Dependence of power and efficiency pump on its volumetric capacity. Layout diagrams of the pump. Technological process for manufacturing the blade.

TNAs are divided into single-shaft and multi-shaft. In single-shaft pumps, the turbine and pumps are located on the same shaft. The advantage of TNAs made according to this design is their simplicity of design and low weight. As a disadvantage, it should be noted that only one of the pumps (usually the oxidizer pump) operates at the optimal speed. In this case, the fuel pump is operated at reduced efficiency values.

The following TNA layout diagrams are distinguished, Fig. 57.

With a three-shaft pump design, the speeds of the pumps and turbine are independent of each other and can be selected from the conditions for optimal pump operation. However, the presence of gearboxes operating in difficult conditions (high peripheral speeds, difficulty in providing an effective lubrication and cooling system) in some cases minimizes the benefits from increasing pump efficiency values.

Single-shaft

Three-shaft

TNA layout diagrams

Single-shaft TPU designs are most widespread in liquid propellant rocket engines.

5.3. Centrifugal pump device

In liquid propellant rocket engines, centrifugal pumps are usually used as the main ones. The main advantages that determine the primary use of these types of pumps in liquid propellant engines are:

Ensuring high supply pressures and productivity with small dimensions and weight;

Ability to work with aggressive and low-boiling components;

The ability to work with a high number of revolutions and the convenience of using a turbine to drive them.

Figure 58 shows a diagram of a single-stage centrifugal pump. The liquid is supplied through inlet pipe 1 to the rotating wheel (impeller) 2. In the pump wheel, the liquid moves through a channel formed by the walls of the wheel and the blades 3. The force exerted by the wheel blades on the liquid causes it to move so that the energy reserve per unit mass of the liquid increases. In this case, there is an increase in both potential energy (static pressure) and kinetic energy of the liquid.

Fig.58

Centrifugal pump diagram:

1 - inlet pipe; 2 - pump wheel (impeller); 3 - shoulder blades;

4 - diffuser; 5 - diffuser blades; 6 - collection or snail; 7 - front seal;

8 - shaft bearing; 9 - bearing seal

At the exit of the wheel, the liquid enters the diffuser 4, where its absolute speed decreases and the pressure additionally increases. The simplest diffuser consists of smooth disks that make up its walls and is called vaneless. The vane diffuser has fixed blades 5 (shown in dotted lines in Fig. 58), which contribute to faster damping of the flow velocity. After passing through the diffuser, the liquid enters the spiral channel (cochlea) 6, the purpose of which is to collect the fluid coming out of the wheel and also reduce its speed. The liquid is supplied to the network through the discharge pipe.

To reduce the flow of liquid from the high-pressure cavity (diffuser, volute) to the low-pressure area, seals 7 are made in the pump.

Fig.59

Schemes of centrifugal pumps:

a-c axial entrance; b- with spiral entry;

V- with two-way entrance; G-multistage pump

Centrifugal pumps are available with axial, spiral and double inlet, single and multi-stage. Choice of axial or spiral input (Fig. 59, a, b) determined primarily by the layout conditions of the pump and propulsion system. Double input (Fig. 59, V) are performed at high flow rates to reduce the inlet speed and thereby improve the anti-cavitation properties of the pump. Multistage pumps (Fig. 59, G) are used when it is necessary to obtain particularly high pressures.

Typically, pump housings are made of casting from high-strength aluminum alloys, and in the case of high pressures - from steel. The number of profiled impeller blades is no more than 8, and their thickness is in the range of 2 ¸ 5 mm.

5.4. Pump impellers

There are impellers of open and closed types, Fig. 60 (a, b).

An open impeller is used in pumps with low flow and low pressure components. An impeller of this type is characterized by significant losses caused by the flow of a component from an area of ​​high pressure (at the outlet of the pump) to an area of ​​low pressure (at the inlet to the pump). The impeller consists of a disk 1 and blades 2 mounted on it.

In closed impellers, a cover 3 is installed on the end surfaces of the blades, which can be made integral with the impeller. In impellers of this type, losses due to component flow are significantly less than in open impellers. Typically, impellers are made by casting. The number of profiled blades, as a rule, does not exceed 8, and their thickness is less than 5 mm. The impellers shown in Fig. 60 are impellers with a one-way component supply.

To reduce the flow rate of the component through the blade channel of the impeller (in order to eliminate the occurrence of the cavitation process), impellers with two-way supply of the component are used, Fig. 61.

Fig.60

Single-sided impellers:

a- open type; b – closed type

Fig.61

Double-sided impeller

8.5. Impeller seals

In order to reduce fluid leakage, the following types of seals are installed in pump impellers: slotted, labyrinth and floating, Fig. 62 a, b, c, respectively.

The operating principle of gap seals is based on ensuring high hydraulic resistance of the annular gap between the graphite liner installed in the pump body and the groove made in the inlet section of the disk. The design of this seal allows up to 15% of leakage from the volume of pumped liquid, while the labyrinth, Fig. 62 b, and floating (a set of fluoroplastic and aluminum washers installed in the inlet section of the impeller), Fig. 62 c, - up to 10% and 5%, respectively.

Fig.62

Impeller seals:

a – slotted; b – labyrinthine; in - floating

5.5. TNA turbine

One of the main elements of the pump is the gas turbine. In a turbine, the potential energy of combustion products from a gas generator or coolant vapor is converted into mechanical work of the turbine. The turbine is designed to drive TNA pumps into rotation. The turbine consists of a nozzle apparatus 1, an impeller 2 with two rows of working blades 3 and 4, a guide vane 5 and a turbine housing 6 with an outlet pipe 7, Fig. 75.

The first stage of the turbine is a combination of the nozzle apparatus 1 and the blades of the impeller 3, the second is formed by the fixed blades of the guide vane 5 and the second row of working blades 4.

The conversion of the enthalpy of the gas flow into the mechanical energy of shaft rotation is carried out in two stages: the enthalpy of the gas flow - into the kinetic energy of the jet (in the nozzle apparatus); kinetic energy of the jet - into mechanical energy of shaft rotation (on the impeller).

Fig.75

TNA turbine design

The shafts of turbopump units (TPA) operate under high loads and high speeds. To lighten their weight, they are made hollow. The highest alternating stresses in the shaft metal occur on its outer surface. In this case, any kind of sharp transitions, marks from a cutting tool and other surface defects are stress concentrators. Cracks may form in these places during operation, which will lead to shaft breakage. Therefore, special attention is paid to the cleanliness of the shaft surface with the introduction of hardening operations in some cases. Not only the areas for bearings, seals, and fits are subject to finishing, but also all other areas of the shaft that are not mated to other parts.

High speeds (10000-20000 rpm and more) force the designer to set very tight tolerances for the alignment of journals and seats, the accuracy of the location of the axial hole, the difference in thickness and other dimensions. The slightest geometric errors lead to uneven distribution of the rotating masses of metal, which causes vibration and shaking of the pump.

5.6. Requirements for gas generators

The thrust value of a liquid-propellant rocket engine is known to be a linear function of fuel consumption per second. The second fuel consumption for each specific engine with a pump component supply system depends on the power developed by the turbine. The power of the turbine is completely determined by the second flow rate and the parameters of the working fluid at the entrance to the turbine, i.e., at the exit from the gas generator. Therefore, the gas generator is a device that sets the operating mode of the entire propulsion system. This circumstance determines the special requirements for this part of the fuel supply system (in addition to general requirements requirements for all LRE units, regardless of the specifics of their operation). These requirements boil down to the following.

1. High stability. This means that the gas generator in all engine operating modes must provide the specified second gas flow rate as accurately as possible, and at the same time, the values ​​of gas parameters (composition, pressure, temperature, etc.) should not go beyond certain (permissible) limits. The more stable the operation of the gas generator, the less load the engine control systems experience in flight, and this increases engine reliability and shooting accuracy.

The stability of the gas generator is especially important for rockets with unregulated rocket engines and rockets whose flight range is controlled only by the flight speed at the end of the active part of the trajectory. In the latter case, the deviation of the coordinates of the end of the active section of the trajectory, caused by the deviation of the engine thrust from the calculated value, due to the unstable operation of the gas generator, will entirely turn into a deviation of the missile's impact point from the target.

2. Ease of managing the workflow in a wide range of changes in its parameters. This requirement is also due to the regulating effect of the gas generator on the engine and the need to change the engine operating mode during one start (when adjusting thrust during take-off and in flight, when transitioning from the main stage of thrust to the final stage, etc.).

3. High efficiency of generator gas, which determines either minimum energy consumption (and, accordingly, minimum fuel consumption) for the TPU drive, or an increase in the TPU power. This requirement is put forward due to the fact that the specific impulse of the engine is determined by the ratio of thrust to the entire second consumption of the ejected mass. The concept of “discarded mass” includes both the products of fuel combustion in the chamber and the exhaust gas after the turbine. For liquid propellant engines, in which this gas is emitted into the atmosphere and develops a specific impulse less than the fuel combustion products flowing out of the engine chamber, the decisive condition for increasing engine efficiency is to reduce fuel consumption for the pump drive. For a liquid-propellant rocket engine with afterburning of generator gas, the main thing is to increase the power of the turbocharger, since this makes it possible to increase the pressure in the chamber and, at a given value of pressure at the nozzle exit, to increase the degree of expansion of the ejected combustion products, i.e., to increase the thermal efficiency of the chamber. Reducing fuel consumption for the TPU drive and increasing the TPU power depend on the amount of energy supplied to the turbine by one kilogram of the working fluid. This energy is equal, as is known, to the product of the relative effective efficiency of the turbine and the available adiabatic heat drop.

5.7. Classification of gas generators

The basis for the classification of gas generators is the method of producing generator gas. Currently, three methods of gas generation are common.

1. Decomposition (with or without catalysts) a substance that, after an external initiating influence, can proceed to further stable spontaneous decomposition, accompanied by the release of a significant amount of thermal energy and gaseous decomposition products. Such a substance can be either a component of the main engine fuel or a special gas-generating agent stored on board the rocket only for this purpose. Gas generators in which this process is implemented are called single-component. In the future, they are distinguished mainly by the type of decomposing substance (hydrogen peroxide, hydrazine, solid fuel, etc.).

2. Incineration liquid fuel consisting of two components. It is best to use the main engine fuel for this purpose, since this greatly simplifies its supply to the gas generator and improves the operating conditions of the rocket. Gas generators of this type are called two-component.

3. Liquid evaporation in the cooling path of the engine chamber. With this method of obtaining the turbine working fluid, the problem of cooling the walls of the engine chamber is simultaneously solved. Gas generators of this type are called steam generators, and engine circuits are called generatorless. Steam generator circuits are divided into circulation and with a change of working fluid. In the first, an arbitrary working fluid (for example, water) circulates through a closed loop “chamber cooling path - turbine - condenser - pump - chamber cooling path,” turning alternately into steam and then into liquid in its various parts. In schemes with a change of working fluid, this circulation is absent. The working fluid after the turbine is removed from the cycle. Obviously, the direct release of exhaust gas into the atmosphere would significantly worsen the engine's efficiency, since the specific thrust of the exhaust pipes is always less than the specific thrust of the engine chamber. To eliminate these losses, one of the fuel components is usually sent to the chamber cooling path. After evaporation and activation in the turbine, it is sent to the engine chamber, where it is burned along with the second component. Thus, generatorless engines are made according to the scheme with afterburning of the turbine working fluid.

The design of gas generation systems differs significantly from each other, but nevertheless, the following common basic elements can be distinguished in each of them:

Gas generator;

Fuel supply devices;

Automation.

In the gas generator (sometimes called a reactor), the working fluid of the turbine is directly formed - gas or steam of specified parameters. Fuel supply devices ensure the flow of gas generation means (starting materials) into the reactor. Automation regulates the work process, as well as starting and shutting down the gas generator. Sometimes (for example, when operating on main fuel) the gas generation system does not have independent fuel supply devices. In this case, the fuel supply to the gas generator is provided by the engine supply system.

The following types of gas generators (GG) are used in liquid propellant engines:

Solid fuel (TGG);

Hybrid (THG);

One-component liquid (one-component ZhGG);

Two-component liquid (two-component ZhGG);

Evaporative liquid (evaporative LGG);

1) Study of the diagram and operating principle of a liquid-propellant rocket engine (LPRE).

2) Determination of changes in the parameters of the working fluid along the path of the liquid-propellant rocket engine chamber.

1. GENERAL INFORMATION ABOUT LPRE

2.1. Composition of the rocket engine

A jet engine is a technical device that creates thrust as a result of the outflow of a working fluid from it. Jet engines provide acceleration for moving vehicles of various types.

A rocket engine is a jet engine that uses only substances and energy sources available on board a moving vehicle.

A liquid rocket engine (LPRE) is a rocket engine that uses fuel (the primary source of energy and working fluid) in a liquid aggregate state for operation.

In general, the rocket engine consists of:

2- turbopump units (TNA);

3- gas generators;

4- pipelines;

5- automation units;

6- auxiliary devices

One or more liquid propellant engines, together with a pneumatic-hydraulic system (PGS) for supplying fuel to the engine chambers and auxiliary units of the rocket stage, constitute a liquid-propellant rocket propulsion system (LPRE).

Liquid rocket fuel (LRF) is a substance or several substances (oxidizer, fuel) that are capable of forming high-temperature combustion (decomposition) products as a result of exothermic chemical reactions. These products are the working fluid of the engine.

Each LRE chamber consists of a combustion chamber and a nozzle. In the liquid propellant engine chamber, the primary chemical energy of the liquid fuel is converted into the final kinetic energy of the gaseous working fluid, as a result of which the reactive force of the chamber is created.

A separate turbopump unit of a liquid-propellant rocket engine consists of pumps and a turbine driving them. The TNA ensures the supply of liquid fuel components to the chambers and gas generators of the liquid propellant rocket engine.

A liquid propellant rocket engine gas generator is a unit in which the main or auxiliary fuel is converted into gas generation products used as the working fluid of the turbine and the working fluids of the pressurization system for tanks with liquid rocket engine components.

The LRE automation system is a set of devices (valves, regulators, sensors, etc.) of various types: electrical, mechanical, hydraulic, pneumatic, pyrotechnic, etc. Automation units provide startup, control, regulation and shutdown of the LRE.

LRE parameters

The main traction parameters of the rocket engine are:

The reactive force of a rocket engine - R - is the resultant of gas and hydrodynamic forces acting on the internal surfaces of the rocket engine when matter flows out of it;

Thrust of the rocket engine - R - the resultant of the reactive force of the rocket engine (R) and all pressure forces environment, which act on the external surfaces of the engine with the exception of external aerodynamic drag forces;

LRE thrust impulse - I - integral of the LRE thrust over its operating time;

Specific thrust impulse of a liquid-propellant rocket engine - I y - the ratio of thrust (P) to mass fuel consumption () of a liquid-propellant rocket engine.

The main parameters that characterize the processes occurring in the liquid rocket engine chamber are pressure (p), temperature (T) and flow rate (W) of the combustion (decomposition) products of liquid rocket fuel. In this case, the values ​​of the parameters at the inlet to the nozzle (section index “c”), as well as in the critical (“*”) and outlet (“a”) sections of the nozzle are especially highlighted.

The calculation of the parameter values ​​in various sections of the liquid-propellant rocket engine nozzle path and the determination of the engine's traction parameters are carried out using the corresponding thermogasdynamics equations. An approximate method of such calculation is discussed in section 4 of this manual.

1. DIAGRAM AND PRINCIPLE OF OPERATION OF THE RD-214 LPRE

3.1. general characteristics Liquid rocket engine "RD-214"

The RD-214 liquid rocket engine has been used in domestic practice since 1957. Since 1962, it has been installed on the 1st stage of multi-stage Cosmos launch vehicles, with the help of which many satellites of the Cosmos and Interkomos series have been launched into low-Earth orbits.

The RD-214 liquid-propellant rocket engine has a pump fuel supply system. The engine runs on a high-boiling nitric acid oxidizer (a solution of nitrogen oxides in nitric acid) and hydrocarbon fuel (kerosene processing products). A special component is used for the gas generator - liquid hydrogen peroxide.

The main engine parameters have the following meanings:

Thrust in void P p = 726 kN;

Specific thrust impulse in vacuum I pack = 2590 N×s/kg;

Gas pressure in the combustion chamber p k = 4.4 MPa;

Gas expansion ratio in the nozzle e = 64

Liquid rocket engine "RD-214", (Fig. 1) consists of:

Four cameras (item 6);

One turbopump unit (TPU) (items 1, 2, 3, 4);

Gas generator (item 5);

Pipelines;

Automation units (items 7, 8)

The engine THA consists of an oxidizer pump (item 2), a fuel pump (item 3), a hydrogen peroxide pump (item 4) and a turbine (item 1). The rotors (rotating parts) of the pumps and turbines are connected by one shaft.

The units and components that supply components to the engine chamber, gas generator and turbine are combined into three separate systems - lines:

Oxidizer supply system

Fuel supply system

Hydrogen peroxide steam and gas generation system.

Fig.1. Liquid rocket engine diagram

1 – turbine; 2 – oxidizer pump; 3 – fuel pump;

4 – hydrogen peroxide pump; 5 – gas generator (reactor);

6 – engine chamber; 7, 8 – automation elements.

3.2. Characteristics of the RD-214 liquid rocket engine units

3.2.1. LRE chamber

The four LRE chambers are connected into a single block along two sections using bolts.

Each LRE chamber (item 6) consists of a mixing head and a housing. The mixing head includes upper, middle and lower (fire) bottoms. A cavity for the oxidizer is formed between the upper and middle bottoms, and a cavity for fuel is formed between the middle and fire bottoms. Each of the cavities is connected with the internal volume of the engine housing by means of the corresponding injectors.

In the process of LRE operation, liquid fuel components are supplied, sprayed and mixed through the mixing head and its nozzles.

The liquid rocket engine chamber housing includes a part of the combustion chamber and a nozzle. The liquid-propellant rocket engine nozzle is supersonic and has a converging and diverging part.

The LRE chamber housing is double-walled. The inner (fire) and outer (power) walls of the body are interconnected by spacers. At the same time, with the help of spacers, channels of the housing liquid cooling path are formed between the walls. Fuel is used as a coolant.

During engine operation, fuel is supplied to the cooling path through special manifold pipes located at the end of the nozzle. Having passed through the cooling path, the fuel enters the corresponding cavity of the mixing head and is introduced into the combustion chamber through nozzles. At the same time, through another cavity of the mixing head and the corresponding nozzles, the oxidizer enters the combustion chamber.

In the volume of the combustion chamber, atomization, mixing and combustion of liquid fuel components occurs. As a result, a high-temperature gaseous working fluid of the engine is formed.

Then, in the supersonic nozzle, the thermal energy of the working fluid is converted into the kinetic energy of its jet, at the end of which the thrust of the rocket engine is created.

3.2.2. Gas generator and turbopump unit

The gas generator (Fig. 1, item 5) is a unit in which liquid hydrogen peroxide, as a result of exothermic decomposition, is converted into a high-temperature vapor working fluid of the turbine.

The turbopump unit provides pressure supply of liquid fuel components to the engine chamber and gas generator.

TNA consists of (Fig. 1):

Screw centrifugal oxidizer pump (item 2);

Screw centrifugal fuel pump (item 3);

Hydrogen peroxide centrifugal pump (item 4);

Gas turbine (item 1).

Each pump and turbine has a stationary stator and a rotating rotor. The rotors of pumps and turbines have a common shaft, which consists of two parts, which are connected by a spring.

The turbine (item 1) drives the pumps. The main elements of the turbine stator are the housing and the nozzle apparatus, and the rotor is the shaft and impeller with blades. During operation, peroxide gas is supplied to the turbine from the gas generator. When steam gas passes through the nozzle apparatus and blades of the turbine impeller, its thermal energy is converted into mechanical energy of rotation of the wheel and turbine rotor shaft. The exhaust steam gas is collected in the outlet manifold of the turbine housing and discharged into the atmosphere through special waste nozzles. In this case, some additional thrust of the rocket engine is created.

Oxidizer (item 2) and fuel (item 3) pumps are of screw-centrifugal type. The main elements of each pump are the casing and the rotor. The rotor has a shaft, a screw and a centrifugal wheel with blades. During operation, mechanical energy is supplied from the turbine to the pump through a common shaft, ensuring rotation of the pump rotor. As a result of the action of the screw blades and the centrifugal wheel on the liquid pumped by the pumps (fuel component), the mechanical energy of rotation of the pump rotor is converted into potential energy fluid pressure, which ensures the supply of the component to the engine chamber. An auger in front of the centrifugal wheel of the pump is installed to preliminarily increase the fluid pressure at the inlet into the inter-blade channels of the impeller in order to prevent cold boiling of the fluid (cavitation) and disruption of its continuity. Violations of the continuity of the flow of a component can cause instability of the fuel combustion process in the engine chamber, and, consequently, instability of the operation of the liquid propellant engine as a whole.

A centrifugal pump (item 4) is used to supply hydrogen peroxide to the gas generator. The relatively low flow rate of the component creates conditions for cavitation-free operation of the centrifugal pump without installing a screw prepump in front of it.

3.3. Engine operating principle

The engine is started, controlled and stopped automatically by electrical commands from the rocket to the corresponding automation elements.

For the initial ignition of fuel components, a special starting fuel is used, which is self-igniting with an oxidizer. The starting fuel initially fills a small section of the pipeline in front of the fuel pump. At the moment of launching the liquid-propellant rocket engine, starting fuel and oxidizer enter the chamber, their self-ignition occurs, and only then the main components of the fuel begin to be supplied to the chamber.

During engine operation, the oxidizer sequentially passes through the elements and assemblies of the main line (system), including:

Dividing valve;

Oxidizer pump;

Oxidizer valve;

Engine chamber mixing head.

The fuel flows through a pipeline that includes:

Dividing valves;

Fuel pump;

Manifold and cooling path of the engine chamber;

Mixing head of the chamber.

Hydrogen peroxide and the resulting steam gas sequentially pass through the elements and units of the steam and gas generation system, including:

Dividing valve;

Hydrogen peroxide pump;

Hydraulic reducer;

Gas generator;

Turbine nozzle apparatus;

Turbine impeller blades;

Turbine manifold;

Waste nozzles.

As a result of the continuous supply of fuel components by the turbopump unit to the engine chamber, their combustion with the formation of a high-temperature working fluid and the outflow of the working fluid from the chamber, thrust of the rocket engine is created.

Variation of the engine thrust value during its operation is ensured by changing the flow rate of hydrogen peroxide supplied to the gas generator. At the same time, the power of the turbine and pumps changes, and, consequently, the supply of fuel components to the engine chamber.

The liquid propellant engine is stopped in two stages using automatic elements. From the main mode, the engine is first transferred to the final operating mode with less thrust and only then is turned off completely.

1. WORK PROCEDURE

4.1. Scope and order of work

In the process of performing the work, the following actions are performed sequentially.

1) The design of the RD-214 liquid-propellant rocket engine is being studied. The purpose and composition of the liquid-propellant rocket engine, the design of the units, and the principle of engine operation are considered.

2) The geometric parameters of the liquid-propellant rocket engine nozzle are measured. The diameter of the inlet (“c”), critical (“*”) and outlet (“a”) sections of the nozzle (D c, D *, D a) is found.

3) The value of the parameters of the liquid-propellant rocket engine working fluid in the inlet, critical and outlet sections of the liquid-propellant rocket engine nozzle is calculated.

Based on the calculation results, a generalized graph of changes in temperature (T), pressure (p) and speed (W) of the working fluid along the nozzle path (L) of the liquid-propellant engine is constructed.

4) The traction parameters of the liquid-propellant rocket engine are determined at the design operating mode of the nozzle ().

4.2. Initial data for calculating the parameters of the RD-214 rocket engine

Gas pressure in the chamber (see option)

Temperature of gases in the chamber

Gas constant

Isoentropic exponent

Function

It is assumed that the processes in the chamber proceed without energy loss. In this case, the energy loss coefficients in the combustion chamber and nozzle are respectively equal

Nozzle operating mode is calculated (index " r»).

By measuring the following is determined:

Diameter of the critical section of the nozzle;

Diameter of the nozzle exit section.

4.3. Sequence of calculating the parameters of the rocket engine

A) The parameters in the nozzle exit section (“a”) are determined in the following sequence.

1) Nozzle exit area

2) Critical section area of ​​the nozzle

3) Geometric degree of gas expansion

CONTROL QUESTIONS

1. What is the significance of the RD-214 rocket engine?

2. List the main systems of the studied liquid propellant rocket engine.

3. What is the purpose of the rocket engine chamber, what parts does it consist of?

4. What is the purpose of the TNA, list its main units?

5. What is the purpose and composition of the steam and gas generation system of the liquid propellant rocket engine "RD-214"?

6. Describe the sequence of passage of the turbine working fluid.

7. List the main traction parameters of the rocket engine; what are their values ​​for the liquid propellant rocket engine "RD-214".

UDC 62-762

ANALYSIS OF THE DYNAMICS OF CHANGES IN THE RADIAL CLEARANCES OF PUMPS AND TURBINES OF THE LPRE HEAT

©2011 A. V. Ivanov Voronezh State Technical University

The article examines the factors influencing the change in gaps in the seals of high-speed turbo machines, and proposes approximating dependencies for analyzing the change in the gap during operation of the unit. It is shown that for high-speed units it is undesirable to use the assumption of constant clearance in all operating modes when calculating and analyzing the operation of turbomachines.

Seal, rotor, stator, gap, turbomachine, deformation.

When creating high-speed turbomachines, one of the key points is the choice of the gap between the rotor and stator seal elements. The choice of optimal values ​​and analysis of changes in gaps in the seals of the flow part play an important role when creating a sealing unit, since it is the gaps that largely determine the efficiency and performance of the structure. This task is especially relevant for turbopump units of liquid rocket engines, the structural elements of which are subject to significant force and temperature deformations (pressure drops across the seal elements up to 60 MPa, temperatures up to 1000 K, peripheral speeds of the rotor seal elements up to 600 m/s). The importance of the gap selection problem is due to the following:

Reducing the gap leads to a decrease in leakage through the seals, that is, an increase in the efficiency of the turbomachine;

Reducing the gap leads to an increase in the likelihood of frictional or impact contact between the rotor and stator elements of the seal, that is, damage to the sealing surfaces and, possibly, failure of the unit.

In turbopump units, stationary and self-aligning seals with guaranteed clearance are most widely used.

For non-contact seals, three types of gaps can be distinguished - installation, working and minimum guaranteed gaps. Assembly gaps are the gaps between the rotor and stator elements of the seal during assembly, defined as half the difference in diameters, based on the assumption of a concentric relative position of the rotor and stator. Working clearances - for-

gaps between the rotor and stator, taking into account force and temperature deformations, obtained from the condition of axisymmetric deformations, determining the flow through the seal. Minimum guaranteed clearances - clearances determined taking into account force and temperature deformations, as well as possible mutual installation and operational displacement of structural elements, which determine the performance of the seal.

In general, there are two types of reasons that cause a change in the gap between the rotor and stator parts of the seal:

Mounting displacements, that is, displacements of the axes of the sealing surfaces relative to the geometric axis, which are present in the assembled unit before its startup, are influenced by three groups of factors: the design layout of the unit, features of the technological process and actual errors in the manufacture of parts, the technological process of assembly and control of seal units;

Operational displacements caused by the operating conditions of the unit as part of the engine - temperature and force deformations, shaft bending due to hydraulic and gas forces, loads from imbalances, vibrations, etc.

Nominal values ​​of radial clearances in seals are assigned based on design experience and operating statistics of similar units or by calculation. Typically, a combination of these two methods is used. Typically, for each seal, the stress-strain state is calculated at the nominal operating mode. Calculations of the dynamics of changes in the thermal state of the structure during the process are also performed.

unit operation. These calculations are performed in specialized CAE systems using the finite element method in order to determine the nominal values ​​of deformations of the seal elements, and assign nominal mounting and working clearances. Calculation of installation and operational displacements is carried out according to the maximum, most unfavorable combinations of tolerances in terms of dimensions, shape and location of surfaces from the point of view of performance. Performing finite element calculations for each operating mode (start, stop, transition from mode to mode) is a complex, time-consuming and labor-intensive process. In this regard, it is advisable to calculate the dynamics of changes in the radial clearance using simplified dependencies. Such dependencies must satisfy the following requirements:

1) universality - must provide the ability to calculate the values ​​of the radial minimum guaranteed and working clearances for any blade machine: pump, turbine, compressor;

2) simplicity - should not require the use of additional calculations using CAE systems;

3) high accuracy - must take into account all data available when designing the rotor and stator seal elements on the deformations of the seal elements, tolerances of dimensions, shape and location of surfaces.

Let's look at rotor seals with guaranteed clearance. The radial working gap in the AKR seal is equal to the difference between the AKM installation gap and the sum of the magnitudes of force and thermal deformations of the AKR rotating and stationary sealing elements:

LYA^L^-L^. (1)

Local minimum clearance

A^^A^-A^-bYa.-v, (2)

where = 8R¡b + 8Rlab - local

reduction in the radius of the sealing surface of the body caused by deviations in its shape during manufacturing and assembly (8 7? fb) and shape deviations due to the influence of force and temperature loading

tions during work (8 R£a6);

s=£c6+sPa6 - displacement of the axis of the sealing surface of the rotor relative to the axis of the sealing surface of the stator during assembly (vsb) and due to force and temperature loading during operation (£work) -

The mounting gap in the seal is equal to the difference in the radii of the sealing surfaces of the stator Ry.c and rotor i?y p, measured during assembly:

The size of the installation gap is selected from the condition that the condition ARmin >0 is met in all operating modes. The total value of force and thermal deformations is determined as

ARM=5RCM+5Rvn-5RVM+

where 5 Ms [ is the deformation of the housing sealing element due to the pressure drop across the seal;

8 Rpс - deformation of the rotor sealing element due to centrifugal forces;

8 Rpд - deformation of the sealing

rotor element from pressure drop across the seal;

8 Rpt - thermal deformation of the rotor sealing element;

8 Rct - thermal deformation of the housing sealing element.

Thermal deformation 8 Rpt, 8 Rct

have a positive value if the temperature of the structure is higher than the temperature of the parts during assembly, and a negative value if the temperature of the structure is lower than the temperature of the parts during assembly.

The amount of displacement of the axes of the sealing surfaces

S - S + £ + £ + £ + £ + £ + £ ,

r R s s.i pr d k.t p "

s - mounting displacement of the axis

the surface of the rotor sealing element relative to the axis of its rotation, caused by gaps along the seating surfaces of parts, deviations in the relative position of the surfaces of parts during manufacturing, and clearances in bearings;

£рс - mounting displacement of the axes of the stator sealing elements during assembly

unit, caused by gaps in the fit of parts and deviations in the relative position of the surfaces of parts during their manufacture;

all - installation displacement of the axes of the sealing elements caused by deformations of the unit housings during the assembly of the unit and the engine;

Radius of precession of the rotor during operation;

£д - displacement of the sealing axes

elements during operation due to force and thermal deformations of the unit housings; VKT - displacement of the sealing axes

elements during operation, caused by deformations of the unit housings under the influence of connected pipelines and engine fasteners;

£п - displacement of the sealing axes

elements caused by rotor deflection under the influence of hydrodynamic forces in the cavities of the unit.

From equations (1), (2) it follows:

The given dependencies are valid for any type of non-contact seals.

As follows from dependencies (2), (3), (4), choosing a minimum but sufficient for safe operation value of the installation gap is a difficult task, since it is necessary to take into account a number of components of deformation and displacement of the axes of the sealing elements. This task is further complicated by the fact that the magnitudes and vector directions of deformations and displacements of the axes are probabilistic in nature.

In accordance with dependence (5), the minimum value of the working gap AKr in the seal is ensured at minimum values<5 и е. Таким образом, одним из направлений обеспечения минимального значения рабочего зазора является повышение точности изготовления деталей агрегата, повышение качества сборки агрегата и двигателя, увеличение жесткости ротора и корпусов агрегата. Более радикальным направлением является использование уплотнений с плавающими кольцами. Схема расчета зазоров в уплотнении с фиксированной гладкой стенкой приведена на рис. 1.

Rice. 1. Scheme for calculating gaps in a seal with a fixed smooth wall

In a seal with a floating ring, the displacement of the rotor axis relative to the housing axis is compensated by the radial displacement of the floating ring. In addition, due to the absence of a rigid connection between the ring and the body, the possibility of changing the shape of the sealing elements during assembly and operation is eliminated. During operation, the floating ring, due to the action of hydrodynamic forces in the sealing gap, which in all operating modes exceed the friction force at the end of the ring, self-aligns relative to the sealing surface of the rotor. In this case, the working gap in the seal is equal to the local minimum gap - A Ш = А11ты. A diagram for calculating gaps in a self-aligning seal with a floating ring is shown in Fig. 2, a.

TNA seals operate at high pressure drops, as a result of which an increased pressing force acts on the floating ring against the end of the housing, which does not allow it to self-align during precession of the axis of the rotor sealing surface. These seals are classified as semi-moving seals. In semi-moving seals, the ring self-aligns relative to the sealing surface of the rotor, compensating for axis displacements and deflections of the rotor, but at the same time, the mounting runout of the sealing surface of the rotor and its runout associated with the precession of the rotor during operation are not compensated. It should be noted that during installation, the semi-moving ring can be displaced relative to the rotor within the mounting gap and, as a result, contact between the ring and the rotor is possible. When starting (stopping), when hydrodynamic forces are less than the friction forces at the end of the ring, the semi-moving ring is aligned relative to the rotor due to collisions

between them. When operating in the mode, the semi-moving ring is aligned relative to the rotor due to hydrodynamic forces in the sealing gap, since they exceed the friction force at the end of the ring. During operation of the unit, the semi-moving ring does not monitor the rotor beats, but monitors the position of the rotor when switching from mode to mode. In a semi-moving seal, the working gap is determined by the relation

LA = LA +8. (6)

r.p tgp r pr V J

The working gap in the semi-moving seal (Fig. 2, b) is less than in the slot seal by the amount

5 Нр Н + £Р с + £с и + £д + т + £п. (7)

I > 3 w ^ £

Rice. 2. Scheme for calculating gaps in self-aligning seals: a - with a floating ring; b - with a semi-moving ring

This is the main advantage of a seal with a semi-moving ring compared to a gap seal, which ensures reduced leakage of the working medium. In gap seals, due to the fact that the magnitude of the axis displacement and rotor deflection are difficult to predict, with small installation gaps there is a possibility of the rotor jamming before it operates or the sealing surfaces wear out during operation. A seal with a semi-moving ring has higher reliability, since it does not have this disadvantage.

It should be noted that force and thermal deformations of the sealing elements and rotor deflection can be determined by calculation with a certain error. In addition, force deformations and rotor deflection change depending on the operating mode, and thermal deformations change over time as stationary temperature values ​​of the structure are reached. Therefore, it is necessary to strive to achieve minimum values ​​of deformation and deflection of the rotor. When LYAD = 0, the working gap in the gap seal is AY^ =LKm, and in the seal with a floating ring

The difference in thermal deformations of the sealing elements of the housing and rotor can be equal to zero at the same temperature values ​​and the same structural materials of the sealing elements, and also provided that the operating temperature of the structure differs little from the temperature at which the assembly is carried out.

Force deformations in the THA seals of engines without afterburning were small. The main contribution was made by temperature deformations, since aluminum alloys were often used for pump impellers. In engines with afterburning, force deformations of seal elements have increased significantly, especially in oxygen-hydrogen rocket engines, in which increased deformations are due to higher structural tension. Currently, when creating reusable reusable liquid propellant rocket engines, it is important to maintain the stability of deformations and gaps from launch to launch of the pump.

Approximating dependencies for determining the components of deformation of the rotor and stator elements of the seal, assuming that the pressure drops across the seal elements depend on the rotor rotation speed, can be presented as follows:

5ys.d(t) =<5ДСН°М (п(т)/пном)2 - силовые деформации статорного элемента уплотнения в произвольный момент времени т;

Force deformations of the stator seal element at the nominal operating mode;

u(t) - rotor rotation frequency in pro-

arbitrary moment in time t; other - rated rotor speed;

5Dr.d(t) = °m (p(t)/pnom)2 - force deformations of the rotor element of the seal due to the action of pressure drop at an arbitrary moment of time t;

<5/?р °м - силовые деформации роторного элемента уплотнения от действия перепада давления на номинальном режиме работы;

gYar.ts(t) = 5yr°m (p(t)/pnom)2 - force deformations of the rotor element of the seal from the action of centrifugal forces at an arbitrary moment of time t;

5yr °m - force deformations of the rotor element of the seal due to the action of centrifugal forces at the nominal operating mode;

temperature deformations of the stator seal element at an arbitrary time t;

(t) = ^ (t) - ^ sb - change in temperature of the stator seal element;

tc(t) is the temperature of the stator seal element at an arbitrary time t; tccb is the temperature of the stator element during seal assembly;

ac(t)) is the temperature coefficient of linear expansion of the material of the stator seal element depending on its temperature at an arbitrary time t, obtained from the approximating dependence;

"Mt) = *p(tK("P(T))LR "them"

temperature deformations of the rotor seal element at an arbitrary time t;

¿Ts,(t) = *p(t)-*rsb - change in temperature of the rotor element of the seal;

/p (t) - temperature of the rotor element of the seal at an arbitrary moment; time-temperature of the rotor element -

when assembling the seal;

ar (t)) - temperature coefficient of linear expansion of the material of the rotor element of the seal depending on its temperature at an arbitrary moment

time t, obtained from the approximating dependence.

Generalized dependence for determining the working gap:

chaN/ chaN/

r.ts (Shch2 - (gyas.s(t) -<5Др.,(т)) .

For minimum guaranteed clearance:

MnY = DDM - (<5Д£б + "

\ 4 /"-nom/ / ^"^nom"

^* "■nom" ^

In the given dependencies, the components of deformations, deviations in shape and location of the rotor and stator elements of the seal have a positive value if they lead to a decrease in the installation gap, negative if they lead to an increase in the radial gap.

As an example, we present the results of calculating the dynamics of changes in the working and local guaranteed clearances during testing of a high-speed turbomachine (Fig. 3). All parameters on the graph, except time, are normalized, that is, related to the nominal value of the corresponding parameter.

Rice. 3. Changing parameters during testing:

1 - minimum guaranteed clearance; 2 ^working gap; 3 - temperature of seal elements; 4 - rotor speed

When calculating the dynamics of changes in radial clearances, the following assumptions were made: the temperature of the rotor and stator seal elements is the same; displacement of the axes of the sealing surfaces and local reduction of the radius

The sealing surface of the housing is constant, regardless of the operating mode of the unit; the force and temperature deformations of the sealing elements are axisymmetric in nature.

It can be seen that the minimum guaranteed gap in some operating modes is up to 15% of the installation gap, the working gap is up to 30% of the installation gap.

During operation of the unit, the radial working gap in the seal can change by 2-4 times compared to the installation one, and the minimum guaranteed gap can change by 2-10 times. Thus, the commonly used methods of application in test analysis, TNA calculations under the assumption of constant radial clearance are not always acceptable.

Bibliography

1. Dmitrenko, A.I. Seal analysis

flow part of pumps and turbines of TNA LRE [Text] / A.I. Dmitrenko, A.B. Ivanov // Scientific and technical anniversary collection. Design Bureau of Chemical Automation. - Voronezh: IPF "Voronezh". - 2001. - P. 364-370.

One of the most complex mechanical engineering structures is the gas turbine.

The development of gas turbines is determined, first of all, by the development of aviation gas turbine engines for military purposes. In this case, the main thing is to increase specific thrust and reduce specific gravity. Economic and resource problems for such engines are secondary.

One of the most loaded parts that limit the time between overhauls is the uncooled turbine blades, made of wrought nickel alloy EI893. Due to limitations in long-term strength, blades made from this alloy have a service life of 48,000 hours. Currently, there is quite high competition in the production of turbine blades, so the issues of reducing cost and increasing the service life of blades are very relevant.

This graduation project examines a relatively new technology for the domestic industry for the production of long-length uncooled turbine blades (more than 200 mm). As a blade blank, a casting from the TsNK-7P material is used without allowance for mechanical processing of the blade, subjected to hot isostatic pressing. To reduce the labor intensity of blade manufacturing, deep-feed grinding of the lock is used, and to increase fatigue resistance, the blade lock after grinding is subjected to hydro-shot peening.

This graduation project examines the production technology of a turbine blade. Since this technical process is universal for blades of various sizes, it can be used both for the manufacture of low-pressure turbine blades of a gas turbine engine (or gas turbine engine) and a turbocharger turbine of a liquid propellant rocket engine. This work examines the blade for the fuel pump of the RD-180 liquid-propellant rocket engine. However, due to the versatility of the blade material and technological process, we also pay increased attention to the service life of the product. The process of creep-feed grinding for parts made of heat-resistant alloys, such as a turbine blade, is examined in detail, and the production technology and properties of diamond rollers used in creep-feed grinding for dressing grinding wheels are described. The project is designed for the accuracy and clamping force of the “pike mouth” device, which is widely used in creep-feed grinding operations during the blade production process. The research part examines the process of increasing fatigue strength by blasting the blade lock with shot in a liquid medium (hydro-shot peening), and describes methods for determining residual stresses and conducting fatigue tests of the blade. The work also describes the CATIA design automation system and the creation of a part model and design documentation in this system. In terms of labor protection, measures have been developed to improve production safety and environmental protection. The efficiency of implementing this blade production process in relation to the previous one was also calculated.

Brief description of TNA RD-180

*Description is given without a gas generator.

The turbopump unit is made according to a single-shaft design and consists of an axial single-stage jet turbine, a single-stage centrifugal screw oxidizer pump and a two-stage centrifugal screw fuel pump (the second stage is used to supply part of the fuel to gas generators).

On the main shaft with the turbine there is an oxidizer pump, coaxially with which two stages of the fuel pump are located on another shaft. The shafts of the oxidizer and fuel pumps are connected by a gear spring to unload the shaft from thermal deformations that arise as a result of the large temperature difference between the working bodies of the pumps, as well as to prevent freezing of the fuel.

To protect the angular contact shaft bearings from excessive loads, effective auto-unloading devices are used.

The turbine is an axial single-stage jet turbine. To prevent fire due to breakdowns of structural elements or friction of rotating parts against stationary ones (due to the selection of gaps from deformations or work hardening on the mating surfaces from vibration), the gap between the blades of the nozzle apparatus and the rotor is made relatively large, and the edges of the blades are made relatively thick.

To prevent fire and destruction of turbine gas path parts, nickel alloys are used in the design, including heat-resistant ones for hot gas lines. The turbine stator and exhaust tract are forcibly cooled with cold oxygen. In areas of small radial or end clearances, various types of heat-protective coatings are used (nickel for the rotor and stator blades, metal-ceramic for the rotor), as well as silver or bronze elements, which prevent fire even if there is a possible contact with the rotating and stationary parts of the turbopump unit.

To reduce the size and mass of foreign particles that could lead to a fire in the gas path of the turbine, a filter with a cell of 0.16 * 0.16 mm is installed at the engine inlet.

Oxidizer pump. The high pressure of liquid oxygen and, as a result, an increased risk of fire determined the design features of the oxidizer pump.

Thus, instead of floating sealing rings on the impeller collars (usually used on less powerful pumps), fixed gap seals with a silver lining are used, since the process of “floating” of the rings is accompanied by friction at the points of contact of the impeller with the housing and can lead to fire of the pump.

The screw, impeller and torus outlet require particularly careful profiling, and the rotor as a whole requires special measures to ensure dynamic balance during operation. Otherwise, due to large pulsations and vibrations, destruction of pipelines and fires at joints occur due to the mutual movement of parts, friction and hardening.

To prevent fire due to breakdowns of structural elements (screw, impeller and guide vanes) under conditions of dynamic loading with subsequent fire due to rubbing of debris, means were used such as increasing structural perfection and strength due to geometry, materials and cleanliness of mining, and also the introduction of new technologies: isostatic pressing of cast billets, the use of granular technology and other types.

The oxidizer booster pump consists of a high-pressure screw and a two-stage gas turbine, which is driven by oxidizing gas taken after the main turbine with its subsequent bypass to the inlet of the main pump.

The fuel booster pump consists of a high-pressure auger and a single-stage hydraulic turbine operating on kerosene taken after the main pump. Structurally, the fuel booster pump is similar to the oxidizer booster pump with the following differences:

· a single-stage hydraulic turbine operates on fuel taken from the outlet of the fuel pump of the main HPU;

· high pressure fuel is removed to relieve the auger from axial actions from the inlet manifold of the BNAG hydraulic turbine.

Table 1: TTX TNA

Parameter	Meaning
Oxidizer
Pump outlet pressure
Component flow through the pump
Pump efficiency
Shaft power
Shaft rotation speed
Turbine power
Turbine inlet pressure
Number of steps
Turbine pressure reduction ratio
Turbine inlet temperature
Turbine efficiency

The invention relates to rocket technology, specifically to liquid rocket engines running on a cryogenic oxidizer and hydrocarbon fuel. A turbopump unit (TPA) of a liquid rocket engine contains an oxidizer pump impeller, a fuel pump impeller and a turbine impeller installed on the shaft of the turbopump unit rotor parts, an additional fuel pump impeller located in the body of the turbopump unit with a shaft and an additional fuel pump impeller, according to the invention between the impeller The turbine and oxidizer pump impeller are equipped with a magnetic coupling and a multiplier. A magnetic coupling and a multiplier can be installed between the oxidizer pump and the fuel pump. A magnetic coupling and a multiplier can be installed between the fuel pump and the additional fuel pump. The invention improves the reliability of the TNA. 2 salary f-ly, 3 ill.

The invention relates to rocket technology, specifically to liquid propellant rocket engines running on a cryogenic oxidizer and hydrocarbon fuel.

A liquid rocket engine is known according to RF patent for invention No. 2095607, intended for use as part of space upper stages, launch vehicle stages and as a propulsion engine of spacecraft, includes a combustion chamber with a regenerative cooling path, a turbopump unit - TNA. The TNA contains pumps for supplying components - fuel and oxidizer with a turbine on the same shaft, into which a capacitor is inserted. The condenser outlet is connected via the refrigerant line to the entrance to the combustion chamber and to the entrance to the regenerative cooling path of the combustion chamber. The outlet from the condenser is connected via a coolant line to the inlet of the pump of one of the components. The pump outlet of the same component is connected to the condenser inlet via a refrigerant line. The second input of the condenser is connected to the turbine output. The output of the pump of another component is in communication with the inlet of the combustion chamber.

The disadvantage of the TNA engine is the deterioration of the cavitation properties of the pump when condensate is bypassed. This property of the pump inevitably leads to a decrease in the consumption of one of the fuel components through the pump, a drop in rocket thrust several times and disruption of the rocket’s flight program or to a disaster.

The method of operation of a liquid-propellant rocket engine and a liquid-propellant rocket engine are known according to RF patent for invention No. 2187684. The method of operation of a liquid-propellant rocket engine consists of supplying fuel components to the combustion chamber of the engine, gasifying one of the components in the cooling path of the combustion chamber, supplying it to the turbine of the turbopump unit, and then discharging it into the nozzle head of the combustion chamber. Part of the flow rate of one of the fuel components is sent to the combustion chamber, and the remaining part is gasified and sent to the turbines of turbopump units. The gaseous component exhausted from the turbines is mixed with a liquid component entering the engine at a pressure exceeding the saturated vapor pressure of the resulting mixture. The liquid rocket engine contains a combustion chamber with a regenerative cooling path, pumps for supplying fuel components and a turbine. Pumps and turbines are arranged in two pumps: main and booster. The engine contains a pump of the booster turbopump unit and a mixer installed in series in front of the supply pump of one of the fuel components of the main turbopump unit. The pump output of the main turbopump unit is connected both to the injector head of the combustion chamber and to the regenerative cooling path of the combustion chamber. The regenerative cooling path, in turn, is connected to the turbines of the main and booster turbopump units, the outputs of which are connected to the mixer.

The disadvantage of this scheme is that the thermal energy removed during cooling of the combustion chamber may not be enough to drive the turbopump unit of a very high-power engine.

The liquid propellant rocket engine is known under RF patent for invention No. 2190114, IPC 7 F02K 9/48, publ. 09.27.2002 This liquid-propellant engine includes a combustion chamber with a regenerative cooling path, a THA turbopump unit with oxidizer and fuel pumps, the output lines of which are connected to the head of the combustion chamber, the main turbine and the main turbine drive circuit. The drive circuit of the main turbine includes a fuel pump connected in series with each other and a regenerative cooling path of the combustion chamber connected to the inlet of the main turbine. The outlet from the turbocharger turbine is connected to the input of the second stage of the fuel pump.

This engine has a significant drawback. Bypassing the fuel heated in the regenerative cooling path of the combustion chamber to the inlet of the second stage of the fuel pump will lead to cavitation and the consequences indicated above. Most liquid rocket engines use fuel components such that the oxidizer consumption is almost always greater than the fuel consumption. Consequently, for powerful liquid-propellant rocket engines with high thrust and high pressure in the combustion chamber, this scheme is unacceptable, because Fuel consumption will not be enough to cool the combustion chamber and drive the main turbine.

In addition, the launch system for the liquid-propellant engine, the system for igniting fuel components and the system for turning off the liquid-propellant engine and cleaning it from fuel residues in the regenerative cooling path of the combustion chamber have not been developed.

A liquid rocket engine and a method for starting it are known according to the Russian Federation patent for invention No. 2232915, publ. 09/10/2003 (prototype), which contains a combustion chamber, a turbopump unit, a gas generator, a starting system, means for igniting fuel components and fuel lines. The output of the oxidizer pump is connected to the inlet of the gas generator. The output of the first stage of the fuel pump is connected to the regenerative cooling channels of the chamber and to the mixing head. The output of the second stage of the fuel pump (additional fuel pump) is connected to an electrically driven flow regulator. The other input of the regulator is connected to the starting tank with standard fuel. The output from the regulator is connected to the gas generator. The outlet from the gas generator is connected to the inlet of the turbine of the turbopump unit, the outlet of which is connected to the mixing head. The flow regulator is equipped with a preliminary stage hydraulic drive, which is connected to the starting tank with standard fuel through a cavitating jet and a hydraulic relay. The hydraulic relay is connected to the second stage of the fuel pump. The throttle installed at the outlet of the first stage of the fuel pump is made in conjunction with the controlled valve of the preliminary stage.

The disadvantage of this scheme is a fire or explosion of the pump and rocket at launch or in flight due to the low reliability of the seal between the turbine and the oxidizer pump, between the oxidizer and fuel pump, as well as between the fuel pump and the additional fuel pump due to the action of a large pressure drop on them: 300...400 kgf/cm 2 for modern liquid propellant engines. For example, when hydrogen and oxygen are used as rocket fuel components, the slightest leaks of these components lead to the formation of an “explosive mixture” and almost always to a rocket explosion.

Objectives of creating the invention: preventing the explosion of a fuel pump or rocket at launch or in flight.

The solution to this problem is achieved due to the fact that the turbopump unit of a liquid rocket engine, containing parts of the rotor of the turbopump unit installed on the shaft: an oxidizer pump impeller, a fuel pump impeller and a turbine impeller, located in the housing of the turbopump unit, an additional fuel pump impeller with a shaft and an additional impeller fuel pump, differs in that a magnetic coupling is installed between the turbine impeller and the oxidizer pump impeller. A magnetic coupling can also be installed between the oxidizer pump and the fuel pump. A magnetic coupling can also be installed between the fuel pump and the additional fuel pump.

Conducted patent studies showed that the proposed technical solution has novelty, inventive step and industrial applicability. The novelty is confirmed by patent research, the inventive step is the achievement of a new effect - absolute tightness of connections between the turbine and pumps, as well as between pumps and prevention of explosion of the pump and rocket at launch or in flight.

Industrial applicability is due to the fact that all elements included in the pump assembly are known from the prior art and are widely used in engine building.

The essence of the invention is illustrated in figures 1...3, where:

Figure 1 shows a diagram of the first version of the TNA,

Figure 2 shows a diagram of the second version of the TNA,

Figure 3 shows a diagram of the third version of the TNA.

The turbopump unit of the TNA liquid rocket engine 1 (Fig. 1) contains a fuel pump shaft 2, an oxidizer pump shaft 3. The oxidizer pump impeller 4 is installed on the oxidizer pump shaft 3, the fuel pump impeller 5 is installed on the fuel pump shaft 2. The turbine impeller 6 is installed at the top of the TNA. All parts of the THA rotor are located inside the THA housing 7. An additional fuel pump 8, which has an additional fuel pump impeller 9 and an additional fuel pump shaft 10, is made coaxially with the THA 1 and is installed on the side opposite the turbine impeller 6. The impeller of the additional fuel pump 9 is installed in the housing of the additional fuel pump 11, the cavity of which “B” is sealed relative to the cavity of the fuel pump “A”. Between the impeller of the fuel pump 5 and the additional fuel pump 8 in the THA housing 7, a magnetic coupling 12 and a multiplier 13 are installed. Magnetic coupling 12 and all other magnetic couplings (if they are used in the design) consist of a driving disk of a magnetic clutch, a driven disk of a magnetic clutch, and between The magnetic coupling discs form a partition made of non-magnetic material, for example, non-magnetic steel (not shown in Figs. 1...3). The turbine impeller is mounted on the turbine shaft 14.

The gas generator 15 is installed coaxially with the THA 1 above the nozzle apparatus of the turbine 16. The gas generator 15 contains a gas generator head 17, inside of which there is an outer plate 18 and an inner plate 19 with a cavity “B” above them and a cavity “G” between them. Inside the head of the gas generator 17, oxidizer nozzles 20 and fuel nozzles 21 are installed. The oxidizer nozzles 20 communicate with the cavity “B” with the internal cavity of the gas generator “D”, and the fuel nozzles 21 communicate with the cavity “G” with the internal cavity of the gas generator “D”. On the outer surface of the gas generator 15 there is a fuel manifold 22, to which a high-pressure fuel line 23 from an additional fuel pump 8 is connected. In the line of the high-pressure pipeline 23 there is a high-pressure valve 24 and a flow regulator 25 with a drive for the flow regulator 26. The outlet from the impeller of the fuel pump 5 connected by pipeline 27 to the entrance to the additional fuel pump 8 and to the combustion chamber (the combustion chamber is not shown in Fig. 1).

The outlet from the impeller of the oxidizer pump 4 is connected by the oxidizer pipeline 28 through the oxidizer valve 29 to the cavity “B” of the gas generator 15. One or more ignition devices 30 are installed on the gas generator 15. The control unit 31 is electrically connected to the ignition devices 30, the high pressure valve 24, the valve oxidizer 29 and flow regulator drive 26.

When the liquid-propellant rocket engine is started, electrical signals are sent from the control unit 31 to valves 24 and 29 and ignition devices 30. The oxidizer and fuel from the impellers of pumps 4, 5 and 8 flow by gravity into the gas generator 15, where it is ignited, the combustion products spin the turbine impeller 6 , mounted on shaft 14.

In the first option (Fig. 1), the shaft of the oxidizer pump 3 is spun through the magnetic coupling 12 and the multiplier 13. The pressure at the outlet of the impellers of pumps 4 and 5 increases. Part of the fuel (about 10%) enters the additional fuel pump 8, where its pressure increases significantly. The additional fuel pump 8 is driven into rotation and has the same rotation speed as the impeller of the oxidizer pump 4 and the impeller of the fuel pump 5 (Fig. 1).

According to the second option (Fig. 2), the torque from the shaft of the oxidizer pump 3 is transmitted to the shaft of the fuel pump 2 through a magnetic coupling 12 and a multiplier 13. In this case, the impeller of the fuel pump 5 will have a higher speed than the impeller of the oxidizer pump 4. The shaft of the additional pump fuel pump 10 is connected to the shaft of fuel pump 2 directly.

According to the third option (Fig. 3), in addition to two magnetic couplings with multipliers, a third magnetic coupling with a multiplier is used in the design of the TNA. As a result, due to the absence of a seal on the shaft of the additional fuel pump 10, its reliability increases. At a pressure at the inlet to the impeller of the fuel pump 4 of the order of P 1 =4...5 kgf/cm 2 , at the outlet from the impeller of the fuel pump 4 P 2 =300 kgf/cm 2 and at a pressure at the outlet of the additional fuel pump 8 approximately P 3 =900 kgf/cm 2 the pressure difference of approximately 600 kgf/cm2 that arises between them is perceived by a partition made of non-magnetic material 14. Pressure at the inlet to the oxidizer pump P 4 =4...5 kgf/cm2, at the outlet of the oxidizer pump P 5 =400 kgf/cm 2 , at the entrance to the combustion chamber P 6 =300 kgf/cm 2 . The presence of magnetic couplings between the pumps and the oxidizer pump and the turbine ensures complete tightness of all modules relative to each other, the presence of multipliers ensures coordination of rotation speeds of the turbine and pumps and at the same time modularity of the design.

As a result, there was a real opportunity to design all the main components of the turbocharger: turbine and pump for optimal parameters, including rotation speeds, and to coordinate the rotation speeds through the use of one multiplier between the turbine and pumps or several multipliers, and this made it possible to minimize the weight of the pump, which is of decisive importance in rocket technology.

The use of the invention allowed:

1. To prevent the explosion of the pump and rocket at launch or in flight due to contact of the oxidizer and fuel in the cavity between the pumps or the penetration of combustion products from the turbine into one of the fuel components, if oxygen and hydrogen or other aggressive components are used as rocket fuel components.

2. Ensure modularity of the TNA design.