Low-noise turbine for geared turbofan engine

FIELD: engines and pumps.

SUBSTANCE: bypass gas turbine engine is used in combination with step-down reduction gearbox to reduce fan rotation speed relative to low-pressure turbine rotation speed. Disclosed also are turbine module and method of its designing. Gas turbine engine is made so, that number of blades in low-pressure turbine, multiplied by low-pressure turbine rotation speed, leads to formation of working noise, which lies outside of human hearing sensitivity range.

EFFECT: enabling audible noise reducing during landing approach.

20 cl, 1 dwg



Same patents:

FIELD: transport.

SUBSTANCE: proposed method comprises feed of air and creation of azimuthal and axial flow to be compressed by compressor, flow heating and discharge at the rate higher than the turbine blades azimuthal rpm and feed of extra air volume. First, hot flow pulse is transferred to said extra air volume. Then, turbine torque is developed to change the flow azimuthal moment to axial force. At compression said flow is directed in azimuthal and axial directions from the axis. Azimuthal vortex is developed to force air flow beyond the compressor blade edges. Said flow is turned azimuthally and axially to the axis to develop the azimuthal vortex. This job is reiterated more than once. Note here that proposed turbojet comprises low-pressure compressor with wing-like elements of blade profiles, turbine, azimuthal combustion chamber, teeth rows with blades upstream and downstream of said combustion chamber. Flow ejection mixing chamber is arranged downstream of combustion chamber. Screws of low-pressure compressor and turbine are made integral and composed of lengthwise surfaces including the sections with profile structure with wing-like elements and centrifugal sections. Fixed row with blades is arranged downstream of turbine, trailing edges of said blades being directed along the axis of engine. The latter comprises extra high-pressure compressor. Compressor rotor incorporates more than one section of variable profile reaching the screw blade outer edge at section edges and having the minimum at its centre. Screw blades are secured at rotor minimums while there leading and trailing edges are directed along rotor spinning. Blades of the last row are directed at outlet along rotor axis. Flow exhaust device is arranged above the rotor is sectionalized and comprises shaped blades fitted at the profile maximums. Blades leading and trailing edges are directed against the rotor rotation.

EFFECT: simplified aircraft design, expanded range of applications.

2 cl, 1 dwg

FIELD: engines and pumps.

SUBSTANCE: fixing structure for attachment of a guide blade to a frame or a casing of a fan of an aircraft engine. The guide blade is formed of composite material. The guide blade is intended to straighten an air flow. The fan casing is arranged outside the fan frame. The fixing structure includes the following: a mating surface, the first installation section, a supporting element, a supporting mating surface and the second installation section. The mating surface is formed in an end section of the guide blade. The first installation section is formed in the mating surface of the guide blade. The supporting element is made from metal as its component material, with that, the supporting element is connected in the form of an integral unit of the fan frame or the fan casing. The supporting mating surface is subject to conjugation with the mating surface of the guide blade and formed in the supporting element. The second installation section has a possibility of wedge-like adhesion to the first installation section of the guide blade and formed in the supporting mating surface of the supporting element. Rigidity and strength of conjugation is increased between the guide blade and the frame or the casing of the fan.

EFFECT: improved design.

9 cl, 9 dwg

FIELD: aircraft engineering.

SUBSTANCE: rotor crossing of combined zone with mutual entry of blades in interblade space of jet stream with simultaneous stay of other rotor blades in ambient air space. Development of rotor torque from jet stream by one part allows, at a time, the other part to develop rotor thrust along with producing of airflow of one direction with lessened jet stream to increase power output in exchange for rate. Removal of rotors from jet stream recovers jet principle of motion. Lateral method of torque production in zone of partially aligned rotors allows the other free parts to implement the thrust without mutual negative influence.

EFFECT: lower costs of cooling, higher safety and efficiency.

21 dwg

Turbojet // 2494271

FIELD: engines and pumps.

SUBSTANCE: proposed turbojet comprises low-pressure turbine and adjustable leaf mixer including conical shell at its outlet. Conical shell is arranged between turbine and mixer to form mid annular chamber between gas channel of inner duct and air channel of outer duct. Annular chamber outlet is communicated via annular outlet slot with gas channel and air-gas flows mixing zone while its inlet is communicated with air intake composed by radial annular rig of turbine housing an mixer annular shell to cover flow section by axially displacing radial ring. There is a circular rib ot the outer side of the last working blade of the turbine, the radisl rib is installed at the external hausing of the turbine.

EFFECT: higher efficiency of mixing inner and outer duct flows, reduced hydraulic resistance, higher thrust at transonic conditions.

4 dwg

FIELD: machine building.

SUBSTANCE: fan blade of a jet turbine engine with a double flow, which contains internal aerodynamic part (11) and external aerodynamic part (12), which are combined in radial direction (Z) and separated with platform (10). Internal aerodynamic part (11) includes one aerodynamic profile (13), and external aerodynamic part (12) includes at least two aerodynamic profiles (14); at that, leading edges of the above aerodynamic profiles (14) of the above aerodynamic part (12) of the blade are axially aligned in a row.

EFFECT: reduction of the number of fan blades at maintaining satisfactory quality owing to maintaining an increased relative pitch of the internal aerodynamic part of the fan blade.

8 cl, 4 dwg

FIELD: engines and pumps.

SUBSTANCE: gas turbine engine is designed with aft location of an open propeller fan with a gas generator and a gas channel in a sleeve of the propeller fan at outlet from the gas generator, and with shanks of blades of the propeller fan located in hollow racks. On external side from the gas channel there is an annular air cavity connected at inlet to outlet from the first stage of the gas generator compressor, and at outlet - to atmosphere via forward and aft labyrinth seals located before rotors of the propeller fan and also via racks and an additional nozzle in an aft cone of the sleeve of the propeller fan. Labyrinth seals are connected at outlet to directed opposite to incoming air flow additional annular forward and aft air intakes located on the sleeve between forward and aft rotor of the propeller fan correspondingly. An additional nozzle is designed with possibility of adjustment of its flow area.

EFFECT: increasing reliability and economy of the engine by eliminating parasitic leakages from the gas annular channel of the propeller fan.

5 dwg

FIELD: engines and pumps.

SUBSTANCE: double-flow jet turbine engine includes cylindrical cover installed at intermediate casing outlet and restricting on the outside the annular space of secondary flow. Cylindrical cover is formed with lattice-like frame and removable cowling panels fixed on the frame. Lattice-like frame includes inlet annular flange to be attached to intermediate casing; outlet flange to be attached to exhaust cowling, and rigid beams attaching both flanges to each other.

EFFECT: invention allows reducing the weight and simplifying the engine maintenance.

10 cl

FIELD: engines and pumps.

SUBSTANCE: invention relates to axial-flow fans of bypass fanjets. Proposed fan comprises gas generator with reducer and fan impeller. Said impeller comprises vanes fitted in barrel with their outer end edge tapered surfaces. Said tapered surfaces feature smaller base with diameter equal to that of vane end surface front edge. Diameter of vane end surface rear edge exceeds that of end surface front edge to make cone larger surface. Cone features height equal to vane end axial width.

EFFECT: higher efficiency, pressure and thrust.

9 cl, 8 dwg

FIELD: engines and pumps.

SUBSTANCE: gas turbine propfan engine with rear arrangement of propfan rotors comprises a gas generator with a compressor, a gas channel in a propfan bushing, a nozzle. In the gas channel there are hollow stands of propfan rotors. At the outer side from the gas channel of the propfan bushing there is an outer circular air cavity connected at the inlet to the intermediate stage of the low pressure compressor, and at the outlet - via hollow stands of propfan rotors with an additional axial nozzle in a bushing fairing. The outer circular air cavity is separated from atmosphere with an outer shell with radial ribs directed towards the bushing axis, where labyrinth seals are arranged. Ratio of the nozzle area at the outlet from the propfan bushing gas channel to the area of the additional axial nozzle in the propfan bushing fairing makes 5…20. The ratio of the outer diameter of the propfan bushing to the average diameter of labyrinth seals of propfan rotors makes 1.05…1.20.

EFFECT: higher cost-effectiveness and reliability of a gas turbine propfan engine by reduction of parasite leakages of gas from a propfan bushing into atmosphere and also by use of air spent in a cooling system of a propfan bushing to develop gas turbine engine thrust.

3 dwg

FIELD: engines and pumps.

SUBSTANCE: extreme values of cold flow expansion coefficient, which correspond to the beginning and the end of cruise flight phase, are determined. Reference value (VR) of expansion coefficient is chosen between the above extreme values. For that reference value (VR) of expansion coefficient there determined is theoretical value (Ath) of surface area of outlet hole (6) of cold flow as per additional theoretical value. Additional theoretical value represents the ratio between theoretical surface area of cold flow outlet hole and nominal cross-section area of the nozzle throat. Outlet hole (6) is located along longitudinal axis (L-L) so that its surface area corresponds to the above theoretical value (Ath). Inner casing (13) of fan preferably has the shape which at least approaches the barrel; at that, throat (T) of cold flow nozzle is located behind the largest cross section (23) of fan inner casing (13).

EFFECT: achieving preliminary adaptation of nozzle to cruise flight phase.

7 cl, 3 dwg

FIELD: manufacture of gas-turbine engines.

SUBSTANCE: proposed gas-turbine engine includes low-pressure compressor with load-bearing intermediate casing and high-pressure compressor with swivel inlet guide apparatus and rotor mounted on bearing on side of first impeller of high-pressure compressor. Provided at outlet of intermediate casing on its bush is taper surface directed towards flow section of high-pressure compressor, thus forming slotted cavity with front baffle plate of ring of internal inlet guide apparatus of high-pressure compressor; this slotted cavity is connected with oil cavity and intermediate inlet cavity at outlet through inter-labyrinth cavity and labyrinth seals. Angle of generatrix of cone of taper surface of bush of load-bearing intermediate casing ranges from 10 to 40 deg. Ratio of height of blade of swivel inlet guide apparatus of high-pressure compressor to height of slotted cavity between taper surface of bush of load-bearing intermediate casing and baffle plate of inner ring of swivel guide apparatus of high-pressure compressor is equal to 30-70.

EFFECT: enhanced reliability of engine due to increased service life of oil support of front bearing of high-pressure compressor and avoidance of leaks.

2 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention can be used in designing of fairings of propeller fan rotor bushing for double-flow turbojet engines with reversing of thrust by turning propeller fan blades. Proposed fairing of propeller fan rotor bushing contains projections on outer surface under each blades separated from blades by minimum possible clearance. Fairing is provided with bushings for accommodating blade trunnions for turning when propeller fan changes from forward to reverse thrust and vice versa. Projections are made to suit root section of each blade and are located under front and tail parts of root section of each blade installed for operation at forward thrust. Height of projections is limited by plane square to axis of bushing for trunnion and passing through point on outer surface of rotor bushing fairing located after blade in plane of axes of bushing of trunnion and fan rotor at a distance from axis of trunnion bushing same as rear edge of projection under tail part of blade profile.

EFFECT: reduced overflow of air from higher pressure surface of blade to lower pressure surface.

4 dwg

FIELD: engines and pumps.

SUBSTANCE: aviation bypass turbojet engine comprises fan, high-pressure compressor, combustion chamber, high and low pressure turbines, mixer and afterburner and nozzle common for both circuits. Downstream the first stage of high-pressure compressor that provides for full pressure increase extent at takeoff regime of not more than πIext*=1.4…1.5, permanently open annular channel is arranged with honeycomb straightener, through which at all modes of engine operation, some air is relieved downstream stage into cocurrent flow of air in external circuit downstream fan.

EFFECT: increased extent of engine bypass-ratio and increased efficiency at non-augmented modes of operation.

1 dwg

FIELD: engines and pumps.

SUBSTANCE: three-body double-circuit turbojet engine with high by-pass ratio comprises front fan and back fan in front part of intermediate tray equipped with external profile grate in circuit of secondary flow and internal profile grate in circuit of primary circuit. Fan blades are directed radially outside and reach fan tray, which limits secondary flow circuit outside. Engine also comprises low pressure compressor intended for compression of air coming to primary flow channel. Front and back fans are rotated directly and separately by means of two coaxial shafts. Blades of back fan are installed in secondary flow circuit, starting from disc joined with its driving shaft via back wheel of low pressure compressor movable blades. Low pressure compressor additionally comprises at least one front wheel of movable blades rotated by driving shaft of front fan, and external stator, in opening of which fixed blade grates are installed, which are inserted between wheels of movable blades. External stator is installed on fan tray via the second intermediate tray that comprises the second external profile grate in secondary flow circuit between blades of front fan and blades of back fan and the second internal profile grate in primary flow circuit. In the first intermediate circuit shaft is installed for rotation of back fan on axial thrust bearing together with driving shaft of front fan on intershaft bearing. Additional bearings are installed between intermediate tray and shafts.

EFFECT: reduction of turbine length and weight.

7 cl, 2 dwg

FIELD: engines and pumps.

SUBSTANCE: helical fan gas-turbine engine comprises turbocompressor with body, compressor, combustion chamber, output of which is connected by gas track to turbine, and helical fan. Helical fan is connected with compressor via magnetic coupling, which comprises slave half-coupling installed in compressor, for instance, on its external blades and slave half-coupling installed on turbocompressor body. Compressor is arranged as double-cascade, with the possibility of cascades rotation in opposite sides. Helical fan is arranged as double-step and comprises front and back steps, which are arranged with the possibility of rotation in opposite sides. Steps of helical fan are installed inside fairing.

EFFECT: increased efficiency and reliability of aviation engine.

3 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: screw blower aircraft gas turbine engine comprises turbo compressor with a housing, compressor, combustion chamber with its outlet communicating, via gas channel, with the turbine, and a screw-type blower linked up with the turbine via a magnetic coupling. The said magnetic coupling comprises a driving half-coupling arranged in the turbine, for example, on its vanes and driven half-coupling mounted on the compressor housing. The two-cascade turbo compressor allows rotating the cascades in opposite directions. The two-stage screw blower includes the front and rear stage arranged to revolve in opposite directions.

EFFECT: higher efficiency and reliability.

2 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed propfan engine comprises turbo compressor with housing, compressor, combustion chamber with its outlet communicated, via das duct, with turbine and two-stage propfan. One stage of the latter is coupled with compressor via magnetic coupling, while another stage is coupled with the former via reversible reduction gear. Magnetic coupling comprises half-coupling arranged in compressor, for example on its vanes, and drive half-coupling mounted on turbo compressor housing. Propfan stages are arranged inside fairing.

EFFECT: higher efficiency and reliability.

3 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed method consists in determining fan channel critical zone starting in nozzle neck and passing forward. Note that in this zone no variations in said channel geometry and, hence, inner tubular surface of acoustic attenuation coat rear part are possible without changing parametres of the nozzle. Note also that aforesaid coat is applied on fan outer fairing inner part. Inner tubular surface in acoustic attenuation coat front part convergent zone is varied in direction of gradual increase in thickness of aforesaid coat to wards aforesaid rear part of this coat. Gradual variation in thickness of the coat rear part inner tubular surface is continued till formation of greater thickness zone in said coat rear part. Greater thickness zone rear end is jointed to critical zone front end by inner tubular surface with curved profile. Another invention of proposed set relates to two-stage gas turbine engine that comprises acoustic attenuation coat with circular section applied onto fan outer fairing and comprising curved profile. The latter is arranged between larger thickness zone and nozzle critical zone. Said thickness approximates to that of the coat front part and nozzle critical zone.

EFFECT: reduced noise in two-stage gas turbine engine rear part with no deterioration in engine operation.

5 cl, 4 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises air intake (2) with cylindrical inner wall and fan (3) with cylindrical casing and air fed from said air intake. Air intake inner cylindrical wall and fan casing form an integral cylindrical component (30) made from composite material consisting of resin and fibrous filler. Said composite material can be based on carbon fiber. Aforesaid cylindrical component is jointed, via flange (32) and bolts, to outer hood (22) of fan duct (23). Rear end (30R) of cylindrical component (30) represents a homogeneous member. Flange (32) is attached to said rear end (30R) and fixed thereon.

EFFECT: ruling out lamination of composite material.

2 cl, 5 dwg

FIELD: engines and pumps.

SUBSTANCE: aircraft bypass gas turbine engine comprises hollow nacelle with front air inlet and rear air outlet. Central hot air generator is arranged axially in said nacelle and fan is mounted ahead of it to generate cold airflow for gas turbine engine. Outer and inner fairings are arranged inside nacelle to embrace central generator and to constrict cold air flow fan duct with circular cross section. Inner fairing with central generator constrict intermediate chamber with circular cross section that embraces central generator and is furnished with at least one rear bore at hot air periphery. Precooler comprises inlet to take in hot air from central generator and outlet of cooled hot air produced using cold air. Precooler is arranged intermediate chamber in thermal contact with rear part of inner fairing to form intermediate channel between precooler and central generator. There is at least one air intake arranged ahead of precooler that passes through inner fairing to bleed cooling air flow from cold flow to at least partially cool hot air flow coming into precooler.

EFFECT: ruling out aerodynamic disturbances in fan duct.

10 cl, 11 dwg