Turboramjet combined power and thermal management multi-modal cooperative system for an aeroengine

Abstract

The present disclosure relates to the technical field of aircraft design, and particularly provides a turbo ram combined power and thermal management multi-modal collaborative system for an aero-engine, comprising a first shell, a second shell, a ducted fan unit and a liquid metal cooling unit; the first shell is arranged at the outer periphery of the aero-engine; a turbine engine is arranged in the first shell; the second shell is arranged below the turbine engine, and an air inlet channel is formed between the upper side of the second shell and the first shell, and a ram channel is formed between the lower side of the second shell and the first shell; the ducted fan unit is arranged in the ram channel and can be unfolded and folded, and when unfolded, the ducted fan unit can increase the bypass ratio of the aero-engine, and when folded, the ducted fan unit is withdrawn into the second shell; the liquid metal cooling unit is connected to the first shell through a first pipeline and can cool the first shell. Through the collaborative design of the ducted fan unit, the liquid metal cooling unit and the liquid metal cooling unit, the present disclosure can effectively solve the problems of thrust gap and thermal protection of the aero-engine.

Description

Technical Field

[0001] This disclosure relates to the field of aircraft design technology, and in particular to a multimodal collaborative system for turbine-ramjet combined power and thermal management of aero engines. Background Technology

[0002] Turbine-based combined cycle (TBCC) systems, through the combination of turbine and ramjet engines, achieve a wide speed range of power coverage from low-speed takeoff to high-speed cruise, making them a core power solution for hypersonic vehicles. However, existing technologies suffer from the following significant drawbacks:

[0003] (1) Turbine-ramjet mode transition gap: Ramjet engines are limited by aerodynamic characteristics. The minimum operating Mach number (Ma) is usually higher than 2.5, and the effective operating Mach number (Ma) is usually higher than 3. This means that the turbine engine needs to bear the thrust alone in the Ma=0-2.5 range. However, the efficiency of the turbine engine drops sharply in the high-speed range. There is a thrust breakpoint in the mode transition zone of Ma=2~3, which leads to the discontinuity of the flight envelope and forms a "thrust gap".

[0004] (2) Challenges of long-term high-temperature thermal protection: Unlike the afterburner of military turbine engines, which operates for only a few minutes, the TBCC's integrated afterburner and ramjet combustor design (referred to as "integrated afterburner / ramjet combustor") needs to operate for extended periods within the Ma range of 1.5-4.0, with wall temperatures exceeding 1200℃. Traditional air-cooling solutions struggle to balance cooling efficiency with thrust loss. Inadequate thermal protection measures can lead to overheating failures, affecting engine reliability and lifespan.

[0005] (3) Low-speed thrust insufficiency: Due to the wide speed range of the engine, it is usually difficult to take into account the thrust requirements of the aircraft at both high and low speeds during the design process. Insufficient thrust during takeoff limits the takeoff performance of the aircraft and makes it difficult to meet the requirements of modern aerospace missions for high performance and high reliability of the engine.

[0006] In the existing technology, the above problems are attempted to be solved by auxiliary power devices with fixed structures or simple heat exchangers, but new problems such as structural redundancy, low thermal management efficiency and surge in cross-modal drag exist. There is an urgent need for an integrated design scheme that takes into account both aerodynamic efficiency and thermal management capabilities. Summary of the Invention

[0007] This disclosure is made in view of the above-mentioned problems. This disclosure provides a multimodal cooperative system for turbo-ramjet combined power and thermal management of aero engines.

[0008] According to one aspect of this disclosure, a multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine is provided, comprising:

[0009] A first housing, the first housing being disposed on the outer periphery of the aircraft engine;

[0010] A turbine engine, wherein the turbine engine is disposed in the first housing;

[0011] The second housing is disposed below the turbine engine. An air intake passage is formed between the upper side of the second housing and the first housing, and a stamping passage is formed between the lower side of the second housing and the first housing.

[0012] A ducted fan unit is disposed in the stamping channel and can be unfolded and folded. When unfolded, it can increase the bypass ratio of the aero-engine, and when folded, it retracts into the second housing.

[0013] A liquid metal cooling unit is connected to the first housing via a first pipe and is capable of cooling the first housing.

[0014] Furthermore, according to one aspect of this disclosure, the multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine also includes:

[0015] An afterburning / ramjet integrated combustion chamber is disposed in a first housing and located behind the turbine engine and the ramjet passage. The first pipeline connects to the first housing at the position corresponding to the afterburning / ramjet integrated combustion chamber.

[0016] Furthermore, according to one aspect of this disclosure, a multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine, wherein the turbojet engine is a turbojet engine or a turbofan engine, includes a power extraction shaft connected to the ducted fan unit via a transmission, capable of driving the ducted fan unit to rotate.

[0017] Furthermore, according to one aspect of this disclosure, in a multimodal collaborative system for turbo-ramjet combined power and thermal management of an aero-engine, the liquid metal cooling unit is also connected to the fuel tank of the aero-engine via heat exchange pipelines, enabling it to exchange heat with the fuel.

[0018] Furthermore, according to one aspect of this disclosure, in a multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine, the ducted fan unit deploys when Ma < 0.9, and the bypass ratio of the aero-engine is 1.8-2.2;

[0019] When 0.9 ≤ Ma < 2, the ducted fan unit folds and retracts into the second housing;

[0020] When 2≤Ma≤3, the ducted fan unit deploys, and the bypass ratio of the aero-engine is 1.2-1.5;

[0021] When Ma > 3, the ducted fan unit folds and retracts into the second housing.

[0022] Furthermore, according to one aspect of this disclosure, the multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine also includes:

[0023] A heat dissipation and airflow guiding unit is disposed in the stamping channel and located downstream of the ducted fan unit. It can be unfolded and folded. When unfolded, it can exchange heat with the airflow flowing through the stamping channel. When folded, it retracts into the second housing. The liquid metal cooling unit is connected to the second housing through a second pipeline and can cool the second housing and the heat dissipation and airflow guiding unit.

[0024] Furthermore, according to one aspect of this disclosure, a multimodal collaborative system for turbine-ramjet combined power and thermal management of an aero-engine is provided with a microchannel cooling structure for the first housing and the second housing, wherein the coolant of the liquid metal cooling unit can circulate in the microchannel cooling structure, and the microchannel cooling structure of the second housing is also connected to the heat dissipation guiding unit through a pipeline.

[0025] Furthermore, according to one aspect of the present disclosure, a multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine, the ducted fan unit includes multi-stage adjustable pitch fan blades, a first retraction actuator, and a nested duct housing, wherein the ducted fan unit is deployed and folded in the ramjet channel by being driven by the first retraction actuator.

[0026] The fan blades are made of titanium alloy, and the angle of the fan blades is adjusted by an electric actuator;

[0027] The nested duct shell is made of carbon-ceramic composite material;

[0028] The heat dissipation and flow guiding unit includes heat dissipation fins, flow guiding blades, and a second retraction actuator. The density of the heat dissipation fins is 20 fins / cm, and the thermal conductivity is ≥400W / (m·K). The angle of the flow guiding blades is adjusted by an electric actuator, and the second retraction actuator drives the heat dissipation and flow guiding unit to unfold and fold in the stamping channel.

[0029] Furthermore, according to one aspect of this disclosure, in a multimodal collaborative system for turbo-ramjet combined power and thermal management of an aero-engine, the first and second pipelines are made of Inconel 718 alloy and the inner walls of the pipelines are gold-plated.

[0030] Furthermore, according to one aspect of this disclosure, in a multimodal collaborative system for turbine-ramjet combined power and thermal management of an aero-engine, the coolant material of the liquid metal cooling unit is a gallium-indium-tin alloy, with 5%-8% volume fraction of graphene nanosheets or silicon carbide nanoparticles added to form a highly thermally conductive nanofluid.

[0031] According to the embodiments of the present disclosure, a multimodal collaborative system for turbine-ramjet combined power and thermal management for aero-engines addresses the technical bottlenecks of existing TBCC systems by providing a combined power scheme of a scalable nested ducted fan unit and a liquid metal thermal management system. This system broadens the lower limit of the operating Mach number of the ramjet engine to Ma=2, eliminating the thrust gap during turbine-ramjet mode transitions; constructs an integrated thermal management network to solve the problems of combustion chamber thermal protection and airborne heat sink optimization under long-term high-temperature environments; and improves the thrust output of the aero-engine during the takeoff phase of the aircraft to meet the power requirements of a high thrust-to-weight ratio.

[0032] It should be understood that both the foregoing general description and the following detailed description are exemplary and intended to provide further illustration of the claimed technology. Attached Figure Description

[0033] The above and other objects, features, and advantages of this disclosure will become more apparent from the more detailed description of the embodiments thereof in conjunction with the accompanying drawings. The drawings are provided to further illustrate the embodiments of this disclosure and form part of the specification. They are used together with the embodiments of this disclosure to explain the disclosure and do not constitute a limitation thereof. In the drawings, the same reference numerals generally represent the same components or steps.

[0034] Figure 1 This is a schematic diagram of a first operating state of a multimodal cooperative system according to an embodiment of the present disclosure, in which the ducted fan unit and the heat dissipation and airflow guiding unit are both in the deployed state;

[0035] Figure 2 This is a schematic diagram of a second working state of a multimodal cooperative system according to an embodiment of the present disclosure. In the diagram, the ducted fan unit is in a folded state, and the heat dissipation and airflow guiding units are all in an unfolded state.

[0036] Figure 3 This is a schematic diagram of a third operating state of a multimodal cooperative system according to an embodiment of the present disclosure, in which the ducted fan unit and the heat dissipation and airflow guiding unit are both in a folded state.

[0037] Explanation of reference numerals in the attached figures:

[0038] 100: Aircraft engine; 1: Air intake; 2: Ram air passage; 3: Second casing; 4: Turbine engine; 5: Ducted fan unit; 6: Heat dissipation and airflow guiding unit; 7: Liquid metal cooling unit; 8: First pipeline; 9: Second pipeline; 10: First casing; 11: Afterburner / ram air combustion chamber; 12: Tail nozzle; 13: Heat exchange pipeline. Detailed Implementation

[0039] To make the objectives, technical solutions, and advantages of this disclosure more apparent, exemplary embodiments according to this disclosure will now be described in detail with reference to the accompanying drawings. Obviously, the described embodiments are merely some embodiments of this disclosure, and not all embodiments of this disclosure. It should be understood that this disclosure is not limited to the exemplary embodiments described herein.

[0040] This disclosure provides a multimodal collaborative system for turbine-ramjet combined power and thermal management of aero engines. Through the collaborative design of the ducted fan unit, the liquid metal cooling unit, and the liquid metal cooling unit, it can effectively solve the thrust gap and thermal protection problems of aero engines.

[0041] The embodiments of this disclosure will now be described in detail with reference to the accompanying drawings.

[0042] like Figure 1 , Figure 2 As shown, this disclosure provides a multimodal collaborative system for turbine-ramjet combined power and thermal management of an aero-engine, including: a first housing 10, a turbine engine 4, a second housing 3, a ducted fan unit 5, and a liquid metal cooling unit 7;

[0043] The first housing 10 is disposed on the outer periphery of the aircraft engine 100. The first housing 10 is usually composed of multiple cylindrical parts, also known as a casing, and its coverage extends from the front air intake 1 to the rear tail nozzle 12.

[0044] The turbine engine 4 is disposed in the first housing 10, and the turbine engine 4 may include main components such as a compressor, a combustion chamber, and a turbine;

[0045] The second housing 3 is located below the turbine engine 4. An air intake duct 1 is formed between the upper side of the second housing 3 and the first housing 10, and a ramming channel 2 is formed between the lower side of the second housing 3 and the first housing 10.

[0046] The upper side of the second housing 3 and the first housing 10 also form a turbine engine installation space and other structures. The turbine engine 4 is located between the upper side of the second housing 3 and the first housing 10. The air intake 1 is located in front of the turbine engine 4. The airflow enters the turbine engine 4 from the air intake 1, and after being pressurized, burned, and powered by the turbine engine 4, it is ejected from the tail nozzle 12.

[0047] The ducted fan unit 5 is located in the stamping channel 2 and can be unfolded and folded. When unfolded, it can increase the bypass ratio of the aero-engine 100. When folded, it retracts into the second housing 3. The second housing 3 is provided with a first receiving cavity. When folded, the ducted fan unit 5 enters the first receiving cavity. The ducted fan unit 5 does not protrude from the outer surface of the second housing 3. Therefore, after folding, the ducted fan unit 5 does not generate resistance in the stamping channel 2.

[0048] The liquid metal cooling unit 7 is connected to the first housing 10 through the first pipe 8 and is capable of cooling the first housing 10. The first pipe 8 includes a coolant delivery pipe and a return pipe, and the coolant can form a circulation path in the housing structure of the first housing 10.

[0049] The multimodal collaborative system for turbine-ramjet combined power and thermal management of aero-engines disclosed in this embodiment broadens the lower limit of the operating Mach number of the ramjet engine to Ma=2, eliminating the thrust gap between turbine and ramjet modes; constructs an integrated thermal management network to solve the problems of combustion chamber thermal protection and airborne heat sink optimization under long-term high-temperature environments; and improves the thrust output of aero-engines during the takeoff phase of the aircraft to meet the power requirements of high thrust-to-weight ratio.

[0050] In some possible implementations, such as Figure 1 , Figure 2 As shown, the multimodal collaborative system for turbine-ramjet combined power and thermal management of aero engines also includes: an afterburner / ramjet integrated combustor 11;

[0051] The afterburning / ramjet integrated combustion chamber 11 is disposed in the first housing 10 and located behind the turbine engine 4 and the ramjet channel 2. The first pipeline 8 connects to the first housing 10 at the position corresponding to the afterburning / ramjet integrated combustion chamber 11. The afterburning / ramjet integrated combustion chamber 11 is an integrated design of the afterburning combustion chamber and the ramjet combustion chamber, which needs to operate for a long time in the range of Ma=1.5-4.0, with a wall temperature exceeding 1200℃.

[0052] In some possible implementations, such as Figure 1 , Figure 2 As shown, the turbine engine 4 is a turbojet engine or a turbofan engine, including a power extraction shaft. The power extraction shaft is connected to the ducted fan unit 5 via a gearbox and can drive the ducted fan unit 5 to rotate. When the ducted fan unit 5 is deployed in the ramjet channel 2, the turbine engine 4 drives the ducted fan unit 5 to rotate via the power extraction shaft, generating additional thrust and increasing the overall equivalent bypass ratio of the aero engine 100. When the ducted fan unit 5 is folded in the ramjet channel 2, the turbine engine 4 no longer drives the ducted fan unit 5 to rotate.

[0053] In some possible implementations, such as Figure 1As shown, the liquid metal cooling unit 7 is also connected to the fuel tank of the aircraft engine 100 via heat exchange pipe 13, enabling heat exchange with the fuel. A fuel-liquid metal heat exchanger is installed in the heat exchange pipe 13, allowing for non-contact heat exchange between the fuel and liquid metal. The fuel tank is equipped with an insulation device, constituting a heat sink that absorbs heat.

[0054] In some possible implementations, when Ma < 0.9, the ducted fan unit 5 is deployed, and the bypass ratio of the aero-engine 100 is 1.8-2.2. At this time, the aircraft is in the takeoff phase or the low-speed cruise phase.

[0055] When 0.9≤Ma<2, the ducted fan unit 5 folds and retracts into the second shell 3. At this time, the aircraft is in the transonic / supersonic phase. After the ducted fan unit 5 folds, it does not generate drag in the ram air passage 2.

[0056] When 2≤Ma≤3, the ducted fan unit 5 is deployed, and the bypass ratio of the aero-engine 100 is 1.2-1.5. At this time, the aircraft is in the mode transition phase.

[0057] When Ma > 3, the ducted fan unit 5 folds and retracts into the second housing 3. At this time, the aircraft is in the high-speed cruise phase. After the ducted fan unit 5 folds, it does not generate resistance in the ram air passage 2.

[0058] In some possible implementations, such as Figure 1 , Figure 2 , Figure 3 As shown, the multimodal collaborative system for turbo-ramjet combined power and thermal management of aero engines also includes: a heat dissipation and airflow guiding unit 6;

[0059] The heat dissipation and airflow guiding unit 6 is located in the stamping channel 2 and downstream of the ducted fan unit 5. It can be unfolded and folded. When unfolded, it can exchange heat with the airflow flowing through the stamping channel 2. When folded, it retracts into the second housing 3. After folding, it does not generate resistance in the stamping channel 2. The liquid metal cooling unit 7 is connected to the second housing 3 through the second pipe 9 and can cool the second housing 3 and the heat dissipation and airflow guiding unit 6.

[0060] like Figure 3 As shown, a second receiving cavity is provided in the second housing 3 corresponding to the heat dissipation guiding unit 6. The second receiving cavity is adjacent to the first receiving cavity and the two can also be connected. When folded, the heat dissipation guiding unit 6 enters the second receiving cavity. The heat dissipation guiding unit 6 does not protrude from the outer surface of the second housing 3, so the heat dissipation guiding unit 6 will not generate resistance in the stamping channel 2 after folding.

[0061] After absorbing heat from the wall of the afterburner / ramjet integrated combustion chamber 11, the liquid metal flows into the heat dissipation and flow guiding unit 6. The heat dissipation and flow guiding unit 6 is equipped with heat exchange fins. The liquid metal exchanges heat with the airflow at the outlet of the ducted fan unit 5 through the heat exchange fins. The airflow is heated and then accelerated through the expansion nozzle to generate additional thrust.

[0062] In some possible implementations, the first housing 10 and the second housing 3 are provided with microchannel cooling structures, in which the coolant of the liquid metal cooling unit 7 can circulate. The microchannel cooling structure of the second housing 3 is also connected to the heat dissipation guiding unit 6 via pipes, and the heat dissipation guiding unit 6 is also provided with a corresponding microchannel cooling structure to allow liquid metal to flow. The microchannel cooling structures in the first housing 10 and the second housing 3 can be multiple cooling channels arranged around the housing, with an inner diameter of 0.5 mm.

[0063] In some possible implementations, such as Figure 1 , Figure 2 As shown, the ducted fan unit 5 includes multi-stage adjustable pitch fan blades, a first retraction actuator, and a nested duct housing. The ducted fan unit 5 is deployed and folded in the stamping channel 2 by being driven by the first retraction actuator. The multi-stage adjustable pitch fan blades can be three or more stages, and the first retraction actuator can be in pneumatic drive mode or electric drive mode.

[0064] In some possible implementations, the fan blades are made of titanium alloy, and the angle of the fan blades is adjusted by an electric actuator.

[0065] In some possible implementations, when the ducted fan unit 5 is folded, the nested duct housing can enter the second housing 3, with a gap of ≤2mm between it and the inner wall of the first receiving cavity, forming a tight nested structure with the first receiving cavity. The nested duct housing is made of carbon-ceramic composite material.

[0066] In some possible implementations, such as Figure 1 , Figure 2 , Figure 3 As shown, the heat dissipation and airflow guiding unit 6 includes heat dissipation fins, airflow guiding blades, and a second retraction actuator. The density of the heat dissipation fins is 20 fins / cm, and the thermal conductivity is ≥400W / (m·K). The angle of the airflow guiding blades is adjusted by an electric actuator, and the expansion and folding of the heat dissipation and airflow guiding unit 6 in the stamping channel 2 is achieved by driving the second retraction actuator. The second retraction actuator can be in pneumatic drive mode or electric drive mode.

[0067] In some possible implementations, the first conduit 8 and the second conduit 9 are made of Inconel 718 alloy, and the inner walls of the conduits are gold-plated. This configuration effectively reduces the oxidation of the liquid metal and extends the service life of the first conduit 8 and the second conduit 9.

[0068] In some possible implementations, the coolant of the liquid metal cooling unit 7 is a gallium-indium-tin alloy, with 5%-8% volume fraction of graphene nanosheets or silicon carbide nanoparticles added to form a highly thermally conductive nanofluid.

[0069] A gallium-indium-tin alloy (mass ratio 68:22:10, melting point -19℃, boiling point 2200℃) is used to add 5%-8% by volume of graphene nanosheets (thickness 5-10nm) or silicon carbide nanoparticles (particle size 50nm) to form a highly thermally conductive nanofluid (increasing thermal conductivity by 25%-30%), which further enhances the cooling efficiency of the wall microchannels of the afterburner / ram-combustion integrated combustion chamber 11, while suppressing the oxidation rate of liquid metal at high temperatures.

[0070] The multimodal cooperative system for turbo-ramjet combined power and thermal management of aero-engines disclosed in this embodiment is divided into four operating stages according to the flight Mach number, and the states and energy flows of each component are as follows:

[0071] a. Takeoff phase (Ma<0.3)

[0072] • Pneumatic component status: The ducted fan unit is fully deployed (duct ratio is 1.8-2.2), the heat dissipation and airflow guiding unit guide vanes are installed at a 45° angle, and the "afterburner / ramjet integrated combustion chamber" is in operation;

[0073] • Energy flow path: The turbine engine power extraction shaft drives the ducted fan to generate additional thrust (accounting for 15%-20% of the total thrust); after absorbing heat from the wall of the "afterburner / ramjet integrated combustion chamber", the liquid metal flows into the heat dissipation and flow guiding unit, and exchanges heat with the airflow at the outlet of the ducted fan unit through the heat dissipation fins. After being heated, the airflow is accelerated through the expansion nozzle to generate additional thrust.

[0074] • Core function: By using ducted fan thrust enhancement and waste heat recovery, the takeoff thrust is increased by 15%-20%, meeting the takeoff requirements of wide-speed-range aircraft.

[0075] b. Low-speed cruising phase (0.3≤Ma<0.9)

[0076] • Pneumatic component status: The ducted fan unit is fully deployed (bypass ratio of 1.8-2.2), the heat dissipation and airflow guiding unit remains deployed, and the "afterburner / ramjet integrated combustion chamber" is not in operation;

[0077] • Energy flow path: The turbine engine power extraction shaft drives the ducted fan unit to generate additional thrust (accounting for 15%-20% of the total thrust), improving the overall equivalent bypass ratio; the "afterburner / ramjet combustor" is not working, and the liquid metal switches to cold source mode, flowing through the combustion chamber wall and heat dissipation guide unit, and the front wall of the ramjet channel, outputting heat to the ambient airflow (approximately -30°C at Ma0.8); the liquid metal will continuously reduce the fuel temperature in the fuel tank, increasing the heat sink;

[0078] • Core function: By driving the ducted fan unit, the equivalent bypass ratio of the whole aircraft is increased, thereby improving the propulsion efficiency for subsonic flight. At the same time, the low-temperature zone of the ramjet channel wall is used to increase the airborne heat sink reserve.

[0079] c. Transonic / Supersonic Phase (0.9 ≤ Ma < 2)

[0080] • Aerodynamic component status: The ducted fan unit is folded into the nested position, the heat dissipation and airflow guiding unit remains deployed, and the "afterburner / ramjet integrated combustion chamber" is in operation;

[0081] • Energy flow path: The ramjet channel is in a natural flow state (no fan boost); the "afterburner / ramjet integrated combustion chamber" is working. After the liquid metal absorbs the heat from the wall of the "afterburner / ramjet integrated combustion chamber", it flows into the heat dissipation and flow guiding unit. It exchanges heat with the ramjet channel through the heat dissipation fins. The heated airflow is then guided by the exhaust airflow of the turbine engine, thereby increasing the thrust.

[0082] • Core function: Cooling the "afterburner / ramjet combustor" by natural flow through the ramjet channel, while increasing thrust by using airflow from the turbine channel to eject air from the ramjet channel.

[0083] d. Mode transition stage (2≤Ma≤3)

[0084] • Pneumatic component status: Duct fan unit is deployed twice (duct ratio is 1.2-1.5), heat dissipation and flow guiding unit guide vanes are adjusted to a 30° installation angle, and "afterburner / ramjet integrated combustion chamber" is in operation;

[0085] • Energy flow path: The turbine engine continuously drives the ducted fan to pressurize the airflow of the ramjet engine (effectively reducing the lower operating limit of the ramjet engine); the liquid metal simultaneously undertakes the cooling and waste heat utilization of the "afterburner / ramjet integrated combustion chamber": the heated airflow is accelerated through the nozzle to supplement the thrust during the mode transition period; as the aerodynamic heat load increases (the temperature of the adiabatic wall rises to 400-600K), the heat is mainly absorbed by the heat sink of the fuel tank.

[0086] • Core function: Through pre-compression and waste heat recovery of the ducted fan unit, a smooth transition between the turbine and ramjet is achieved, eliminating the thrust gap in the Ma2-3 range.

[0087] e. High-speed cruising phase (Ma>3)

[0088] • Aerodynamic component status: The ducted fan unit and heat dissipation guide unit are completely folded (aerodynamic resistance ≤0.5kN), and the stamping channel forms a straight flow channel;

[0089] • Energy flow path: The ramjet engine operates independently, with liquid metal only responsible for cooling the combustion chamber walls (using a closed-loop system); excess heat is gradually dissipated through the fuel heat sink (fuel temperature rise is controlled below 100°C).

[0090] • Core function: Enables efficient operation in pure stamping mode with no resistance components in the flow channel.

[0091] The above description, with reference to the accompanying drawings, describes a multimodal cooperative system for turbo-ramjet combined power and thermal management of an aero-engine according to embodiments of the present disclosure, which has the following advantages:

[0092] (1) Solving the thrust gap problem: By using a scalable ducted fan unit to assist in the compression during the mode transition phase, the lower limit of the working Ma of the ramjet engine can be effectively reduced, so as to achieve a smooth relay between the turbine and the ramjet engine, eliminate the thrust gap, and improve the power continuity and stability of the aircraft when switching between different speed domains.

[0093] (2) Optimize thermal protection performance: With the liquid metal thermal management system as the core, it comprehensively manages the cooling of the wall of the "afterburner / ram-integrated combustion chamber", the heat dissipation of the heat dissipation guide unit and the heat sink of the fuel tank, etc., to ensure the structural integrity and reliability of the combustion chamber under long-term high-temperature operation. At the same time, it makes full use of the thermal characteristics of each component to achieve efficient utilization and balanced distribution of heat, and extend the overall service life of the engine.

[0094] (3) Increase takeoff thrust: During the takeoff phase, the deployment of the retractable ducted fan unit, combined with waste heat recovery, provides additional thrust to the engine, improving the insufficient thrust of the TBCC system during takeoff, enabling the aircraft to complete the takeoff process more smoothly, and expanding the aircraft's mission adaptability and takeoff performance.

[0095] (4) Compact structure: Utilizing the space of the stamping channel in the low-speed stage to achieve comprehensive performance optimization - the scalable nested design improves the utilization rate of the stamping channel in the low-speed stage.

[0096] The basic principles of this disclosure have been described above with reference to specific embodiments. However, it should be noted that the advantages, benefits, and effects mentioned in this disclosure are merely examples and not limitations, and should not be considered as essential features of each embodiment of this disclosure. Furthermore, the specific details disclosed above are for illustrative and facilitative purposes only, and are not limitations. These details do not limit the scope of this disclosure to the necessity of employing the aforementioned specific details for implementation.

[0097] The block diagrams of devices, apparatuses, devices, and systems disclosed herein are merely illustrative examples and are not intended to require or imply that they must be connected, arranged, or configured in the manner shown in the block diagrams. As those skilled in the art will recognize, these devices, apparatuses, devices, and systems can be connected, arranged, and configured in any manner. Words such as “comprising,” “including,” “having,” etc., are open-ended terms meaning “including but not limited to,” and are used interchangeably with them. The terms “or” and “and” as used herein refer to the terms “and / or,” and are used interchangeably with them unless the context clearly indicates otherwise. The term “such as” as used herein refers to the phrase “such as but not limited to,” and is used interchangeably with it.

[0098] Additionally, as used herein, the "or" used in a list of items beginning with "at least one" indicates a separate list, such that a list of, for example, "at least one of A, B, or C" means A or B or C, or AB or AC or BC, or ABC (i.e., A and B and C). Furthermore, the word "exemplary" does not imply that the described example is preferred or better than other examples.

[0099] It should also be noted that in the systems and methods of this disclosure, the components or steps can be decomposed and / or recombined. These decompositions and / or recombinations should be considered as equivalent solutions to this disclosure.

[0100] Various changes, substitutions, and modifications can be made to the technology described herein without departing from the teachings defined by the appended claims. Furthermore, the scope of the claims of this disclosure is not limited to the specific aspects of the processes, machines, manufactures, events, means, methods, and actions described above. Currently existing or later-developed processes, machines, manufactures, events, means, methods, or actions that perform substantially the same function or achieve substantially the same result as the corresponding aspects described herein can be utilized. Therefore, the appended claims include such processes, machines, manufactures, events, means, methods, or actions within their scope.

[0101] The above description of the disclosed aspects is provided to enable any person skilled in the art to make or use this disclosure. Various modifications to these aspects will be readily apparent to those skilled in the art, and the general principles defined herein may be applied to other aspects without departing from the scope of this disclosure. Therefore, this disclosure is not intended to be limited to the aspects shown herein, but rather to be carried out within the widest scope consistent with the principles and novel features disclosed herein.

[0102] The above description has been given for purposes of illustration and description. Furthermore, this description is not intended to limit the embodiments of this disclosure to the forms disclosed herein. Although numerous exemplary aspects and embodiments have been discussed above, those skilled in the art will recognize certain variations, modifications, alterations, additions, and sub-combinations thereof.

Claims

1. A turbo-ram combined power and thermal management multi-modal synergic system for an aero-engine, characterized in that, include: A first housing (10) is disposed on the outer periphery of the aircraft engine (100); A turbine engine (4) is disposed in the first housing (10); The second housing (3) is disposed below the turbine engine (4). An air intake (1) is formed between the upper side of the second housing (3) and the first housing (10), and a stamping channel (2) is formed between the lower side of the second housing (3) and the first housing (10). Ducted fan unit (5), the ducted fan unit (5) is disposed in the stamping channel (2), and can be unfolded and folded. When unfolded, it can increase the bypass ratio of the aero-engine (100), and when folded, it is retracted into the second housing (3). A liquid metal cooling unit (7) is connected to the first housing (10) through a first pipe (8) and is capable of cooling the first housing (10). The heat dissipation and airflow guiding unit (6) is disposed in the stamping channel (2) and located downstream of the duct fan unit (5). It can be unfolded and folded. When unfolded, it can exchange heat with the airflow flowing through the stamping channel (2). When folded, it is retracted into the second housing (3). The liquid metal cooling unit (7) is connected to the second housing (3) through the second pipe (9) and can cool the second housing (3) and the heat dissipation and airflow guiding unit (6). The ducted fan unit (5) includes multi-stage adjustable pitch fan blades, a first retraction actuator, and a nested duct housing. The ducted fan unit (5) is deployed and folded in the stamping channel (2) by being driven by the first retraction actuator. The fan blades are made of titanium alloy, and the angle of the fan blades is adjusted by an electric actuator; The nested duct shell is made of carbon-ceramic composite material; The heat dissipation and flow guiding unit (6) includes heat dissipation fins, flow guiding blades, and a second retraction actuator. The density of the heat dissipation fins is 20 fins / cm, and the thermal conductivity is ≥400W / (m·K). The angle of the flow guiding blades is adjusted by an electric actuator. The second retraction actuator drives the heat dissipation and flow guiding unit (6) to unfold and fold in the stamping channel (2).

2. The multimodal collaborative system for turbine-ramjet combined power and thermal management of aero engines according to claim 1, characterized in that, Also includes: An afterburning / ramjet integrated combustion chamber (11) is disposed in the first housing (10) and located behind the turbine engine (4) and the ramjet channel (2). The first pipeline (8) is connected to the first housing (10) at the position corresponding to the afterburning / ramjet integrated combustion chamber (11).

3. The multimodal collaborative system for turbine-ramjet combined power and thermal management of aero engines according to claim 1, characterized in that, The turbine engine (4) is a turbojet engine or a turbofan engine, including a power extraction shaft, which is connected to the ducted fan unit (5) via a gearbox and can drive the ducted fan unit (5) to rotate.

4. The multimodal collaborative system for turbine-ramjet combined power and thermal management of aero engines according to claim 1, characterized in that, The liquid metal cooling unit (7) is also connected to the fuel tank of the aircraft engine (100) through a heat exchange pipeline (13), and can exchange heat with the fuel.

5. The multimodal collaborative system for turbine-ramjet combined power and thermal management of aero-engines according to claim 1, characterized in that, When Ma < 0.9, the ducted fan unit (5) is deployed, and the bypass ratio of the aero-engine (100) is 1.8-2.2; When 0.9≤Ma<2, the ducted fan unit (5) folds and retracts into the second housing (3); When 2≤Ma≤3, the ducted fan unit (5) is deployed, and the bypass ratio of the aero-engine (100) is 1.2-1.5; When Ma > 3, the ducted fan unit (5) folds and retracts into the second housing (3).

6. The multimodal cooperative system for turbine-ramjet combined power and thermal management of aero-engines according to claim 1, characterized in that, The first housing (10) and the second housing (3) are provided with microchannel cooling structures, and the coolant of the liquid metal cooling unit (7) can circulate in the microchannel cooling structure. The microchannel cooling structure of the second housing (3) is also connected to the heat dissipation and flow guiding unit (6) through a pipeline.

7. The multimodal cooperative system for turbine-ramjet combined power and thermal management of aero engines according to claim 1, characterized in that, The first pipe (8) and the second pipe (9) are made of Inconel 718 alloy and the inner wall of the pipe is gold plated.

8. The multimodal cooperative system for turbine-ramjet combined power and thermal management of aero engines according to claim 1, characterized in that, The coolant material of the liquid metal cooling unit (7) is a gallium-indium-tin alloy, with 5%-8% volume fraction of graphene nanosheets or silicon carbide nanoparticles added to form a highly thermally conductive nanofluid.

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