Lifting return reusable rocket substage and rocket

Abstract

The application provides a lift-type return reusable rocket substage and rocket, wherein the substage comprises a rocket body assembly, a rib wing assembly, an oblique grid fin assembly and a landing leg fairing assembly; the rib wing assembly comprises two rib wings; the rib wings are arranged on the outer side of the rocket body of the substage and satisfy G1 = the distance from the center of pressure of the rocket body of the substage to the front end face of the rocket body of the substage, L1 = 2.25D, W1 = 0.225D, 0.1D ≤ C1 ≤ 0.125D, λ1 = λ2 = 45°; wherein G1 is the distance from the geometric center of the rib wing to the front end face of the rocket body of the substage; L1 is the length of the rib wing; W1 is the width of the rib wing; C1 is the root chord of the rib wing; λ1 is the front edge sweepback angle of the rib wing; λ2 is the rear edge sweepforward angle of the rib wing; D is the diameter of the rocket body of the substage; the oblique grid fin assembly comprises grid fins; the connecting lines of two groups of oppositely arranged grid fins are not perpendicular; the landing leg fairing assembly comprises a landing leg fairing. The substage can effectively improve the return lateral maneuvering capability of the rocket substage.

Description

Technical Field

[0001] This invention relates to the field of reusable launch vehicle technology, specifically to a lift-type reusable rocket stage and the rocket itself. Background Technology

[0002] Reusable launch vehicle technology was first realized and commercialized on a large scale by SpaceX's Falcon 9 rocket. The vertical reverse thrust recovery technology used by the Falcon 9 has become the mainstream reference solution in this field. This solution relies on engine reverse thrust ignition and aerodynamic drag for deceleration during the first stage's reentry into the atmosphere, and uses deployable grid fins for attitude control, ultimately achieving a vertical landing via landing legs.

[0003] Despite its historic success, the technology's physics, based on ballistic / high-drag return, inherently presents three interconnected technical bottlenecks that are difficult to completely resolve through optimization: 1. The aerodynamic layout does not support lift-based return, severely limiting maneuverability; 2. The return payload environment is extremely harsh, resulting in high reuse costs; 3. Limited longitudinal control capabilities impair flight efficiency.

[0004] In conclusion, the vertical thrust reverse recovery technology adopted by the Falcon 9, especially its aerodynamic layout structure for reentry, has also caused losses in other rocket performance aspects, particularly in terms of maneuverability, reentry environment, and aerodynamic efficiency, which limits its potential to achieve lower cost, greater flexibility, and higher reliability.

[0005] To improve the lateral maneuverability of the rocket's first stage during reentry, designing a lift-type reusable rocket first stage is particularly important. Summary of the Invention

[0006] The purpose of this invention is to overcome the shortcomings of the prior art and provide a lift-type reusable rocket first stage.

[0007] This invention provides a first stage of a lift-type reusable rocket, comprising: a rocket body assembly, a strake wing assembly, an obliquely mounted grid fin assembly, and a landing leg fairing assembly; the rocket body assembly includes a first stage rocket body, which is cylindrical; the strake wing assembly includes two strake wings symmetrically arranged on the outer surface of the first stage rocket body; each strake wing is arranged along the axis of the first stage rocket body and satisfies: G1 = distance from the center of pressure of the first stage rocket body to the front end face of the first stage rocket body, L1 = 2.25D, W1 = 0.225D, 0.1D ≤ C1 ≤ 0.125D, λ1 = λ2 = 45° Wherein, L1 is the length of the strake wing; W1 is the width of the strake wing; C1 is the root chord length of the strake wing; λ1 is the leading edge sweep angle of the strake wing; λ2 is the trailing edge sweep angle of the strake wing; D is the diameter of the first stage rocket body; the oblique grid fin assembly includes four grid fins arranged near the front end face of the first stage rocket body and along the circumferential direction of the first stage rocket body, wherein the line connecting two sets of oppositely arranged grid fins is not perpendicular; the landing leg fairing assembly includes four landing leg fairings arranged near the bottom of the first stage rocket body and along the circumferential direction of the first stage rocket body.

[0008] According to one embodiment of the present invention, the arrow body of the first stage arrow body satisfies: L=10D; where L is the length of the arrow body.

[0009] According to one embodiment of the present invention, the angle θ1 = 60° between the central axes of the two grid fins that are circumferentially close together along the circumference of the first-stage rocket body.

[0010] According to one embodiment of the present invention, the two leading edge fins are respectively disposed on the vertical line of the line connecting the two grid fins that are circumferentially close to each other along the circumferential line of the first stage rocket body.

[0011] According to one embodiment of the present invention, the grid rudder satisfies: G2=1.1D, L2=0.36D, W2=0.3D, C2=0.04D, E2=0.96D, F2=0.05D; wherein, G2 is the distance from the rudder axis of the grid rudder to the front end face of the first stage rocket body; L2 is the length of the grid rudder; W2 is the width of the grid rudder; C2 is the chord length of the grid rudder; E2 is the distance from the short side of the grid rudder to the axis of the first stage rocket body when the grid rudder is deployed at 90°; and F2 is the width of the grid of the grid rudder.

[0012] According to one embodiment of the present invention, the landing leg fairing satisfies: C3≤0.1D, L3 / C3≥10; wherein, C3 is the vertical height between the highest point of the outer surface of the landing leg fairing and the circumferential outer surface of the first stage rocket body; L3 is the length along the axis of the first stage rocket body when the landing leg fairing is retracted into the state of the first stage rocket body.

[0013] According to one embodiment of the present invention, L3 = 1.3D.

[0014] According to one embodiment of the present invention, when the landing leg fairing is retracted into the first stage rocket body, the outer surface of the landing leg fairing includes a left plane, a right plane, and a front plane; the outer surface of the landing leg fairing is formed by splicing the left plane, the right plane, and the front plane together; the left plane and the right plane away from the front plane form the trailing edge of the landing leg fairing, and the trailing edge of the landing leg fairing is used to connect with the first stage rocket body.

[0015] According to one embodiment of the present invention, four landing leg fairings are uniformly arranged circumferentially along the first stage rocket body; when the landing leg fairings are retracted into the first stage rocket body, the edge of the front plane away from the bottom of the first stage rocket body is the leading edge of the landing leg fairing, and the landing leg fairings satisfy: G3=8.2D, C3=0.1D, φ1=15°, φ2=80°, φ4=145°, φ5=75°, S1=0.025S, S2=S3=0.35S; G3 is the distance from the front end of the landing leg fairing to the bottom edge of the first stage rocket body. The distance between the front end faces of the first stage rocket body; φ1 is the angle between the two ends of the leading edge of the landing leg fairing along the circumference of the rocket body; φ2 is the angle between the two ends of the trailing edge of the landing leg fairing along the circumference of the first stage rocket body; φ4 is the angle between the left plane and the right plane; φ5 is the angle between the front plane of the landing leg fairing and the axis of the first stage rocket body; S1 is the area of ​​the front plane of the landing leg fairing; S2 and S3 are the areas of the left plane and the right plane of the landing leg fairing; S is the cross-sectional area of ​​the first stage rocket body.

[0016] On the other hand, the present invention provides a lift-type reusable rocket, including a stage as described above.

[0017] The lifting-type reusable rocket first stage according to the present invention solves the problem of poor lateral maneuverability of the rocket first stage during reentry by configuring large-sized leading-edge extensions.

[0018] It should be understood that the above general description and the following specific embodiments are merely exemplary and illustrative, and do not limit the scope of the invention. Attached Figure Description

[0019] The accompanying drawings, which are part of the specification of this invention, illustrate exemplary embodiments of the invention. The drawings, together with the description in the specification, serve to illustrate the principles of the invention.

[0020] Figure 1This is a perspective view of the first stage of a lift-type reusable rocket according to an embodiment of the present invention; Figure 2 This is a front view of a first stage of a lift-type reusable rocket according to an embodiment of the present invention; Figure 3 This is a side view of a first stage of a lift-return reusable rocket according to an embodiment of the present invention; Figure 4 This is a schematic diagram of a grid fin according to an embodiment of the present invention; Figure 5 This is a bottom view of a sub-stage when the grid rudder is deployed according to an embodiment of the present invention; Figure 6 This is a bottom view of a first stage of a lift-type reusable rocket according to an embodiment of the present invention; Figure 7 This is a schematic diagram of the landing leg fairing according to an embodiment of the present invention; Figure 8 This is a flowchart illustrating the determination of the geometric center of a leading edge wing according to an embodiment of the present invention; Figure 9 This is a three-dimensional digital model of a rocket first stage according to an embodiment of the present invention; Figure 10 This is a three-dimensional digital model bottom view of a rocket first stage according to an embodiment of the present invention.

[0021] Explanation of reference numerals in the attached figures: 100-Rocket body assembly; 200-Leading slat fin assembly; 300-Slanted grid fin assembly; 400-Landing leg fairing assembly; 101 First stage rocket body; 201 Leading slat fin; 301 Grid fin; 401 Landing leg fairing; 402 Left plane; 403 Right plane; 404 Forward plane. Detailed Implementation

[0022] The features and exemplary embodiments of various aspects of the present invention will now be described in detail. To make the objectives, technical solutions, and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only configured to explain the present invention and to exemplify the principles of the present invention, and are not configured to limit the present invention. In addition, the structural components in the drawings are not necessarily drawn to scale. For example, the dimensions of some structural components or regions in the drawings may be enlarged for other structural components or regions to aid in the understanding of the embodiments of the present invention.

[0023] The directional terms used in the following description refer to the directions shown in the figures and are not intended to limit the specific structure of the embodiments of the present invention. In the description of the present invention, it should be noted that, unless otherwise stated, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand the specific meaning of the above terms in the present invention according to the specific circumstances.

[0024] Furthermore, the terms "comprising," "including," "having," or any other variations thereof are intended to cover non-exclusive inclusion, such that a structure or component that includes a list of elements includes not only those elements but also other structural elements that are not expressly listed or inherent to the structure or component. Without further limitations, an element defined by the phrase "comprising..." does not exclude the presence of other identical elements in the article or apparatus that includes the element.

[0025] Spatial relation terms such as "below," "under," "under," "low," "above," "on," and "high" are used for descriptive convenience to explain the positioning of one element relative to a second element, indicating that these terms are intended to cover different orientations of the device, in addition to those different from those shown in the figure. Furthermore, phrases such as "one element on / below another element" can indicate that two elements are in direct contact, or that there are other elements between the two elements. In addition, terms such as "first" and "second" are also used to describe individual elements, areas, parts, etc., without specifically indicating order or sequence, and should not be considered restrictive. Similar terms are used throughout the description to represent similar elements.

[0026] In the following description of the present invention, the terms "rocket," "launch vehicle," "spacecraft," "space launch vehicle," or "missile" may be used in certain scenarios for ease of description only, and their connotations are not limited to the specific terms used. Generally, the rockets of the present invention include space launch vehicles used to launch satellites, spacecraft, or other probes, as well as various missiles, rockets, and other weapons used to carry payloads, and similar products capable of sending payloads into the air. Those skilled in the art, when interpreting the above specific terms, should not limit the rocket or launch vehicle to only one type of space launch vehicle or missile based on the specific terms used in the description, thereby narrowing the scope of protection of the present invention.

[0027] It will be apparent to those skilled in the art that the present invention can be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention.

[0028] Figure 1 This is a perspective view of the first stage of a lift-type reusable rocket according to an embodiment of the present invention; Figure 2 This is a front view of a first stage of a lift-type reusable rocket according to an embodiment of the present invention; Figure 3 This is a side view of a first stage of a lift-return reusable rocket according to an embodiment of the present invention; Figure 4 This is a schematic diagram of a grid fin according to an embodiment of the present invention; Figure 5 This is a bottom view of a sub-stage when the grid rudder is deployed according to an embodiment of the present invention; Figure 6 This is a bottom view of a first stage of a lift-type reusable rocket according to an embodiment of the present invention; Figure 7 This is a schematic diagram of the landing leg fairing according to an embodiment of the present invention; Figure 8 This is a flowchart illustrating the determination of the geometric center of a leading edge wing according to an embodiment of the present invention; Figure 9 This is a three-dimensional digital model of a rocket first stage according to an embodiment of the present invention; Figure 10 This is a three-dimensional digital model bottom view of a rocket first stage according to an embodiment of the present invention.

[0029] like Figures 1-3 As shown, this invention provides a first stage of a lift-type reusable rocket, comprising: a rocket body assembly 100, a strake fin assembly 200, an obliquely mounted grid fin assembly 300, and a landing leg fairing assembly 400. The rocket body assembly 100 includes a first-stage rocket body 101, which is cylindrical. The strake fin assembly 200 includes two strakes 201 symmetrically arranged on the outer surface of the first-stage rocket body 101. Each strake fin 201 is arranged along the axis of the first-stage rocket body 101 and satisfies: G1 = the distance from the center of pressure of the first-stage rocket body 101 to the front end face of the first-stage rocket body 101 (i.e., the distance from the geometric center of the strake fin 201 to the front end face of the first-stage rocket body 101 - e.g., ...). Figure 1 (The distance in the S4 direction shown), L1=2.25D, W1=0.225D, 0.1D≤C1≤0.125D, λ1=λ2=45°. Where, G1, L1 is the length of the strake wing 201. W1 is the width of the strake wing 201. C1 is the root chord length of the strake wing 201. λ1 is the leading edge sweep angle of the strake wing 201. λ2 is the trailing edge sweep angle of the strake wing 201. D is the diameter of the first stage rocket body 101.

[0030] The angled grid fin assembly 300 includes four grid fins 301 arranged near the front end face of the first stage rocket body 101 and along the circumferential direction of the first stage rocket body 101, wherein the line connecting two sets of opposing grid fins 301 is not perpendicular. The landing leg fairing assembly 400 includes four landing leg fairings 401 arranged near the bottom of the first stage rocket body 101 and along the circumferential direction of the first stage rocket body 101.

[0031] Specifically, the vertical reverse thrust recovery technology used by the Falcon 9 has three technical bottlenecks: 1. The aerodynamic layout does not support lift-based return, severely limiting maneuverability. Its return trajectory is essentially a controlled ballistic trajectory, meaning the rocket's lateral maneuverability is extremely limited. According to publicly available optimization analyses of the return trajectory of this type of rocket, to avoid excessive loss of payload capacity, its effective lateral range is typically limited to 100-200 kilometers. This limitation necessitates precise pre-planning of the landing site (whether on a sea platform or a land-based site) before launch, imposing stringent requirements on launch windows, weather conditions, and the deployment accuracy of sea-based recovery vessels, significantly reducing mission flexibility and the ability to respond to unforeseen circumstances.

[0032] 2. The reentry environment is extremely harsh, and reusability is costly. The steep trajectory of ballistic reentry causes the first stage to undergo severe aerodynamic deceleration in the dense atmosphere, generating extremely high aerodynamic overloads and heat fluxes. Peak aerodynamic overloads during ballistic reentry typically exceed 5G, reaching as high as 8G in some mission profiles. The bottom of the ballistic reentry rocket directly impacts the high-speed incoming flow, forming a complex bow-shaped shock wave. This causes the engine nozzle and bottom thermal protection system to experience intense aerodynamic heating, with peak heat fluxes easily exceeding 500 kW / m², and even higher in some areas. Such harsh force and thermal load environments not only require additional reinforcement of the rocket structure, increasing dry weight, but also necessitate the use of large-area, high-performance heat shield materials. This makes post-recovery inspection, maintenance, and component replacement extremely complex and expensive, directly limiting further reductions in reusability costs.

[0033] 3. Limited longitudinal control capability, resulting in compromised flight efficiency. The Falcon 9's cross-shaped grid rudder layout means that only two control surfaces are involved in longitudinal control. If flying with an angle of attack, the longitudinal trim requirement will lead to a loss of 15%-20% in longitudinal control efficiency, directly affecting the accuracy and stability of attitude control.

[0034] The longitudinal static stability margin of a rocket during its return flight is crucial for trimming the large angle-of-attack pitching moment of a lift-type return phase. It is a prerequisite for achieving controllable and stable gliding flight and the basis for a low-G (e.g., below 3G) return.

[0035] The first stage provided in this embodiment innovatively proposes that the geometric center of the strake 201 coincides with the pressure center (i.e., pressure core) of the maximum dynamic pressure point of the rocket's reentry phase through joint optimization of the strake 201's profile, installation position, and area. This solves the problem of trimming and longitudinal static stability of the first stage during high angle-of-attack gliding, ensuring a moderate longitudinal static stability margin during reentry (which is crucial for trimming the high angle-of-attack pitching moment of the lift-off reentry phase and is the basis for achieving low-G reentry, such as below 3G). This improves aerodynamic control efficiency and ensures the stability of the first stage in challenging flight attitudes. For example, the two strakes 201 can be symmetrically arranged relative to the axis of the first stage body 101, that is, the two strakes 201 are symmetrically arranged on both sides of the first stage body at an angle of 180 degrees apart in the circumferential direction.

[0036] The first stage, equipped with fixed large-size leading-edge extensions 201, effectively improves the lift-to-drag ratio of the rocket's first-stage return flight, ensuring sufficient lift during the reentry phase. This enables the rocket to possess a wide range of gliding maneuvers, achieving a mode transition from "ballistic return" to "lift-based return," and effectively reducing propellant consumption. Verification shows that the reentry phase achieves a Mach number of 0.9-3, a lift-to-drag ratio of 0.2-0.4, and a designed lateral range exceeding 500 kilometers, significantly enhancing mission flexibility and giving the rocket's first stage a high-performance lift-based return capability. Furthermore, the inclined grid fin assembly 300 of the first stage adopts an "X"-shaped inclined grid fin 301 installation layout. Compared with the traditional "+" shaped 90° layout (i.e., the line connecting two sets of opposing grid fins 301 is vertically arranged), it can more efficiently decompose the normal force generated by fin deflection into the pitch channel, improve the longitudinal trim capability of the rocket during angle-of-attack flight, ensure precise control of the pitch, roll, and yaw attitude after the rocket's lift reentry into the atmosphere, improve the control accuracy of the aerodynamic control phase, and ensure that the first stage can return and land with high precision. For example, C1=0.125D.

[0037] The first stage adopts a cylindrical first stage body 101 structure, and integrates aerodynamic surfaces such as leading-edge slat wing assembly 200, oblique grid rudder assembly 300 and landing leg fairing assembly 400 for optimized design, which effectively improves the first stage's lateral maneuverability during reentry, breaks through the lateral stroke limit, and reduces peak aerodynamic overload and peak aerodynamic heat flux, meeting the requirements of high lift, high maneuverability and low aerodynamic interference during reentry.

[0038] Verification has shown that this substage has increased the lateral range of the existing ballistic return trajectory from less than 200 kilometers to the 500-kilometer level; reduced the peak aerodynamic overload from the existing 5-8G to the 3G level; and reduced the peak aerodynamic heat flux from the existing level of more than 500 kW / m² by more than 40%.

[0039] In addition, the root chord length of the edge wing 201 satisfies 0.1D≤C1≤0.125D, which can further reduce the structural weight while ensuring that the edge wing area and the edge wing installation position remain unchanged.

[0040] like Figure 4 As shown, according to an embodiment of the present invention, the leading edge and trailing edge of the grid rudder 301 (i.e., the edges of the first stage rocket body that are close to and far away from the grid rudder 301 when it is unfolded at 90°) are generally arc-shaped, and its radius of curvature is basically the same as the outer radius of the first stage rocket body 101, thereby achieving a zero-gap or minimal-gap fit between the grid rudder 301 and the first stage rocket body 101.

[0041] like Figure 2 As shown, according to one embodiment of the present invention, the arrow body 101 of a first-stage arrow body satisfies: L = 10D. Wherein, L is the length of the arrow body.

[0042] like Figure 5 As shown, according to one embodiment of the present invention, the angle θ1 = 60° between the central axes of the two grid fins 301 that are circumferentially closer together and the circumference of the first stage rocket body 101, and correspondingly, the angle θ1 = 120° between the central axes of the two grid fins 301 that are circumferentially farther apart and the circumference of the first stage rocket body 101.

[0043] In this embodiment, with Figure 5 S5 is the reference longitudinal symmetry plane. The angle θ1 = 60° between the central axes of two adjacent grid rudders 301 on the same side of the reference longitudinal symmetry plane and the circumference of the first-stage rocket body 101. The circumferential angle θ2 = 60° between the grid rudders 301 and the reference longitudinal symmetry plane of the first-stage rocket body 101. This first stage, through a 60° / 60° oblique "X"-shaped grid rudder layout, achieves efficient distribution of the normal torque generated by rudder deflection in the pitch and yaw control channels, especially enhancing pitch control capability. Verification shows that the Mach number during the return phase is 0.9-3.20. Furthermore, during the return phase, the rudder deflection angle can achieve trim within 15 degrees of the flight angle of attack, improving pitch control capability by 20% compared to the "+"-shaped grid rudder, and solving the problem of 15%-20% loss in longitudinal control efficiency when the "+"-shaped grid rudder is in flight at an angle of attack.

[0044] The 60° / 60° angled "X" grid fin layout of the first stage complements the stability provided by the leading-edge extensions, improving aerodynamic efficiency, enhancing the efficiency and capability of the pitch control channel, optimizing attitude control efficiency, and ensuring precise attitude control of the first stage of the high-lift rocket.

[0045] like Figure 5 As shown, according to one embodiment of the present invention, two leading edge winglets 201 are respectively disposed on the vertical line of the line connecting the two grid rudders 301 that are circumferentially close together along the circumferential line of the first stage rocket body 101.

[0046] like Figure 2 , Figure 3 and Figure 5 As shown, according to an embodiment of the present invention, the grid rudder 301 satisfies: G2=1.1D, L2=0.36D, W2=0.3D, C2=0.04D, E2=0.96D, F2=0.05D. Wherein, G2 is the distance from the rudder axis of the grid rudder 301 to the front end face of the first-stage rocket body 101 (i.e., the grid rudder is designed to be close to the front end face of the first-stage rocket body). L2 is the length of the grid rudder 301. W2 is the width of the grid rudder 301. C2 is the chord length of the grid rudder 301. E2 is the distance from the short side of the grid rudder 301 (i.e., the edge away from the first-stage rocket body 101 when the grid rudder is deployed 90°) to the axis of the first-stage rocket body 101. F2 is the width of the grid of the grid rudder 301.

[0047] In this embodiment, for example, such as Figure 5 As shown, the angle θ3 between the grid of the grid rudder 301 and the long side of its outer frame can be 45° (e.g., Figure 4 and Figure 5 (as shown) or 90° (i.e., the rudder grid can be installed perpendicular to the outer frame of the grid rudder, not shown), to accommodate different manufacturing processes.

[0048] like Figure 2 and Figure 6 As shown, according to one embodiment of the present invention, the landing leg fairing 401 satisfies: C3 ≤ 0.1D, L3 / C3 ≥ 10. Wherein, C3 is the vertical height between the highest point of the outer surface of the landing leg fairing 401 and the circumferential outer surface of the first-stage rocket body 101. L3 is the length of the landing leg fairing 401 along the axis of the first-stage rocket body 101 when it is retracted into its retracted state.

[0049] Specifically, the landing leg fairings protect the landing leg retraction mechanism from airflow impact. The Falcon 9 rocket has a diameter of 3.8 meters, and the landing leg fairings are relatively large compared to the rocket's diameter, generating additional aerodynamic drag and resulting in approximately 1.3% of payload capacity loss during the ascent phase. These landing leg fairings also cause significant shock waves and turbulent wakes to form on the first-stage rocket during reentry. These disruptive flow fields directly affect the grid fins 301 located behind them, further reducing the effectiveness of the grid fins 301.

[0050] In this embodiment, by limiting the height and length-to-height ratio of the landing leg fairing (ensuring C3 ≤ 0.1D and L3 / C3 ≥ 10), it achieves an ultra-low profile and a large length-to-height ratio, reducing aerodynamic drag during ascent and thus minimizing motion loss. Furthermore, the landing leg fairing of this first-stage rocket reduces aerodynamic interference and additional drag across the entire flight profile to the lowest engineering-allowed level, ensuring the weakest local shock wave intensity and smallest wake turbulence range in its retracted configuration, resulting in low drag. This minimizes interference with the circumferential flow field of the first-stage rocket body, reduces wake interference during reentry, and avoids contamination of the leading-edge extension wing assembly 200 and grid fin assembly 300 by incoming flow, thereby reducing the impact on grid fin control efficiency, ensuring the efficiency of each aerodynamic surface, and enabling all components of the aerodynamic layout to work efficiently and collaboratively within their designed flow field environment, effectively reducing the loss of carrying capacity. For example, C3 = 0.1D ensures sufficient internal space for the landing leg fairing.

[0051] Furthermore, L3 = 1.3D.

[0052] like Figure 7 As shown, according to one embodiment of the present invention, when the landing leg fairing 401 is retracted into the first stage rocket body 101, the outer surface of the landing leg fairing 401 includes a left plane 402, a right plane 403, and a front plane 404. The outer surface of the landing leg fairing 401 is formed by splicing the left plane 402, the right plane 403, and the front plane 404 together. The left plane 402 and the right plane 403 are located away from the front plane 404, forming the trailing edge of the landing leg fairing 401, which is used for connection with the first stage rocket body 101.

[0053] The landing leg fairing 401 of the first stage provided in this embodiment adopts a three-plane splicing scheme, which effectively reduces the difficulty of molding, processing and production and final assembly compared with overall curved surface forming, and improves production efficiency.

[0054] like Figure 2 and Figure 6As shown, according to one embodiment of the present invention, four landing leg fairings 401 are uniformly arranged circumferentially along the first stage rocket body 101. When the landing leg fairings 401 are retracted into the first stage rocket body 101, the edge of the front plane 404 away from the bottom of the first stage rocket body 101 is the leading edge of the landing leg fairing. The landing leg fairings 401 satisfy: G3=8.2D, C3=0.1D, φ1=15°, φ2=80°, φ4=145°, φ5=75°, S1=0.025S, S2=S3=0.35S. G3 is the distance from the front end of the landing leg fairing 401 to the front end face of the first stage rocket body 101. φ1 is the angle between the two ends of the leading edge of the landing leg fairing 401 along the circumferential direction of the rocket body. φ2 is the angle between the two ends of the trailing edge of the landing leg fairing 401 along the circumference of the first-stage rocket body 101. φ4 is the angle between the left plane 402 and the right plane 403. φ5 is the angle between the front plane 404 of the landing leg fairing 401 and the axis of the first-stage rocket body 101. S1 is the area of ​​the front plane 404 of the landing leg fairing 401. S2 and S3 are the areas of the left plane 402 and the right plane 403 of the landing leg fairing 401. S is the cross-sectional area of ​​the first-stage rocket body 101, S = .

[0055] In this embodiment, with Figure 5 S5 in the figure is the reference longitudinal symmetry plane. The angle φ3 = 45° between the centerline of the landing leg fairing (i.e., the line connecting the left plane 402 and the right plane 403) and the longitudinal symmetry plane of the rocket body along the circumference of the first-stage rocket body 101. For example, while ensuring that the area of ​​each planar plate of the landing leg fairing 401 (the area of ​​each plane on the outer side of the landing leg fairing 401 when it is closed to the first-stage rocket body) remains unchanged, multiple small planar plates can be processed first, and then the multiple small planar plates can be spliced ​​together to form each planar plate, which can further reduce the processing and production difficulty. For example, the left plane can be divided into n planar plates, and after splicing, it should satisfy S21 + S22 + ... S2n = S2.

[0056] like Figure 8 As shown, according to one embodiment of the present invention, the process of determining the geometric center of the strake wing by pressing the center is as follows: S1: Initial value for the centrifugal force; S2: Move the edge wing so that its geometric center coincides with the pressure center; S3: Calculate the actual pressure center; S4: Determine whether the geometric center coincides with the actual center of pressure; S5: If they do not overlap, repeat S2-S4.

[0057] This process determines the geometric center of the strake wing through multiple iterations. Taking a typical return trajectory with a maximum dynamic pressure of 100 kPa (altitude 15 km, Mach number 3.4, angle of attack 5°) as an example: First, given an initial pressure center value of 7D from the leading edge of the first stage, the strake wing is moved so that G1 = 7D. Then, the pressure center is calculated. If the geometric center does not coincide with the pressure center, the strake wing is moved until the geometric center is located at the pressure center. The pressure center is calculated again, and the relationship between the geometric center and the pressure center is determined. Then, the strake wing is moved until the geometric center coincides with the pressure center. The designed G1 can be obtained after 3-4 iterations.

[0058] On the other hand, the present invention provides a lift-type reusable rocket, including the aforementioned first stage.

[0059] like Figure 9 and Figure 10 As shown, according to one embodiment of the present invention, an example of designing a rocket first stage is as follows: Arrow body assembly 100: The arrow body diameter D = 4.5m is selected, and the arrow body length L = 10D = 45m is obtained through calculation. The cross-sectional area of ​​the arrow body S = 15.904㎡; Fixed wing assembly 200: Through the above iteration, G1=33.75m is obtained, and through calculation, L1=2.25D=10.125m, W1=0.225D=1.0125m, C1=0.125D=0.5625m, λ1=λ2=45° are obtained; 300 slanted grid rudder assembly: G2=1.1D=4.95m, L2=0.36D=1.62m, W2=0.3D=1.35m, C2=0.04D=0.18m, E2=0.96D=4.32m, F2=0.05D=0.225m, θ1=60°, θ2=60°, θ3=45°.

[0060] Landing leg fairing assembly 400: G3=8.2D=36.9m, L3=1.3D=5.85m, C3=0.1D=0.45m, φ1=15°, φ2=80°, φ3=45°, φ4=145°, φ5=75°, S1=0.3976㎡, S2=S3=5.566㎡.

[0061] The above embodiments of the present invention can be combined with each other and have corresponding technical effects.

[0062] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A first stage of a lift-type reusable rocket, characterized in that, include: Rocket body assembly, strake wing assembly, angled grid fin assembly, and landing leg fairing assembly; The arrow body assembly includes a sub-stage arrow body, which is cylindrical; The wing assembly includes two winglets symmetrically arranged on the outer side of the first-stage rocket body; each winglet is arranged along the axis of the first-stage rocket body and satisfies: G1 = distance from the center of gravity of the first-stage rocket body to the front end face of the first-stage rocket body, L1 = 2.25D, W1 = 0.225D, 0.1D ≤ C1 ≤ 0.125D, λ1 = λ2 = 45°; where L1 is the length of the winglet; W1 is the width of the winglet; C1 is the root chord length of the winglet; λ1 is the leading edge sweep angle of the winglet; λ2 is the trailing edge sweep angle of the winglet; and D is the diameter of the first-stage rocket body. The inclined grid fin assembly includes four grid fins arranged near the front end face of the first stage rocket body and along the circumferential direction of the first stage rocket body, wherein the line connecting two sets of grid fins arranged opposite each other is not perpendicular. The landing leg fairing assembly includes four landing leg fairings located near the bottom of the first stage rocket body and arranged along the circumferential direction of the first stage rocket body.

2. The sub-level according to claim 1, characterized in that, The first-stage arrow body satisfies: L=10D; where L is the length of the arrow body.

3. The sub-level according to claim 1, characterized in that, The angle θ1 = 60° between the central axes of the two grid fins that are circumferentially close together along the circumference of the first-stage rocket body.

4. The sub-stage according to claim 3, characterized in that, The two leading edge fins are respectively positioned on the vertical line of the line connecting the two grid fins that are circumferentially close to each other along the circumferential line of the first-stage rocket body.

5. The sub-level according to claim 1, characterized in that, The grid rudder satisfies: G2=1.1D, L2=0.36D, W2=0.3D, C2=0.04D, E2=0.96D, F2=0.05D; where G2 is the distance from the rudder axis of the grid rudder to the front end face of the first stage rocket body; L2 is the length of the grid rudder; W2 is the width of the grid rudder; C2 is the chord length of the grid rudder; E2 is the distance from the short side of the grid rudder to the axis of the first stage rocket body when the grid rudder is deployed at 90°; and F2 is the width of the grid of the grid rudder.

6. The sub-level according to claim 1, characterized in that, The landing leg fairing satisfies: C3≤0.1D, L3 / C3≥10; where C3 is the vertical height between the highest point of the outer surface of the landing leg fairing and the circumferential outer surface of the first stage rocket body; L3 is the length of the landing leg fairing along the axis of the first stage rocket body when it is retracted into the state of the first stage rocket body.

7. The sub-stage according to claim 6, characterized in that, L3=1.3D.

8. The sub-stage according to claim 6, characterized in that, When the landing leg fairing is retracted into the first stage rocket body, the outer surface of the landing leg fairing includes a left plane, a right plane, and a front plane; the outer surface of the landing leg fairing is formed by splicing the left plane, the right plane, and the front plane together; the left plane and the right plane away from the front plane form the trailing edge of the landing leg fairing, and the trailing edge of the landing leg fairing is used to connect with the first stage rocket body.

9. The sub-stage according to claim 8, characterized in that, The four landing leg fairings are evenly arranged around the circumference of the first stage rocket body; when the landing leg fairings are retracted into the first stage rocket body, the edge of the front plane away from the bottom of the first stage rocket body is the leading edge of the landing leg fairing, and the landing leg fairings satisfy: G3=8.2D, C3=0.1D, φ1=15°, φ2=80°, φ4=145°, φ5=75°, S1=0.025S, S2=S3=0.35S; G3 is the distance from the front end of the landing leg fairing to the front end face of the first stage rocket body; φ1 is the angle between the two ends of the leading edge of the landing leg fairing along the circumference of the first stage rocket body; φ2 is the angle between the two ends of the trailing edge of the landing leg fairing along the circumference of the first stage rocket body; φ4 is the angle between the left plane and the right plane; φ5 is the angle between the front plane of the landing leg fairing and the axis of the first stage rocket body; S1 is the area of ​​the front plane of the landing leg fairing; S2 and S3 are the areas of the left plane and the right plane of the landing leg fairing; S is the cross-sectional area of ​​the first stage rocket body.

10. A lift-type reusable rocket, characterized in that, Includes a sub-level as described in any one of claims 1 to 9.

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