Jet flow impact type active and passive composite cooling structure of scramjet engine

Abstract

The invention relates to the technical field of high-temperature thermal protection of scramjet engines, in particular to a jet flow impact type active and passive composite cooling structure of a scramjet engine. Wherein a jet flow and regeneration channel combined structure is adopted for active cooling, supercritical hydrocarbon fuel serves as a medium for active cooling, and a closed-loop design of flow dividing, flow converging, jet flow and regeneration is adopted; passive cooling is achieved by laying an ultra-high-temperature ceramic matrix composite coating on the wall face of the combustion chamber, heat in the combustion chamber can be effectively prevented from being conducted to a wall face matrix, the combination performance with a metal or composite matrix is good, and the passive cooling device is suitable for the thermal environment under the high Mach number working condition. And through the synergistic effect of active and passive cooling, safe and stable operation of the scramjet engine under the high Mach number working condition is guaranteed.

Description

Technical Field

[0001] This invention relates to the field of high-temperature thermal protection technology for scramjet engines, specifically to a jet-impact type active-passive composite cooling structure for scramjet engines. Background Technology

[0002] Hypersonic vehicles represent a core development direction for future long-range rapid transportation, deep space exploration, and national defense. Their core propulsion system, the scramjet engine (using hydrocarbon fuels as propellants), has become the mainstream choice for hypersonic propulsion systems due to its wide availability of fuel, high energy density, and convenient storage and transportation. However, when the vehicle's flight speed reaches Mach 5 or higher, the combustion chamber faces the dual challenges of combustion heat release and aerodynamic heating, resulting in a dramatic increase in heat load and extremely high wall temperatures and heat flux densities. Traditional passive heat insulation technologies (such as heat-insulating coatings and high-temperature alloys) can only block heat but cannot actively remove heat from the walls. Once the heat flux density exceeds the load-bearing limit, it can easily lead to localized overheating of the walls, decreased structural strength, or even ablation failure, seriously threatening the stable operation of the engine and flight safety. To solve this problem, "active regenerative cooling technology using hydrocarbon fuels as coolants" has become the mainstream solution. Fuel flows through the cooling channels on the combustion chamber walls, absorbing heat from the walls through convection heat transfer before entering the combustion chamber to participate in combustion, achieving energy recycling. However, the operating pressure inside the combustion chamber (3.5-7 MPa) is higher than the critical pressure of hydrocarbon fuels (2-3 MPa), meaning the fuel remains in a supercritical state. Its thermal properties undergo drastic changes within the pseudo-critical temperature range. Furthermore, as heat absorption increases, thermal decomposition occurs when the fuel temperature rises to 600-800 K (depending on fuel type and pressure), generating small-molecule products and altering thermal properties. Carbon deposits may also accumulate on the channel walls. The combination of these two factors leads to frequent heat transfer deterioration in supercritical hydrocarbon fuels under both non-cracking and cracking conditions. Existing combustion chambers generally employ single-layer cooling channels. While simple in design and easy to manufacture, their single flow path and heat exchange method cannot cope with the heat transfer fluctuations of supercritical fuels. Uneven flow and low heat exchange efficiency occur under non-cracking conditions, while exacerbated heat transfer fluctuations under cracking conditions make it difficult to achieve uniform heat exchange across the entire combustion chamber, failing to effectively curb heat transfer deterioration and thus failing to meet the thermal control requirements for conditions above Mach 5. Summary of the Invention

[0003] This invention aims to effectively reduce the temperature of the combustion chamber of high-temperature components in hypersonic vehicles, solving the problems of insufficient heat exchange efficiency and difficulty in controlling heat transfer deterioration in existing technologies. It provides a jet-impact type active-passive composite cooling structure for scramjet engines. The active cooling adopts a structure combining jet and regeneration channels, while the passive cooling is achieved by applying an ultra-high temperature ceramic matrix composite material coating to the combustion chamber wall. This effectively blocks the conduction of heat from inside the combustion chamber to the wall substrate and has good adhesion to the metal or composite material matrix, making it suitable for the thermal environment under high Mach number conditions. Through the synergistic effect of active and passive cooling, the safe and stable operation of the scramjet engine under high Mach number conditions is ensured.

[0004] To solve the above-mentioned technical problems, the technical solution adopted by the present invention is as follows: a jet impact type active-passive composite cooling structure for a scramjet engine, comprising an active cooling structure and a passive cooling structure; the active cooling structure includes multiple cooling channels located on the outer wall of the combustion chamber of the scramjet engine, a manifold is provided above the cooling channels, multiple longitudinal baffles are provided inside the manifold, the longitudinal baffles divide the manifold into multiple manifold sub-cavities, each manifold sub-cavity has a manifold inlet at its top, and a jet hole is provided at the bottom of the manifold corresponding to the position of the cooling channel; The cooling channel is provided with a fuel inlet and a fuel outlet. The fuel inlet and the manifold inlet are both connected to a fuel pump and a fuel tank through fuel pipelines. An adjustment device is installed on the fuel pipelines. The fuel outlet is connected to the interior of the combustion chamber through a pipeline. The passive cooling structure is that the outer wall of the combustion chamber is coated with a wall coating, which is an ultra-high temperature ceramic matrix composite material.

[0005] As a further limitation of the technical solution of the present invention, the cooling channel is a single-layer rectangular flow channel, and multiple cooling channels are evenly distributed along the combustion chamber axis.

[0006] As a further limitation of the technical solution of the present invention, the jet hole points vertically downward toward the inner wall of the cooling channel.

[0007] As a further limitation of the technical solution of the present invention, the wall coating is applied to the surface of the combustion chamber wall substrate using a plasma spraying process, fully covering the combustion chamber wall, from the intake outlet to the tail nozzle inlet, and the wall coating extends to the edge of the cooling channel inlet.

[0008] As a further limitation of the technical solution of the present invention, the ultra-high temperature ceramic matrix composite material is a silicon carbide (SiC), hafnium carbide (HfC), zirconium carbide (ZrC) ceramic fiber reinforced ceramic matrix material.

[0009] Compared with the prior art, the present invention has the following beneficial effects: This invention addresses the problem that traditional passive insulation technology cannot actively remove heat from the combustion chamber walls: Under Mach 5 and above operating conditions, the combustion chamber is subjected to both combustion heat release and aerodynamic heating, resulting in a dramatic increase in heat load. Traditional passive insulation technology can only block heat and cannot actively remove the heat accumulated on the walls, which can easily lead to local overheating of the walls, reduced structural strength, or even ablation failure, affecting engine stability and flight safety. This invention adopts an active cooling subsystem: an integrated jet-regeneration channel structure. Active cooling uses supercritical hydrocarbon fuel as the medium and adopts a closed-loop design of "diversion-convergence-jet-regeneration". The fuel pressure in the convergence chamber is higher than the mainstream pressure in the cooling channel, forming a pressure difference to drive the jet. The high-speed jet can impact the mainstream fuel boundary layer in the cooling channel and destroy the heat transfer barrier layer formed by the abrupt change in the thermophysical properties of the supercritical fuel.

[0010] This invention addresses the problem that existing single-layer cooling channels cannot meet the complex heat exchange requirements of supercritical fuels. Existing combustion chambers using single-layer cooling channels are limited by a single flow path and heat exchange method, making them unable to cope with heat exchange fluctuations in supercritical fuels. Under non-pyrolysis conditions, uneven flow and low heat exchange efficiency are common. Under pyrolysis conditions, heat exchange fluctuations are further aggravated, making it difficult to achieve uniform heat exchange across the entire combustion chamber. The active cooling subsystem designed in this invention, after being pressurized by a fuel pump, is divided into two paths (mainstream path - cooling channel, jet path - manifold) by a regulating device. The flow ratio of the two paths can be dynamically adjusted according to the combustion chamber heat load to achieve optimal heat exchange under different operating conditions.

[0011] This invention addresses the problems of insufficient active-passive synergistic cooling mechanism and inadequate thermal control in existing technologies: In existing technologies, passive insulation and active regenerative cooling are mostly used separately without synergistic effects, failing to fully leverage the advantages of both cooling methods and making it difficult to meet the high-efficiency thermal control requirements of the combustion chamber under Mach number 5 and above. This invention adopts an active-passive synergistic mechanism, with heat blocking and removal working in tandem: The passive coating first reduces the heat input from the high-temperature combustion gas to the wall surface, reducing the total heat that active cooling needs to remove; Active cooling enhances convective heat transfer through jets, quickly removing the heat absorbed by the wall surface, forming a highly efficient combination of "less heat absorption + faster heat dissipation". Attached Figure Description

[0012] Figure 1 This is a schematic diagram of the circulating principle of the jet impact type active-passive composite cooling structure for scramjet engines of the present invention.

[0013] Figure 2 A cross-sectional view of the side view of the active and passive cooling regeneration channel.

[0014] Figure 3 This is a cross-sectional view of the main and passive cooling regeneration channel.

[0015] The markings in the image are as follows: 1-Aircraft, 2-Fuel tank, 3-Fuel pump, 4-Regulating device, 5-Air intake, 6-Combustion chamber, 7-Tail nozzle, 8-Cooling passage, 9-Jet hole, 10-Manifold, 11-Manifold inlet, 12-Wall coating. Detailed Implementation

[0016] The present invention will be further described below with reference to specific embodiments. Example 1

[0017] like Figure 1 As shown, a jet-impact type active-passive composite cooling structure for a scramjet engine includes an active cooling structure and a passive cooling structure. The active cooling structure includes multiple cooling channels 8 located on the outer wall of the combustion chamber 6 of the scramjet engine. A manifold 10 is provided above the cooling channels 8. Multiple longitudinal baffles are provided inside the manifold 10, which divides the manifold 10 into multiple manifold sub-cavities. Each manifold sub-cavity has a manifold inlet 11 at its top and a jet hole 9 at the bottom of the manifold 10 corresponding to the position of the cooling channel 8. The cooling channel 8 is provided with a fuel inlet and a fuel outlet. The fuel inlet and the manifold inlet 11 are both connected to the fuel pump 3 and the fuel tank 2 through fuel pipelines. An adjustment device 4 is installed on the fuel pipelines. The fuel outlet is connected to the interior of the combustion chamber 6 through a pipeline. The passive cooling structure is that the outer wall of the combustion chamber 6 is sprayed with a wall coating 12, which is an ultra-high temperature ceramic matrix composite material.

[0018] The core of this invention is a combined active and passive combustion chamber wall cooling system, integrating "jet-regeneration channel composite active cooling" with "ultra-high temperature ceramic matrix composite coating passive cooling," and combining... Figure 1 (Circuit principle diagram) Figure 2 (Side view sectional view) Figure 3 (Main view sectional view), achieving a precise match between structure and function, the specific principle is as follows: (1) Active cooling subsystem: integrated structure of jet-regeneration channel Active cooling uses supercritical hydrocarbon fuel as the medium and adopts a closed-loop design of "splitting-combining-jetting-regeneration".

[0019] Channel and manifold layout: Cooling channel 8 is a single-layer rectangular flow channel located on the wall of combustion chamber 6. Figure 2 The cross-section shows the channel section), and multiple cooling channels 8 are evenly distributed along the combustion chamber axis. Figure 3 Multiple cooling channels 8 are visible in the cross-section; the manifold 10 is integrated above the cooling channels 8. Figure 2 Located at the top of the channel), it is connected to the fuel pipeline through the manifold inlet 11, and the manifold 10 is equipped with a longitudinal baffle ( Figure 3 As can be seen, the baffle divides the manifold into multiple independent sub-cavities, ensuring a uniform distribution of fuel pressure.

[0020] Jet enhancement design: An array of jet holes 9 are opened at the bottom of each manifold cavity corresponding to the cooling channel 8. Figure 2 The cross-section shows that the jet hole 9 points vertically downwards to the inner wall of the channel, and the fuel pressure in the manifold 10 is higher than the mainstream pressure in the cooling channel 8, forming a pressure difference to drive the jet. The high-speed jet can impact the mainstream fuel boundary layer in the cooling channel and destroy the heat transfer barrier layer formed by the abrupt change in the thermal properties of the supercritical fuel.

[0021] Dynamic flow control: After being pressurized by fuel pump 3, the fuel is divided into two paths by regulating device 4 (main stream path - cooling channel 8, jet path - junction cavity 10). The flow ratio of the two paths can be dynamically adjusted according to the heat load of the combustion chamber to achieve optimal heat exchange under different operating conditions.

[0022] (2) Passive cooling subsystem: Ultra-high temperature ceramic matrix composite coating Passive cooling serves as a "heat insulation layer" on the combustion chamber walls, corresponding to the attached... Figure 1 The combustion chamber wall coating 12, and Figure 2 , Figure 3 The coating structure on the outer side of the substrate in the middle wall.

[0023] Coating Structure and Performance: The coating is applied to the substrate surface of the combustion chamber wall using plasma spraying technology, fully covering the combustion chamber wall 6 (from the intake duct 5 outlet to the tailpipe 7 inlet), and... Figure 2 , Figure 3 As can be seen, the coating extends to the edge of the inlet of cooling channel 8 to prevent localized overheating in the channel inlet area due to heat concentration.

[0024] (3) Active and passive coordination mechanism Synergistic heat barrier and removal: The passive coating 12 first reduces the heat input from the high-temperature combustion gas in the combustion chamber to the wall, reducing the total heat that active cooling needs to remove; active cooling enhances convective heat transfer through jets, quickly removing the heat absorbed by the wall, forming a highly efficient combination of "less heat absorption + fast heat dissipation".

[0025] The working process of the jet-impact active-passive composite cooling structure for scramjet engines of this invention is as follows: The active cooling process involves fuel from the fuel tank 2 entering the manifold 10 and cooling channel 8 via the fuel pump 3. The flow rates of the manifold 10 and cooling channel 8 are controlled by the regulating device 4, ensuring that the pressure entering the manifold 10 is greater than the pressure in the cooling channel 8. Part of the fuel enters the manifold through the inlet 11, which is divided into multiple sub-manifolds by a partition, ensuring that the fuel entering the sub-manifolds is uniformly fed into the cooling channel through the jet orifice (i.e., the jet). Part of the fuel enters through the cooling channel 8 (the main flow), working in conjunction with the jet fluid to cool the combustion chamber wall. After absorbing heat from the combustion chamber 6, the high-temperature fuel is pumped into the combustion chamber for combustion. The passive cooling process involves a wall coating 12 applied to the combustion chamber wall, effectively preventing heat transfer from inside the combustion chamber to the wall substrate. Based on the synergistic effect of active and passive cooling, the temperature of the combustion chamber wall is effectively reduced.

Claims

1. A jet-impact type active-passive composite cooling structure for a scramjet engine, characterized in that, It includes an active cooling structure and a passive cooling structure; the active cooling structure includes multiple cooling channels (8) located on the outer wall of the combustion chamber (6) of the scramjet engine, a manifold (10) is provided above the cooling channels (8), multiple longitudinal baffles are provided in the manifold (10), the longitudinal baffles divide the manifold (10) into multiple manifold sub-cavities, each manifold sub-cavity has a manifold inlet (11) at the top, and a jet hole (9) is provided at the bottom of the manifold (10) at the position corresponding to the cooling channel (8); The cooling channel (8) is provided with a fuel inlet and a fuel outlet. The fuel inlet and the manifold inlet (11) are both connected to the fuel pump (3) and the fuel tank (2) through fuel pipelines. An adjustment device (4) is installed on the fuel pipeline. The fuel outlet is connected to the interior of the combustion chamber (6) through a pipeline. The passive cooling structure is a combustion chamber (6) with a wall coating (12) sprayed on the outer wall surface, and the wall coating (12) is an ultra-high temperature ceramic matrix composite material.

2. The jet-impact type active-passive composite cooling structure for scramjet engines according to claim 1, characterized in that, The cooling channel (8) is a single-layer rectangular flow channel, and multiple cooling channels (8) are evenly distributed along the axial direction of the combustion chamber (6).

3. The jet-impact type active-passive composite cooling structure for scramjet engines according to claim 1, characterized in that, The jet hole (9) points vertically downward toward the inner wall of the cooling channel (8).

4. The jet-impact type active-passive composite cooling structure for scramjet engines according to claim 1, characterized in that, The wall coating (12) is applied to the surface of the combustion chamber (6) wall substrate using a plasma spraying process, fully covering the combustion chamber (6) wall. The wall coating (12) extends from the outlet of the intake duct (5) to the inlet of the tail nozzle (7), and extends to the edge of the inlet of the cooling channel (8).

5. The jet-impact type active-passive composite cooling structure for scramjet engines according to claim 1, characterized in that, The ultra-high temperature ceramic matrix composite material is a silicon carbide, hafnium carbide, or zirconium carbide ceramic fiber reinforced ceramic matrix material.

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