CN113032911A - Aircraft heat-proof tile gap structure, design method and heat-proof tile - Google Patents

Aircraft heat-proof tile gap structure, design method and heat-proof tile Download PDF

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Publication number
CN113032911A
CN113032911A CN202110411523.5A CN202110411523A CN113032911A CN 113032911 A CN113032911 A CN 113032911A CN 202110411523 A CN202110411523 A CN 202110411523A CN 113032911 A CN113032911 A CN 113032911A
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heat
seam
flow
diversion trench
aircraft
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CN113032911B (en
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廖军好
张志刚
赵金山
石义雷
陈挺
粟斯尧
肖雨
罗万清
余嘉
彭治雨
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Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center
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Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center
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    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/02—Reliability analysis or reliability optimisation; Failure analysis, e.g. worst case scenario performance, failure mode and effects analysis [FMEA]
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/08—Thermal analysis or thermal optimisation
    • 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 an aircraft heat-proof tile gap structure, a design method and a heat-proof tile, wherein the gap structure comprises a T-shaped gap, the top of the windward side wall at the joint of the T-shaped gap is provided with a diversion trench, the bottom surface of the diversion trench is a curved surface, and the depth of the diversion trench is gradually reduced along the direction of flowing to the T-shaped gap; the edges of the aircraft heat-proof tile gap structure are provided with chamfers. The three-dimensional chamfer configuration adopted by the invention can effectively avoid high pneumatic heating of the windward side of the T-shaped seam, and solve the problem of local high heat flow of the heat-proof tile seam of the hypersonic aircraft.

Description

Aircraft heat-proof tile gap structure, design method and heat-proof tile
Technical Field
The invention relates to the technical field of hypersonic aircraft design and thermal protection, in particular to an aircraft heat-proof tile gap structure, a design method and a heat-proof tile.
Background
At present, complex hypersonic aircrafts such as space shuttles and the like mainly select and install heat-proof tiles as main means for pneumatic heat protection. Due to factors such as thermal expansion, aerodynamic load, assembly tolerance and the like, transition areas such as micro steps, gaps and the like inevitably occur at the joints of the heat-proof tiles. And the phenomena of shock wave/boundary layer interference, separation reattachment, transition from laminar flow to turbulent flow and the like accompanying the gap flow enable the pneumatic thermal environment at the gap to be extremely bad. The research result shows that: the local high heat flow area of the heat-proof tile gap is mainly concentrated on the windward side wall of the T-shaped gap (as shown in figure 1), and the peak heat flow can reach dozens to twenty times of the value of the local non-interference heat flow. On one hand, if the design margin is insufficient, a local structure can be burnt out, and the safety of the whole aircraft is threatened; on the other hand, if the design margin of the thermal protection system is increased, the payload of the aircraft is reduced, and the performance of the aircraft is reduced.
Therefore, in view of the above disadvantages, it is necessary to provide an effective means for solving the problem of local high heat flow in the gaps of the thermal tile of the hypersonic aircraft.
Disclosure of Invention
Technical problem to be solved
The invention aims to solve the technical problem that the T-shaped seam windward side wall of the heat-proof tile in the prior art is poor in pneumatic thermal environment and high in local peak heat flow.
(II) technical scheme
In order to solve the technical problem, the invention provides an aircraft heat-proof tile gap structure which comprises a T-shaped gap, wherein a diversion trench is arranged at the top of the windward side wall at the joint of the T-shaped gap, the bottom surface of the diversion trench is a curved surface, and the depth of the diversion trench is gradually reduced along the direction of the flow-direction gap; the edges of the aircraft heat-proof tile gap structure are provided with chamfers.
Preferably, the radius of the curved surface of the diversion trench is not less than half of the width of the flow direction slot in the T-shaped slot.
Preferably, the bottom surface of the diversion trench is a cylindrical surface, and the included angle α formed between the axis of the cylindrical surface and the upper surface of the heat-proof tile is in the range of 30-45 °.
Preferably, the initial depth of the guide groove is determined according to the air flow deflection angle β of the main impact air flow flowing into the flow direction slot and the position of the air flow stagnation point, and the maximum depth does not exceed the curved surface radius of the groove bottom surface of the guide groove.
The invention also provides a design method of the aircraft heat-proof tile gap structure, which comprises the following steps:
s1, performing aerodynamic thermal environment numerical simulation calculation on the T-shaped seam of the heat-proof tile seam structure to obtain a first calculation result, and determining the position and the size of a local high heat flow area of the T-shaped seam, and the air flow deflection angle beta and the air flow stagnation point position of main impact air flow flowing into the seam;
s2, according to the first calculation result, arranging a diversion trench on the top of the windward side wall at the joint of the T-shaped seam, wherein the bottom surface of the diversion trench is a curved surface, the depth of the diversion trench is gradually reduced along the direction of the flow-direction seam, and chamfers are arranged on the edges of the diversion trench;
s3, carrying out aerodynamic thermal environment numerical simulation calculation on the T-shaped seam with the diversion trench to obtain a second calculation result;
s4, correcting the angle, the depth and the edge chamfer of the diversion trench according to the second calculation result;
s5, performing aerodynamic heating numerical simulation calculation on the corrected T-shaped seam; and if the heat flow distribution is smoothly transited from the upper surface of the heat-proof tile to the inside of the T-shaped seam, ending the iteration, otherwise, updating the second calculation result according to the calculation result of the step S5, returning to the step S4, continuing to correct, and repeating the iteration to obtain the optimal shape.
Preferably, in step S2, when the guiding groove is formed, the radius of the curved surface of the guiding groove is not less than half of the width of the flow-direction slit in the T-shaped slit.
Preferably, in the step S2, when the guiding groove is opened, the bottom surface of the guiding groove is a cylindrical surface, and an included angle α formed between an axis of the cylindrical surface and the upper surface of the heat-proof tile is 30 ° to 45 °.
Preferably, in step S1, when performing the aerodynamic thermal environment numerical simulation calculation, a finite volume method is used to solve the hypersonic aerodynamic thermal numerical calculation method and the flow field display method of the three-dimensional NS equation, so as to obtain the position and size of the local high heat flow region of the T-shaped seam, and the cross-sectional pressure, density and streamline distribution of the flow direction seam center line in the T-shaped seam, and further determine the position of the stagnation point of the airflow on the windward wall of the T-shaped seam and the deflection angle β of the airflow in the flow direction seam.
Preferably, in step S2, according to the airflow stagnation point position and the airflow deflection angle β determined by the first calculation result, the initial depth of the guiding groove and the included angle α formed between the axis of the cylindrical surface and the upper surface of the heat protection tile are set, and the initial depth of the guiding groove does not exceed the radius of the curved surface of the groove bottom surface of the guiding groove at most.
The invention also provides an aircraft heat-proof tile which is provided with the aircraft heat-proof tile gap structure or is designed by adopting the design method of the aircraft heat-proof tile gap structure.
(III) advantageous effects
The technical scheme of the invention has the following advantages: the invention provides an aircraft heat-proof tile gap structure, a design method and a heat-proof tile, and provides a three-dimensional chamfer configuration for greatly reducing local high pneumatic heating of a windward side wall of a T-shaped gap through a sunken diversion trench according to the flowing characteristics of the gap and the main mechanism of high pneumatic heating of the windward side wall of the T-shaped gap by combining the influence rule of geometric parameter variables on the pneumatic heating of the T-shaped gap, so that the local high pneumatic heating of the heat-proof tile gap on the outer surface of a hypersonic aircraft can be effectively reduced, and the performance of the aircraft is further improved.
Drawings
FIG. 1 is a set of flat plate T-shaped slot model aerodynamic heat numerical calculation heat flow cloud chart (unit kW/m)2)；
FIG. 2(a) is a partial enlarged view of a conventional T-shaped seam profile;
FIG. 2(b) is a cloud of aerodynamic heat flow calculated from the aerodynamic heat values of FIG. 2(a) (in kW/m2)；
FIG. 3(a) is a two-dimensional flow diagram of flow into a conventional T-slot, where β is the air flow deflection angle and the air flow stagnation point is marked;
FIG. 3(b) is a schematic diagram of the construction of a "three-dimensional chamfer" in an embodiment of the present invention;
FIG. 3(c) is a schematic diagram showing a two-dimensional cut with the airflow stagnation point as the cylinder center;
FIG. 3(d) is a schematic structural view of a gap between heat-proof tiles of an aircraft according to an embodiment of the present invention;
FIG. 4 is a hypersonic aerothermic calculated heat flux cloud (in kW/m) for the slot structure shown in FIG. 3(d) in an embodiment of the present invention2)；
FIG. 5(a) is a hypersonic aerothermic calculated heat flux cloud for a conventional T-slot at an edge chamfer radius R of 0.1 mm;
FIG. 5(b) is a hypersonic aerothermic calculated heat flux cloud for a conventional T-slot at an edge chamfer radius R of 0.3 mm;
FIG. 5(c) is a hypersonic aerothermic calculated heat flux cloud for a conventional T-slot at an edge chamfer radius R of 0.5 mm;
FIG. 6 is a graph comparing the heat flux on the centerline of the windward sidewall of the slot shown in FIG. 3(d) and the conventional T-slot shown in FIGS. 5(a) to 5(c) in accordance with an embodiment of the present invention;
FIG. 7 is a schematic step diagram of a method for designing an aircraft heat protection tile gap structure according to an embodiment of the invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, 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 some, but not all, embodiments of the present invention. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
As shown in fig. 1 to 2(b), the present invention utilizes the prior art aerodynamic thermal numerical simulation means (including a gas flow calculation program developed by using a finite volume method, numerical discrete solution of a three-dimensional NS equation system, etc.) to study the aerodynamic thermal environment of a conventional T-shaped seam. As shown in fig. 1, a group of conventional T-shaped seams is generated on the upper surface of a blunt front edge flat plate as a carrier, a single T-shaped seam structure is shown in fig. 2(a), the T-shaped seam includes an "|" type flow direction seam (namely, a vertical seam) and a "-" type windward side seam (namely, a transverse seam), the end of the flow direction seam is communicated with the windward side seam, and the flow direction seam points to the windward side wall (the flow direction seam of the T-shaped seam is 3mm wide than the windward side seam, and the edge is chamfered by 1mm), and the pneumatic heat value calculation is performed on the T-shaped seam by using a high-precision calculation method, so that as shown in fig. 1 and fig. 2(b), the peak heating area of the T-shaped seam is concentrated on the top of the windward side wall (namely, the side where the heat-proof tile is in contact with the external flow field), and the severe pneumatic heating is mainly due to the flow direction seam existing at the front part of the windward side wall of the T-shaped seam, on, on the other hand, due to the boundary layer effect of the flow to the boundary of the object plane of the T-shaped slot, only the top area of the windward side wall of the T-shaped slot is severely pneumatically heated.
As shown in fig. 3(a) to 6, an aircraft heat protection tile gap structure provided in an embodiment of the present invention includes a T-shaped slot, wherein a (recessed) diversion trench is disposed at a top of a windward side wall at a junction of the T-shaped slot (a flow direction slot and a windward side slot), the diversion trench corresponds to the flow direction slot in the T-shaped slot, as shown in fig. 3(d), a bottom surface of the diversion trench is a curved surface, a depth of the diversion trench is gradually reduced along a direction of the flow direction slot, so that the diversion trench is inclined, and a diversion direction of the diversion trench is coplanar with a direction of the flow direction slot, so as to guide an incoming flow entering the flow direction slot and reduce impact of the incoming flow on an upper portion of the windward side wall. The edges of the aircraft heat-proof tile gap structures are provided with chamfers, so that smooth transition between the diversion trench and the outer wall of the heat-proof tile is ensured. As shown in fig. 4 to 5(c), under the same heat flow scale, the appearance of the aircraft heat-proof tile gap structure adopting the three-dimensional chamfer angle is free of a local high heat flow area on the windward side wall, so that the pneumatic heating amount is effectively reduced.
In the prior art, in order to avoid the strong impact of the air flow outside the flowing direction seam boundary layer, the pneumatic heating is reduced by adopting a 'edging' mode of a two-dimensional chamfer angle to the windward side seam generally, but actually the contribution of the scheme to reducing the pneumatic heating is very limited, and the overall roughness of the surface of the aircraft is possibly greatly improved due to the overlarge radius of the chamfer angle, so that the disturbance in the boundary layer near the wall surface is increased to cause the transition in advance, the overall pneumatic heating of the aircraft is increased, and the performance of the aircraft is influenced. The invention adopts the structure that the flow guide groove is arranged at the top part in front of the flow direction seam (namely one side impacted by an incoming flow field), and the three-dimensional chamfer angle is adopted, so that the high pneumatic heating of the windward side of the T-shaped seam can be effectively avoided, the problem of local high heat flow of the heat-proof tile seam of the hypersonic aircraft is solved, and the blank of the prior art is filled.
Preferably, to ensure the flow guiding effect of the flow guiding groove, the radius of the curved surface of the flow guiding groove is not less than half of the width of the flow direction seam in the T-shaped seam.
Preferably, as shown in fig. 3(a) to 3(d), the three-dimensional chamfer configuration in the aircraft thermal tile gap structure provided by the invention can be realized by obliquely cutting a cylinder at the joint of the flow direction seam and the windward side seam, that is, the bottom surface of the diversion trench is a cylindrical surface, and the axis of the cylindrical surface forms an included angle α with the upper surface of the thermal tile. The included angle alpha can be determined by the air flow deflection angle beta of the incoming flow (namely the main impact air flow) flowing into the flow gap and the position of an air flow stagnation point, further, the included angle alpha formed by the axis of the cylindrical surface (namely the axis of the cutting cylinder) and the upper surface of the heat-proof tile ranges from 30 degrees to 45 degrees, more preferably from 30 degrees to 40 degrees, and after the cylinder is cut, a cone-like opening is formed at the top of the windward side wall. Within a certain range, the larger the radius of the cutting cylinder, the lower the heat flow after cutting.
Further, the initial depth of the diversion trench (i.e., the depth of the diversion trench near one end of the flow slot) can be determined according to the air flow deflection angle β of the main impact air flow flowing into the slot and the position of the air flow stagnation point formed by the air flow on the windward wall, the position of the air flow stagnation point is a function of the depth h, the length L and the air flow deflection angle β of the flow slot, and the larger the air flow deflection angle β is, the more the air flow stagnation point is detected, the deeper the initial depth of the corresponding diversion trench is. The intersection point position of the axis of the cutting cylinder and the windward side wall is preferably not higher than the airflow stagnation point so as to ensure that the airflow stagnation point is eliminated after cutting, thereby effectively reducing the impact of downward detection of the incoming flow on the windward side of the T-shaped seam. Meanwhile, in order to avoid introducing new disturbance, the initial depth of the diversion trench should not exceed the curved surface radius of the trench bottom surface of the diversion trench to the maximum.
The flow deflection angle β and the specific position of the stagnation point of the main impingement flow flowing into the slot can be calculated by using aerodynamic thermal numerical simulation in the prior art, as shown in fig. 3 (a). The shielding of the main impingement airflow of the flow slot is cut off by a cylinder with the same width as the flow slot, so that the main impingement airflow can smoothly pass through the position, thereby reducing pneumatic heating, as shown in fig. 3(b) to 3 (d).
In order to further quantitatively analyze the effectiveness of the pneumatic heating of the invention, as shown in fig. 4 to 6, the invention also carries out the local high pneumatic heating numerical simulation comparison analysis of the conventional T-shaped seam with different edge (two-dimensional) chamfer radiuses and the aircraft heat-proof tile seam structure provided by the invention, fig. 5(a), 5(b) and 5(c) respectively show hypersonic aerodynamic heat calculation heat flow cloud graphs of the conventional T-shaped seam when the edge chamfer radiuses are R equal to 0.1mm, 0.3mm and 0.5mm, fig. 3(d) shows the aircraft heat-proof tile seam structure provided by the invention, the conventional T-shaped seam is realized by cutting a cylinder at the top of the windward side wall of the T-shaped seam, the cut-off outer shape edge is further chamfered, the edge chamfer radius R equal to 0.5mm, the diameter d of the cutting cylinder is the same as the width of the flow direction seam in the T-shaped seam (d equal to 3mm), the included angle alpha between the axis of the cylindrical surface and the upper surface of the heat-proof tile is 30 degrees, the initial depth of the diversion trench is determined according to the air flow deflection angle beta of the main impact air flow in the flow direction slot and the position of the air flow stagnation point, preferably, the axis of the cylindrical surface passes through the air flow stagnation point, and fig. 4 is a hypersonic aerodynamic heat calculation heat flow cloud diagram of the gap structure of the heat-proof tile of the aircraft shown in fig. 3 (d). As shown in fig. 6, it can be known from comparison of simulation results that the conventional two-dimensional chamfer configuration can reduce the pneumatic heating of the windward side of the T-shaped seam, but the effect is not obvious, and even if the chamfer radius is increased to 0.5mm, the peak heat flow is only reduced by 30.6%, whereas the three-dimensional chamfer configuration provided by the invention can effectively avoid the high pneumatic heating of the windward side of the T-shaped seam, and compared with the conventional T-shaped seam with the chamfer radius R of 0.1mm at the edge, the peak heat flow heat reduction is as high as 73.9%.
As shown in fig. 7, the invention further provides a design method of the aircraft heat-proof tile gap structure, which specifically comprises the following steps:
s1, performing aerodynamic thermal environment numerical simulation calculation on the T-shaped seam of the heat-proof tile seam structure to obtain a first calculation result, and determining the position and the size of a local high heat flow area of the T-shaped seam, and the air flow deflection angle beta and the air flow stagnation point position of main impact air flow flowing into the seam;
s2, according to the first calculation result, arranging a diversion trench on the top of the windward side wall at the joint of the T-shaped seam, wherein the bottom surface of the diversion trench is a curved surface, the depth of the diversion trench is gradually reduced along the direction of the flow-direction seam, and chamfers are arranged on the edges of the diversion trench;
s3, carrying out numerical simulation calculation on the aerodynamic thermal environment of the T-shaped seam with the flow guide groove to obtain a second calculation result;
s4, correcting the angle, the depth and the edge chamfer angle of the guide groove according to the second calculation result;
s5, performing aerodynamic heating numerical simulation calculation on the corrected T-shaped seam; and if the heat flow distribution is smoothly transited from the heat-proof tile to the inside of the T-shaped seam, ending the iteration, otherwise, updating the second calculation result according to the calculation result of the step S5, returning to the step S4, continuing to correct, and repeating the iteration to obtain the optimal shape.
As shown by the inner curve of the elliptical area in fig. 6, for the T-shaped seam after the guiding gutter is formed, in the result of the aerodynamic heat value simulation calculation, the heat flow distribution in the smooth transition from the upper surface of the heat-proof tile to the T-shaped seam is significantly reduced relative to the two-dimensional edge chamfer, but a new peak value may occur, in order to obtain an optimal appearance, the aerodynamic heat value simulation calculation and the parameter correction of the guiding gutter may be performed through multiple iterations, the guiding gutter is optimized, and when the heat flow distribution is in the smooth transition from the upper surface of the heat-proof tile to the T-shaped seam, that is, the inner curve of the elliptical area shown in fig. 6 becomes smooth, the formed guiding gutter may be considered to be optimal, and the optimal appearance of the seam.
Preferably, in step S2, when the guiding groove is opened, the radius of the curved surface of the guiding groove is not less than half of the width of the flow-direction slit in the T-shaped slit.
Preferably, in step S2, when the guiding groove is opened, the bottom surface of the guiding groove is a cylindrical surface, that is, the guiding groove can be realized by a cutting cylinder, and the included angle α formed between the axis of the cylindrical surface and the upper surface of the heat-proof tile is in the range of 30 ° to 45 °.
Preferably, the design method adopts a hypersonic aerodynamic thermal numerical calculation method for solving a three-dimensional NS equation by a finite volume method to carry out the numerical simulation calculation of the aerodynamic thermal environment.
Further, in step S1, when performing the aerodynamic thermal environment numerical simulation calculation, a hypersonic velocity aerodynamic thermal numerical calculation method and a flow field display method for solving a three-dimensional NS equation by using a finite volume method are used to obtain the position and size of the local high heat flow region of the T-shaped seam, and the cross-sectional pressure, density and streamline distribution of the flow direction seam center line in the T-shaped seam, so as to determine the position of the stagnation point of the airflow on the windward wall of the T-shaped seam and the deflection angle β of the airflow in the flow direction seam.
Furthermore, in step S2, according to the airflow stagnation point position and the airflow deflection angle β determined by the first calculation result, the initial depth of the guiding groove and the included angle α formed between the axis of the cylindrical surface and the upper surface of the heat protection tile, that is, the angle of the axis of the cutting cylinder for cutting the guiding groove, are set, and the initial depth of the guiding groove does not exceed the radius of the curved surface of the groove bottom surface of the guiding groove at most.
Preferably, in steps S3 to S5, the three-dimensional chamfered T-shaped seam is designed as an input condition, and profile parameters such as the angle, the depth, the edge chamfer radius and the like of the flow guide groove are iteratively modified one by one based on a prediction result obtained by a hypersonic aerodynamic thermal value calculation method for solving a three-dimensional NS equation by using a finite volume method, so as to obtain an optimal profile. Further, in step S4, when the shape parameter of the T-shaped seam is corrected, the angle of the guiding groove and the edge chamfering radius are fixed, the depth of the guiding groove is corrected, the angle of the guiding groove is set, and the edge chamfering radius is set.
The invention also provides an aircraft heat-proof tile which is provided with the aircraft heat-proof tile gap structure in any one of the above embodiments or is designed by adopting the aircraft heat-proof tile gap structure design method in any one of the above embodiments.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention, but not to limit it; although the present invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equivalently replaced; and such modifications or substitutions do not depart from the spirit and scope of the corresponding technical solutions of the embodiments of the present invention.

Claims (10)

1. The utility model provides an aircraft heat protection tile crack structure which characterized in that:
the T-shaped seam is included, a diversion trench is arranged at the top of the windward side wall at the joint of the T-shaped seam, the bottom surface of the diversion trench is a curved surface, and the depth of the diversion trench is gradually reduced along the direction of the flow direction seam;
the edges of the aircraft heat-proof tile gap structure are provided with chamfers.
2. The aircraft thermal tile gap structure of claim 1, wherein: the radius of the curved surface of the diversion trench is not less than half of the width of the flow direction seam in the T-shaped seam.
3. The aircraft thermal tile gap structure of claim 1, wherein: the bottom surface of the diversion trench is a cylindrical surface, and the included angle alpha formed by the axis of the cylindrical surface and the upper surface of the heat-proof tile ranges from 30 degrees to 45 degrees.
4. The aircraft heat protection tile slot structure of claim 3, wherein: the initial depth of the flow guide groove is determined according to the air flow deflection angle beta of the main impact air flow in the flow direction slot and the air flow stagnation point position, and the maximum depth of the flow guide groove does not exceed the curved surface radius of the groove bottom surface of the flow guide groove.
5. A design method for an aircraft heat protection tile gap structure is characterized by comprising the following steps:
s1, performing aerodynamic thermal environment numerical simulation calculation on the T-shaped seam of the heat-proof tile seam structure to obtain a first calculation result, and determining the position and the size of a local high heat flow area of the T-shaped seam, and the air flow deflection angle beta and the air flow stagnation point position of main impact air flow flowing into the seam;
s2, according to the first calculation result, arranging a diversion trench on the top of the windward side wall at the joint of the T-shaped seam, wherein the bottom surface of the diversion trench is a curved surface, the depth of the diversion trench is gradually reduced along the direction of the flow-direction seam, and chamfers are arranged on the edges of the diversion trench;
s3, carrying out aerodynamic thermal environment numerical simulation calculation on the T-shaped seam with the diversion trench to obtain a second calculation result;
s4, correcting the angle, the depth and the edge chamfer of the diversion trench according to the second calculation result;
s5, performing aerodynamic heating numerical simulation calculation on the corrected T-shaped seam; and if the heat flow distribution is smoothly transited from the upper surface of the heat-proof tile to the inside of the T-shaped seam, ending the iteration, otherwise, updating the second calculation result according to the calculation result of the step S5, returning to the step S4, continuing to correct, and repeating the iteration to obtain the optimal shape.
6. The design method according to claim 5, wherein:
in the step S2, when the guiding groove is opened, the radius of the curved surface of the guiding groove is not less than half of the width of the flow direction slot in the T-shaped slot.
7. The design method according to claim 5, wherein:
in the step S2, when the guiding groove is opened, the bottom surface of the guiding groove is a cylindrical surface, and an included angle α formed between an axis of the cylindrical surface and the upper surface of the heat-proof tile is in a range of 30 ° to 45 °.
8. The design method according to claim 7, wherein:
in step S1, when performing the aerodynamic thermal environment numerical simulation calculation, a finite volume method is used to solve the hypersonic aerodynamic thermal numerical calculation method and the flow field display method of the three-dimensional NS equation, so as to obtain the position and size of the local high heat flow region of the T-shaped seam, and the cross-sectional pressure, density and streamline distribution of the flow direction seam center line in the T-shaped seam, thereby determining the position of the stagnation point of the airflow on the windward wall of the T-shaped seam and the deflection angle β of the airflow in the flow direction seam.
9. The design method according to claim 8, wherein:
in the step S2, according to the airflow stagnation point position and the airflow deflection angle β determined by the first calculation result, the initial depth of the guiding groove and the included angle α formed between the axis of the cylindrical surface and the upper surface of the heat-proof tile are set, and the initial depth of the guiding groove does not exceed the radius of the curved surface of the groove bottom surface of the guiding groove at most.
10. An aircraft thermal tile, characterized in that:
having an aircraft insulation tile gap structure as claimed in any of claims 1 to 4, or
The design method of the aircraft thermal tile gap structure according to any one of claims 5 to 9 is used for designing the thermal tile gap structure.
CN202110411523.5A 2021-04-16 2021-04-16 Aircraft heat-proof tile gap structure component, design method and heat-proof tile Active CN113032911B (en)

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