CN114483612B - Aerodynamic turbine compression system - Google Patents
Aerodynamic turbine compression system Download PDFInfo
- Publication number
- CN114483612B CN114483612B CN202210207659.9A CN202210207659A CN114483612B CN 114483612 B CN114483612 B CN 114483612B CN 202210207659 A CN202210207659 A CN 202210207659A CN 114483612 B CN114483612 B CN 114483612B
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- China
- Prior art keywords
- compressor
- bleed air
- turbine
- compression system
- aerodynamic
- Prior art date
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Links
- 230000006835 compression Effects 0.000 title claims abstract description 28
- 238000007906 compression Methods 0.000 title claims abstract description 28
- 239000002828 fuel tank Substances 0.000 claims abstract description 11
- 239000000463 material Substances 0.000 claims description 9
- 239000012530 fluid Substances 0.000 claims description 2
- 238000007789 sealing Methods 0.000 claims 1
- 230000000694 effects Effects 0.000 description 6
- 230000002411 adverse Effects 0.000 description 3
- 238000011144 upstream manufacturing Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 229910001220 stainless steel Inorganic materials 0.000 description 2
- 239000010935 stainless steel Substances 0.000 description 2
- 230000007812 deficiency Effects 0.000 description 1
- 238000002347 injection Methods 0.000 description 1
- 239000007924 injection Substances 0.000 description 1
- 230000035484 reaction time Effects 0.000 description 1
- 229920002379 silicone rubber Polymers 0.000 description 1
- 239000004945 silicone rubber Substances 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
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
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Life Sciences & Earth Sciences (AREA)
- Sustainable Development (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention provides an aerodynamic turbine compression system. The system comprises an aerodynamic turbine, a compressor and a shaft connected between the aerodynamic turbine and the compressor, wherein the aerodynamic turbine compression system further comprises a bleed air pipeline, the bleed air pipeline is provided with a lead-in end and a lead-out end, the lead-in end of the bleed air pipeline is communicated with an outlet part of the compressor, the bleed air 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. The aerodynamic turbine compression system according to the invention is particularly suitable for bleed air pressurization of aircraft fuel tank inerting systems. The bleed air pipeline can rapidly discharge high-pressure gas when the compressor surge occurs, so that the surge is obviously lightened or eliminated, and the flight safety of the aircraft 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 bleed air on a fuel tank inerting system.
Background
The aircraft fuel tank inerting system may bleed air in a bleed air pressurization manner, and the air preparation system may be configured with an air turbine compression system in order to pressurize the bleed air. However, because the aircraft fuel tank inerting system cannot allow high temperature gas to enter the fuel tank, the bleed air passage downstream of the turbocompressor may be undesirably shut down, which, once it occurs, is highly susceptible to causing the turbocompressor to surge, thereby damaging the compressor and the upstream and downstream piping and causing aircraft failure.
In view of this problem, some system solutions prevent the internal pressure of the compressor from continuously increasing by reducing the reaction time of the valves upstream and downstream of the turbocompressor, and thus by rapidly reducing the gas entering the compressor, which can slow down the continuous expansion of the damage caused by the surge, but cannot fundamentally solve the problem.
By referring to the disclosed patent, the existing product schemes are mostly a mode of constructing gas flow internal circulation in the compressor, guiding high-pressure gas in the compressor to carry out limited flow, avoiding gas stagnation, and the measures can reduce the influence of surge to a certain extent, but have limited action effect.
Depending on the existing internal circulation mode, the adverse effect caused by surge is difficult to be eliminated greatly, the shaft of the turbine compressor needs to be additionally reinforced, and the weight and the cost are possibly increased, so that the aerodynamic turbine compression system used in the current aircraft fuel tank inerting system still needs to be improved, and the problem of surge of the turbine compressor is solved.
Disclosure of Invention
To overcome the deficiencies in the prior art, the present invention provides an aerodynamic turbine compression system comprising: an aerodynamic turbine comprising a turbine inlet portion through which the air enters and a turbine outlet portion through which the air exits; a compressor including a compressor inlet portion through which the gas enters and a compressor outlet portion through which the gas is discharged after being pressurized; and a shaft connected between the aerodynamic turbine and the compressor; the air turbine compression system further comprises a bleed air pipeline, the bleed air pipeline is provided with a lead-in end and a lead-out end, the lead-in end of the bleed air pipeline is communicated with the outlet part of the compressor, the bleed air 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.
According to one aspect of the invention, the outlet end of the bleed air line is connected to the turbine outlet portion or to an extension line in fluid communication with the turbine outlet portion.
According to a further aspect of the invention, a throttle device is provided in the bleed air line downstream of the one-way flap. Preferably, the restriction device comprises a restrictor ring.
According to a further aspect of the invention, the bleed air line has a smaller pipe diameter than the compressor outlet and the aerodynamic turbine outlet. Preferably, the bleed air line has a diameter in the range of 2cm to 5 cm.
According to a further aspect of the invention, a seal is provided between the check valve and the bleed air line.
According to a further aspect of the invention, 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 high temperatures above 300 ℃.
According to a further aspect of the invention, the diameter of the check valve is the same as the diameter of the bleed air line.
The aerodynamic turbine compression system according to the invention is preferably used in an aircraft fuel tank inerting system.
By adopting the aerodynamic turbine compression system for air entraining and pressurizing of the aircraft fuel tank inerting system, provided 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 rapidly discharged when the surge of the compressor occurs, the surge is obviously lightened or eliminated, the compressor and the upstream and downstream pipelines thereof are prevented from being damaged, and the flight safety of an 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 arranging the throttle downstream of the one-way flap of the bleed air line, in this way the outlet of the bleed air line is as close to the connection to the turbine outlet as possible, but does not have a negative effect on the turbine, the length of the bleed air line can be made very short, and correspondingly the weight of the added bleed air line can be as small 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 illustrates a schematic diagram of an aerodynamic turbine compression system according to a preferred embodiment of the present invention.
Reference numerals
1. Aerodynamic turbine
11. Turbine inlet
12. Turbine outlet
2. Compressor with a compressor body having a rotor with a rotor shaft
21. Compressor outlet
22. Compressor inlet
3. Shaft
4. Bleed air 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 will be further described with reference to specific embodiments and drawings, in which more details are set forth in the following description in order to provide a thorough understanding of the present invention, but it will be apparent that the present invention can be embodied in many other forms than described herein, and that those skilled in the art may make similar generalizations and deductions depending on the actual application without departing from the spirit of the present invention, and therefore should not be construed to limit the scope of the present invention in terms of the content of this specific embodiment. Like parts in the drawings are designated by the same or similar reference numerals.
FIG. 1 illustrates a schematic diagram of an aerodynamic turbine compression system 100 according to a preferred embodiment of the present invention. The anti-surge aerodynamic turbine compression system 100 according to the 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 rotates the rotating parts in the compressor 2 by the shaft 3.
The aerodynamic turbine 1 has a turbine housing, an impeller (not shown) located inside the turbine housing, a turbine inlet portion 11, and a turbine outlet portion 12. The turbine inlet 11 is generally disposed in the circumferential direction of the turbine casing, the turbine outlet 12 is generally disposed in the axial direction of the turbine casing, and the impeller is disposed around the shaft inside the turbine casing. In use, air flowing at a high speed enters the inside of the turbine shell through the turbine inlet 11, is discharged through the turbine outlet 12, and performs work by volume expansion of the air 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 provided in the axial direction of the compressor housing, the compressor outlet 21 may be provided in the circumferential direction of the compressor housing, and the impeller may be provided in the compressor housing along the circumference of the shaft. Gas enters the compressor 2 through the compressor inlet portion 22, and the gas rotates at a high speed within the compressor 2 as the impeller of the compressor 2 rotates, with the gas being 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 turbine compression 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 constituting the air flow inlet through which a portion of the air from the compressor outlet 21 is led out when surge occurs. Another part of the gas at the compressor outlet 21 will then be sent out along the compressor outlet line 7.
In particular, the inlet end of the bleed air line 4 can be attached directly to the outlet section at the compressor housing or to the outlet line 7 which is connected to the compressor outlet section.
The bleed air line 4 also 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 line connected to the turbine outlet 12, so that the exhaust air flow in the bleed air line 4 merges with the exhaust air flow of the turbine outlet 12 and is discharged through the turbine exhaust line 8 which is connected to the turbine outlet 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 pipe diameter of the bleed air line 4 should be smaller than the diameter of the compressor outlet 21. On the other hand, in the case of the outlet end of the bleed air line 4 being arranged via an extension line which connects the turbine outlet 12, the diameter of the bleed air line 4 should also be smaller than the diameter of the turbine outlet 12 turbine discharge line 8. Preferably, the diameter of the bleed air line 4 can be set in the range of 2cm to 5 cm.
Since the ambient temperature at the location of the aircraft where the aerodynamic turbine compression system 100 is located is typically in the range of 200-300 ℃, the material from which the bleed air line 4 is made should be able to withstand high temperatures of 300 ℃, the material of the bleed air line 4 may for example comprise stainless steel to ensure that the bleed air line 4 is able to adapt to such ambient temperatures.
The bleed air 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 against the elasticity of the spring under the gas pressure, and when the gas pressure is smaller than the elasticity of the spring, the valve is freely closed. The opening pressure of the check valve 5 should be determined based on the surge margin and turbine inlet pressure of the type of turbocompressor, which is different for different types of aircraft, and typically the opening pressure of the check valve 5 is 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 so that a portion of the gas is rapidly discharged through the bleed air line 4 towards the turbine outlet 12 and merges with the gas flow from the turbine outlet 12.
The diameter of the non-return flap 5 is preferably the same as the diameter of the bleed air line 4. Likewise, in order 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 the high temperature of 300 ℃. The material of the check valve preferably comprises stainless steel.
In addition, a seal is provided between the non-return flap 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 flap 5 during normal operation. The seal should be able to withstand high temperatures of 300 c and the material of the seal may be silicone rubber.
In order to prevent the turbine from being adversely affected by the high-pressure gas guided through the bleed air line 4, a restrictor ring 6 is preferably provided downstream of the check valve 5 to reduce the pressure and flow of the gas discharged through the bleed air line 4, so that the injection of the high-pressure gas stream into the turbine outlet exhaust gas is avoided. Typically, the design principle of the restrictor ring 6 is that the outlet pressure of the restrictor ring 6 must not be 2 times higher than the turbine outlet pressure, and the restrictor ring outlet air flow must not affect the exhaust of the turbine exhaust. On the other hand, since the flow restriction ring 6 reduces the adverse effect of the compressed gas on the turbine, the outlet end of the bleed air line 4 can thus be arranged close to the turbine body, so that the length of the increased bleed air line 4 can be as short as possible, with a corresponding weight saving, which is particularly advantageous for aircraft.
Hereinafter, the operation of the aerodynamic turbine compression system 100 with bleed air line 4 according to a preferred embodiment of the invention will be 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 automatically opened under the action of high-pressure gas, the high-pressure gas in the compressor 2 is rapidly released through the bleed air pipeline 4, the high-pressure gas is discharged to a position close to the turbine outlet 12 after being limited by the flow limiting ring 6, 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.
With the continuous discharge of high pressure gas, the pressure inside the compressor 2 is reduced continuously, and the compressor 2 gradually exits from surge, so that the damage to equipment comprising the compressor 2 and front and rear pipelines is avoided. After the compressor 2 exits 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 turbo compressor returns to normal operation.
In the normal operating mode of the system 100, the non-return flap 5 is always in a closed state, and the seal between the non-return flap 5 and the bleed air line 4 ensures that the gas between the compressor 2 and the turbine 1 is in an isolated state.
When the compressor works normally, air enters the turbine 1 through the turbine inlet 11 and is discharged through the turbine outlet 12, the volume expansion acts externally to drive the shaft 3 to rotate, and then the compressor 2 is driven to rotate, gas enters the compressor 2 through the compressor inlet 22 and rotates at a high speed in the compressor 2, and after the gas is pressurized, the gas is discharged through the compressor outlet 21.
In other alternative embodiments, the outlet end of the bleed air line 4 may also be connected not to the turbine outlet 12 but directly outside the aircraft body, the gas in the bleed air line 4 being discharged.
In other alternative embodiments, the outlet of the bleed air line 4 is also connected to the outlet line of the turbine remote from the turbine outlet 12, the restrictor ring 6 also being omitted in the event that the distance between the outlet and the turbine outlet 12 is sufficiently great that the air flow from the bleed air line 4 has negligible effect on the turbine.
The aerodynamic turbine compression system according to the invention is particularly suitable for bleed air pressurization of aircraft fuel tank inerting systems. By providing the bleed air line 5 with a check valve 5 at the compressor outlet 21, high-pressure gas can be rapidly discharged when a surge of the compressor 2 occurs, and the surge is significantly reduced or eliminated, 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 has low cost and convenient arrangement.
By arranging the throttle downstream of the one-way flap of the bleed air line, in this way the outlet of the bleed air line is as close to the connection to the turbine outlet as possible, but does not have a negative effect on the turbine, the length of the bleed air line can be made very short, and correspondingly the weight of the added bleed air line can be as small as possible, which is particularly important for aircraft.
While the invention has been described in terms of preferred embodiments, it is not intended to be limiting, but rather to the invention, as will occur to those skilled in the art, without departing from the spirit and scope of the invention. Therefore, any modification, equivalent variation and modification of the above embodiments according to the technical substance of the present invention fall within the protection scope defined by the claims of the present invention.
Claims (7)
1. An aerodynamic turbine compression system for an aircraft fuel tank inerting system, comprising:
an aerodynamic turbine comprising a turbine inlet portion through which the air enters and a turbine outlet portion through which the air exits;
a compressor including a compressor inlet portion through which gas enters and a compressor outlet portion through which the gas is discharged after being pressurized; and
A shaft connected between the aerodynamic turbine and the compressor;
the air turbine compression system is characterized by further comprising a bleed air pipeline, wherein the bleed air pipeline is provided with a lead-in end and a lead-out end, the lead-in end of the bleed air pipeline is communicated with the outlet part of the compressor, the bleed air pipeline is provided with a one-way valve, when the compressor enters surge, the one-way valve is opened, when the compressor exits surge, the one-way valve is automatically closed,
the outlet end of the bleed air line is connected to the turbine outlet portion or an extension line in fluid communication with the turbine outlet portion, and a throttle device is provided in the bleed air line downstream of the one-way flap.
2. The aerodynamic turbine compression system of claim 1 wherein the throttle device comprises a restrictor ring.
3. The aerodynamic turbine compression system of claim 1,
the pipe diameter of the air entraining pipeline is smaller than the pipe diameters of the outlet part of the compressor and the outlet part of the aerodynamic turbine.
4. The aerodynamic turbine compression system of claim 3, wherein the bleed air line has a diameter in the range of 2cm to 5 cm.
5. The aerodynamic turbine compression system of claim 1,
and a sealing piece is arranged between the one-way valve and the air entraining pipeline.
6. The aerodynamic turbine compression system of claim 5,
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 high temperatures above 300 ℃.
7. The aerodynamic turbine compression system of claim 1 wherein the one-way valve has an identical path to the bleed air line pipe diameter.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210207659.9A CN114483612B (en) | 2022-03-04 | 2022-03-04 | Aerodynamic turbine compression system |
Applications Claiming Priority (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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| Publication Number | Publication Date |
|---|---|
| CN114483612A CN114483612A (en) | 2022-05-13 |
| CN114483612B true CN114483612B (en) | 2024-01-05 |
Family
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210207659.9A Active CN114483612B (en) | 2022-03-04 | 2022-03-04 | Aerodynamic turbine compression system |
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| CN103244258A (en) * | 2013-05-13 | 2013-08-14 | 宁波威孚天力增压技术有限公司 | Turbocharging system applicable to single cylinder diesel |
| CN104832221A (en) * | 2015-03-24 | 2015-08-12 | 清华大学 | Anti-surge system for turbocharging |
| 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 |
| CN209087991U (en) * | 2018-11-27 | 2019-07-09 | 广州汽车集团股份有限公司 | A kind of fuel cell and its air supply system |
| 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 |
| CN111952635A (en) * | 2020-08-18 | 2020-11-17 | 东风汽车集团有限公司 | Adjustable air supply device for fuel cell system |
| 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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