CN112607040A - Wall surface staggered inclined hole jet cooling technology for high-temperature part of aircraft - Google Patents
Wall surface staggered inclined hole jet cooling technology for high-temperature part of aircraft Download PDFInfo
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- CN112607040A CN112607040A CN202011625281.1A CN202011625281A CN112607040A CN 112607040 A CN112607040 A CN 112607040A CN 202011625281 A CN202011625281 A CN 202011625281A CN 112607040 A CN112607040 A CN 112607040A
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- inclined hole
- hole jet
- staggered
- jet cooling
- wall surface
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/08—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
Abstract
The invention relates to the technical field of high-temperature thermal protection, and particularly discloses a wall surface staggered inclined hole jet cooling technology for an aircraft high-temperature part, which is used for preventing the aircraft key part from working failure in a high-temperature environment. The cooling system comprises a fuel supply system, a wall surface staggered inclined hole jet flow cooling system and a fuel flow regulating and controlling system. Fuel in the fuel tank enters the jet cooling channel with the staggered inclined holes on the wall surface through the fuel pump, and the flow in the channel can be controlled by an automatic adjusting valve. The wall surface staggered inclined hole jet flow cooling channel can effectively inhibit the failure rate of high-temperature components of the aircraft, prevent the heat transfer deterioration phenomenon in the traditional cooling channel, and provide a new idea for the design of an aircraft cooling system.
Description
Technical Field
The invention relates to the technical field of high-temperature thermal protection, in particular to a wall surface staggered inclined hole jet flow cooling channel which utilizes staggered inclined holes to weaken the nonuniformity of heat flow, improve the heat transfer capacity and efficiently cool high-temperature parts of an aircraft.
Background
The hypersonic aircraft is a new concept air combat platform which is vigorously developed by all military and strong countries in the world due to the advantages of good stealth performance, high flight speed (more than Mach 5), wide attack range, large effective load and the like, the selection of a power device is crucial, and the specific impulse of a rocket engine, a turbojet engine, a sub-combustion ramjet engine and a super-combustion ramjet engine under different Mach numbers is compared, so that the specific impulse of a hydrogen fuel and hydrocarbon fuel air-breathing type engine is higher than that of a non-air-breathing type rocket engine; the flying speed of the turbine turbofan engine is generally lower than Mach 3 due to the temperature resistance of turbine blades; supersonic air flow in the scramjet engine is compressed through an air inlet channel to reduce the efficiency of the circulation process, and only the flight within the range of Mach 3-5 can be realized; further, hydrogen fuel is highly demanded in terms of storage and transportation, etc., while hydrogen fuel of low density causes an increase in structural volume and weight, and hydrocarbon fuel does not have the above-mentioned drawbacks.
Scramjet engines using hydrocarbon fuels as propellants are widely considered as core components of hypersonic aircraft. The scramjet engine flies at a hypersonic speed, the heat load of a combustion chamber is increased sharply due to the dual functions of combustion heat release and pneumatic heating, and when the flying Mach number reaches more than 8, the temperature and the heat flux density of the combustion chamber are respectively close to 4000K and 10-20MW/m2The heat-resisting temperature of the most advanced high-temperature-resistant composite materials (C/C and SiC) at present is only 2200K; therefore, the safety of the scramjet engine operating at high flight mach numbers cannot be guaranteed only by means of passive thermal insulation, and an active regenerative cooling technology using self-fuel as a coolant is considered to be one of the most effective approaches for solving the effective thermal control of the combustion chamber.
The pressure in the combustion chamber is generally between 3.5 and 7MPa and exceeds the critical pressure (2 to 3MPa) of the hydrocarbon fuel; therefore, the regenerative cooling absorbs the heat transferred to the wall surface of the combustion chamber by utilizing the heat convection of the supercritical hydrocarbon fuel in the cooling channel, so as to achieve the purpose of reducing the wall surface temperature of the combustion chamber.
The thermal physical property of the supercritical hydrocarbon fuel is changed violently along with the rise of the temperature, particularly near the quasi-critical temperature; along with the increase of the heat absorption capacity, the temperature of the hydrocarbon fuel is increased, the cracking reaction is carried out, small molecular products are generated, and the thermophysical property is changed again. The supercritical hydrocarbon fuel is often accompanied with heat transfer deterioration phenomenon in the convection heat exchange process under the non-cracking/cracking working condition due to the two points, and the existing single-layer cooling channel cannot effectively reduce the wall surface temperature of a combustion chamber or effectively inhibit the temperature rise of key parts of an aircraft.
Disclosure of Invention
In order to overcome the defects of the prior art, the invention provides a wall surface staggered inclined hole jet flow cooling technology for an aircraft high-temperature part, which utilizes wall surface staggered inclined holes to strengthen the heat transfer of the high-temperature part and effectively reduce the temperature of the aircraft high-temperature part.
In order to achieve the above purposes, the invention adopts the technical scheme that: the wall surface staggered inclined hole jet flow cooling technology for the high-temperature part of the aircraft comprises a fuel supply system, a wall surface staggered inclined hole jet flow cooling system and a fuel flow regulating and controlling system, wherein the fuel supply system is connected with the wall surface staggered inclined hole jet flow cooling system, and the fuel flow regulating and controlling system is also connected with the wall surface staggered inclined hole jet flow cooling system and distributes the fuel flow of the whole circulation loop.
Further, the driving power of the fuel pump in the fuel supply system is a battery device or a fluid expansion device.
Further, the channels in the wall surface staggered inclined hole jet flow cooling system refer to smooth channels, corrugated channels or other combinations of related enhanced heat exchange channels.
Further, the shape of the channel in the wall surface staggered inclined hole jet flow cooling system can be rectangular, triangular, circular and other relevant geometric shapes.
Further, the angle of the inclined holes in the wall surface staggered inclined hole jet flow cooling system is not limited to 30 degrees, 45 degrees, 60 degrees and can be other angles.
Further, the cross section of the inclined hole in the wall surface staggered inclined hole jet flow cooling system can be circular, oval, rectangular and other relevant geometric shapes.
Further, the inclined holes in the wall surface staggered inclined hole jet flow cooling system can be symmetrically distributed or asymmetrically distributed.
Compared with the prior art, the invention has the beneficial effects that:
through the reasonable configuration and layout of the wall surface staggered inclined hole jet flow cooling channels, the heat transfer deterioration phenomenon is prevented, the wall surface temperature of the combustion chamber is effectively reduced or the temperature rise of key parts of the aircraft is effectively restrained, and therefore the aim of high-Mach-number flight is achieved.
Drawings
FIG. 1 is a schematic diagram of a cooling cycle for a combustion chamber wall of a hypersonic aircraft.
FIG. 2 is a layout view of staggered jet inclined holes on the wall surface of a cooling channel
FIG. 3 is a wall shape of a wall staggered inclined hole jet cooling channel.
FIG. 4 is a cross-sectional view of a wall-staggered inclined hole jet cooling channel.
FIG. 5 is a symmetrical arrangement of wall staggered inclined hole jet cooling channels.
FIG. 6 is an asymmetric arrangement of wall staggered angled hole jet cooling channels.
In the figure:
1. a fuel tank; 2. a fuel pump; 3. a regulating device; 4. wall surface staggered inclined hole jet cooling channels; 5 regulating and controlling the device; 6. a combustion chamber; 7. a single channel; 8. wall fins; 9. staggered inclined holes; 10. smoothing the cooling channel; 11. a corrugated cooling channel; 12. a circular channel cross-section; 13. an elliptical channel cross-section; 14. a rectangular channel cross-section.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
The invention is described in further detail below with reference to the attached drawing figures:
FIG. 1 shows a cooling cycle diagram for the combustion chamber wall of a hypersonic aircraft. As shown in the figure, fuel in a fuel tank 1 enters a wall surface staggered inclined hole jet flow cooling channel 4 through a fuel pump 2, the flow rate of the channel is controlled by a regulating device 3, the operation pressure is controlled by a regulating device 5, low-temperature fuel absorbs wall surface heat, and high-temperature fuel is pumped into a combustion chamber for combustion.
The shape of the wall of the jet cooling channel with the wall surface of the staggered inclined holes in the embodiment is shown in figure 3, namely a smooth cooling channel 10, a corrugated cooling channel 11 or other combinations of related enhanced heat exchange channels.
The cross-sectional shape of the wall-staggered inclined hole jet cooling channel in this embodiment is shown in fig. 4, i.e., a circular channel cross-section 12, an elliptical channel cross-section 13, a rectangular channel cross-section 14, or other relevant geometries.
The symmetrically arranged wall surface staggered inclined hole jet flow cooling channels in the embodiment are shown in the attached figure 5, namely, staggered inclined holes on the wall surface rib piece 8 are regularly distributed.
The jet cooling channels with staggered inclined holes on the wall surface in the asymmetric arrangement in the embodiment are shown in fig. 6, that is, the staggered inclined holes on the wall surface fins 8 are distributed irregularly.
The invention provides a wall surface staggered inclined hole jet flow cooling technology for high-temperature components of an aircraft, which prevents heat transfer deterioration through reasonable design and layout of wall surface staggered inclined holes, effectively reduces the wall surface temperature of a combustion chamber or effectively restrains the temperature rise of key components of the aircraft, thereby achieving the aim of high-Mach-number flight.
The above is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and various modifications and changes will occur to those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (8)
1. A wall staggered inclined hole jet cooling technique for high temperature components of an aircraft, comprising: the fuel supply system consists of a fuel tank and a fuel pump and provides working media for the wall surface staggered inclined hole jet flow cooling system; the wall surface staggered inclined hole jet cooling system consists of a plurality of channels, and staggered inclined holes are formed in wall surface ribs and are used for absorbing heat of high-temperature components; the fuel flow regulating and controlling system consists of an automatic control valve or a related flow regulating and controlling device.
2. The staggered wall hole jet cooling technique for high temperature components of an aircraft of claim 1, wherein said fuel supply system is connected to the staggered wall hole jet cooling system, and the driving power of the fuel pump is a battery device or a fluid expansion device.
3. The technology of claim 1, wherein the channels in the wall-staggered inclined hole jet cooling system are smooth channels, corrugated channels or a combination of other related heat exchange enhancement channels.
4. The staggered wall inclined hole jet cooling technique for high-temperature components of aircraft according to claim 1, wherein the shape of the channels in the staggered wall inclined hole jet cooling system can be rectangular, triangular, circular or other relevant geometric shapes.
5. The technique of claim 1, wherein the angle of the inclined holes in the wall surface staggered inclined hole jet cooling system is not limited to 30 °, 45 °, 60 ° or other angles.
6. The staggered wall inclined hole jet cooling technology for high-temperature components of aircraft according to claim 1, wherein the cross section of the inclined holes in the staggered wall inclined hole jet cooling system can be circular, oval, rectangular or other relevant geometric shapes.
7. The staggered wall inclined hole jet cooling technology for high-temperature components of aircraft according to claim 1, wherein the arrangement of inclined holes in the staggered wall inclined hole jet cooling system can be symmetrical distribution or other asymmetrical distribution.
8. The staggered wall inclined hole jet cooling technology for high-temperature components of aircrafts as claimed in claim 1, wherein the channels in the staggered wall inclined hole jet cooling system can be made by brazing, etching, 3D printing or other related technical means.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011625281.1A CN112607040A (en) | 2020-12-31 | 2020-12-31 | Wall surface staggered inclined hole jet cooling technology for high-temperature part of aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011625281.1A CN112607040A (en) | 2020-12-31 | 2020-12-31 | Wall surface staggered inclined hole jet cooling technology for high-temperature part of aircraft |
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| CN112607040A true CN112607040A (en) | 2021-04-06 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202011625281.1A Pending CN112607040A (en) | 2020-12-31 | 2020-12-31 | Wall surface staggered inclined hole jet cooling technology for high-temperature part of aircraft |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102216159A (en) * | 2008-10-07 | 2011-10-12 | 空中客车运营简化股份公司 | Air intake arrangement for an aircraft |
| CN203147824U (en) * | 2011-10-26 | 2013-08-21 | 斯奈克玛 | Annular wall of combustion chamber of turbo engine, combustion chamber of the turbo engine and the turbo engine |
| CN108223022A (en) * | 2018-01-04 | 2018-06-29 | 沈阳航空航天大学 | A kind of turbulence structure in array jetting cooling |
| CN109469512A (en) * | 2019-01-04 | 2019-03-15 | 西北工业大学 | A kind of chiasma type X air film hole cooling structure for turbo blade |
| CN111578310A (en) * | 2020-04-30 | 2020-08-25 | 南京理工大学 | Air film cooling hole structure for turboshaft engine |
| CN111878238A (en) * | 2020-07-23 | 2020-11-03 | 西北工业大学 | Double-layer cooling channel for reducing temperature of aircraft component |
-
2020
- 2020-12-31 CN CN202011625281.1A patent/CN112607040A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102216159A (en) * | 2008-10-07 | 2011-10-12 | 空中客车运营简化股份公司 | Air intake arrangement for an aircraft |
| CN203147824U (en) * | 2011-10-26 | 2013-08-21 | 斯奈克玛 | Annular wall of combustion chamber of turbo engine, combustion chamber of the turbo engine and the turbo engine |
| CN108223022A (en) * | 2018-01-04 | 2018-06-29 | 沈阳航空航天大学 | A kind of turbulence structure in array jetting cooling |
| CN109469512A (en) * | 2019-01-04 | 2019-03-15 | 西北工业大学 | A kind of chiasma type X air film hole cooling structure for turbo blade |
| CN111578310A (en) * | 2020-04-30 | 2020-08-25 | 南京理工大学 | Air film cooling hole structure for turboshaft engine |
| CN111878238A (en) * | 2020-07-23 | 2020-11-03 | 西北工业大学 | Double-layer cooling channel for reducing temperature of aircraft component |
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Application publication date: 20210406 |