Composite cooling method and cooling device for aircraft

Abstract

The invention discloses a composite cooling method and cooling device for an aircraft, and the composite cooling method comprises the following steps: respectively injecting a gas medium at the leading edge of the aircraft and a fuselage located at the downstream of the leading edge, and arranging the gas medium injected at the leading edge to push shock waves away from the outer surface of the leading edge so as to reduce the heat at the leading edge, the mixed gas is formed after the mixed gas is mixed with the incoming flow; and the mixed gas flows in the downstream fuselage direction along with the incoming flow and is mixed with the gas medium sprayed at the fuselage again, so that boundary layer transition of the fuselage is inhibited and delayed. According to the composite cooling method, the capability of inhibiting transition of the fuselage boundary layer can be remarkably improved, and transition of the boundary layer can be effectively delayed, so that heat flow of the fuselage area can be remarkably reduced, and heat accumulation of the long-endurance aircraft at the fuselage is reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft cooling, and in particular to a composite cooling method and a cooling device for an aircraft. Background Art

[0002] In order to obtain high lift-to-drag ratio characteristics, high-speed aircraft usually adopt a pointed cone structure. However, when the above-mentioned high-speed aircraft flies for a long time in the atmosphere, not only will the pointed cone on the aircraft and the small-radius blunt leading edge structure including the wing leading edge, the air inlet lip and the vertical tail leading edge be subjected to severe aerodynamic heating, but also a large area of ​​the fuselage will be subjected to long-term heat accumulation, which seriously affects the service life of the thermal protection structure on the aircraft. Therefore, it is urgent to reduce the aerodynamic heating of the surface of high-speed aircraft.

[0003] In the prior art, in order to reduce the aerodynamic heating temperature of high-speed aircraft, a reverse jet cooling method is often used, that is, a nozzle is set at the head of the aircraft or the windward passivated leading edge, and the nozzle sprays gas in the opposite direction of the incoming flow. On the one hand, the injected gas can push the shock wave away from the surface of the object to reduce drag and reduce heat flow. On the other hand, the injected low-temperature gas can exchange heat with the surrounding to absorb part of the heat, thereby improving the aerodynamic heating environment of the area covered by the jet on the aircraft. The above-mentioned reverse jet cooling method has a good cooling effect on the small radius passivated leading edge structure, and can be applied to high-speed aircraft with a nose cone structure.

[0004] Although the above method can reduce the temperature of the small-radius blunt leading edge structure on the aircraft, the airflow temperature of the small-radius blunt leading edge structure will increase when the gas flows to the large area of ​​the fuselage located downstream. The large area of ​​the fuselage is still subjected to severe aerodynamic heating and long-term heat accumulation. The aerodynamic heating problem of the aircraft has not been effectively solved. Summary of the invention

[0005] The present invention provides a composite heat reduction device for an aircraft, which is used to solve the aerodynamic heating problem of the aircraft in the prior art, and is particularly used to improve the aerodynamic heating environment of a small-radius blunt leading edge facing the wind on the aircraft and a large area on the fuselage downstream of the leading edge.

[0006] In one aspect, the present invention provides a composite heat reduction method for an aircraft, the composite heat reduction method comprising the following steps:

[0007] The gas medium is sprayed at the leading edge of the aircraft and the fuselage downstream of the leading edge respectively, and the gas medium sprayed at the leading edge is configured to push the shock wave away from the outer surface of the leading edge to reduce the heat at the leading edge, and to form a mixed gas after mixing with the incoming flow;

[0008] The mixed gas flows along the incoming flow toward the downstream fuselage and is mixed again with the gas medium injected at the fuselage, thereby suppressing and delaying the boundary layer transition of the fuselage.

[0009] Optionally, the gas medium sprayed at the leading edge of the aircraft has a spraying direction that is opposite to the incoming flow direction.

[0010] Optionally, the gas medium sprayed onto the fuselage of the aircraft has a spraying direction along the normal direction of the fuselage wall at the spraying location.

[0011] Optionally, the gas medium is a heavy gas having a density greater than that of air.

[0012] Preferably, the gas medium is carbon dioxide or nitrogen or a mixture thereof.

[0013] In another aspect, the present invention provides a heat reduction device for implementing the composite heat reduction method for the above-mentioned aircraft, the heat reduction device comprising:

[0014] A first injection hole is arranged at the leading edge of the aircraft, and the gas medium injected by the first injection hole is suitable for pushing the shock wave away from the outer surface of the leading edge and mixing with the incoming flow to form a mixed gas;

[0015] A second injection hole is arranged on the fuselage of the aircraft, and along the flow direction of the mixed gas, the second injection hole is located downstream of the first injection hole, and the mixed gas upstream flows through the second injection hole and is mixed again with the gas medium injected by the second injection hole;

[0016] An air supply component is connected to the first injection hole and the second injection hole, and is used to supply air to the first injection hole and the second injection hole.

[0017] Furthermore, the spraying direction of the first spray hole is set to be opposite to the incoming flow direction.

[0018] Optionally, the spraying direction of the second spray hole is set to be along the normal direction of the fuselage wall where the second spray hole is located.

[0019] Optionally, the air supply assembly comprises:

[0020] A gas storage tank, used for storing the gas medium;

[0021] A power pump, whose air inlet is connected to the air storage tank, and whose air outlet is connected to the delivery pipeline and respectively connected to the first injection hole and the second injection hole through the delivery pipeline.

[0022] Preferably, a control valve is provided on the delivery pipeline.

[0023] Compared with the prior art, the advantages of the present invention are:

[0024] In the embodiment of the present invention, the composite heat reduction method of the aircraft is as follows: a heavy gas, such as carbon dioxide, is sprayed at the leading edge of the aircraft in the direction opposite to the incoming flow. The sprayed carbon dioxide can push the shock wave away from the outer surface of the leading edge to reduce the aerodynamic heat at the leading edge; then, the carbon dioxide sprayed at the leading edge is mixed with the incoming flow and flows downstream with the incoming flow. Since the concentration of carbon dioxide in the mixed gas is higher than the concentration of carbon dioxide in the incoming flow before mixing, the mixed gas has the ability to suppress the transition when flowing to the fuselage downstream. Therefore, it can suppress the transition of the fuselage boundary layer to a certain extent and reduce the heat flow in the fuselage area. On this basis, carbon dioxide is sprayed along the normal direction of the fuselage wall at the fuselage of the aircraft and at a position downstream of the leading edge. The carbon dioxide sprayed at the fuselage is mixed again with the incoming flow mixed with carbon dioxide from the leading edge, so that the concentration of carbon dioxide in the mixed gas is increased again, and its ability to suppress the transition of the fuselage boundary layer is further improved, and the boundary layer transition is delayed, thereby significantly reducing the heat flow in the fuselage area and reducing the heat accumulation at the fuselage of the long-flight aircraft.

[0025] The above-mentioned composite heat reduction method combines reverse jet technology and heavy gas, which can not only reduce the structural heat flux at the leading edge, but also reduce the heat flux in a large area of ​​the fuselage downstream of the leading edge on the aircraft by suppressing the boundary layer transition. Therefore, it can take into account the heat accumulation of the local high heat flux area at the leading edge and the large area of ​​the fuselage, thereby reducing the overall aerodynamic heating of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Hereinafter, the present invention will be described in more detail based on embodiments and with reference to the accompanying drawings.

[0027] Figure 1 is a schematic structural diagram of a composite heat reduction device for an aircraft in an embodiment of the present invention;

[0028] Figure 2 It is a schematic diagram of the principle that the leading edge ejects gas medium in the opposite direction of the incoming flow to push the shock wave away from the outer surface of the leading edge;

[0029] Figure 3 It is a schematic diagram of the ejection of heavy gas in the normal direction of the fuselage wall over a large area of ​​the fuselage downstream of the leading edge;

[0030] Reference numerals:

[0031] 1. Leading edge; 2. Fuselage;

[0032] 3. The first injection hole; 4. The second injection hole;

[0033] 5. Gas storage tank; 6. Power pump;

[0034] 71. a first delivery branch pipe; 72. a second delivery branch pipe;

[0035] 81. a first drainage tube; 82. a second drainage tube;

[0036] 91. First control valve; 92. Second control valve. DETAILED DESCRIPTION

[0037] The present invention will be further described below in conjunction with the accompanying drawings.

[0038] An embodiment of the present invention provides a composite heat reduction method for an aircraft, the composite heat reduction method comprising the following steps:

[0039] Gas medium is respectively sprayed at the leading edge 1 of the aircraft and the fuselage 2 located downstream of the leading edge. The gas medium sprayed at the leading edge 1 is configured to push the shock wave away from the outer surface of the leading edge 1 to reduce the heat at the leading edge 1, and to form a mixed gas after mixing with the incoming flow; the mixed gas flows along the incoming flow toward the downstream fuselage 2, and is mixed again with the gas medium sprayed at the fuselage 2, thereby suppressing and delaying the boundary layer transition of the fuselage 2.

[0040] The gas medium injected at the leading edge 1 of the aircraft is injected in the direction opposite to the incoming flow. The leading edge 1 of the aircraft includes the leading edge of the wing, the lip of the airway, the vertical tail and other small radius blunt leading edges.

[0041] The gas medium sprayed on the fuselage 2 of the aircraft has a spraying direction along the normal direction of the wall surface of the fuselage 2 at the spraying location.

[0042] The gas medium is a heavy gas with a density greater than that of air.

[0043] The gas medium is carbon dioxide or nitrogen or a mixture thereof.

[0044] In an embodiment of the present invention, a composite heat reduction method for an aircraft is roughly as follows: a heavy gas, such as carbon dioxide, is sprayed at the leading edge 1 of the aircraft in a direction opposite to the incoming flow. The sprayed carbon dioxide can push the shock wave away from the outer surface of the leading edge 1 to form a more obvious detached shock wave, isolating the shock wave from the leading edge 1 to a certain extent, so as to reduce the aerodynamic heat at the leading edge 1; then, the carbon dioxide sprayed at the leading edge 1 is mixed with the incoming flow and flows downstream with the incoming flow. Since the carbon dioxide concentration in the mixed gas is higher than the carbon dioxide concentration in the incoming flow before mixing, the mixed gas has the ability to suppress transition when flowing to the downstream fuselage 2. Therefore, it can suppress the transition of the boundary layer of the fuselage 2 to a certain extent, and reduce the heat flow in the fuselage 2 area. The transition position is as follows: Figure 3On this basis, carbon dioxide is injected along the normal direction of the wall of the fuselage 2 at the fuselage 2 downstream of the leading edge 1, and the carbon dioxide injected at the fuselage 2 is mixed again with the incoming flow mixed with carbon dioxide from the leading edge 1, so that the carbon dioxide concentration in the mixed gas is increased again, and its ability to suppress the transition of the boundary layer of the fuselage 2 is further improved, and the boundary layer transition is delayed. The delayed transition position is as follows Figure 3 The second turning point in .

[0045] The expression of transition Reynolds number is: Among them, ρ * , μ * are the density and viscosity at Eckert reference temperature, respectively; U e is the air velocity at the outer edge of the boundary layer, x tr The transition position is the transition Reynolds number. The effect of carbon dioxide on the transition delay of the boundary layer can be characterized by the transition Reynolds number.

[0046] When the aircraft head cone angle is 5° and the molar fraction of carbon dioxide after mixing reaches about 14%, the transition Reynolds number increases by more than about 75%, thereby significantly reducing the heat flow in the fuselage 2 area and reducing the heat accumulation in the fuselage 2 of the aircraft during long flight.

[0047] The above-mentioned composite heat reduction method combines reverse jet technology and heavy gas, which can not only reduce the structural heat flux of the leading edge 1, but also reduce the heat flux of a large area of ​​the aircraft fuselage 2 by suppressing the boundary layer transition. Therefore, it can take into account the heat accumulation of the local high heat flux area at the leading edge 1 and the large area of ​​the fuselage 2, thereby reducing the overall aerodynamic heating of the aircraft.

[0048] The composite heat reduction method combining reverse jet technology and heavy gas in this embodiment is applicable to long-flight high-speed aircraft. The high-speed aircraft in the embodiment of the present invention flies at a speed greater than 5 Ma and at an altitude in near space 20 to 100 kilometers from the ground.

[0049] An embodiment of the present invention further provides a heat reduction device for an aircraft, which is used to implement the composite heat reduction method for the aircraft. The heat reduction device includes a first injection hole 3, a second injection hole 4 and an air supply assembly.

[0050] The first injection hole 3 is located at the leading edge 1 of the aircraft. The injection direction of the first injection hole 3 is set to be opposite to the incoming flow direction. The gas medium injected by the first injection hole 3 can push the shock wave away from the outer surface of the leading edge 1 and mix with the incoming flow to form a mixed gas.

[0051] The second injection hole 4 is located on the outer wall of the fuselage 2 of the aircraft. Along the flow direction of the mixed gas, the second injection hole 4 is located downstream of the first injection hole 3. The upstream mixed gas flows through the second injection hole 4 and is mixed again with the gas medium injected by the second injection hole 4.

[0052] Preferably, a plurality of first injection holes 3 are provided, and the plurality of first injection holes 3 are arranged on the leading edge 1 at intervals along the axial direction of the leading edge 1 .

[0053] Preferably, the spray direction of the second spray hole 4 is set to be along the normal direction of the wall surface of the fuselage 2 where the second spray hole 4 is located. Preferably, the second spray hole 4 is set in plurality, and the plurality of second spray holes 4 are arranged in intervals along the airflow direction and the circumferential direction of the fuselage 2, thereby forming a spray hole matrix defined by the plurality of second spray holes 4. Such a setting can more effectively reduce the heat flow in the fuselage 2 area.

[0054] The first injection hole 3 and the second injection hole 4 are both connected to an air supply component, and the air supply component is used to supply air to the first injection hole 3 and the second injection hole 4 .

[0055] The air supply assembly includes an air storage tank 5 and a power pump 6.

[0056] The gas storage tank 5 is used for gaseous medium, such as nitrogen dioxide, nitrogen or a mixture thereof.

[0057] The air inlet of the power pump 6 is communicated with the air storage tank 5, and the air outlet of the power pump 6 is connected to the delivery pipeline and is respectively communicated with the first injection hole 3 and the second injection hole 4 through the delivery pipeline.

[0058] Control valves are respectively arranged on the conveying pipelines.

[0059] The delivery pipeline includes a first delivery branch pipe 71 and a second delivery branch pipe 72 arranged in parallel, the air inlet of the first delivery branch pipe 71 and the air inlet of the first delivery branch pipe 71 are both connected to the air outlet of the power pump 6, and the air outlet of the first delivery branch pipe 71;

[0060] The gas outlet of the first delivery branch pipe 71 is connected to the first drainage pipe 81. The outlet end of the first drainage pipe 81 is located at the leading edge 1. The outlet end of the first drainage pipe 81 is provided with a first nozzle. The first injection hole 3 is defined by the injection hole of the first nozzle. The first nozzle injects the gas medium at the leading edge 1. A first control valve 91 is provided between the first drainage pipe 81 and the first delivery branch pipe 71.

[0061] Similarly, the gas outlet of the second delivery branch pipe 72 is connected to the second drainage pipe 82, the outlet end of the second drainage pipe 82 is located on the fuselage 2, the outlet end of the second drainage pipe 82 is provided with a second nozzle, the second injection hole 4 is defined by the injection hole of the second nozzle, and the gas medium is injected by the second nozzle at the fuselage 2. A second control valve 92 is provided between the second drainage pipe 82 and the second delivery branch pipe 72.

[0062] Although the present invention has been described with reference to preferred embodiments, various modifications may be made thereto and parts thereof may be replaced by equivalents without departing from the scope of the present invention. In particular, the various technical features mentioned in the various embodiments may be combined in any manner as long as there are no structural conflicts. The present invention is not limited to the specific embodiments disclosed herein, but includes all technical solutions falling within the scope of the claims.

Claims

1. A composite heat reduction method for an aircraft, characterized in that: spraying gaseous medium at the leading edge of the aircraft and at the fuselage downstream of the leading edge respectively, wherein the gaseous medium sprayed at the leading edge is configured to push the shock wave away from the outer surface of the leading edge to reduce the heat at the leading edge, and to form a mixed gas after mixing with the incoming flow; The mixed gas flows along with the incoming flow toward the downstream fuselage and is mixed again with the gas medium injected at the fuselage, thereby suppressing and delaying the boundary layer transition of the fuselage.

2. The composite heat reduction method for an aircraft according to claim 1, characterized in that: The gas medium ejected at the leading edge has an ejection direction that is opposite to the incoming flow direction.

3. The composite heat reduction method for an aircraft according to claim 1, characterized in that: The gas medium sprayed on the fuselage of the aircraft has a spraying direction along the normal direction of the fuselage wall at the spraying point.

4. The composite heat reduction method for an aircraft according to any one of claims 1 to 3, characterized in that: The gas medium is a heavy gas with a density greater than that of air.

5. The composite heat reduction method for an aircraft according to claim 4, characterized in that: The gas medium is carbon dioxide or nitrogen or a mixture thereof.

6. A cooling device for implementing the composite cooling method for an aircraft according to any one of claims 1 to 5, characterized in that: include: A first injection hole is arranged at the leading edge of the aircraft, and the gas medium injected by the first injection hole is suitable for pushing the shock wave away from the outer surface of the leading edge and mixing with the incoming flow to form a mixed gas; A second injection hole is arranged on the fuselage of the aircraft, and along the flow direction of the mixed gas, the second injection hole is located downstream of the first injection hole, and the mixed gas upstream flows through the second injection hole and is mixed again with the gas medium injected by the second injection hole; An air supply component is connected to the first injection hole and the second injection hole, and is used to supply air to the first injection hole and the second injection hole.

7. The heat reduction device according to claim 6, characterized in that: The spraying direction of the first spray hole is set to be opposite to the incoming flow direction.

8. The heat reduction device according to claim 6, characterized in that: The spraying direction of the second spray hole is set to be along the normal direction of the fuselage wall where the second spray hole is located.

9. The heat reduction device according to claim 6, characterized in that: The gas supply assembly comprises: A gas storage tank, used for storing the gas medium; A power pump, whose air inlet is connected to the air storage tank, and whose air outlet is connected to the delivery pipeline and respectively connected to the first injection hole and the second injection hole through the delivery pipeline.

10. The heat reduction device according to claim 9, characterized in that: A control valve is arranged on the delivery pipeline.

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