Multistage compressor of gas turbine engine |
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IPC classes for russian patent Multistage compressor of gas turbine engine (RU 2235912):
   
 Axial multistage compressor / 2180054
The invention relates to a compressor engineering, in particular to the design of multistage axial compressor
 The fan / 2174194
The invention relates to ventilyatorostroeniya and can be used as part of systems management products aviation and rocketry
 Screw compressor / 2173406
The invention relates to devices for mixing, pumping and compression of gases, can be used as a portable source of compressed air for aircraft on-Board tool in workshops, garages, workshops
 Axial three-stage compressor / 2136974
The invention relates to a compressor engineering, in particular to multi-stage axial compressors and allows to increase the generated pressure and the efficiency of the compressor
 The axial compressor of a gas turbine engine / 2111385
The invention relates to the field of aircraft engine industry, namely the high pressure compressor turbojet engines, mainly with a large by-pass ratio
 Three-stage multistage compressor / 2097604
The invention relates to a compressor engineering, namely the multi-stage axial compressors
 Vertical pump / 2005919
 Compressor housing (versions) and compressor impeller blade / 2247867
Proposed housing of compressor includes axially convex inner surface located around row of impeller blades with radial clearance between surface and blades. Edges at tip of blades add to housing contour, thus reducing losses on ends at blade tips and blocking of flow.
 Method of increasing efficiency of axial multi-stage compressor operation / 2359160
Invention relates to compressor production and can be used in, for example, gas turbine plants incorporating axial multi-stage compressors. The proposed method comprises injecting water into, at least, two compression stages that allows maximum increase in efficiency of the axial multi-stage compressor along with minimum water flow rate. The above effect is provided for by a mathematical expression to calculate the compressor efficiency that allows for steam content and steam enthalpy in air behind the compressor, and to calculate the optimum amount of water injected into the compressor stages required for the said increase at preliminary stages of compressor operation.
 Birotary srew-type blower / 2367822
Invention relates to aircraft engine production, particularly to aircraft gas turbine engine blowers. Proposed birotary blower consists of two consecutive impellers running in opposite directions. To produce required profiles of both impellers, necessary distributions of blades angles are corrected by algebraic summation of their designed angles. Birotary screw-type blower can be cowled.
 Birotary screw-type blower / 2367823
Invention relates to aircraft engine production, particularly to blowers of aircraft gas turbine engines. Proposed screw-type blower comprises 1st and 2nd turbine wheels arranged one behind the other and furnished with vanes running in disks relative to radial axes. Front edges of the 1st turbine wheel vanes feature a departure from radial rotational axis increasing from the hub to periphery to create the vane sweep forward. 2nd turbine wheel vane front edges feature a departure from radial rotational axis decreasing from the hub to periphery to create the vane sweep forward. Birotary stew-type blower can be cowled.
 Impeller of axial fan / 2422681
Holes are made in the hub located at an angle of 90 degrees to its axis as per the number of blades; pin is fixed in each hole; slot for arrangement and turn of the base is made in each pin; each rotation axis is installed in the holes made in pin on opposite sides of the slot and passes through the slot; springs are made in the form of torsion springs arranged on axes of bases; at that, one free end of each spring is connected to the pin, and the other end is connected to the base, and edges of pins protruding to outer surface of hub are made in the form of cylindrical surface coaxial to hub, the diameter of which is equal to outer diameter of the hub.
 Gas turbine engine birotary screw fan / 2439376
Birotary screw fan comprises structural ring of rear suspension and rear bearing housing. Note here that said structural ring is jointed with rear bearing housing flange wherein
 "maxinio" standard technology of vehicle manufacturing and operation, no-run take-off and landing electric aircraft (versions), lifting device, turbo-rotary engine (versions), multistep compressor, fan cowling, turbo-rotary engine operation method and method of electric aircraft lifting force creation method / 2457153
Set of invention relates to aircraft engineering. In the first version, no-run take-off and landing electric aircraft contains fuselage (1) with pillars (3) and helical fan in cowlings (2), lifting planes and chassis. Turbo-rotary engines (4) have generator units of the first and the second screws connected by electric circuit with electric motors (12,13) of the first (23) and the second (28) screw respectively of screw fan (11) installed on the front upper pillar (10). In the second version, no-run take-off and landing electric aircraft contains fuselage with upper and lateral pillars with fork at the end. In the fork of front upper pillar (10), turbo-rotary engine (4) with generator units having rotors on shafts of screw fan screws is installed, and in forks of front lateral pillars (48), electric motors (50) driven fans (49) are installed. Invention also covers lifting device, turbo-rotary engine, multistep compressor, fan cowling, turbo-rotary engine operation method and lifting force creation method.
 Air bleed device, compressor stage with said device, compressor with said stage, and turbojet with said compressor / 2467209
Proposed device comprises moving wheel with moving vanes and stationary wheel with fixed vanes. It comprises also moving wheel manifold for collection of airflow sucked off moving vanes and stationary wheel manifold to collect airflow forced onto fixed vanes. Moving wheel manifold is located on compressor stage casing outer edge, opposite the moving wheel. Stationary wheel manifold is located above moving wheel manifold.
 Fan for local ventilation of wells / 2509894
Fan includes two basic modules of the first and the second stages adjacent to each other by a connection box-type insert so that each module includes a housing, an electric motor, an impeller installed immediately on the electric motor shaft. Impellers of the first and the second stages are made as per a counter-rotation scheme as all-welded impellers with non-rotational double plate blades of S-shape with a variable along the impeller radius by a geometry calculated by means of a single vortex method as jointly operating without any directing vanes based on minimum acoustic power (noise) of the fan, maximum efficiency, pressure and capacity.
 Water injection system of axial multistage compressor / 2524594
Water injection system of an axial multistage compressor, having tubes and outlet channels, further comprises a fairing, in this case the fairing is located in the area of the front edge of each guide blade of the axial multistage compressor with a possibility to form a slot-type channel. Each tube is positioned in a longitudinal cavity formed in the area of the front edge of the indicated blade, and has holes made by the blade height to provide a uniform vapour flow by the section of an air flow, and outlet channels are made at the front edge of each indicated blade, in this case the slot-type channel, and the outlet channels are designed in each guide blade of the axial multistage compressor with a possibility to provide a nonseparated water flow and air flow. In addition, each tube has a heat-protective material.
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The invention relates to compressors of gas turbine engines for aircraft and ground application. The technical problem on which the invention is directed, is to increase the reliability of the multi-stage compressor by eliminating the imbalance drives installed between the first and last stages. This technical result is achieved by the fact that according to the invention the rotor disks placed between the first and last disks made with the balancing ring tabs on the radius of the transition from the blade to the rim of the CDs made with the balancing projections on the side of the entrance to the compressor, and the remaining disks, starting from n-th to the penultimate made with the balancing tabs on the output side of the compressor, in this case n=(0,4...0,6)z N=(1...4)R; L=(0,8...4)R; h=(0,2 0,05...) R; I=(0,2 0,05...) R, where n is the sequence number of the disk from the side of the entrance to the compressor; z is the number of stages of the compressor; R is the radius of the transition from the blade to the disk rim by balancing the ledge; N is the radial height of the balancing protrusion from the side of the rim of the disk; L is the axial length of the balancing ledge on the side of the road disc 1 - the thickness of the balancing protrusion from the side of the blade disk; h is the thickness of the ball to the compressors of gas turbine engines for aircraft and ground application. Known multi-stage compressor of a gas turbine engine, the rotor of the disc type which consists of a spline shaft and installed it drives the rotor blades, and discs are tied together in the axial direction nuts [1]. The disadvantage of this design is the low reliability due to excessive vibration due to imbalance of the rotor. Closest to the claimed is a multi-stage compressor rotor disc-type, spline shaft which is installed drives the rotor blades and drive the last stage has decompressing maze with an annular radial rib for mounting the balancing weights [2]. The disadvantage of this design is adopted for the prototype, is the increased level of vibrations of the compressor due to the increased imbalance of the disk between the first and last stages of the compressor. The technical problem on which the invention is directed, is to increase the reliability of the multi-stage compressor by eliminating the imbalance drives installed between the first and last steps. The essence of the technical solutions is that in multi-stage compressor of a gas turbine engine, the drive last stage has decompressing maze radial edge for balancing weights, according to the invention the rotor disks placed between the first and last disks made with the balancing ring tabs on the radius of the transition from the blade to the rim of the CDs made with the balancing projections on the side of the entrance to the compressor, and the remaining disks, starting from n-th to the penultimate made with the balancing tabs on the output side of the compressor, with n=(0,4...0,6)z N=(1...4)R; L=(0,8...4)R; h=(0,2 0,05...) R; l=(0,2 0,05...) R, where n is the sequence number of the disk from the side of the entrance to the compressor; z is the number of compressor stages; R is the radius of the transition from the blade to the disk rim by balancing the ledge; N is the radial height of the balancing protrusion from the side of the disk rim; L is the axial length of the balancing protrusion from the side of the blade disk; 1 - the thickness of the balancing protrusion from the side of the blade disk; h - thickness of the balancing protrusion from the side of the rim of the disk. Performing on the rotor disk, is placed between the first and last disks, balancing annular projections on the radius of the transition from the blade to the rim possibly due to the low level voltage in this part of the drive to minimize the weight of the compressor. Balancing the protrusion is located at a maximum distance from the axis of the collet with the balancing of the ledge. Mass balancing of the protrusion may be minimal. The placement of the balancing tabs in some drives, starting with n=0,4...0,6 z and the following, except the last, on the output side of the compressor, and the rest of them, except the first, from the input to the compressor, allows the balancing disks in the rotor of the compressor pairs, i.e., the front and rear sides, starting from the middle of the rotor of the compressor, which facilitates the production of disks on the spline projections of the shaft and balancing. When n<0.4 to z will be difficult to set up and balancing disk on the output side of the compressor, and in the case of n>0,6 z - setting and balancing discs from the input to the compressor. If N<R, the balancing disks will be difficult due to the reduced radial height of the balancing ledge. However, when N>4R unnecessarily increase the weight of the discs and the rotor of the compressor as a whole. When L<0,8 R will be difficult balancing disks due to the reduction of the axial length of the balancing ledge and reduce its weight, but if L>4R - excessive weight gain disks. When h<0,05 R too will decrease the mass of the balancing ledge and will be difficult balancing disk, and if h>0,2 R, will increase weight balansirovochnye complicates the balancing disk and in the case of 1>0,05 R increases the mass of the balancing ledge and weight of the compressor rotor. In Fig.1 shows a longitudinal section of the compressor of the gas turbine engine of Fig.2 shows the element I of Fig.1 in an enlarged view, Fig.3 shows the element II of Fig.1 in an enlarged view, Fig.4 shows the element III of Fig.1 in an enlarged view, Fig.5 shows the element IV of Fig.3 in an enlarged view. Multistage compressor 1 gas turbine engine consists of a stator 2 and rotor 3, on the spline shaft 4 slots 5 installed disks 6, strapped on their hubs 7 front 8 and rear 9 nuts. Between the rims 10 disc 6 mounted operating ring 11 forming the sleeve 12 of the rotor 3 of the compressor 1. The first stage disc 13 is made from the entrance 14 to the compressor with an annular radial rib 15 on which are mounted the balancing weights 16. On the drive last stage 17 is installed decompressing labyrinth 18, which on the output side 19 of the compressor is made annular radial rib 20 for setting the balancing weights 21. On the rest, except the first and last disks, at a radius of 22 value R of the transition from the blade 23 to the rim 10 is made of the balancing ring protrusions 24 of the radial height H of the axial length, L, is to record any disks 6. To ensure the removal and placement of disk 6 when balancing on the spline shaft 4, the portion of the disk, starting with n=0,4...0,6 z from the entrance 14 to the compressor, made with the balancing ledge on the output side of the compressor, and the remaining disks are made with the balancing ledge by the entrance to the compressor. Does this device as follows. When the engine is operating in a multi-stage compressor 1 level of vibration is minimized. Drives the first and last stages are balanced at the expense of the balancing weights 16, 21. Additionally balanced in this part of the disk, starting with n and ending with the penultimate, by removing part of the metal with the balancing of the projections 24. Balancing disks in the rotor of the compressor is thus carried out in pairs, i.e., the front and rear sides, starting from the middle of the rotor of the compressor. This facilitates the formulation of the disk 6 in the slots 5 of the shaft 4 and their balancing. Sources of information 1. C. A. Vunow. Construction and design of aircraft gas turbine engines. - M.: Mashinostroenie, 1981, S. 89, Fig.3.27. 2. RF patent №2176331, F 04 D 17/00, 2000. Claims Multi-stage compressor of a gas turbine engine rotor disk first stolen decompressing maze radial edge for balancing weights, characterized in that the rotor disks placed between the first and last disks made with the balancing ring tabs on the radius of the transition from the blade to the rim of the CDs made with the balancing projections on the side of the entrance to the compressor, and the remaining disks, starting from n-th to the penultimate made with the balancing tabs on the output side of the compressor, with n=(0,4...0,6)z; H=(1...4)R; L=(0,8...4)R; h=(0,2 0,05...) R; I=(0,2 0,05...) R, where n is the sequence number of the disk from the side of the entrance to the compressor; z is the number of compressor stages; R is the radius of the transition from the blade to the disk rim by balancing the ledge; N is the radial height of the balancing protrusion from the side of the disk rim; L is the axial length of the balancing protrusion from the side of the blade disk; 1 - the thickness of the balancing protrusion from the side of the blade disk; h - thickness of the balancing protrusion from the side of the disk rim.
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