Integrated active cooling and propelling system of hypersonic gliding aircraft, nose cone assembly and aircraft

Abstract

The invention provides an integrated active cooling and propelling system of a hypersonic gliding aircraft, a nose cone assembly and an aircraft, and belongs to the technical field of thermal protection and propelling of hypersonic aircrafts. The problems of insufficient working medium heat sink, single energy utilization and incompatibility with propulsion in traditional active cooling are solved. The passive thermal protection layer is attached to the outer side surface of the aircraft nose cone and used for blocking external extreme aerodynamic heat flow. The active cooling flow channel structure is arranged between the passive thermal protection layer and the nose cone; the ammonia conveying module comprises a first ammonia storage tank, a throttling valve and a second check valve which are sequentially connected through a pipeline, and an outlet of the ammonia conveying module communicates with an inlet of the active cooling flow channel structure; an outlet of the active cooling flow channel structure is communicated with an inlet of the auxiliary propulsion spray pipe module; and the auxiliary propulsion spray pipe module is used for spraying out the working medium from the active cooling flow channel structure in an accelerated manner. The method is mainly used in the field of aerospace equipment.

Description

Technical Field

[0001] This invention belongs to the field of thermal protection and propulsion technology for hypersonic vehicles, and in particular relates to an integrated active cooling and propulsion system for hypersonic gliders. Background Technology

[0002] Hypersonic glide vehicles, as a new generation of aerospace equipment, fly at Mach numbers exceeding 5, enabling long-distance, highly maneuverable gliding in near space. They demonstrate immense strategic and application value in areas such as rapid long-range delivery and space access. However, behind their outstanding performance lie two intertwined core technological challenges: reliable thermal protection under extreme aerodynamic and thermal environments, and effective attitude and trajectory control during the unpowered gliding phase.

[0003] Firstly, regarding thermal protection, when an aircraft traverses the dense atmosphere at high Mach numbers, its nose cone, particularly the stagnation point, will experience extremely severe aerodynamic heating. The heat flux density is extremely high, with temperatures reaching 2000-5000K. At Mach 10, the temperature of the airflow after the shock wave exceeds 10000K, far exceeding the melting point and long-term operating limits of most traditional engineering materials. To address this challenge, current technologies primarily rely on passive thermal protection systems. These systems typically employ high-performance ablation-resistant or heat-insulating materials such as ceramic matrix composites and carbon-carbon composites, increasing material thickness to block heat transfer to the internal structure. However, this purely passive protection method has inherent drawbacks: firstly, to withstand prolonged, high-heat-flux heating, the required material thickness and weight increase significantly, severely limiting the aircraft's payload and maneuverability; secondly, its protective capability has an upper limit. When the heat flux density exceeds the material's tolerance limit or the heating time is too long, it can easily lead to structural ablation, thermal deformation, or even failure, resulting in insufficient reliability.

[0004] Secondly, in terms of propulsion and control, hypersonic gliders are primarily in a powerless gliding state after separation, lacking a continuous and controllable thrust source. This makes it difficult to make precise attitude adjustments and trajectory corrections when performing complex maneuvers or responding to disturbances, limiting their mission adaptability and flight stability. Although there are means of control using aerodynamic control surfaces, the efficiency and reliability of these surfaces themselves face severe challenges in extreme thermal environments.

[0005] In the search for more efficient thermal management solutions, active cooling technology is considered a potential supplementary or alternative approach, such as convection cooling and perspiration cooling. Convection cooling, in particular, involves introducing a cooling medium into channels beneath the aircraft skin, achieving precise temperature control of specific areas through the medium's physical heat absorption. However, existing active cooling technologies still have significant shortcomings. For example, commonly used cooling media, such as aviation kerosene, have limited specific heat capacity and low heat sink density, resulting in insufficient cooling efficiency under extreme heat fluxes. While water absorbs a large amount of latent heat during vaporization, the specific heat of its vaporization products is relatively low, limiting subsequent physical heat sinking and potentially causing instability in two-phase flow. Liquid hydrogen, despite possessing extremely high specific heat and latent heat of vaporization, requires storage at extremely low temperatures (around 20K), placing extremely high demands on tank insulation, making the system complex, heavy, and significantly increasing engineering difficulty and cost. More critically, the vast majority of existing active cooling systems are designed with only a single thermal protection function in mind. After completing their heat absorption task, the cooling medium is usually discharged or collected as a "waste heat" carrier, and the thermodynamic energy it carries is not further utilized. This "disposable" usage mode results in huge energy waste. Furthermore, the cooling system is completely independent of the propulsion system, which adds extra piping, tanks, and control units, leading to a complex overall system structure, increased weight, and low overall energy utilization efficiency.

[0006] In recent years, researchers have explored using media with high specific heat and endothermic decomposition properties as cooling media. Liquid ammonia, when heated, not only absorbs sensible heat through physical heating but also undergoes endothermic decomposition at high temperatures, producing a mixture of hydrogen and nitrogen. This chemical endothermic process provides an additional, substantial heat sink. Although this property has attracted considerable theoretical attention, its efficient and reliable application in hypersonic vehicles still faces many unresolved technical bottlenecks: for example, how to design cooling channel structures to achieve precise matching and efficient coupling between the physical and chemical endothermic processes of liquid ammonia and the non-uniform heat flow in the nose cone? How to smoothly introduce the cooled liquid ammonia and its high-temperature decomposition products into the propulsion system and design matching nozzles to efficiently generate thrust? In other words, how to achieve seamless integration and synergistic optimization from thermal protection to propulsion energy utilization—current technologies have not yet provided mature and complete systematic solutions.

[0007] In summary, in the current field of hypersonic glide vehicles, thermal protection systems and propulsion systems are typically designed separately or in a simple parallel configuration. Passive thermal protection faces the contradiction between temperature limits and weight, while traditional active cooling suffers from insufficient heat sink capacity of the working fluid, limited energy utilization, and incompatibility with propulsion. Applying promising working fluids to integrated systems still lacks effective structural design methods and pathways for cascaded energy utilization. Summary of the Invention

[0008] In view of this, the present invention aims to propose an integrated active cooling and propulsion system, nose cone assembly and aircraft for hypersonic gliders, in order to solve the problems of insufficient working fluid heat sink, single energy utilization and incompatibility with propulsion in traditional active cooling.

[0009] To achieve the above objectives, the present invention adopts the following technical solution: an integrated active cooling and propulsion system for a hypersonic glider, the system comprising a passive thermal protection layer, an active cooling channel structure, an ammonia delivery module, an auxiliary propulsion nozzle module, and a nose cone; The passive thermal protection layer is attached to the outer surface of the aircraft's nose cone to block external extreme aerodynamic heat flow. The active cooling channel structure is arranged between the passive thermal protection layer and the nose cone; The ammonia delivery module includes a first ammonia storage tank, a throttle valve, and a second check valve connected in sequence by pipelines. The outlet of the ammonia delivery module is connected to the inlet of the active cooling channel structure for delivering liquid ammonia to the active cooling channel structure. The outlet of the active cooling channel structure is connected to the inlet of the auxiliary propulsion nozzle module; The auxiliary propulsion nozzle module is used to accelerate the ejection of the working propellant from the active cooling channel structure.

[0010] Furthermore, a preferred embodiment is proposed, wherein the ammonia delivery module further includes a spring accumulator, the spring accumulator comprising a spring, a piston, and an ammonia preheating decomposition cylinder; the spring acts on the piston, the piston dividing the ammonia preheating decomposition cylinder into two non-communicating parts; the outlet of the first ammonia storage tank is connected to the ammonia preheating decomposition cylinder through a first check valve.

[0011] Furthermore, a preferred embodiment is proposed, wherein the ammonia delivery module further includes a second ammonia storage tank and a pressure limiting valve; the inlet of the second ammonia storage tank is connected to the ammonia preheating and decomposition cylinder, and its outlet is connected to the inlet of the throttle valve; the pressure limiting valve is disposed on the second ammonia storage tank.

[0012] Furthermore, a preferred embodiment is proposed, wherein the ammonia delivery module further includes: an active cooling channel structure comprising a spiral channel arranged on the surface of the nose cone, and the distribution density of the spiral channel on the surface of the nose cone is non-uniform.

[0013] Furthermore, a preferred embodiment is proposed, wherein the cross-section of the spiral flow channel is circular, and its inner wall is provided with a turbulence enhancement structure.

[0014] Furthermore, a preferred embodiment is proposed, wherein the working fluid transported by the ammonia delivery module is liquid ammonia; after the liquid ammonia flows through the active cooling channel structure, it enters the auxiliary propulsion nozzle module in the form of a mixed working fluid containing liquid ammonia and its decomposition products.

[0015] Furthermore, a preferred embodiment is proposed, wherein the auxiliary propulsion nozzle module is a Laval nozzle structure.

[0016] Furthermore, a preferred method is proposed, wherein the system's workflow is as follows: liquid ammonia is output from the ammonia delivery module, flows through the active cooling channel structure, and then enters the auxiliary propulsion nozzle module to be accelerated and ejected.

[0017] Based on the same inventive concept, the present invention also proposes a nose cone component for an aircraft, including an integrated active cooling and propulsion system for hypersonic gliders as described in any one of the above-mentioned claims.

[0018] Based on the same inventive concept, the present invention also proposes a hypersonic gliding vehicle, including the nose cone component of the vehicle as described above.

[0019] Compared with the prior art, the beneficial effects of the present invention are: This invention overcomes the limitations of traditional hypersonic vehicles' separate thermal management and propulsion systems, which are often functionally isolated. By using liquid ammonia simultaneously as both the cooling medium and propulsion fuel, and designing a continuous flow path between the cooling channel and the nozzle, a single working medium and a single system process can sequentially complete the two core functions of active nose cone cooling and auxiliary thrust generation. This solves the problems of independent cooling and propulsion systems, discontinuous energy utilization, and system redundancy in existing technologies.

[0020] This invention does not employ a single protective method, but innovatively integrates a triple protection mechanism: passive thermal shielding, physical heat absorption through liquid ammonia convection, and chemical heat absorption through the high-temperature decomposition of liquid ammonia. The passive thermal shielding first blocks and dissipates some of the extreme aerodynamic heat flow; any residual heat that is not completely blocked is absorbed by the liquid ammonia within the flow channel through convection; in the extremely high heat flow region, the liquid ammonia further undergoes an endothermic decomposition reaction, absorbing a huge amount of heat through a chemical reaction heat sink. This gradient, coupled protection system of passive, active physical, and active chemical heat transfer can cope with more extreme and sustained aerodynamic heating, effectively preventing structural thermal deformation and failure of the nose cone due to overheating.

[0021] This invention changes the energy waste model of traditional active cooling systems where the cooling working fluid is discarded after a single use. The aerodynamic heat absorbed by liquid ammonia in the cooling channel is not wasted, but converted into the internal energy and pressure energy of the working fluid itself. Subsequently, this energy-carrying working fluid—a mixture of liquid ammonia and its decomposition products—is directly transported to the propulsion nozzle for accelerated discharge, converting part of the absorbed heat energy into propulsion kinetic energy. This continuous energy utilization model of cooling and propulsion maximizes the utilization of the limited-load working fluid energy, significantly improving the overall efficiency of the entire aircraft energy system.

[0022] The active cooling channel of this invention adopts a non-uniformly distributed spiral channel layout. Its channel density matches the non-uniformly distributed heat flux density on the nose cone surface, enabling targeted cooling of high heat flux areas and achieving optimized allocation of thermal protection resources and precise thermal control. Simultaneously, by adjusting parameters such as the liquid ammonia delivery flow rate, the cooling intensity and the working fluid state entering the nozzle can be controlled synchronously, thereby achieving dynamic and coordinated control of cooling efficiency and thrust output. This allows the system to adapt to the differentiated needs of the aircraft under various flight conditions. Attached Figure Description

[0023] The accompanying drawings, which form part of this invention, are used to provide a further understanding of the invention. The illustrative embodiments of the invention and their descriptions are used to explain the invention and do not constitute an undue limitation of the invention. In the drawings: Figure 1 This is a schematic diagram of the overall structure of an integrated active cooling and propulsion system for a hypersonic glider according to the present invention; Figure 2 This is a schematic diagram showing the distribution of the cooling channels described in this invention.

[0024] In the picture: 1-Spring, 2-Piston, 3-First ammonia storage tank, 4-First check valve, 5-Ammonia preheating and decomposition cylinder, 6-Heating source, 7-Second ammonia storage tank, 8-Pressure limiting valve, 9-Throttle valve, 10-Second check valve, 11-Auxiliary propulsion nozzle, 12-Passive thermal protection layer, 13-High heat flux zone of nose cone, 14-Nose cone. Detailed Implementation

[0025] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. It should be noted that, unless otherwise specified, the embodiments and features in the embodiments of the present invention can be combined with each other, and the described embodiments are only some embodiments of the present invention, not all embodiments.

[0026] Implementation Method 1: This implementation method addresses the problems of insufficient working fluid heat sink, single energy utilization, and incompatibility with propulsion in traditional active cooling. It proposes an integrated active cooling and propulsion system for hypersonic gliders. The system includes a passive thermal protection layer 12, an active cooling channel structure, an ammonia delivery module, an auxiliary propulsion nozzle module 11, and a nose cone 14. The passive thermal protection layer 12 is attached to the outer surface of the nose cone 14 of the aircraft to block external extreme aerodynamic heat flow. The active cooling channel structure is arranged between the passive thermal protection layer 12 and the nose cone 14; The ammonia delivery module includes a first ammonia storage tank 3, a throttle valve 9, and a second check valve 10 connected in sequence by pipelines. The outlet of the ammonia delivery module is connected to the inlet of the active cooling channel structure and is used to deliver liquid ammonia to the active cooling channel structure. The outlet of the active cooling channel structure is connected to the inlet of the auxiliary propulsion nozzle module 11; The auxiliary propulsion nozzle module 11 is used to accelerate the ejection of the working propellant from the active cooling channel structure.

[0027] In this embodiment, the ammonia delivery module further includes a spring accumulator, which includes a spring 1, a piston 2, and an ammonia preheating and decomposition cylinder 5. The spring 1 acts on the piston 2, and the piston 2 divides the ammonia preheating and decomposition cylinder 5 into two non-communicating parts. The outlet of the first ammonia storage tank 3 is connected to the ammonia preheating and decomposition cylinder 5 through a first check valve 4.

[0028] In practical applications, the mechanical potential energy pre-stored in spring 1 is converted into pressure energy of the working fluid in ammonia preheating and decomposition cylinder 5 via piston 2, providing a stable and controllable initial flow rate and pressure for the system. Compared with pumping systems that rely on continuous electricity or external air sources, this simplifies the system configuration and improves robustness in the complex environment of high-speed aircraft.

[0029] In this embodiment, the ammonia delivery module further includes a second ammonia storage tank 7 and a pressure limiting valve 8; the inlet of the second ammonia storage tank 7 is connected to the ammonia preheating and decomposition cylinder 5, and its outlet is connected to the inlet of the throttle valve 9; the pressure limiting valve 8 is disposed on the second ammonia storage tank 7.

[0030] A second ammonia storage tank 7 and a pressure relief valve 8 were added to achieve pressure buffering and overload protection for the conveyed working medium. The effect is to make the pressure of liquid ammonia delivered to the cooling channel more stable, avoid pressure fluctuations affecting the uniformity of cooling, and at the same time, the pressure relief valve prevents the system from being damaged by overpressure.

[0031] In this embodiment, the ammonia delivery module further includes: the active cooling channel structure is a spiral channel arranged on the surface of the nose cone 14, and the distribution density of the spiral channel on the surface of the nose cone 14 is non-uniform.

[0032] By defining the active cooling channel as a non-uniformly distributed spiral channel, the high heat flux density of the external aerodynamic heat load at the nose cone stagnation point and leading edge is used as the design input. The channel space topology that matches it is optimized in reverse, namely a spiral adaptive surface with non-uniform density matching heat flux. In principle, this optimizes the utilization efficiency of the cooling medium and avoids regional failures caused by overcooling or undercooling.

[0033] In this embodiment, the cross-section of the spiral flow channel is circular, and its inner wall is provided with a turbulence enhancement structure.

[0034] Circular cross-sections are a common choice for reducing flow resistance, while turbulence-enhancing structures such as internal threads and turbulence-increasing columns aim to disrupt the flow boundary layer, causing strong mixing of the fluid and thus greatly enhancing the rate of heat transfer from the nose cone to liquid ammonia.

[0035] In this embodiment, the working medium transported by the ammonia delivery module is liquid ammonia; after the liquid ammonia flows through the active cooling channel structure, it enters the auxiliary propulsion nozzle module 11 in the form of a mixed working medium containing liquid ammonia and its decomposition products.

[0036] Liquid ammonia undergoes a physical phase change, sensible heat absorption, and chemical transformation within the flow channel. This process not only removes a massive amount of heat, achieving cooling, but more importantly, it alters the chemical composition and energy form of the working fluid, producing highly reactive hydrogen gas. This results in the working fluid exiting the cooling end having a much higher propulsion potential than undecomposed pure liquid ammonia.

[0037] In this embodiment, the auxiliary propulsion nozzle module 11 is a Laval nozzle structure.

[0038] The propulsion nozzle is designed as a Laval nozzle, which enables efficient expansion and acceleration of the high-temperature mixed gas produced by ammonia decomposition from subsonic to supersonic speeds, thereby maximizing the conversion of the thermal and pressure energy of the working fluid into directional thrust.

[0039] In this embodiment, the system's workflow is as follows: liquid ammonia is output from the ammonia delivery module, flows through the active cooling channel structure, and then enters the auxiliary propulsion nozzle module 11 to be accelerated and ejected.

[0040] The design principle of existing technologies is function-oriented and specialized, meaning that the cooling system uses a dedicated coolant and the propulsion system uses a dedicated fuel, with the two systems operating independently. The core principle of the system proposed in this embodiment lies in establishing a new paradigm of working fluid-oriented, functionally sequential operation. It uses liquid ammonia as a single medium as the energy carrier throughout the entire process, allowing it to play two roles sequentially within the system: first, as a cooling working fluid within the nose cone channel, absorbing aerodynamic heat through physical and chemical processes; then, seamlessly converting into propulsion fuel, carrying the absorbed heat energy into the nozzle to perform work. This design creates a closed-loop energy utilization system from heat absorption to work generation, breaking down the functional barriers between cooling and propulsion.

[0041] Traditional convective cooling primarily relies on the physical heat capacity and latent heat of vaporization of the working fluid, resulting in limited heat sink capacity. This embodiment proposes a system with a deeper thermal management principle: a sequential and combined utilization of physical and chemical heat sinks. Liquid ammonia first utilizes its high specific heat capacity for convective heat transfer. When the heat flux density is extremely high and physical cooling is insufficient, the high-temperature environment within the flow channel triggers an endothermic decomposition reaction of the liquid ammonia. This chemical reaction process itself absorbs a large amount of heat, resulting in a heat sink value far exceeding that of simple physical heating. This principle significantly increases the heat load capacity per unit mass of working fluid, making it particularly suitable for hypersonic extreme thermal environments.

[0042] Existing thermal protection designs often homogenize the protected area. The system proposed in this embodiment is based on a more refined principle: the configuration strength of the cooling structure must dynamically match the spatial distribution function of the external heat load. Specifically, the principle is as follows: first, the non-uniform heat flow field on the three-dimensional surface of the nose cone is identified, i.e., the high heat flow region is near the stagnation point; then, based on this, a spiral cooling channel network with a non-uniform distribution density is designed, with denser and more optimized channels in the high heat flow region and sparser channels in the low heat flow region. This coupled design principle of distributing cooling power on demand achieves optimal local temperature control and overall thermal protection performance while minimizing working fluid consumption and system weight, which is the core of system efficiency optimization.

[0043] Implementation Method 2, see below Figure 1 and Figure 2 This embodiment describes a complete implementation of the integrated active cooling and propulsion system for hypersonic gliders described in Embodiment 1, including: An integrated active cooling and propulsion system for a hypersonic glide vehicle, the system The system includes a spring 1, a piston 2, a first ammonia storage tank 3, a first check valve 4, an ammonia preheating and decomposition cylinder 5, a heating source 6, a second ammonia storage tank 7, a pressure limiting valve 8, a throttle valve 9, a second check valve 10, an auxiliary propulsion nozzle 11, a passive thermal protection layer 12, a high-heat flux zone 13 in the nose cone, and a nose cone 14. The spring accumulator 1 is connected to the liquid ammonia storage tank 2 to provide initial pressure for liquid ammonia delivery. The outlet of the piston 2 is connected to the inlet of the ammonia preheating and decomposition cylinder 5 via the first check valve 4. The ammonia preheating and decomposition cylinder 5 has heat exchange coils wound around its outer side and is heated by the heating source 6 to achieve preheating and preliminary decomposition of liquid ammonia. The outlet of the ammonia preheating and decomposition cylinder 5 is connected to the second ammonia storage tank 7. The second ammonia storage tank 7 stabilizes the working fluid delivery pressure through a spring and piston structure. The outlet of the second ammonia storage tank 7 is connected to the active cooling channel inside the nose cone 14 after passing through the pressure limiting valve 8, the throttle valve 9, and the second check valve 10. The nose cone 14 is wrapped around the outside of the high heat flow zone 13 of the nose cone, and the outside is covered with a passive thermal protection layer 12. The active cooling channel inside is connected to the auxiliary propulsion nozzle 11. The passive thermal protection layer 12 is used to block the external extreme aerodynamic heat flow. The liquid ammonia working fluid in the active cooling channel achieves precise temperature control of the nose cone through convective heat transfer and chemical decomposition heat absorption. The liquid ammonia and decomposition products that have completed the cooling function are accelerated and discharged through the auxiliary propulsion nozzle 11 to generate auxiliary thrust.

[0044] This embodiment employs liquid ammonia as a dual-function working fluid, simultaneously achieving active cooling and auxiliary propulsion of the nose cone. This overcomes the limitations of traditional separate designs for thermal protection and propulsion systems, providing an integrated power and thermal protection solution for hypersonic gliders. A triple thermal protection system is integrated, combining passive thermal protection layer barrier, liquid ammonia convective heat transfer physical heat absorption, and liquid ammonia high-temperature decomposition chemical heat absorption. This significantly improves thermal protection efficiency under extreme aerodynamic thermal environments, preventing thermal deformation and failure of the nose cone structure. After heat absorption and decomposition within the cooling channel, the liquid ammonia working fluid is directly used as propulsion fuel, eliminating the need for a separate fuel supply system. This allows for cascaded utilization of the working fluid's energy, improving the overall system energy efficiency and structural integration. A non-uniform spiral flow channel layout precisely matches the non-uniform heat flow distribution of the nose cone, combined with a customized Laval nozzle design, enabling dynamic control of cooling efficiency and thrust output to adapt to different flight conditions.

[0045] In this embodiment, the throat diameter and expansion ratio of the Laval nozzle are optimized by the gas dynamic parameters of ammonia decomposition products to meet the thrust requirements of hypersonic gliders under different flight conditions. The thrust output changes linearly with the liquid ammonia decomposition rate, achieving precise control of propulsion performance.

[0046] In practical applications, the system sets up three-dimensional evaluation indicators: cooling efficiency, thrust efficiency, and overall system efficiency. Through weighted allocation, it achieves synergistic optimization of cooling and propulsion performance. The overall system efficiency, with cooling and thrust efficiency as its core, incorporates thermo-mechanical coupling characteristics to construct a quantitative evaluation system, ensuring balanced system performance under extreme thermo-mechanical coupling environments. The evaluation system includes the thermal insulation performance of the passive thermal protection layer as a core consideration, constructing a quantitative model for the efficiency of passive-active coupled thermal protection. This evaluation system achieves real-time monitoring of the system's operating status through quantitative calculations, providing data support for system optimization.

[0047] Furthermore, this embodiment also proposes an integrated active cooling and propulsion design method for hypersonic gliders, including the following steps: Step 1: Determine the heat flux load and thrust requirements of the aircraft's nose cone, and clarify the core performance indicators of integrated active cooling and propulsion; Step 2: Design the material and thickness of the passive thermal protection layer to match the initial thermal barrier requirements of the nasal cone; Step 3: Lay out the active cooling channel assembly. Determine the number, distribution, and size of the channels based on the heat flow distribution of the nose cone, and configure the inner wall of the channels with high thermal conductivity. Step 4: Determine the transport parameters and decomposition reaction conditions for the liquid ammonia working fluid to achieve simultaneous cooling and decomposition. Step 5: Design the propulsion nozzle module and optimize the nozzle structure parameters to adapt to the characteristics of ammonia decomposition products; Step 6: Establish a comprehensive system efficiency evaluation system and complete the overall system assembly and performance debugging.

[0048] Implementation Method 3: This implementation method proposes a nose cone component for an aircraft, including an integrated active cooling and propulsion system for hypersonic gliders as described in any one of Implementation Methods 1 to 2.

[0049] Implementation Method 4: This implementation method proposes a hypersonic gliding vehicle, including the nose cone component of the vehicle as described in Implementation Method 1 to Implementation Method 2.

[0050] The specific embodiments of the present invention disclosed above are merely illustrative of the invention. These embodiments do not exhaustively describe all details, nor do they limit the invention to the specific embodiments described. Many modifications and variations can be made based on the content of this specification. This specification selects and specifically describes these embodiments to better explain the principles and practical applications of the invention, thereby enabling those skilled in the art to better understand and utilize the invention.

Claims

1. An integrated active cooling and propulsion system for a hypersonic glider, characterized in that, The system includes a passive thermal protection layer (12), an active cooling channel structure, an ammonia delivery module, an auxiliary propulsion nozzle module (11), and a nose cone (14). The passive thermal protection layer (12) is attached to the outer surface of the nose cone (14) of the aircraft to block external extreme aerodynamic heat flow; The active cooling channel structure is arranged between the passive thermal protection layer (12) and the nose cone (14); The ammonia delivery module includes a first ammonia storage tank (3), a throttle valve (9), and a second check valve (10) connected in sequence by pipelines. The outlet of the ammonia delivery module is connected to the inlet of the active cooling channel structure and is used to deliver liquid ammonia to the active cooling channel structure. The outlet of the active cooling channel structure is connected to the inlet of the auxiliary propulsion nozzle module (11); The auxiliary propulsion nozzle module (11) is used to accelerate the ejection of the working fluid from the active cooling channel structure.

2. The integrated active cooling and propulsion system for a hypersonic glider according to claim 1, characterized in that, The ammonia delivery module also includes a spring accumulator, which includes a spring (1), a piston (2) and an ammonia preheating decomposition cylinder (5); the spring (1) acts on the piston (2), and the piston (2) divides the ammonia preheating decomposition cylinder (5) into two non-communicating parts; the outlet of the first ammonia storage tank (3) is connected to the ammonia preheating decomposition cylinder (5) through a first check valve (4).

3. The integrated active cooling and propulsion system for a hypersonic glider according to claim 2, characterized in that, The ammonia delivery module also includes a second ammonia storage tank (7) and a pressure limiting valve (8); the inlet of the second ammonia storage tank (7) is connected to the ammonia preheating decomposition cylinder (5), and its outlet is connected to the inlet of the throttle valve (9); the pressure limiting valve (8) is installed on the second ammonia storage tank (7).

4. The integrated active cooling and propulsion system for a hypersonic glider according to claim 1, characterized in that, The ammonia delivery module further includes: the active cooling channel structure is a spiral channel arranged on the surface of the nose cone (14), and the distribution density of the spiral channel on the surface of the nose cone (14) is non-uniform.

5. The integrated active cooling and propulsion system for a hypersonic glider according to claim 4, characterized in that, The spiral flow channel has a circular cross-section, and its inner wall is provided with a turbulence enhancement structure.

6. The integrated active cooling and propulsion system for a hypersonic glider according to claim 1, characterized in that, The working medium transported by the ammonia delivery module is liquid ammonia; after the liquid ammonia flows through the active cooling channel structure, it enters the auxiliary propulsion nozzle module (11) in the form of a mixed working medium containing liquid ammonia and its decomposition products.

7. The integrated active cooling and propulsion system for a hypersonic glider according to claim 1, characterized in that, The auxiliary propulsion nozzle module (11) is a Laval nozzle structure.

8. The integrated active cooling and propulsion system for a hypersonic glider according to claim 1, characterized in that, The system's workflow is as follows: liquid ammonia is output from the ammonia delivery module, flows through the active cooling channel structure, and then enters the auxiliary propulsion nozzle module (11) to be accelerated and ejected.

9. A nose cone component for an aircraft, characterized in that, Including the integrated active cooling and propulsion system for hypersonic gliders as described in any one of claims 1-8.

10. A hypersonic gliding vehicle, characterized in that, Includes the aircraft nose cone component as described in claim 9.

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