Jet-propulsion plants (F02K)

F02K              Jet-propulsion plants (arrangement or mounting of jet-propulsion plants in land vehicles or vehicles in general b60k; arrangement or mounting of jet-propulsion plants in waterborne vessels b63h; controlling aircraft attitude, flight direction, or altitude by jet reaction b64c; arrangement or mounting of jet-propulsion plants in aircraft b64d; plants characterised by the power of the working fluid being divided between jet propulsion and another form of propulsion, e.g. propeller, f02b, f02c; features of jet-propulsion plants common to gas-turbine plants, air intakes or fuel supply control of air-breathing jet-propulsion plants f02c)(2873)

Solid-propellant rocket engine (versions) // 2642764
FIELD: engine devices and pumps.SUBSTANCE: solid-propellant rocket engine in the first version comprises a body with a solid fuel arranged therein, a nozzle unit installed on the rear bottom of body, an igniter includes a solid fuel igniter mounted in the front and/or rear bottom of the body. The igniter is mounted in the front bottom of the body along the axis of a central channel. The igniter includes a laser connected by a cable to a power supply source and directed by the laser beam focus of given pulse shape to a layer of flammable substance which is applied to the end of the igniter, and the igniter is placed in a perforated shell and installed in a casing with radial clearance. In the igniter installed in the front bottom of the body, the casing with its one end is airtightly connected to the laser, and on its other end there is a flare, on which a deflector of radially directed ejection of flame from the igniter is installed to the surface of through solid fuel central channel. In the igniter installed in the rear bottom of the body, the casing with its one end is airtightly connected to the laser, and the other end of the casing there is a branch pipe which to which the igniter is mounted eccentrically to the central channel in the rear bottom of the body with the possibility of concentrated ejection of flame from the igniter through a prechamber to the solid fuel end. In the second version the solid-propellant rocket engine contains armouring coatings on the ends of solid fuel, and the igniter is mounted in the front bottom of the body and is a laser. The laser drives ignition of solid fuel of directed laser beam radiation focus of specified shape of pulse through the prechamber to solid fuel end, for this purpose in the armouring coating of the solid fuel end there is a local zone of ignition with the possibility for combustion transition into the solid fuel central channel. In the third version of the rocket engine the laser is mounted eccentrically in the rear bottom of the body eccentrically to the nozzle, driving ignition of solid fuel by directed laser beam radiation focus of specified shape of pulse through the prechamber into one of the cells of pyramidal shape made on the end of solid fuel, with subsequent transition of combustion over entire honeycomb pyramidal-cellular surface of solid fuel end.EFFECT: enhancement of reliability and reduction of ignition time of solid fuel.3 cl, 8 dwg

Blisk of cooled fuel feed pylons // 2642718
FIELD: engines and pumps.SUBSTANCE: each pylon has three channels plugged from side, open ends of two of these channels are closed by plugs, channels are connected by holes, and in central channel there is a damper with several holes.EFFECT: elimination of pylons burnout at high thermal loads, enhanced reliability of blisk of fuel supply, expanded performance range of fuel consumption at practically constant pressure drop on nozzles for improvement of combustion efficiency of fuel-air mixture.2 cl, 4 dwg

Afterburner combustion chamber of turbojet engine // 2642712
FIELD: engines and pumps.SUBSTANCE: afterburner combustion chamber of turbojet engine contains chassis, connected to the turbine, nozzle, fuel, or fuel-air collectors, injectors are connected with the jets. The spray nozzles are equipped with micro-swirlers. Each of the micro-swirlers is a cone-shaped nozzle casing with roundings of a small radius R at the base of the cone located near the atomizer. Nozzles with a micro-swirl are connected behind the manifolds by the flow of air so that the injected fuel or fuel-air mixture coincides in the direction of travel with the flow of air and combustion products behind the turbine in the afterburner combustion chamber. At the base of the cone of the nozzle body is a cylindrical spray nozzle.EFFECT: invention allows to reduce emissions of harmful substances into the atmosphere and decrease the length of the afterburner combustion.2 cl, 3 dwg

Fuel supply circuit and cooling method // 2642711
FIELD: aviation.SUBSTANCE: invention relates to a power supply circuit (6) for supplying a rocket motor (2) with at least the first liquid fuel. The mentioned power supply circuit comprises at least one buffer tank (20) for the first liquid fuel and the first heat exchanger (18), which is embedded in the mentioned buffer tank (20) and suitable for connecting to the cooling circuit (17) for cooling at least one power source to cool the mentioned heat source by transferring the heat to the first fuel.EFFECT: improved cooling of aircraft heat sources.11 cl, 3 dwg

System and method for feeding fuel to rocket engine // 2641802
FIELD: engines and pumps.SUBSTANCE: invention relates to a system for feeding rocket fuel to a rocket engine (2) comprising a first tank (3), a second tank (4), a first power supply system (6) connected to the first tank (3), and a second power supply system (7) connected to the second tank (4). For cooling the rocket fuel contained in the second tank (4), the first power supply system (6) includes a branch (12) passing through a first heat exchanger (14) built in the second tank (4). The invention also relates to a method of feeding rocket fuel to the rocket engine (2).EFFECT: maintaining pressure in the tanks above the minimum limit.14 cl, 9 dwg

ethod and device for rocket engine power supply // 2641791
FIELD: engines and pumps.SUBSTANCE: invention relates to a device for feeding the chambers of rocket engines (100) with the first and second propellant components. The first power supply circuit (16) of the thrust-generating chamber (10) includes a turbo pump (22) having at least one pump (22a) for pumping the first propellant component from a first tank (12) and a turbine (22b) mechanically connected to said pump (22a). The first power circuit connects the pump output to turbine input through a heat exchanger (24) configured for heating the first rocket fuel component by heat created by thrust-generating chamber in order to operate the turbine. In accordance with the present invention, the second power supply circuit (18) is arranged to supply the second propellant component creating the thrust chamber by the second propellant component from the second tank (14) which is configured to maintain high pressure. The present invention also provides a method for feeding the thrust-generating chamber of the rocket engine chamber with the first and second propellant components.EFFECT: increased pressure in tanks containing rocket fuel components.13 cl, 2 dwg

Low-thrust rocket engine on gaseous hydrogen and oxygen with jet-type injectors in air cross-flow // 2641785
FIELD: engines and pumps.SUBSTANCE: low-thrust rocket engine on hydrogen gas and oxygen comprises a fuel ignition spark plug, a mixing head providing mixing of fuel and internal cooling of the combustion chamber wall, combustion chambers and nozzles, in the air cross-flow of oxygen there are jet nozzles of jet type provided in the mixing head of the engine, the total flow vectors of which are directed in a plane perpendicular to the axis of the engine towards each other.EFFECT: improved combustion of gaseous hydrogen and oxygen.4 cl, 1 dwg

Rotary axisymmetric nozzle of turbojet engine // 2641425
FIELD: engines and pumps.SUBSTANCE: nozzle comprises a fixed body with a spherical hollow tip and a rotary device mounted for rotating about the axis that is transverse to the longitudinal axis of the engine. The fixed body with the spherical tip and rotary device are provided with screens on the side of their inner surfaces forming a cooling channel. Inside the spherical hollow tip there are Z-shaped frames rigidly attached to the inner and outer surfaces of the tip. There are holes for passing cooling air configured on free surfaces of the frames and on the rear flange of the spherical hollow tip, and on the inner and outer surfaces of the spherical tip, as well as between the frames on its inner surface.EFFECT: invention makes it possible to increase reliability of nozzle operation by providing more uniform cooling of the nozzle body spherical hollow tip.3 dwg
ethod of modelling of process of gasification of liquid rocket fuel in tank of carrier-rocket and device for its implementation // 2641424
FIELD: aviation.SUBSTANCE: method for modelling the process of gasification of the liquid component of rocket fuel in the tank of the carrier-rocket, based on the supply of heat to the experimental model set (EMS), carrying out measurements of temperature, pressure at various points of EMS, discharge of vapour-gas mixture (VGM) through the drain line (DL), while the charging of the boost gas and the conductive supply of heat to the EMS are carried out, the amount of which is determined from the condition of equality of the partial pressures of the pressurant gas and liquid vapour in the EM and the fuel tank, and the total pressure corresponds to the beginning of the discharge of the VGM in the DL, the diameter of the DL is determined from the condition for resetting the preset excess pressure for the same time as in the real tank. Actuation pressure of a draining valve is selected first from a predetermined interval, the lower limit of which is the minimum boost pressure in the tank, and the upper one is the maximum pressure at which the strength of the design of the EMS remains, determine the range of parameters of the gasification process, under which condensate appears on the inner surface of the DL and crystallization, perform additional heat supply to the DL to prevent its freezing. It is considered the device for the realization of the method, which includes the EMS in the form of a model tank containing a tray for liquid, a temperature sensor, a pressure sensor, an inlet nozzle, DL, a drain valve, and a gas analyzer, while heating elements for liquid and DL are introduced in the EMS, an equipment for recording condensate and its crystallization is installed in the DL, and EMS and DL are made of a material similar to the material of the fuel tank of the carrier rocket under investigation.EFFECT: detection of the conditions of condensate in the drainage line with subsequent crystallization during refuelling of the launch vehicle with cryogenic components of fuel or parking in the primed state at the start with the thermal loading of the fuel tank from the environment.2 cl, 1 dwg

Output device aircraft engine and group of aircraft engines of power unit (versions) // 2641341
FIELD: engines and pumps.SUBSTANCE: aircraft engine output device contains an outlet nozzle (1), a main engine outlet nozzle (2), at least one rotary damper (3), a control gear drive and a noise silencer (4). The device is provided with additional outlet nozzle (5) and a distributing branch pipe (6) coaxially connected to outlet branch pipe (1) of the engine and divided in the zone of rotary damper installation into two branch pipes connected to each other from the side of outlet branch pipe, each of which is provided with outlet nozzle, the main one (2) is in the form of the main outlet nozzle of the engine, and the other one is additional (5). The noise silencer (4) is located in one of the branches before the additional outlet nozzle (5). Actuator of the control mechanism is connected to the damper (3). The rotary damper (3) is installed for alternate closure of each of the branches.EFFECT: invention improves efficiency of noise suppression during take-off and landing, reduces pressure losses of the outlet jet and fuel consumption during cruising operation.12 cl, 6 dwg

Low-thrust liquid-propellant rocket engine // 2641323
FIELD: engines and pumps.SUBSTANCE: liquid rocket engine of low thrust (LRELT) contains a chamber 1, a mixing head with an inner bottom 2, an axial centrifugal nozzle 3, a peripheral belt of the jet nozzles 4, the annular conical deflector 5 between them, wherein the centrifugal nozzle slice 6 is recessed from the outlet edge 7 of the deflector surface forming surface toward the peripheral belt of the jet nozzles 4, a cavity of the combustion chamber 8 above the outer surface 9 of the deflector and a cavity 10 under the inner surface 11 of the deflector and an inner bottom of the mixing head are communicated by channels 12, which are offset from the nozzle holes by half a step (α/2). With this design, the nozzles 4 do not experience perturbations when the engine is running.EFFECT: increase of power characteristics of the engine, reliable cooling of the combustion chamber and the mixing head.5 dwg

Afterburner of the two-convention turboreactive engine // 2641191
FIELD: engines and pumps.SUBSTANCE: afterburner combustion chamber of turbojet engine contains chassis, connected to the turbine, nozzle, fuel, or fuel-air collectors, injectors are connected with the jets. Before the mixer in the second circuit, an additional collector with sprayers is installed. Sprays of the additional collector are located in the middle of the pockets of the mixer, which ensures the creation of a fuel-air mixture with an air excess factor equal to the same coefficient in the hot gas behind the turbine.EFFECT: invention makes it possible to obtain a gas flow with the required excess air factor at any point in the section ahead of the front-mounted device of the afterburner, which simplifies the problem of uniform distribution of afterburner fuel in cross section.2 dwg

ethod of forming reactive forces of motion from air-dynamic part of jet and device for realization of this method // 2641178
FIELD: engines and pumps.SUBSTANCE: method is carried out by means of transverse extrusion of the required air mass for a combustion chamber from accelerating air column by jet from the channel formed by a blade rotor in a cylindrical body with subsequent replacement of accelerated air mass onto newly formed air mass with subsequent replacement of already cut part of reactive jet on column of air, which is carried out in the blind position of the channel by the lateral filling of the released space with air from the surrounding space through a screw window of a body in the period of its movement for getting into the jet from the other side. Lateral extrusion of the air mass from its volume is possible at the formation of a considerable internal pressure from the head of the jet to the counter-acting force of the resistance to acceleration of the free air mass in the channel, which requires the absence of the initial movement in the associated direction and its excess in the volume, which are allocated for the traction force increase.EFFECT: increased efficiency of the air-jet engine by the steady supply of air mass under any operating conditions with a concomitant increase in traction force.2 cl, 6 dwg

Liquid-propellant rocket engine chamber with controlled nozzle // 2640903
FIELD: engine devices and pumps.SUBSTANCE: chamber comprises a cooled part of nozzle and an uncooled header of carbon-carbon composite material, steering assemblies and a frame, according to the invention, in the uncooled header, there are niches in which several sections of a detachable earth nozzle are disposed, having rotation shafts tangential in the area of the joint of uncooled header cooled part of the nozzle, mounted in brackets fixed to the cooled part of nozzle and connected by steering units to the engine frame.EFFECT: increasing the efficiency and reliability of work along the entire rocket flight path.3 dwg

Combustion chamber of liquid-propellant engine with afterburning of generator gas // 2640893
FIELD: engine devices and pumps.SUBSTANCE: combustion chamber comprising a gas conduit, a mixing head with mixing elements, a chamber body and fuel supply lines is described, according to the invention, in the region of chamber minimum section, a toroidal gas conduit is made, the cavity of which, by means of a finned path made on the outer wall of the chamber body and head outer bottom, is connected to head mixing elements.EFFECT: increasing the reliability of combustion chamber and reducing the linear size of combustion chamber.4 dwg

Reduction gear for high-speed and small-size fan drive turbine // 2639821
FIELD: engines and pumps.SUBSTANCE: gas turbine engine comprises a flexible support for a fan drive gear set. The first turbine section has a first output area and is capable to rotate at first speed. The second turbine section has a second output area and is capable to rotate at second speed which exceeds the first speed. The first characterizing parameter is the product of the square of the first speed and the first output area, the second characterizing parameter is the product of the square of the second speed and the second area of the output. The ratio of the first characterizing parameter to the second characterizing parameter ranges from 0.5 to 1.5. Frame has a lateral rigidity of the frame and a transverse rigidity of the frame, and flexible support has transverse rigidity of flexible support and lateral rigidity of flexible support. The lateral rigidity of the flexible support is less than said lateral rigidity of the frame, and said transverse rigidity of the flexible support is less than said transverse rigidity of the frame. The ratio of the thrust force provided by said engine to the volume of the turbine section including said high pressure turbine and said low-pressure turbine is greater than or equal to 1.5 and less than or equal to 5.5 lbs-force/inm3, at that said thrust is designed static thrust at take-off from sea level.EFFECT: reduction of turbine section concerning both diameter and axial length, increase of efficiency and extension of service life of gas turbine engine.18 cl, 11 dwg, 1 tbl

Application of thermal protection coating on outer surface of caseworks // 2639417
FIELD: process engineering.SUBSTANCE: method of applying a thermal protection coating on the outer surface of the preform fabrication includes a cover, the inner surface dimensions of which match the dimensions of the outer surface of the body. Application is performed by putting a donning cover on the product, and impregnating it with the binder and sealing the heat-protective coating with the ensuing polymerization. Case is seamless and its length exceeds the length of the article on the process over-measure. The ends of the cover are fixed and locked around the perimeter with the help of clamping rings. Clamping rings are pushed to shape the envelope with subsequent fixation and centered with the casework axis. Then the ring with cover, facing the flat end of a product, is pulled over the outer surface of the body until the act of putting cover on the product is completed.EFFECT: invention improves the technological aspects of thermal barrier coating application on the outer surface of the large-sized caseworks.2 cl, 1 dwg
ethod of increasing reactive thrust of valveless pulsejet engine // 2639279
FIELD: engines and pumps.SUBSTANCE: method includes cyclic exhaust of combustion products and absorption of atmospheric air in the inlet port with simultaneous generation of two annular vortices of multidirectional spin, which is carried out in the front part of the combustion chamber in the cycle of expansion of the combustion products flowing towards the inlet duct. A portion of the above combustion product stream is directed through an annular torus-shaped tapered channel to provide flow acceleration and create an ejecting effect at the inlet of the combustion chamber of the engine.EFFECT: increase in reactive thrust due to the intensification of mass transfer, which is carried out by the generation of two annular vortices with a multidirectional spin.3 dwg

Gas turbine engine // 2638709
FIELD: engine devices and pumps.SUBSTANCE: gas turbine engine contains a fan, a compressor section, a combustion chamber in fluid communication with the compressor section, a fan drive turbine communicating with the combustion chamber, a geared system, a flexible support and a lubrication system. The geared system is designed to provide a speed reduction between the fan drive turbine and the fan and to transmit the input power from the fan drive turbine with an efficiency that exceeds 98% and less than 100% to the fan. The flexible support provides support for parts of the geared system. The support moves away from the fixed engine structure to compensate for at least radial movement between the geared system and the fixed structure. The lubrication system is designed to supply grease to the gear system and remove heat energy from the geared system. The fan drive turbine has the first outlet sectional area and is rotatable at the first speed and the engine further comprises of the second turbine having the second outlet sectional area and rotatable at a speed greater than the first speed of rotation. The first characterizing parameter is defined as the product of the square of the first speed and the first area, and the second characterizing parameter is defined as the product of the square of the second speed and the second area. The ratio of the first characterizing parameter to the second characterizing parameter is from 0.5 to 1.5.EFFECT: invention makes it possible to increase the efficiency of a gas turbine engine.11 cl, 1 tbl, 5 dwg

Liquid miscellaneous engine with perfect pneumon pump unit // 2638705
FIELD: engines and pumps.SUBSTANCE: invention can be used in the development of a launch vehicle (LV) for light loads. Liquid rocket engine (LRE) includes the combustion chamber, four pneumatic pump units for feeding fuel and oxidizer, a tank with high pressure helium, a tank with liquid methane, with each pneumatic pump unit containing two outlets for evacuation of the gaseous and liquid components. Gaseous components of methane and oxygen are diverted to the combustion control chambers for subsequent afterburning.EFFECT: reducing the mass of the liquid rocket engine and increasing its efficiency, simplifying the design of the liquid rocket engine.1 dwg

Combustion chamber of liquid-propellant engine (lpe) without generator // 2638420
FIELD: engines and pumps.SUBSTANCE: combustion chamber of LPE, running on a non-generator circuit comprising of fuel and oxidizer supply lines, a chamber block with a supersonic nozzle, with the combustion chamber being annular in shape, parallel to the chamber block rigidly connected by an outer convex and inner curved body of the rotary device to the chamber block and the supersonic nozzle, and the cooling channel of the annular combustion chamber is connected by a cooling channel in the inner curved body of the rotary device to the cooling channel of the chamber block with the supersonic nozzle, and through the cooling channel in the outer convex bottom and through a pipeline the cooling channel of the annular chamber is connected to the pipeline at the outlet of the supersonic nozzle.EFFECT: reducing the linear size of the engine, obtaining an additional set of heat to improve the engine's power characteristics and reducing the hydraulic resistance in the cooling channel of the engine.3 dwg

Combustion chamber of lpe with electroplasma ignition // 2638418
FIELD: engines and pumps.SUBSTANCE: combustion chamber of a liquid-propellant rocket engine, fuelled with liquid oxygen and liquid hydrogen or liquid oxygen and liquefied natural gas, contains an ignition device, a chamber housing with fuel supply lines for cooling, a mixing head with fuel supply lines, a gas main with a supply line for oxidative generator gas, according to the invention, the supply of the generator gas through the gas main of the mixing head is carried out along the axis of the combustion chamber, and the ignition devices, fixed to the profile gas duct between the mains of the generator gas and fuel supply, are installed in the bushings located between the rows of mixing elements from the periphery of the fire bottom in the places of the mixing elements.EFFECT: increasing the reliability of ignition of fuel components in high-thrust LPE combustion chambers, reducing labour intensity and time for maintenance of the combustion chamber and ignition devices.2 dwg
Direct flow turboretactive detonation engine (dftde) // 2638239
FIELD: engines and pumps.SUBSTANCE: direct flow turboreactive detonation engine consists of an inlet part, a middle part and an outlet part. The inlet part includes a fan and a compressor. The middle part includes a device for supplying fuel to the mixing section, a section for mixing fuel with air, a system for burning an inflammable mixture, and a combustion chamber. The output part includes a turbine and an outlet nozzle, as well as a fuel supply system, device for attachment to an external casing, and an engine management system. Inlet and outlet parts are made in the form of axisymmetric circular hollow rotating cones, interconnected through a narrow middle part by their narrow parts, having blades mounted on the inner surfaces of the cones that do not overlap the central part of the channel completely and forming spirals twisted around the common central axis of the channel. The inlet cone with blades serves as a fan/compressor, and the outlet cone with blades is a turbine and an outlet nozzle. The middle part and the outlet cone are combined into one integral part. The device for supplying fuel to the mixing section is made in the form of a centripetal pump. The burning system, by creating short high-voltage electrical impulses, provides combustion of the combustible mixture in the detonation mode. The rotating parts of the engine are attached to the outer casing via bearings fixed to the outer surfaces of the rotating parts.EFFECT: providing an independent horizontal start and the possibility of changing, alternating speeds in the ranges from subsonic to hypersonic.10 cl, 2 dwg

ethod of modelling heat and mass exchange processes with environment of aircraft construction element and device for its implementation // 2638141
FIELD: aviation.SUBSTANCE: method of modelling heat and mass exchange processes of aircraft construction element (ACCE) with the environment under conditions of absolute pressure reduction is based on the introduction of gas stream into the experimental model installation (EMI), providing conditions for interaction of gas stream in the contact zone with ACCE, measuring temperature, pressure, speed. An additional amount of heat is supplied to the ACCE by burning pyrotechnic mixture fixed to the ACCE. The gas stream, pressure and gas composition parameters in the ACCE are selected in accordance with the atmospheric parameters at the current altitude when the ACCE is moving. An additional amount of heat is supplied by heating the ACCE with thermal equivalent of pyrotechnic mixture, for example an electric heater. The ACCE heating zone is additionally supplied with energy in the form of acoustic, laser action, the parameters of which are determined from the condition for increasing the efficiency of ACCE heating. Device for implementing the method includes in its composition an experimental stand, in the form of a closed volume for creating lower absolute pressure, an ECI comprising ACCE fixing system, temperature sensors, pressure, inlet and outlet nozzles, gas analyzer for determining the percentage of gases at the inlet and outlet. The EMI also includes a pyrotechnic mixture with ignition system, a high-speed video camera, a gas stream preparation system, an ACCE rotation system with a fixed source of heat input relative to the direction of gas stream, acoustic, laser emitters, an electric heater.EFFECT: invention makes it possible to expand the boundaries of modelling heat and mass exchange processes of aircraft construction element with environment in conditions of absolute pressure reduction.4 cl, 1 dwg

Structure based on superplastic forming/diffusion binding for reducing noise from air flow // 2637276
FIELD: aviation.SUBSTANCE: inner wall of nacelle includes a monolithic layered structure based on superplastic forming and diffusion binding, monolithic layered structure comprises a core disposed between the first and second facing sheets with formation of a layered structure. Moreover, the core includes a plurality of cells, and the first facing sheet has a plurality of wholes to ensure the arrival of noise and air into the cells.EFFECT: higher structural strength and high resistance to damage, heat resistance, noise reduction without increasing the weight of the structure are achieved.15 cl, 14 dwg
Pulse plasma heat actuator of ejector type // 2637235
FIELD: aviation.SUBSTANCE: invention relates to aircraft flow control systems at subsonic and nearsonic flight speeds. The pulse plasma heat actuator of the ejector type comprises an underwater channel with a check valve, a discharge chamber with built-in needle electrodes, an ejector nozzle, a mixing chamber, a rarefaction cavity with a slit connecting the rarefaction cavity with the surface of the wing, and an output diffuser. The actuator allows to create a high-speed pulsating jet of gas flowing from the nozzle without overheating of the working area in one area of the flow and simultaneously carry out the boundary layer exhaust into the other one.EFFECT: expanded possibility of controlling the aircraft wing flow.2 dwg

Gas turbine engine with high-speed turbine section of low pressure // 2637159
FIELD: engine devices and pumps.SUBSTANCE: gas turbine engine comprises a very high-speed low-pressure turbine, wherein the ratio of the parameter which is determined by product of outlet section area of the low-pressure turbine by square of rotation speed of the low-pressure turbine, the same parameter of the high-pressure turbine is about 0.5 to about 1.5.EFFECT: improvement of performance characteristics of the gas turbine engines, including increase of engine efficiency, and reduction of turbine section dimensions without degrading its efficiency.20 cl, 2 dwg

ethod of operation of three-circuit turbojet engine // 2637153
FIELD: engines and pumps.SUBSTANCE: compressed air from the adaptive fan is divided into three streams. The stream of the first circuit is fed to the gas generator, the exhaust gases from which are fed into the low pressure turbine, and from it through the mixer and the afterburner into the main jet nozzle. The stream of the second circuit is fed through the afterburner into the main jet nozzle. The stream of the third circuit is fed to the nozzle of the third circuit. The operation of the engine is controlled by switching from a three-circuit operation scheme to a two-circuit operation scheme and back, and also by changing the degree of the two-circuit engine by switching the direction of the compressed air streams by means of the distribution devices and engaging the afterburner. In the maximum and transient operation modes with engine forcing, the compressed air stream of the third circuit is fed directly from the third circuit channel through the afterburner into the main jet nozzle. Opening and closing of switchgears for connecting and disconnecting the channel of the third circuit is carried out according to the values of the reduced low pressure rotor speed.EFFECT: increase of the maximum flight thrust of the turbojet engine in maximum and transient modes with engine forcing while maintaining fuel consumption parameters.4 dwg
Electrothermal micromotor // 2636954
FIELD: engines and pumps.SUBSTANCE: electrothermal micromotor comprises an outer and an inner cylindrical bodies, arranged coaxially with the formation of toroidal cavity between their walls, a swirler of inlet fuel, a pipeline for fuel supply into the swirler, a gas duct with a jet nozzle, a cylindrical heating element and a tubular thermocouple located at the inlet of the jet nozzle, current leads of the heating element and thermocouples brought out through the end face of the inner housing by means of a sealing heat-resistant sealant. At one end the outer and inner bodies are hermetically connected to each other by means of flanges, and at the other end on the side surface of the outer body, there is a pipeline for feeding fuel into the toroidal cavity, inside which there is a swirler of the inlet fuel flow in the form of a screw channel, in the form of a two-start thread on the external surface of the internal body, the outer surface of which is in contact with the inner surface of the outer body which outlet is connected to the cavity of the inner body at the inlet of the gas duct in the form of the screw channel, formed by the outer surface of the cylindrical heating element, the tubular body of the thermocouple, placed along the helical line on surface of the heating element and contacting with internal surface of the internal body. Jet nozzle is mounted on end of the inner body and is provided with an outer flange hermetically connected with the flange of outer body. The thermocouple sensitive element is located near the inlet to the nozzle critical section, and on the side opposite to the nozzle, the length of the outer body exceeds the length of the inner body, on which the shoulder contacting the inner surface of the outer body is made, the sealing heat-resistant sealant is placed in the cavity of the outlet of the current leads of the thermocouple and the heating element, which is formed by the free inner surface of the outer body and a stop washer fitted on the cylindrical heating element.EFFECT: increased tightness degree, reduction of weight and increase of specific impulse of the micromotor thrust.3 dwg

Dual-mode igniter and two-mode method of injection for ignition of rocket engine // 2636357
FIELD: engine devices and pumps.SUBSTANCE: in accordance with the invention, the igniter comprises a supply element (21) for feeding a first propellant (A), a feed element (31) for supplying a second propellant (B), a feed element (41) for supplying high-pressure fluid medium (F), a first buffer tank (22), a second buffer tank (32), a first switching device (50), a second switching device (60) and a flare-forming combusion chamber (10); an opening located downstream of the first buffer tank (22) and the second buffer tank (32) - both open into the chamber (10), the first switching device (50) and the second switching device (60) are designed so as to connect respectively the upstream opening of the first buffer tank (22) either with the feed element (21) for feeding the first propellant or with the feed element (41) for supplying the high-pressure fluid (F), and to connect the upstream opening of the second buffer tank (32) or to the feed element (31) for feeding the second propellant or with the feed element for feeding the high-pressure fluid (F).EFFECT: invention ensures launching of the rocket engine both under low pressure conditions and under high pressure conditions.10 cl, 2 dwg

Pulsating gas turbine engine // 2635953
FIELD: engine devices and pumps.SUBSTANCE: pulsating gas turbine engine comprises a body, a rotor equipped with jet engines with a compressor on a shaft, and a gas turbine fitted coaxially on the rotor shaft. The rotor with tangentially installed pulsating jet engines is built into the gas turbine with blades which is bifurcated in form of a fork and mounted coaxially on the rotor shaft enclosing it symmetrically on both sides. The blades of the turbine are provided with shaped cuts with slight clearance along the contour of nozzles of pulsating jet engines made in form of parabolic chambers. Spark plugs of fuel-air mixture coming from flow channels through check valves are installed in foci of parabolic chambers arranged in apexes of the parabolic chambers, into which by means of fuel channels with the help of conical air intakes mounted on the rear sides of the parabolic chambers which act as compressors and forming jet pumps are supplied with fuel in the form of fuel-air mixture (aerosol). From outlet nozzles of the parabolic chambers, the focused flows of combustion products of fuel-air mixture are directed to the of gas turbine blades. The opposite torques on the rotor shaft and on coaxial turbine shaft are summed with the help of differential.EFFECT: increased efficiency of the pulsating gas turbine engine.3 dwg

Convergent-divergent nozzle of turbomachine, bypass turbojet engine and turboprop engine // 2635863
FIELD: engine devices and pumps.SUBSTANCE: convergent-divergent nozzle of the turbomachine comprises an annular central structural element and an annular housing coaxially arranged around central structural element so as to limit an annular channel of the engine gas flow there with it. Between critical section of the nozzle and nozzle outlet section, an outer profile of central structural element and an inner profile of the housing are formed in longitudinal section, by means of curves of lines, whose radii of curvature have value of second order derivative of function y (X) determining the shape of said curves of lines relative to axial position along the corresponding curve of the line. The corresponding radii of curvature of curves of the lines are identical in absolute value. Other inventions of the group refer to the bypass turbojet engine and the turboprop comprising said nozzle.EFFECT: increased aerodynamic properties of the nozzle.5 cl, 3 dwg

System for controlling supersonic ramjet engine // 2635758
FIELD: engine devices and pumps.SUBSTANCE: on surface of the central body front part there are two to four receivers of air pressure and a receiver of total pressure of undisturbed flow, pressure sensors are positioned inside the central body and connected by air line with a receiver of full pressure of undisturbed flow, receivers of the air pressure on the central body and in the front part of the central body, on the other hand, with control unit consisting of a processor module, control module and power key module for outputting signal to nozzle control unit depending on mach number, pressure drop, angle of attack, angle of sliding.EFFECT: increased accuracy for maintaining antisurge reserves, restoration coefficient of total pressure, increased range and creation of possibility for selection of different flight paths.2 dwg

ethod of control of ramjet engine of winged rocket // 2635757
FIELD: aviation.SUBSTANCE: in the event of malfunction of command pressure indicators, a command for the execution of the backup control algorithm for the ramjet engine is issued. The predetermined height of the winged rocket is reached and the speed of the winged rocket is maintained, corresponding to the altitude of the flight of the winged rocket. At the same time, the fuel consumption is regulated by the speed and altitude parameters of the winged rocket, and the height and speed of the winged rocket are measured using satellite navigation equipment.EFFECT: increase of reliability of work and increase of survivability and safety of flight.2 cl, 1 dwg

Driving device for moving movable hood of thrust reverse // 2635748
FIELD: transportation.SUBSTANCE: driving device for moving a movable hood of a thrust reverse comprises a drive, a locking device, and a locking retention device. The drive includes a first element and a second element mounted movably relative to the first element. One of the first or the second element is a screw, and the other is a nut made with the possibility to interact with the screw, so that the rotation of the first element relative to the second element results in the translational movement of the second element relative to the first element. The locking device comprises a locking part and a drive mechanism. The locking part is made rotatable relative to the first element between the locked position, in which the locking part prevents the rotation of the first element, and the unlocked position, in which the locking part provides the rotation of the first element. The drive mechanism is configured to rotate the locking part into the unlocked position. The locking retention device comprises a retaining piece configured to be progressively moved relative to the first element between the first and the second position. In the first position, the retaining piece provides the rotation of the locking part between the locked position and the unlocked position. In the second position, the retaining piece prevents the locking part from rotating into the locked position.EFFECT: invention allows to reduce the dimensions of the locking device and to reduce its sensitivity to sharp accelerations.19 cl, 16 dwg

Solid-propellant rocket engine // 2635427
FIELD: engines and pumps.SUBSTANCE: engine contains a body with bottoms, a channel charge fastened to the body and equipped with a combustion surface compensator in the form of an annular slit located at the front bottom, a nozzle. The charge channel is blind, and the annular slit is inclined. The charge is equipped with a second combustion surface compensator formed by a portion of the charging channel above the nozzle and a partially burning end face of the charge at the rear bottom. The surface of the channel between the combustion surface compensators is fastened to the armour cover. In this case, the cover is connected to the inner surface of the body by means of flexible cords, arranged evenly along the circumference in a plane equally spaced from the ends of the armoured section of the channel. Each cord is placed in a protective shell.EFFECT: invention makes it possible to increase the effective summary by a factor of 1,5-2 due to the creation of conditions for excluding the ignition of part of the surface of the charge channel during the entire engine operating time while maintaining the overall dimensions of the engine and achieving the possibility of using high-energy fuels, existing in the industry, without changing the quantitative and qualitative composition thereby increasing the effectiveness of their use.2 cl, 2 dwg

Gas turbine engine (versions) and method for increasing gas turbine engine performance // 2635181
FIELD: engines and pumps.SUBSTANCE: fan section contains a fan which is rotatable about an axis. A reducer interacts with a fan, wherein the mentioned reducer comprises a planetary gear of a fan drive with a gear ratio of the planetary gear of at least 2.5 and less or equal to 5.0. The bypass ratio of the engine lies in the range from 11.0 to 22.0. The planetary gear contains a sun pinion, a lot of satellite pinions, a crown pinion and a pinion cage. Each of the plurality of satellite pinions comprises at least one bearing. The crown pinion is fastened to prevent rotation. The low pressure turbine is mechanically attached to the sun pinion. The fan section is mechanically attached to the pinion cage. The fan blade tip speed is less than 1,400 feet per second.EFFECT: ways to increase the thermal, traction efficiencies and transmission efficiency are achieved by reducing the load on its bearings and pinions.7 cl, 3 dwg

Solid-propellant rocket engine body // 2635171
FIELD: engine devices and pumps.SUBSTANCE: body of the solid fuel rocket engine comprises a bottom with a central hole and a sleeve having a portion bent inside the body located in the central hole region and configured to mount a process wedge between the bottom and the sleeve portion bent inside the body. There are radial bands installed between the bottom and the sleeve. The bands are fixed by one edge portion on the sleeve portion bent inside the body from the side of its surface facing the bottom, and the second edge portion is fastened to the bottom. The middle portion of the bands forms loops located between the bottom and the sleeve. There are hinges for installing the process wedge therein and having a length equal to the perimeter of the cross-section of the process wedge that is interfaced therewith. The edge portion of the bands can be provided with tails, the length of which allows them to be fastened together.EFFECT: increasing the reliability of the solid fuel rocket engine.2 cl, 5 dwg

Inter-blade air-substituting method of increasing thrust of jet engine and device for its implementation // 2634976
FIELD: engines and pumps.SUBSTANCE: method is carried out by increasing the resistance of release body by further resistance from interaction with a cross-inserted new body in the form of a column of air in place of waste air, formed by cross-lamellar filling of the released space by the air from the surrounding space as the cut-off of the waste body. Distinctive feature of proposed design from the cylindrical rotary mechanism forming channels of axial direction is a one-way closed cylindrical body with side screw window, a dividing cylinder between one end window at all end windows opened on opposite side of the cylinder. The generated thrust force is comparable to the lifting force generated by the aircraft wing with the difference, which is created in associated direction by forced supply of fixed air mass under engine nozzle, but not imparting motion to the wing of the aircraft at the intersection of the air space.EFFECT: increased thrust force of the jet engine.2 cl, 15 dwg
Three-tier working blade of turbo-fan // 2634509
FIELD: machine engineering.SUBSTANCE: three-tier working blade of the turbo-fan comprises a fan blade and a turbine blade which are arranged in series from a turbo fan casing to a rotor disc, and a turbine blade connected to each other by means of an intermediate element with formation of three flow-through gas channels. The intermediate element is made in form of working blade of a turbo-expander with formation of smooth transition from a profile to the profile of all three working blades. The flow part of the working blade gas channel of the turbo expander is limited by shelves. The fan working blade is connected to the turbo expander working blade by means of a detachable hinged joint.EFFECT: intensive engine cooling, increased engine thrust, reduced weight and increased strength properties of the three-tier working blade of the turbo-fan, as well as its reliability as a whole.1 dwg
Solid-propellant charge // 2633980
FIELD: engine devices and pumps.SUBSTANCE: charge to the start-up jet engine comprises a bundle of tubes of high nitrogen pyroxylin powder fixed to the engine bottom, and an igniter located on the end of the charge. In the composition of the powder containing the complex combustion catalyst, the content of residual solvent is 0.2…0.7%, and the channel diameter dk tube at a length l corresponding to the length of the cylindrical part of the engine chamber, meets the ratio , where æ=240…300. The value of the temperature difference of the initial velocities at the level of 4…8% is provided.EFFECT: invention makes it possible to increase maximum possible rate of reactive grenade.1 cl

Solid fuel gas generator // 2633976
FIELD: machine engineering.SUBSTANCE: gas generator comprises a central hollow cylinder closed at one end and open in the form of a converging nozzle from the other end, local gas-generating parts of solid fuel charge located in the cylinder with separating partitions therebetween, and an ignition system of the charge parts. The gas generator further comprises a cylindrical body with peripheral and central filler charges, a resonance chamber and two circular grids. The cylindrical body is provided with a bottom on one side and a conical adapter with a branch pipe on the other side. The resonance chamber is made in the form of a cup with a piston located therein and installed in the center of the body bottom. The filler charges are installed in the body between inlet and outlet circular grids with formation of a circular channel therebetween and coaxial the hollow cylinder. The hollow cylinder is in axial hole of the central filler charge and faces the nozzle along the axis towards the resonance chamber with formation of an intermediate cavity therebetween communicating through the circular channel between the filler charges with the branch pipe.EFFECT: possibility for repeated activation of the gas generator and controlling the mass flow rate of gaseous low-temperature working medium, increased reliability of the gas generator.5 cl, 3 dwg

Centrifugal turbine // 2633974
FIELD: engines and pumps.SUBSTANCE: centrifugal turbine comprises a housing, am impeller disk of a centrifugal turbine with a blade ring and a band, a nozzle cascade, according to the invention, a counter-protrusion 5 is formed on the impeller disk of the centrifugal turbine 3 on the opposite side of the blade ring 6, forming a labyrinth seal 11 with the housing of the turbine 1 and equal in weight to the blade ring with the bandage.EFFECT: simplified design, reduced weight of the turbo-pump unit, increased reliability for long-term operation of turbo-pump units at high rotational speeds and at high temperatures of the working body after the gas generator, eliminated axial force influencing the turbine.1 dwg

Solid fuel jet engine with single changeable thrust vector // 2633973
FIELD: engine devices and pumps.SUBSTANCE: solid fuel jet engine with a single changeable thrust vector comprises a charge body, an ignition device, and an oblique cut nozzle. The nozzle is divided into parts by the junction plane passing through the point of intersection of minimum nozzle generatrix with the plane of oblique cut. The parts of nozzle are interconnected by a heat-resistant ring with predictable entrainment of ring material from the action of combustion products charge jet.EFFECT: invention makes it possible to simplify the development of rocket engines for correcting the flight of carrier rocket and elements detachable from it.5 cl, 2 dwg
ethod of working process organizing in direct-flow air jet engine // 2633730
FIELD: engines and pumps.SUBSTANCE: method includes supplying a powder of metal fuel to the combustion chamber, igniting it and burning it in a stream of air from the air inlet. The powder in the form of a uniformly mixed suspension in a liquefied fuel gas placed in a fuel tank is preloaded with a displacement pressure, heated and supplied to the combustion chamber through a nozzle. The maximum diameter of the powder particles, the extrusion pressure and the heating temperature of the slurry are determined from the protected ratios.EFFECT: increased power characteristics and reliability of the direct-flow air-jet engine.2 cl, 2 tbl, 2 dwg

Design of gear turbofan gas turbine engine // 2633498
FIELD: engine devices and pumps.SUBSTANCE: gas turbine engine comprises a fan, a compressor section, a combustion chamber in fluid communication with the compressor section, a turbine section in fluid communication with the combustion chamber, and a speed adjustment system. The turbine section includes a fan drive turbine and the second turbine, where the fan drive turbine comprises of a plurality of turbine stages. The fan comprises of a plurality of blades configured to rotate around the axis. The ratio between the number of fan blades and the number of stages of the fan drive turbine is from 2.5 to 8.5. The speed adjustment system is driven by the fan drive turbine to rotate the fan around the axis. The fan drive turbine comprises of the first rear rotor connected to the first shaft, and the second turbine comprises second rear rotor connected to the second shaft. Between the first shaft and the second shaft, an annular gap is formed. The first bearing assembly is located axially behind the first connection between the first rear rotor and the first shaft and the second bearing assembly is located in the annular gap formed between the first shaft and the second shaft.EFFECT: invention makes it possible to eliminate the need for load-bearing structures connected to the fixed structure via an intermediate power frame, to reduce the length of the shafts, to support the external shaft coaxially with the coupling of the high-pressure turbine rotor with the outer shaft, to provide a more compact turbine section, and to reduce its weight and fuel consumption.20 cl, 13 dwg

Design of gearbox of turbofan gas turbine engine // 2633495
FIELD: engines and pumps.SUBSTANCE: gas turbine engine comprises of a fan rotatable about the axis, a compressor section, a combustion chamber in fluid communication with the compressor section, and a turbine section in fluid communication with the combustion chamber. The turbine section contains a fan drive turbine and the second turbine. The second turbine is located before the fan drive turbine. The fan drive turbine contains at least three rotors, wherein at least one rotor has a radius (R) of the channel and the effective radius (r) of the rim, and r/R ratio ranges from about 2.00 to about 2.30.EFFECT: speed changing system is powered by a fan drive turbine to ensure the fan rotation around the axis.27 cl, 1 tbl, 12 dwg

Gear fan-type gas-turbine motor arrangement // 2633218
FIELD: engines and pumps.SUBSTANCE: gear unit, for example an epicyclic wheel train, can be used to drive the fan section, so as to allow the fan section to rotate at a speed different from that of the turbine section. The turbine section is in fluid medium communication with the combustion chamber and comprises a fan drive turbine and a second turbine. The fan drive turbine comprises a plenty of turbine stages. The fan has a plurality of vanes rotated about an axis. The ratio of fan vanes quantity to a number of fan drive turbine stages is from 2.5 to 8.5. The fan drive turbine comprises the first backside rotor which is attached to the first shaft, and the second turbine comprises the second backside rotor which is attached to the second shaft. Axially behind the first connection between the first backside rotor and the first shaft is the first bearing assembly and axially before the second connection between the second backside rotor and the second shaft is the second bearing assembly. The second turbine has at least two stages and is capable of operating at a first pressure, and the fan drive turbine has more than two stages and is capable of operating at a second pressure which is less than the first pressure.EFFECT: increase working characteristics and productivity by using the required combinations of the disclosed design features of the various components of the described gas turbine engine.19 cl, 12 dwg

Suspension hinge of rotary nozzle made of composite materials and method for its manufacture // 2632393
FIELD: machine engineering.SUBSTANCE: suspension hinge consists of an elastic part (1) with fastening elements (2, 3). In it, the elastic part (1) is made of braided (4) and wound (5) thread layers with their dense arrangement and formed around the fastening elements (2, 3). The fastening elements (2, 3) are made with the inclusion of force rings in their structure with bending of same threads of the elastic part (1) around them. The method is also claimed for manufactuing said suspension hinge of rotary nozzle and consisting in execution of successively performed assembly and winding operations. In the course of preliminary manufacture of structural and technological carcass of hinge made from braided (4) and wound (5) threads with use of metal rings (6, 7) in areas of fastening elements followed by its circular piercing and impregnation with silicon.EFFECT: creation of suspension hidge for connection of links of structures with provision of relative movement, the operation principle is not related to shear deformation of material and method of its manufacture.2 cl, 5 dwg

Reactive engine, method for shooting with rocket ammunition and rocket ammunition // 2631958
FIELD: weapons and ammunition.SUBSTANCE: jet engine includes a housing, a cantilever bar, a hollow central body, a means for controlling the movement of the hollow central body when the nozzle is opened, and means for moving the hollow central body to close the nozzle at a predetermined point of time by providing the predetermined resultant of pressure forces of the gaseous combustion products of propellant to the central body. The housing has a front part, a combustion chamber adapted to accommodate the propellant charge, and a nozzle. The cantilever bar is fixed in the front part of the housing with a free end portion and protrudes outward from the nozzle. The hollow central body is movable along the cantilever bar in the direction of the effluent of the gaseous combustion products of the propellant to open the nozzle and move in a direction reverse to the mentioned one to close the nozzle. The hollow central body covers the cantilever bar along its lateral surface. The means for controlling the movement of the hollow central body when the nozzle is opened is located on the cantilever bar inside the hollow central body. Another invention of the group relates to a rocket ammunition comprising a head part, the above mentioned jet engine connected to the head part on the side of the front part of its housing and a stabiliser. When firing with the rocket ammunition comprising the mentioned jet engine from a mobile launcher with a launch tube, the gaseous combustion products of the propellant are applied to the hollow central body to open the nozzle in the launch tube, allowing the mentioned products to flow out of the open nozzle to create a driving force acting on the rocket ammunition. The position of the hollow central body in the open nozzle is adjusted depending on the charge temperature of the propellant. Simultaneous supply of gaseous products of propellant combustion inside the hollow central body is provided. The nozzle is closed by means of the hollow central body at a predetermined moment of time by providing the predetermined resultant of pressure forces of the gaseous combustion products of the propellant to the hollow central body.EFFECT: increasing accuracy of shooting, preventing the impact of a jet of powder gases on the shoote and its barotrauma when the initial velocity of the rocket ammunition is increased.18 cl, 4 dwg
 
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