JPH08232680A - Combustor, turbine, axial compressor and gas turbine - Google Patents

Abstract

PURPOSE: To increase thermal efficiency of a gas turbine so as to reduce a thermal load, by arranging a water pipe in a multistreak screw shape along in an inner cylinder, and extending a steam injection pipe so as to perform injection directly to an outer side turbine moving blade group. CONSTITUTION: A combustion path 6a is extended in an arbitrary shape including a helical shape. A water pipe 2 is arranged along an internal surface of an inner cylinder 1 into a multiscrew shape by providing a suitable space, thickness and length. An outside water injection pump pressure and many superheated steam controls 61 are calculated to be controlled to supply high pressure water and superheated steam. A steam injection pipe 8 is suitably extended to the downstream, to inject steam 3 from many of the steam injection pipes 8 including injection directly to an outer side turbine moving blade group 1 stage 12. High temperature combustion gas of ending fuel injection complete combustion from a fuel injector 4 is large converted into superheated steam energy, to make an exhaust temperature approach 100 deg.C. In this way, thermal efficiency of a gas turbine is rapidly increased, so that a thermal load can be actively reduced.

Description

Detailed Description of the Invention

[0001]

TECHNICAL FIELD This invention relates to various gas turbines including combustors and turbines for gas turbines and axial compressors.

[0002]

2. Description of the Related Art Since the exhaust temperature of a gas turbine is as high as about 500 ° C., the exhaust heat loss is also very large, and even when it is used for both heat and electricity supply, the maximum output of the generator is due to the characteristics of the gas turbine. Is decreased due to the rise in the outside air temperature, and the thermal efficiency at low load is significantly reduced, which results in a very expensive fuel cost when used at low load, and the turbine inlet temperature is raised to a high temperature. Therefore, it requires an extremely expensive material and manufacturing technology, so it becomes an expensive gas turbine, but there is a steam injection cycle etc.In addition, in a gas turbine for aeronautical use, turbine vanes and axial compressor vanes are partially operated. It is used by converting it to a blade, but the structure becomes very complicated and the strength is insufficient at high speed rotation, and in addition, the structure that makes high speed contact with the outside air and air cooling is small and the manufacturing cost becomes expensive. There are drawbacks. (Japanese Patent Publication 1-33348) (Japanese Patent Publication 5-2)
0571)

[0003]

DISCLOSURE OF THE INVENTION The first object of the present invention is to
The objective is to improve the gas turbine that enables the combined use of heat and electricity or the circulation of waste heat along with the rotational power, and to provide various improved parts. A second object of the present invention is to improve various aviation gas turbines and to provide improved various aviation gas turbine parts.

[0004]

When the present invention is used for the combined supply of heat and electricity, or when it is possible to circulate waste heat together with rotation power, the exhaust temperature can be greatly converted to low-temperature superheated steam energy. To 100 ° C to greatly reduce exhaust heat loss, and to greatly increase the fuel injection amount up to about 3 times, greatly increase the maximum output of the generator to about 3 times, and increase or decrease the output significantly. Even from a normal air-fuel ratio of 60 to a theoretical air-fuel ratio of close to 15, mainly by controlling the fuel injection amount and the superheated steam injection amount, it is possible to increase and decrease the load efficiently and drastically. C
By easily obtaining a large amount of hot water and hot water close to
It becomes possible to circulate the waste heat and supply it optimally as heat. In addition, the mass injected into the turbine greatly increases with the mass of air + mass of superheated steam.
It is possible to dramatically increase the output, and the total thermal efficiency can exceed 90%.

When the present invention is carried out as a gas turbine for various aircraft, the turbine vanes that do not generate output or the axial flow compressor vanes that do not perform compression work are completely abolished, and an outer turbine rotor blade group that produces output or By converting it to a group of outer compressor blades that perform compression work and simplifying the structure and rotating in opposite directions, the rotation speed is divided by half and the input is greatly reduced, and further input The outer compressor blades are fitted to the outer compressor blade group to reduce the total number of cooling fins on the outer circumference, and the outer compressor blade group is completely rotated by contacting with the cold outside air at high speed. The rapid heat transfer cooling and the contact cooling of the compressed air at all stages further reduce the input, and the efficient cooling and compression increases the air density to increase the fuel injection amount and thus the large output. Then Turbine side example and supplies cooling air also to the possible effective cooling due to the low temperature. Further, in order to extend the life of the outer turbine blade group, a large number of cooling fins are provided on the outer periphery of the integral outer turbine shell or the like that is externally fitted and fixed to the outer turbine blade group to make high-speed rotating contact with cold outside air. As a result, the outer turbine blade group is rapidly cooled by heat transfer from the root side of the blades in all stages for efficient cooling and long life.

[0006]

[Operation] A large amount of low temperature superheated steam energy is converted to coexist with the combustion gas to bring the exhaust temperature close to 100 ° C, greatly reduce exhaust heat loss, and reduce condensate loss to a minimum. Using the earth flow part that completes the complete combustion of various combustors for use, the part where the combustion gas is diluted by the downstream air is mainly used only in the inner cylinder and extended to any shape according to the purpose. The water injection pipe is selected from the upstream side of the water pipe to the downstream side on the downstream side of the combustion gas that has completed complete combustion by connecting the water pipe along the inner surface of Injecting superheated steam, including direct injection into the outer turbine blade group 1 stage, and coexisting with high temperature combustion gas to make high temperature superheated steam and eliminate condensate loss, Turbine exhaust temperature is also 1 The exhaust heat loss is greatly reduced by approaching to 00 ° C, and the theoretical air-fuel ratio of 15 is operated at a normal air-fuel ratio of 60 or more. Therefore, it is possible to greatly increase the fuel injection amount by three times or more. Since the mass injected into the air also greatly increases with the mass of air + mass of superheated steam + mass of fuel, the maximum output of the generator also greatly increases by three times or more, and a large increase or decrease in output mainly depends on the fuel injection amount. By controlling the amount of superheated steam injection, it is possible to efficiently increase and decrease the load significantly. Also, if the exhaust gas is cooled with water, most of the exhaust heat amount can be recovered as a large amount of hot water and hot water close to 100 ° C, so most of the exhaust heat amount can be recovered. In addition to being able to circulate, it is possible to supply heat optimally as a total heat efficiency of 90% or more, and the effect of making the turbine production cost comparable to that of a steam turbine occurs.

In order to keep the turbine inlet temperature constant and increase the thermal efficiency, output and pressure ratio of the aviation gas turbine, the smaller the input of the axial compressor, the better, and the larger the pressure ratio of the compressed air and the mass of the fuel. Good, the less cooling air the turbine has, the better the structure is simple, so the turbine stationary blades that do not generate output and the axial flow compressor stationary blades that do not perform compression work are completely abolished, and the outer turbine blades that generate output and The outer compressor blades that perform compression work are largely converted into two or more blade groups, respectively, and the structure is simplified to achieve a small, lightweight and large output, and the rotation speed is 2 minutes by rotating in opposite directions. Each one is divided into 1 to greatly reduce the input,
Further, a large number of cooling fins are provided on the outer periphery of the integral outer compressor body etc. externally fitted and fixed to the outer compressor blade group for further reducing the input, and the outer compressor blade is brought into contact with cold outside air at high speed and low pressure. Rapid heat transfer cooling of all the groups to effectively cool the compressed air in all rows of the outer compressor blades by high-speed contact heat transfer cooling to further reduce the input power and increase the number of stages of multi-stage blades. Since it compresses linearly and efficiently, a large output is possible by increasing the pressure ratio, fuel injection amount, and intake air mass, and the cooling air supplied to the turbine side is also low in temperature, so a small amount is required. A large number of cooling fins are provided on the outer circumference of the outer turbine shell, etc., which is externally fitted and fixed to the blade group. Cooling, so effective cooling There is an effect that can be in the long-life Te.

[0008]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS A first embodiment will be described with reference to FIG. 1, which shows a gas turbine capable of supplying heat and electricity together to circulate exhaust heat. A turbine blade group 10 and a compressor blade group 11 fixed to a normal simple cycle gas turbine main shaft 9 are appropriately provided, and a normal turbine vane group is largely converted into an outer turbine blade group 12 of the present invention. Therefore, the outer turbine moving blade group 12 is radially divided into a plurality of groups in the divided turbine shell 15 or is fixed integrally as a unit, and the outer turbine moving blade group first stage 13 and the outer turbine moving blade group final stage 18 are also discs. A large number of axial directions in which the portions 14a and 14b are provided and integrally fixed to each other and are pivotally supported by the gas turbine main shaft 9 respectively, and the integral outer turbine barrel 16 having a taper is provided outside the split turbine barrel 15 2 along the irregularities 17a and 17b
Since the structure is made to be a heavy structure and is fixed to the outer turbine moving blade group 13 and the outer turbine moving blade group final stage 18 respectively, and all the normal compressor stationary blade groups are also the outer compressor moving blade group 31, the split type compressor body 32 The multistage outer compressor blade group 31 is divided into a plurality of radial directions or fixedly assembled as a unit, and the outer compressor blade group 1 stage 33 and the outer compressor blade group final stage 34 are assembled.
And, if necessary, the intermediate outer compressor blade group 31 is integrally formed with disk portions 14c, 14d, 14e and the like to pivotally support the gas turbine main shaft 9 and to have a taper on the outer side of the split type compressor body 32. Outer compressor blade group 1st stage 33 and outer side, which are externally fitted along a large number of axial irregularities 17c and 17d respectively provided with the integral outer compressor body 35 and are pivotally supported by the gas turbine main shaft 9 as a double structure. Mainly fixing each of them to the compressor blade group final stage 34 to provide the outer compressor blade group final stage 34.
The outer turbine blade group 1 stage 13 is directly connected to the outer gas turbine main shaft 58 to form the combustor 6x (6a, 6b) of the present invention.
An annular injection port 38 for injecting compressed air to 6c) or the like is provided so as to project annularly from the inside and outside of the outer compressor rotor blade final stage 34 toward the combustor 6x side, and is rotatably fitted to the outside thereof. A labyrinth seal 65 is appropriately provided between the intake port 39 and the receiving port 39, and is fixed to the combustor 6x side by the flange 29b, and the outer turbine rotor blade first stage 13 and the like are rotated on the 6x by injection of combustion gas and superheated steam. An annular combustion gas injection port group 28 is fixedly provided by a flange 29a so as to project, and an annular sheath 27 for receiving combustion gas or the like injected from the annular combustion gas injection port group 28 is provided on the outer turbine blade group 1st stage 13 side. Is rotatably fitted to the ring-shaped combustion gas injection port group 28 and provided with a labyrinth seal 65 at its fitting portion.
The gas turbine main shaft 9 and the outer gas turbine main shaft 58 are provided with a counter-rotating gear device 41 or a counter-rotating reduction gear device 51 on the upstream side of the axial compressor 30 or on the downstream side of the turbine 5.
The rotary power generated in the opposite directions are concentrated on the gas turbine main shaft 9 side.

High-efficiency gas turbines have a turbine inlet temperature of about 1000 ° C, so the high-temperature combustion gas is diluted with air so that the air-fuel ratio is 60 or more and the air is excessive. According to the present invention, the high temperature combustion gas temperature is largely converted to the superheated steam mass to bring the exhaust temperature close to 100 ° C, so that the fuel injection amount can be greatly increased. By approaching the theoretical air-fuel ratio of 15, the output is tripled or more. However, since the pressure of water, which is much easier to compress than that of air, is increased to a high pressure, the combustion gas temperature with the minimum input is converted into steam energy with a low temperature, and the turbine output is tripled mainly due to the large increase in the injection mass. For the above reasons, it is small, lightweight, large output, low noise, exhaust heat loss is dramatically reduced, cooling of the turbine is not required, and the structure is simple. Thermal efficiency is dramatically improved by circulating exhaust heat quantity. Expenses are also possible.

With reference to FIGS. 1 and 2, the combustor 6a of the present invention is shown.
And 6b, in order to add a large steam generation portion to various combustors for a gas turbine to make a normal gas turbine into a composite gas turbine, a can-type combustor 6a is used.
Alternatively, the cannula-shaped combustor 6b may be extended to an arbitrary shape including a spiral shape, and a portion that generates heat and uses the upstream portion to be diluted with downstream air may be deleted to remove the inner cylinder 1a. As the only 1b, the water pipes 2 are arranged along the inner surface in the shape of a rectangular round multi-threaded screw or the like at appropriate intervals, thicknesses and lengths,
External water injection pump pressure and multiple superheated steam control valves 61
Is calculated and controlled based on the numerical value of an arbitrary measurement point to supply high-pressure water and superheated steam, and the water pipe 2 is provided at the downstream side of the portion where the complete combustion inside the many inner cylinders 1a and 1b is almost completed. The steam injection pipe 8 selected from the upstream side toward the downstream side is appropriately extended to the downstream side, and the steam 3 such as superheated steam is directly injected into the outer turbine rotor blade group 1 stage 12, and a large number of steam injection pipes 8 are included. From the normal fuel injector 4, the temperature of the high-temperature combustion gas, which has been completely burned by the normal fuel injector 4, is largely converted into superheated steam energy, and the temperature is drastically lowered to reduce the exhaust gas temperature to 1
By approaching the temperature to 00 ° C, it becomes possible to approach the stoichiometric air-fuel ratio combustion, and by greatly increasing the fuel injection amount, the steam mass is greatly increased and the output heat loss is greatly reduced as a large output, and the combustion gas and superheated steam coexist. By doing so, there is no condensate loss, and if necessary, a large amount of hot water and hot water can be easily obtained by cooling the exhaust gas to complete the recycling use of waste heat and the mass injected to the turbine. The speed also greatly increases, but especially by greatly increasing the mass as the mass of the combustion gas + the mass of superheated steam, the thermal efficiency of the gas turbine is dramatically increased and the heat load is dramatically reduced, making it an inexpensive gas turbine. You will also be able to get

1 and 5, the combustor 6C of the present invention is shown.
In order to add a large steam generation part to various combustors for gas turbines to make a normal gas turbine into a composite gas turbine, an annular type combustor 6C is extended to an arbitrary shape, A portion diluted with air downstream is deleted by using the upstream portion that generates A large number of water pipes 2 are arranged in a multi-threaded manner at appropriate intervals, thicknesses and lengths, and an appropriate number of water injection pumps are provided outside the pump pressure and a large number of superheated steam control valves. 61 is calculated and controlled on the basis of the numerical value of an arbitrary measurement point to supply high-pressure water and superheated steam, and the downstream side of the portion inside the inner cylinder 1C where almost complete combustion is completed and the upstream side of the water pipe Select further downstream The steam injection pipe 8 is appropriately extended to the downstream side to directly inject the outer turbine blade group 1 stage 12, and the steam 3 such as superheated steam is injected in accordance with the opening degree of the superheated steam control valve 61. It is possible to approach the stoichiometric air-fuel ratio combustion by drastically lowering the temperature of the high-temperature combustion gas that has been completely burned by the normal fuel injector 4 by injecting from 8 and making the exhaust temperature approach 100 ° C. By increasing the amount of fuel injection, the mass of water vapor is largely increased to produce a large output, and the exhaust heat loss is greatly reduced by circulating the exhaust heat amount by cooling the exhaust gas, so that the combustion gas and the superheated steam coexist. There is no condensate loss, and if necessary, a large amount of hot water and hot water can be easily obtained to facilitate exhaust heat utilization, and the mass and speed injected into the turbine also greatly increase. Burn As a small low-noise light high power high thermal efficiency by a large increase as the mass of gas and superheated steam, the heat load is for the purpose of obtaining a cheap gas turbine by reducing dramatically.

Referring to FIGS. 1, 2, 3, 5, and 7, the case where various combustors are used in combination with the outer compressor rotor blade group final stage 34 of the present invention will be described. Compressor 30
In order to easily secure airtightness when injecting compressed air from the outside to the combustor side, an annular injection port 38 having labyrinth seals 65 on both outer peripheries is formed in an annular shape from the inside and outside in the radial direction of the outer compressor rotor blade final stage 34. Since it projects to the outside, an annular receiving port 39 and a flange 29b, which are rotatably fitted to the outside and whose inner surface forms one side of the labyrinth seal, are provided, and inside the outermost tube 7 of the most upstream of each mallet combustor. The axial flow compressor 30 and various combustors are used in combination by providing a large number of guide vanes 59 and fixing the flange 29b to the outside and combining the outer compressor rotor blade final stage 34 and various combustors.

Referring to FIG. 1, FIG. 2, FIG. 4, FIG. 5, and FIG. 7, the case where various combustors are used in combination with the outer turbine rotor blade first stage 13 of the present invention will be described. On the device side, a ring-shaped combustion gas injection port group 28 for rotating the outer turbine blade group 1 stage 13 and the like by injection of combustion gas and the like is fixedly provided by a flange 29a so as to project.
On the step 13 side, an annular sheath 27 for receiving combustion gas or the like injected from the annular combustion gas injection port group 28 is rotatably fitted to the annular combustion gas injection port group 28 having a labyrinth seal 65 outside. Then, one side of the labyrinth seal 65 is also provided in the fitting part, and the turbine 5 and various combustors are used in combination by combining the outer turbine blade group 1 stage 13 and various combustors.

The upstream portion of the combustor inner cylinder 1d of the present invention will be described with reference to FIG. 6. To cool the high temperature combustion heat to create high temperature superheated air, high speed injection stirring combustion, high speed complete combustion is completed, and therefore compression is performed. The bank 68, which converts high-speed air injected from the machine side into high-pressure air, is made to project a circular shape on the outer circumference of the inner cylinder 1d, etc. An appropriate number of superheated air generation cylinders 66 are provided so as to surround the periphery, and the superheated air injection holes 6 are provided in each of them.
A large number of holes to be injected through 7 to increase the number of holes to be injected toward the center side, and the injection direction is also inclined to the upstream side, so that a large amount of high-temperature heated air is injected, stirred and mixed, and a lean fuel that is difficult to burn Also completes combustion and mixes fuel and air evenly.

Referring to FIGS. 8 and 9, the outer turbine rotor blade group used for the non-cooling application will be explained. To simplify the structure, the normal turbine stator blade group that does not generate an output is completely eliminated and the first stage In order to largely convert all odd-numbered stages from the last stage to the outer turbine rotor blade group 12, a large number of axial irregularities 17a are provided in the axial direction of the annular split turbine shell 15 having the split portions 69b, and the intermediate outer turbine Moving blade group 12
Are divided in the radial direction by dividing portions 69a or are integrally inserted into and fixed between the divided turbine shells 15, and the outer turbine moving blade group first stage 13 and the outer turbine moving blade group final stage 18 are respectively The disk portions 14a and 14b are provided and integrally fixed to each other in the radial direction and the axial direction to pivotally support the gas turbine main shaft 9 including the bearings, respectively, and to burn from the radially inner side of the outer turbine rotor blade first stage 13 A large number of axial irregularities 17a are provided in which a labyrinth seal 65 is provided by projecting the inside of an annular sheath 27 in an annular shape on the container side, and an integral outer turbine shell 16a having a taper is provided outside the split turbine shell 15 in each case.・ By providing a large number of spaces 71 by external fitting along 17b, the heat insulating effect is increased and
As a heavy structure, a labyrinth seal 65 is provided by projecting one of the annular sheaths 27 on the combustor side from the integral outer turbine shell 16a to form the outer turbine blade group 12.

Referring to FIGS. 11 and 12, the outer turbine rotor blade group 12 used for heat transfer cooling by the outside air will be described. The structure is simplified and an output is generated to cool the turbine blade by the outside air. Since the normal turbine stationary blade group is completely abolished and all the odd-numbered stages from the first stage to the final stage are largely converted to the outer turbine rotor blade group, the dividing portion 69b
A plurality of axial irregularities 17a are provided in the axial direction of the annular split turbine shell 15 having a plurality of intermediate turbine outer blade groups 12 in the radial direction to provide a split portion 69a (FIG. 8). The outer turbine moving blade group first stage 13 and the outer turbine moving blade group final stage 18 are integrally provided in the divided turbine shells 15 and fixed to each other with the disk portions 14a and 14b, respectively, in the radial direction and The gas turbine main shaft 9 is axially fixed and pivotally supported by the gas turbine main shaft 9 including bearings, and a cooling air inlet is provided if necessary, and an annular sheath is provided on the combustor side from the radial inner side of the outer turbine rotor blade first stage 13 A labyrinth seal 65 is provided by projecting the inside of 27 in an annular shape, one of the annular sheaths 27 is projected on the combustor side, the labyrinth seal 65 is provided on the inner diameter side, and a large number of cooling fins 60 are provided on the outer circumference. Of the outer turbine blade group 12 and the integral outer turbine barrel 16b by externally fitting the tapered outer turbine barrel 16b having a taper degree along a number of axial irregularities 17a and 17b provided on each. By contacting a wide area, heat is dissipated faster than the cooling fin 60, which is in high-speed contact with the outside air, and the double structure makes the outer turbine blade group easy to assemble.

The outer turbine blade group for obtaining the best outside air cooling will be described with reference to FIGS. 14 and 15.
To simplify the structure and completely abolish the normal turbine vanes that do not generate output to cool the turbine blades by the outside air, and greatly convert all the odd stages from the first stage to the final stage to the outer turbine rotor blade group, The outer turbine rotor blade group 1st stage 13 pivotally supports the cooling fin 60, the annular sheath 27 and the disk portion 14a integrally with the gas turbine main shaft 9 including bearings. The outer turbine blade group final stage 18 is integrally formed with the cooling fin 60 and the cooling fin 60 and the disc portion 14b so as to be pivotally supported including the bearing on the gas turbine main shaft 9, and each of them is appropriately numbered. Tighten with bolts 72 to form the outer turbine rotor blade group.

Referring to FIG. 10, the outer turbine rotor blade group 12
The case where the fan 73 is attached to and used as a fan will be described.

Although it is almost the same as the description of the above, the difference is that a large number of fans 73 are provided on the outer surface of the integral outer turbine shell 16a so as to project downstream to the turbine side. Therefore, it is used for moving a large amount of low pressure air.

Referring to FIG. 13, a case where a prop fan is attached to the outer turbine blade group and is also used as a prop fan will be described.

The difference is almost the same as described above, but the difference is that a large number of prop fans 80 are annularly connected to each other at the downstream side of the cooling fins 60 on the outer surface of the integral outer turbine shell 16b by projecting them together as an integral unit. A cooling fin 60 is installed at the same angle as the prop fan at the place selected. Therefore, it is used for moving a large amount of low-pressure air, and at the same time, effectively cools the turbine blades of the outer turbine blade group 12 by the outside air from the root of the blades.

Referring to FIG. 16, the case where the prop fan 80 is attached to the outer turbine rotor blade group and is also used as a prop fan will be described.

Although it is almost the same as the explanation of the above, the difference is that the prop fan 8
It is a place where 0 is divided into each and integrated into each stage. That is, the outer turbine rotor blade group 1st stage 13 integrally fixes the cooling fin 60, the annular sheath 27 and the disc portion 14a to the latter stage and pivotally supports the gas turbine main shaft 9 including the bearings to support the outer turbine rotor blades. The middle stage of the blade group 12 is integrally fixed to the cooling fin 60 and is fixed to the rear stage by bolts 72, or the divided parts of the prop fan 80 and the cooling fin 60 are integrally fixed to the rear stage so that the outer turbine rotor blade group end. The step 18 is united with the part of the prop fan 80, the cooling fins 60 sheets, and the disc portion 14b which are divided, and is pivotally supported by the gas turbine main shaft 9 including the bearing, and the tip end of the prop fan 80 is connected in an annular shape. It is used for moving a large amount of low-pressure air, and at the same time, effectively cools the turbine blades of the outer turbine blade group 12 by the outside air from the root of the blade.

Referring to FIGS. 17 and 18, the outer compressor rotor blade group 31 used for non-cooling applications will be described.
In order to simplify the structure and greatly reduce the input, we abolished the normal axial flow compressor vanes that do not do much compression work,
All the odd-numbered stages from the first stage to the final stage are largely converted into the outer compressor rotor blade group 31, and a large number of axial irregularities 17c are provided in the axial direction of the ring-shaped split compressor body 32 having the split portions 69b. The outer compressor blade group 31 in the middle is divided by providing the dividing portions 69a in the radial direction, respectively, or is inserted and fixed as a unit between the divided compressor body 32, and the outer compressor blade group 1 stage 33 and the outer side. The compressor blade group final stage 34 and the intermediate selected outer compressor blade group 31 are provided with disc portions 14c and 14d and 14e and 14f, etc., as appropriate for the application, respectively, and the bearings are integrally provided on the gas turbine main shaft 9. It is pivotally supported including the outer compressor blade group final stage 34 in a radial direction on both sides of the combustor, and an annular injection port 38 is provided in an annular shape so that labyrinth seals 65 are provided on both the inner and outer circumferences. Outside of 32 Integral outer compressor cylinder 35 with the path
Many axial irregularities 17c and 17d, each provided with a
A plurality of spaces 71 are provided by fitting along the outer circumference to increase the heat insulation effect, and the outer compressor moving blade group first stage 33 and the outer compressor moving blade group final stage 34 are fixed in the radial direction and the axial direction to form a double structure. By doing so, the outer compressor rotor blade group 31 will be easy to assemble and will have high strength and high speed rotation resistance.

Referring to FIGS. 20 and 21, the outer compressor blade group used for cooling by the outside air will be described. The structure is simplified and the outer compressor blade group 31 is heat-transfer-cooled by the outside air. In order to significantly reduce the input, the normal axial flow compressor stationary blades that do not do much compression work are completely abolished, and all the odd-numbered stages from the first stage to the final stage are largely converted to the outer compressor rotor blade group 31, and the division unit A large number of projections and depressions 17e are provided in the axial direction of the annular split compressor body 32 having 69b, and the middle outer compressor blade group 31 is split by providing split portions 69a (FIG. 17) in the radial direction. The outer compressor blade group 1 stage 33, the outer compressor blade group final stage 34, and the selected outer compressor blade group 31 in the middle are respectively inserted into the disc portion. 14c
And 14d, and 14e and 14f, etc., are provided appropriately according to the purpose, and are integrally pivotally supported together with the gas turbine main shaft 9 including the bearings, and are annular from the radial direction of the outer compressor rotor blade final stage 34 to the combustor side. Nozzle 38 is projected annularly and labyrinth seals 65 are provided on both outer and inner circumferences, and a large number of cooling fins 60 are provided on the outer circumference of the split type compressor body 32 at an appropriate angle according to the purpose. The outer compressor blade group 31 is formed by externally fitting the tapered outer compressor body 35b along a large number of axial projections and depressions 17e and 17f, respectively.
The outer compressor blade group 1 stage 33 and the outer compressor blade group final stage 34 are fixed in the radial and axial directions by facilitating heat conduction by contact with the outer compressor blade group, and all the odd-numbered stages of the axial compressor are exposed to the outside air. High-speed heat transfer cooling reduces the volume of compressed air and divides the rotation speed in half, thereby dramatically reducing the input and forming an outer compressor blade group 31 with high strength and high rotation speed.

The outer compressor blade group for obtaining the best outside air cooling will be described with reference to FIGS. 23 and 24. The structure of the outer compressor blade group is simplified and the outer compressor blade group 3 is controlled by the outside air.
1 heat-transfer cools all the compressed air to cool the compressed air, so the normal axial flow compressor vanes, which are inefficient in compression work, are completely abolished, and all odd-numbered stages from the first stage to the final stage are outside compressor blade groups. The outer compressor blade group 1st stage 33 is integrated with the cooling fin 60 and the disc portion 14g to pivotally support the gas turbine main shaft 9 including the bearings, and the middle portion of the outer compressor blade group 31. The stages are integrated with the cooling fins 60, respectively, and are appropriately fixed to the front stages by bolts 72, and the disc portion 14i is provided in the selected intermediate stage.
Etc. are provided to pivotally support the gas turbine main shaft 9 including the bearings, and annular injection ports 38 are projecting annularly on both sides of the outer compressor moving blade group final stage 34 in the radial direction from the both sides in the combustor side to form inner and outer peripheries. The labyrinth seal 65 is provided so that the disc portion 1 is directed inward in the radial direction.
4h is provided and the gas turbine main shaft 9 is pivotally supported including the bearings and is fixed to the front stage by an appropriate number of bolts 72, and the compressed air is heat-transfer-cooled by the outside air in all the odd stages from the first stage to the final stage. It constitutes the outer compressor rotor blade group 31 of the axial compressor.

Referring to FIG. 19, a case in which a fan 73 is attached to the outer compressor rotor blade group 31 and is also used as a fan will be described.

The explanation is almost the same as the above, but the difference is that a large number of fans 73 are provided on the outer surface of the upstream side of the integral outer compressor body 35a.
It is a place where it is projected as one or more rows. Therefore, it is used for moving a large amount of low pressure air.

Referring to FIG. 22, a case will be described in which the fan 73 is attached to the outer compressor rotor blade group 31 and is also used as a fan.

Although it is almost the same as the description of the above, the difference is that a large number of fans 73 are integrally projected in one or more rows together with the cooling fins 60 on the upstream side outer surface of the integrated outer compressor body 35b. Therefore, it is used for multi-stage heat transfer cooling of compressed air and moving large amounts of low pressure air.

Referring to FIG. 25, description will be given of a case where a fan 73 is attached to the outer compressor rotor blade group 31 and is also used as a fan.

Although it is almost the same as the explanation of the above, the difference is that the fan 73 for integrating each stage from the first stage to the final stage is divided into each and integrated into each stage. That is, in the first stage 33 of the outer compressor blade group, the divided fan 73 and the disk portion 14g are integrally supported to pivotally support the gas turbine main shaft 9 including the bearings. On the downstream side, a fan divided in accordance with the application and a cooling fin 60 in the same direction are integrated and fixed to the upstream stage with bolts 72, and the following can also be fixed according to the application so that more than one row of fans can be mounted Is about to It is also used for multistage heat transfer cooling of compressed air and moving large amounts of low pressure air.

Referring to FIGS. 17, 18, 19, and 27, the cooling air 74 for cooling the turbine side by using the outer compressor rotor blade group 31 used for non-cooling applications will be described.

[0021]

The difference is almost the same as the description of the above, but the difference is that the spaces 71a and 71b of the outer compressor blade group 31 used for the purpose of not mainly cooling the compressed air are used as the cooling air passages 75 and the outer compressor blade group end is used. As shown in FIG. 27, a part of the compressed air ejected from the step 34 is connected to a large number of spaces 71a (71c) radially outward and moved in the space 71a (71c) in an invisible part in the direction of a dotted arrow. Then, the outer compressor blade group 1 stage 33 and the joint portion of the split compressor body 32, etc. are connected to the space 71b (71d) in front and moved in the space 71b (71d) indicated by a solid arrow to the outside. Injecting from the outer side in the radial direction of the compressor rotor blade group final stage 34 to the combustor side,
This is a place where the turbine side is cooled as low-temperature cooling air 74 by the integrated outer compressor body 35a that makes high-speed low-pressure contact with the outside air by high-speed rotation. Therefore, it has a great effect to obtain the low temperature turbine side cooling air 74.

20, FIG. 21, FIG. 22, FIG. 26, and FIG.
The cooling air 74 for cooling the turbine side using the outer compressor blade group 31 used for cooling compressed air will be described with reference to FIG.

[0022]

The explanation is almost the same as that of the above, but the difference is that the space 7 of the outer compressor rotor blade group 31 is used to efficiently cool the compressed air.
A part of the compressed air ejected from the outer compressor rotor blade group final stage 34 is partly shown in FIG. As described above, the plurality of spaces 71c are communicated outward in the radial direction, and an invisible portion is moved in the direction of the dotted arrow in the space 71c, and the outer compressor rotor blade first stage 33
And the space between the space 71d in front of the space by the joint portion of the split type compressor body 32 and the like, and move in the space 71d shown by the solid line arrow to the combustor side from the outside in the radial direction of the outer compressor rotor blade final stage 34 It is a place where the turbine side is cooled as low-temperature cooling air 74 by the cooling fins 60 on the outer periphery of the integral outer compressor body 35b which is injected into the air and makes high-speed low-pressure contact with the outside air at high speed and low pressure.
Therefore, it has a great effect in obtaining the dramatically low temperature turbine side cooling air 74.

Referring to FIGS. 23, 24, 25 and 28, the cooling air 74 for cooling the turbine side by using the outer compressor rotor blades for optimally cooling the compressed air will be described.

[0023]

Although it is almost the same as the explanation of the above, the difference is that the outer compressor blade group 31 is divided from the first stage to the last stage and assembled integrally in order to cool the compressed air efficiently, This is a configuration where 75 is added to the cooling fin 60. That is, as shown in FIG. 28, a part of the compressed air ejected from the final stage 24 of the outer compressor blade group is connected to the plurality of cooling air passages 75a in the radially outward direction, and the inside of the cooling air passages 75a is indicated by the solid arrow direction. To the cooling air passage 7 at the joint portion of the outer compressor moving blade group 1st stage 33 and the same 3rd stage.
5b, move in the cooling air passage 75b in the invisible part in the direction of the dotted arrow, and eject from the radial direction outside of the outer compressor blade group final stage 34 to the combustor side for high speed rotation. This makes it possible to obtain very low temperature cooling air 74 on the turbine side by passing through a large number of cooling fins 60 that are in high speed and low pressure contact with the outside air. Therefore, it has a great effect to obtain the cooling air 74 on the turbine side with a dramatically low temperature.

An example of blade cooling on the turbine side by the cooling air 74 will be described with reference to FIG. 27. A cooling air passage 75 is provided in the outer compressor rotor blade group 31 to recool a part of the compressed air to cool the cooling air 74. In such a case, the cooling air 74 is passed from the cooling air passage 75 of the outer compressor rotor blade group 31 through the annular receiving port 39 to the annular air reservoir 7 near the uppermost stream of the combustor 6.
7a into the annular air reservoir 77b in the vicinity of the most downstream side of the combustor 6, and a plurality of combustion gas injection port groups 28 inside the flush 2
9 is supplied to the cooling air passage 75 from the outer side in the radial direction to perform cooling in accordance with the application in the same manner as a normal nozzle, and is connected to the cooling air inlet 78a on the inner side in the radial direction to form the outer turbine blade group 1 The cooling of the stage 13 is a large blade for transmitting the power of the outer turbine rotor blade group. During operation, the cooling fin 60 on the outer periphery is brought into high-speed low-pressure contact with the outside air to bring the outer turbine rotor blade group from the outer diameter side to high-speed heat transfer. Including the annular sheath 27 and the integral joint 76, which have been cooled and have a greatly expanded cooling area, they are greatly curved to ensure heat resistance and to be expandable and contractible, and a combination of different materials or integrally cast rotor blades. The cooling air 74 is supplied to the internal cooling air passage 75 to cool the moving blades mainly from the inner diameter side and further from the inner diameter side. Therefore, the cooling air inlet 78b for cooling the turbine moving blade group 10 side is substantially axially arranged. Provided a number of outer turbine Dotsubasagun one stage provided to the outer turbine Dotsubasagun 1 stage 13 to achieve savings balance air is provided in an annular radially inwardly integral joint 76 a labyrinth seal 65.

The double reversing gear device 41 will be described with reference to FIGS. 1 and 29. When the double reverse gear device 41 is provided, the gas turbine main shaft 9 that rotates in mutually opposite directions when the outer turbine rotor blade group 12 is provided.
Since there are two main shafts, the outer gas turbine main shaft 58, a gear unit is required to extract the power of the two shafts from the one shaft.
Therefore, in order to obtain the double reversing gear device 41, the outer compressor blade group side 47 on the upstream side of the outer compressor blade group 31 of the outer gas turbine main shaft 58 or the downstream side of the outer turbine blade group 12 on the outer gas turbine main shaft 58. A plurality of main body side support shafts having an internal gear 40 on the turbine rotor blade group side 48, a main shaft side main drive gear 42 fixed to a gas turbine main shaft side 49, and rotatably supported on a gas turbine main body 43. The main body side support shaft 44 is engaged with the plurality of first driven small gears 45 fixed to the main shaft 44.
The first main driving small gear 46 fixed to the other end of the above is meshed with the inner gear 40 provided on the outer compressor moving blade group side 47 or the outer turbine moving blade group side 48 to form the outer compressor moving blade group 31 or the outer side. As the double reversing gear device 41 in which the turbine rotor blade group 12 and the gas turbine main shaft 9 rotate in mutually opposite directions, the shaft outputs of the inner side and the outer side are combined by accurately performing double reversal at a selected rotation ratio. The rotational power of is also taken mainly from the inner side of the gas turbine main shaft 9.

The double inversion reduction gear device 51 will be described with reference to FIGS. 1 and 30. The outer turbine rotor blade group 12
Since the two axes of the gas turbine main shaft 9 and the outer gas turbine main shaft 58, which rotate in opposite directions, are rotated at a high speed, the double reversing reduction gear is required to decelerate and extract the power of the two shafts from the one shaft. Equipment 51 etc. are required. Therefore, the planetary reduction gear units 50 and 50 are combined from both sides of the double reversal gear unit 41 to form a double reversal reduction gear unit 51. That is, the plurality of planetary gears 5 are provided with the sun gears 52a and 52b at the outer end of the outer compressor blade group 31 or the outer turbine blade group 12 and the end of the gas turbine main shaft 9, respectively.
3a and 53b are meshed with each other to form the planetary gear 53a.
.53b meshes with the internal gears 54a and 54b,
Planetary gear support shafts 55a5 of the plurality of planetary gears 53a and 53b
On the right side of the circular plate 56a which pivotally supports 5b from both sides, a cylindrical outer compressor rotor blade group side 47 or outer turbine rotor blade group side 4
8, an internal gear 40 is provided, and similarly, a main spindle side driving large gear 42 is fixed to the left side of the circular plate 56b as a gas turbine main spindle side 49, and a plurality of first driven small gears 45 are formed on the main spindle side driving large gear 42. A plurality of first shafts fixed to the other end of the main body side support shaft 44 pivotally supported by the gas turbine main body 43
The main driving small gear 46 is meshed with the internal gear 40 to connect the outer compressor moving blade group side 47 or the outer turbine moving blade group side 48 and the gas turbine main shaft side 49, and to connect the power shaft 57 to the right center of the circular plate 56b. By projecting, the power of the two shafts is decelerated and taken out from this power shaft 57.

Referring to FIG. 7, a case where the outer turbine rotor blade group 12 is used in combination with a normal gas turbine will be described. In the case where the double reversing gear device 41 and the double reversing reduction gear device 51 are used. In addition, when the gear device is configured to be separable by a horizontal joint, the outer turbine rotor blade group 12 is rotated in the opposite direction of the gas turbine main shaft 9 at a selected rotation ratio. The first driven small gear 22b fixed to the other end of the support shaft 21b pivotally supported by the gas turbine main body 43 is equipped with the outer driven main gear 19 and meshes with the first driven small gear 20b. It meshes with the second driven small gear 24b fixed to the second support shaft 23b pivotally supported by the turbine body 43,
The other second drive small gear 25 fixed to the second support shaft 23b
B is meshed with the driven large gear 26 on the main shaft side so that the normal turbine rotor blade group 10 and the outer turbine rotor blade group 12 rotate in opposite directions to each other at a selected rotation ratio, and are combined with a normal gas turbine. .

Referring to FIG. 7, a case where the outer compressor blade group 31 is used in combination with a simple cycle gas turbine will be described. When the gear unit is configured to be separable by a horizontal joint and combined, the outer compressor operation is performed. The gas turbine main shaft 9 on the downstream side of the blade group 31 is attached to the outer compressor main driving gear 3
6 is fixed and meshed with the first driven small gear 20a, and the two-stage first driving small gears 22a and 22c fixed to the other end of the support shaft 21a pivotally supported by the gas turbine main body 43 Two-stage second driven small gears 24a, 24c slidably fitted onto a spline support shaft 23a pivotally supported by the turbine body 43.
To the outer compressor driven large gear 37 fixed to the outer compressor rotor blade group 31. The other second main driving small gear 25a fixed to the spline support shaft 23a is meshed with the outer compressor driven large gear 37 fixed to the outer compressor rotor blade group 31 The normal compressor blade group 11 and the outer compressor blade group 31
Is selected from two rotation ratios in opposite directions and rotated to combine with a normal simple cycle gas turbine.

12, FIG. 15, FIG. 21, FIG. 24, FIG.
The case where the outer turbine rotor blade group 12 and the outer compressor rotor blade group 31 are used in a turbojet engine will be described with reference to the second embodiment shown in FIGS. 27, 28, and 31. Therefore, the outer turbine rotor blade group 1
Both 2 and the outer compressor blade group 31 are equipped with the cooling fins 60 shown in FIGS. 12, 15, 21, 24, and 31 which perform heat transfer cooling by the outside air, and further select as necessary. The configuration for recooling the turbine-side cooling air described in FIGS. 26, 27, and 28 by the outside air is combined with the configuration in FIG. 31, and the outer compressor blade group 31 and the outer turbine blade group 12 are connected to the outer gas turbine main shaft 58. And is rotatably fitted to the turbine moving blade group 10, the compressor moving blade group 11 and the gas turbine main shaft 9 of the normal simple cycle gas turbine, and the normal combustor is appropriately pivoted to the gas turbine main body 43. In addition, a turbojet engine that is compact, lightweight, and has high output and high efficiency by enabling the relative speed between moving blades to be twice the conventional relative speed between moving blades and stationary blades.

12, FIG. 15, FIG. 22, FIG. 25, and FIG.
28. The case of using the outer turbine rotor blade group 12 and the outer compressor rotor blade group 31 in a turbofan engine will be described with reference to FIGS. 28 and 32. Both 12 and the outer compressor blade group 31 are provided with cooling fins 60 for heat transfer cooling by the outside air as shown in FIGS. 12 and 15, and the outer compressor blade group 31 has a large number of fans 73 on the upstream side. The configuration shown in FIG. 32 is provided by selecting the provided FIG. 22 and FIG. 25, etc., and further recooling the turbine side cooling air shown in FIG. 27 and FIG. In combination with the configuration, the outer turbine moving blade group 12 and the outer compressor moving blade group 31 are connected by the outer gas turbine main shaft 58, and the gas turbine moving blade group 10 and the pressure of the normal simple cycle gas turbine are combined. It is rotatably fitted over the moving blade group 11 and the gas turbine main shaft 9, and is equipped with an ordinary combustor as appropriate to pivotally support the two shafts with various bearings and the like. It is a turbofan engine that is pivotally supported by 43 to make it possible to make the relative speed between the moving blades close to twice the conventional relative speed between the moving blade and the stationary blade, making it easy to select the rotational speed.

FIG. 13, FIG. 16, FIG. 21, FIG. 24, FIG.
When the case where the outer turbine rotor blade group 12 and the outer compressor rotor blade group 31 are used in a prop fan engine will be described with reference to FIGS. 27 and 28, the outer turbine rotor blade group will be small, lightweight, and have high output and high efficiency. 12 is shown in FIG.
Cooling fin 60 for cooling the heat transfer blades by the outside air as described in 6.
And a fan 73 are selected, and the outer compressor blade group 31 is also cooled by heat transfer blades by high-speed contact between the outside air and the cooling fins 60, and compressed air is subjected to all-stage heat transfer cooling of odd-numbered blade stages. 21 and FIG. 24, the outer compressor blade group 31 described in FIG. 24 is selected, and as shown in FIGS.
-Combining the configuration for re-cooling the compressed air for cooling the turbine side described with reference to FIG. 28 with the outside air, connecting the outer turbine moving blade group 12 and the outer compressor moving blade group 31 with the outer gas turbine main shaft 58, It is rotatably externally fitted to the turbine moving blade group 10, the compressor moving blade group 11 and the gas turbine main shaft 9 of the simple cycle gas turbine, and is equipped with a normal combustor or the like as appropriate to pivot the two shafts with various bearings or the like. In addition to supporting, the two shafts are appropriately pivoted to the gas turbine body 43 so that the relative speed between the moving blades is 2 times that of the conventional relative speed between the moving blade and the stationary blade.
A gas turbine that enables the speed to be doubled and makes it easy to select the rotation speed.

A case in which the outer turbine rotor blade group 12, the outer compressor rotor blade group 31, and the outer gas turbine main shaft 58 are used in a double-reversal turbofan engine will be described with reference to FIGS. 32 and 33.

Although it is almost the same as the explanation of the above, the difference is that it is a double reversal turbofan engine in which the fan 73b is also provided at the most upstream side of the gas turbine main shaft 9 and the fans 73a and 73b rotate in opposite directions.

With reference to FIGS. 31 and 34, the outer turbine rotor blade group 12 and the outer compressor rotor blade group 31 are connected by the outer gas turbine main shaft 58, and the turbine rotor blade group 10 and the compression of the ordinary simple cycle gas turbine are compressed. A double-reversal prop fan engine, which is rotatably fitted onto the moving blade group 11 and the gas turbine main shaft 9 and includes a normal combustor, will be described.

Although the explanation is almost the same as that of 1., the difference is that the disk portion 14g of the outer compressor rotor blade first stage 31 is extended to the upstream side and equipped with the prop fan 80a, and the rear portion thereof is pivoted to the gas turbine main body 43. While supporting the appropriate portion of the arm portion 81 as a stationary blade, the gas turbine main shaft 9 is also extended to the most upstream side, and the prop fan 80b is provided upstream of the prop fan 80a.
Each tip is projected in a reverse L-shape from the front surface in the rotational direction to the downstream side, and the airflow is compressed and deflected inward in the radial direction to suppress the generation of shock waves and rotate in opposite directions. Double reversal prop fan 8
It is about 0a and 80b.

The combined use of the ramjet propulsion device will be described with reference to FIG. 31. The outer turbine moving blade group 12 and the outer compressor moving blade group 31 are connected by the outer gas turbine main shaft 58 to form a turbine of a normal simple cycle gas turbine. A gas turbine equipped with a normal combustor and the like, which is rotatably fitted over the rotor blade group 10, the compressor rotor blade group 11, and the gas turbine main shaft 9, and is provided with a ram combustor 83 with the gas turbine main shaft 9 being thicker. The portion of the ram fuel injector 8 to which the ram combustor 83 is appropriately fixed to the portion is expanded.
It is equipped with 2 etc. and uses the Ramjet propulsion device together.

The adiabatic uncooled theoretical air-fuel ratio combustion gas turbine will be described with reference to FIGS. 1, 2, 3, 4, 5, 5, 8, 9, 17, and 18. The low-temperature superheated steam will be described below. In order to convert the energy into a large amount of energy and perform adiabatic uncooled theoretical air-fuel ratio combustion, the combustor is a combustor 6a, 6 shown in FIG. 1, FIG. 2, FIG.
b, 6c, etc., and using FIG. 1, FIG. 2, FIG. 3, FIG.
Including the connection method shown in, the outer turbine blade group 12 and the outer compressor blade group 31 side are contacted. As the rotor blade group 12, the outer turbine rotor blade group 12 used for non-cooling purposes shown in FIGS. 8 and 9 is used, and all the generated heat is largely converted to low-temperature superheated steam energy. 31 is also FIG.
・ The outer compressor blade group 31 used for non-cooling applications shown in Fig. 18 is used to approach the adiabatic uncooled theoretical air-fuel ratio combustion.

The above

Since the gas turbine described in (1) is used as an ultra-compact gas turbine generator for loading an electric vehicle, a large number of magnetic bearings and the like are used to enable high-speed rotation and downstream of the gas turbine main shaft 9 and the outer turbine rotor blade group 12. Alternatively, a publicly known generator is provided upstream of the outer compressor rotor blade group 31, and a well known heat exchange condenser is provided for cooling exhaust gas from the outer turbine rotor blade group final stage 12 to circulate and use waste heat. And secure hot water and make-up water for in-combustor injection.

The above

Since the gas turbine described in (1) is used for both heat and electricity supply, the double reversing gear device 41 is provided downstream of the gas turbine main shaft 9 and the outer turbine rotor blade group 12 or upstream of the outer compressor rotor blade group 31. The double reversing reduction gear device 51 is configured to communicate with a load such as a generator, and also includes a known heat exchange condenser for cooling exhaust gas from the outer turbine rotor blade group 12 to circulate exhaust heat. Then, by supplying electricity and obtaining a large amount of hot water and hot water close to 100 ° C, the waste heat is circulated and supplied as heat.

The above

In order to use the gas turbine described in (1) as a gas turbine for propulsion of an ultra-high speed ship, a double reversing reduction gear unit 51 is constructed downstream of the gas turbine main shaft 9 and the outer turbine rotor blade group 12 to form a single shaft. Known heat exchange condensate for converting exhaust gas injected from the outer turbine rotor blade group final stage 12 and circulating it as condensate by converting the exhaust gas to a downstream water jet propulsion machine or the like. It is equipped with a vessel to secure hot water to be injected into the combustor, its makeup water, and so on.

The above

In order to use the gas turbine described in (1) as a gas turbine for propulsion of an ultrahigh-speed ship, ordinary reduction gear units are provided downstream of the gas turbine main shaft 9 and the outer turbine rotor blade group 12, respectively. Combustion is provided with a known heat exchange condenser for collecting and cooling the exhaust gas injected from the outer turbine rotor blade final stage 12 and circulating it as condensate, each of which is connected to a heavy-duty reverse jet propulsion machine or the like. Secure hot water to be sprayed into the container and its supplementary water, etc.

The adiabatic uncooled theoretical air-fuel ratio combustion gas turbine will be described with reference to FIGS. 1, 2, 3, 4, 5, 5, 8, 9, 17, 17, and 33. In order to obtain a rotary power and compressed air at the same time as adiabatic uncooled theoretical air-fuel ratio combustion by converting it to superheated steam energy
The combustors are the combustors 6a, 6b, 6 shown in FIGS. 1, 2, and 5.
Using c, etc., including the connection method shown in Fig. 1, Fig. 2, Fig. 3, Fig. 4, and Fig. 5, it is connected to the outer turbine rotor blade group 12 and outer compressor rotor blade group 31 side. If it is converted into energy, there is no condensate loss and cooling is strictly prohibited. Therefore, the outer turbine blade group 12 shown in FIGS. 8 and 9 is the outer turbine blade group 12 used for non-cooling applications. At the same time, the outer compressor rotor blade group 31 largely converts all the generated heat into low-temperature superheated steam energy.
-A diagram for simultaneously obtaining rotational power and compressed air as a gas turbine for propulsion of an ultrahigh-speed ship by connecting the outer compressor blade group 31 used for non-cooling use shown in FIG. 19 with the outer gas turbine main shaft 58 A fan 73b shown in FIG. 33 is provided upstream of the fan 73 shown in FIG. 19 so as to be rotatably fitted over the simple cycle gas turbine main shaft 9, the turbine rotor blade group 10 and the compressor rotor blade group 11 to form a gas turbine main shaft. 9 and the outer turbine rotor blade group 12 are connected to the water jet propulsion machine and the like via a double reversing reduction gear device 51 and the like to collect and cool exhaust gas injected from the outer turbine rotor blade group final stage 12. It is a gas turbine for propulsion of ultra-high-speed vessels, which is equipped with a well-known heat exchange condenser to circulate it as condensate and secures the hot water injected into the combustor and its supplementary water.

[0047]

[Effects of the Invention] Due to the characteristics of the gas turbine, the output of the generator is affected by the outside air temperature, and the maximum output of the generator decreases due to the rise of the outside air temperature, but if it is converted to low-temperature steam energy, the turbine inlet temperature Becomes closer to the steam turbine inlet temperature, and a large increase in the fuel injection amount is possible. Therefore, by generating a gas turbine with a super-lean combustion ratio of 60 or more due to the characteristics of the gas turbine, it is possible to generate power by approaching the theoretical air-fuel ratio to the 15 side. It is also possible to triple the maximum output of the machine, and therefore, it is possible to dramatically increase the range in which the output can be increased or decreased by air-fuel ratio control, that is, fuel control. That is, by incorporating a water tube boiler in the combustor, the above effect is generated, and the mass injected to the turbine side is greatly increased to the mass of air + mass of superheated steam + mass of fuel. Yes, by cooling the exhaust water with water, most of the exhaust heat can be easily recovered as a large amount of hot water and hot water close to 100 ° C, so most of the exhaust heat can be circulated and optimally used as heat. This has the great effect of dramatically increasing the overall thermal efficiency.

When a normal axial flow compressor vane is completely abolished and the outer compressor blade group 31 for compression work is converted into a large number, the number of blade stages for compression work is greatly increased by more than double. If the relative speed between the blades is the conventional speed between the rotor blade and the stator blade, the rotational speed between the rotor blades will be reduced in half, and it will be possible to bring the input and centrifugal forces close to one-quarter side. By combining the outer compressor rotor blade group 31 and various combustors including the combustor of the present invention with the outer compressor rotor blade group final stage 31 for use in a normal gas turbine, the pressure ratio of the gas turbine can be efficiently increased. It has a great effect.

In addition to abolishing a normal gas turbine stationary blade that does not generate output to a large conversion to an outer turbine rotor blade group 12 that generates output, the number of blade stages that generate output greatly increases by more than double. , If the relative speed between the moving blades is set to the conventional speed between the moving blade and the stationary blade, the rotating speed of each moving blade can be decelerated by half, so the centrifugal force can be reduced by 1/4 and the weight can be reduced. When the respective speeds are set to the conventional speeds, the relative speed between the blades can be doubled and the pressure ratio of the gas turbine can be dramatically increased, which has the great effect of simplifying the structure, reducing the size and weight, and increasing the output. By using the group 12 and various combustors including the combustor of the present proposal in combination with the outer turbine blade group 1 stage 12 for use in a normal gas turbine, it is possible to obtain various gas turbines having a simple structure, small size, lightweight and large output. There is.

The most upstream edge of the inner cylinder of the combustor is appropriately projected to convert high-speed air into high-pressure air, and the cup-shaped superheated air generating cylinder 66 is provided so as to surround the periphery of the normal fuel injector 4. A large number of superheated air injection holes 67 are provided in each of them, the number of holes for injecting toward the center is increased, and the injection direction is inclined toward the upstream side. Since the reverse injection is performed during the mixed combustion to carry out the agitation mixed combustion, there is a great effect that the lean fuel, which is difficult to burn, can be easily completed completely burned.

The double structure of the split turbine shell 15 and the integral outer turbine shell 16 has the effect of being easy to assemble and being lightweight and strong in response to a large centrifugal force. 13 and the outermost turbine blade group 18 can be pivotally supported on the gas turbine main shaft 9 at the most upstream and most downstream of the outer turbine blade group 18, which also has the effect of suppressing vibration.

Since the normal turbine stationary blades are completely abolished and all the odd-numbered stages are largely converted to the outer turbine rotor blade group 12, the outer peripheral portion is brought into contact with the outside air at a high speed during operation, and the outside air is cooled at a high speed, a low pressure and a low temperature. If a large number of cooling fins 60 are provided in order to dramatically increase the temperature, the cooling speed will be further increased and the outer turbine blade group 12 will be subjected to high-speed heat transfer cooling from the outer peripheral side, which has a great effect of saving cooling air.

Since the normal turbine stationary blades are completely abolished and all odd-numbered stages are largely converted into the outer turbine moving blade group 12, the outer peripheral portion is brought into contact with the outside air at a high speed during operation, and the outside air is cooled at a high speed, a low pressure and a low temperature. Since the cooling fin 60 is integrally formed with the outer turbine rotor blade group 12 at each stage, the heat transfer cooling rate is further increased, and the outer turbine rotor blade group 12 is further increased from the outer peripheral side. High-speed heat-transfer cooling produces a great effect of further reducing cooling air.

The outer turbine rotor blade group 12 pivotally supports the first and last stages on the gas turbine main shaft 9, and the outer turbine rotor blade group 1 stage 13 is connected to the outer gas turbine main shaft 58 for power transmission. Therefore, it is preferable to use a large blade, and the high-speed gas mass for obtaining the rotational power tends to collect on the outer turbine rotor blade group 12 side due to centrifugal force. The effect of being able to design easily is by sticking etc.

Even when the outer turbine blade group 12 is used for aviation, the first and last stages are pivotally supported on the gas turbine main shaft 9, and the outer turbine blade group 1 stage 13 is connected to the outer gas turbine main shaft 58. Since it has a role of power transmission such as a large blade, it is preferable to use a large-sized blade. Even when a prop fan 80 is provided as a load, a high-speed gas mass for obtaining rotational power is easily collected on the outer turbine rotor blade group 12 side by centrifugal force. Easy to design outer turbine blade group 12 with prop fan 8
By fixing 0 etc. and connecting it in a ring shape at the tip, in addition to the effect of reducing the thickness of the blade, it also has the effect of suppressing the shock check wave.

The outer turbine rotor blade group 12 is assembled by integrating the rotor blade stage and the cooling fin 60, the rotor blade stage, the cooling fin 60, the prop fan 80, etc., in each stage so that the outside air and the high speed are high during operation. When contacting and radiating heat, heat is transferred without boundaries, which has the great effect of heat-transfer cooling the outer turbine blade group 12 more efficiently than the root portion of the outer periphery.

When the normal axial flow compressor vanes are completely abolished and all odd-numbered stages from the first stage to the final stage are converted into the outer compressor rotor vane group 31, the number of rotor stages performing compression work increases more than double. In addition to the above, when the assembly structure is a double structure of the split type compressor body 32 and the integrated type outer compressor body 35a, the outer compressor blade group 12 that is lightweight, easy to assemble, high strength and capable of high speed rotation is obtained. It has a great effect.

When all the normal axial flow compressor vanes are completely abolished and all the odd stages from the first stage to the final stage are converted into the outer compressor rotor blade group 31, the number of rotor stages for compression work is increased to more than double. In addition to this, since the integrated outer compressor body 35b comes into contact with the outside air at high speed during operation, a large number of cooling fins 60 are provided on the outer periphery of the integrated outer compressor body 35b. There is a great effect that all 30 odd stages can be cooled by high-speed heat transfer by outside air.

When the normal axial flow compressor vanes are completely abolished and all odd-numbered stages from the first stage to the final stage are largely converted into the outer compressor rotor blade group 31, the number of rotor stages performing compression work is greatly increased to more than double. In addition, since the outer compressor rotor blade group 31 is formed integrally with the cooling fins 60 in each stage, the cooling fins 60 come into contact with the outside air at a high speed during operation, so that the heat of the rotor blades does not have a boundary. Since the heat is transferred to the cooling fins 60 side, a large effect is produced by the large number of cooling fins 60, which enables efficient and high-speed heat transfer cooling of all odd-numbered stages of the axial flow compressor 30 by the outside air.

When all the normal axial flow compressor vanes are completely abolished and all the odd stages from the first stage to the final stage are converted into the outer compressor rotor blade group 31, the number of rotor stages for compression work is increased to more than double. In addition to this, since the integrated outer compressor body 35a rotates at high speed during operation, a large number of low-pressure air is generated by providing a large number of fans 73 on the outer peripheral surface of the integrated outer compressor body 35a on the upstream side. It has the effect that it can be used as a group of outer compressor blades for moving.

When the normal axial flow compressor vanes are completely abolished and all the odd stages from the first stage to the final stage are largely converted to the outer compressor rotor blade group 31, the number of rotor stages performing compression work is greatly increased to more than double. In addition, since the integrated outer compressor body 35b rotates at high speed during operation, a large number of fans 73 are provided on the upstream outer peripheral surface of the integrated outer compressor body 35b, and a large number of them are formed on the entire downstream outer surface. Since the cooling fin 60 of the above is projected, if it is used in combination with an aeronautical gas turbine as appropriate, it has the effect of being able to use all-stage heat transfer cooling of compressed air as well as the effect of being used as a fan of a turbofan engine.

When the normal axial flow compressor vanes are completely abolished and all odd-numbered stages from the first stage to the final stage are converted into the outer compressor rotor blade group 31, the number of rotor stages performing compression work is greatly increased to more than double. In addition, since the outer compressor moving blade group 31 is integrally formed with the cooling fin 60 and the fan 73 for each moving blade stage, the cooling fin 60 and the like contact the outside air at high speed during operation, In order to dissipate the amount of heat that is transferred from the boundary without any boundaries, when used in combination with an aviation gas turbine as appropriate, the compressed air can be efficiently transferred and cooled by the outer compressor rotor blade group 31 in all stages, and in addition, the turbofan There is also the effect that it can be used efficiently as an engine.

When the normal axial flow compressor vanes are completely abolished and all odd-numbered stages from the first stage to the final stage are largely converted into the outer compressor rotor blade group 31, the number of rotor stages performing compression work is greatly increased to more than double. In addition to the above, it is possible to obtain the cooling air 74 for cooling the turbine side even if the outer compressor blade group 31 used for the application where the cooling of the compressed air is not used is used in combination with various gas turbines. There is an effect that can be.

If the normal axial flow compressor vanes are completely abolished and all the odd stages from the first stage to the final stage are converted into the outer compressor rotor blade group 21, the number of rotor stages performing compression work is increased to more than double. In addition to this, if a cooling fin 60 is provided on the entire outer peripheral surface of the outer compressor rotor blade group 31, the compressed air is heat-transfer-cooled at all odd-numbered rotor blade stages to significantly reduce the input and to produce high-density compressed air. However, if the cooling air on the turbine side is re-cooled by recirculating the cooling fin 60, the turbine side can be effectively cooled, and there is a great effect that the cooling air can be saved and the thermal efficiency can be increased.

When the stationary blades of the normal axial flow compressor are completely abolished and all odd-numbered stages from the first stage to the final stage are largely converted into the outer compressor rotor blade group 31, the number of rotor blade stages performing compression work is increased more than double. In addition, in order to efficiently cool the compressed air, the outer compressor blade group 31 is integrally formed and assembled for each stage from the first stage to the final stage, so that it becomes low temperature compressed air, and the compressed air is Since it recirculates in the cooling air passage 75 provided in the cooling fin 60, it has a great effect of obtaining the significantly cooled turbine side compressed cooling air 74.

The outer turbine rotor blade group 1 stage 13 is connected to the annular sheath 2
The cooling area is expanded by 7 and is cooled from the outer diameter side by the outside air through the cooling fin 60, and the cooling area is expanded from the inner diameter side by the annular sheath 27 and the integral joint 76, and the cooling air is supplied from the cooling air inlet 78a. Since 74 is supplied to cool the blade mainly from the inner diameter side, there is a great effect that the outer turbine moving blade group 1st stage 13 can be cooled.

In the present invention, the normal turbine stationary blades that do not generate output are completely abolished and the odd-numbered stages are largely converted to the outer turbine moving blade group 12 that generates output. The double reversing gear device 41 has a great effect, and in the present invention, since the normal axial flow compressor vanes are completely abolished and all the odd stages from the first stage to the final stage are largely converted to the outer compressor rotor blade group 31, The double reversing gearbox 41 is also very effective for driving the compressor blade group 31.

In the present invention, the normal turbine stationary blades that do not generate output are completely abolished, and the odd-numbered stages are largely converted to the outer turbine moving blade group 12 that generates output. Therefore, in order to obtain rotational power that is decelerated in one direction. The double inversion reduction gear device 51 has a great effect, and in the present invention, the normal axial flow compressor vanes are completely abolished and all the odd stages from the first stage to the final stage are converted into the outer compressor rotor blade group 31. Therefore, the double reversing reduction gear device 51 has a great effect in driving the outer compressor rotor blade group 31 and also in obtaining the rotational power reduced.

When a normal turbine stationary blade that does not generate an output is completely abolished and all odd-numbered stages are converted into an outer turbine rotor blade group 12 that generates an output and used in combination with a normal gas turbine or in a normal axial flow When the stationary vanes of the compressor are completely abolished and all the odd stages from the first stage to the final stage are largely converted into the outer compressor rotor blade group 31 and used in combination with a normal gas turbine, the double reversing gear device or the double reversing gear unit is used. Since the reduction gear unit is configured to be separable, it has a great effect mainly for use in combination with a simple cycle gas turbine.

A normal gas turbine stationary blade that does not generate output is completely abolished, and all odd-numbered stages are largely converted to a group of outer turbine blades 12 for external heat transfer cooling that generate output and a normal axial compressor. When the stationary blades are completely abolished and the outer compressor blade group 31 that performs compression and cooling work in all stages at odd stages from the first stage to the final stage is largely converted and used as a turbojet engine, the blade stages that generate output are doubled. Since the jet gas flow path resistance is drastically reduced as well as the above, there is a great effect that a large thrust can be obtained by a high jet gas velocity, and the turbine output on the side of the axial compressor 30 is dramatically increased. In addition to the above, outer compressor blade group 3 that performs compression and cooling work in all stages in odd stages 3
With 1 etc., it has a great effect to dramatically increase the pressure ratio and increase the thermal efficiency.

The normal turbine stationary blades that do not generate output are completely abolished and all odd-numbered stages are largely converted to the outer turbine blade group 12 for external air heat transfer cooling that generates output. The blades were also completely abolished, and all stages of compression cooling work were performed in odd stages from the first stage to the final stage, and the turbo compressor was converted into an outer compressor rotor blade group 31 equipped with a large number of fans 73 on the entire outermost circumference. When it is used, the blade stages that generate the output are greatly increased by more than double, and the injected gas flow path resistance is drastically reduced. Therefore, there is a great effect that a multistage turbine and large thrust can be obtained by increasing the injection gas velocity at high speed Since the turbine output on the side of the axial compressor 30 increases remarkably, the compression ratio is remarkably increased by the outer compressor blade group 31 which performs compression and cooling work in all stages in odd stages to increase the thermal efficiency. Turbo It is effective to obtain a An'enjin.

In addition to producing no output, it does not move even if a large rotational force is applied, so that the normal turbine stationary blades with large energy loss are completely abolished, and all odd-numbered stages are cooled by heat transfer to the outside air producing all stages. The outer turbine rotor blade group 12 provided with a prop fan 80 on the downstream side outer periphery is largely converted, and the normal axial flow compressor vanes are completely abolished to perform compression and cooling work in all stages in odd stages from the first stage to the final stage. When it is used as a prop fan engine by largely converting it to the outer compressor blade group 31 that operates, the number of blade stages that generate output greatly increases and the jet gas flow path resistance decreases drastically. The large increase in speed has a great effect of dramatically obtaining a multi-stage turbine, and the turbine output on the side of the axial flow compressor 30 increases remarkably. Therefore, the outer compressor performing all-stage compression cooling work in odd stages. There is a large effect in order to obtain dramatically elevated in thermal efficiency and prop fan engine to raise the output like the pressure ratio and the air density by blade groups 31, and the like.

The normal turbine stationary blades that do not generate output are completely abolished and all odd-numbered stages are largely converted to the outer turbine blade group 12 that cools all stages of outside air heat transfer that generates output. The blades are also completely abolished, and all stages are subjected to compression and cooling work in odd stages from the first stage to the final stage, and are largely converted into an outer compressor rotor blade group 31 having a large number of fans 73a on the outermost surface of the uppermost stream, and the gas turbine main shaft 9 When the fan 73b is provided even in the most upstream of the projecting and is used as a double-reversal turbofan engine, the number of blade stages that generate the output is greatly increased more than double and the injection gas flow path resistance is dramatically reduced. Therefore, there is a great effect to obtain a multi-stage turbine dramatically due to a large increase in the injection gas velocity, and the turbine output on the side of the axial flow compressor 30 dramatically increases, so that all-stage compression cooling work is performed in odd-numbered stages. Outside pressure And dramatically increasing the pressure ratio and the air density by blades group 31, etc., it has a large effect in order to obtain increased the thermal efficiency and output 2 counterrotating turbofan engine or the like.

The normal turbine stationary blades that do not generate output are completely abolished, and all odd-numbered stages are largely converted into the turbine moving blade group 12 that cools all stages of external air heat transfer that generates output. Is also completely abolished and is largely converted to the outer compressor blade group 31 which performs compression and cooling work in all stages in odd stages from the first stage to the final stage, and the disc portion 14g of the first stage is extended to the upstream side to prop fan 80a. And further comprises a prop fan 80b extending the gas turbine main shaft 9 on the upstream side thereof.
When used as a double-reversal prop fan engine, the number of blade stages that generate output greatly increases by a factor of two or more, and the jet gas flow path resistance decreases dramatically. There is a great effect of obtaining a turbine, and the turbine output on the side of the axial flow compressor 30 increases dramatically. Therefore, the compression ratio and air density can be increased by the outer compressor blade group 31 that performs compression and cooling work in all stages in odd stages. It has a great effect to obtain a double reversal prop fan etc. that has dramatically increased its thermal efficiency and output.

The outer ends of the prop fans 80a and 80b in the radial direction are projected from the entire surface in the rotational direction toward the downstream side in an inverted L-shape, and the air flows are compressed and deflected inward in the radial direction. By using a rotating double reversal prop fan, it is possible to suppress the generation of shock waves.

The characteristic of the gas turbine provided with the outer turbine rotor blade group 12 and the outer compressor rotor blade group 31 is that the number of turbine stages is dramatically increased in order to dramatically increase the pressure ratio on the axial compressor side. However, since there are no stationary blades, it is possible to expand the combustion gas, etc. almost linearly, and all stages, which have been dramatically stepped up, rotate, so there is very little damping of the injection speed and a large rotational force is obtained. In addition, the relative speed between the moving blades can be made close to twice the conventional speed of the moving blades and the stationary blades, resulting in a small and large output, and the axial flow compressor side also compresses almost linearly, which is a breakthrough. Even with multiple stages, the input becomes small and the pressure ratio can be dramatically increased. The heat transfer cooling by the outside air of the compressed air and the pressure ratio further increase as the flight speed increases. To be more efficient There is a large effect that can increase the number of rows Mach.

The feature of the gas turbine provided with the outer turbine rotor blade group 12 and the outer compressor rotor blade group 31 is that the number of turbine stages for dramatically increasing the pressure ratio on the axial flow compressor side is multistage. However, since there are no stationary blades, it is possible to expand combustion gas etc. almost linearly and the damping of the injection speed is very small, so in addition to being able to obtain a large rotational force by further configuring in multiple stages, Since the relative speed between the moving blades can be made close to twice the conventional speed between the moving blades and the stationary blades, the output is small, lightweight and large, and the axial flow compressor side also compresses almost linearly. Since the input is small, the pressure ratio can be dramatically increased, and the pressure ratio can be dramatically increased. When the combustors 6a, 6b and 6c are provided, the high temperature combustion gas temperature can be changed to the low temperature superheated steam energy. To a steam turbine Possible for the exhaust gas temperature 100 ° C
In addition to the large effect that the adiabatic non-cooling can be achieved and the fuel injection amount can be greatly increased to approach the stoichiometric air-fuel ratio combustion, the gas turbine is a fully rotating engine, and the thermal efficiency can be greatly increased. It has an effect.

When the present invention is used as an ultra-compact gas turbine generator for driving an automobile, the relative speed between the moving blades is the conventional speed of the moving blades and the stationary blades because it is composed of two shafts rotating in opposite directions. In addition to being close to twice the size of the above, it is possible to greatly increase the fuel injection amount by converting it to low-temperature superheated steam energy, which makes it possible to use a microminiature gas turbine generator for automobiles There is a great effect to get.

When the present invention is used for the combined supply of heat and electricity, a high temperature combustion gas temperature can be largely converted into a low temperature superheated steam energy to be synthesized in a steam turbine. Therefore, a large increase in fuel injection amount results in stoichiometric air-fuel ratio combustion. The maximum output can be greatly increased by about 3 times because it is close to the above, and the load adjustment can greatly increase and decrease the fuel injection amount and the superheated steam injection amount. In addition, since a large amount of hot water and hot water close to 100 ° C can be obtained by cooling the exhaust gas, most of the exhaust heat amount can be circulated and used, and in addition, there is a great effect that it can be easily supplied as heat. It also has a great effect to enable more than%.

When the present invention is used as a gas turbine for propulsion of an ultra-high speed ship, a high temperature gas temperature can be largely converted into a low temperature heating steam energy and can be synthesized with a steam turbine.
Due to the large increase in the fuel injection amount, the maximum output can be tripled by approaching the stoichiometric air-fuel ratio combustion, resulting in a large output of small size, light weight, and large output. Load adjustment also greatly increases the fuel injection amount and superheated steam injection amount. Since it can be increased / decreased, it has a great effect that it can be easily controlled. Since hot water and hot water close to 100 ° C can be obtained by cooling the exhaust gas, exhaust heat can be circulated and used as condensate water, and in addition it can be supplied as heat. In addition to the extremely small loss, the gas turbine is a fully rotating engine, which has the great effect of obtaining a gas turbine for propulsion of ultra-high-speed vessels with dramatically higher thermal efficiency.

When the high temperature combustion gas temperature is largely converted to the low temperature superheated steam mass by the combustors 6a, 6b, 6c and the like and synthesized with the steam turbine, the fuel injection amount is greatly increased to approach the stoichiometric air-fuel ratio combustion and the size and weight are reduced. It has a large effect of making it possible to make a large output, and it is optimal to use it as a velocity type energy mainly because it is small and lightweight and has a large output due to the large increase in superheated steam mass. There is a great effect that it can be dramatically increased, and a large increase in mass also has a large effect of reducing noise,
The large conversion to the outer turbine rotor blade group has the great effect of linearly expanding the velocity-type energy to efficiently convert it into rotational power.The large conversion to the outer compressor rotor blade group allows the air to be compressed linearly and efficiently. Since there is a great effect of obtaining a large pressure ratio with the minimum input because the pressure ratio is raised, it is very effective to efficiently obtain low pressure compressed air if the fans 73a and 73b are provided as double reversing fans in the most upstream. .

[Brief description of drawings]

FIG. 1 is a partial cross-sectional view showing a first embodiment.

FIG. 2 is a cross-sectional view showing a cannula type combustor 6b and a joint.

FIG. 3 is a view of a canula type combustor as seen from the axial flow compressor side.

FIG. 4 is a view of a canula type combustor as seen from a turbine side.

FIG. 5 is a cross-sectional view showing an annular combustor 6c and a joint.

FIG. 6 is a cross-sectional view showing an embodiment of the most upstream portion of the combustor inner cylinder.

FIG. 7 is a cross-sectional view for explaining a case where the outer compressor rotor blade group or the outer turbine rotor blade group is used separately.

FIG. 8 is a partial cross-sectional view showing an example of an outer turbine blade group.

FIG. 9 is a partial vertical cross-sectional view showing an example of an outer turbine rotor blade group.

FIG. 10 is a partial vertical cross-sectional view showing an example of an outer turbine rotor blade group.

FIG. 11 is a partial cross-sectional view showing an example of an outer turbine rotor blade group.

FIG. 12 is a partial vertical cross-sectional view showing an example of an outer turbine blade group.

FIG. 13 is a partial vertical cross-sectional view showing an example of an outer turbine rotor blade group.

FIG. 14 is a diagram of an example of an outer turbine rotor blade group as viewed from the combustor side.

FIG. 15 is a cross-sectional view showing an example of an outer turbine rotor blade group integrally provided with a cooling fin.

FIG. 16 is a cross-sectional view showing an example of an outer turbine rotor blade group integrally provided with a fan.

FIG. 17 is a cross-sectional view showing an example of an outer compressor blade group that does not like cooling.

FIG. 18 is a vertical cross-sectional view showing an example of an outer compressor rotor blade group that does not like cooling.

FIG. 19 is a vertical cross-sectional view of an example of an outer compressor rotor blade group including a fan that does not like cooling.

FIG. 20 is a cross-sectional view of an example of a split compressor body for cooling the outside air and an example of an integrated outside compressor body.

FIG. 21 is a vertical cross-sectional view of an example of a group of outer compressor blades for cooling the outside air.

FIG. 22 is a vertical cross-sectional view of an example of a group of outside compressor blades for cooling the outside air, which includes a fan.

[Fig. 23] Fig. 23 is a diagram of an example of a group of outside compressor blades for cooling the outside air viewed from the combustor side.

FIG. 24 is a vertical cross-sectional view showing an example of a group of outer compressor blades for cooling the outside air.

FIG. 25 is a vertical cross-sectional view showing an example of a group of outside compressor blades for cooling the outside air equipped with a fan.

FIG. 26 is a cross-sectional view of an example of an integrated outer compressor body for recooling cooling air.

FIG. 27 is a vertical cross-sectional view showing an example of the flow of cooling air in the vicinity of the outer compressor blade group final stage example and in the outer turbine blade group first stage example.

28A and 28B are a longitudinal sectional view and a front view in the vicinity of an example of the final stage of the outer compressor rotor blade group.

FIG. 29 is a vertical sectional view showing a double inversion gear device.

FIG. 30 is a vertical cross-sectional view showing a double inversion reduction gear device.

FIG. 31 is a sectional view showing a second embodiment.

FIG. 32 is a cross-sectional view showing an outer compressor rotor blade group integrally provided with a fan or a cooling fin and an outer turbine rotor blade group integrally provided with a cooling fin.

FIG. 33 is a cross-sectional view of a double reversal turbofan.

FIG. 34 is a cross-sectional view of a double reversal prop fan.

[Explanation of symbols]

1: Inner cylinder 2: Water pipe 3: Water vapor 4: Fuel injector
5: Turbine 6: Combustor 7: Outer cylinder 8: Steam injection pipe 9: Gas turbine main shaft 10: Turbine blade group 1
1: Compressor blade group 12: Outer turbine blade group 13:
Outer turbine blade group 1st stage 14: Disc part 15: Divided turbine shell 16: Integrated outer turbine barrel 17: Axial unevenness 18: Outer turbine blade group final stage 19: Outer turbine main gear 20: 1st Driven small gear 21: Support shaft
22: 1st drive small gear 23: Spline support shaft (2nd support shaft) 24: 2nd driven small gear 25: 2nd drive small gear
26: Main shaft side driven large gear 27: Annular sheath 28: Combustion gas injection port group 29: Flange 30: Axial flow compressor 3
1: Outer compressor blade group 32: Split compressor barrel 33:
Outer compressor blade group 1st stage 34: Outer compressor blade group final stage
35: Integrated outer compressor barrel 36: Outer compressor main driving gear 37: Outer compressor driven large gear 38: Annular injection port
39: Annular socket 40: Internal gear 41: Double reversing gear device 42: Main shaft side main drive gear 43: Gas turbine main body 44: Main body side support shaft 45: First driven small gear 46:
1st main driving small gear 47: Outer side compressor blade group side 48: Outer side turbine blade group side 49: Gas turbine main shaft side 5
0: Planetary reduction gear unit 51: Double reversal reduction gear unit
52: sun gear 53: planetary gear 54: internal gear 5
5: Planetary gear support shaft 56: Circular plate 57: Power shaft 5
8: Outer gas turbine main shaft 59: Guide vanes 60: Cooling fins 61: Superheated steam control valve 62: Annular combustion chamber 6
3: Inner surface of outer inner cylinder 64: Inner surface of inner inner cylinder 65: Labyrinth seal 66: Overheated air generation cylinder 67: Overheated air injection hole 68: Levee 69: Dividing portion 70: Shroud
71: space 72: bolt 73: fan 74: cooling air 75: cooling air passage 76: integral joint 77: annular air reservoir 78: cooling air inlet 80: prop fan 81: arm 82: ram fuel injector 83: ram Combustor 84: Ramjet propulsion device

─────────────────────────────────────────────────── ─── Continuation of the front page (51) Int.Cl. ⁶ Identification code Office reference number FI technical display location F02K 3/06 F02K 3/06 F23R 3/00 F23R 3/00 A 3/44 3/44 3 / 46 3/46

Claims (40)

[Claims] 1. Various types of combustors for gas turbines, wherein, for example, a can-shaped combustor (6a) is extended to an arbitrary shape and an upstream part for generating heat is used to dilute with a downstream air. Is removed and only the inner cylinder (1a) including the quadrangular shape is provided, and the water pipe (2) is arranged along the inside of the multi-threaded screw with an appropriate thickness, length and interval, and the external water injection pump pressure is increased. And a large number of superheated steam control valves (61) are calculated and controlled on the basis of numerical values at arbitrary measurement points to supply high-pressure water and superheated steam to almost completely burn the inside of the large number of inner cylinders (1a). The steam injection pipe (8) selected from the upstream side to the downstream side of the water pipe (2) on the downstream side of the completed portion is extended to include direct injection to the outer turbine rotor blade group 1 stage (12) downstream. Then, steam (3) such as superheated steam is injected from many steam injection pipes (8). Te, normal fuel injectors (4) those exhaust temperature greatly lowers the combustion gas temperature of the high-temperature ended fuel injection complete combustion was 500 ° C before and after from 10
By approaching 0 ° C, it becomes possible to approach the stoichiometric air-fuel ratio combustion, the output is greatly increased by the increase of the fuel injection amount, the exhaust heat loss is greatly reduced, and the combustion gas and the superheated steam coexist to condense water. The loss is eliminated, and the mass injected into the turbine is greatly increased as the mass of combustion gas + mass of superheated steam to increase the thermal efficiency of the gas turbine and dramatically reduce the heat load to obtain an inexpensive turbine. Combustor for various gas turbines. 2. Various types of combustors for gas turbines, wherein a combustor (6b) of, for example, a cannula type is extended to an arbitrary shape and an upstream part for generating heat is used to dilute with a downstream air. Inner cylinder (1b) with squares removed to eliminate
As a chisel, a plurality of water pipes (2) are arranged along the inner surface thereof in the shape of multiple threads with appropriate intervals, thicknesses and lengths, and external water injection pump pressure and a large number of superheated steam control valves (61) are provided. Is calculated and controlled based on the numerical value of an arbitrary measurement point to supply high-pressure water and superheated steam, and the water pipe (2 ) Selected from upstream to downstream steam injection pipe (8)
Is extended to include the direct injection to the outer turbine blade group 1 stage (12) downstream to inject steam (3) such as superheated steam from a large number of steam injection pipes (8) for normal fuel injection. (4) The temperature of the high-temperature combustion gas after the complete fuel injection is drastically reduced and the exhaust temperature is brought close to 100 ° C, which makes it possible to approach the stoichiometric air-fuel ratio combustion, and the output is increased by a large increase in the fuel injection amount. Is greatly increased, exhaust heat loss is greatly reduced, and condensate loss is eliminated by co-existing combustion gas and superheated steam, and a large amount of hot water and hot water can be easily obtained by cooling the exhaust gas if necessary. In addition to facilitating the use of exhaust heat, the mass injected into the turbine is greatly increased as the mass of combustion gas + mass of superheated steam to increase the thermal efficiency of the gas turbine and dramatically reduce the heat load, making it an inexpensive gas turbine. Eye to get And the combustor of a variety of gas turbines. 3. In various combustors for gas turbines, for example, an annular type combustor (6c) is extended to an arbitrary shape and an upstream part for generating heat is used to dilute with a downstream air. To remove only the inner cylinder (1c) and form a large number of water pipes (2) along the outer inner cylinder inner surface (63) and the inner inner cylinder inner surface (64) of the annular combustion chamber (62). Each of them is arranged with an appropriate interval, thickness and length, and the external water injection pump pressure and a large number of superheated steam control valves (61) are controlled by calculation based on the numerical values at arbitrary measurement points to achieve high pressure. Of the inner cylinder (1
The steam injection pipe (8) selected from the upstream side to the downstream side of the water pipe (2) on the downstream side of the portion where the complete combustion inside c) is almost completed is directly injected downstream to the first stage of the outer turbine rotor blade group. Water vapor such as superheated steam (3)
Is injected from a large number of steam injection pipes (8), and the high temperature combustion gas temperature at which the complete fuel injection is completed by the normal fuel injector (4) is drastically reduced to bring the exhaust temperature close to 100 ° C. By making it possible to approach stoichiometric air-fuel ratio combustion, the output greatly increases due to the large increase in fuel injection amount, exhaust heat loss is greatly reduced, and there is no condensate loss by coexisting combustion gas and superheated steam. If necessary, a large amount of hot water and hot water can be easily obtained by cooling the exhaust gas to facilitate the use of exhaust heat, and the mass injected into the turbine can be the mass of the combustion gas +
A combustor for various gas turbines whose purpose is also to obtain an inexpensive gas turbine by increasing the thermal efficiency of the gas turbine by dramatically increasing the mass of superheated steam and dramatically reducing the heat load. 4. A combustor for a gas turbine is used in combination with an outer compressor rotor blade group final stage (31), so that it projects annularly from the outer compressor blade group final stage (31) and has a labyrinth seal ( 65) has an annular injection port (38) rotatably fitted to the outside and has a labyrinth seal (6) inside.
An annular receiving port (39) and a flange (29) forming one side of 5) are provided, and a large number of guide vanes (59) are provided inside the outermost outer cylinder (7) of various combustors to provide an annular outside. The outer compressor blade group final stage (31) by fixing the flange (29b)
A group of outer compressor blades including an annular receiving port, which can be combined with various combustors. 5. A combustor for a gas turbine is used in combination with a first stage (12) of an outer turbine rotor blade group. An annular combustion gas injection port group (28) that is rotated by the injection of
On the outer turbine blade group 1st stage (12) side, an annular sheath (27) for receiving combustion gas or the like injected from the annular combustion gas injection port group (28) has a Rabins seal (65) on the outside. A gas injection port group (28) is rotatably fitted to the outside and one of the Rabins seals (65) is also provided at the fitting portion to combine the outer turbine rotor blade group 1 stage (12) with various combustors. An outer turbine rotor blade group including a combustion gas injection port group, which is characterized by combining a turbine (5) and various combustors. 6. An embankment (68) for converting high-speed air injected from the compressor side into high-pressure air for cooling high-temperature combustion heat to produce high-temperature superheated air for high-speed agitation combustion.
Ordinary fuel injector (4)
A cup of overheated air generating cylinder (6
6) with an appropriate number of superheated air injection holes (6
By enlarging a number of 7) to increase the number of holes to be injected toward the center and inclining the injection direction to the upstream side as well, a large amount of high-temperature superheated air is agitated and mixed to completely burn a lean fuel that is difficult to burn. A combustor for a gas turbine, characterized by being terminated. 7. A group of outer turbine blades used for non-cooling purposes is largely converted to an outer turbine blade group (12) by eliminating all the ordinary turbine stationary blade groups and all odd-numbered stages from the first stage to the final stage. Then, a large number of axial irregularities (17a) and (17b) are provided in the axial direction of the split turbine shell (15) to divide the outer turbine blade group (12) in the middle in the radial direction or as a unit. The outer turbine rotor blade group first stage (13) and the outer turbine rotor blade group final stage (18) are inserted into and fixed to the shaped turbine shell (15), respectively.
a) and (14b) are provided and integrally fixed in the radial direction and the axial direction, respectively, and are pivotally supported on the gas turbine main shaft (9), respectively, and burned from the radially inner side of the outer turbine rotor blade first stage (13). A Rabins seal (65) is provided by projecting an inner side of an annular sheath (27) in an annular shape on the container side, and an integrated turbine body (16) having a taper is provided on the outer side of the split type turbine body (15). An annular sheath (27) is provided on the combustor side to form a double structure by externally fitting along the multiple axial irregularities (17a) and (17b) to provide multiple spaces (71).
One of the two is protruded to provide a labyrinth seal (65) and fixed to the outer turbine rotor blade group first stage (13) and the outer turbine rotor blade group final stage (18) in the radial direction and the axial direction, and An outer turbine blade group for a gas turbine, which is provided as a blade group. 8. A turbine blade is heat-transfer-cooled by the outside air, and a normal turbine vane is completely abolished, and all odd-numbered stages from the first stage to the final stage are largely converted into an outer turbine rotor blade group (12) and divided. A large number of axial irregularities (17a) (17) in the axial direction of the annular split turbine shell (15) having the portion (69b).
b) is provided to divide the intermediate outer turbine rotor blade group (12) in the radial direction, or to insert and fix the intermediate outer turbine rotor blade group (12) between the divided turbine shells (15) as a unit. (13) and outer turbine blade group final stage (18)
Also provided with disk portions (14a) and (14b), respectively, integrally fixed in the radial direction and the axial direction, respectively, and pivotally supported on the gas turbine main shaft (9), respectively, and the outer turbine rotor blade first stage (13) A labyrinth seal (65) is provided by annularly projecting the inside of the annular sheath (27) from the inner side in the radial direction to the combustor side, and one of the annular sheaths (27) is installed inside the integral turbine shell (16). Protrusion and labyrinth seal (6
5) and a plurality of cooling fins (60) provided on the outer periphery at an appropriate angle and having a taper, the integrated outer turbine shell (16)
b) a large number of axial irregularities (17a) each provided
The outer turbine blade group (1
2) and the integrated outer turbine shell (16b) contact in a wide range to dissipate heat faster than the cooling fin (60) that is in high-speed contact with the outside air, and the double structure makes the outer turbine blade group easy to assemble. An outer turbine rotor blade group for a gas turbine, which is provided as (12). 9. A high speed cooling of the turbine blades by the outside air, the conventional turbine stationary blades are completely abolished, and all the odd stages from the first stage to the final stage are converted into the outer turbine rotor blade group (12). The rotor blade group 1st stage (13) is a cooling fin (60)
And the annular sheath (27) and the disc portion (14a) are integrally formed and pivotally supported on the gas turbine main shaft (9), and the middle stage of the outer turbine rotor blade group (12) serves as a cooling fin (60). The outer turbine group final stage (18) is integrally configured with the cooling fin (60) and the disc portion (14b) and is pivotally supported on the gas turbine main shaft (9), and each is bolt (72). An outer turbine rotor blade group for a gas turbine, which is equipped with an outer turbine rotor blade group (12) that is appropriately tightened by the above to rapidly heat-transfer and cool the turbine blades by the outside air. 10. The integral outer turbine shell (16) for using the outer turbine blade group (12) also as a fan.
The outer turbine rotor blade group for a gas turbine according to claim 7, wherein a large number of fans (73) are integrally provided in one or more rows on the outer surface of (a). 11. The outer turbine blade group (12) is also used as a prop fan, so that a large number of prop fans (80) are annularly connected to the outer surface of the integral outer turbine shell (16b) at the tip thereof to project. The outer turbine blade group for a gas turbine according to claim 8, wherein a cooling fin (60) is provided at the selected and installed location at the angle of the fan. 12. The outer turbine rotor blade group (12) is also used as a prop fan, so that the normal turbine stator blade group is completely abolished and all odd-numbered stages from the first stage to the final stage are used as the outer turbine rotor blade group. After major conversion, the outer turbine rotor blade first stage (13) is pivotally supported on the gas turbine main shaft (9) by integrating the cooling fin (60), the annular sheath (27) and the disc portion (14a). The middle stage of the outer turbine rotor blade group (12) is integrated with the cooling fin (60) or integrated with the part of the prop fan (80) and the cooling fin (60) which are separated from each other. The final stage (18) is integrated with the part of the prop fan (80), the cooling fin (60), and the disc portion (14b) which are separated from each other, and the gas turbine main shaft (9).
The propeller (80) is pivotally supported at each stage and is tightened inside the prop fan (80) with bolts (72) and is annularly connected at the tip of the prop fan (80) to form an outer turbine blade group and a prop fan (80). ) And a cooling fin (60), which are provided as an outer turbine blade group, and an outer turbine blade group for a gas turbine. 13. An outer compressor blade group for use in a non-cooling application, the normal axial flow compressor vane is completely abolished, and an odd number of outer compressor blade groups (31) from the first stage to the final stage. And a large number of axial irregularities (17c) (17) on the ring-shaped split compressor cylinder (32) having the split portion (69b).
d) is provided to divide the outer compressor blade group (31) in the middle by radially providing the dividing portions (69a), or is inserted and fixed as a unit between the divided compressor cylinders (32). The outer compressor blade group 1 stage (33), the outer compressor blade group final stage (34), and the intermediate selected outer compressor blade group (31) have disk portions (14c) and (14d), respectively.
And (14c), (14f), etc. are appropriately provided and integrally supported to the gas turbine main shaft (9) to form an annular shape on both sides of the outer compressor rotor blade final stage (34) in the radial direction from the radial side. An integral outer compressor body (35) having an injection port (38) projecting annularly and providing labyrinth seals (65) on the outer and inner peripheries, and having a taper inside on the outer side of the split type compressor body (32).
a) Multiple axial irregularities (17c) each provided with a)
The outer compressor blade group 1 (33) and the outer compressor blade group final stage (34) are fixed to each other in the radial direction and the axial direction by externally fitting along (17d) to provide a large number of spaces (71). An axial compressor outer compressor blade group for a gas turn, comprising an outer compressor blade group (31) having a heavy structure. 14. An outer compressor blade group (31) for cooling by outside air, wherein a normal axial flow compressor vane is completely abolished and all odd-numbered stages from the first stage to the final stage are arranged. )
And a large number of axial irregularities (17e) are provided on the ring-shaped split compressor body (32) having the split portion (69b), and the intermediate outer compressor rotor blade groups (31) are respectively arranged in the radial direction. The divided portion (69a) is divided into two parts, or is inserted and fixed as a unit between the divided type compressor body (32), and the outer compressor blade group 1 stage (33) and the outer compressor blade group final stage (34) and an intermediate selected outer compressor blade group (31)
Disc portions (14d) and (14e) are appropriately provided and integrally supported to the gas turbine main shaft (9) from both sides in the radial direction of the outer compressor rotor blade final stage (34). A ring-shaped injection port (38) is formed in a ring shape on the combustor side, and labyrinth seals (65) are provided on both the inner and outer circumferences, and a large number of cooling fins (60) are arranged on the outer circumference of the split compressor body (32).
Of outer compressor blades by externally fitting along a number of axial irregularities (17e) (17f) each provided with an integral outer compressor body (35b) having a taper inside at an appropriate angle. First stage (33) and outer compressor blade group final stage (34)
An outer compressor blade group of an axial flow compressor for a gas turbine, comprising an outer compressor blade group (31) having a double structure fixed to the radial direction and the axial direction. 15. The outer compressor blade group (31) for obtaining the best compressed air multi-stage outside air cooling, the normal axial flow compressor vanes are completely abolished, and all the odd stages from the first stage to the final stage are outside. Largely converted to the compressor blade group (31), the outer compressor blade group 1
The stage (33) is pivotally supported on the gas turbine main shaft (9) integrally with the cooling fin (60) and the disc portion (14g), and the intermediate stage of the outer compressor blade group (31) is the cooling fin (60). And a disk portion (14i) or the like provided at the intermediate stage which is fixed to the front stage by bolts (72) and is pivotally supported by the gas turbine main shaft (9), and the outer compressor blade group final stage A ring-shaped mound (3
8) Protruding and labyrinth seals (65) on both inside and outside
And a cooling fin (60) provided radially outward and a disk portion (14h) provided radially inward and pivotally supported on the gas turbine main shaft (9) and fixed to the preceding stage by a bolt (27). An outer compressor blade group of an axial compressor for a gas turbine, comprising an outer compressor blade group (31). 16. The integral outer compressor barrel (35a) for using the outer compressor blade group (31) as a fan as well.
The outer compressor blade group of the axial flow compressor for a gas turbine according to claim 13, wherein one or more rows of fans (73) are integrally provided on the outer circumference of the upstream side of the fan. 17. The integral outer compressor barrel (35b) for using the outer compressor blade group (31) also as a fan.
15. A plurality of fans (73) are integrally provided with one or more rows of cooling fins (60) in the same direction so as to integrally project on the outer circumference of the upstream side of the outside of the axial compressor for a gas turbine according to claim 14. Compressor blade group. 18. The outer compressor rotor blade group (31) is also used as a fan, so that the outer compressor rotor blade group one stage (33).
Is the separated fan (73) and disk (14g)
Is integrally pivoted to the gas turbine main shaft (9), and the middle stage and below of the outer compressor blade group (31) are divided according to the application to the part of the fan (73) and the cooling fin (in the same direction as the fan). 16. The fan of two or more rows can be mounted by fixing the above-mentioned 60) and the like integrally with the bolts (72) to the upstream stage and fixing the downstream stage according to the application. Of the outer compressor blades of the axial compressor for gas turbines in Japan. 19. A space (71a) (71b) of the outer compressor blade group (31) is cooled by a cooling air passage (71a) to obtain cooling air (74) on the turbine side using the outer compressor blade group (31). 75) used as the outer compressor blade group final stage (3
4) A part of the compressed air blown out from 4) is connected to a large number of spaces (71a) outward in the radial direction, and is moved in the space (71a) to form a split type with the outer compressor rotor blade first stage (33). Connect to the space (71b) at the joint of the compressor body (32),
The cooling air (74) on the turbine side is obtained by moving in the space (71b) and ejecting it toward the combustor side from the outside of the outer compressor blade group final stage (34) in the radial direction. The outer compressor blade group of the axial-flow compressor for a gas turbine according to claim 13 or 16. 20. A space (71) of the outer compressor blade group for efficiently cooling the compressed air for obtaining cooling air (74) on the turbine side by using the outer compressor blade group (31).
a) (71b) is deleted and only (71b) is restored and (71c) (71d) is used as a large number as the cooling air passage (75), and the outer compressor rotor blade final stage (3)
4) A part of the compressed air ejected from 4) is connected to a large number of spaces (71c) in the radial direction outward, and is moved in the space (71c) to form a split type with the outer compressor rotor blade first stage (33). Connect to the space (71d) at the joint of the compressor body (32),
The cooling air (74) on the turbine side is obtained by moving the air in the space (71d) and ejecting it toward the combustor side from the radial outer side of the outer compressor blade group final stage (34). The outer compressor blade group of the axial compressor for a gas turbine according to claim 14 or claim 17. 21. In order to obtain low-temperature cooling air (74) on the turbine side by using the outer compressor blade group (31), the outer compressor blade group (3) for efficiently cooling the compressed air.
Since 1) is integrally formed and assembled for each blade stage from the first stage to the final stage, it is best to newly install the cooling air passage (75) in the cooling fin (60). )
Then, a part of the compressed air ejected from the final stage (34) of the outer compressor blade group is connected to the large number of cooling air passages (75a) in the radially outward direction to move in the cooling air passages (75a). The outer compressor blade group 1 stage (33) and the same three-stage joint portion are connected to the cooling air passage (75b) and moved in the cooling air passage (75b) to move the outer compressor blade group final stage (3).
The low temperature turbine side cooling air (74) is obtained by jetting from the radially outer side of 4) to the combustor side, to obtain an axial flow compressor for a gas turbine according to claim 15 or claim 18. Outer compressor blades. 22. The low temperature cooling air (74) obtained by cooling the compressed air by using the outer compressor blade group (31) again is supplied to the cooling air inlet (78a) to form the outer turbine blade group first stage (13). ) For cooling the outer turbine blade group 1 stage (1
3) is a large wing, and the outer cooling fin (60) is brought into high-speed contact with the outside air to conduct heat transfer cooling from the outer diameter side, and the annular sheath (27) and the integral joint (76) have greatly expanded the cooling area.
And the cooling air passage (7
The cooling air (74) is supplied to 5) to mainly cool from the inner diameter side, and further to cool the blade from the inner diameter side, the cooling air inlet (78b) that cools the turbine moving blade group (10) side is almost axially arranged. A large number of labyrinth seals (6
The gas turbine according to claim 8 or 9, characterized in that 5) is annularly provided inward of the integral joint (76) as a first stage (13) of a group of outer turbine blades. Outer turbine blades. 23. An outer compressor blade group (31) comprising a double reversing gear device (41) upstream of the outer compressor blade group (31) or downstream of the outer turbine blade group (12). Or outer turbine rotor blade group (12) and gas turbine main shaft (9)
In order to accurately rotate the two in opposite directions at a selected rotation ratio, an internal gear (40) is provided on the outer compressor blade group side (47) or the outer turbine blade group side (48), and the gas turbine main shaft side is provided. A plurality of first driven small gears fixed to a plurality of main body side support shafts (44) rotatably supported by a gas turbine main body (43) by fixing a main shaft side main driving gear (42) to (49). The first main driving small gear (46) meshing with (45) and fixed to the other end of the main body side support shaft (44) is connected to the outer compressor moving blade group side (47) or the outer turbine moving blade group side (48). The internal compressor (40) provided to the outer compressor blade group (3
1) or the outer turbine rotor blade group (12) and the gas turbine main shaft (9) accurately rotate in opposite directions at a selected rotation ratio, the outer compressor rotor blade group (31) or the outer turbine rotor blade. 2 for gas turbines including group (12)
Double inversion gear device. 24. A double reversing reduction gear device (51) is provided upstream of the outer compressor blade group (31) or downstream of the outer turbine blade group (12) to provide the outer compressor blade group (31). )
Alternatively, in order to accurately rotate the outer turbine rotor blade group (12) and the gas turbine main shaft (9) in opposite directions at a selected rotation ratio and reduction ratio, the outer compressor rotor blade group (31) or the outer turbine rotor blade group ( 12) and the sun gear (52a) (52a) (52) at the outer end of the gas turbine main shaft (9) and the like.
b) and meshes with the plurality of planetary gears (53a) (53b), respectively.
b) are meshed with the internal gears (54a) (54b), respectively, and the circular plates (5) that pivotally support the planetary gear support shafts (55a) (55b) of the plurality of planetary gears (53a) (53b) from both sides.
On the right side of 6a), a cylindrical outer compressor blade group side (47)
Alternatively, as the outer turbine rotor blade group side (48), the internal gear (4
0), the main shaft side main driving gear (42) is fixed to the left side of the circular plate (56b), and the gas turbine main shaft side (4
As a 9), a plurality of first driven gears (45) are meshed with the main drive gear (42) on the main shaft side to form a gas turbine body (43).
A plurality of first main driving small gears (46) fixed to the other end of the main body side supporting shaft (44) pivotally supported by the inner gear (40) are meshed with the outer compressor blade group side (47). Alternatively, the outer turbine blade group side (48) and the gas turbine main shaft side (49) are connected to each other, and a power shaft (57) is projectingly provided at the center of the right side of the circular plate (56b) from the power shaft (57). 2 includes an outer compressor rotor blade group (31) or an outer turbine rotor blade group (12) characterized by decelerating and extracting biaxial power and rotating them in opposite directions to each other at a selected rotation ratio and reduction ratio. Double inversion reduction gear device. 25. The outer turbine blade group (12) for use in combination with a conventional simple cycle gas turbine,
A turbine main drive gear (19) is provided on the upstream side of the outer turbine blade group (12), meshes with the first driven small gear (20b), and is pivotally supported by the gas turbine body (43). The first driving pinion (22) fixed to the other end of the shaft (21b).
b) is meshed with a second driven small gear (24b) fixed to a second support shaft (23b) pivotally supported by the gas turbine body (43), and fixed to the second support shaft (23b). The other second main driving small gear (25b) is meshed with the main shaft side driven large gear (26) so that the outer turbine blade group (12) of the present invention can be used in combination with a normal simple cycle gas turbine. A double inversion gear device characterized by the above. 26. The outer compressor blade group (31) is used in combination with an ordinary simple cycle gas turbine, so that the outer compressor main shaft (9) is provided downstream of the outer compressor blade group (31). First with the large gear (36) fixed
A two-stage first driving small gear (22a) (22c) that is meshed with the driven small gear (20a) and is fixed to the other end of the support shaft (21a) pivotally supported by the gas turbine body (43). Is meshed with a two-stage second driven small gear (24a) (24c) slidably fitted onto a spline support shaft (23a) pivotally supported by the gas turbine body (43) so that the spline can be changed. The other second main driving small gear (25a) fixed to the support shaft (23a) is meshed with the outer compressor driven large gear (37) fixed to the outer compressor moving blade group (31) to perform normal compression. The outer compressor blade group (31) and the outer compressor blade group (31) of the present invention are rotated in opposite directions by selecting from two rotation ratios so that the outer compressor blade group (31) is a normal simple cycle gas. Double reversing gear, characterized by being used in combination with a turbine Location. 27. An outer turbine blade group (12) and an outer compressor blade group (31) are connected by an outer gas turbine main shaft (58) to provide a turbine blade group (10) and an ordinary simple cycle gas turbine. The outer turbine rotor blade group (12) for rotatably externally fitting to a compressor rotor blade group (11) and a gas turbine main shaft (9) and having a normal combustor or the like to form a turboset engine. Alternatively, the outer turbine rotor blade group (12) according to claim 9 is used, and the outer compressor rotor blade group (31) according to claim 14 or 15 is used as the outer compressor rotor blade group (31). By converting all of the outside into moving blades, rapid heat transfer cooling by outside air is possible and the relative speed between moving blades can be made close to twice the relative speed between moving blades and stationary blades. A gas turbine that has been characterized. 28. An outer turbine blade group (12) and an outer compressor blade group (31) are connected by an outer gas turbine main shaft (58) to form a turbine blade group (10) and a turbine blade group (10) of a normal simple cycle gas turbine. The outer turbine rotor blade group (12) for rotatably externally fitting to a compressor rotor blade group (11) and a gas turbine main shaft (9) and having a normal combustor or the like to form a turboset engine. Alternatively, the outer turbine moving blade group (12) according to claim 9 is used, and the combination of the outer turbine moving blade group (12) and the outer compressor moving blade group (31) is used as claim 20 or 21 or 22. By using the outer compressor rotor blade group (31) and the outer turbine rotor blade group (12) described in 1 above, the outside is largely converted into rotor blades to enable rapid heat transfer cooling by outside air and Cooling air While a low temperature by cooling the gas turbine it possible to bring the relative speed between the moving blade to twice the relative speed between the rotor blade pairs vanes. 29. An outer turbine blade group (12) and an outer compressor blade group (31) are connected by an outer gas turbine main shaft (58) to form a turbine blade group (10) and a turbine blade group (10) of an ordinary simple cycle gas turbine. The outer turbine rotor blade group (12) for rotatably externally fitting to a compressor rotor blade group (11) and a gas turbine main shaft (9) to have a normal combustor or the like to form a turbofan engine. Alternatively, the outer turbine rotor blade group (12) according to claim 9 is used, and the outer compressor rotor blade group (31) according to claim 17 or 18 is used as the outer compressor rotor blade group (31). Depending on the application, the outer turbine rotor blade group (12) and the outer compressor rotor blade group (31)
By adding the outer turbine rotor blade group (12) and the outer compressor rotor blade group (31) according to claims 20 to 22 as a combination of the above, the outer side in the radial direction is largely converted into rotor blades. Enables rapid heat transfer cooling by outside air and re-cools the cooling compressed air on the turbine side by outside air as low temperature cooling air, and the relative speed between moving blades is double the conventional relative speed between moving blades and stationary blades. A gas turbine that makes it possible to bring them closer to each other and facilitates selection of rotational speed. 30. An outer turbine blade group (12) and an outer compressor blade group (31) are connected by an outer gas turbine main shaft (58) to form a turbine blade group (10) of a normal simple cycle gas turbine. The outer turbine rotor blade group (12) for rotatably fitting the compressor rotor blade group (11) and the gas turbine main shaft (9) to a prop fan engine equipped with an ordinary combustor. Or claim 1
The outer turbine rotor blade group (12) according to claim 2 is used, and the outer compressor rotor blade group (31) is used as the outer compressor rotor blade group (31).
The outer compressor blade group (31) according to claim 1 is used for an application using low-temperature cooling air.
As a combination of 2) and the outer compressor blade group (31), the turbine rotor blade group (12) and the outer compressor blade group (31) according to claims 20 to 22 are added,
By converting all of the radial outside into moving blades, rapid heat transfer cooling by outside air is possible, and the compressed compressed air on the turbine side is re-cooled by outside air to be low temperature cooling air. A gas turbine characterized by a dramatic increase. 31. The outer turbine blade group (12) and the outer compressor blade group (31) are connected to the outer gas turbine main shaft (5).
8) is connected to the turbine rotor blade group (10), compressor rotor blade group (11) and gas turbine (9) of an ordinary simple cycle gas turbine so as to be rotatably externally fitted and provided with an ordinary combustor or the like. In the above gas turbine, a fan (73b) is also provided at the most upstream side of the gas turbine main shaft (9) so that the fans (73a) and (73b) rotate in mutually opposite directions.
30. The gas turbine according to claim 29, which is a double-reversal turbofan engine. 32. The outer turbine blade group (12) and the outer compressor blade group (31) are connected to the outer gas turbine main shaft (5).
8) and is rotatably externally fitted to the turbine moving blade group (10), the compressor moving blade group (11) and the gas turbine main shaft (9) of a normal simple cycle gas turbine to connect a normal combustor or the like. In the equipped gas turbine, the first stage (31) of the outer compressor rotor blade group is extended to the upstream side, the prop fan 80a is provided at the tip, and the rear part is rotatably supported by the gas turbine body (43). With its arms (81)
Using the appropriate parts of the above as the stationary blades, the gas turbine main shaft (9) is also extended to the most upstream side and provided as a prop fan (80b) on the upstream side of the prop fan (80a) to rotate in mutually opposite directions. Double reversal prop fan (80
28. A) (80b) is provided.
Alternatively, the gas turbine according to claim 28. 33. The double reversal prop fan (80a).
(80b) The outer tips in the radial direction are projected in an inverted L shape toward the front side and the downstream side in the rotational direction, respectively, and the airflow is compressed and deflected inward in the radial direction to suppress the generation of shock waves. The double reversal prop fan according to claim 32, wherein: 34. The outer turbine blade group (12) and the outer compressor blade group (31) are connected to the outer gas turbine main shaft (5).
8) and is rotatably externally fitted to the turbine moving blade group (10), the compressor moving blade group (11) and the gas turbine main shaft (9) of a normal simple cycle gas turbine to connect a normal combustor or the like. In the equipped gas turbine, a portion of the gas turbine main shaft (9) having a ram combustor (83) is appropriately enlarged and the ram combustor (83) is provided in the portion.
29. The gas turbine according to claim 27 or 28, further comprising a ram jet propulsion device (84) provided with a ram fuel injector (82) and the like, which are fixedly attached to the ram jet injector. 35. An outer turbine blade group (12) according to claim 7 and an outer compressor blade group (31) according to claim 13 for use as an adiabatic uncooled theoretical air-fuel ratio combustion gas turbine. The turbine main shaft (58) is connected to the turbine rotor blade group (10), the compressor rotor blade group (11) and the gas turbine main shaft (9) of an ordinary simple cycle gas turbine so as to be rotatably fitted on the outer periphery thereof. 4. A gas turbine comprising the combustor according to claim 3. 36. A gas turbine main shaft (9) of said gas turbine and a downstream or outer compressor rotor blade group (31) of said outer turbine rotor blade group (12) for use as a micro gas turbine generator for driving an automobile. 36. A publicly known generator is provided upstream, and a publicly known heat exchange condenser for cooling and circulating exhaust gas from the outer turbine blade group final stage (12) is provided. The gas turbine described in 1. 37. A gas turbine main shaft (9) of the gas turbine, downstream of the outer turbine blade group (12) or upstream of the outer compressor blade group (31) for use as a gas turbine for supplying heat and electricity together. Double reversing gear unit (41)
Alternatively, a double inversion reduction gear device (51) is configured to communicate with a load such as a generator, and the outer turbine rotor blade group final stage (1
The gas turbine according to claim 35, further comprising a known heat exchange condenser for cooling and circulating the exhaust gas from 2). 38. A double-reversing reduction gear unit (5) downstream of the gas turbine main shaft (9) and the outer turbine rotor blade group (12) for use as a gas turbine for propulsion of an ultrahigh-speed ship.
1) A well-known heat exchange condenser for cooling and circulating exhaust gas from the outer turbine blade group final stage (12) by connecting it to a well-known water jet propulsion machine or the like downstream thereof The gas turbine according to claim 35, further comprising: 39. For use as a gas turbine for propulsion of an ultrahigh-speed ship, ordinary reduction gears are provided at the downstream of the gas turbine main shaft (9) and the outer turbine rotor blade group (12), respectively, and a double gear downstream thereof is provided. A known heat exchange condenser or the like for cooling and circulating the exhaust gas from the outer turbine blade group final stage (12) is provided, which is connected to each of the reverse water jet propulsion machines and the like. Item 35. The gas turbine according to Item 35. 40. An outer turbine rotor blade group (12) according to claim 7 and an outer compressor rotor blade according to claim 16 for simultaneously obtaining rotational power and compressed air as a gas turbine for propulsion of an ultra high speed ship. The group (31) is connected to the outer gas turbine main shaft (5
8) A gas turbine main shaft (9) provided with a turbine blade group (10) and a compressor blade group (11) of an ordinary simple cycle gas turbine, and a fan (73b) according to claim 31 by being connected by 8). A rotatably fitted externally equipped combustor according to any one of claims 1 to 3, wherein a double reversing reduction gear is installed downstream of the gas turbine main shaft (9) and the outer turbine rotor blade group (12). (51) A known heat exchange condenser for connecting to a known water jet propulsion machine downstream thereof via (51) or the like to cool and circulate exhaust gas from the outer turbine blade group final stage (12) A gas turbine characterized in that it is equipped with.