A matching modular air-breathing engine's axisymmetric inlet

Abstract

This invention discloses an axisymmetric air intake for a modular air-breathing engine, relating to the field of air intake technology. The axisymmetric air intake for a modular air-breathing engine includes: a central body; a movable lip, coaxially fitted outside the central body; and multiple expansion sections. The central body and the movable lip form a single annular compression channel shared by all engines. Downstream of this single annular compression channel is a throat, and the annular cross-section at the throat is circumferentially divided into multiple annular fan-shaped inlets. Each annular fan-shaped inlet connects to an independent expansion section. This axisymmetric air intake for a modular air-breathing engine simplifies the integrated design of the air intake and the aircraft. By adopting a layout of "a shared single annular airflow channel before the throat," this invention requires only one integrated air intake inlet and one shared compression channel to simultaneously serve multiple downstream engines.

Description

Technical Field

[0001] This invention relates to the field of air intake technology, specifically to an axisymmetric air intake that is compatible with a modular air-breathing engine. Background Technology

[0002] The propulsion system is crucial for multi-stage-to-orbit (MSO) vehicles, which are currently primarily powered by rocket engines. However, rocket engines have low specific impulse and poor economic efficiency. If air-breathing engines were used as the primary propulsion during the atmospheric flight phase, the weight of the oxidizer carried could be significantly reduced, thereby lowering the overall weight of the vehicle and providing more space for mission payloads. Furthermore, using air-breathing engines as the first-stage propulsion system enables conventional horizontal takeoff and landing, eliminating the need for a vertical launch tower and greatly reducing launch infrastructure costs.

[0003] Multistage-to-orbit (MTO) vehicles require sufficient thrust during flight. However, the thrust of a single air-breathing engine is currently limited. Combining multiple air-breathing engines into a modular engine is the most direct and effective way to solve the thrust shortage problem. The air intake is an important component of an air-breathing engine, and it needs to provide the engine with high-quality airflow of specific flow rate, pressure, and temperature during operation.

[0004] Currently, air-breathing engine inlet design technology is typically developed for the needs of a single engine. The design process uses the flow rate, pressure ratio, and other requirements of a single engine as inputs to determine parameters such as the intake capture area, compression surface profile, and throat dimensions, ultimately forming an independent inlet design to supply air to a single engine. This "one engine, one inlet" design approach is widely used in single-engine or twin-engine aircraft, and the technology is relatively mature. However, when the power system uses multiple engines combined into a module, following the above design approach requires designing an independent inlet for each engine.

[0005] Against this backdrop, when using a modular propulsion system consisting of multiple identical engines operating simultaneously as the power source for a multi-stage-to-orbit vehicle, the core challenge of the existing "one engine, one intake" approach is not simply airflow distribution, but rather ensuring complete consistency in key parameters such as total temperature, total pressure, Mach number, and flow rate at the inlet of each engine in a parallel configuration. This "consistency in air supply" is a prerequisite for ensuring coordinated power output from each module and avoiding thrust eccentricity, but it has never been addressed in existing single-engine matching designs.

[0006] Existing intake manifold design technologies are typically developed for the needs of a single engine. When a powertrain consists of multiple engines forming a module, following this approach would require matching each engine with an independent intake manifold, which leads to the following drawbacks and shortcomings:

[0007] First, the design of multiple air intakes increases the difficulty of integrating the air intakes with the aircraft, and brings more complex structural problems, including additional installation space requirements, increased weight, and structural interference between the air intakes.

[0008] Second, while multiple air intakes can theoretically provide airflow to the engine independently, mutual interference exists between them in the external flow field of the aircraft, especially under supersonic flight conditions. The shock waves generated by each air intake intertwine, forming a complex and unpredictable interference flow field. This interference can cause each air intake to malfunction as expected, and due to the complexity of the interference flow field, it is difficult to accurately assess its performance. Consequently, it is impossible to guarantee the provision of high-quality airflow to the downstream engine, or even to guarantee the normal start-up and operation of the air intakes.

[0009] Third, when multiple engines are operating together, each engine requires a high degree of consistency in the quality of the incoming airflow. If multiple independent air intakes are used to supply air separately, it is difficult to ensure that the airflow quality entering each engine is consistent due to the different flow field positions and degrees of interference of each air intake. This results in inconsistent thrust output from each engine, making it impossible for them to output sufficient thrust as expected.

[0010] In summary, the existing design technology of matching a single air intake with a single engine is difficult to apply directly to modular air-breathing engine power systems. There is an urgent need for an air intake design technology that can solve the above-mentioned air supply problems in a comprehensive manner at the system level. Summary of the Invention

[0011] To address the shortcomings of existing technologies, this invention provides an axisymmetric air intake duct that matches a modular air-breathing engine, thus solving the problems in the background technology.

[0012] To achieve the above objectives, the present invention provides the following technical solution: an axisymmetric intake duct for a modular air-breathing engine, comprising: a central body; a movable lip, coaxially fitted outside the central body; and multiple expansion sections; wherein the central body and the movable lip form a single annular compression channel shared by all engines, the downstream of the single annular compression channel is a throat, the annular cross-section of the throat is equally divided into multiple annular fan-shaped inlets along the circumference, each annular fan-shaped inlet is connected to an independent expansion section, and each expansion section delivers airflow to the corresponding downstream engine.

[0013] Furthermore, the outer contour surface of the central body forms a single-level or sequentially connected multi-level conical surface along the airflow direction.

[0014] Furthermore, the areas of each of the aforementioned annular fan-shaped inlets are equal, and the number of the annular fan-shaped inlets is equal to the number of downstream engines.

[0015] Furthermore, each expansion section has a connection interface at its downstream end, the shape and size of which are consistent with the engine inlet, and it directly connects to the front end of the engine casing.

[0016] Furthermore, the end of the lip cover along the airflow direction is cylindrical, and the range of movement of the lip cover is limited to the axial length of the cylinder; each annular fan-shaped inlet forms a ring structure so that the lip cover can still align with the annular fan-shaped inlet when it moves along the airflow direction.

[0017] Furthermore, the lip cover has an inner surface and an outer surface. A single annular compression channel is formed by the inner surface extending from the front edge to the rear and together with the outer conical surface of the central body. The inner surface and the outer surface meet at the front edge to form the lip cover front edge line, and the specific shape of the front edge line is a pointed front edge or a blunt front edge.

[0018] Furthermore, the area change rate, length, outlet cross-sectional shape, and area expansion ratio of each expansion segment are exactly the same, only the spatial azimuth angle is different.

[0019] Furthermore, a vent hole or vent slit is provided near the throat of the tail of the central body.

[0020] Furthermore, the throat Mach number is between 1.3 and 2.0.

[0021] Furthermore, the axisymmetric air intake is installed on the nose, belly, sides, or back of the aircraft.

[0022] The present invention has the following beneficial effects:

[0023] (1) The axisymmetric air intake of this modular air-breathing engine simplifies the integrated design difficulty of the air intake and the aircraft. By adopting a layout of "a single annular airflow channel shared in front of the throat", this invention only requires the design of one integral air intake inlet and one set of shared compression channels to simultaneously serve multiple downstream engines. Compared with the traditional approach of designing independent air intakes for each engine, this invention significantly reduces the number of air intakes, reduces the structural coupling complexity between the air intake and the aircraft body, reduces structural weight, and saves installation space, thereby significantly reducing the difficulty of integrated design of the air intake and the aircraft.

[0024] (2) The axisymmetric air intake of this modular air-breathing engine avoids the problem of mutual interference of external shock waves between multiple air intakes. This invention has only one air intake inlet within the entire flight envelope, and the entire power system generates only one set of external shock waves. Compared with the traditional scheme where multiple air intakes generate shock waves that intertwine and form complex interference flow fields during supersonic flight, this invention fundamentally eliminates the problem of shock wave interference between air intakes. The external flow field structure of a single inlet is clear and controllable, and the working state of the air intake can be accurately evaluated, ensuring that the air intake can stably capture sufficient high-quality airflow under various flight conditions, avoiding the risk of the air intake failing to start or becoming unstable due to shock wave interference.

[0025] (3) The axisymmetric intake duct of this modular air-breathing engine achieves absolute consistency in the air supply conditions of multiple engines without control delay. In this invention, all incoming flows undergo deceleration and pressurization through the same compression channel, and are passively diverted to each expansion section at the throat through a completely symmetrical circumferentially divided structure. Since all airflows before the throat undergo the same compression process, the airflow parameters at the throat are uniform in the circumferential direction; and since the geometric configuration of each branch after the throat is completely symmetrical, the flow loss and diffusion characteristics of each branch are completely identical. The combination of these two factors ensures that the total temperature, total pressure, flow rate, and velocity field distribution obtained by each engine are completely consistent within the engineering error range. More importantly, this diversion process is completely passive, requiring no sensor feedback or actuator response, and there is no control delay or adjustment deviation problem, fundamentally guaranteeing the synchronization of air supply to each engine. Compared with the traditional solution where multiple independent intake ducts result in differences in air supply quality due to different flow field positions and degrees of interference, the advantages of this invention are particularly prominent. Attached Figure Description

[0026] Figure 1 This is an overall diagram of the present invention;

[0027] Figure 2 This is the front view of the present invention;

[0028] Figure 3 This is an axial half-sectional view of the present invention;

[0029] Figure 4 This is a graph showing the variation of the flow coefficient of different channels with the incoming Mach number according to the present invention;

[0030] Figure 5 This is a graph showing the variation of Mach number at the exit of different channels with the incoming Mach number according to the present invention;

[0031] Figure 6 This is a graph showing the variation of the outlet pressure ratio of different channels with the incoming Mach number according to the present invention;

[0032] Figure 7 This is a graph showing the variation of the total pressure recovery coefficient at the outlet of different channels as a function of the incoming Mach number.

[0033] In the diagram, 1 is the central body, 2 is the lip cover, 3 is the dilation section, 4 is the throat, 5 is the #1 channel outlet, 6 is the #2 channel outlet, 7 is the #3 channel outlet, and 8 is the #4 channel outlet. Detailed Implementation

[0034] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0035] It should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," "counterclockwise," "axial," "radial," and "circumferential" indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the present invention or simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the present invention.

[0036] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified.

[0037] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments.

[0038] Furthermore, to better illustrate this application, numerous specific details are provided in the following detailed embodiments. Those skilled in the art should understand that this application can be implemented without certain specific details. In some instances, methods, means, components, and circuits well-known to those skilled in the art have not been described in detail in order to highlight the main points of this application.

[0039] The following is based on Figures 1-7 This invention describes an axisymmetric air intake duct for a modular air-breathing engine, as provided in an embodiment of the present invention.

[0040] Please refer to Figures 1-7This invention provides an axisymmetric air intake duct for a modular air-breathing engine, comprising a central body 1, a lip cover 2, and an expansion section 3. The air intake duct is axisymmetric in overall configuration, with its axis of symmetry extending along the airflow direction of the air intake itself.

[0041] The movable lip cover 2 is coaxially sleeved on the outside of the central body 1; wherein, the central body 1 and the movable lip cover 2 form a single annular compression channel shared by all engines, and the downstream of the single annular compression channel is the throat 4. The annular cross section at the throat 4 is divided into multiple annular fan-shaped inlets along the circumference, and each annular fan-shaped inlet is connected to an independent expansion section 3. Each expansion section 3 delivers airflow to the corresponding downstream engine.

[0042] The installation orientation (including the relative angle between its axis of symmetry and the longitudinal axis of the aircraft) of the axisymmetric air intake for matching modular air-breathing engines provided by the present invention is determined by the overall layout of the aircraft and can be installed at the nose, belly, sides or back of the aircraft, etc. The present invention does not limit this.

[0043] When viewed along the airflow direction, the air intake channel of the present invention exhibits a layout feature of first "combining" and then "separating".

[0044] The annular compression channel is a shared compression channel section before the throat 4. The central body 1 is fixedly installed at the front end of the air intake. Its outer contour surface forms a single-stage or sequentially connected multi-stage conical surface along the airflow direction (the figure shows a single-stage conical surface). The multi-stage conical surface is used to compress the supersonic incoming flow in multiple stages (i.e., external pressure section compression). It should be noted that the number of stages of the conical surface (single, second, or third stage and above) and the specific cone angle value of each stage depend on the overall compression requirements of the aircraft and can be freely selected by the designer according to the specific operating conditions. This invention does not impose a unique limitation on this.

[0045] The lip cover 2 is an annular shell structure, coaxially fitted onto the outside of the central body 1, with its leading edge forming the physical entrance of the air intake. The lip cover 2 has an inner surface and an outer surface: the inner surface extends rearward from the leading edge and together with the outer conical surface of the central body 1 forms an annular compression channel, and its geometric profile (such as the contraction angle) is determined according to the expected compression requirements of the inner pressure section, which is the key surface for realizing the core function of this invention; the outer surface is designed according to the overall aerodynamic layout and structural strength requirements of the aircraft to control the external flow shock waves and flight drag of the lip cover 2. This invention does not limit the specific shape of the outer surface. The inner surface and the outer surface meet at the leading edge to form the leading edge line of the lip cover 2. The specific shape of the leading edge line (such as a sharp leading edge or a blunt leading edge; at relatively low flight speeds, when the aerodynamic thermal environment will not cause significant ablation or deformation of the leading edge material, a sharp leading edge can be used; at higher flight speeds, when the thermal load is high, a blunt leading edge can be used) is determined comprehensively based on the requirements of thermal protection and structural strength. The wall thickness of the lip cover 2 is reduced as much as possible while ensuring structural strength and rigidity to reduce structural weight. However, this invention does not limit the specific wall thickness value of the lip cover 2, nor does it require it to be a structure with uniform wall thickness. All incoming flow passes through the single annular compression channel formed by the conical surface of the central body 1 and the inner surface of the lip cover 2, and then enters the throat 4. Before the throat 4, the airflow channel is an integral single channel shared by all downstream engines.

[0046] It should be noted that the end of the lip cover 2 along the airflow direction is cylindrical, and the range of movement of the movable lip cover 2 is limited to the axial length of the cylinder; each annular fan-shaped inlet forms a ring structure so that the lip cover 2 can still align with the annular fan-shaped inlet when it moves along the airflow direction.

[0047] In addition, the throat 4 mentioned above is located inside the cylindrical structure at the end of the lip cover 2 along the airflow direction. It is a channel, and the length of the throat 4 channel changes when the lip cover 2 moves.

[0048] Starting from the throat section 4, the process enters the "splitting" stage, namely the independent expansion section 3. The annular airflow channel is circumferentially divided into multiple annular fan-shaped inlets (the number of divisions equals the number of downstream engines), and each annular fan-shaped inlet connects to an independent expansion section 3. Each expansion section 3 gradually transitions the airflow from the annular fan-shaped section at the throat 4 to the inlet section of the downstream engine, and completes subsonic diffusion within the expansion section 3, ultimately delivering the compressed, high-quality airflow to each engine. In summary, after the throat 4, the single annular compression channel splits into multiple independent sub-channels (i.e., expansion sections 3), each serving a downstream engine. The downstream end of each expansion section is equipped with a connection interface, the shape and size of which are consistent with the engine inlet, allowing direct docking with the front end of the engine casing.

[0049] Preferably, each expansion section 3 is located between the throat 4 and the corresponding downstream engine, smoothly transitioning the airflow from the annular fan-shaped section to the corresponding engine inlet section.

[0050] The core technical concept of this invention lies in the following: before the intake throat 4, there is a single annular compression channel. After all the incoming airflow passes through the same annular compression channel, it is divided into multiple annular fan-shaped inlets along the circumference at the throat 4. Each inlet is distributed to each downstream engine via an independent expansion section. This "combined-divided" flow channel layout exists independently of specific angles and area values, constituting the core of the protection of this invention.

[0051] Design point profile parameters determination (total flow capture and throat 4 design)

[0052] The cruise state is used as the design point for the air intake. The specific Mach number and cruise altitude are determined by the overall mission profile of the aircraft, and this invention does not impose a unique limitation on them. As an example, a flight Mach number Ma∞ = 4.0 and an altitude H = 25km can be selected.

[0053] Step 1: Determine the capture area and lip cover position

[0054] Based on the total flow requirements of all downstream engines (single engine flow requirement × number of engines) and the incoming flow parameters at the design point, the intake duct inlet capture area A0 is determined, and then the leading edge radius of the lip cover 2 is determined. The design point follows the full flow capture principle, that is, the flow captured at the intake duct inlet is exactly equal to the total air supply flow requirement of all downstream engines at the design point, thereby determining the axial reference position of the lip cover 2 at the design point.

[0055] Step 2: Determine the parameters of the central cone surface

[0056] The central body 1 employs one or more stages of conical compression. The number of stages and the semi-cone angles of each stage are determined based on the comprehensive optimization results of the expected compression ratio and the total pressure recovery coefficient. As a preferred example and not a limitation, when a three-stage conical surface is used, the semi-cone angles of the three stages can increase sequentially from front to back, for example, taking 8°, 12°, and 16° respectively, to achieve gradual compression from weak to strong, thereby reducing the total pressure loss of the external shock wave. However, those skilled in the art should understand that if the overall compression requirements of the aircraft are different, a two-stage conical surface (such as 15° and 20°) or a four-stage conical surface can also be used, or an isentropic compression surface can be used instead of a polygonal conical surface. All of the above variations do not depart from the core concept of the present invention of "shared external pressure section".

[0057] Step 3: Determine the contraction angle of the inner surface of the lip cover 2 and the area of ​​the throat passage 4.

[0058] The inner surface of the lip cover 2 has a compression angle that turns the airflow inward after the leading edge of the inlet, forming an internal shared compression channel section together with the conical surface of the central body 1, which further compresses the incoming flow to the throat 4. Based on the expected Mach number of the throat 4 (e.g., Ma≈1.3~2.0, preferably around 1.5), the required cross-sectional area A of the throat 4 is determined using a one-dimensional isentropic relationship. th The specific value of the compression angle inside the lip cover 2 is adjusted according to the design conditions, and the present invention does not limit the value. As an illustrative range, when the design Mach number is between 3 and 6, the compression angle is usually selected in the range of 3° to 10°, but other values ​​outside this range may also be applicable to specific design requirements and do not affect the essence of the present invention.

[0059] (III) Design of the posterior branch channel and dilation section of the larynx

[0060] The throat passage 4 section is an axisymmetric annular section. Its inner diameter is determined by the radius of the tail of the central body, and its outer diameter is determined by the radius of the inner wall of the lip mask throat passage 4.

[0061] The annular cross-section of throat 4 is divided into N equal parts along the circumference, where N is equal to the number of downstream modular engines (N=4 in this embodiment). Each part corresponds to an annular fan-shaped inlet, and its area is A. th / N ensures that each downstream engine receives the same flow share. If the number of downstream engines changes (e.g., 3 or 6), the flow is divided into 3 or 6 equal parts accordingly. The design method is completely analogous and does not depart from the protection scope of this invention.

[0062] Starting from the throat section 4, each annular fan-shaped inlet extends rearward to form an independent expansion section 3, gradually transitioning the airflow from the annular fan-shaped section at the throat 4 to the inlet section of the downstream engine. The downstream engine inlet is usually circular (determined by the engine's own design), but if the engine inlet is rectangular, elliptical, or other shapes, the outlet section of the expansion section 3 can be matched accordingly. Only the transition section profile needs to be adjusted, without affecting the basic design concept of this invention.

[0063] The airflow channel area of ​​each expansion section 3 gradually increases along the flow direction, achieving subsonic diffusion. The area expansion ratio of each expansion section 3 is determined by the ratio of the annular fan-shaped area of ​​the throat 4 to the corresponding engine inlet area. The specific value varies depending on the downstream engine inlet size, and this invention does not impose a unique limitation on this. The geometric configuration of the expansion section 3 adopts a variable cross-section flow channel design that smoothly transitions from the annular fan shape to the target cross-sectional shape. The wall profile can be generated using polynomial curves, B-spline curves, or any smooth curve optimized according to the area law, as long as the area change rate and curvature are continuous and flow separation is avoided. The specific algorithm for generating the wall profile is a conventional choice for designers and does not affect the essence of this invention.

[0064] Each expansion section 3 has a connection interface at its downstream end, the shape and size of which are consistent with the engine inlet, and can be directly connected to the front end of the engine casing.

[0065] Those skilled in the art should understand that the geometric configurations (including area change rate, outlet shape, expansion ratio, etc.) of each expansion section 3 after the throat 4 in this invention are completely identical, differing only in spatial azimuth angle. This is a key structural feature to ensure consistent air supply to each engine. This design concept is fundamentally different from the asymmetric branch channel design used in the intake duct of a combined engine to adapt to different thermodynamic cycle requirements. The latter typically has different geometric parameters for each branch channel and requires an active adjustment mechanism to achieve mode switching.

[0066] (iv) Flow regulation under non-design conditions

[0067] During climb, acceleration, or maneuvering flight, the incoming Mach number and altitude deviate from the design point, and the downstream engine's demand for incoming airflow also changes accordingly. At this time, the actual airflow at the air intake is changed by shifting the lip cover 2, thereby achieving airflow matching.

[0068] The specific implementation method is as follows: During the design phase, for several typical non-design operating conditions within the flight envelope (including different Mach numbers, different altitudes, and their combinations), the axial displacement ΔL of the lip cover 2 relative to the design point position is calculated through CFD numerical simulation or wind tunnel testing. During flight, the control system, based on the real-time flight status, translates the lip cover 2 to a predetermined position via an actuator (hydraulic actuator, electric actuator, or any mechanical device capable of axial displacement; this invention does not limit the driving method; the lip cover 2 is generally fixed to a mechanical device capable of axial displacement, and generally not fixed to the central body). This translation distance can be positive or negative. During the translation of the lip cover 2, its external surface position changes accordingly, which may cause changes in the external flow state and local drag of the lip cover. This drag change can be calculated and evaluated during the design phase and included in the aircraft trim drag budget, without affecting the realization of the core function of this invention.

[0069] (v) Central body 1 auxiliary measures

[0070] As a further optimization measure, several vent holes or vent slots can be provided near the throat 4 at the tail of the central body 1 to discharge part of the boundary layer airflow when the inlet has difficulty starting, thus assisting the inlet to overcome the critical starting state. The number, diameter, and arrangement of the vent holes can be determined based on specific flow simulation results, and this invention does not impose specific limitations on them. It should be emphasized that this venting structure is not a necessary feature for realizing the "combined-separated" flow channel layout of this invention. Even without vent holes, the basic technical solution of this invention remains complete.

[0071] (vi) Optional range and versatility of design parameters

[0072] It should be particularly emphasized that all the specific values ​​recorded in the above embodiments, including but not limited to the design Mach number of 4.0, altitude of 25km, single-unit flow rate of 34kg / s, three-stage cone angles of 8°, 12°, and 16°, lip cover contraction angle of 3° to 10°, throat Mach number of 1.5, and four engines, are merely exemplary parameters given for the purpose of understanding the present invention and do not constitute any limitation on the scope of protection of the present invention.

[0073] The core of this invention lies in the layout design of the intake duct flow channel. Specifically, before the throat 4, there is a single annular airflow channel shared by all engines. At the throat 4, the annular cross-section is circumferentially divided into multiple annular fan-shaped inlets. After the throat 4, each inlet expands independently and connects to each engine. This "combined-divided" flow channel layout exists independently of any specific design parameters such as Mach number, height, flow rate, angle, or area.

[0074] In practical engineering applications, designers can freely adjust various design variables based on factors such as the specific mission requirements of the aircraft, the type and number of engines, and overall aerodynamic layout constraints. For example, the number of stages of the cone surface of the central body 1 can be single, double, triple, or more, and the values ​​of the cone angles of each stage can be re-optimized according to the expected compression ratio; the contraction profile of the inner surface of the lip cover 2 can adopt broken lines, arcs, or spline curves, and its turning angle can be adjusted accordingly with the design Mach number; the aerodynamic shape and wall thickness distribution of the outer surface of the lip cover can be flexibly designed according to the overall structure and thermal protection requirements. The target Mach number of the throat 4 can be selected in the range of Ma = 1.2 to 2.5, and can even be extended to a wider range according to special requirements, not limited to the preferred value of 1.5. The number of downstream engines can be 2, 3, 4, 6, or any integer number, and the number of equal divisions of the annular fan after the throat 4 changes accordingly. The shape of the expansion section exit section depends on the downstream engine interface and can be circular, rectangular, elliptical, or irregular, as long as a smooth transition from the annular fan inlet is ensured. The lip cover 2 can be driven by any of the following methods: hydraulic actuation, electric drive, pneumatic drive, or mechanical linkage. This invention does not limit the method of driving the translation.

[0075] As long as the flow channel layout follows the basic technical route of "shared inlet, single channel before throat 4, equal division of throat 4 into annular fans, and independent expansion of throat 4 to each engine", it falls within the protection scope of this invention. The changes in the aforementioned parameters are merely routine engineering adjustments made by those skilled in the art based on specific operating conditions after reading this invention, requiring no creative effort and not constituting a deviation from the essence of this invention.

[0076] The core of this invention lies in its flow channel geometry topology: through a shared compression channel section before throat 4 + circumferentially symmetrically divided throat 4 + fully symmetrically independent expansion after throat 4, the airflow parameters of each branch after the diversion are naturally consistent, and this diversion process is completely passive, requiring no active adjustment components. This topology has the following three essential characteristics:

[0077] (i) Symmetrical and evenly distributed structure. The annular cross-section at throat 4 is divided into geometrically identical annular fan-shaped inlets along the circumference. Each inlet has an equal area and is rotationally symmetrical in its spatial orientation relative to the intake axis. This symmetry ensures that each branch has identical inlet boundary conditions at the compression surface outlet, which is the physical basis for achieving passive even distribution.

[0078] (ii) Passive flow diversion without active adjustment elements. From the throat 4, which is divided into equal sections, to each engine inlet, there are no moving parts, valves, or guide vanes within the flow channel. After passing through the throat 4, the airflow enters the expansion sections 3 of each branch based on its own flow inertia. The flow diversion process does not rely on any external control commands or actuators. This means that the consistency of air supply is not affected by sensor accuracy, control algorithms, or actuator response speed, fundamentally ensuring the synchronization and stability of air supply conditions for multiple engines.

[0079] (iii) The geometric parameters of the expansion sections of each branch are completely symmetrical. The area change rate, length, outlet cross-sectional shape and area expansion ratio of all expansion sections are exactly the same, only the spatial azimuth angle is different. This ensures that the diffusion process experienced by the airflow in each branch is completely consistent, and the outlet airflow parameters (total temperature, total pressure and velocity distribution) are exactly the same within the engineering error range.

[0080] The above three features fully define the technical solution of the present invention from the perspective of flow channel topology. The technical differences between it and the known combined engine intake flow splitting scheme will be further explained in the advantages section below.

[0081] To address the mismatch between engine flow requirements and actual intake flow under off-design conditions, the geometric opening area of ​​the intake inlet is altered by axially shifting the lip cover 2, thereby changing the actual airflow entering the intake and achieving a match with engine requirements. During the intake design phase, the required lip cover displacement is pre-calculated and a control law is established for each typical off-design condition within the flight envelope. During flight, the lip cover is driven to translate according to the predetermined law, without relying on real-time flow feedback.

[0082] Please refer to the appendix for detailed experimental results. Figure 4-7 The above parameters can be obtained by measurement, reading or direct calculation in CFD numerical simulation or wind tunnel test.

[0083] Figure 4The graph shows the variation of the flow coefficient of different channels in this invention with the incoming Mach number. The flow coefficient is defined as the ratio of the actual physical flow rate of the channel to the intake capture flow rate, i.e.: , i = 1~4; where, This represents the flow coefficient of the channel. This represents the physical flow of the channel, which can be directly read from the simulation results. This indicates captured traffic, where i is the channel number; captured traffic Where K is a constant with a value of 0.04042. For the Mach number of the incoming flow, This represents the flow function value for the incoming Mach number. For the total incoming pressure, The total temperature of the incoming flow is the capture area. The capture area is the area projected from the air intake onto the direction perpendicular to the incoming flow. The total pressure, total temperature, and flow rate of the incoming flow are calculated using the following formula:

[0084]

[0085]

[0086]

[0087] in, For incoming static pressure, For incoming static temperature; The specific heat ratio of air is 1.4.

[0088] Figure 5 This is a graph showing the variation of Mach number at the exit of different channels with the incoming Mach number. The exit Mach number uses the mass average value, which can be directly read from the calculation results.

[0089] Figure 6 This is a graph showing the variation of the outlet pressure ratio of different channels with the incoming Mach number, according to the present invention. Defined as the ratio of the static pressure at the intake outlet to the static pressure of the incoming flow, expressed as... Where i represents the channel number; The average static pressure at the outlet of this channel is obtained directly from the calculation results.

[0090] Figure 7 This is a graph showing the variation of the total pressure recovery coefficient at the outlet of different channels as a function of the incoming Mach number. (Outlet total pressure recovery coefficient) Defined as the ratio of the total pressure at the intake outlet to the total pressure of the incoming flow, expressed as... Where i represents the channel number; The average total pressure at the outlet of this channel is obtained directly from the calculation results.

[0091] Appendix Figure 4-7 In the diagrams, #1, #2, #3, and #4 represent the data corresponding to the exits of channels #1, #2, #3, and #4, respectively. Please refer to the appendix for the specific locations of each channel exit. Figure 1 Exit 5 of Channel #1, Exit 6 of Channel #2, Exit 7 of Channel #3, and Exit 8 of Channel #4. From Figure 4-7 It can be seen that the flow coefficient, average outlet Mach number, outlet pressure ratio, and total pressure recovery coefficient curves for the four channels are basically identical, indicating that the expansion channels at different circumferential positions achieve nearly consistent flow distribution, velocity state, pressure level, and total pressure loss under the same incoming flow conditions. Therefore, without setting independent flow control valves, guide vanes, or active distribution mechanisms, the consistency of inlet air supply parameters for multiple modular engines can be improved.

[0092] This technical solution serves multiple engines simultaneously through a single air intake, ensuring that the airflow entering each engine is compressed from the same common compression channel section, resulting in consistent total temperature and pressure. This provides unified incoming flow conditions for the coordinated operation of each engine. At the same time, it avoids the structural complexity and shock wave interference problems caused by designing separate air intakes for each engine.

[0093] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. An axisymmetric air intake duct for a modular air-breathing engine, characterized in that, include: A central body (1); a translational lip (2) coaxially fitted on the outside of the central body (1); and multiple expansion sections (3); wherein the central body (1) and the translational lip (2) form a single annular compression channel shared by all engines, the downstream of the single annular compression channel is a throat (4), the annular cross section at the throat (4) is divided into multiple annular fan-shaped inlets along the circumference, each annular fan-shaped inlet is connected to an independent expansion section (3), and each expansion section (3) delivers airflow to the corresponding downstream engine.

2. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The outer contour surface of the central body (1) forms a single-level or sequentially connected multi-level conical surface along the airflow direction.

3. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The areas of each of the aforementioned annular fan-shaped inlets are equal, and the number of the annular fan-shaped inlets is equal to the number of downstream engines.

4. The axisymmetric intake manifold for a modular air-breathing engine according to claim 3, characterized in that, Each expansion section (3) has a connection interface at its downstream end. The shape and size of the connection interface are consistent with the engine inlet and are directly connected to the front end of the engine casing.

5. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The end of the lip cover (2) along the airflow direction is cylindrical, and the range of movement of the lip cover (2) is limited to the axial length of the cylinder. Each annular fan-shaped inlet forms a circular structure so that the lip cover (2) can still align with the annular fan-shaped inlet when it moves along the airflow direction.

6. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The lip cover (2) has an inner surface and an outer surface. A single annular compression channel is formed by the inner surface extending from the front edge to the rear and together with the outer conical surface of the central body (1). The inner surface and the outer surface meet at the front edge to form the lip cover (2) front edge line. The specific shape of the front edge line is a pointed front edge or a blunt front edge.

7. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The area change rate, length, outlet cross-sectional shape and area expansion ratio of each expansion segment (3) are exactly the same, only the spatial azimuth angle is different.

8. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The central body (1) has an air vent or air vent near the tail and throat (4).

9. The axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The Mach number of the throat (4) is between 1.3 and 2.

0.

10. An axisymmetric intake manifold for a modular air-breathing engine according to claim 1, characterized in that, The axisymmetric air intake is installed on the nose, belly, sides or back of the aircraft.

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