Active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle

Abstract

According to the active thermal protection and thermal power generation integrated system based on the supercritical Brayton cycle, the two active cooling modes of convection cooling and sweating cooling are combined, and the energy utilization rate is increased; according to the gradient cooling structure of the nose cone front edge of the high-speed aircraft, the nose cone front edge is divided into a top end section, a middle section and a low end section, heat flows of the three sections are sequentially decreased, and porous metal matrix layers are arranged on the three sections of the nose cone front edge in a gradient mode according to the porosity; a two-stage regenerative-shunt recompression closed Brayton cycle power configuration obtained through calculation and optimization is adopted, and the characteristics of high energy recovery efficiency and compact system structure of the closed Brayton cycle are achieved; and a wing rib array is arranged in the heat exchange channel, so that a larger heat exchange surface area is provided.

Description

Technical Field

[0001] The present invention belongs to the technical field of thermal protection and heat utilization of aerospace vehicles, and in particular, relates to a thermal protection-power generation integrated system. Background Art

[0002] When a high-speed aircraft is flying in the atmosphere, the aerodynamic heating caused by the friction with the incoming atmosphere will cause the temperature of the aircraft's outer surface to rise rapidly, resulting in a decrease in the strength and stiffness of the aircraft structure. For high-speed aircraft that need to operate for a long time, their outer surface is in a high-temperature environment for a long time, with a large heat flow. In order to ensure their normal operation, the thermal protection requirements are very high.

[0003] Traditional aircraft thermal protection system designs usually remove heat by ablation or use the material's own properties to resist heat, with poor repeatability and sustainability, and cannot meet the long-term working requirements of aircraft. Although active thermal protection technologies such as regenerative cooling, air film cooling and sweat cooling are more effective, the use of a single active thermal protection method often requires a large amount of cooling fluid, which inevitably leads to excessive weight of high-speed aircraft. At the same time, long-term operation of the aircraft will inevitably increase the requirements for the power supply carried, resulting in large losses. Taking targeted and multi-thermal protection methods for different key parts of the aircraft, and flexibly and efficiently utilizing the advantages of different thermal protection technologies, is an effective way to solve the thermal protection problem of high-speed aircraft under high heat load. In addition, the high-temperature thermal energy generated by aerodynamic heating is a high-grade energy. The high temperature of aerodynamic heat is transferred and dissipated to the outside world through the cooling medium, and this part of energy cannot be truly utilized. Among the traditional active cooling methods, only convection cooling can effectively utilize this part of energy, converting heat energy into electrical energy, replacing the power supply required to be carried by the aircraft, and reducing the required coolant consumption, but it requires additional circulating fluids and heat exchange equipment, thereby increasing the weight of the aircraft. Therefore, it is necessary to establish a mathematical model of such a pneumatic heat-to-electricity conversion cycle system, describe different functions and corresponding index parameters, and optimize the cycle configuration to obtain the most suitable cycle configuration.

[0004] The above-mentioned convection cooling thermoelectric conversion technology using aerodynamic heat can meet the cooling needs of most areas of the engine and high-speed aircraft fuselage, but it is difficult to effectively cool the ultra-high temperature components at the leading edge of the nose cone that are subjected to severe aerodynamic heating. The coolant used for cooling the circulating working fluid in the above-mentioned thermoelectric conversion system is transported to the leading edge of the aircraft nose cone to form an open loop. Through the sweat cooling method, the coolant absorbs the extreme temperature of the leading edge of the nose cone and evaporates and discharges to the external space and forms an air film on the surface of the leading edge of the nose cone, which can effectively reduce the high-temperature heat flow to the leading edge of the aircraft nose cone. The pyrolyzable materials currently used for filling cannot be reused, and it is impossible to distribute the coolant in a targeted manner to regulate the unevenly distributed thermal load on the leading edge of the nose cone, which may cause insufficient coolant in some areas and excessive coolant in other areas, which cannot be discharged in time, resulting in deterioration of heat transfer and excessive coolant consumption. Changing the sweat cooling structure of the leading edge of the nose cone and reasonably distributing the coolant are effective means to improve cooling efficiency and ensure the safe operation of the nose cone of the aircraft.

[0005] In summary, there is still room for development in the application of thermal protection measures and aerodynamic heat recovery during the flight of high-speed aircraft. Summary of the invention

[0006] The purpose of the present invention is to solve the problems existing in the above-mentioned prior art. The present invention proposes a thermal protection-power generation integrated system, which adopts different cooling methods according to different parts of the aircraft. The present invention provides an active thermal protection and thermal power generation integrated system based on the supercritical Brayton cycle, which combines two active cooling methods: convection cooling and sweat cooling. The aerodynamic heat of the aircraft surface is used for thermoelectric conversion through convection cooling, which achieves thermal protection while replacing part of the power source carried by the aircraft. At the same time, since part of the heat is used for expansion work and converted into electricity, the heat that needs to be cooled by the coolant is correspondingly reduced, and the required coolant consumption is reduced. The high-temperature coolant after absorbing the exhaust steam heat is transported to the gradient cooling structure at the leading edge of the aircraft nose cone through a pipeline, and the extremely high temperature part of the nose cone is cooled by sweat cooling, thereby improving energy utilization.

[0007] The present invention adopts a gradient cooling nose cone leading edge cooling structure. The nose cone leading edge is divided into three sections: the top, the middle and the bottom. The heat flow in the three regions decreases in sequence. The porous metal matrix layer is arranged in a gradient manner in the three sections of the nose cone leading edge according to the porosity. The pore size of the gradient arranged porous metal matrix layer corresponds to the heat flow, which avoids the blockage of excess coolant in the pores, better removes heat, reduces the consumption of sweating coolant, and is more conducive to the long-term operation of high-speed aircraft.

[0008] In order to improve the thermal efficiency and work capacity of the circulation system and reduce the heat loss in the circulation process, the present invention applies the closed Brayton cycle to the recovery and utilization of aerodynamic heat of high-speed aircraft, and adopts the recuperator shunt recompression closed Brayton cycle power configuration obtained through calculation and optimization, which has the characteristics of high energy recovery efficiency of the closed Brayton cycle and compact system structure, and maximizes the collection and utilization of aerodynamic heat energy and reduces energy loss under the premise of ensuring the thermal protection of the aircraft cabin.

[0009] In addition, the present invention adopts a rib array structure microchannel with strong turbulence effect and low pressure loss as a heat exchange channel. The circulating working fluid flows through the heat exchange channel and performs convective heat exchange with the high-temperature wall of the aircraft cabin. Due to the presence of the rib array in the microchannel, more heat exchange surface area is provided, and the working fluid generates strong disturbance when flowing in the channel, thereby improving the heat exchange performance; at the same time, thanks to the streamlined structure of the rib surface, the pressure loss of the working fluid when flowing through the rib array is relatively small, which greatly reduces the negative effect of pressure loss caused by adding a turbulent body in the channel.

[0010] The present invention provides a thermal protection-power generation integrated system, which comprises a thermal protection module, a power generation module, a coolant tank and connecting pipes.

[0011] Furthermore, the thermal protection module includes a gradient sweating cooling structure and a convection cooling heat exchange channel, the gradient sweating cooling structure is arranged at the leading edge of the nose cone of the high-speed aircraft, and the convection cooling heat exchange channel is arranged on the wall of the aircraft cabin; the circulating working fluid in the convection heat exchange channel absorbs the heat from the aircraft wall and then passes into the power generation module; the power generation module is arranged in the high-speed aircraft cabin, and the circulating working fluid converts the absorbed heat into electricity in the power generation module; the coolant tank is connected to the power generation module and the thermal protection module in turn through the connecting pipeline, the coolant enters the power generation module through the coolant tank for heat exchange, and after heating, enters the gradient sweating cooling structure in the thermal protection module to cool the extreme heat flow at the leading edge of the nose cone of the high-speed aircraft;

[0012] The power generation module adopts a closed Brayton cycle power configuration with regenerator diversion and recompression, including: a turbine, a generator, a cooler, a high-temperature regenerator, a low-temperature regenerator, a main compressor, a recompressor, a booster pump, and a control valve; the circulating working fluid absorbs the heat from the surface of the aircraft cabin through the convection cooling heat exchange channel to undergo overheating expansion, and then enters the turbine to link with the engine to generate electricity; the circulating working fluid after doing work enters the high and low-temperature regenerators for heat exchange respectively, and is divided into a mainstream and a secondary flow at the hot end outlet of the low-temperature regenerator, the mainstream part enters the cooler for heat exchange with the coolant, then enters the main compressor for pressurization, and then enters the low-temperature regenerator for heat exchange, the secondary flow part enters the recompressor for pressurization, then merges with the mainstream part at the cold end outlet of the low-temperature regenerator, enters the high-temperature regenerator for heat exchange, and finally enters the heat exchange channel to start a new cycle;

[0013] The thermal protection module introduces a gradient sweating cooling structure, including: a porous matrix section with large porosity, a porous matrix section with medium porosity, a porous matrix section with small porosity, a porous matrix inner layer, a tapered delivery pipe, a diversion impact hole and a water storage tank; the porous metal matrix nose cone cooling structure is filled with porous metal with a graded porosity gradient reduction from the top to the bottom of the nose cone front edge, and the coolant penetrates through the matrix layers with different porosities to each temperature interval for sweating cooling;

[0014] The convection cooling heat exchange channel adopts a strong disturbance low pressure loss wing rib heat exchange channel, and a wing rib array structure is arranged in the channel. The working fluid flows through the wing rib array turbulence structure to enhance heat exchange and absorb heat from the surface of the aircraft cabin;

[0015] As a further improvement of the technical solution of the present invention, the porous matrix with large, medium and small porosity in the gradient sweating cooling structure is respectively distributed and filled in the top, middle and bottom of the gradient sweating cooling structure in a gradient manner. For the bottom area with smaller heat flow, the porous matrix section with small porosity reduces the coolant consumption and the accumulation of coolant in the pores by controlling the coolant sweating cooling rate; the coolant is transported to the cooling cabin and accelerated to the inner surface of the porous matrix at the top of the gradient sweating cooling structure through the tapered delivery pipe, and at the same time, the coolant is uniformly and quickly transported to the leading edge of the head cone of the gradient sweating cooling structure through the diversion impact hole;

[0016] As a further improvement of the technical solution of the present invention, the high-disturbance low-pressure-loss rib heat exchange channel optimizes the number and arrangement of the ribs by adding a smooth rib structure in the channel, thereby forming a staggered rib array; when the fluid flows through the rib array, the fluid collision disturbance is enhanced and the heat exchange area is increased; at the same time, when passing through the rib array, the streamlined surface of the rib structure reduces the resistance encountered by the fluid during the flow process, thereby suppressing the flow pressure loss and suppressing the pump power consumption;

[0017] As a further improvement of the technical solution of the present invention, the circulating working fluids that can be used in the closed Brayton cycle power configuration with regenerator splitting and recompression include but are not limited to supercritical carbon dioxide (sCO2) and helium (He);

[0018] As a further improvement of the technical solution of the present invention, the recommended circulating working fluid split ratio of the closed Brayton cycle power configuration with regenerator split recompression at the outlet of the low-temperature regenerator is in the range of 0.2 to 0.4;

[0019] As a further improvement of the technical solution of the present invention, the porosity of the porous matrix segment with large porosity is x1, the porosity of the porous matrix segment with medium porosity is x2, the porous matrix segment with small porosity is x3, the porosity of the permeable layer is x4, and x1=x4>x2>x3;

[0020] As a further improvement of the technical solution of the present invention, the diversion impact holes are evenly distributed along the wall of the tapered conveying pipe, and the apertures of the diversion impact holes are in the millimeter range.

[0021] The beneficial effects of the present invention are as follows:

[0022] 1. The present invention relates to a thermal protection and thermoelectric conversion integrated system that couples a closed convection cooling system with an open transpiration cooling system; it combines two active cooling methods, convection cooling and transpiration cooling. The aerodynamic heat of the surface of a high-speed aircraft is absorbed by convection cooling and used for thermoelectric conversion, which achieves thermal protection while converting heat into electricity, thereby reducing or even replacing the battery power source carried by the aircraft. At the same time, since part of the heat is used for expansion work and converted into electricity, the amount of heat that needs to be cooled by the coolant is correspondingly reduced, and the required coolant consumption is reduced. The high-temperature coolant that has absorbed the exhaust steam heat is transported through a pipeline to the gradient cooling structure at the leading edge of the aircraft nose cone, and the extremely high temperature part of the nose cone is gradient cooled by transpiration cooling, thereby improving energy utilization.

[0023] 2. The present invention relates to a gradient cooling structure for the leading edge of a nose cone of a high-speed aircraft; the leading edge of the nose cone is divided into three sections: the top, the middle and the bottom, and the heat fluxes of the three regions decrease in sequence. The porous metal matrix layers are arranged in a gradient manner in the three sections of the leading edge of the nose cone according to the porosity; the pore size of the gradient-arranged porous metal matrix layer corresponds to the heat flux, and the porous matrix section with small porosity reduces the consumption of coolant and the accumulation of coolant in the pores by controlling the sweating cooling rate of the coolant, thereby reducing the consumption of sweating coolant and being more conducive to the long-term operation of the high-speed aircraft; the tapered delivery pipe preferentially delivers to the inner surface of the porous matrix at the top of the leading edge of the gradient-cooled nose cone, and at the same time, the diversion impact holes thereon evenly and quickly deliver the coolant to other sections of the leading edge of the gradient-cooled nose cone to better take away the heat;

[0024] 3. The present invention relates to a closed Brayton cycle power configuration with two-stage heat recovery-divided flow recompression; the closed Brayton cycle power configuration with two-stage heat recovery-divided flow recompression obtained by calculation and optimization has the characteristics of high energy recovery efficiency of the closed Brayton cycle and compact system structure, and improves the thermal efficiency and work capacity of the cycle system under the premise of ensuring the thermal protection of the aircraft cabin, maximizes the collection and utilization of aerodynamic thermal energy, and reduces energy loss;

[0025] 4. The present invention relates to a heat exchange microchannel with a rib array. The rib array is arranged in the heat exchange channel by utilizing micro-surface processing design, and the circulating working fluid flows through the heat exchange channel to perform convective heat exchange with the high-temperature wall of the aircraft cabin. Due to the presence of the rib array in the microchannel, more heat exchange surface area is provided, and the working fluid generates strong disturbance when flowing in the channel, so the heat exchange performance is improved, and more heat is absorbed by the circulating working fluid, thereby ensuring that the temperature of the aircraft wall is within a safe range while improving the power generation capacity of the thermoelectric conversion subsystem. At the same time, thanks to the streamlined structural design of the rib surface, the pressure loss of the working fluid when flowing through the rib array is relatively small, thereby reducing the pressure loss caused by adding a turbulent structure in the heat exchange channel and suppressing the negative effect of increased pump power consumption. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] The accompanying drawings constituting a part of the present application are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention:

[0027] Figure 1 It is a schematic diagram of a thermal protection-power generation integrated system according to an embodiment of the present invention;

[0028] Figure 2 Schematic diagram of the cooling structure of the porous metal matrix head cone with gradient distribution according to an embodiment of the present invention;

[0029] Figure 3 2 is a schematic diagram of a high-interference low-pressure-loss rib heat exchange channel structure according to an embodiment of the present invention;

[0030] Figure 4 Schematic diagram of a simple Brayton cycle power configuration (S-BC) according to an embodiment of the present invention;

[0031] Figure 5 Schematic diagram of a regenerative Brayton cycle power configuration (R-BC) according to an embodiment of the present invention

[0032] Figure 6 Schematic diagram of a regenerator split-flow recompression Brayton cycle power configuration (RR-BC) according to an embodiment of the present invention;

[0033] Figure 7The mass flow power density of the three power configuration layouts of the present invention varies with the compressor pressure ratio;

[0034] Figure 8 The mass power density of the three power configuration layouts of the present invention varies with the compressor pressure ratio;

[0035] Fig. 9 The total mass of the three power configuration layouts of the present invention varies with the compressor pressure ratio.

[0036] Among them: 1. Circulating fluid pump; 2. Gradient cooling structure at the leading edge of the nose cone; 3. High-disturbance low-pressure loss wing rib heat exchange channel; 4. Heat exchange channel guide; 5. Turbine; 6. Generator; 7. High-temperature regenerator; 8. Low-temperature regenerator; 9. Cooler; 10. Compressor; 11. Recompressor; 12. Control valve; 13. Pressure pump; 14. Coolant tank; 15. Aircraft cabin wall; 21. Large-porosity porous matrix section; 22. Medium-porosity porous matrix section; 23. Small-porosity porous matrix section; 24. Liquid storage tank; 25. Permeation layer; 26. Reduction delivery pipe; 27. Diversion impact hole; 31. Microchannel inlet; 32. Microchannel outlet; 33. Wing rib structure. DETAILED DESCRIPTION

[0037] In order to enable those skilled in the art to better understand the scheme of the present invention, the technical scheme in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work should fall within the scope of protection of the present invention.

[0038] This embodiment provides a thermal protection-power generation integrated system, and its composition diagram is as follows: Figure 1As shown, it includes an aircraft nose cone leading edge gradient cooling structure 2 and an aircraft cabin aerodynamic heat and thermoelectric conversion subsystem; the thermoelectric conversion subsystem includes a circulating working fluid pump 1, a wing rib heat exchange channel 3, a heat exchange channel guide port 4, a turbine 5, a generator 6, a high-temperature regenerator 7, a low-temperature regenerator 8, a cooler 9, a compressor 10, a recompressor 11, a control valve 12, a booster pump 13 and a coolant tank 14; the specific working process is: under the action of the circulating working fluid pump 1, the circulating working fluid enters the strong disturbance low pressure loss wing rib heat exchange channel 3 through the heat exchange channel guide port 4, exchanges heat with the high-temperature wall of the high-speed aircraft cabin, absorbs heat, and flows into the turbine 5 through the heat exchange channel guide port 4, the working fluid expands to do work, and the generator 6 is linked to generate electricity, The exhaust steam after work is passed into the high-temperature regenerator 7 and the low-temperature regenerator 8 for heat exchange, and the working fluid is diverted for heat exchange, and the mainstream part is passed into the cooler 9. The coolant is pressurized and pumped into the cooler 9 from the coolant tank 14 through the working fluid pump 13 to exchange heat with the mainstream working fluid, and after taking away the heat of the mainstream working fluid, it is transported to the gradient cooling structure 2 at the leading edge of the nose cone of the aircraft. The mainstream working fluid flows from the cooler 9 into the compressor 10 for compression, and after compression, it enters the low-temperature regenerator 8 for heat exchange with the undivided working fluid. The secondary flow working fluid is directly passed from the outlet of the low-temperature regenerator 8 into the compressor 11 for pressurization, and then merges with the mainstream working fluid, enters the high-temperature regenerator 7 to absorb heat and heat up, and finally, under the action of the circulating working fluid pump 1, a new round of circulation begins.

[0039] like Figure 2 As shown, the gradient-cooled aircraft nose cone leading edge structure of the present invention comprises a porous matrix section 21 with a high porosity, a porous matrix section 22 with a medium porosity, a porous matrix section 23 with a low porosity, a permeable layer 25, a tapered delivery pipe 26 and a diversion impact hole 27; according to the different heat flow from the top to the bottom of the leading edge structure, the porous matrix section 21 with a high porosity, the porous matrix section 22 with a medium porosity and the porous matrix section 23 with a low porosity are used in sequence to take away the extreme heat flow of the nose cone leading edge in the form of sweat cooling; the matrix section 21 with a high porosity is conducive to timely replenishing the coolant to ensure the safety of the top of the nose cone leading edge; the matrix sections 22 and 23 with a medium and low porosity control the pore size to avoid the blockage of the excess coolant in the pores, thereby better taking away the heat and reducing the sweat coolant consumption, which is more beneficial to the long-term operation of the high-speed aircraft;

[0040] like Figure 3 As shown, the high-disturbance low-pressure-loss heat exchange channel in the present invention includes a microchannel inlet 31, a microchannel outlet 32 ​​and a rib structure 33; the rib structure 33 is staggered in the heat exchange channel to form a rib array, which significantly increases the heat exchange area. The circulating working fluid flows into the heat exchanger through the microchannel inlet 31, and generates strong disturbances when flowing through the rib array, thereby enhancing the heat exchange effect between the working fluid and the high-temperature compartment wall. At the same time, the streamlined structure of the rib surface suppresses the pressure loss of the working fluid when it flows through, reduces the power consumption required by the circulating working fluid pump, and finally flows out through the microchannel outlet 32.

[0041] Preferred Embodiments

[0042] The following embodiments are provided by using the regenerator split flow recompression power configuration and the comparative configuration mentioned in the present invention:

[0043] The cooling performance and quality performance of the regenerator split recompression power configuration (RR-BC) used in the pneumatic heat-to-electricity conversion subsystem are compared with the regenerator configuration (R-BC) and the basic configuration (S-BC) as the research object. The schematic diagrams of the three configurations are shown in Figures 4 to 6 As shown. Given the turbine inlet temperature (the outlet temperature of the circulating fluid in the heat exchange channel) of 673K, the compressor inlet pressure of 7800kPa, the compressor inlet temperature of 315K, the compressor pressure ratio of 2.1-2.5, the high and low temperature regenerator pinch point temperature of 10K, the turbine isentropic efficiency of 0.9, the compressor and recompressor isentropic efficiency of 0.89, the cooling water inlet temperature of 293K, the inlet pressure of 200kPa, and for the convenience of comparison, the cooling water outlet temperature is given as 353K. Based on the closed Brayton cycle energy analysis equation and the thermodynamic calculation equation, the results of the two key evaluation indicators of the mass power density and mass flow power density of the aerodynamic thermal thermoelectric conversion subsystem under different power configurations are obtained through numerical calculation.

[0044] (1) Mass flow power density Wfr

[0045] In order to analyze the cooling performance of different power configurations in the aerodynamic heat and heat conversion subsystem, it is necessary to establish a thermodynamic model of the heat and heat conversion subsystem and conduct parameter discussion and analysis. The basic equations for closed Brayton cycle energy analysis are shown in Table 1.

[0046] Table 1 Basic energy equation of working fluid in the cycle

[0047] Heat load and power consumption of main components Thermodynamic equations Heat exchanger capacity <![CDATA[Q=m h C p,h (T h,o -T h,i )=m h C p,h (T c,o -T c,i )=UAΔT]]> Turbine shaft power <![CDATA[W t =m w ·(h tout -h tin ) <!-- 5 -->]]> Compressor power consumption <![CDATA[W c =m w ·(h cout -h cin )]]> Net power generation <![CDATA[W net =In t -IN c ]]>

[0048] In the traditional thermal cycle power generation process operating on the ground, it is usually assumed that the amount of cooling water is unlimited and is not used as a system evaluation indicator. However, for high-speed aircraft, cooling water is the only cold source to maintain the temperature of the aircraft, and the cooling water consumption is directly related to the mass load of the long-flight high-speed aircraft system. In order to objectively evaluate the power generation corresponding to each unit mass flow of cooling water of different Brayton power configurations under different working conditions, the evaluation indicator of mass flow power density is introduced, which is defined as:

[0049]

[0050] Where mf is the cooling water mass flow rate required by the thermoelectric conversion subsystem.

[0051] The cooling water mass flow rate is defined as:

[0052]

[0053] Where Q total is the total heat absorbed by the circulating medium through the heat exchange channel, C f is the specific heat capacity of cooling water, T f0 and T f1 is the inlet and outlet temperature of the cooling water in the cooler; η is the net power of the thermoelectric conversion subsystem.

[0054] The net power of the base configuration, regenerator configuration, and regenerator split recompression configuration are defined as:

[0055]

[0056]

[0057]

[0058] Where h is the enthalpy value at each point, and α is the split ratio of the working fluid in the regenerator split recompression configuration.

[0059] (2) Mass power density Ws

[0060] In order to estimate the mass performance of different power configurations in the thermoelectric conversion subsystem, a mass estimation model was proposed based on the thermodynamic model to calculate the total mass of the power configuration used.

[0061] The total mass of the power configuration used in the thermoelectric conversion subsystem is defined as:

[0062] M total =M re +M c +M rc +M tb

[0063] Among them, M re is the total mass of the regenerator (including high-temperature regenerator and low-temperature regenerator), M c is the mass of the cooler, M rc and M tb They are the masses of the compressor (including the re-compressor) and the turbine respectively.

[0064] In the integrated thermal protection-power generation system, the turbine and compressor are both turbine equipment, and their total mass per unit power is assumed to be a constant value. In the aerodynamic thermal thermoelectric conversion subsystem, the mass of the compressor and expander is 50 kg. The regenerator and cooler are both heat exchange equipment, and their mass is related to parameters such as their own heat transfer coefficient and compactness, and are calculated using the interpolation method.

[0065] Thermodynamic model calculations of heat transfer capacity (UA) and pressure loss (Δp)

[0066] Q=UAΔT

[0067] Δp=P i -P o

[0068] Where ΔT is the logarithmic mean temperature difference, P i is the fluid inlet pressure, P o is the fluid outlet pressure.

[0069] The known parameters such as UA, Δp, working fluid flow rate mw, logarithmic mean temperature difference ΔT of the heat exchanger, and wall temperature Tw are input into the model together, and the thermal physical parameters of the fluid used, thermal conductivity (λ), density (ρ), dynamic viscosity (μ), and constant pressure specific heat capacity (Cp) of the fluid are calculated according to the current state. The number of circular tubes (n) and the length of the circular tube (L) are linearly averaged to form a two-dimensional matrix of n and L, and based on the above-mentioned fluid thermal physical parameters, the Reynolds number kinematic viscosity (v), Reynolds number (R e ) and coefficient of friction (Cf):

[0070]

[0071]

[0072]

[0073] in ρ is the fluid density, dh is the hydraulic diameter of the heat exchanger pipe, and n is the number of pipes used in the heat exchanger.

[0074] According to the formula, the corresponding pressure loss approximate value matrix (Δpx) and thermal conductivity approximate value matrix (UAx) are calculated. The pressure loss approximate value matrix (Δpx) is defined as

[0075]

[0076] The thermal conductivity approximation matrix (UAx) can be defined as:

[0077] UA x =h c A s

[0078] Where hc is the heat transfer coefficient in the heat exchanger. As is the heat transfer area, which are defined as:

[0079]

[0080] A s =3.14d h nL

[0081] Where, for the hot end of the heat exchanger, Ti and Tw are the outlet temperature and outlet wall temperature of the heat exchanger fluid, respectively; for the cold end of the heat exchanger, Ti and Tw are the inlet temperature and inlet wall temperature of the heat exchanger fluid, respectively. λ is the thermal conductivity of the fluid.

[0082] By introducing a two-dimensional interpolation function, the heat exchanger pressure loss Δp and UA values ​​obtained by thermodynamic calculations are interpolated and matched with the Δpx matrix and UAx matrix respectively. When the values ​​of Δp and UA are within the value range of the Δpx matrix and UAx matrix, the values ​​closest to Δp and UA in the two matrices are selected as approximate values, and the specific number of round tubes (n) and length (L) are returned. The approximate volume Vx of the heat exchanger is calculated from the specific n and L values ​​obtained as the approximate value of the actual heat exchanger volume V. The heat conductivity approximate value matrix (UAx) can be defined as:

[0083]

[0084] Where σP is the porosity of the heat exchanger material used.

[0085] The mass M of the heat exchanger can be obtained by multiplying the approximate volume Vx by the density ρ of the material used in the heat exchanger. In order to further measure the relationship between the system power generation and the additional mass of the power configuration, the mass power density is further proposed and defined as:

[0086] W s =W net / M total

[0087] The high mass power density means that the mass load of the power configuration is utilized more efficiently, that is, the introduction of the aerodynamic thermal and thermoelectric conversion subsystem produces higher benefits.

[0088] like Figure 7 and Figure 8 As shown: In the gas-powered thermal-to-electric conversion subsystem, the maximum mass flow power density of the regenerative configuration (R-BC) and the regenerator split-flow recompression configuration (RR-BC) are 89.5 kJ / kg and 164 kJ / kg, respectively, which are about 114% and 292% higher than the 41.8 kJ / kg of the basic configuration (S-BC). The maximum mass power density of the regenerative configuration and the regenerator split-flow recompression configuration are 42.2 W / kg and 31.5 W / kg, respectively, which are about 24.8% higher and 6.80% lower than the 33.8 W / kg of the basic configuration. The two key performances of the regenerative configuration are better than those of the basic configuration, while the regenerator split-flow recompression configuration has a more complex configuration and a larger total mass of components, such as Fig. 9As shown, the result is that the maximum mass flow rate is slightly lower than the basic configuration, but its maximum mass flow power density is nearly three times that of the basic configuration. The optimal range of the diversion ratio calculated is within the range of 0.2 to 0.4. When the compressor pressure ratio π is 2.2, the mass power density of the basic, reheat, and reheater diversion recompression configurations are similar, which are 28.6W / kg, 28.3W / kg, and 27.9W / kg, respectively, but the mass flow power density is 34.5kJ / kg, 61.3kJ / kg, and 146kJ / kg, respectively. Compared with the basic configuration, the mass flow power density of the reheat configuration and the reheater diversion recompression configuration is increased by 77.7% and 323%, respectively. A thermal protection-power generation integrated system of the present invention adopts the reheater diversion recompression closed Brayton cycle power configuration. Compared with the traditional basic closed Brayton cycle configuration, it has good mass power density, and the mass flow power density is always the largest, that is, the performance is particularly outstanding in terms of power generation and reducing cooling water consumption.

[0089] The present invention aims to propose an integrated system for efficiently utilizing aerodynamic heat to achieve thermal protection and power generation. Through a transpiring cooling nose cone structure combined with gradient cooling, aerodynamic heat recovery is achieved and converted into electrical energy to supply internal equipment of the aircraft, thereby reducing battery power consumption. By converting part of the heat into electricity and inputting cooling water after heat exchange in the cooler for transpiring cooling, the required cooling water consumption is reduced, and the aircraft has greater potential when it has long-duration operation requirements. With the joint participation of optimized cycle power configuration and heat exchange channels, the performance of the integrated system is improved.

[0090] The above description of the disclosed embodiments enables one skilled in the art to implement or use the present invention. Various modifications to these embodiments will be apparent to one skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the scope of the present invention. Therefore, the present invention will not be limited to the embodiments shown herein, but rather to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. An integrated system of active thermal protection and thermal power generation based on supercritical Brayton cycle, characterized in that: The system includes a thermal protection module, a power generation module, a coolant tank and connecting pipes; The thermal protection module includes a gradient sweating cooling structure and a convection cooling heat exchange channel, wherein the gradient sweating cooling structure is arranged at the leading edge of the nose cone of the high-speed aircraft, and the convection cooling heat exchange channel is arranged on the wall of the aircraft cabin; the circulating working fluid in the convection cooling heat exchange channel absorbs the heat from the aircraft wall and then passes into the power generation module; The power generation module is arranged in the cabin section of the high-speed aircraft, and the circulating working fluid converts the absorbed heat into electricity in the power generation module; the power generation module adopts a closed Brayton cycle power configuration with regenerator shunt recompression; The coolant tank is connected to the power generation module and the thermal protection module in sequence through the connecting pipeline. The coolant enters the power generation module through the coolant tank for heat exchange, and after being heated, enters the gradient sweating cooling structure in the thermal protection module to cool the extreme heat flow at the leading edge of the nose cone of the high-speed aircraft.

2. According to claim 1, an active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle is characterized in that: The power generation module adopts a closed Brayton cycle power configuration with regenerator splitting and recompression, including: a turbine, a generator, a cooler, a high-temperature regenerator, a low-temperature regenerator, a main compressor, a recompressor, a booster pump, and a control valve.

3. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 2 is characterized in that: The circulating working fluid absorbs the heat from the surface of the aircraft cabin through the convection cooling heat exchange channel to expand due to overheating, and then enters the turbine to link with the engine to generate electricity; the circulating working fluid after doing work enters the high and low temperature regenerators for heat exchange respectively, and is divided into a mainstream and a secondary flow at the hot end outlet of the low temperature regenerator. The mainstream part enters the cooler for heat exchange with the coolant, then enters the main compressor for pressurization, and then enters the low temperature regenerator for heat exchange. The secondary flow part enters the re-compressor for pressurization, then merges with the mainstream part at the cold end outlet of the low temperature regenerator, enters the high temperature regenerator for heat exchange, and finally enters the heat exchange channel to start a new round of circulation.

4. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 1 is characterized in that: The thermal protection module is a gradient sweating cooling structure, including: a large-porosity porous matrix section, a medium-porosity porous matrix section, a small-porosity porous matrix section, a porous matrix inner layer, a tapered delivery pipe, a permeation layer, a diversion impact hole and a water storage tank; the gradient sweating cooling structure is a porous metal matrix nose cone cooling structure, which is filled with porous metal with a graded porosity gradient from the top to the bottom of the nose cone front edge, and the coolant penetrates through the matrix layers of different porosities to each temperature interval for sweating cooling.

5. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 4 is characterized in that: The porosity of the porous matrix segment with large porosity is x1, the porosity of the porous matrix segment with medium porosity is x2, the porosity of the porous matrix segment with small porosity is x3, the porosity of the permeation layer is x4, and x1=x4>x2>x3.

6. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 1 is characterized in that: The convection cooling heat exchange channel adopts a strong disturbance low pressure loss wing rib heat exchange channel, and a smooth wing rib array structure is added in the channel. The working fluid flows through the wing rib array structure to enhance heat exchange and absorb heat from the surface of the aircraft cabin.

7. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 6, characterized in that: The high-disturbance low-pressure-loss wing-rib heat exchange channel comprises a microchannel inlet, a microchannel outlet and a wing-rib structure. The wing-rib structure is staggeredly arranged in the heat exchange channel to form a wing-rib array.

8. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 1 is characterized in that: The closed Brayton cycle power configuration with regenerator split flow and recompression can use circulating working fluids including supercritical carbon dioxide (sCO2) and helium (He).

9. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 2, characterized in that: The closed Brayton cycle power configuration with regenerator split recompression has a circulating working fluid split ratio at the outlet of the low-temperature regenerator in the range of 0.2 to 0.

4.

10. The active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle according to claim 3, characterized in that: The flow splitting impact holes are evenly distributed along the wall of the tapered conveying pipe, and the apertures of the flow splitting impact holes are in the order of millimeters.

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