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

Abstract

The application provides an active thermal protection and heat power generation integrated system based on a supercritical Brayton cycle, which combines convection cooling and sweating cooling, and improves energy utilization; and relates to a high-speed aircraft head cone front edge gradient cooling structure, the head cone front edge is divided into a top end, a middle part and a low end, the heat flow of the three areas decreases in turn, and the porous metal matrix layer is arranged in the three sections of the head cone front edge in gradient according to the size of porosity; a two-stage regenerative-shunt recompression closed Brayton cycle power configuration is adopted, which has the characteristics of high energy recovery efficiency of the closed Brayton cycle and compact system structure; and a wing rib array is arranged in the heat exchange channel, so that more heat exchange surface area is provided.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of aerospace vehicle thermal protection and thermal utilization, and particularly relates to a thermal protection-power generation integrated system. BACKGROUND

[0002] During the flight of a high-speed vehicle in the atmosphere, aerodynamic heating caused by friction with the incoming atmosphere will rapidly increase the temperature of the outer surface of the vehicle, resulting in a decrease in the strength and rigidity of the vehicle structure. For a high-speed vehicle that needs to operate for a long time, its outer surface is in a high-temperature environment for a long time, and the heat flux is large. In order to ensure normal operation, the thermal protection requirement is very high.

[0003] Traditional aircraft thermal protection system design usually removes heat by ablation or uses the properties of the material itself to resist heat, which has poor repeatability and sustainability and does not meet the long-time operation requirement of the aircraft. Although the use of regenerative cooling, gas film cooling and sweating cooling and other active thermal protection technologies is more effective, the use of a single active thermal protection method often requires a large amount of cooling medium, which inevitably leads to excessive mass of the high-speed vehicle, and the requirement for the power supply carried by the aircraft for long-time operation is increased, resulting in large loss. It is an effective way to solve the problem of thermal protection of high-speed vehicles under high heat load to adopt a targeted and combined method of multiple thermal protection modes for different key parts of the aircraft, and to flexibly and efficiently utilize the advantages of different thermal protection technologies. In addition, the high-temperature thermal energy generated by aerodynamic heating is a high-grade energy. By transferring the high temperature of aerodynamic heat to the outside through a cooling medium, this part of energy cannot be truly utilized. In the traditional active cooling method, only the convection cooling can effectively utilize this part of energy to convert thermal energy into electrical energy, replace the power supply carried by the aircraft, and reduce the consumption of cooling liquid, but it is necessary to additionally increase the circulating medium and heat exchange equipment, thereby increasing the weight of the aircraft. Therefore, it is necessary to establish a mathematical model of such an aerodynamic heat thermoelectric conversion cycle system, describe different functions and corresponding index parameters, and optimize the cycle configuration to obtain the most suitable cycle configuration.

[0004] The convective cooling thermoelectric conversion technology using aerodynamic heat can meet the cooling demand of most areas of the engine and the fuselage of the high-speed aircraft, but it is difficult to effectively cool the super-high-temperature components of the nose cone leading edge part subjected to severe aerodynamic heating. The cooling liquid for cooling the circulating working medium in the above-mentioned thermoelectric conversion system is delivered to the nose cone leading edge part of the aircraft to form an open loop, and the cooling liquid is evaporated and discharged to the outside space in the form of sweating cooling after absorbing the extreme temperature of the nose cone leading edge, and a gas film is formed on the surface layer of the nose cone leading edge, which can effectively reduce the high-temperature heat flow received by the nose cone leading edge part of the aircraft. The pyrolyzable material currently filled and used cannot be repeatedly applied, and it is also impossible to control the distribution of the cooling liquid according to the uneven thermal load of the nose cone leading edge, which may cause insufficient cooling liquid in part of the area and excessive cooling liquid in another part of the area, and the cooling liquid cannot be discharged in time, resulting in heat transfer deterioration and excessive consumption of the cooling liquid. Changing the sweating cooling structure of the nose cone leading edge and reasonably distributing the cooling liquid are effective means to improve the cooling efficiency and ensure the safe operation of the nose cone part of the aircraft.

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

[0006] The purpose of the present application is to solve the problems existing in the prior art, and the present application provides a thermal protection-power generation integrated system, which adopts different cooling methods according to different parts of the aircraft. The present application provides an active thermal protection and thermal power generation integrated system based on supercritical Brayton cycle, which combines convective cooling and sweating cooling two active cooling methods. The aerodynamic heat on the surface layer of the aircraft is used for thermoelectric conversion through the convective cooling method, so as to realize thermal protection and replace part of the power carried by the aircraft. At the same time, since part of the heat is used for expansion work and converted into electric quantity, the amount of heat that needs to be cooled by the cooling liquid is reduced, and the consumption of the cooling liquid required is reduced. The high-temperature cooling liquid after absorbing the exhaust heat is delivered to the gradient cooling structure of the nose cone leading edge of the aircraft through the pipeline, and the extreme high-temperature part of the nose cone is cooled through the sweating cooling method, thereby improving the energy utilization rate.

[0007] The present application adopts a gradient cooling nose cone leading edge cooling structure, the nose cone leading edge is divided into a top end, a middle part and a low end, the heat flow of the three regions decreases in turn, and the porous metal matrix layer is arranged in the three sections of the nose cone leading edge in gradient according to the porosity. The pore diameter of the gradient arranged porous metal matrix layer corresponds to the heat flow, avoids the blockage of the excess cooling liquid in the pores, better carries away the heat, reduces the sweating cooling liquid consumption, and is more conducive to the long-time operation of the high-speed aircraft.

[0008] In order to improve the thermal efficiency and power of the circulating system, and reduce the heat loss in the circulating process. The closed Brayton cycle is applied to the recycling of the aerodynamic heat of the high-speed aircraft, and the regenerator shunt recompression closed Brayton cycle power configuration obtained by calculation optimization is adopted, which has the characteristics of high energy recovery efficiency of the closed Brayton cycle and compact system structure, so that the collection and utilization of the aerodynamic heat energy are maximized under the premise of ensuring the heat protection of the aircraft cabin section, and the energy loss is reduced.

[0009] In addition, the wing rib array structure microchannel with strong disturbance effect and low pressure loss is used as the heat exchange channel, the circulating working medium flows through the heat exchange channel and exchanges heat with the high-temperature wall surface of the cabin section of the aircraft by convection, and due to the existence of the wing rib array in the microchannel, more heat exchange surface area is provided, and the working medium is disturbed when flowing in the channel, so that the heat exchange performance is improved; meanwhile, due to the streamline structure of the wing rib surface, the pressure loss of the working medium flowing through the wing rib array is relatively small, which greatly reduces the negative effect of the pressure loss caused by adding the disturbance body in the channel.

[0010] The application provides a heat protection-power generation integrated system, which comprises a heat protection module, a power generation module, a cooling liquid tank and a connecting pipeline.

[0011] Further, the heat protection module comprises 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 head cone of the high-speed aircraft, and the convection cooling heat exchange channel is arranged on the wall surface of the cabin section of the aircraft; the circulating working medium in the convection heat exchange channel absorbs heat from the wall surface of the aircraft and then enters the power generation module; the power generation module is arranged in the cabin section of the high-speed aircraft, and the circulating working medium converts the absorbed heat into electric energy in the power generation module; the cooling liquid tank is connected with the power generation module and the heat protection module in sequence through the connecting pipeline, the cooling liquid enters the power generation module through the cooling liquid tank for heat exchange, is heated and then enters the gradient sweating cooling structure of the heat protection module to cool the extreme heat flow of the leading edge of the head cone of the high-speed aircraft.

[0012] The power generation module adopts a closed Brayton cycle power configuration with regenerator diversion and recompression, including: turbine, generator, cooler, high-temperature regenerator, low-temperature regenerator, main compressor, recompressor, pressurization pump, and control valve; the circulating working fluid absorbs heat from the surface of the aircraft compartment through the convection cooling heat exchange channel and undergoes superheated expansion, then enters the turbine and drives the engine to generate electricity; the working fluid after doing work enters the high-temperature and low-temperature regenerators for heat exchange, and at the hot end outlet of the low-temperature regenerator, it is diverted into a main stream and a secondary stream. The main stream enters the cooler to exchange heat with the coolant, then enters the main compressor for pressurization, and then enters the low-temperature regenerator for heat exchange. The secondary stream enters the recompressor for pressurization, then merges with the main stream 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 incorporates a gradient sweating cooling structure, including: a high-porosity porous matrix section, a medium-porosity porous matrix section, a low-porosity porous matrix section, a porous matrix inner layer, a tapered delivery pipe, a diversion impact hole, and a water storage tank; the porous metal matrix head cone cooling structure is filled with porous metal with a graded decrease in porosity from the top of the head cone to the bottom, and the coolant permeates through the matrix layers with different porosities to each temperature range for sweating cooling;

[0014] The convective cooling heat exchange channel adopts a strong disturbance and low pressure loss rib heat exchange channel. A rib array structure is arranged in the channel. The working fluid flows through the rib array disturbance structure to enhance heat exchange and absorb the heat on the surface of the aircraft compartment.

[0015] As a further improvement to the technical solution of the present invention, the porous matrix with large, medium and small porosities in the gradient sweating cooling structure is respectively gradient distributed and filled at the top, middle and bottom of the gradient sweating cooling structure. For the bottom region with lower heat flux, the small porosity porous matrix section reduces the consumption of coolant and the accumulation of coolant in the pores by controlling the sweating cooling rate of the coolant. The coolant is transported to the cooling chamber and accelerated to the inner surface of the porous matrix at the top of the gradient sweating cooling structure via the tapered delivery pipe. At the same time, the coolant is uniformly and rapidly transported to the leading edge of the head cone of the gradient sweating cooling structure via the diversion impact hole.

[0016] As a further improvement to 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 inside the channel, 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, suppresses the flow pressure loss, and suppresses the pump power consumption.

[0017] Further, as an improvement of the technical scheme of the present application, the circulating working medium that can be used in the regenerator split-flow recompression closed Brayton cycle power configuration includes but is not limited to supercritical carbon dioxide (sCO2) and helium (He);

[0018] Further, as an improvement of the technical scheme of the present application, the recommended split-flow ratio of the circulating working medium at the outlet of the low-temperature regenerator of the regenerator split-flow recompression closed Brayton cycle power configuration ranges from 0.2 to 0.4.

[0019] Further, as an improvement of the technical scheme of the present application, the porosity of the large-porosity porous base section is x1, the porosity of the medium-porosity porous base section is x2, the porosity of the small-porosity porous base section is x3, and the porosity of the permeation layer is x4, and x1=x4>x2>x3.

[0020] Further, as an improvement of the technical scheme of the present application, the split-flow impact holes are uniformly distributed along the wall of the tapered delivery pipe, and the diameter of the split-flow impact holes is in the order of millimeters.

[0021] The present application has the following beneficial effects:

[0022] 1. The present application relates to a heat protection and thermoelectric conversion integrated system coupled with a convection cooling closed system and a sweating cooling open system; two active cooling methods of convection cooling and sweating cooling are combined. The convection cooling method is used to absorb the aerodynamic heat of the surface layer of a high-speed aircraft and convert the heat into electricity, so that the heat protection is realized, and the battery power carried by the aircraft is reduced or even replaced; 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 cooling liquid is reduced, and the consumption of the cooling liquid is reduced; the high-temperature cooling liquid after absorbing the heat of the exhaust steam is delivered to the gradient cooling structure of the leading edge of the aircraft nose cone through a pipeline, and the gradient cooling structure is used to cool the extremely high-temperature part of the nose cone through the sweating cooling method, thereby improving the energy utilization rate.

[0023] 2. The present application relates to a gradient cooling structure of the leading edge of a high-speed aircraft nose cone; the leading edge of the nose cone is divided into a top end, a middle part and a low end, the heat flux of the three regions decreases in turn, and the porous metal base layers are arranged in the three regions of the leading edge of the nose cone in gradient according to the porosity; the pore diameter of the gradient arranged porous metal base layer corresponds to the heat flux, the small-porosity porous base section controls the sweating cooling rate of the cooling liquid, reduces the consumption of the cooling liquid and the accumulation of the cooling liquid in the pores, reduces the consumption of the sweating cooling liquid, and is more conducive to the long-time operation of the high-speed aircraft; the tapered delivery pipe is preferentially delivered to the inner surface of the porous base layer at the top end of the gradient cooling leading edge of the nose cone, and the split-flow impact holes on the tapered delivery pipe uniformly and rapidly deliver the cooling liquid to other sections of the gradient cooling leading edge of the nose cone, so that the heat is better taken away.

[0024] 3.The application relates to a closed Brayton cycle power configuration with two-stage regenerative split recompression; the two-stage regenerative split recompression closed Brayton cycle power configuration is obtained through calculation optimization, and has the characteristics of high energy recovery efficiency of the closed Brayton cycle and compact system structure, improves the thermal efficiency and power capacity of the cycle system under the premise of guaranteeing the thermal protection of the cabin section of the aircraft, maximizes the collection and utilization of aerodynamic heat, and reduces energy loss.

[0025] 4.The application relates to a heat exchange microchannel with a wing rib array; a wing rib array is arranged in the heat exchange channel through micro-surface processing design, the circulating working medium flows through the heat exchange channel and exchanges heat with the high-temperature wall surface of the cabin section of the aircraft in a convection mode, the wing rib array in the microchannel provides more heat exchange surface area, and the working medium generates strong disturbance when flowing in the channel, so that the heat exchange performance is improved, more heat is absorbed by the circulating working medium, the wall surface temperature of the aircraft is ensured to be in a safe range, and the power generation capacity of the thermoelectric conversion subsystem is improved; meanwhile, the wing rib surface is designed in a streamlined structure, the pressure loss of the working medium flowing through the wing rib array is relatively small, the pressure loss caused by adding the disturbance structure in the heat exchange channel is reduced, and the negative effect of the increase of pump power consumption is inhibited. BRIEF DESCRIPTION OF DRAWINGS

[0026] The accompanying drawings, which form a part of this application, are included to provide a further understanding of the application, and are incorporated in and constitute a part of this application, illustrate embodiments of the present application and serve to explain the present application, and do not limit the present application in any manner:

[0027] Figure 1 is a schematic diagram of a heat protection-power generation integrated system according to an embodiment of the application;

[0028] Figure 2 is a schematic diagram of a gradient-distributed porous metal matrix nose cone cooling structure according to an embodiment of the application;

[0029] Figure 3 is a schematic diagram of a strong disturbance and low pressure loss wing rib heat exchange channel structure according to an embodiment of the application;

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

[0031] Figure 5 is a schematic diagram of a regenerative Brayton cycle power configuration (R-BC) according to an embodiment of the application

[0032] Figure 6 is a schematic diagram of a regenerator split recompression Brayton cycle power configuration (RR-BC) according to an embodiment of the application;

[0033] Figure 7is the mass flow power density of the three power configuration layouts of the application varying with the compressor pressure ratio;

[0034] Figure 8 is the mass flow power density of the three power configuration layouts of the application varying with the compressor pressure ratio;

[0035] Figure 9 is the total mass of the three power configuration layouts of the application varying with the compressor pressure ratio.

[0036] 1. Circulating working medium pump; 2. Head cone front edge gradient cooling structure; 3. Strong disturbance low pressure loss wing rib heat exchange channel; 4. Heat exchange channel guide opening; 5. Turbine; 6. Generator; 7. High temperature regenerator; 8. Low temperature regenerator; 9. Cooler; 10. Compressor; 11. Re-compressor; 12. Control valve; 13. Pressurizing pump; 14. Cooling liquid tank; 15. Aircraft cabin section wall surface; 21. Large porosity porous base section; 22. Medium porosity porous base section; 23. Small porosity porous base section; 24. Liquid storage cabin; 25. Permeation layer; 26. Shrinkage conveying pipe; 27. Split impact hole; 31. Micro-channel inlet; 32. Micro-channel outlet; 33. Wing rib structure. DETAILED DESCRIPTION

[0037] In order to make the personnel in the art better understand the application scheme, the technical solutions in the embodiments of the application will be described clearly and completely in conjunction with the drawings in the embodiments of the application. Obviously, the described embodiments are only a part of the embodiments of the application, not all the embodiments. Based on the embodiments in the application, all other embodiments obtained by those skilled in the art without creative labor should belong to the scope of protection of the application.

[0038] The embodiment provides a heat protection-power generation integrated system, and a component schematic diagram thereof is as shown in Figure 1As shown, the aircraft nose cone leading edge gradient cooling structure 2 and the aircraft cabin aerodynamic heat thermoelectric conversion subsystem; thermoelectric conversion subsystem includes circulating working medium pump 1, wing rib heat exchange channel 3, heat exchange channel guide 4, turbine 5, generator 6, high temperature regenerator 7, low temperature regenerator 8, cooler 9, compressor 10, re-compressor 11, control valve 12, pressurizing pump 13 and cooling liquid tank 14; the specific working process is: under the action of circulating working medium pump 1, the circulating working medium enters the strong disturbance low pressure loss wing rib heat exchange channel 3 through the heat exchange channel guide 4, exchanges heat with the high temperature wall surface of the high speed aircraft cabin, and flows into the turbine 5 after absorbing heat through the heat exchange channel guide 4, the working medium expands and does work, the generator 6 is linked to generate electricity, the spent steam after work is introduced into the high temperature regenerator 7 and the low temperature regenerator 8 to exchange heat, the working medium is divided into two parts, the main stream is introduced into the cooler 9, the cooling liquid from the cooling liquid tank 14 is pumped into the cooler 9 by the working medium pump 13 and the pressurizing pump to exchange heat with the main stream working medium, and then is transported to the aircraft nose cone leading edge gradient cooling structure 2 after absorbing the heat of the main stream working medium, the main stream working medium flows into the compressor 10 from the cooler 9 to be compressed, and then enters the low temperature regenerator 8 to exchange heat with the working medium that has not been divided, the secondary flow working medium is directly introduced into the re-compressor 11 from the outlet of the low temperature regenerator 8 to be pressurized, and then is merged with the main stream working medium to enter the high temperature regenerator 7 to absorb heat and be heated, and finally starts a new cycle under the action of the circulating working medium pump 1.

[0039] As shown in the figure, Figure 2 The aircraft nose cone leading edge gradient cooling structure in the application includes a high porosity porous matrix segment 21, a medium porosity porous matrix segment 22, a small porosity porous matrix segment 23, a permeation layer 25, a tapered delivery pipe 26 and a flow dividing impact hole 27; according to the different heat flow from the top end to the bottom end of the leading edge structure, the high porosity porous matrix segment 21, the medium porosity porous matrix segment 22 and the small porosity porous matrix segment 23 are used in turn to take away the extreme heat flow of the nose cone leading edge in the form of perspiration cooling, the high porosity porous matrix segment 21 is beneficial to timely supplement of cooling liquid to ensure the safety of the top end of the nose cone leading edge, the medium and small porosity porous matrix segments 22 and 23 control the size of the pores to avoid blockage of the excess cooling liquid in the pores, better take away the heat, reduce the consumption of perspiration cooling liquid and are more beneficial to long time operation of the high speed aircraft.

[0040] As shown in the figure, Figure 3 The strong disturbance low pressure loss heat exchange channel in the application includes a micro-channel inlet 31, a micro-channel outlet 32 and a wing rib structure 33; the wing rib structure 33 is arranged in the heat exchange channel in a staggered manner to form a wing rib array, which significantly increases the heat exchange area, the circulating working medium flows into the heat exchanger through the micro-channel inlet 31, generates strong disturbance when flowing through the wing rib array, strengthens the heat exchange effect between the working medium and the high temperature cabin wall, and at the same time, the streamline structure of the wing rib surface suppresses the pressure loss when the working medium flows through, reduces the power consumption required by the circulating working medium pump, and finally flows out through the micro-channel outlet 32.

[0041] Preferred embodiments

[0042] The following embodiments are provided using the recuperator split recompression power configuration and the comparative configuration in the present application:

[0043] The cooling performance and mass performance of the recuperator split recompression power configuration (RR-BC) of the aerothermodynamic heat to electricity conversion subsystem are studied in comparison with the recuperation configuration (R-BC) and the basic configuration (S-BC), as shown in the schematic diagrams of the three configurations in Figures 4-6 The turbine inlet temperature (the outlet temperature of the working medium in the heat exchange channel) is given as 673 K, the compressor inlet pressure is 7800 kPa, the compressor inlet temperature is 315 K, the compressor pressure ratio is 2.1-2.5, the pinch point temperatures of the high-temperature and low-temperature recuperators are 10 K, the turbine isentropic efficiency is 0.9, the compressor and recompressor isentropic efficiencies are 0.89, the cooling water inlet temperature is 293 K, the inlet pressure is 200 kPa, and the cooling water outlet temperature is given as 353 K for the convenience of comparison. Based on the closed Brayton cycle energy analysis equations and the thermodynamic calculation equations, the mass power density and the mass flow power density of the aerothermodynamic heat to electricity 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 aerothermodynamic heat to electricity conversion subsystem, a thermodynamic model of the heat to electricity conversion subsystem needs to be established for parameter discussion and analysis. The basic energy equations of the working medium in the cycle are shown in Table 1.

[0046] Table 1 Basic energy equations of the working medium in the cycle

[0047] Main components heat load and power consumption Thermodynamic equations Heat exchanger heat exchange amount Q = m h C p,h (T h,o -T h,i ) = m h C p,h (T c,o -T c,i ) = UAΔT Turbo shaft work power W t = m w ·(h tout -h tin ) <!-- 5 -->]]> Compressor power consumption W c = m w ·(h cout -h cin )]]> Net power generation [CDATAW net = W t - W c ]]>

[0048] In the conventional thermodynamic cycle power generation process operating on the ground, the cooling water is generally considered to be unlimited and is not used as a system evaluation index. However, for a high-speed aircraft, the cooling water is the only cold source for maintaining the temperature of the aircraft, and the cooling water consumption is directly related to the mass load of the long-endurance high-speed aircraft system. In order to objectively evaluate the power generation per unit mass flow of cooling water of different Brayton power configurations under different working conditions, the mass flow power density is introduced as an evaluation index, which is defined as:

[0049]

[0050] Where mf is the cooling water mass flow required by the heat to electricity conversion subsystem.

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

[0052]

[0053] where Q total is the total heat absorbed by the cycle working fluid through the heat transfer channels, C f is the specific heat capacity of the cooling water, T f0 and T f1 are the inlet and outlet temperatures of the cooling water in the cooler; η is the net power of the thermoelectric conversion subsystem.

[0054] The net power of the basic configuration, the regenerative configuration and the regenerator split recompression configuration are defined as follows, respectively:

[0055]

[0056]

[0057]

[0058] where h is the enthalpy 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 is proposed based on the thermodynamic model, and the total mass of the power configurations used is calculated.

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

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

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

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

[0065] The heat transfer capacity (UA) and the pressure loss (Δp) are calculated by the thermodynamic model

[0066] Q = UAΔT

[0067] Δp=P i -P o

[0068] In the formula, ΔT is the logarithmic mean temperature difference, and P i P is the fluid inlet pressure. o This is the fluid outlet pressure.

[0069] The model is input with known parameters such as UA, Δp, working fluid flow rate mw, logarithmic mean temperature difference of the heat exchanger ΔT, and wall temperature Tw. Based on the current state, the thermal properties of the fluid are calculated: thermal conductivity (λ), density (ρ), dynamic viscosity (μ), and specific heat capacity at constant pressure (Cp). The number of circular tubes (n) and the length of the circular tubes (L) are linearly divided to form a two-dimensional matrix of n and L. Based on the aforementioned fluid thermal properties, the Reynolds number kinematic viscosity (v) and Reynolds number (R) are calculated using the formula. e ) and coefficient of friction (Cf):

[0070]

[0071]

[0072]

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

[0074] The corresponding approximate pressure loss matrix (Δpx) and approximate heat transfer capacity matrix (UAx) are calculated using the formulas. The approximate pressure loss matrix (Δpx) is defined as follows:

[0075]

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

[0077] UA x =h c A s

[0078] In the formula, hc is the heat transfer coefficient within the heat exchanger. As is the heat transfer area, defined as follows:

[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 of the heat exchanger fluid and the outlet wall temperature, respectively, and for the cold end of the heat exchanger, Ti and Tw are the inlet temperature of the heat exchanger fluid and the inlet wall temperature, 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 value calculated by thermodynamic calculation are respectively interpolated and matched with the Δpx matrix and the UAx matrix. When the values of Δp and UA are each within the numerical range of the Δpx matrix and the UAx matrix, the values closest to Δp and UA in the two matrices are selected as the actual values, and the values of the specific number of circular tubes (n) and the length (L) are returned. The approximate volume Vx of the heat exchanger is calculated from the specific n and L values, and is used as an approximate value of the actual heat exchanger volume V. The heat transfer capacity approximation matrix (UAx) can be defined as:

[0083]

[0084] In the formula, σP is the porosity of the material used for the heat exchanger.

[0085] The approximate volume Vx is multiplied by the density ρ of the material used for the heat exchanger to obtain the mass M of the heat exchanger. To further measure the relationship between the power generation of the system and the additional mass of the power configuration, the mass power density is further proposed as an index, which is defined as:

[0086] W s = W net / M total

[0087] A larger mass power density means that the mass load of the power configuration is more efficiently utilized, i.e., the benefits generated by the introduction of the aero-thermal thermoelectric conversion subsystem are higher.

[0088] As shown in Figure 7 and Figure 8 : In the aero-thermal thermoelectric conversion subsystem, the maximum mass flow power densities of the regenerative configuration (R-BC) and the regenerator split 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 densities of the regenerative configuration and the regenerator split recompression configuration are 42.2 W / kg and 31.5 W / kg, respectively, which are about 24.8% higher and about 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 recompression configuration has a more complex configuration and a larger total mass of components, such as Figure 9As shown, the result is slightly lower maximum mass flow than the basic configuration, but its maximum mass flow power density is nearly three times that of the basic configuration, and the calculated optimal range of split ratio is in the range of 0.2-0.4. When the compressor pressure ratio π is 2.2, the mass power densities of the basic, regenerative, and regenerative split recompression configurations are similar, being 28.6 W / kg, 28.3 W / kg, and 27.9 W / kg, respectively, but the mass flow power densities are 34.5 kJ / kg, 61.3 kJ / kg, and 146 kJ / kg, respectively. Compared with the basic configuration, the mass flow power densities of the regenerative configuration and the regenerative split recompression configuration are increased by 77.7% and 323%, respectively. The heat protection-power generation integrated system of the present application adopts the regenerative split recompression closed Brayton cycle power configuration, and compared with the traditional basic closed Brayton cycle configuration, has good mass power density, the mass flow power density is always the largest, that is, the performance is particularly outstanding in terms of power generation and reduction of cooling water consumption.

[0089] The present application aims to provide an integrated system for efficiently utilizing aerodynamic heat to achieve heat protection and power generation, which recovers and utilizes aerodynamic heat through a gradient-cooled sweat-cooling nose cone structure, converts the aerodynamic heat into electrical energy to supply internal devices of the aircraft, and reduces battery power consumption; by converting part of the heat into electrical energy and inputting the cooled water after heat exchange in the cooler for sweat cooling, the required cooling water consumption is reduced, and the aircraft has more potential when it has long endurance operation requirements; under the joint participation of the optimized cycle power configuration and the heat exchange channel, the performance of the integrated system is improved.

[0090] The above description of disclosed embodiments enables one of ordinary skill in the art to make or use the application. Various modifications to these embodiments will be apparent to those skilled in the art, and the generic principles defined herein can be applied to other embodiments without departing from the scope of the application. Thus, the present application is not intended to be limited to the embodiments shown herein but is to be accorded 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 comprises a thermal protection module, a power generation module, a cooling liquid tank and a connecting pipeline; The thermal protection module comprises 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 head cone of the high-speed aircraft, and the convection cooling heat exchange channel is arranged on the wall surface of the cabin section of the aircraft; the circulating working medium in the convection cooling heat exchange channel absorbs the heat of the wall surface of the aircraft and is then introduced into the power generation module; The power generation module is arranged in the cabin section of the high-speed aircraft, and the circulating working medium converts the absorbed heat into electric energy in the power generation module; the power generation module adopts a closed Brayton cycle power configuration with a regenerator, a shunt and a recompression; The cooling liquid tank is connected with the power generation module and the thermal protection module in sequence through the connecting pipeline, the cooling liquid enters the power generation module to exchange heat through the cooling liquid tank, is heated and then enters the gradient sweating cooling structure of the thermal protection module to cool the extreme heat flow of the leading edge of the head cone of the high-speed aircraft; The power generation module adopts a closed Brayton cycle power configuration with a regenerator, a shunt and a recompression, and comprises a turbine, a generator, a cooler, a high-temperature regenerator, a low-temperature regenerator, a main compressor, a recompression compressor, a pressurizing pump and a control valve. The circulating working medium absorbs the heat of the surface layer of the cabin section of the aircraft through the convection cooling heat exchange channel, is overheated and expanded, is introduced into the turbine and drives the generator to generate electricity, and the working medium after work is introduced into the high-temperature regenerator and the low-temperature regenerator to exchange heat, is divided into a main flow and a secondary flow at the hot end outlet of the low-temperature regenerator, is introduced into the cooler to exchange heat with the cooling liquid after the main flow is introduced into the main compressor to be pressurized, is introduced into the low-temperature regenerator to exchange heat again, the secondary flow is introduced into the recompression compressor to be pressurized, is combined with the main flow at the cold end outlet of the low-temperature regenerator, is introduced into the high-temperature regenerator to exchange heat and is finally introduced into the convection cooling heat exchange channel to start a new cycle; The thermal protection module is a gradient sweating cooling structure, which comprises a large-porosity porous base section, a medium-porosity porous base section, a small-porosity porous base section, a porous base inner layer, a tapered delivery pipe, a permeation layer, a shunt impact hole and a water storage cabin; the gradient sweating cooling structure is a porous metal base head cone cooling structure, the head cone is filled with the porous metal with a porosity gradient and a hierarchical decrease from the top end to the low end, and the cooling liquid is permeated through the base layers with different porosities to sweat and cool in different temperature intervals.

2. The integrated system of active thermal protection and thermal power generation based on supercritical Brayton cycle according to claim 1, characterized in that, The porosity of the large-porosity porous base section is x1, the porosity of the medium-porosity porous base section is x2, the porosity of the small-porosity porous base section is x3, the porosity of the permeation layer is x4, and x1=x4>x2>x3.

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

4. The integrated system of active thermal protection and thermal power generation based on supercritical Brayton cycle according to claim 3, characterized in that, The strong disturbance and low pressure loss wing rib heat exchange channel comprises a micro-channel inlet, a micro-channel outlet and a wing rib structure, the wing rib structure is arranged in the heat exchange channel in a staggered manner to form a wing rib array.

5. The integrated system of active thermal protection and thermal power generation based on supercritical Brayton cycle according to claim 1, characterized in that, The circulating working medium of the closed Brayton cycle power configuration with a regenerator, a shunt and a recompression can comprise supercritical carbon dioxide (sCO2) and helium (He).

6. The integrated system of active thermal protection and thermal power generation based on supercritical Brayton cycle according to claim 1, characterized in that, The split ratio of the cycle working medium at the outlet of the low-temperature regenerator of the closed Brayton cycle power configuration of the regenerative split recompression is 0.2-0.

4.

7. The integrated system of active thermal protection and thermal power generation based on supercritical Brayton cycle according to claim 1, characterized in that, The split impact holes are uniformly distributed along the wall of the tapered delivery pipe, and the diameter of the split impact holes is in millimeter level.

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