A wing system applied to a hypersonic vehicle

Abstract

The present invention discloses a wing system applied to a hypersonic aircraft, which includes a wing structure unit and a thermal sensing unit. The thermal sensing unit is arranged on the wing structure unit. The load-bearing structure in the wing structure unit includes a number of load-bearing I-beams, and a number of high thermal conductivity material blocks are inlaid on the load-bearing I-beams. The thermal sensing unit includes a signal acquisition and processing device and thin-film thermoelectric materials. The thin-film thermoelectric materials cover the surfaces of the high thermal conductivity material blocks, and the signal acquisition and processing device is installed on the load-bearing structure. The present invention combines force / thermal coupling to optimize the design of the wing structure, and utilizes the Seebeck effect generated by the thin-film thermoelectric materials to be able to detect the temperature of the wing surface, having the advantages of simple mechanical structure and integrated integration.

Description

Technical Field

[0001] The present invention belongs to the fields of the design of thermal protection systems for aerospace vehicles and the design of multifunctional thermoelectric materials, and in particular, to a wing system applied to hypersonic vehicles. Background Art

[0002] During the cruise of hypersonic vehicles, there is a severe thermal environment. Therefore, a large amount of effort is required to design the thermal protection system during the design process. The wing itself also needs to have the structural function of load-bearing, which requires considering the connection and coupling between systems, inevitably resulting in problems such as large self-weight, complex structure, and redundant structure. In addition, when the vehicle is in the cruise state, it is difficult to detect the temperature of the wing surface. Once a local heat transport failure occurs, it is difficult to locate its position. All of the above bring difficulties to the design of the wing system. Summary of the Invention

[0003] The purpose of the present invention is to solve the problems in the prior art, and provide a wing system applied to hypersonic vehicles, which integrates the functions of load-bearing, heat transport, and heat sensing by applying thin-film thermoelectric materials, and solves the problems such as large self-weight, complex structure, and redundant structure of the wing.

[0004] To achieve the above purpose, the present invention adopts the following technical solutions:

[0005] A wing system applied to hypersonic vehicles includes a wing structure unit and a heat sensing unit, and the heat sensing unit is arranged on the wing structure unit;

[0006] The load-bearing structure in the wing structure unit includes a number of load-bearing I-beams, and a number of high thermal conductivity material blocks are inlaid on the load-bearing I-beams;

[0007] The heat sensing unit includes a signal acquisition and processing device and thin-film thermoelectric materials. The thin-film thermoelectric materials cover the surface of the high thermal conductivity material blocks, and the signal acquisition and processing device is installed on the load-bearing structure.

[0008] Further, the load-bearing I-beam includes a main load-bearing beam, branch load-bearing beams, and a tree-like connection structure. The main load-bearing beam is connected to the tree-like connection structure, and the tree-like connection structure is connected to the branch load-bearing beams.

[0009] Further, the materials of the main load-bearing beam and the branch load-bearing beams are both TC4 titanium alloy, and the material of the tree-like connection structure is TB3 titanium alloy.

[0010] Further, the load-bearing structure further includes a load-bearing frame. A number of the load-bearing I-beams are arranged in the load-bearing frame, and the ends of the load-bearing I-beams are connected to the load-bearing frame.

[0011] Further, the high thermal conductivity material blocks include tree-branch-shaped embedded high thermal conductivity material blocks and parallel embedded high thermal conductivity material blocks. The tree-branch-shaped embedded high thermal conductivity material blocks are parallelepiped-shaped materials, and one side surface of the tree-branch-shaped embedded high thermal conductivity material blocks is embedded in the side surface of the load-bearing I-beam in a tree-branch shape, and one side surface of the parallel embedded high thermal conductivity material blocks is embedded in the side surface of the load-bearing I-beam in parallel.

[0012] Further, the surface of the load-bearing structure is enveloped by a wing skin, and the surface of the wing skin is enveloped by a heat protection layer.

[0013] Further, the material of the wing skin is titanium alloy, and the heat protection layer adopts an oxidation-resistant C / C composite material.

[0014] Further, the signal acquisition and processing device includes a temperature sensor, an electrical signal sensor, and a signal processing center;

[0015] The temperature sensor is installed in the wing root area of the load-bearing frame, the electrical signal sensor is installed at both ends of the tree-branch-shaped embedded high thermal conductivity material block, and the temperature sensor and the electrical signal sensor are signal-connected to the signal processing center.

[0016] Further, the thickness of the thin-film thermoelectric material is 100 - 1000 nm.

[0017] Further, the number of the load-bearing I-beams gradually increases along the horizontal chord and the vertical chord of the leading edge of the wing.

[0018] Compared with the prior art, the present invention has the following beneficial effects:

[0019] The present invention provides a wing system for a hypersonic aircraft, which can play a bearing role while completing the heat transport of the aircraft wing surface by embedding a high thermal conductivity material block on the side of a bearing I-beam, and the bearing I-beam is densely packed in the wing leading edge area of the bearing frame to accelerate the heat transport of the wing leading edge. The thin film thermoelectric material is covered on the surface of the high thermal conductivity material block embedded in a tree-like shape, and the Seebeck effect generated by the thin film thermoelectric material due to the temperature gradient is used. The temperature distribution of the entire wing surface can be obtained with the temperature of the wing root area of the bearing frame as the initial condition, and the temperature of the aircraft wing surface can be thermally sensed. When a local fault occurs in the heat transport of the aircraft wing surface, the fault area can be quickly located to detect whether the cruising state of the hypersonic aircraft is good at all times. In addition, the thin film thermoelectric material has the characteristics of light weight and simple structure, which can ensure that the aerodynamic heat generated by the hypersonic aircraft in the cruising state can be transported rapidly in a multi-channel and networked manner through a bearing structure with branched heat conduction, and the thin film thermoelectric material is used to realize the real-time dynamic detection of the wing surface temperature, so as to achieve "thermal perception" so that the fault point can be quickly found when the local wing surface temperature is abnormal. The present invention integrates the load-bearing, heat transport and heat sensing functions into one, and is lighter, simpler in structure and has fewer mechanical connection devices than traditional hypersonic aircraft.

[0020] Furthermore, no mechanical connection device is required between the thin-film thermoelectric material and the high thermal conductivity material block. The main load-bearing beam 101 and the branch load-bearing beam 102 are both made of TC4 titanium alloy material, the branch-like connecting structure 103 is made of TB3 titanium alloy material, the heat protection layer is made of antioxidant C / C composite material, and the wing skin is made of titanium alloy material. The mechanical structure is simple and hardly adds additional mass, size and load kinetic energy to the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings required for use in the embodiments are briefly introduced below. It should be understood that the following drawings only show certain embodiments of the present invention and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other related drawings can be obtained based on these drawings without creative work.

[0022] Figure 1 It is a schematic diagram of applying the wing system applied to a hypersonic aircraft of the present invention to the aircraft.

[0023] Figure 2 It is a schematic diagram of the interior of the wing structure unit of the present invention.

[0024] Figure 3 This is a schematic diagram of the structure in which the high thermal conductivity material block of the present invention is embedded in the load-bearing I-beam.

[0025] Figure 4Schematic diagram of the heat transport path of the wing system of the present invention.

[0026] Figure 5 Schematic diagram of the sensor positions of the signal acquisition and processing device of the present invention.

[0027] Figure 6 Schematic diagram when there is a temperature fault point on the wing surface of the present invention.

[0028] Wherein: 1 - load-bearing structure, 2 - high thermal conductivity material block, 3 - temperature sensor, 4 - electrical signal sensor, 11 - load-bearing I-beam, 12 - load-bearing frame, 21 - high thermal conductivity material block in a dendritic mosaic, 22 - high thermal conductivity material block in a parallel mosaic, 101 - main load-bearing beam, 102 - branch load-bearing beam, 103 - dendritic connection structure. Detailed implementation manners

[0029] To make the objectives, 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 accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. Generally, the components of the embodiments of the present invention described and illustrated in the accompanying drawings here can be arranged and designed in various different configurations.

[0030] Therefore, the following detailed description of the embodiments of the present invention provided in the accompanying drawings is not intended to limit the scope of the claimed present invention, but merely represents selected embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0031] It should be noted that: similar reference numerals and letters denote similar items in the following drawings. Therefore, once an item is defined in one drawing, it does not need to be further defined and explained in subsequent drawings.

[0032] In the description of the embodiments of the present invention, it should be noted that if terms such as "upper", "lower", "horizontal", "inner", etc. are used to indicate the orientation or positional relationship, it is based on the orientation or positional relationship shown in the accompanying drawings, or the orientation or positional relationship in which the product of the present invention is usually placed during use. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus should not be construed as a limitation of the present invention. In addition, terms such as "first", "second", etc. are only used for distinguishing descriptions and cannot be understood as indicating or implying relative importance.

[0033] In addition, when the term "horizontal" appears, it does not mean that the component is required to be absolutely horizontal, but it can be slightly inclined. For example, "horizontal" only means that its direction is more horizontal relative to "vertical", and it does not mean that the structure must be completely horizontal, but it can be slightly inclined.

[0034] In the description of the embodiments of the present invention, it should also be noted that unless otherwise clearly specified and limited, when terms such as "arranged", "installed", "connected", and "coupled" appear, they should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0035] The present invention will be further described in detail below with reference to the accompanying drawings:

[0036] See Figures 1 to 6 , the present invention provides a wing system applied to a hypersonic vehicle, integrating load-bearing / heat transport / heat sensing functions, mainly composed of a load-bearing / heat transport integrated wing structure unit and a heat sensing unit based on thin-film thermoelectric materials. It includes a load-bearing structure 1, a high thermal conductivity material block 2, a temperature sensor 3, an electrical signal sensor 4, a load-bearing I-beam 11, a load-bearing frame 12, a dendritically embedded high thermal conductivity material block 21, a parallelly embedded high thermal conductivity material block 22, a main load-bearing beam 101, a branch load-bearing beam 102, and a dendritically connecting structure 103.

[0037] The wing structure unit mainly includes a load-bearing structure 1, a high thermal conductivity material block 2, a wing skin, and a thermal protection layer. The load-bearing structure 1 includes a load-bearing I-beam 11 and a load-bearing frame 12. The load-bearing I-beam 11 includes a main load-bearing beam 101, a branch load-bearing beam 102, and a dendritically connecting structure 103. The high thermal conductivity material block 2 is divided into a dendritically embedded high thermal conductivity material block 21 and a parallelly embedded high thermal conductivity material block 22 according to the embedding position. The dendritically embedded high thermal conductivity material block 21 is a parallelepiped block material, which is adhered to the side concave part of the load-bearing I-beam 11 with a smaller side area by a high-temperature resistant adhesive, so that it is arranged in a dendritic shape. The parallelly embedded high thermal conductivity material block 22 is adhered to the side concave part of the load-bearing I-beam 11 with a larger side area by a high-temperature resistant adhesive. After embedding, the high thermal conductivity material block 2 needs to fully fill the side concave part of the load-bearing I-beam 11 and be closely combined up and down. The high thermal conductivity material block 2 can be made of graphite, silicon, alumina ceramic, or silicon carbide ceramic material, and has the characteristics of good thermal conductivity, high temperature resistance, high mechanical strength, and stable chemical properties. The load-bearing I-beam 11 and the load-bearing frame 12 are connected by a high-temperature resistant adhesive. The wing skin envelopes the upper and lower surfaces of the load-bearing structure 1, and the thermal protection layer envelopes the upper and lower surfaces of the wing skin.

[0038] The thermal sensing unit includes thin-film thermoelectric materials and a signal acquisition and processing device. The signal acquisition and processing device includes a temperature sensor 3, an electrical signal sensor 4, and a signal processing center. As shown Figure 5 in the figure, the temperature sensor 3 is installed in the wing root area of the bearing frame 12, and the electrical signal sensor 4 is installed at the cold end and the hot end of the high thermal conductivity material block 21 in a tree-branch-like inlay. Generally, the end close to the bearing I-beam 11 is the hot end, and the other end is the cold end. The thin-film thermoelectric materials cover the surface of the high thermal conductivity material block 21 in a tree-branch-like inlay. When the aircraft is in a cruising state, the temperature distribution on the wing surface is uneven, and the high thermal conductivity material block 21 in a tree-branch-like inlay has a hot end and a cold end. Due to the Seebeck effect generated by the thin-film thermoelectric materials, a potential difference will be generated between the hot and cold ends of the high thermal conductivity material block 21 in a tree-branch-like inlay. This potential difference is collected by the electrical signal sensor 4 and processed by the signal processing center located in the aircraft cockpit to convert the potential difference into a temperature difference. It should be noted that the temperature in the wing root area of the bearing frame 12 is measured by the temperature sensor 3 as an initial condition, and based on the calculated temperature difference, the entire temperature of the wing surface can be obtained.

[0039] Since the load-bearing conduction direction and the heat transport direction of the wing have a high degree of similarity, the force transmission and heat transfer are coupled. A number of bearing I-beams 11 are arranged along the horizontal chord direction and the vertical chord direction of the load-bearing structure 1. The high thermal conductivity material blocks 2 inlaid in the side recesses of each bearing I-beam 11 form multiple heat transport channels, greatly increasing the heat transfer area of the wing surface of the aircraft. As shown Figure 6 in the figure, the bearing I-beams 11 are arranged more densely at the leading edge of the wing surface of the bearing frame 12 to increase the tree-branch-like branched heat transport channels, in order to achieve rapid heat transport from the high-temperature area to the low-temperature area and reduce the temperature in the high-temperature area.

[0040] The thickness of the thin-film thermoelectric materials in the thermal sensing unit is between 100 and 1000 nm, and there is no need for a mechanical connection device between the thin-film thermoelectric materials and the high thermal conductivity material block 2. The main load-bearing beam 101 and the branch load-bearing beam 102 are made of TC4 titanium alloy material, the tree-branch-like connection structure 103 is made of TB3 titanium alloy material, the thermal protection layer is made of oxidation-resistant C / C composite material, and the wing skin is made of titanium alloy material, hardly adding extra mass, size, and load kinetic energy to the aircraft.

[0041] The working principle of the wing system applied to the hypersonic aircraft provided by the present invention is as follows:

[0042] The wing system applied to the hypersonic aircraft of the present invention mainly includes a load-bearing / heat transport integrated wing structure unit and a thermal sensing unit based on thin-film thermoelectric materials. The wing structure unit is the structural basis of the wing system, and the thermal sensing unit mainly realizes the function of real-time monitoring of the wing surface temperature and quickly locating local fault areas.

[0043] The wing structure unit has both load-bearing and heat transport functions. The load-bearing I-beam 11 covers the wing surface. Heat is transferred from the high-temperature area to the low-temperature area through the parallelly inlaid high-thermal-conductivity material blocks 22 and spreads through the tree-branch-shaped inlaid high-thermal-conductivity material blocks 21. Due to the existence of the high-thermal-conductivity material blocks 2, the wing surface is almost covered with branched heat conduction channels, and the heat transport effect is greatly enhanced.

[0044] The heat sensing unit includes thin-film thermoelectric materials and a signal acquisition and processing device. The signal acquisition and processing device needs to collect the potential difference generated at the cold and hot ends of the thin-film thermoelectric materials and the temperature in the wing root area of the load-bearing frame 12. The collected electrical signals are processed by the signal processing center, and the temperature difference between the cold and hot ends of the thin-film thermoelectric materials can be obtained. Taking the temperature in the wing root area of the load-bearing frame 12 as the initial condition, through step-by-step iteration by the built-in calculator in the signal processing center, the temperature distribution of the entire wing surface is gradually calculated from the wing root area of the load-bearing frame 12 to the wing tip area of the load-bearing frame 12. It should be noted that due to the existence of a temperature gradient in the thin-film thermoelectric materials, a cold end and a hot end are generated, and according to the Seebeck effect of the thermoelectric materials, a potential difference can be generated. The electrical signal sensors 4 are located at the cold and hot ends of the thin-film thermoelectric materials, and thus the potential difference can be detected. When the aircraft is in good cruising condition, the wing surface temperature shows a gradually changing distribution with different gray levels, and the gray level in the high-temperature area is greater than that in the low-temperature area. When a local heat transport failure occurs on the wing surface, the temperature gray level at the location is abnormal ( Figure 6 represented by a circle), and it can be judged that this local area is in an abnormal state. Therefore, the heat sensing unit based on the thin-film thermoelectric materials can not only monitor the wing surface temperature in real time but also quickly locate the position of the fault point.

[0045] Based on the integrated and systematic wing design concept, and facing the requirements of the wing of a supersonic aircraft for load-bearing, heat transfer, and heat sensing, starting from the overall design discipline of the aircraft, a wing system applied to hypersonic aircraft is proposed. This system has the following characteristics: First, the wing system integrates the functions of load-bearing, heat transfer, and heat sensing. It is lighter in mass and simpler in structure compared to traditional hypersonic aircraft, and the connections of various parts of the wing system mostly use high-temperature-resistant adhesives, which are lighter in mass compared to traditional mechanical connection devices. Second, there is a high similarity in the distribution of thermal loads and force loads in the wing system. A high thermal conductivity material block 2 is inlaid in the concave part of the load-bearing beam I-beam 11 to complete the role of heat transfer while supporting the wing structure. The tree-branch-shaped inlaid high thermal conductivity material blocks 21 are distributed on both sides of the load-bearing beam I-beam 11 to achieve the role of branched heat conduction, increasing the heat conduction channels and significantly improving the heat transfer area of the wing surface. At the same time, due to more branched heat conduction channels arranged at the leading edge of the wing, more heat energy can be transported. Third, the thin-film thermoelectric material covers the high thermal conductivity material block 2, and the signal acquisition and processing device can collect the electrical signals generated by the Seebeck effect and output the temperature distribution of the aircraft wall surface, which has guiding significance for studying the true state of the aircraft.

[0046] The wing system of the present invention applied to hypersonic aircraft, which integrates the functions of load-bearing, heat transfer, and heat sensing, can combine with the true state of the wing thermal protection system during flight and is used to guide the optimization design of the wing thermal protection system and the demonstration of the flight ability of the aircraft.

[0047] The above are only the preferred embodiments of the present invention and are not used to limit the present invention. For those skilled in the art, the present invention can have various changes and modifications. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included within the protection scope of the present invention.

Claims

1. A wing system applied to a hypersonic vehicle, characterized in that It includes a wing structure unit and a heat sensing unit, and the heat sensing unit is arranged on the wing structure unit; The load-bearing structure (1) in the wing structure unit includes a number of load-bearing I-beams (11), and a number of high thermal conductivity material blocks (2) are inlaid on the load-bearing I-beams (11); The heat sensing unit includes a signal acquisition and processing device and thin film thermoelectric materials. The thin film thermoelectric materials cover the surface of the high thermal conductivity material blocks (2), and the signal acquisition and processing device is installed on the load-bearing structure (1); The load-bearing I-beam (11) includes a main load-bearing beam (101), a branch load-bearing beam (102) and a dendritic connection structure (103). The main load-bearing beam (101) is connected to the dendritic connection structure (103), and the dendritic connection structure (103) is connected to the branch load-bearing beam (102); The high thermal conductivity material blocks (2) include dendritically inlaid high thermal conductivity material blocks (21) and parallelly inlaid high thermal conductivity material blocks (22). The dendritically inlaid high thermal conductivity material blocks (21) are parallelepiped-shaped materials, and one side of the dendritically inlaid high thermal conductivity material blocks (21) is dendritically inlaid on the side surface of the load-bearing I-beam (11), and one side of the parallelly inlaid high thermal conductivity material blocks (22) is parallelly inlaid on the side surface of the load-bearing I-beam (11).

2. The wing system for a hypersonic vehicle according to claim 1, characterized in that The materials of the main load-bearing beam (101) and the branch load-bearing beam (102) are both TC4 titanium alloy, and the material of the dendritic connection structure (103) is TB3 titanium alloy.

3. The wing system for a hypersonic vehicle according to claim 1, wherein The load-bearing structure (1) further includes a load-bearing frame (12). A number of the load-bearing I-beams (11) are arranged in the load-bearing frame (12), and the ends of the load-bearing I-beams (11) are connected to the load-bearing frame (12).

4. The wing system for a hypersonic vehicle according to claim 1, characterized in that The surface of the load-bearing structure (1) is enveloped by a wing skin, and the surface of the wing skin is enveloped by a heat protection layer.

5. The wing system for a hypersonic vehicle according to claim 4, characterized in that The material of the wing skin is titanium alloy, and the heat protection layer adopts oxidation-resistant C / C composite material.

6. The wing system for a hypersonic vehicle according to claim 1, characterized in that The signal acquisition and processing device includes a temperature sensor (3), an electrical signal sensor (4) and a signal processing center; The temperature sensor (3) is installed in the wing root area of the load-bearing frame (12), the electrical signal sensor (4) is installed at both ends of the dendritically inlaid high thermal conductivity material blocks (21), and there is a signal connection between the temperature sensor (3) and the electrical signal sensor (4) and the signal processing center.

7. The wing system for a hypersonic vehicle according to claim 1, characterized in that, The thickness of the thin film thermoelectric material is 100~1000nm.

8. The wing system for a hypersonic vehicle according to claim 1, characterized in that, The number of the load-bearing I-beams (11) gradually increases along the horizontal chord and the vertical chord of the wing leading edge.

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