CN112193401B - Thermal protection method for front edge of hypersonic aircraft - Google Patents
Thermal protection method for front edge of hypersonic aircraft Download PDFInfo
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- CN112193401B CN112193401B CN202010264045.5A CN202010264045A CN112193401B CN 112193401 B CN112193401 B CN 112193401B CN 202010264045 A CN202010264045 A CN 202010264045A CN 112193401 B CN112193401 B CN 112193401B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/38—Constructions adapted to reduce effects of aerodynamic or other external heating
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C30/00—Supersonic type aircraft
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Abstract
The invention relates to a method for protecting the leading edge of a hypersonic aerocraft from heat. Designing a pneumatic support rod, and fixing the pneumatic support rod at the foremost end of an aircraft; designing a dredging type thermal protection structure of the aircraft; obtaining key parameters influencing peak interference heat flow; optimizing the local shape of the aircraft leading edge area; performing shock tunnel heat measurement test verification on the obtained aircraft leading edge region; completing the design of the sparse heat protection structure, obtaining a cooling effect and a rule influencing the cooling effect parameters, and obtaining an optimal sparse heat protection structure; performing an electric arc wind tunnel thermal examination test on the obtained optimal dredging type thermal protection structure; and determining whether the design of the sparse heat protection structure is finished according to the result of the heat assessment test. The method can ensure the feasibility of the thermal protection scheme of the leading edge region without reducing the overall performance indexes such as lift-drag ratio of the aircraft and the like, and can effectively solve the problem of thermal protection of the leading edge region of the hypersonic aircraft.
Description
Technical Field
The invention belongs to the technical field of aircraft heat reduction and drag reduction, and particularly relates to a thermal protection method for a front edge of a hypersonic aircraft.
Background
For a hypersonic aircraft with a complex shape in a high Mach number near space, the radius of the front edge of an elastic body/protective cover cannot be large due to the high lift-drag ratio shape and the working requirement of an air-breathing engine, the front edge area bears severe pneumatic heating, and meanwhile, the high enthalpy dissociation incoming flow and a thermal protection material generate complex nonlinear coupling action, so that extremely strict requirements on the temperature resistance, the oxidation ablation resistance, the reliability and the like of a thermal protection structure in the front edge area are provided. The conventional thermal protection material is difficult to ensure that the leading edge area is not greatly ablated, the feasibility of a thermal protection scheme cannot be ensured by a means of the hard resistance of the thermal protection material alone, and the heat reduction optimization design of the thermal protection scheme of the leading edge area becomes a bottleneck problem of aircraft development.
For the hypersonic speed aircraft in the near space with the flight Mach number of more than 15 for a long time, the temperature of the front edge area of the hypersonic speed aircraft may exceed 3000 ℃, the front edge area is difficult to be ensured not to be greatly ablated only by means of the hard resistance of the thermal protection materials such as C/SiC, antioxidant C/C and ultrahigh temperature ceramics which are commonly used at present, and the adoption of the active thermal protection modes such as air film/sweating cooling and the like can bring extra air sources and control equipment, so that the technical route is still immature. Therefore, the feasibility of ensuring the thermal protection scheme of the leading edge region while not reducing the overall performance indexes such as lift-drag ratio of the aircraft has great technical difficulty, and a novel integrated heat and drag reduction technology needs to be developed in a targeted manner.
Disclosure of Invention
The invention aims to overcome the defects in the prior art and provides a method for thermally protecting the leading edge of a hypersonic aerocraft. The scheme of the invention can solve the problems in the prior art.
The technical solution of the invention is as follows:
according to one aspect of the invention, the pneumatic strut comprises a strut main body and a fixing device, wherein the strut main body is rod-shaped, one end of the strut main body is fixed at the foremost end of an aircraft through the fixing device, the other end of the strut main body is a free end, and the length and the shape of the pneumatic strut are obtained according to the optimal design of the appearance of an aircraft precursor, the overall performance index, the peak interference heat flow and the flight trajectory parameter.
Furthermore, the length of the pneumatic support rod is larger than the thickness of the material completely ablated in the full flight track time.
Furthermore, the pneumatic strut is in smooth connection transition with the aircraft forebody region.
According to a second aspect of the invention, a method for protecting the leading edge of a hypersonic aircraft from heat is provided, which comprises the following steps:
designing a pneumatic support rod, and fixing the pneumatic support rod at the foremost end of the aircraft;
designing a dredging type thermal protection structure of the aircraft;
analyzing the influence of the front edge incident shock waves on peak interference heat flow according to the influence of the pneumatic supporting rod on the front edge incident shock waves, and obtaining key parameters influencing the peak interference heat flow;
optimizing the local shape of the aircraft leading edge region through the obtained critical parameters to minimize the peak interference heat flow of the aircraft surface;
carrying out shock tunnel heat measurement test verification on the obtained aircraft leading edge region, finishing the design of the aircraft leading edge region if the verification result meets the design requirement, and adjusting according to key parameters influencing peak interference heat flow until the verification result meets the design requirement if the verification result does not meet the design requirement;
completing the design of the sparse heat protection structure on the optimized local appearance of the front edge area of the aircraft, and comparing by selecting heat sparse schemes of different materials to obtain a cooling effect and a rule influencing parameters of the cooling effect and obtain an optimal sparse heat protection structure;
performing an electric arc wind tunnel thermal examination test on the obtained optimal dredging type thermal protection structure, and verifying the cooling performance, the oxidation resistance, the thermal matching between the thermal dredging and the oxidation resistant materials and the like of the dredging type thermal protection structure;
and determining whether the design of the sparse conductive type thermal protection structure is finished according to the result of the thermal assessment test, if the result of the thermal assessment test meets the requirement, finishing the design of the sparse conductive type thermal protection structure, and if the result of the thermal assessment test does not meet the requirement, returning to adjust the pneumatic support rod or the sparse conductive type thermal protection structure until the requirement of the thermal assessment test is met.
Furthermore, the dredging type thermal protection structure is formed by coating an anti-oxidation material on the surface of a high-thermal-conductivity material.
Further, the key parameters affecting the peak disturbance heat flow are: the incident laser angle of the leading edge of the pneumatic strut, the object plane inclination angle of the interference part of the aircraft forebody and the curvature radius. The more parallel the object plane of the interference part of the aircraft precursor and the incident shock wave is, the smaller the peak interference heat flow is; the larger the radius of curvature of the aircraft forebody interference site, the smaller the peak dry heat flow.
Compared with the prior art, the invention has the beneficial effects that:
according to the invention, by selecting and combining the pneumatic support rod and the dredging type thermal protection scheme, the feasibility of the thermal protection scheme of the leading edge region can be ensured while the overall performance indexes such as lift-drag ratio of the hypersonic aerocraft are not reduced, and the problem of thermal protection of the leading edge region of the hypersonic aerocraft can be effectively solved.
Drawings
The accompanying drawings, which are included to provide a further understanding of the embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 illustrates a schematic flow chart provided in accordance with an embodiment of the present invention;
FIG. 2 is a schematic view of an aircraft leading edge with aerodynamic struts added and flow field disturbance provided according to an embodiment of the invention;
fig. 3 is a schematic structural diagram illustrating a heat channeling scheme according to an embodiment of the present invention;
fig. 4 is a schematic diagram illustrating a cooling effect of a heat dredging scheme according to an embodiment of the present invention.
The reference numbers in the drawings are as follows:
1. attaching shock waves; 2. conical shock waves; 3. a shear layer; 4. bow shock waves; 5. anti-oxidation C/C; 6. high thermal conductivity C/C.
Detailed Description
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It is noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of example embodiments according to the present application. As used herein, the singular forms "a", "an" and "the" are intended to include the plural forms as well, and it should be understood that when the terms "comprises" and/or "comprising" are used in this specification, they specify the presence of stated features, steps, operations, devices, components, and/or combinations thereof, unless the context clearly indicates otherwise.
The relative arrangement of the components and steps, the numerical expressions and numerical values set forth in these embodiments do not limit the scope of the present invention unless specifically stated otherwise. Meanwhile, it should be understood that the sizes of the respective portions shown in the drawings are not drawn in an actual proportional relationship for the convenience of description. Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.
According to the embodiment of the invention, the pneumatic supporting rod comprises a supporting rod main body and a fixing device, wherein the supporting rod main body is in a rod shape, one end of the supporting rod main body is fixed at the foremost end of an aircraft through the fixing device, the other end of the supporting rod main body is a free end, and in one embodiment, the top end of the free end is in an arc shape, so that the resistance of the aircraft is reduced; the length and the shape of the pneumatic strut are obtained according to the shape of the aircraft forebody, the overall performance index, the peak interference heat flow and the flight path parameter optimization design.
Further, in one embodiment, the length of the pneumatic strut is larger than the thickness of the ablation material in the full flight track time, so that the complete appearance of the aircraft body of the aircraft is ensured after the flight of the aircraft is finished.
In a further embodiment, the pneumatic strut is in smooth connection transition with the front body area of the aircraft, so that the overall lift-drag ratio performance of the aircraft can meet the index requirement.
In one embodiment, the peak interfering heat flow brought to the aircraft body by the pneumatic struts is as small as possible within the full flight trajectory range to meet the structural thermal protection requirements of the aircraft body.
According to a second embodiment of the invention, a method for thermally protecting a leading edge of a hypersonic flight vehicle is provided, which comprises the following steps:
designing a pneumatic support rod, fixing the pneumatic support rod at the foremost end of an aircraft, and optimally designing the length and the shape of the pneumatic support rod according to the appearance of a front body of the aircraft, the overall performance index, the peak interference heat flow and the flight path parameters;
the length of the pneumatic support rod is larger than the thickness of an ablation material in the full flight track time, so that the complete appearance of an aircraft body of the aircraft is ensured after the aircraft flies;
the pneumatic supporting rod is in smooth connection transition with a front body area of the aircraft, so that the overall lift-drag ratio performance of the aircraft can meet the index requirement;
in the full flight trajectory range, the peak value interference heat flow brought to the aircraft body by the pneumatic supporting rod needs to be as small as possible so as to meet the structural thermal protection requirement of the aircraft body.
And step two, designing a dredging type thermal protection structure of the aircraft, scientifically managing the flow direction of heat inside the thermal protection layer by adopting an energy management concept according to the characteristic that the heat flow distribution on the surface of the pneumatic support rod is extremely uneven, so that the heat in a high-temperature area is quickly transferred to a low-temperature area, effectively reducing the temperature and the temperature gradient of a front edge area, reducing the ablation retreating amount and the thermal stress of the pneumatic support rod, and not bringing serious pneumatic interference while realizing the ablation control of the pneumatic support rod.
In one embodiment, a design of a sparse thermal protection scheme is implemented by using a high thermal conductive material, and as the thermal conductivity and the oxidation resistance of the material are difficult to be considered, a layer of oxidation resistant material needs to be coated on the surface of the high thermal conductive material to improve the ablation performance. In a specific embodiment, according to the heat conductivity, the oxidation resistance and the matching between different materials, a high conductivity C/C material is used as a dredging material, and a layer of oxidation resistance C/C material is prepared on the surface of the material, so that the oxidation resistance is improved while the temperature of the front edge of the pneumatic support rod is reduced, and the structural schematic diagram is shown in FIG. 3. FIG. 4 is a schematic diagram illustrating the cooling effect of the thermal dispersion scheme, wherein the thermal dispersion scheme can reduce the temperature of the stagnation point of the leading edge of the pneumatic strut by more than 300 ℃;
analyzing the influence of the front edge incident shock wave on the peak value interference heat flow according to the influence of the pneumatic strut on the front edge incident shock wave, and obtaining key parameters influencing the peak value interference heat flow;
in one embodiment, based on a numerical simulation method, the influence rule of the interference of the incident shock wave of the front edge of the pneumatic strut on the peak interference heat environment of the surface of the aircraft is studied in detail, and key parameters influencing the peak interference heat flow are obtained, wherein the key parameters comprise the incident shock wave angle of the front edge of the pneumatic strut, the object plane inclination angle and the curvature radius of the interference part of the front body of the aircraft, and the object plane of the interference part of the front body of the aircraft is more parallel to the incident shock wave, so that the peak interference heat flow is smaller; the larger the radius of curvature of the aircraft forebody interference site, the smaller the peak dry heat flow.
Optimizing the local shape of the aircraft leading edge region through the obtained key parameters to reduce the peak interference heat flow on the surface of the aircraft as much as possible;
the peak interference heat flow of the aircraft surface is reduced as much as possible through the local shape optimization of the aircraft leading edge area, so as to ensure that the optimal heat reduction and drag reduction effects are achieved.
Step five, the obtained aircraft leading edge area is verified through a shock tunnel heat measurement test, if the verification result meets the design requirement, the design of the aircraft leading edge area is finished, and if the verification result does not meet the design requirement, the adjustment is carried out according to the key parameters in the step three until the verification result meets the design requirement;
sixthly, designing a sparse thermal protection structure for the optimized local appearance of the front edge area of the aircraft, and comparing the design by selecting thermal sparse schemes of different materials to obtain a cooling effect and rules influencing cooling effect parameters and obtain an optimal sparse thermal protection structure;
performing an electric arc wind tunnel thermal examination test on the obtained optimal dredging type thermal protection structure, and verifying the cooling performance, the oxidation resistance, the thermal matching between thermal dredging and oxidation resistant materials and the like of the dredging type thermal protection structure;
and step eight, determining whether the design of the sparse and conductive type thermal protection structure is finished according to the result of the thermal assessment test, finishing the design of the sparse and conductive type thermal protection structure if the result of the thermal assessment test meets the requirement, and returning to adjust the pneumatic support rod or the sparse and conductive type thermal protection structure until the requirement of the thermal assessment test is met if the result of the thermal assessment test does not meet the requirement. In one embodiment, if the pneumatic support rod does not pass a thermal assessment test, the thermal dredging and heat conducting protection structure needs to be adjusted, the cooling effect and the oxidation resistance of the thermal dredging scheme are comprehensively considered, and the overall ablation resistance is improved by adjusting the parameters such as the thickness of the oxidation resistant material; if the shock wave interference zone material of the aircraft body does not pass the thermal assessment test, the appearance and the geometric dimension of the pneumatic support rod need to be continuously optimized, and the peak interference thermal environment of the aircraft body is further reduced.
In conclusion, compared with the prior art, the hypersonic aircraft leading edge thermal protection method disclosed by the invention at least has the following advantages: according to the invention, by selecting and combining the pneumatic support rod and the dredging type thermal protection scheme, the feasibility of the thermal protection scheme of the leading edge region can be ensured while the overall performance indexes such as lift-drag ratio of the hypersonic aerocraft are not reduced, and the problem of thermal protection of the leading edge region of the hypersonic aerocraft can be effectively solved.
The above is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and various modifications and changes will occur to those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (2)
1. A thermal protection method for a front edge of a hypersonic aircraft is characterized by comprising the following steps: the method comprises the following implementation steps:
designing a pneumatic supporting rod and fixing the pneumatic supporting rod at the foremost end of an aircraft, wherein the pneumatic supporting rod comprises a supporting rod main body and a fixing device, the supporting rod main body is rod-shaped, one end of the supporting rod main body is fixed at the foremost end of the aircraft through the fixing device, the other end of the supporting rod main body is a free end, the length and the shape of the pneumatic supporting rod are obtained according to the optimal design of the appearance, the overall performance index, the peak interference heat flow and the flight track parameter of an aircraft precursor, the length of the pneumatic supporting rod is greater than the thickness of a completely ablated material in the time of a full flight track, and the pneumatic supporting rod is required to be smoothly connected and transited with an aircraft precursor region;
designing a dredging type thermal protection structure of the aircraft;
analyzing the influence of the front edge incident shock waves on peak interference heat flow according to the influence of the pneumatic supporting rod on the front edge incident shock waves, and obtaining key parameters influencing the peak interference heat flow;
optimizing the local shape of the aircraft leading edge region through the obtained critical parameters to minimize the peak interference heat flow of the aircraft surface;
the obtained aircraft leading edge area is verified through a shock tunnel heat measurement test, if the verification result meets the design requirement, the design of the aircraft leading edge area is completed, and if the verification result does not meet the design requirement, the adjustment is carried out according to key parameters influencing peak interference heat flow until the verification result meets the design requirement;
completing the design of the sparse heat protection structure on the optimized local appearance of the front edge area of the aircraft, and comparing by selecting heat sparse schemes of different materials to obtain a cooling effect and a rule influencing parameters of the cooling effect and obtain an optimal sparse heat protection structure;
performing an electric arc wind tunnel thermal examination test on the obtained optimal dredging type thermal protection structure, and verifying the cooling performance, the oxidation resistance, the thermal matching between the thermal dredging and the oxidation resistant materials and the like of the dredging type thermal protection structure;
and determining whether the design of the sparse conductive type thermal protection structure is finished according to the result of the thermal assessment test, if the result of the thermal assessment test meets the requirement, finishing the design of the sparse conductive type thermal protection structure, and if the result of the thermal assessment test does not meet the requirement, returning to adjust the pneumatic support rod or the sparse conductive type thermal protection structure until the requirement of the thermal assessment test is met.
2. The method for thermally protecting a leading edge of a hypersonic aircraft as claimed in claim 1, wherein: the key parameters affecting the peak interference heat flow are as follows: the incident laser angle of the leading edge of the pneumatic strut, the object plane inclination angle and the curvature radius of the interference part of the aircraft forebody.
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| CN113656949A (en) * | 2021-07-30 | 2021-11-16 | 深圳市中金岭南有色金属股份有限公司凡口铅锌矿 | Cooling effect analysis method, device, equipment and storage medium of cooling system |
| CN114722543B (en) * | 2022-06-09 | 2022-08-12 | 中国飞机强度研究所 | Design method for heat reflecting screen in structural heat strength test of hypersonic aircraft |
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