CN114483612A - Aerodynamic turbine compression system - Google Patents
Aerodynamic turbine compression system Download PDFInfo
- Publication number
- CN114483612A CN114483612A CN202210207659.9A CN202210207659A CN114483612A CN 114483612 A CN114483612 A CN 114483612A CN 202210207659 A CN202210207659 A CN 202210207659A CN 114483612 A CN114483612 A CN 114483612A
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- CN
- China
- Prior art keywords
- compressor
- aerodynamic
- turbine
- air
- bleed
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 230000006835 compression Effects 0.000 title claims abstract description 15
- 238000007906 compression Methods 0.000 title claims abstract description 15
- 239000002828 fuel tank Substances 0.000 claims abstract description 11
- 239000000463 material Substances 0.000 claims description 9
- 238000007789 sealing Methods 0.000 claims description 3
- 239000012530 fluid Substances 0.000 claims description 2
- 230000002411 adverse Effects 0.000 description 5
- 230000000694 effects Effects 0.000 description 4
- 230000006378 damage Effects 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 238000011144 upstream manufacturing Methods 0.000 description 3
- 229910001220 stainless steel Inorganic materials 0.000 description 2
- 239000010935 stainless steel Substances 0.000 description 2
- 230000007547 defect Effects 0.000 description 1
- 230000035484 reaction time Effects 0.000 description 1
- 229920002379 silicone rubber Polymers 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D25/00—Pumping installations or systems
- F04D25/02—Units comprising pumps and their driving means
- F04D25/04—Units comprising pumps and their driving means the pump being fluid-driven
- F04D25/045—Units comprising pumps and their driving means the pump being fluid-driven the pump wheel carrying the fluid driving means, e.g. turbine blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D15/00—Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby
- F01D15/08—Adaptations for driving, or combinations with, pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/02—Surge control
- F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
- F04D27/0215—Arrangements therefor, e.g. bleed or by-pass valves
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/02—Surge control
- F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
- F04D27/0223—Control schemes therefor
Abstract
The invention provides an aerodynamic turbine compression system. The system comprises an air power turbine, a compressor and a shaft connected between the air power turbine and the compressor, wherein the air power turbine compression system further comprises an air bleed pipeline, the air bleed pipeline is provided with an introducing end and an extracting end, the introducing end of the air bleed pipeline is communicated with an outlet part of the compressor, the air bleed pipeline is provided with a one-way valve, when the compressor enters surge, the one-way valve is opened, and when the compressor exits the surge, the one-way valve is automatically closed. The aerodynamic turbocompression system according to the invention is particularly suitable for bleed air pressurization of an aircraft fuel tank inerting system. The bleed air line can rapidly discharge high-pressure gas when compressor surge occurs, so that the surge is remarkably relieved or eliminated, and the flight safety of the airplane is ensured.
Description
Technical Field
The invention relates to the technical field of aviation, in particular to an aerodynamic turbine compression system for carrying out air entraining on a fuel tank inerting system.
Background
The aircraft fuel tank inerting system can bleed air by adopting a bleed air pressurization mode, and an air preparation system needs to be provided with an air power turbine compression system in order to pressurize the bleed air. However, since the aircraft fuel tank inerting system cannot allow high-temperature gas to enter the fuel tank, the bleed air channel downstream of the turbo compressor may be closed in a non-commanded manner, which, once it occurs, may easily cause the turbo compressor to surge, which may damage the compressor and the piping upstream and downstream, and cause aircraft failure.
In order to solve the problem, some system schemes prevent the internal pressure of the compressor from continuously increasing by reducing the reaction time of the upstream and downstream valves of the turbocompressor and then rapidly reducing the gas entering the compressor, and the measures can slow down the continuous expansion of the harm caused by surge, but cannot fundamentally solve the problem.
By consulting the published patent, the existing product scheme is mostly a mode of constructing gas flow internal circulation inside the compressor, guiding high-pressure gas inside the compressor to perform limited flow and avoiding gas stagnation, and the measure can reduce the influence of surge to a certain extent, but has limited action effect.
The adverse effect caused by surging is difficult to be greatly eliminated by depending on the existing internal circulation mode, the shaft of the turbocompressor needs to be additionally reinforced, and the increase of the weight and the cost is possibly caused, so that the aerodynamic turbocompressor used in the current aircraft fuel tank inerting system still needs to be improved, and the surging problem of the turbocompressor is solved.
Disclosure of Invention
In order to overcome the defects in the prior art, the invention provides an aerodynamic turbine compression system, which comprises: an aerodynamic turbine comprising a turbine inlet through which the air enters and a turbine outlet through which the air exits; the compressor comprises a compressor inlet part and a compressor outlet part, the gas enters through the compressor inlet part, and the gas is discharged through the compressor outlet part after being pressurized; and a shaft connected between the aerodynamic turbine and the compressor; the air-powered turbine compression system further comprises an air bleed pipeline, the air bleed pipeline is provided with an introducing end and a leading-out end, the introducing end of the air bleed pipeline is communicated with the outlet portion of the compressor, the air bleed pipeline is provided with a one-way valve, when the compressor enters surge, the one-way valve is opened, and when the compressor exits the surge, the one-way valve is automatically closed.
According to an aspect of the invention, the outlet of the bleed line is connected to the turbine outlet or an extension line in fluid communication with the turbine outlet.
According to a further aspect of the invention, a throttle device is provided in the bleed air line downstream of the non-return valve. Preferably, the restriction device comprises a restrictor ring.
According to a further aspect of the invention, the bleed line has a smaller pipe diameter than the compressor outlet and the aerodynamic turbine outlet. Preferably, the diameter of the bleed air duct is in the range 2cm to 5 cm.
According to a further aspect of the invention, a seal is provided between the non-return valve and the bleed air line.
According to a further aspect of the invention, the material from which one or more of the non-return valve, the bleed air line and the seal is made comprises a high temperature resistant material that is resistant to high temperatures above 300 ℃.
According to a further aspect of the invention, the diameter of the one-way valve is the same as the diameter of the bleed air line.
The aerodynamic turbocompression system according to the present invention is preferably used in an aircraft fuel tank inerting system.
By adopting the air-powered turbine compression system for air entraining and pressurization of the aircraft fuel tank inerting system, which is disclosed by the invention, the air entraining pipeline with the one-way valve is arranged at the outlet part of the compressor, so that high-pressure gas can be quickly discharged when the compressor surges, the surge is remarkably relieved or eliminated, the damage to the compressor and the pipelines at the upstream and the downstream of the compressor is avoided, and the flight safety of the aircraft is ensured.
In addition, the bleed air pipeline is adopted to eliminate surge, and compared with other schemes, the scheme has low cost and convenient arrangement.
By providing the throttle device downstream of the non-return valve of the bleed air line, the bleed portion of the bleed air line is connected as close as possible to the outlet portion of the turbine without adversely affecting the turbine, the length of the bleed air line can be set short and, correspondingly, the weight of the added bleed air line can be as low as possible, which is particularly important for aircraft.
Drawings
The invention will be elucidated on the basis of exemplary embodiments shown in the schematic drawings, in which:
FIG. 1 shows a schematic view of an aerodynamic turbine compression system according to a preferred embodiment of the present invention.
Reference numerals
1 aerodynamic turbine
11 turbine inlet section
12 turbine outlet section
2 compressor
21 compressor outlet part
22 compressor inlet part
3 shaft
4 air-leading pipeline
5 one-way valve
6 flow-limiting ring
7 compressor outlet pipeline
8 turbine exhaust line
100 aerodynamic turbine compression system
Detailed Description
The present invention is further described in the following description with reference to specific embodiments and the accompanying drawings, wherein the details are set forth in order to provide a thorough understanding of the present invention, but it is apparent that the present invention can be embodied in many other forms different from those described herein, and it will be readily appreciated by those skilled in the art that the present invention can be implemented in many different forms without departing from the spirit and scope of the invention. Like parts in the drawings are indicated with the same or similar reference numerals.
FIG. 1 shows a schematic view of an aerodynamic turbine compression system 100 according to a preferred embodiment of the present invention. The anti-surge aerodynamic turbo compression system 100 according to the present invention mainly comprises an aerodynamic turbine 1 and a compressor 2. The rotating parts of the turbine 1 and the compressor 2 are connected by a shaft 3, so that the rotating movement of the rotating parts of the turbine 1 causes the rotating parts in the compressor 2 to rotate via the shaft 3.
The aerodynamic turbine 1 has a turbine shell, an impeller (not shown) located inside the turbine shell, a turbine inlet portion 11 and a turbine outlet portion 12. The turbine inlet portion 11 is generally disposed circumferentially of the turbine shell, the turbine outlet portion 12 is generally disposed axially of the turbine shell, and the impeller is disposed about the shaft within the turbine shell. When in use, air flowing at high speed enters the interior of the turbine shell through the turbine inlet part 11 and is discharged through the turbine outlet part 12, and the volume of the air expands to do work to drive the shaft 3 to rotate.
The compressor 2 has a compressor housing, an impeller (not shown) located inside the compressor housing, a compressor inlet portion 22, and a compressor outlet portion 21. The compressor inlet 22 may be disposed in an axial direction of the compressor housing, the compressor outlet 21 may be disposed in a circumferential direction of the compressor housing, and the impeller may be disposed along a circumference of the shaft inside the compressor housing. The gas enters the compressor 2 through the compressor inlet portion 22, and the gas rotates at a high speed in the compressor 2 as the impeller of the compressor 2 rotates, and the gas is pressurized. The pressurized gas is discharged through the compressor outlet 21 and is conveyed to the aircraft fuel tank inerting system through a connecting pipeline.
In order to reduce the surge of the compressor, in the aerodynamic turbocompressor system 100 according to the invention, a bleed air line 4 is provided which communicates with the compressor outlet, the end of the bleed air line 4 connected to the compressor outlet 21 forming an air flow intake, a portion of the air from the compressor outlet 21 being drawn off via the bleed air line 4 in the event of a surge. Another part of the gas in the compressor outlet section 21 will be sent out along the compressor outlet line 7.
In particular, the intake of the bleed air line 4 can be attached directly to the outlet at the compressor housing or to an outlet line 7 connected to the compressor outlet.
The bleed air line 4 further comprises an air flow outlet along the lower end of the bleed air line 4, the outlet of the bleed air line 4 preferably being connected to an exhaust air line connected to the turbine outlet portion 12, so that the exhaust air flow in the bleed air line 4 and the exhaust air flow of the turbine outlet portion 12 are merged together and discharged through a turbine exhaust air line 8 connected to the turbine outlet portion 12.
According to a preferred embodiment of the invention, in order to minimize unnecessary energy losses, it is required that the diameter of the bleed air line is not too large, so that an excessive flow of exhaust gas is avoided. The diameter of the bleed air line 4 should be smaller than the diameter of the compressor outlet portion 21. On the other hand, in the case where the outlet end of the bleed air line 4 is provided on an extension line connecting the turbine outlet portion 12, the diameter of the bleed air line 4 should also be smaller than the diameter of the turbine discharge line 8 of the turbine outlet portion 12. Preferably, the diameter of the bleed air line 4 may be set in the range of 2cm to 5 cm.
Since the ambient temperature at the location where the aerodynamic turbine compression system 100 is arranged in an aircraft is typically in the range of 200-300 ℃, the bleed air line 4 should be made of a material that can withstand high temperatures of 300 ℃, and the material of the bleed air line 4 may comprise, for example, stainless steel, to ensure that the bleed air line 4 can be adapted to such ambient temperatures.
The gas guiding pipeline 4 is also provided with a one-way valve 5 with a spring, the valve in the one-way valve 5 can be automatically opened under the gas pressure against the elastic force of the spring, and when the gas pressure is smaller than the elastic force of the spring, the valve can be freely closed. The opening pressure of the check valve 5 should be determined according to the surge boundary of the turbo compressor of the type and the turbine inlet pressure, the surge boundary of the turbo compressor is different for different types of aircraft, and the opening pressure of the check valve 5 is usually in the range of 50psig to 120 psig. Thus, when surge occurs, the pressure of the high pressure gas inside the compressor will open the check valve, allowing a portion of the gas to be rapidly discharged through the bleed line 4 toward the turbine outlet 12 and combined with the flow from the turbine outlet 12 and flow out.
The diameter of the non-return valve 5 is preferably the same as the diameter of the bleed air line 4. Likewise, to ensure that the use of the check valve 5 is not affected by the ambient temperature, the material from which the check valve 5 is made should also be able to withstand high temperatures of 300 ℃. The material of the check valve preferably comprises stainless steel.
Furthermore, a seal is provided between the non-return valve 5 and the bleed air line 4, so that high-pressure gas is prevented from leaking between the bleed air line 4 and the non-return valve 5 during normal operation. The sealing member should be able to withstand a high temperature of 300 c, and the material of the sealing member may be silicon rubber.
In order to prevent the high-pressure gas guided through the bleed air line 4 from adversely affecting the turbine, a restrictor ring 6 is preferably provided downstream of the non-return valve 5 to reduce the pressure and flow of the gas discharged through the bleed air line 4, thereby preventing the high-pressure gas stream from rapidly injecting exhaust gas at the turbine outlet. In general, the design principle of the restrictor ring 6 is that the outlet pressure of the restrictor ring 6 is not higher than 2 times of the turbine outlet pressure, and the outlet gas flow of the restrictor ring does not affect the exhaust gas emission of the turbine. On the other hand, since the restrictor ring 6 reduces the adverse effect of the compressed gas on the turbine, the outlet end of the bleed air duct 4 can thus be arranged close to the turbine body, so that the increased bleed air duct 4 can be as short as possible and correspondingly lighter in weight, which is particularly advantageous for aircraft.
In the following, the operation of the aerodynamic turbocompressor system 100 with an air bleed line 4 according to a preferred embodiment of the invention is exemplarily described.
When surge occurs in the compressor 2, the gas pressure in the compressor 2 rises sharply, the check valve 5 on the bleed air pipeline 4 is opened automatically under the action of high-pressure gas, the high-pressure gas in the compressor 2 is released rapidly through the bleed air pipeline 4, is limited by the flow limiting ring 6 and then is discharged to a position close to the turbine outlet 12, and the exhaust gas of the bleed air pipeline 4 is discharged to the outside of the aircraft body through the turbine discharge pipeline 8 connected to the turbine outlet 12.
As the high-pressure gas is continuously discharged, the pressure inside the compressor 2 is continuously reduced, and the compressor 2 gradually exits surge, thereby preventing the equipment including the compressor 2 and the front and rear pipes from being damaged. After the compressor 2 is out of surge, the pressure of the gas acting on the one-way valve 5 is reduced, the valve is automatically closed under the action of the spring, and the turbocompressor returns to normal operation.
In the normal operating mode of the system 100, the check valve 5 is always closed, and the seal between the check valve 5 and the bleed air line 4 ensures that the air between the compressor 2 and the turbine 1 is isolated.
The surge is not serious, and when the compressor normally works, air enters the turbine 1 through the turbine inlet 11 and is discharged through the turbine outlet 12, the volume expansion does work outwards to drive the shaft 3 to rotate, further the compressor 2 is driven to rotate, the gas enters the compressor 2 through the compressor inlet 22 and rotates at a high speed in the compressor 2, and after the gas pressurization is completed, the gas is discharged through the compressor outlet 21.
In other alternative embodiments, the outlet end of the bleed air line 4 may not be in communication with the turbine outlet portion 12, but may be communicated directly to the outside of the aircraft body to vent the bleed air line 4.
In other alternative embodiments, the outlet end of the bleed air line 4 is also connected to the outlet line of the turbine away from the turbine outlet 12, and the flow-restricting ring 6 may also be omitted in cases where the distance between the outlet end and the turbine outlet 12 is sufficiently far that the flow from the bleed air line 4 has a negligible effect on the turbine.
The aerodynamic turbocompression system according to the invention is particularly suitable for bleed air pressurization of an aircraft fuel tank inerting system. By providing the bleed air duct 5 with the non-return valve 5 at the compressor outlet portion 21, high-pressure gas can be discharged quickly when surge occurs in the compressor 2, and surge is reduced or eliminated significantly, thereby ensuring the flight safety of the aircraft.
In addition, the bleed air pipeline is adopted to eliminate surge, and compared with other means, the scheme is low in cost and convenient to arrange.
By providing the throttle device downstream of the non-return valve of the bleed air line, the bleed portion of the bleed air line is connected as close as possible to the outlet portion of the turbine without adversely affecting the turbine, the length of the bleed air line can be set short and, correspondingly, the weight of the added bleed air line can be as low as possible, which is particularly important for aircraft.
Although the present invention has been disclosed in terms of the preferred embodiment, it is not intended to limit the invention, and variations and modifications may be made by one skilled in the art without departing from the spirit and scope of the invention. Therefore, any modification, equivalent change and modification of the above embodiments according to the technical essence of the present invention are within the protection scope defined by the claims of the present invention, unless the technical essence of the present invention departs from the content of the present invention.
Claims (10)
1. An aerodynamic turbocompressor system comprising:
an aerodynamic turbine comprising a turbine inlet through which the air enters and a turbine outlet through which the air exits;
the compressor comprises a compressor inlet part and a compressor outlet part, gas enters through the compressor inlet part, and is discharged through the compressor outlet part after being pressurized; and
a shaft connected between the aerodynamic turbine and the compressor;
the air-powered turbine compression system is characterized by further comprising an air bleed pipeline, wherein the air bleed pipeline is provided with an introducing end and a leading-out end, the introducing end of the air bleed pipeline is communicated with the outlet portion of the compressor, the air bleed pipeline is provided with a one-way valve, when the compressor enters surge, the one-way valve is opened, and when the compressor exits surge, the one-way valve is automatically closed.
2. The aerodynamic turbocompressor system according to claim 1, wherein the lead-out of the bleed air line is connected to the turbine outlet or an extension line in fluid communication with the turbine outlet.
3. An aerodynamic turbocompression system according to claim 1 or 2, wherein a throttling means is provided in the bleed air line downstream of the check valve.
4. The aerodynamic turbocompressor system according to claim 3, wherein the throttling means comprise a restrictor ring.
5. The aerodynamic turbocompressor system according to claim 1,
and the pipe diameter of the air guide pipeline is smaller than the pipe diameters of the outlet part of the compressor and the outlet part of the aerodynamic turbine.
6. The aerodynamic turbocompressor system according to claim 5, wherein the diameter of the bleed air line is in the range of 2cm to 5 cm.
7. The aerodynamic turbocompressor system according to claim 1,
and a sealing element is arranged between the one-way valve and the bleed air pipeline.
8. The aerodynamic turbocompressor system according to claim 7,
the material from which one or more of the check valve, the bleed air line and the seal is made comprises a high temperature resistant material that is resistant to temperatures above 300 ℃.
9. The aerodynamic turbocompressor system according to claim 1, wherein the diameter of said one-way valve is the same as the diameter of the bleed air line.
10. The aerodynamic turbocompression system of claim 1, wherein said aerodynamic turbocompression system is used in an aircraft fuel tank inerting system.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210207659.9A CN114483612B (en) | 2022-03-04 | 2022-03-04 | Aerodynamic turbine compression system |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210207659.9A CN114483612B (en) | 2022-03-04 | 2022-03-04 | Aerodynamic turbine compression system |
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| CN114483612A true CN114483612A (en) | 2022-05-13 |
| CN114483612B CN114483612B (en) | 2024-01-05 |
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| US20170102148A1 (en) * | 2015-10-09 | 2017-04-13 | Dresser-Rand Company | System and method for operating a gas turbine assembly |
| CN105626266A (en) * | 2016-01-12 | 2016-06-01 | 中国科学院工程热物理研究所 | Anti-surge air escape energy recycling system of gas turbine |
| CN106762155A (en) * | 2016-12-15 | 2017-05-31 | 中国航空工业集团公司西安飞机设计研究所 | A kind of reverse-bootstrap air supply system based on turbocompressor |
| WO2019201033A1 (en) * | 2018-04-17 | 2019-10-24 | 沪东中华造船(集团)有限公司 | System for producing compressed air by utilizing waste heat of ship |
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| CN109921061A (en) * | 2019-03-27 | 2019-06-21 | 重庆长安汽车股份有限公司 | A kind of fuel cell air supply system and air supply method |
| CN111244506A (en) * | 2020-01-17 | 2020-06-05 | 擎能动力科技(苏州)有限公司 | New energy automobile fuel cell system, working method, hydrogen gas inlet flow calculation method and efficiency evaluation method |
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| CN112576366A (en) * | 2020-12-12 | 2021-03-30 | 贵州永红航空机械有限责任公司 | Two-wheeled pressure turbine cooler driven by air dynamic pressure bearing |
| CN113738675A (en) * | 2021-09-17 | 2021-12-03 | 南京磁谷科技股份有限公司 | Main motor air-cooled constant temperature system of magnetic suspension air compressor |
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