degree of fire safety due to the high operating temperatures of the generator gas and the location of the impellers of the oxidizer pump and fuel on the same shaft.

A known method of operating a liquid-propellant rocket engine and a liquid-propellant rocket engine (RU 2116491 C1, publ.: 27.07.1998). The engine is designed for use in rocket science, as well as for single-stage means of launching payloads into near-Earth orbit. It contains a gas generator, turbopump units of fuel and oxidizer, a chamber with a regenerative cooling path, units of an automation system and control of engine parameters. The fuel line units operate sequentially and continuously on methane and hydrogen. All the fuel is converted into a reducing generator gas, a small part of which is supplied to the turbine and then discharged into the chamber nozzle, and the rest is burnt with liquid oxygen. The engine is equipped with a unit for the sequential supply of liquid methane and hydrogen, installed in front of the fuel pump, and the gas generator is located coaxially to the chamber. Moreover, the output from it is simultaneously the input of the camera. The disadvantage of this technical solution is the complexity of the design due to the use of three fuel components, lower costs of generator gas compared to schemes using two gas generators of various types, as well as high operating temperatures of the turbines.

Known liquid rocket engine (RU 2179650, publ. 20.02.2002 Bull. No. 5). A liquid rocket engine includes a combustion chamber with a regenerative cooling path, an evaporator, pumps for supplying fuel and oxidizer components, and a turbine. The outlet from the supply pump of one of the components is connected by means of a line to the inlet to the evaporator through the refrigerant line, and the outlet from the evaporator through the same line is connected to the inlet to the turbine, and the outlet of the pump of the other component is connected to the combustion chamber. The engine also includes a condenser, a valve-operated intercooler source, an intercooler circulation pump with an intercooler turbine. The evaporator inlet through the coolant line is connected to the outlet of the intercooler turbine, the evaporator outlet through the coolant line is connected to the source of the intercooler by means of a controlled valve and to the inlet to the intercooler circulation pump. The output of the intercooler circulation pump is connected to the inlet to the regenerative cooling circuit of the combustion chamber, and the outlet from the duct is connected to the inlet to the intercooler turbine. In this case, the condenser inlet through the refrigerant line is connected to the outlet of the pump of one of the components, and the outlet through the same line is connected to the combustion chamber. The inlet and outlet of the condenser through the coolant line are respectively connected to the outlet of the turbine and the inlet to the pump of the same component. Calculations show that using the proposed LRE power supply scheme, it is possible, for example, to create a pressure of 180 atm in the combustion chamber for an engine with a thrust of 8 tons on oxygen + kerosene fuel at a gasified oxygen temperature of ~ 600 K, while the classical scheme with the afterburning of an oxidizing gas generator gas provides at a gas generator gas temperature of ~ 700 K and other equal conditions, the pressure in the COP is about 120 atm. The disadvantage of this technical solution is the presence of a complex system of heat exchangers and condensers, as well as a source of intercooler, which will impair the mass perfection of the system.

The development of the modern space industry requires the creation of launch vehicles that combine low launch costs, rapid reuse and high payloads in order to launch payloads of more than 100 tons into low Earth orbit. At the same time, the cost and labor costs in preparing the mission should be minimal. In this regard, the primary task is the production of an efficient rocket engine with high thrust and the highest possible technical characteristics.

The technical result provided by the set of features of the claimed group of inventions is the high efficiency of a rocket engine with high thrust and the highest possible technical characteristics.

This technical result is achieved by the fact that the method of operating a closed-cycle liquid-propellant rocket engine with afterburning of oxidizing and reducing generator gases without complete gasification includes gas pressurization from the starting cylinder of the oxidizer tank and fuel tank, the passage of fuel through the main fuel line and through the fuel pump into the regenerative cooling jacket the fire wall of the combustion chamber and the Laval nozzle, part of which is discharged through the jet nozzles of the combustion chamber, and the other part of the fuel, which has made a partial phase transition, is sent from the regenerative cooling jacket to the reduction and oxidizing gas generators, the passage of the oxidizer through the main oxidizer fuel line and through the oxidizer pump, one part of which is supplied to the reduction and oxidizing gas generators, the second part of the oxidizer is sent directly to the combustion chamber, combustion in the reduction and oxidizing gas generators of fuel mixed with the oxidizer, forming reducing and oxidizing generator gases, which are sent through the reducing gas turbines and the oxidizer pump to the mixing head , from where they are discharged into the combustion chamber for ignition, while the remaining non-gasified fuel residue is sent through pipelines to the fuel heat exchanger-evaporator, where it is gasified and pressurizes the fuel tank in an autogenous way, the third part of the oxidizer is sent to the oxidizer heat exchanger-evaporator, where it pressurizes the oxidizer tank in an autogenous way.

Also, the technical result is achieved by the fact that a liquid-propellant rocket engine containing a combustion chamber with a regenerative cooling jacket, fuel and oxidizer tanks, a starting cylinder for boosting tanks, oxidizer and fuel turbopump units, control and regulation devices, main pipelines, reduction and oxidizing gas generators installed on at the inlet of the corresponding turbine, the fuel heat exchanger-evaporator and the oxidizer heat exchanger-evaporator, while the remaining non-gasified fuel residue is supplied to the inlet to the fuel heat exchanger-evaporator from the regenerative cooling jacket of the combustion chamber, and the outlet is connected to the fuel tank, carrying out autogenous pressurization of the fuel tank, to the inlet to the heat exchanger-evaporator of the oxidizer after the pump receives a part of the oxidizer, and the outlet is connected to the oxidizer tank, carrying out the pressurization of the oxidizer tank in an autogenous way.

Technical solutions disclose the use of a scheme with two gas generators and afterburning of reducing and oxidizing generator gases without complete gasification. The non-gasified residue is fed into heat exchangers and used to pressurize the fuel tanks, which increases the payload mass by reducing the mass of the tank pressurization systems and reduces the overall cost of the installation. The regulation of the "oxidizer/fuel" ratio in the combustion chamber, which is carried out by changing the flow rates of the fuel pair components using throttle valves and control valves installed on the supply lines of each of the components, contributes to a rapid change in the thrust parameter.

The increased expansion ratio of the Laval nozzle and the high pressure in the combustion chamber (which is achieved by converting 95% of the mass flow rate into reducing and oxidizing generator gases) increase thrust and specific impulse, therefore, the overall efficiency of the entire system and the efficiency of the engine.

This development is an atmospheric version of a rocket engine with an SL (sea level) optimized Laval hypersonic nozzle. Structurally, the main block of the rocket engine consists of a combustion chamber, a Laval nozzle, turbopump units (TPU), gas generators and automation - electronic elements (strain gauge sensors, electric drives for thrust vector control, other system components) that provide a regular start, access to the nominal mode, execution of a cyclogram operation and shutdown of the propulsion system.

The combustion chamber is made according to a cylindrical scheme. In view of the relatively small size, the combustion chamber design involves a swinging suspension; thrust vector control is carried out using hydraulic pistons or electric actuators. In a liquid rocket engine, a scheme is implemented with the afterburning of oxidizing and reducing generator gases without complete gasification. The engine cycle is closed.

According to the technological solution, both fuel components pass through gas generators enriched with fuel and oxidizer. Before entering the gas generators, the fuel entering the combustion chamber, being heated, passes through the regenerative cooling jacket, which lowers the temperature of the walls of the combustion chamber and the Laval nozzle to acceptable values, at which the deterioration in the strength properties of the fire wall of the combustion chamber due to heating makes it possible to keep working loads with a sufficient margin strength. However, 2-4% of the regenerative jacket inlet flow will be used for curtain cooling, making the outlet flow about 96-98% of the jacket inlet flow. For the present solution, the indicator will be about 800 K. A curtain of fuel is fed through the nozzles into the Laval nozzle. As the fuel passes through the regenerative cooling jacket, it is partially gasified. At the moment it enters the gas generators, the oxidizer is in a liquid state of aggregation. Part of the fuel (2-4% of the flow rate going to the propulsion system) used for screen cooling of the walls of the regenerative gas generator may not be fed into the regenerative cooling jacket, but be directed through small nozzles to the gas generator. Part of the non-gasified balance of fuel components (in rocket science, fuel and oxidizer are considered to be fuel) passes through heat exchangers (evaporators) and is gasified, supplied to fuel tanks for pressurization and supply of oxidizer and fuel to the propulsion system.

The supply of fuel components is carried out through turbopumps.

Refueling of the apparatus with the proposed propulsion system is carried out immediately before the start. Thrust control is controlled by the on-board control system by changing the flow rate of the fuel pair components through the chamber and gas generators, as well as by changing the oxidizer-fuel ratio by regulating the non-gasified flow rates supplied to the combustion chamber.

The most appropriate ignition method is electric.

Due to the increase in the consumption of gases exhausted by HP turbines, it is possible to achieve high pressures in the combustion chamber (about 30 MPa), which will give maximum efficiency at the geometric degrees of expansion of the Laval nozzle ≈40, which can also lead to overexpansion of the flow when operating at sea level.

The supercritical part of the Laval nozzle is profiled according to the parabola method without the implementation of an ideal contour with a reduced angle at the cut in order to avoid separation of the gas flow from the fire wall. For profiling the subcritical part, the rounding of the inlet to the nozzle and the criticism of the gas channel is taken into account.

The nozzle head can be made of steel. A spherical head is used, which has high strength. Generator gas nozzles - jet, non-gasified fuel components - centrifugal, two-component, mixing - external.

Speeds of circulation of pump shafts ≈30000 rpm. The number of stages for the oxidizer and fuel pumps is two. A multi-turbine scheme is used. The number of turbines is two.

The essence of the invention is illustrated by the scheme of the claimed installation, which shows:

1 - fuel tank

2 - oxidizer tank

3 - starting cylinder pressurization tanks

4 - heat exchanger-evaporator of fuel

5 - fuel pump impellers

6 - fuel pump auger

7 - reducing gas turbine

8 - recovery gas generator

9 - fuel pump shaft

10 - oxidizer pump shaft

11 - oxidizer pump turbine

12 - oxidizing gas generator

13 - oxidizer pump impellers

14 - oxidizer pump screw

15 - heat exchanger-evaporator of the oxidizer

16 - mixing head

17 - regenerative cooling jacket

18 - combustion chamber and Laval nozzle.

The principle of operation of the installation using the scheme with afterburning of reducing and oxidizing gases without complete gasification is as follows.

Pressurization gas (for example, helium He/nitrogen N ₂ ) enters the oxidizer tank 2 and fuel tank 1 from the starting pressurization tank of tanks 3, which starts the operation of the propulsion system (LPRE). Then the turbopump unit spins up and ignites. The propulsion system goes into mode. The fuel passes through the main fuel line and enters the screw 6 of the multistage fuel pump. The screw 6 of the fuel pump increases the initial pressure of the fuel without cavitation breakdown, after which the fuel passes into the impellers of the fuel pump 5. At the outlet, the fuel pressure rises to values of tens of MPa, after which the fuel flow is directed to the regenerative cooling jacket 17 of the fire wall of the combustion chamber and nozzle Laval 18 to mitigate the conditions of her work. The regenerative cooling jacket 17 consists of an inner fire wall of the combustion chamber and a Laval nozzle, an outer casing and internal stiffening ribs. The coolant (fuel) is pumped through the internal channel formed by the cavity between the outer casing and the fire wall of the combustion chamber and the Laval nozzle 18. Part of the fuel (2-4% of the flow rate going into the regenerative cooling jacket) exits through the air curtain jet nozzles located on the inner side of the fire wall in the critical sections of the Laval nozzle 18. The remaining fuel flow rate (96-98% of the total mass flow rate of the fuel entering into the jacket), which has made a partial phase transition, leaves the regenerative cooling jacket 17. Part of the remaining fuel consumption is supplied to the combustion chamber 18, to the jet nozzles of the chamber curtain. Most of the fuel consumption (about 90%), which has passed through the regenerative cooling jacket, in different quantities enters the reduction 8 (RGG) and oxidizing 12 (OGG) gas generators, where it burns in a mixture with an oxidizer, forming reducing and oxidizing generator gases.

In turn, the reducing and oxidizing generator gases rotate the reducing gas turbine 7 and the oxidizer pump turbine 11 located on the fuel pump shaft 9 and on the oxidizer pump shaft 10. The reducing gas turbine 7 is located on the fuel pump shaft 9. The oxidizing gas turbine 11 is located on oxidizer pump shaft 10. After that, the reducing and oxidizing generator gases, the pressure of which was reduced as a result of work on the turbines, go to the mixing head 16, where they exit the jet nozzles of the combustion chamber, and ignite.

The remaining non-gasified fuel residue passes through pipelines to the fuel heat exchanger-evaporator 4, where it is gasified and inflates the fuel tank 1 in an autogenous way.

Pressurization gas (for example, helium He/nitrogen N2), which is also used to pressurize oxidizer tank 2, is supplied from the start-up cylinder for pressurization of tanks 3 at the moment the main pressurization valve begins to open. Gas flows are separated using a tee. To obtain a constant outlet pressure, it is advisable to install a reducer. After that, the oxidizer passes through the main oxidizer fuel line and enters the auger 14 of the oxidizer pump. The oxidizer pressure rises to values of several MPa to avoid cavitation failure, after which the oxidizer flow enters the impellers of the oxidizer pump 13, where its pressure increases to values of tens of MPa. An additional line is separated from the main line of the oxidizer, supplying the oxidizer to the reduction (VGT) 8 and oxidizing (OGG) 12 gas generators, which ensures their operation. Part of the oxidizer, which is 7-10% of the total flow rate (oxidizer) entering the propulsion system, is not supplied to gas generators 8 and 12, but is sent to the heat exchanger and to the nozzle head of the combustion chamber 18, being injected through special nozzles located on the nozzle head and serving to change the "oxidizer/fuel" ratio by throttling the flow rate through the supply pipe to the combustion chamber 18, which contributes to a rapid change in thrust and specific impulse of the engine. In particular, this makes it possible to simplify the change in thrust at the stage of entry of the conditional first stage of a launch vehicle with a liquid-propellant rocket engine into the dense layers of the Earth's atmosphere to return to the launch (landing) site.

Part of the oxidizer is supplied to the oxidizer evaporator 15, and pressurizes the oxidizer tank 2 in an autogenous way.

During the start-up procedure, the oxidizer supply is opened first, then the fuel supply. The interval between feeds depends on the individual technical characteristics of the system and the sequence diagram for starting the propulsion system. On average, the interval is several hundred milliseconds. After the engine enters the mode, the pressurization of the tanks of fuel components 1 and 2 with gas from the start-up cylinder for pressurizing tanks 3 is turned off, and the propulsion system switches to the pressurization mode with autogenous vapors.

The total mass fraction of non-gasified residue from the total second mass flow rate of fuel components is ~5%. This indicator allows you to simultaneously increase the consumption of components going to the gas generators, and leave a sufficient margin for autogenous pressurization of tanks. The remainder of the non-gasified fuel will also be used to control the oxidizer/fuel ratio in the combustion chamber and to create a curtain. Thus, the proportion of non-gasified residue going to the combustion chamber will be changed by means of component flow controllers, which will lead to a change in the oxidizer/fuel ratio in the combustion chamber. Such a combination will allow using the proposed method with maximum benefit.

The proposed scheme makes it possible to combine high efficiency, thrust, pressure in the combustion chamber and engine safety of full gasification schemes with the possibility of autogenous pressurization, simplified throttling and liquid-propellant rocket cooling, which affects the overall technological process, cost and profitability of the propulsion system as a whole.

The claimed technical solution when using the CH ₄ +O ₂ fuel pair can have the following characteristics:

- increased specific impulse (s) - 355 (Vac) / 330 (SL);

- high thrust (N) - 2181662 (Vac) / 2024484 (SL);

- high pressure in the combustion chamber ≥30 MPa;

- reduced mass of the fuel tank pressurization system and, as a result, an increase in the characteristic speed of the launch vehicle;

- the possibility of scaling (changing thrust indicators and product dimensions) and modernization (updating the product, bringing it into line with new requirements and standards, technical conditions, quality indicators);

- low cost of the product, which fully meets the modern tasks of the functioning of the orbital launch market.

Example 1. Methane CH4 is used as a fuel. The oxidizing agent is oxygen O2. Despite the fact that this fuel pair (CH4 + O2) is inferior in specific impulse to the hydrogen/oxygen mixture (H2 + O2), the main prerequisite for using methane is the low cost of the product. The boiling point of liquid methane is cryogenic, but higher than the boiling point of hydrogen. The combustion of methane occurs more “smoothly”, which cannot be said about the fuel pair “kerosene + oxygen”. The boiling points of liquid oxygen and methane are close, which simplifies the design of filling infrastructure. It is also possible to use supercooled fuel components to increase the specific impulse of the engine and reduce the size of the tanks. Methane is not a self-igniting, carcinogenic and toxic substance, which speaks in favor of the environmental component of the choice.

Example 2. Liquid hydrogen H ₂ is used as a fuel, liquid oxygen O ₂ is used as an oxidizing agent. The pressure in the combustion chamber is 10 MPa, the geometric expansion ratio of the nozzle is 14.5. With these parameters, the specific impulse of the rocket engine will be higher than when using CH ₄ and will be 3778.56 m/s at sea level and 4130.33 m/s in vacuum, which will provide a higher characteristic velocity of the launch vehicle, an increase in the mass of the payload, and a reduction in the required mass of fuel. The required fuel temperature will be ≈20 K, which is due to the boiling of liquid hydrogen. The temperature in the combustion chamber at the optimal oxidizer-fuel ratio of 4.627 will be ≈3194.57 K.

However, in this case, problems may arise with the use of hydrogen due to its cryogenicity, the need to introduce booster pumps into the propulsion system, as well as the low density of the component, which leads to an increase in the required volume of tanks.

At the same time, other technical characteristics of the system (thrust, total second mass fuel consumption, mass flow rates going to gas generators, thrust-to-weight ratio) depend on parameters that are individual for each carrier and are set by the terms of reference.

The scheme (method) presented in the patent can use almost any fuel pair, however, the main application criteria is the technical and economic feasibility of each of the possible solutions.

Thus, the claimed technology will allow to combine the advantages of full-flow liquid rocket engines (that is, LRE systems with full gasification), such as high efficiency compared to the "classic" closed cycle, high pressure in the combustion chamber, lower turbine operating temperature, high degree of explosion - and fire safety, with simplified throttling of oxidizer and fuel flows, the injection of which is carried out by means of special nozzles located on the chamber head, autogenous pressurization of tanks and curtain cooling, which reduces the temperature of the fire wall. EFFECT: technical result makes it possible to simplify the control of LRE thrust, its specific impulse, to obtain high efficiency indicators of LRE along with high reliability and safety of the propulsion system, to simplify the operation of thrust throttling systems by changing the value of non-gasified control supplied to the injectors.

Claims (8)

1. A method of operating a liquid-propellant rocket engine of a closed cycle with afterburning of oxidizing and reducing generator gases without complete gasification, including pressurization with gas from the starting cylinder of the oxidizer tank and the fuel tank, the passage of fuel through the main fuel line and through the fuel pump into the regenerative cooling jacket of the fire wall of the combustion chamber and the Laval nozzle, part of which is removed through the jet nozzles of the combustion chamber, and the other part of the fuel, which has made a partial phase transition, is sent from the regenerative cooling jacket to the recovery and oxidizing gas generators, the passage of the oxidizer through the main oxidizer fuel line and through the oxidizer pump, one part of which is supplied to the reduction and oxidizer gas generators, the second part of the oxidizer is sent directly to the combustion chamber, combustion in the reducing and oxidizing gas generators of fuel mixed with the oxidizer, forming reducing and oxidizing generator gases, which are sent through the reducing gas turbines and the oxidizer pump to the mixing head, from where they are discharged into the combustion chamber for ignition, at the same time, the remaining non-gasified fuel residue is sent through pipelines to the fuel heat exchanger-evaporator, where it is gasified and pressurizes the fuel tank in an autogenous way, the third part of the oxidizer is sent to the heat exchanger-evaporator of the oxidizer, where it pressurizes the oxidizer tank in an autogenous way. 2. A liquid-propellant rocket engine containing a combustion chamber with a regenerative cooling jacket, fuel and oxidizer tanks, a starting cylinder for boosting tanks, oxidizer and fuel turbopump units, control and regulation devices, main pipelines, characterized in that it also includes reduction and oxidizing gas generators , installed at the inlet of the corresponding turbine, a heat exchanger - a fuel evaporator and a heat exchanger - an oxidizer evaporator, while the remaining non-gasified fuel residue is supplied to the inlet to the heat exchanger - fuel evaporator from the regenerative cooling jacket of the combustion chamber, and the outlet is connected to the fuel tank, carrying out autogenous pressurization of the tank fuel, a part of the oxidizer enters the inlet to the heat exchanger - the oxidizer evaporator after the pump, and the outlet is connected to the oxidizer tank, carrying out the pressurization of the oxidizer tank in an autogenous way.