Integrated internal and external flow design method for two-stage compression TBCC inlet duct

Abstract

This paper presents an integrated design method for the internal and external flow of a two-stage compression TBCC inlet. The method integrates the internal waverider TBCC inlet, capable of continuous operation across a wide speed range, with a waverider forebody possessing a high lift-to-drag ratio, achieving an integrated internal / external flow configuration for continuous operation across a broad speed range. The two-stage compression internal waverider TBCC inlet employs streamline-following design based on an internal contraction fundamental flow field with dual incident shock waves. The second-stage compression profile of the inlet is designed as an adjustable profile to control the transition between low-speed turbine mode and high-speed ramjet mode. The coupling design of the fundamental flow field of the internal waverider TBCC inlet with the flow field of the external waverider forebody enables aerodynamic transition between the internal waverider flow within the internal waverider TBCC inlet and the external waverider flow outside the waverider forebody. This integrated configuration ensures continuous operation of the aircraft across a wide speed range and provides good lift-to-drag characteristics at hypersonic speeds, broadening the operating speed range of the waverider and enabling horizontal takeoff and landing and reusability of hypersonic aircraft.

Description

Technical Field

[0001] This invention pertains to the design of hypersonic vehicles, specifically to the integration of internal and external flow in a wide-speed-range adjustable internal wave propulsion (TBCC) combined propulsion system, and particularly to an integrated internal and external flow design method based on a two-stage compression TBCC inlet. Background Technology

[0002] Hypersonic vehicles have become a fiercely contested military domain for nations worldwide. Wide-speed-range, high-speed, horizontally operable, and reusable hypersonic vehicles have become a crucial research area for future aircraft due to their strategic importance, economic viability, and feasibility. For wide-speed-range flight, combined propulsion systems with multiple power systems have emerged and become a research hotspot. These include turbine-based combined cycle (TBCC) and rocket-based combined cycle (RBCC) propulsion systems. Their advantage lies in combining turbine / rocket engines suitable for low-speed operation with subsonic / scramjet engines suitable for high-speed operation, fully utilizing the operational characteristics of each engine within its advantageous speed range to achieve a propulsion system capable of continuous operation across a wide speed range.

[0003] Due to its key characteristics of horizontal takeoff and landing and reusability, the research on air-breathing TBCC (Through-Turn-Off-Cylinder) combined propulsion systems has attracted considerable attention. As a critical component of air-breathing propulsion systems, the inlet's ability to operate continuously across a wide speed range determines the success of the combined propulsion system. Traditional fixed-geometry inlets no longer meet the requirements for wide-speed-range operation; inlets need continuous adjustment during flight to adapt to changes in flight conditions and ensure the engine's airflow requirements are met. In recent years, three-dimensional internal waverider TBCC inlets have become an important research direction for future aircraft inlet systems due to their small wetted area, high compression efficiency, and high flow capture coefficient. Furthermore, waveriders, due to their high lift-to-drag ratio and strong maneuverability, have become the preferred choice for aerodynamic layout design of hypersonic aircraft, and research on the integration of internal turning inlets / waveriders is underway, achieving integrated design through geometric or aerodynamic transitions. However, current research in this area is limited to a relatively small flight speed range. The air intake system in the integrated configuration mostly adopts a fixed geometry structure, but the flight envelope of the aircraft is usually above Ma4+, so it is relatively difficult to achieve horizontal take-off and landing and reusability. Summary of the Invention

[0004] The purpose of this invention is to develop a new generation of wide-speed-range hypersonic vehicles capable of horizontal takeoff and landing and reusability, providing an integrated internal and external flow design method based on a two-stage compression TBCC inlet. By using an internal waverider TBCC propulsion system, the operating speed range of the vehicle is broadened, enabling horizontal takeoff and landing and reusability. The lift-to-drag ratio and maneuverability of the vehicle are improved by using an external waverider forebody, thus developing an integrated configuration design method for internal / external waverider aerodynamic coupling. This invention employs a three-dimensional internal waverider TBCC inlet with strong compression characteristics coupled with a waverider body with a high lift-to-drag ratio, proposing an integrated internal and external flow design method based on a two-stage compression TBCC inlet. This provides technical and solution support for the development of hypersonic vehicles capable of horizontal takeoff and landing, reusability, continuous operation across a wide speed range, and hypersonic cruise.

[0005] The aforementioned integrated internal and external flow design method for a two-stage compression TBCC inlet includes:

[0006] 1) Basic flow field design for two-stage compression and internal contraction;

[0007] 2) External compression basic flow field design;

[0008] 3) Inner / outer wave-riding aerodynamic transition design;

[0009] 4) Two-stage compression internal wave-riding TBCC inlet compression section design;

[0010] 5) Adjustable flow splitting structure design for the intake compression section;

[0011] 6) Adjustable low-speed turbine channel design;

[0012] 7) High-speed stamping channel design;

[0013] 8) External wave compression profile design.

[0014] In step 1), the design of the basic flow field for the two-stage compression internal contraction is as follows: the basic flow field for internal contraction is a necessary condition for the design of the internal wave-riding inlet, and the cruise state of the aircraft is usually used as the design point of the basic flow field; this design method is for Ma7-class wide-speed-range aircraft, so Ma7 is used as the design point of the basic flow field, and the basic flow field for the two-stage compression internal contraction is designed and solved according to the requirements of the inlet compression performance under cruise state.

[0015] In step 2), the design of the basic external compression flow field is as follows: based on the flight requirements of the Ma7 class hypersonic vehicle in this scheme, a three-dimensional curved shock wave surface for wave riding on the lower wall of the hypersonic vehicle is designed. Based on the designed shock wave surface, the basic external compression flow field is solved by combining the local deflection shearing method.

[0016] In step 3), the internal / external wave-riding aerodynamic transition design is as follows: the design intersects the first incident shock wave surface of the internal contraction basic flow field with the shock wave surface of the external compression basic flow field, and the resulting intersection line is the geometric leading edge profile shared by the internal / external wave-riding. The internal / external flow completes the aerodynamic transition at this profile. This leading edge profile is designed as part of the geometric profile of the internal wave-riding inlet and the external wave-riding compression wall, respectively, and is integrated into the structural design.

[0017] In step 4), the design of the two-stage compression inner wave TBCC inlet compression section is as follows: based on the existing partial leading edge profile, the capture shape of the inlet is designed according to the requirements of the aircraft. Then, the inlet compression section is generated in the two-stage compression inner wave basic flow field using streamline tracking technology, and the aerodynamic profile is reconstructed through boundary layer correction technology.

[0018] In step 5), the adjustable flow splitting structure of the intake duct compression section is designed as follows: the intake duct compression section is a shared section for the low-speed turbine passage and the high-speed ramjet passage. To generate the low-speed turbine passage, the second-stage compression profile of the compression section is partially modified and designed as a rotatable flow splitter. The shaft is positioned at the beginning of the second-stage compression profile, and the rotation angle is the second-stage compression angle. Rotating the flow splitter upwards to align it parallel to the first-stage compression profile generates the inlet of the low-speed turbine passage. The opening and closing of the low-speed turbine passage is controlled by this structure, which also allows for adjustment and control of the airflow distribution between the two passages.

[0019] In step 6), the adjustable low-speed turbine channel design is as follows: the low-speed turbine channel needs to be closed under high-speed conditions to ensure the structural safety of the turbine, therefore an adjustable structure is required for control. Thus, the channel consists of an adjustable section and a non-adjustable section. The adjustable section is equipped with a rotatable adjustment plate, which works in conjunction with the adjustable flow divider plate of the compression section to ensure a smooth transition of the aerodynamic profile during adjustment, while also ensuring the degree of channel closure. The non-adjustable section is generated according to the structural requirements, controlling the centerline and area variation of the polynomial.

[0020] In step 7), the high-speed stamping channel is designed as follows: the high-speed stamping channel operates under supercombustion conditions and can be designed with a fixed geometric surface. The center line and area change law can be controlled according to structural requirements, or the channel aerodynamic surface can be generated by maintaining a constant cross-sectional area.

[0021] In step 8), the outer waverider compression profile design involves: based on the existing common leading-edge profile, extending the leading-edge profile outward according to the structural and geometric requirements of the hypersonic vehicle to form a complete waverider FCT profile. The complete leading-edge profile is discretized into a point set and streamlined in the external compression basic flow field to generate the outer waverider compression profile. This profile and the inner waverider inlet complete an aerodynamic transition at the common leading-edge profile, forming a wide-speed-range integrated internal / external flow aerodynamic configuration.

[0022] For the design of future wide-speed-range hypersonic vehicles, the technical solution of this invention has the following advantages:

[0023] Integrating the TBCC (Through-Tube Compression Control) intake system with the waverider forebody vehicle enables horizontal takeoff and landing and reusability of hypersonic vehicles, offering enhanced economic efficiency and scalability. The TBCC intake, based on a two-stage compression internal waverider, offers flexible adjustment, ensuring normal operation across a wide speed range and broadening the vehicle's flight envelope. The internal waverider TBCC intake boasts advantages such as a small frontal area, high compression performance, and controllable inlet and outlet shapes, facilitating integrated design with the external flow waverider forebody. The integrated internal / external flow structure provided by this invention exhibits excellent lift-drag and maneuverability characteristics, meeting the specific requirements of future next-generation hypersonic vehicles. Attached Figure Description

[0024] Figure 1 The Mach number distribution and density gradient contour map of the basic flow field during two-stage compression and contraction.

[0025] Figure 2 This is a schematic diagram of the basic flow field under external compression.

[0026] Figure 3 This is a schematic diagram of the aerodynamic transition between the inner and outer wave-riding phases.

[0027] Figure 4 This is a schematic diagram of an internal wave-driven TBCC intake based on two-stage compression.

[0028] Figure 5 A schematic diagram of the design for the outflow waverider front.

[0029] Figure 6 A bottom view of the integrated internal and external flow configuration of the two-stage compression TBCC inlet.

[0030] Figure 7 This is an isometric drawing of the integrated internal and external flow configuration of the two-stage compression TBCC inlet.

[0031] The markings in the diagram represent the following: 1 represents the first-stage compression wall, 2 represents the second-stage compression wall, 3 represents the first incident shock wave, 4 represents the second incident shock wave, 5 represents the reflected shock wave, 6 represents the shock wave generator, 7 represents the external compression shock wave surface, 8 represents the upper leading edge profile of the inlet, 9 represents the shared leading edge profile for the inner / outer wave-riding aerodynamic transition, 10 represents the first-stage compression profile of the inlet, 11 represents the flow splitting adjustment rotation axis, 12 represents the second-stage compression profile of the inlet, and 13 represents the second-stage compression profile in turbine mode. 14 indicates that the turbine channel adjustable section adjustment plate is in the closed position; 15 indicates that the turbine channel adjustable section adjustment plate is in the open position; 16 indicates the turbine channel adjustable section adjustment plate rotation shaft; 17 indicates the turbine channel non-adjustable section; 18 indicates the ramjet channel isolation section; 19 indicates the turbine channel outlet; 20 indicates the ramjet channel outlet; 21 indicates the inlet lip; 22 indicates the outer leading edge profile of the waverider forebody; 23 indicates the side compression wall of the waverider forebody; 24 indicates the outer lower wall of the waverider generated by the shared leading edge profile of the internal and external flows. Detailed Implementation

[0032] To make the technical problems, technical solutions and specific structures of this invention clearer and more understandable, the invention will be further described in detail below with reference to the accompanying drawings and embodiments.

[0033] The specific design method of this invention includes the following steps:

[0034] Design a two-stage compression-contraction basic flow field: Design as follows Figure 1 The two-stage compression internal contraction basic flow field is shown in this embodiment. Ma7 is used as the basic flow field design point. The first-stage compression wall 1 and the second-stage compression wall 2 are designed, resulting in a basic flow field including a first incident shock wave 3, a second incident shock wave 4, and a reflected shock wave 5. Both the first-stage compression wall 1 and the second-stage compression wall 2 are straight walls. The matching of the two-stage compression angles considers the design requirement of the inlet throat Mach number being between Ma4.0 and 4.5 under the design condition of Ma7, while ensuring that the inlet area generated after the second-stage compression wall 2 rotates meets the flow requirements of the turbine passage. Considering both the throat Mach number and the turbine passage flow requirements, the two-stage compression angles are determined to be 3° and 3.1°, respectively. The external compression basic flow field is designed as follows: Figure 2 As shown, since the design point of the internal contraction basic flow field in this embodiment is Ma7, the design point of the external flow field is also selected as Ma7. The external compression shock surface 7 is designed using a nine-node controlled Bezier surface. This shock surface is the shock surface of the external flow wave-riding surface. The external compression basic flow field after the external compression shock surface 7 is solved by the local deflection kissing method, and the shock generator 6 that generates the external compression shock surface 7 is obtained.

[0035] Inner / outer wave-riding aerodynamic transition design: such as Figure 3As shown, the first incident shock wave 3 of the basic flow field of the two-stage compression internal contraction is extracted and intersects with the outer compression shock wave surface 7. The resulting intersection line is the common leading edge profile 9 of the inner / outer wave-riding aerodynamic transition. The airflow completes the aerodynamic transition of the inner / outer flow at this profile. This leading edge profile is designed as an integrated structure part of the inner wave-riding inlet and the leading edge profile of the wave-riding forebody.

[0036] Design a two-stage compression inlet TBCC intake compression section: such as Figure 4 As shown, based on the shared leading edge profile 9 of the inner / outer wave-riding aerodynamic transition, the upper leading edge profile 8 of the inlet is designed according to the requirements of the aircraft to obtain the overall leading edge profile of the inlet. Then, the inlet compression section is generated in the two-stage compression inner wave-riding basic flow field using streamline tracking technology, and the aerodynamic profile is reconstructed through profile trimming. The first-stage compression profile 10 of the inlet, the flow splitting adjustment rotating shaft 11, and the second-stage compression profile 12 of the inlet constitute the upper wall of the compression section of the inlet. The inlet lip 21 is located at the intersection of the first incident shock wave 3 and the outer compression shock wave surface 7.

[0037] The intake duct compression section is designed with an adjustable flow splitting structure: the intake duct compression section is a shared section for the low-speed turbine passage and the high-speed ramjet passage. In order to generate the low-speed turbine passage, the second-stage compression profile 12 of the intake duct is designed as a rotatable flow splitter plate. The rotation angle is the compression angle of the second-stage compression wall 2, that is, the compression angle of the second-stage compression wall 2 is 3.1°. In turbine mode, the second-stage compression profile 13 is nearly parallel to the first-stage compression profile 10 of the intake duct, thereby generating the inlet of the low-speed turbine passage. The second-stage compression profile 12 of the intake duct rotates around the flow splitting adjustment rotation axis 11 to control the opening and closing of the turbine passage.

[0038] Adjustable low-speed turbine channel: The low-speed turbine channel needs to be closed under high-speed conditions to ensure the structural safety of the turbine, therefore an adjustable structure is required for control. When the adjustable section adjustment plate of the turbine channel is in the closed position 14, the turbine channel is closed. When the adjustable section adjustment plate of the turbine channel is in the open position 15, the turbine channel is open. The adjustment plate moves around the rotation axis 16 of the adjustable section adjustment plate of the turbine channel. The design of the non-adjustable section 17 of the turbine channel is generated by using a fourth-order polynomial to control the center line and area change law, given the turbine channel outlet 19.

[0039] High-speed stamping channel: The stamping channel isolation section 18 operates under supercombustion conditions and has a fixed geometric shape. The stamping channel outlet 20 is given according to requirements and is generated by the control center line and area change law using a fourth-order polynomial.

[0040] External wave compression profile: such as Figure 5As shown, based on the existing shared leading edge profile 9 of the inner / outer waverider aerodynamic transition, according to the structural and geometric requirements of the hypersonic vehicle, the outer leading edge profile 22 of the waverider forebody is designed outward on the outer compression shock wave surface 7 to form the FCT profile of the complete waverider surface. The complete leading edge profile is discretized into a point set and streamlined in the outer compression basic flow field to obtain the side compression wall 23 of the waverider forebody and the outer lower wall 24 generated by the shared leading edge profile of the inner and outer flows.

[0041] like Figure 6 The diagram shows a bottom view of an integrated internal and external flow configuration based on a two-stage compression TBCC inlet. It can be seen that the outer waverider lower wall 24, generated by the shared leading edge profile of the internal and external flows, completes the aerodynamic transition with the internal waverider inlet at the shared leading edge profile 9. The side compression wall 23 of the waverider forebody and the outer waverider lower wall 24, generated by the shared leading edge profile of the internal and external flows, form the overall aerodynamic profile of the external flow waverider. (As shown...) Figure 7 The diagram shown is an isometric view of the integrated internal and external flow configuration of the two-stage compression TBCC inlet. The leading edge profile 8 on the upper side of the inlet and the shared leading edge profile 9 of the inner / outer wave-riding aerodynamic transition together form the leading edge inlet of the inlet, which forms an integrated internal / outer flow configuration with the external flow profile. In the turbine mode, the second-stage compression profile 13 rotates around the split adjustment rotation axis 11, which can control the opening and closing of the turbine passage inlet. In addition, the turbine passage adjustable section adjustment plate, which is in the open position 15, can also rotate downward around the turbine passage adjustable section adjustment plate rotation axis 16 in the ram mode to close the turbine passage.

[0042] This invention provides an integrated internal / external flow design method based on a two-stage compression TBCC inlet. It employs an internal waverider TBCC inlet with a small frontal area and a waverider forebody with a high lift-to-drag ratio for aerodynamic integration. The second-stage compression profile of the two-stage compression TBCC inlet can be continuously adjusted according to operating conditions to ensure the inlet's compression and starting performance across the entire speed range. This variable geometry structure broadens the inlet's operating speed range, enabling continuous operation within a wide Mach 0–7 range, horizontal takeoff and landing, and reusability, offering strong economic advantages for future aircraft development. Furthermore, the waverider effect of the external flow aerodynamic profile allows the aircraft to possess a high lift-to-drag ratio and strong maneuverability. This invention's integrated internal / external flow design method expands the design concept of wide-speed-range variable geometry internal / external flow integration, breaking through the design limitations of previous fixed geometry integration methods that could not achieve wide-speed-range, horizontal takeoff and landing. It has significant reference value for the design of future wide-speed-range hypersonic aircraft.

[0043] This invention addresses the wide-speed-range flight requirements, including horizontal takeoff and landing, reusability, and hypersonic cruise. It integrates a continuous-operational-range internal waverider TBCC inlet with a high lift-to-drag ratio waverider forebody, achieving a unified internal / external flow configuration for continuous operation across the wide speed range. The two-stage compression internal waverider TBCC inlet employs a streamline-following design based on a double-incident shock wave-driven internal contraction flow field. The second-stage compression profile of the inlet is designed as adjustable to control the transition between low-speed turbine and high-speed ramjet modes. Furthermore, the basic flow field of the internal waverider TBCC inlet is coupled with the flow field of the external waverider forebody, enabling aerodynamic transition between the internal waverider flow and the external waverider flow. This integrated configuration ensures continuous operation of the aircraft across a wide speed range and provides good lift-to-drag characteristics at hypersonic speeds, broadening the operating speed range of the waverider and enabling horizontal takeoff and landing and reusability for hypersonic vehicles.

[0044] The above embodiments are merely preferred embodiments of the present invention and should not be considered as limiting the scope of the present invention. All equivalent variations and improvements made within the scope of the present invention should still fall within the patent coverage of the present invention.

Claims

1. An integrated internal and external flow design method for a two-stage compression TBCC inlet duct, characterized in that... Includes the following steps: 1) Basic flow field design for two-stage compression with internal contraction: Using Ma7 as the design point for the basic flow field, the basic flow field for internal contraction of two-stage compression is designed and solved according to the requirements of the inlet compression performance under cruise conditions. 2) Design of basic external compression flow field: Based on the flight requirements of Ma7 hypersonic vehicles, a three-dimensional curved shock wave surface for wave riding on the lower wall of the hypersonic vehicle is designed. Based on the designed shock wave surface, the basic external compression flow field is solved by combining the local deflection and shearing method. 3) Inner / outer wave-riding aerodynamic transition design: The first incident shock wave surface of the inner contraction basic flow field intersects with the shock wave surface of the outer compression basic flow field, and the resulting intersection line is the geometric leading edge profile shared by the inner and outer wave-riding. The inner and outer flows complete the aerodynamic transition at this profile. This leading edge profile is designed as part of the geometric profile of the inner wave-riding inlet and the outer wave-riding compression wall, respectively, and is integrated into the structural design. 4) Design of the compression section of the two-stage compression internal wave-riding TBCC inlet: Based on the existing partial leading edge profile, the capture shape of the inlet is designed according to the needs of the aircraft. The inlet compression section is generated in the basic flow field of the two-stage compression internal wave-riding using streamline tracking technology. The aerodynamic profile is reconstructed through boundary layer correction technology. 5) Adjustable flow splitter design for the intake duct compression section: The intake duct compression section is a shared section for the low-speed turbine passage and the high-speed ramjet passage. In order to generate the low-speed turbine passage, the second-stage compression profile of the compression section is locally modified and designed as a rotatable flow splitter. The shaft position is set at the starting position of the second-stage compression profile, and the rotation angle is the second-stage compression angle. The flow splitter is rotated upward to be parallel to the first-stage compression profile to generate the inlet of the low-speed turbine passage. 6) Adjustable low-speed turbine channel design: The channel consists of an adjustable section and a non-adjustable section. The adjustable section is equipped with a rotatable adjustment plate, which works in conjunction with the adjustable flow divider of the compression section to ensure a smooth transition of the aerodynamic profile during adjustment, while ensuring the degree of channel closure. The non-adjustable section is generated according to the centerline and area change law of the polynomial control according to structural requirements. 7) High-speed stamping channel design: The high-speed stamping channel is designed with a fixed geometric surface. The center line and area change law are controlled according to structural requirements, or the aerodynamic surface of the channel is generated by maintaining a constant cross-sectional area. 8) External waverider compression profile design: Based on the existing common leading edge profile, according to the structural and geometric requirements of the hypersonic vehicle, the leading edge profile is extended outward to form a complete waverider FCT profile. The complete leading edge profile is discretized into a point set and streamlined in the external compression basic flow field to generate the external waverider compression profile. The external waverider compression profile and the internal waverider inlet complete the aerodynamic transition at the common leading edge profile to form a wide-speed-domain integrated internal / external flow aerodynamic configuration.

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