Local cavity design method for internal turning inlet cowl lip shock point thermal protection

Abstract

The application relates to a local cavity design method for internal turning inlet lip point heat protection of a near space hypersonic aircraft. Given an internal contraction reference flow field and a captured inlet shape, an internal turning inlet lip profile is determined according to the intersection line of the reference flow field conical shock wave surface and the captured inlet shape; given an internal turning inlet lip bluntness section, a three-dimensional swept surface of the lip is obtained by sweeping along the internal turning inlet lip profile; the size of a local cavity of the internal turning inlet lip point is specified according to the internal turning inlet lip profile and the lip bluntness section; the swept curve of the internal turning inlet lip point area is reconstructed according to the size of the local cavity, and a local cavity front edge profile of the internal turning inlet lip point is generated; the three-dimensional surface of the internal turning inlet lip point area is reconstructed according to the local cavity front edge profile of the internal turning inlet lip point, and a three-dimensional bluntness lip of the internal turning inlet with the local cavity of the internal turning inlet lip point is generated.

Description

TECHNICAL FIELD

[0001] The present application relates to near space hypersonic vehicles, in particular to a local cavity design method for the cowl lip shock point thermal protection of an internal turning inlet. BACKGROUND

[0002] Since the birth of aircraft, people have been constantly pursuing faster, farther and higher flight. In particular, after the concept of scramjet engine (William H Heiser, David T. Pratt, Daniel H. Daley, et al. Mehta. Hypersonic airbreathing propulsion [M]. AIAA Education Series. Washington, DC: AIAA, 1994) was proposed, air-breathing hypersonic vehicles powered by scramjet engine were favored by all countries for their great strategic significance, and became the frontiers of aerospace field. Air-breathing hypersonic vehicles do not need to carry oxidizer, have long endurance and high maneuverability, and many other potential. In order to realize the breakthrough from engineering concept to practical application, the United States and Europe have successively formulated NASP, Hyper-X, HyTech, HIFiRE, HAWC and LAPCAT programs, and carried out a large number of mechanism research, ground and flight experiments. Among them, X-43A verification machine, X-51A verification machine and HIFiRE all realize the short time flight task above Mach 5 with power, and verify the feasibility of scramjet engine technology, and accelerate the practical process of hypersonic vehicles. However, there are still many factors restricting the long time and long endurance flight of air-breathing hypersonic vehicles. Among them, the complex shock wave interference problem of aircraft aerodynamic components is one of the key factors, which can change the flow characteristics of the internal and external flow of the aircraft and cause the jump of local aerodynamic force / thermal load, seriously affect the flight safety and service life of the aircraft, and become a challenging topic in aircraft design and scientific research.

[0003] To fully expose the shock interference problem of each part of the aircraft, highlight the key points and decompose the difficulties, researchers have targeted to simplify a series of classic aerodynamic configurations (for example, oblique shock incident cylinder configuration, swept cylinder configuration, double wedge configuration and compression corner configuration, etc.), and carried out extensive research in theory, experiment and numerical simulation. So far, the research on classic configurations has been relatively rich and perfect. However, with the refinement of hypersonic vehicle design, new geometric forms of shock interference problems are emerging. In recent years, with the development of scramjet engine, three-dimensional internal turning inlet gradually becomes a research hotspot (Bisek N J. High-fidelity simulations of the Hifire-6 flow path [C]. AIAA paper. 2016, 2016-1115). Compared with two-dimensional oblique shock wave system, three-dimensional shock wave system converges three-dimensional airflow inward, so it has higher compression efficiency and less energy loss, which can bring more net thrust to the engine, so the three-dimensional internal turning inlet with three-dimensional curved shock wave gradually becomes the development direction of the new generation of inlet. Among them, the Falcon HCV scheme published by DARPA (Walker S, Sherk J, Shell D, et al. The DARPA / AF Falcon program: The hypersonic technology vehicle #2 (HTV-2) flight demonstration phase [C], AIAA International Space Planes & Hypersonic Systems & Technologies Conference. 2008), Boeing's Valkyrie scheme (Guy N. Boeing unveils hypersonic airline concept [R], Aviation Week, 2018) all use three-dimensional internal turning inlet, which shows the importance of three-dimensional internal turning inlet in the design of new generation of hypersonic vehicles.

[0004] However, three-dimensional internal turning inlets bring new shock wave interference problems while bringing the advantages of high compression efficiency and strong flow capture performance. Unlike the blunt leading edge structure at the lip position of the traditional two-dimensional inlet (Li Z F. Mechanism of hypersonic inlet starting characteristics[D]. University of Science and Technology of China, 2013), the internal turning inlet lip shape is generally determined by the intersection of the capture inlet shape and the reference flow field class inner cone shock wave surface, usually showing a V-shaped structure. Considering the problem of aerodynamic thermal protection, the V-shaped leading edge will be blunted, at which time the shock wave will not adhere to the leading edge wall, but will be detached. The detached shock wave will intersect at the stagnation point area of the leading edge and cause complex shock wave interference. The shock wave interference of the V-shaped leading edge will exist throughout the supersonic / hypersonic flight phase, continuously generating high heat / power load on the aircraft wall, and may affect the air quality entering the aircraft internal flow passage. In fact, such V-shaped leading edge structure is widely used in internal turning inlet lips, two-dimensional inlet side panel spillage ports, and aircraft wing and engine compartment connections.

[0005] For the typical V-shaped configuration represented by the inward turning inlet cowl lip (spill lip), the Chinese University of Science and Technology, National University of Defense Technology, and others have conducted in-depth research on shock wave interference type classification, aerodynamic heating / power characteristics, and unsteady oscillation (Zhang Z Y. V-shaped blunt lip shock wave interference and aerodynamic heating / power characteristics research[D]. Chinese University of Science and Technology, 2020. DOI:10.27517 / d.cnki.gzkju.2020.000130. Gao W Z, Li Z F, Cao R, et al. Numerical simulation and analysis of the influence of V-shaped lip on the shock wave incident boundary layer flow[J]. Journal of Propulsion Technology, 2019, 40(11): 2488-2497. DOI:10.13675 / j.cnki.tjjs.180761). To facilitate parameterization research and highlight key points while ignoring secondary factors, Xiao F, et al. (Xiao F, Li Z, Zhang Z, et al. Hypersonic shock wave interactions on a V-shaped blunt leading edge[J]. AIAA Journal, 2018, 56(1): 356-367) first extracted a simplified V-shaped blunt leading edge configuration. The main parameters of the model are the leading edge expansion angle β, the leading edge blunt radius r, and the rounded radius R at the root intersection position. Through experiments and numerical simulations, Xiao F found that the special geometric contraction effect of the V-shaped blunt leading edge makes its shock wave interference phenomenon extremely complex, presenting three types of shock wave interference: regular reflection (RR), Mach reflection (MR), and same-side regular reflection (SRR). The complex shock wave interference structure will produce shock waves, shear layers, and supersonic airflow impacting the wall, causing the wall heat flux to rise rapidly. Unexpectedly, simply increasing the leading edge blunt radius does not necessarily reduce the heat load, and on the contrary, it may result in higher heat load. In particular, under high Mach number conditions, the wall heat flux peak has even exceeded the classic Edney's type IV shock wave interference. The research of Xiao F, et al. pointed out that when the shock wave interference structure transitions from MR to sRR, the wall peak heat flux will decrease significantly. Zhang Z Y, et al. (Zhang Z Y. V-shaped blunt lip shock wave interference and aerodynamic heating / power characteristics research[D]. Chinese University of Science and Technology, 2020. DOI:10.27517 / d.cnki.gzkju.2020.000130) systematically studied the wave structure characteristics of the flow field and the transition process of the shock wave interference type under Ma6 conditions, and established the transition criteria between different interference structures.Zhang et al. (Zhang T, Cheng J R, Shi C G, Zhu C X, You Y C. Mach reflection of three-dimensional curved shock waves on V-shape blunt leading edges. J. Fluid Mech. 2023, 975, A45) developed a theoretical model for MR structure, accurately predicted the wave structure of Mach reflection, and derived the transition boundary of MR structure and sRR structure based on the theoretical method.

[0006] The purpose of studying the flow mechanism is to develop an effective heat reduction scheme. In recent years, Beijing University of Aeronautics and Astronautics has applied some classic flow control methods to the V-shaped blunt leading edge of the internal turning inlet to reduce heat flow. Liu Sijia et al. developed a scheme to reduce the heat flow at the lip using counter-flow jets, and found that this method can reduce the local wall heating peak by more than 45%. In order to guide the optimization of this scheme, they also quantitatively analyzed the uncertainty of the incoming flow and jet parameters. At the same time, Kang Dakai et al. (Kang D K, Yan C, Liu S J, Wang Z W, Jiang Z H. Modelling and shock control for a V-shaped blunt leading edge. J. Fluid Mech. 2023, 968, A15) introduced a shock control bump to solve the problem of local heat flow peak caused by shock / boundary layer interaction at the lip. At high incoming Mach number, the shock control bump effectively suppresses boundary layer separation, reducing the local heat flow peak on the outside by 66%. However, the introduction of the shock control bump will cause the heat flow level at the stagnation point to rise, especially at lower incoming Mach numbers. In order to reduce the heat flow at the stagnation point caused by the impact of the lateral supersonic jet, Li Shuai et al. (Li S, Yan C, Kang D K, Liu S J, Jiang Z H. Investigation of flow control methods for reducing heat flux on a V-shaped blunt leading edge under real gas effects. Phys. Fluids. 2023, 35(3), 036113) introduced a stagnation point boss, which changed the head-on collision of two supersonic jets into an oblique collision, significantly reducing the impact intensity and reducing the heat flow peak in the symmetric plane. However, the introduction of the stagnation point boss often leads to the appearance of heat flow peaks in the asymmetric plane, increasing the uncertainty of heat flow generation. Therefore, although some progress has been made in the research of heat reduction schemes suitable for the V-shaped lip of the internal turning inlet, it is still necessary to explore more effective heat reduction schemes. SUMMARY

[0007] The present application aims to solve the problems of heat protection of the inner turning inlet lip and control of the flow structure of the separation zone under high-speed flight conditions, and provides a local cavity design method for heat protection of the inner turning inlet lip.

[0008] The present application comprises the following steps:

[0009] 1) According to the design requirements, the inner contraction reference flow field and the capture inlet shape are given, and the inner turning inlet lip profile is determined according to the intersection line of the reference flow field class inner cone shock wave surface and the capture inlet shape.

[0010] 2) The inner turning inlet lip bluntness section is given, and the lip three-dimensional swept surface is obtained by sweeping along the inner turning inlet lip profile, and the inner turning inlet lip bluntness section adopts a circular design.

[0011] 3) The size of the lip stagnation point local cavity is specified according to the inner turning inlet lip profile and the lip bluntness radius, the stagnation point local cavity size is reconstructed according to the swept curve of the inner turning inlet lip stagnation point region, and the stagnation point local cavity leading edge profile is generated, and the stagnation point local cavity leading edge profile adopts a spline curve design.

[0012] 4) The stagnation point region three-dimensional surface is reconstructed according to the stagnation point local cavity leading edge profile, and the inner turning inlet three-dimensional bluntness lip with the stagnation point local cavity is generated, and the three-dimensional surface of the stagnation point local cavity is designed by curve grid.

[0013] In step 1), the capture inlet shape is projected along the incoming flow direction to the reference flow field class inner cone shock wave surface, and the intersection line generated is the inner turning inlet leading edge profile, wherein a part of the leading edge profile close to the flow field center is the inner turning inlet lip profile.

[0014] In step 1), the inner turning inlet lip profile is V-shaped.

[0015] In step 2), the bluntness radius of the bluntness section of the lip is denoted as r; the lip three-dimensional swept surface is symmetrical about the x-y plane and the x-z plane, and the intersection point of the inner turning inlet lip profile and the x-y symmetrical plane is the lip stagnation point.

[0016] In step 3), the specific steps of generating the stagnation point local cavity leading edge profile include:

[0017] (1) According to the curvature radius of the inner turning inlet lip contour line at the lip stagnation point, the curvature center of the lip contour line at the lip stagnation point is determined, and the circumferential angle of the stagnation point local cavity is pointed to the curvature center; the circumferential angle of the stagnation point local cavity is 5°-20°;

[0018] (2) The maximum depth of the stagnation point local cavity is specified; the maximum depth h of the stagnation point local cavity is 0.05-0.1 times of the lip bluntness radius r of the blunted section;

[0019] (3) According to the circumferential angle of the stagnation point local cavity and the maximum depth of the stagnation point local cavity, the front edge curve of the inner turning inlet lip stagnation point region is reconstructed, and the front edge contour line of the stagnation point local cavity is generated.

[0020] The method has the following advantages: the method realizes the local cavity design for the inner turning inlet lip stagnation point heat protection. By setting the cavity at the blunted lip stagnation point of the inner turning inlet, the flow structure of the separation zone generated by the collision of the opposite supersonic jet at the stagnation point is controlled, the internal airflow of the separation zone is locally expanded, the strength of the backflow airflow impacting the wall surface is weakened, the peak pressure and heat flow of the stagnation point region are comprehensively reduced, the heat flow peak value on the asymmetric surface is avoided, and the safety and stability of the three-dimensional inner turning inlet are improved. BRIEF DESCRIPTION OF DRAWINGS

[0021] Figure 1 The figure is a schematic diagram of the inner turning inlet lip contour line generation method;

[0022] Figure 2 The figure is a schematic diagram of the inner turning inlet lip three-dimensional blunted profile;

[0023] Figure 3 The figure is a schematic diagram of the inner turning inlet lip flow field flow structure;

[0024] Figure 4 The figure is a schematic diagram of the inner turning inlet lip stagnation point local cavity front edge contour line design method;

[0025] Figure 5 The figure is a schematic diagram of the inner turning inlet three-dimensional blunted lip with a stagnation point local cavity;

[0026] Figure 6 The figure is a blunted lip wall surface heat flow distribution diagram of the inner turning inlet with a stagnation point local cavity. In the figure, the horizontal coordinate represents the inverse circular circumferential angle; the vertical coordinate represents the dimensionless wall surface heat flow. Figure 6

[0027] Figure 7 The figure is a blunted lip wall surface heat flow distribution diagram of the inner turning inlet with a stagnation point local cavity. In the figure, the horizontal coordinate represents the inverse circular circumferential angle; the vertical coordinate represents the dimensionless wall surface heat flow. Figure 7 ​​

[0028] The marks in the figure are as follows: 1 represents an internal contraction reference flow field, 2 represents a capture inlet shape, 3 represents an internal turning inlet compression profile, 4 represents an internal cone shock wave surface of the reference flow field, 5 represents an internal turning inlet leading edge profile, 6 represents an internal turning inlet lip profile, 7 represents a lip bluntness section, 8 represents a lip three-dimensional swept surface, 9 represents a lip stagnation point, 10 represents a rear-swept body shock wave, 11 represents a curved shock wave formed by continuous compression of supersonic airflow after the rear-swept body shock wave, 12 represents a bow shock wave formed by convergence of the rear-swept body shock wave and the curved shock wave, 13 represents a shear layer, 14 represents a supersonic jet wrapped by the shear layer 13, 15 represents a flow separation zone generated by collision of the supersonic jet 14, 16 represents a curvature center of the lip profile at the lip stagnation point 9, 17 represents a stagnation point local concave cavity leading edge profile, and 18 represents a stagnation point local concave cavity three-dimensional surface. DETAILED DESCRIPTION

[0029] In order to make the technical problems to be solved by the present application, technical solutions and beneficial effects more clear and explicit, the present application is further described in detail below in combination with the drawings and examples.

[0030] First, the basic design mechanism is introduced for the design method. The design principle of the internal turning inlet is shown in Figure 1 . In the design, a certain internal contraction reference flow field 1 with better performance is first selected, and then flow line tracking is performed in the reference flow field according to a given capture inlet shape 2. The envelope surface formed by the flow lines is the internal turning inlet compression profile 3. Therefore, the internal turning inlet leading edge profile 5 is generally determined by the intersection line of the capture inlet shape 2 and the internal cone shock wave surface 4 of the reference flow field. A part of the leading edge profile close to the center of the reference flow field is referred to as the internal turning inlet lip profile 6, which usually presents a V-shaped structure. Considering the aerodynamic thermal protection problem, the V-shaped leading edge is blunted, as shown in Figure 2 . After blunting, the shock wave will not adhere to the lip wall surface, but will be detached before the stagnation point to form a detached shock wave. The detached shock wave will cause shock wave interference before the lip stagnation point 9.

[0031] Under specific incoming flow and geometric conditions, a same-side conventional reflection structure will be generated on the lip, as shown in Figure 3The swept trailing wave 10 generated at the swept leading edge intersects with the curved wave 11 formed by the continuous compression of the supersonic flow behind the swept trailing wave by the curved wall, and converges into the bow wave 12 formed by the convergence of the swept trailing wave and the curved wave, and emits the shear layer 13. The supersonic jet 14 wrapped by the shear layer 13 is formed between the shear layer 13 and the three-dimensional swept surface 8 of the lip of the inward turning inlet. The supersonic jet 14 wrapped by the shear layer 13 converges from the upper and lower sides of the flow field to the lip stagnation point 9, and collides, so that the flow separation zone 15 generated by the collision of the supersonic jet 14 appears on both sides of the stagnation point. A part of the flow close to the wall on both sides of the stagnation point is deviated from the wall under the influence of the compression wave, and the other part of the flow flows around the recirculation zone and then impacts the wall, and reattaches at the stagnation point, thereby generating a peak heat flux at the stagnation point. At this time, the flow structure at the stagnation point (the collision of the supersonic jet on the opposite sides) is significantly different from the stagnation point characteristics of the classical blunt body, and the peak heat flux generated by the convergence of the supersonic jet is much higher than the stagnation point heat flux q o As shown in Figure 6 By setting a concave cavity near the stagnation point, the flow structure in the separation zone can be effectively controlled, the intensity of the airflow impacting the wall can be reduced, and the peak heat flux at the stagnation point can be reduced.

[0032] As shown in Figures 1 to 6 The local concave cavity design method for the lip stagnation point of the inward turning inlet for thermal protection comprises the following steps:

[0033] 1) As shown in Figure 1 According to the design requirements, the inward turning inlet reference flow field 1 and the capture inlet shape 2 are given, the capture inlet shape 2 is projected onto the reference flow field inner cone shock wave surface 4 along the incoming flow direction, and the intersection line is the inward turning inlet leading edge profile 5. A part of the leading edge profile close to the center of the flow field is the inward turning inlet lip profile 6, which is usually V-shaped.

[0034] 2) The inward turning inlet lip bluntness section 7 is given, and the lip three-dimensional swept surface 8 is obtained by sweeping along the inward turning inlet lip profile 6, as shown in Figure 2 The inward turning inlet lip bluntness section 7 adopts a circular design, and the lip bluntness radius is denoted as r. The lip three-dimensional swept surface 8 is symmetrical about the x-y plane and the x-z plane, and the intersection point of the inward turning inlet lip profile 6 and the x-y symmetrical plane is denoted as the lip stagnation point 9.

[0035] 3) The size of the local concave cavity at the lip stagnation point is specified according to the inward turning inlet lip profile 6 and the lip bluntness radius r. As shown in Figure 4As shown, first, according to the curvature radius R of the inner turning inlet lip contour line 6 at the lip stagnation point 9, the curvature center 16 of the lip contour line at the lip stagnation point 9 is determined, and the circumferential angle Φ of the stagnation point local cavity is 5-20° with the curvature center 16 of the lip contour line at the lip stagnation point 9 as the center. Then, the maximum depth h of the stagnation point local cavity is specified, and the maximum depth h of the stagnation point local cavity is usually 0.05-0.1 times the lip bluntness radius r of the inner turning inlet lip bluntness section 7. Finally, according to the circumferential angle Φ of the stagnation point local cavity and the maximum depth h of the stagnation point local cavity, the inner turning inlet lip stagnation point area leading edge curve is reconstructed, and the leading edge contour line 17 of the stagnation point local cavity is generated. The leading edge contour line 17 of the stagnation point local cavity is designed by using a spline curve, so as to ensure smooth transition at the intersection with the inner turning inlet lip contour line 6.

[0036] 4) According to the stagnation point local cavity leading edge contour line 17, the stagnation point area bluntness profile is reconstructed, and the three-dimensional profile 18 of the stagnation point local cavity is generated. The three-dimensional swept profile of the lip and the three-dimensional profile 18 of the stagnation point local cavity together constitute the three-dimensional bluntness lip of the inner turning inlet with the stagnation point local cavity, as shown. Figure 5 The three-dimensional profile 18 of the stagnation point local cavity is designed by using a curve grid, so as to ensure smooth transition at the intersection line with the three-dimensional swept profile of the inner turning inlet lip 8.

[0037] The present application changes the flow field structure of the stagnation point area of the inner turning inlet lip by designing a local cavity in the stagnation point area, reduces the peak value of the stagnation point heat flow, and thus improves the heat protection performance of the inlet.

[0038] Figure 6 The numerical simulation results of the symmetric plane inner wall surface heat flow distribution of the inner turning inlet bluntness lip with the stagnation point local cavity (VBLE with stagnation cavity) and the original lip (Original VBLE) are shown. It can be seen that, compared with the original model, the central heat flow peak value of the inner turning inlet bluntness lip with the stagnation point local cavity is obviously reduced, and the heat flow reduction amplitude can reach 24%. It is shown that the local cavity design can effectively alleviate the impact of high-temperature gas flow on the wall, so as to reduce the heat load. Figure 7 The wall surface heat flow distribution of the stagnation point local cavity in different z planes is shown. As shown, Figure 7In the formula, z represents the distance perpendicular to the flow field direction, and r represents the lip bluntness radius. Different z / r values represent different positions, z=0 represents the initial position in the flow field direction, z / r=1 / 4 represents the position at a distance of one fourth of the lip bluntness radius, z / r=1 / 2 represents the position at a distance of one half of the lip bluntness radius, and z / r=3 / 4 represents the position at a distance of three fourths of the lip bluntness radius. The heat flow distribution at these positions can be used to evaluate the heat protection effect of the local cavity at different depths. The results show that the wall heat flow at the cavity decreases as the z value increases, which indicates that no heat flow peak value outside the symmetry plane is generated in the local cavity. Therefore, the design of the local cavity can comprehensively reduce the heat load at the lip stagnation point of the internal turning inlet duct and improve the safety and stability of the three-dimensional internal turning inlet duct.

[0039] Experiments show that, by introducing the local cavity design at the stagnation point, the heat load at the lip stagnation point of the internal turning inlet duct can be comprehensively reduced, and the safety and stability of the inlet duct can be improved. This has important significance for the design and application of high-speed aircraft.

[0040] The above embodiments are only preferred embodiments of the present application and cannot be considered as limiting the scope of the present application. Any equivalent changes and improvements made according to the scope of the present application should still belong to the scope of the present application.

Claims

1. A local cavity design method for thermal protection of the inward-turned inlet lip stagnation point, characterized by The following steps are involved: 1) Given the inward-converging reference flow field and capture inlet shape according to design requirements, the inward-turning inlet lip profile is determined based on the intersection of the reference flow field's inner cone shock wave surface and the capture inlet shape; 2) Given a blunt cross-section of the inward-turning inlet lip, sweep along the inward-turning inlet lip profile to obtain a three-dimensional swept profile of the lip. The blunt cross-section of the inward-turning inlet lip is designed to be circular. 3) The size of the local cavity at the stagnation point of the lip is specified based on the inward-turning inlet lip profile and the lip blunting radius. The swept curve of the stagnation point area of ​​the inward-turning inlet lip is reconstructed based on the cavity size to generate the leading edge profile of the local cavity at the stagnation point. The leading edge profile of the local cavity at the stagnation point is designed using a spline curve. 4) The three-dimensional surface of the stagnation area is reconstructed according to the leading edge profile of the local concave cavity at the stagnation point, and a three-dimensional blunted lip of the inward-turning air inlet with the local concave cavity at the stagnation point is generated. The three-dimensional surface of the local concave cavity at the stagnation point is designed through a curved grid.

2. The local cavity design method for thermal protection of the inward-turned air inlet lip according to claim 1 is characterized in that In step 1), the capture inlet shape is projected onto the inner cone shock wave surface of the reference flow field along the incoming flow direction, and the resulting intersection line is the leading edge profile of the inward-turning air inlet, wherein a portion of the leading edge profile close to the center of the flow field is the lip profile of the inward-turning air inlet.

3. The local cavity design method for thermal protection of the inward-turned air inlet lip according to claim 1 is characterized in that In step 1), the inward-turned air inlet lip profile is V-shaped.

4. The local cavity design method for thermal protection of the inward-turned air intake lip stagnation point according to claim 1 is characterized in that In step 2), the three-dimensional swept surface of the lip is symmetrical about the xy plane and the xz plane, and the intersection of the inward-turned air inlet lip profile line and the xy symmetry plane is the lip stationary point.

5. The local cavity design method for thermal protection of the inward-turned air inlet lip according to claim 1 is characterized in that In step 3), the specific steps of generating the leading edge profile of the stagnation point local cavity include: first, according to the curvature radius of the lip profile of the inward-turning air inlet lip at the lip stagnation point, determining the center of curvature of the lip profile at the lip stagnation point, and taking the center of curvature as the center of the circle, indicating the circumferential angle of the stagnation point local cavity; then, specifying the maximum depth of the stagnation point local cavity; finally, reconstructing the leading edge curve of the stagnation point area of ​​the inward-turning air inlet lip according to the circumferential angle of the stagnation point local cavity and the maximum depth of the stagnation point local cavity, to generate the leading edge profile of the stagnation point local cavity.

6. The local cavity design method for thermal protection of the inward-turned air inlet lip according to claim 5 is characterized in that The circumferential angle Φ of the local concave cavity at the stationary point is 5° to 20°.

7. The local cavity design method for thermal protection of the inward-turned air inlet lip according to claim 5 is characterized in that The maximum depth of the local concave cavity at the stagnation point is 0.05 to 0.1 times the radius of the lip passivation section.

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