High-speed aircraft thermal environment calculation method

Abstract

The invention provides a thermal environment calculation method for a high-speed aircraft, and the method comprises the steps: respectively calculating the radiation equilibrium temperature of the bottom of a tail cabin in two states: the tail nozzle of an engine is not extended and extended to the maximum position, and taking out the maximum value; and taking the reference temperature, the maximum radiation equilibrium temperature and a plurality of temperature values between the reference temperature and the maximum radiation equilibrium temperature as a hot wall temperature data set for bottom hot wall heat flow calculation. Through interpolation of heat flow of the heat wall at the bottom of the stern room in two states on time and actual temperature of the bottom of the stern room, the heat flow value of the heat wall in the elongation process of the exhaust nozzle is obtained, so that the influence of elongation of the exhaust nozzle on the thermal environment of the bottom of the stern room is considered during heat transfer calculation of the stern room. Typical moments of aerodynamic heat and exhaust nozzle inner wall temperature and a hot wall temperature value of a tail cabin bottom hot wall heat flow are combined into a feature point data set, preprocessing of three different boundary conditions is achieved, real-time numerical values of the three different boundary conditions are obtained through an interpolation function during heat transfer calculation, the boundary condition processing complexity is reduced, and the calculation efficiency is improved. And the calculation efficiency is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of aerodynamic heat and heat transfer, and in particular relates to a method for calculating the thermal environment of a high-speed aircraft. Background Art

[0002] The external heating effect on the tail cabin of a high-speed aircraft comes from three aspects: aerodynamic heating of the outer surface of the aircraft, convection heating of the bottom of the tail cabin, and heating of the high-temperature tail nozzle of the aircraft. The high-speed incoming flow produces severe aerodynamic heating on the outer surface of the aircraft, causing the temperature of the aircraft structure to rise. The bottom of the tail cabin of the aircraft is in the leeward area and is not directly affected by the high-speed incoming flow, but the pressure in the bottom area of ​​the tail cabin is relatively low. The high-temperature and high-speed gas flowing along the outer surface of the aircraft flows to the low-pressure area at the bottom of the tail cabin under the drive of pressure. At the same time, the tail jet of the engine has a certain blowing effect on the bottom of the tail cabin, and the two together cause convection heating of the bottom of the tail cabin. In addition, the high-temperature tail nozzle of the engine transfers heat to the tail cabin.

[0003] The prior art uses aerodynamic heating of the outer surface of the aircraft, convection heating of the bottom of the tail cabin of the aircraft, and heating of the high-temperature tail nozzle of the aircraft as three boundary conditions for calculating the tail cabin structure and the overall thermal environment in the cabin. The three boundary conditions are obtained by calculation or experiment respectively: by numerical simulation of the external flow field, the cold wall heat flux and recovery enthalpy of the outer surface of the aircraft tail cabin are calculated, and the temperature of the outer surface structure of the aircraft tail cabin is extracted in real time along the time course of flight. The aerodynamic heating boundary condition is applied to the outer surface of the tail cabin by converting the cold wall heat flux and the cold and hot wall heat flux of the recovery enthalpy; the hot wall heat flux of the bottom of the tail cabin is calculated by the overall numerical simulation of the external flow field and the tail jet as the boundary condition of the bottom of the tail cabin; the temperature data of the inner wall of the thermal insulation layer of the engine tail nozzle is obtained by numerical simulation of the flow field and heat transfer of the engine internal flow and the engine bearing structure as the heating boundary condition of the high-temperature tail nozzle of the aircraft. The connection position of the aircraft tail cabin and other parts of the aircraft is set as an adiabatic boundary condition. By carrying out numerical simulation of heat transfer of the aircraft's tail cabin structure and the air domain and equipment inside the cabin, we can obtain the tail cabin structure and the overall thermal environment data inside the cabin.

[0004] The prior art has the following defects:

[0005] (1) When calculating the heat flux of the hot wall at the bottom of the tail cabin, it is necessary to set the hot wall temperature at the bottom. The hot wall temperature value used in the prior art is not determined by a clear criterion, but is given by the experience of the calculator, which may lead to unreasonable hot wall temperature values ​​and cause calculation errors.

[0006] (2) The tail nozzle structure of an aircraft has the property of thermal expansion and contraction. After the engine is ignited and started, the tail nozzle of the aircraft is heated and becomes longer. The elongation of the tail nozzle changes the shape of the aircraft, which will cause the degree of the tail jet blowing on the bottom of the tail cabin to change, affecting the size of the heat flow at the bottom of the tail cabin. The existing technology does not take the influence of the elongation of the tail nozzle into consideration.

[0007] (3) The aerodynamic heating boundary conditions of the tail cabin exterior, the heat flux / convection boundary conditions of the bottom hot wall, and the temperature boundary conditions of the tail nozzle inner wall are obtained through different numerical simulation processes, and the corresponding typical moments of the flight trajectory are different. The existing technology does not combine the typical moments corresponding to the three boundaries. When calculating the heat transfer of the tail cabin, it is necessary to interpolate the boundary conditions through three different functions according to the time history moments in the flight trajectory to calculate the actual boundary conditions at the current moment. The data processing process is complicated and the calculation time is long. Summary of the invention

[0008] The present invention aims to solve at least one of the technical problems existing in the prior art.

[0009] The present invention provides a method for calculating the thermal environment of a high-speed aircraft, the method comprising:

[0010] Step 1: Under the reference temperature, by means of joint numerical simulation of the tail cabin outer flow field and the tail jet flow field, the radiation equilibrium temperature of the outer surface of the tail cabin bottom corresponding to the two states, when the tail nozzle is not extended and when the tail nozzle reaches the maximum extension, is calculated respectively;

[0011] Step 2: construct a typical hot wall temperature value data set based on the radiation equilibrium temperature and reference temperature of the outer surface of the tail cabin bottom;

[0012] Step 3: For the two states of the tail nozzle not being extended and the tail nozzle reaching the maximum extension, all typical hot wall temperature values ​​are set in sequence, and the joint simulation of the tail cabin outer flow field and the tail jet is carried out to obtain two sets of hot wall heat flux data at the bottom of the tail cabin corresponding to the two states;

[0013] Step 4: Construct a set of boundary condition feature points. The set of boundary condition feature points includes all typical moments corresponding to aerodynamic heating and the inner wall temperature of the tail nozzle insulation layer, and typical hot wall temperature values ​​corresponding to the hot wall heat flux at the bottom of the tail cabin in two states. Perform linear interpolation processing on each boundary condition data at each feature point to generate a new boundary condition data set.

[0014] Step 5: Interpolate the new boundary condition data set of aerodynamic heating and the temperature of the inner wall of the tail nozzle insulation layer at the time of heat transfer calculation, and interpolate the new boundary condition data set of the tail cabin bottom hot wall heat flow in two states for the tail cabin bottom temperature, and obtain the boundary condition data of aerodynamic heating and the temperature of the inner wall of the tail nozzle insulation layer at the time of heat transfer calculation, and the two-state tail cabin bottom hot wall heat flow data corresponding to the real-time temperature;

[0015] Step 6, interpolating the heat flux data of the hot wall at the bottom of the tail cabin in two states at the time of heat transfer calculation, and determining the actual heat flux boundary conditions of the hot wall at the bottom of the tail cabin at the time of heat transfer calculation;

[0016] Step seven, divide the tail cabin heat transfer calculation grid, read the heat transfer calculation time and real-time temperature, load the actual tail cabin bottom thermal wall thermal boundary conditions at the tail cabin bottom loading calculation time, load the aerodynamic heating boundary conditions on the surface of the tail cabin outer heat protection layer, load the temperature boundary conditions on the inner wall of the tail nozzle insulation layer, carry out heat transfer numerical calculation, and obtain the overall thermal environment of the tail cabin including the tail cabin structure temperature and the thermal environment inside the tail cabin.

[0017] Furthermore, in step one, corresponding to the different shapes of the tail nozzle when it is not extended and when it reaches the maximum extension, the tail cabin external flow field and the tail jet joint simulation calculation grid are divided respectively, and the joint simulation calculation grid includes the tail cabin external heat protection surface, the tail cabin bottom surface, and the engine tail nozzle surface; the external flow parameters and the flow field parameters at the engine tail nozzle inlet are set, and the NS equation and the radiation transport equation are solved simultaneously through the numerical simulation method to calculate the radiation equilibrium temperature of the tail cabin bottom under the two states respectively.

[0018] Furthermore, in step 2, the maximum radiation equilibrium temperature value of the radiation equilibrium temperature of the outer surface of the bottom of the tail cabin corresponding to the two states is taken, which is recorded as Twr_max; at least one temperature value is selected between the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, and together with the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, a typical hot wall temperature value data set is formed, which is recorded as Twr. Among them, twr is the typical hot wall temperature value, and R is a real number set.

[0019] Furthermore, in step 4, the aerodynamic heating boundary condition is obtained by the cold wall heat flux dataset Qw(T q ,q w ) and the recovery enthalpy data set Hr(T q ,hr),where T q is the typical time selected for the calculation of cold wall heat flux and recovery enthalpy, q w is the typical time T q The corresponding cold wall heat flux, hr is the typical time T q The corresponding recovery enthalpy; the temperature data set of the inner wall of the tail nozzle insulation layer is Tw(Tf ,t w ), where T f is the typical time corresponding to the calculation of the temperature of the inner wall of the tail nozzle insulation layer, t w is the typical time T f The corresponding inner wall temperature of the tail nozzle insulation layer.

[0020] Furthermore, the typical time used for the aerodynamic heating boundary condition and the temperature boundary condition of the inner wall of the tail nozzle insulation layer, as well as the hot wall heat flux Qr at the bottom of the tail cabin corresponding to the tail nozzle not extended, are calculated. s (twr,q r ) and the corresponding thermal wall heat flux Qr at the bottom of the tail cabin when the tail nozzle reaches the maximum extension e (twr,q r ) are combined together to form the boundary condition feature point set T.

[0021] Furthermore, the cold wall heat flux data set Qw(T q ,q w ), recovery enthalpy data set Hr(T q ,hr), the temperature data set of the inner wall of the nozzle insulation layer Tw(T f ,t w ), heat flux value Qr of the hot wall at the bottom of the tail cabin s (twr,q r ) or Qr e (twr,q r ) are linearly interpolated to form a new boundary condition data set, which includes a new cold wall heat flow data set Qw'(T,q w ), new recovery enthalpy data set Hr'(T,hr), new nozzle insulation layer inner wall temperature data set Tw'(T,t w ), new heat flux dataset Qr of the bottom wall of the tail cabin s '(T,q r ), and the new heat flux data set Qr e '(T,q r ).

[0022] Furthermore, in step five,

[0023] according to Interpolate the new boundary condition data set, where T1 and T i_max are the minimum and maximum feature points in the boundary condition feature point set T, T1, T2, ...T i ,…T i_max are the characteristic points in the boundary condition characteristic point set T; when the new cold wall heat flux data set Qw'(T,qw ), a new recovery enthalpy data set Hr'(T,hr) or a new nozzle insulation layer inner wall temperature data set Tw'(T,t w ) is used for interpolation, x i is the heat transfer calculation time t i ,f(x i ) is the heat transfer calculation time t i The corresponding actual cold wall heat flux, actual recovery enthalpy, or actual tail nozzle insulation layer inner wall temperature; when the new tail cabin bottom hot wall heat flux data set Qr s '(T,q r ) or the new heat flux data set Qr of the bottom wall of the tail cabin e '(T,q r ) is used for interpolation, x i is the real-time temperature T of the bottom of the tail tank wi ,f(x i ) is the real-time temperature T wi At the moment when the tail nozzle begins to extend, rs0 The corresponding wall heat flux value Qr s '(T wi ), or when the tail nozzle extends to the maximum position t rsmax The corresponding wall heat flux value Qr e '(T wi ).

[0024] Further, in step six, when t i <t rs0 When the time point is between the start of flight and the engine ignition, the bottom of the tail cabin is not heated. At this time, the heat flux Qr i ′(t i ,T wi )=0;

[0025] When t rs0 ≤t i ≤t rsmax Then at time point t i The actual bottom wall heat flux is QUR s '(T wi ) and Qr e '(T wi ) are the wall temperatures T wi At the moment t when the tail nozzle starts to extend rs0 and the moment t at which it stretches to its maximum position rsmax The corresponding hot wall heat flux value;

[0026] When t i >t rsmax At the time before the engine is turned off, the heat flux on the hot wall is related to t rsmaxAt the same time, at time t i The actual bottom wall heat flux is Qr i '(t i ,T wi )=Qr e '(T wi ).

[0027] Furthermore, in step seven, the tail cabin heat transfer calculation grid includes the grid of the solid domain and the grid of the air domain in the tail cabin, and the solid domain includes the tail cabin outer heat protection layer, the tail nozzle insulation layer and the tail cabin bearing structure.

[0028] By applying the technical solution of the present invention, a method for calculating the thermal environment of a high-speed aircraft is provided. The method calculates the radiation equilibrium temperature of the bottom of the tail cabin in two states: when the tail nozzle of the engine is not extended and when it is extended to the maximum position, takes the maximum value, and takes the reference temperature, the maximum value of the radiation equilibrium temperature and several temperature values ​​between the two as the hot wall temperature data set for calculating the bottom hot wall heat flux; by interpolating the hot wall heat flux of the bottom of the tail cabin in the two states of when the tail nozzle is not extended and when it is extended to the maximum position to the time and the actual temperature of the bottom of the tail cabin, the hot wall heat flux value during the extension of the tail nozzle is obtained, so that the influence of the extension of the tail nozzle on the thermal environment of the bottom of the tail cabin is taken into account when calculating the heat transfer of the tail cabin; by merging the typical moments of aerodynamic heat, the inner wall temperature of the tail nozzle and the hot wall temperature value of the hot wall heat flux of the bottom of the tail cabin into a characteristic point data set, the pre-processing of three different boundary conditions is realized. The method for calculating the thermal environment of a high-speed aircraft of the present invention can improve the accuracy and calculation efficiency of the heat transfer calculation of the tail cabin. Compared with the prior art, the technical solution of the present invention can solve the technical problem of insufficient accuracy and efficiency in the calculation of the thermal environment of the tail cabin in the prior art. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The included drawings are used to provide a further understanding of the embodiments of the present invention, which constitute a part of the specification, are used to illustrate the embodiments of the present invention, and together with the text description, explain the principles of the present invention. Obviously, the drawings in the following description are only some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0030] Figure 1 A schematic diagram of a process flow of a high-speed aircraft thermal environment calculation method provided according to a specific embodiment of the present invention is shown;

[0031] Figure 2 A schematic diagram of the structure and boundary conditions of the tail cabin of an aircraft when the tail nozzle is not extended according to a specific embodiment of the present invention is shown;

[0032] Figure 3A schematic diagram of the tail cabin structure and boundary conditions of an aircraft when the tail nozzle reaches the maximum extension according to a specific embodiment of the present invention is shown. DETAILED DESCRIPTION

[0033] It should be noted that, in the absence of conflict, the embodiments in this application and the features in the embodiments can be combined with each other. The technical solutions 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. The following description of at least one exemplary embodiment is actually only illustrative and is by no means intended to limit the present invention and its application or use. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0034] It should be noted that the terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments according to the present application. As used herein, unless the context clearly indicates otherwise, the singular form is also intended to include the plural form. In addition, it should be understood that when the terms "comprise" and / or "include" are used in this specification, it indicates the presence of features, steps, operations, devices, components and / or combinations thereof.

[0035] Unless otherwise specifically stated, the relative arrangement, numerical expressions and numerical values ​​of the parts and steps set forth in these embodiments do not limit the scope of the present invention. Meanwhile, it should be understood that, for ease of description, the sizes of the various parts shown in the accompanying drawings are not drawn according to the actual proportional relationship. The technology, method and equipment known to those of ordinary skill in the relevant art may not be discussed in detail, but in appropriate cases, the technology, method and equipment should be considered as a part of the specification. In all examples shown and discussed here, any specific value should be interpreted as being merely exemplary, rather than as a limitation. Therefore, other examples of exemplary embodiments may have different values. It should be noted that similar reference numerals and letters represent similar items in the following drawings, and therefore, once a certain item is defined in an accompanying drawing, it does not need to be further discussed in subsequent drawings.

[0036] like Figure 1 As shown, according to a specific embodiment of the present invention, a method for calculating the thermal environment of a high-speed aircraft is provided, the method comprising:

[0037] At the reference temperature, through the joint numerical simulation of the tail cabin outer flow field and the tail jet flow field, the radiation equilibrium temperature of the tail cabin bottom outer surface corresponding to the two states when the tail nozzle is not extended and when the tail nozzle reaches the maximum extension is calculated respectively;

[0038] A typical hot wall temperature data set is constructed based on the radiation equilibrium temperature and reference temperature of the outer surface of the tail cabin bottom;

[0039] For the two states of the tail nozzle not being extended and the tail nozzle reaching the maximum extension, all typical hot wall temperature values ​​are set in sequence, and the joint simulation of the tail cabin outer flow field and the tail jet is carried out to obtain two sets of hot wall heat flux data at the bottom of the tail cabin corresponding to the two states;

[0040] Construct a set of boundary condition feature points, which includes all typical moments corresponding to aerodynamic heating and the inner wall temperature of the tail nozzle insulation layer, and typical hot wall temperature values ​​corresponding to the hot wall heat flux at the bottom of the tail cabin in two states. Perform linear interpolation processing on each boundary condition data at each feature point to generate a new boundary condition data set.

[0041] Interpolate the new boundary condition data set of aerodynamic heating and the temperature of the inner wall of the tail nozzle insulation layer at the time of heat transfer calculation, and interpolate the boundary condition data set of the hot wall heat flow at the bottom of the tail cabin in two states for the temperature of the bottom of the tail cabin, and obtain the boundary condition data of aerodynamic heating and the temperature of the inner wall of the tail nozzle insulation layer at the time of heat transfer calculation, and the hot wall heat flow data of the bottom of the tail cabin in two states corresponding to the real-time temperature;

[0042] The heat flux data of the hot wall at the bottom of the tail cabin in two states are interpolated at the time of heat transfer calculation to determine the actual heat flux boundary conditions of the hot wall at the bottom of the tail cabin at the time of heat transfer calculation;

[0043] Divide the heat transfer calculation grid of the tail cabin and read the heat transfer calculation time t i And real-time temperature T wi , tail tank bottom loading t i The actual thermal flow boundary condition of the hot wall at the bottom of the tail cabin at the moment, the aerodynamic heating boundary condition is loaded on the surface of the heat protection layer outside the tail cabin, and the temperature boundary condition is loaded on the inner wall of the tail nozzle insulation layer. Numerical calculation of heat transfer is carried out to obtain the overall thermal environment of the tail cabin including the tail cabin structure temperature and the thermal environment inside the tail cabin.

[0044] By using this configuration, a method for calculating the thermal environment of a high-speed aircraft is provided. The method calculates the radiation equilibrium temperature of the bottom of the tail cabin in two states: when the tail nozzle of the engine is not extended and when it is extended to the maximum position, takes the maximum value, and takes the reference temperature, the maximum value of the radiation equilibrium temperature, and several temperature values ​​between the two as the hot wall temperature data set for calculating the bottom hot wall heat flux; by interpolating the hot wall heat flux of the bottom of the tail cabin in the two states of when the tail nozzle is not extended and when it is extended to the maximum position to time and the actual temperature of the bottom of the tail cabin, the hot wall heat flux value during the extension of the tail nozzle is obtained, so that the influence of the extension of the tail nozzle on the thermal environment of the bottom of the tail cabin is taken into account when calculating the heat transfer of the tail cabin; by merging the typical moments of aerodynamic heat, the inner wall temperature of the tail nozzle, and the hot wall temperature value of the hot wall heat flux of the bottom of the tail cabin into a characteristic point data set, the pre-processing of three different boundary conditions is realized. The method for calculating the thermal environment of a high-speed aircraft of the present invention can improve the accuracy and calculation efficiency of the heat transfer calculation of the tail cabin.

[0045] Firstly, in the present invention, at the reference temperature, by means of the joint numerical simulation of the tail cabin outer flow field and the tail jet flow field, the radiation equilibrium temperature of the outer surface of the tail cabin bottom corresponding to the two states, when the tail nozzle is not extended and when the tail nozzle reaches the maximum extension, is calculated respectively.

[0046] As a specific embodiment of the present invention, the present invention corresponds to the different shapes of the tail nozzle when it is not extended and when it reaches the maximum extension, and divides the tail cabin external flow field and tail jet joint simulation calculation grid respectively, and the joint simulation calculation grid includes the tail cabin external heat protection surface, the tail cabin bottom surface, and the engine tail nozzle surface. The external flow parameters and the flow field parameters at the entrance of the engine tail nozzle are set, and the NS equation and the radiation transport equation are solved simultaneously by the numerical simulation method to calculate the radiation equilibrium temperature of the bottom of the tail cabin under the two states respectively.

[0047] Preferably, a normal temperature of 280K or 300K is set as the reference temperature, denoted as Twr_c.

[0048] Furthermore, in the present invention, after obtaining the radiation equilibrium temperature of the outer surface of the bottom of the tail cabin, a typical thermal wall temperature value data set is constructed according to the radiation equilibrium temperature of the outer surface of the bottom of the tail cabin and the reference temperature. In the present invention, the obtained typical thermal wall temperature value data set is used to carry out thermal wall heat flux calculation of the outer surface of the bottom of the tail cabin.

[0049] As a specific embodiment of the present invention, the maximum radiation equilibrium temperature value among the radiation equilibrium temperatures of the outer surface of the bottom of the tail cabin corresponding to the two states is taken, and is recorded as Twr_max. At least one temperature value is selected between the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, and together with the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, a typical hot wall temperature value data set is formed, which is recorded as Twr. Among them, twr is the typical hot wall temperature value, and R is a real number set.

[0050] Preferably, when the computer hardware resources are sufficient and the total time required for calculation is acceptable, the number of temperature values ​​selected between the maximum radiation equilibrium temperature and the reference temperature is as large as possible.

[0051] Furthermore, in the present invention, after constructing a typical hot wall temperature value data set, after the engine is ignited, all typical hot wall temperature values ​​are set in sequence for the two states, and a joint simulation of the tail cabin external flow field and the tail jet is carried out to obtain two sets of hot wall heat flux data at the bottom of the tail cabin corresponding to the two states.

[0052] As a specific embodiment of the present invention, the heat flux data of the hot wall at the bottom of the tail cabin corresponding to the tail nozzle not extended is recorded as Qr s (twr,q r ), the thermal wall heat flux data of the tail cabin bottom corresponding to the maximum extension of the tail nozzle is recorded as Qr e (twr,q r ), where twr represents different typical hot wall temperature values, q r is the hot wall heat flux value corresponding to the hot wall temperature value twr. At the time of engine ignition, the tail nozzle is about to start to extend but has not yet extended. This time is recorded as t rs0 The time when the tail nozzle extends to the maximum position is recorded as t rsmax , together forming the moment set Τ r ={t rs0 ,t rsmax}.

[0053] Furthermore, in the present invention, after obtaining two sets of hot wall heat flux data corresponding to the two states at the bottom of the tail cabin, a set of boundary condition feature points including aerodynamic heating, hot wall heat flux at the bottom of the tail cabin, and the temperature of the inner wall of the tail nozzle insulation layer is constructed, and the boundary condition data sets corresponding to each boundary condition feature point are linearly interpolated to generate a new boundary condition data set. In the present invention, the boundary condition feature point set is used for the tail cabin heat transfer calculation.

[0054] As a specific embodiment of the present invention, the aerodynamic heating boundary condition is obtained by using the cold wall heat flow data set Qw(T q ,q w ) and the recovery enthalpy data set Hr(T q ,hr),where T q is the typical time selected for heat flow and recovery enthalpy calculations, q w is the typical time T q The corresponding cold wall heat flux, hr is the typical time T q The corresponding recovery enthalpy. The temperature data set of the inner wall of the tail nozzle insulation layer is Tw(T f ,tw ), where T f is the typical time corresponding to the calculation of the temperature of the inner wall of the tail nozzle insulation layer, t w is the typical time T f The typical time used to combine the aerodynamic heating boundary condition and the temperature boundary condition of the inner wall of the tail nozzle insulation layer, and at the same time, the heat flux Qr s (twr,q r ) and Qr e (twr,q r ) are combined with the typical time to form the boundary condition characteristic point set T.

[0055] Then, the cold wall heat flux data set Qw(T q ,q w ), recovery enthalpy data set Hr(T q ,hr), the temperature data set of the inner wall of the nozzle insulation layer Tw(T f ,t w ), heat flux value Qr of the hot wall at the bottom of the tail cabin s (twr,q r ) or Qr e (twr,q r ) are linearly interpolated to form a new boundary condition data set, which includes a new cold wall heat flow data set Qw'(T,q w ), new recovery enthalpy data set Hr'(T,hr), new nozzle insulation layer inner wall temperature data set Tw'(T,t w ), new heat flux dataset Qr of the bottom wall of the tail cabin s '(T,q r ), and the new heat flux data set Qr e '(T,q r ).

[0056] Furthermore, in the present invention, after a new boundary condition data set is generated, an interpolation process is performed on the new boundary condition data set to obtain boundary condition data corresponding to the heat transfer calculation time and the real-time temperature.

[0057] As a specific embodiment of the present invention,

[0058] according to Interpolate the new boundary condition data set, where T1 and T i_max are the minimum and maximum feature points in the boundary condition feature point set T, T1, T2, ...T i_max are the characteristic points in the boundary condition characteristic point set T; when the new cold wall heat flux data set Qw'(T,qw ), a new recovery enthalpy data set Hr'(T,hr) or a new nozzle insulation layer inner wall temperature data set Tw'(T,t w ) is processed, x i is the heat transfer calculation time t i ,f(x i ) is the heat transfer calculation time t i The corresponding actual cold wall heat flux, actual recovery enthalpy, or actual tail nozzle insulation layer inner wall temperature; when the new tail cabin bottom hot wall heat flux data set Qr s '(T,q r ) or the new heat flux data set Qr of the bottom wall of the tail cabin e '(T,q r ) is processed, x i is the real-time temperature T of the bottom of the tail tank wi ,f(x i ) is the real-time temperature T wi At the moment when the tail nozzle begins to extend, rs0 The corresponding wall heat flux value Qr s '(T wi ), or when the tail nozzle extends to the maximum position t rsmax The corresponding wall heat flux value Qr e '(T wi ).

[0059] Furthermore, in the present invention, after obtaining the boundary condition data corresponding to the heat transfer calculation time and the real-time temperature, the actual heat flux value of the tail cabin bottom hot wall at the heat transfer calculation time is determined according to the heat transfer calculation time and the real-time temperature.

[0060] As a specific embodiment of the present invention, when the heat transfer calculation time t i <The time when the tail nozzle starts to extend t rs0 When the time point is between the start of flight and the engine ignition, the bottom of the tail cabin is not heated. At this time, the heat flux Qr i '(t i ,T wi )=0.

[0061] When t rs0 ≤t i ≤t rsmax Then at time point t i The actual bottom wall heat flux is QUR s '(T wi ) and Qr e '(T wi ) are the wall temperatures T wi At the moment t when the tail nozzle starts to extend rs0and the moment t at which it stretches to its maximum position rsmax The corresponding hot wall heat flux value.

[0062] When t i >t rsmax At the time before the engine is turned off, the heat flux on the hot wall is related to t rsmax At the same time, at time t i The actual bottom wall heat flux is Qr i '(t i ,T wi )=Qr e '(T wi ).

[0063] Furthermore, in the present invention, after determining the actual heat flux value of the hot wall at the bottom of the tail cabin, the heat transfer calculation grid of the tail cabin is divided, and the heat transfer calculation time t is read. i And real-time temperature T wi , load t i The actual heat flux value of the hot wall at the bottom of the tail cabin at the moment is calculated, the aerodynamic heating boundary condition is applied to the surface of the outer heat protection layer of the tail cabin, and the temperature boundary condition is applied to the inner wall of the tail nozzle insulation layer. Numerical calculation of heat transfer is carried out to obtain the overall thermal environment of the tail cabin including the structural temperature of the tail cabin and the thermal environment inside the tail cabin.

[0064] As a specific embodiment of the present invention, the heat transfer calculation grid of the tail cabin includes grids of solid domains such as the tail cabin outer heat protection layer, the tail nozzle insulation layer and the tail cabin bearing structure, and grids of the air domain in the tail cabin.

[0065] Compared with the prior art, the high-speed aircraft thermal environment calculation method of the present invention has the following beneficial effects:

[0066] (1) The present invention proposes a method for determining the hot wall temperature value used in the calculation of the heat flux of the hot wall at the bottom of the tail cabin, thereby solving the problem that the hot wall temperature value is determined unreasonably and calculation errors are caused by the hot wall temperature value given by the calculation personnel based on their experience.

[0067] (2) The present invention proposes a method for processing the heat flow of the hot wall at the bottom of the tail cabin before, during and after the extension of the tail nozzle structure of an aircraft, thereby solving the problem that the prior art cannot consider the impact of the thermal extension of the tail nozzle on the bottom heat flow.

[0068] (3) The present invention proposes a method for merging typical moments and typical thermal wall temperature values ​​of three boundary conditions, namely, the aerodynamic heating boundary condition of the outer surface of the tail cabin, the thermal flow boundary condition of the hot wall at the bottom of the tail cabin, and the temperature boundary condition of the inner wall of the tail nozzle insulation layer. The interpolation processing of three different types of boundary conditions can be realized through the same function, which reduces the complexity of the boundary condition data processing process during the tail cabin heat transfer calculation and improves the calculation efficiency.

[0069] In order to have a further understanding of the present invention, the numerical simulation process of the thermal environment of the tail cabin of a high-speed aircraft is taken as an example to explain in detail the method for calculating the thermal environment of a high-speed aircraft of the present invention.

[0070] In this embodiment, the high-speed aircraft thermal environment calculation method specifically includes the following steps.

[0071] Step 1: The reference temperature is defined as 280K, corresponding to the shape of the tail nozzle when it is not extended (i.e. the original shape, such as Figure 2 ), and the shape of the tail nozzle when it reaches its maximum extension (as shown in Figure 3 As shown in the figure, the tail cabin outer flow field and tail jet joint simulation calculation grids are divided respectively, and the grids include the tail cabin outer heat protection surface, tail cabin bottom surface, and engine tail nozzle surface. The high-speed incoming flow parameters and the flow field parameters at the engine tail nozzle inlet are set, and the NS equation and radiation transport equation are solved simultaneously through the numerical simulation method to calculate the radiation equilibrium temperature of the tail cabin bottom under two states.

[0072] Step 2, take out the maximum radiation equilibrium temperature value of the radiation equilibrium temperature of the bottom of the tail cabin under the two states. In this embodiment, the maximum radiation equilibrium temperature is 800K. Select a temperature value between the maximum radiation equilibrium temperature and the reference temperature. In this embodiment, it is 500K. Together with the maximum radiation equilibrium temperature and the reference temperature, it constitutes a typical thermal wall temperature value data set Twr = {280K, 500K, 800K}, which is used to carry out thermal wall heat flux calculation on the outer surface of the bottom of the tail cabin.

[0073] Step 3: After the engine is ignited, all typical hot wall temperature values ​​are set in sequence for the state where the tail nozzle is not extended and the state where the tail nozzle is extended to the maximum position, that is, 280K, 500K and 800K are set for the bottom of the tail cabin in sequence, and a joint simulation of the flow field outside the tail cabin and the tail jet is carried out to obtain two sets of hot wall heat flux data sets at the bottom of the tail cabin when the tail nozzle is not extended and when the tail nozzle is extended to the maximum position: Qr s (twr,q r ) and Qr e (twr,q r The joint simulation software used is the self-developed software XCFD, and the same type of commercial software can also be used for calculation, such as Fluent. The twr in the obtained hot wall heat flux data set represents different typical hot wall temperature values, including 280k, 500k, 800k, q r is the hot wall heat flux value corresponding to the typical hot wall temperature value.

[0074] In this embodiment, at the time of engine ignition, the tail nozzle is about to extend but has not yet extended, which is recorded as t rs0 , is the 60th second in the flight trajectory, i.e., t rs0 = 60s. The time when the tail nozzle extends to the maximum position is recorded as trsmax , is the 560th second in the flight trajectory, i.e., t rsmax = 160s. Together they form the set Τ r ={60s,560s}.

[0075] Step 4: Construct a set of boundary condition feature points, which includes all typical moments corresponding to aerodynamic heating and the inner wall temperature of the tail nozzle insulation layer, and typical hot wall temperature values ​​corresponding to the heat flux of the bottom hot wall of the tail cabin in two states. Perform linear interpolation processing on each boundary condition data at each feature point to generate a new data set of aerodynamic heating, bottom hot wall heat flux, and inner wall temperature of the tail nozzle insulation layer. The specific method is as follows:

[0076] (1) The aerodynamic heating boundary condition is determined by the cold wall heat flux data set Qw(T q ,q w ), recovery enthalpy data set Hr(T q ,hr),where T q is the typical time selected when carrying out heat flow and recovery enthalpy calculation. In this embodiment, the typical time is 30s, 60s, 350s, 560s and 600s. The temperature data set of the inner wall of the tail nozzle insulation layer is Tw(T f ,t w ), where T f is the typical time corresponding to the calculation of the temperature of the inner wall of the tail nozzle insulation layer. In this embodiment, the typical time is 30s, 60s, 300s, 500s and 650s. The typical time used for the boundary conditions of aerodynamic heating and the inner wall temperature of the tail nozzle insulation layer is combined, and the heat flux Qr of the hot wall at the bottom of the tail cabin is taken into account. s (twr,q r )、Qr e (twr,q r ) are combined with the typical time to form the boundary condition characteristic point set T. That is: T = {30, 60, 280, 300, 350, 500, 560, 600, 650, 800}.

[0077] (2) The cold wall heat flux data set Qw(T q ,q w ), recovery enthalpy data set Hr(T q ,hr), the temperature data set of the inner wall of the nozzle insulation layer Tw(T f ,t w ), heat flux value Qr of the hot wall at the bottom of the tail cabin s (twr,q r )、Qr e (twr,q r ) to perform linear interpolation to form a new boundary condition data set: Qw'(T,qw )、Hr'(T,hr)、Tw'(T,t w )、Qr s '(T,q r )、Qr e '(T,q r ).

[0078] Step 5: Perform interpolation calculation on the new boundary condition data set using the following function:

[0079]

[0080] Among them, 30 and 800 are the minimum and maximum feature points in the boundary condition feature point set T.

[0081] Process Qw'(T,q w )、Hr'(T,hr)、Tw'(T,t w ), the data set and the time t of heat transfer calculation are i Substituting into the above formula, we can obtain the heat transfer calculation time t i The actual cold wall heat flux, actual recovery enthalpy and actual inner wall temperature of the tail nozzle insulation layer.

[0082] Processing Qr s '(T,q r )、Qr e '(T,q r ), extract the real-time temperature T of the bottom of the tail cabin wi , the real-time temperature T wi Substituting into the above formula, we can obtain the time t when the tail nozzle starts to extend rs0 , when it stretches to the maximum position t rsmax , wall temperature T wi The corresponding wall heat flux value Qr s '(T wi ) and Qr e '(T wi ).

[0083] Step 6: Further process the heat flux data of the hot wall at the bottom of the tail cabin and extract the time point t for the heat transfer calculation of the tail cabin i and the real-time temperature of the bottom of the tail compartment T wi , determine t i The actual values ​​of the boundary conditions for the heat transfer calculation at time:

[0084] If t i <60s, the time point is between the start of flight and the engine ignition, the bottom of the tail cabin is not heated, and the heat flux Qr i ′(t i ,T wi )=0;

[0085] If 60s≤t i ≤560s. Through step 4, the nozzle starts to extend at 60s, extends to the maximum position at 560s, and the wall temperature T wi The corresponding wall heat flux value Qr s '(T wi ) and Qr e '(T wi ). Then at time point t i , the actual bottom wall heat flux value is:

[0086]

[0087] If t i >560s, before the engine is turned off, the heat flux of the hot wall is the same as the time when the tail nozzle is extended to the maximum position 560s, then t i The actual heat flow value at the moment is Qr i '(t i ,T wi )=Qr e '(T wi ).

[0088] Step 7: Divide the heat transfer calculation grid of the tail cabin, including the grids of the solid domains such as the tail cabin outer heat protection layer, the tail nozzle insulation layer, and the tail cabin bearing structure, and the grids of the air domain inside the tail cabin. The calculation time t of the heat transfer calculation process is read according to the above calculation method at the bottom of the tail cabin. i And real-time temperature T wi , load t i The actual heat flux value of the hot wall at the moment, the aerodynamic heating boundary condition is loaded on the surface of the heat protection layer outside the tail cabin, and the temperature boundary condition is loaded on the inner wall of the engine insulation layer. After the calculation is completed, the overall thermal environment of the tail cabin including the temperature of the tail cabin structure and the thermal environment inside the tail cabin can be obtained. The numerical calculation is completed using the self-developed software XCFD, and can also be completed using the same type of commercial software, such as Fluent.

[0089] In summary, the present invention provides a method for calculating the thermal environment of a high-speed aircraft, which calculates the radiation equilibrium temperature of the bottom of the tail cabin in two states: when the engine tail nozzle is not extended and when it is extended to the maximum position, takes the maximum value, and takes the normal temperature as the reference temperature, and takes the reference temperature, the maximum value of the radiation equilibrium temperature, and several temperature values ​​between the two as the hot wall temperature data set for calculating the bottom hot wall heat flux. By interpolating the hot wall heat flux of the bottom of the tail cabin in the two states of when the tail nozzle is not extended and when it is extended to the maximum position, the hot wall heat flux value during the extension of the tail nozzle is obtained, so that the influence of the extension of the tail nozzle on the thermal environment of the bottom of the tail cabin is taken into account when calculating the heat transfer of the tail cabin. By merging the typical moments of aerodynamic heat, the inner wall temperature of the tail nozzle, and the hot wall temperature value of the hot wall heat flux of the bottom hot wall of the tail cabin into a characteristic point data set, the pre-processing of three different boundary conditions is realized, so that the real-time values ​​of the three different boundary conditions can be obtained through an interpolation function during the heat transfer calculation, which reduces the complexity of the boundary condition processing and improves the calculation efficiency.

[0090] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for calculating the thermal environment of a high-speed aircraft, characterized in that: The high-speed aircraft thermal environment calculation method comprises: Step 1: Under the reference temperature, by means of joint numerical simulation of the tail cabin outer flow field and the tail jet flow field, the radiation equilibrium temperature of the outer surface of the tail cabin bottom corresponding to the two states, when the tail nozzle is not extended and when the tail nozzle reaches the maximum extension, is calculated respectively; Step 2: construct a typical hot wall temperature value data set based on the radiation equilibrium temperature and reference temperature of the outer surface of the tail cabin bottom; Step 3: For the two states of the tail nozzle not being extended and the tail nozzle reaching the maximum extension, all typical hot wall temperature values ​​are set in sequence, and the joint simulation of the tail cabin outer flow field and the tail jet is carried out to obtain two sets of hot wall heat flux data at the bottom of the tail cabin corresponding to the two states; Step 4: Construct a set of boundary condition feature points. The set of boundary condition feature points includes all typical moments corresponding to aerodynamic heating and the inner wall temperature of the tail nozzle insulation layer, and typical hot wall temperature values ​​corresponding to the hot wall heat flux at the bottom of the tail cabin in two states. Perform linear interpolation processing on each boundary condition data at each feature point to generate a new boundary condition data set. Step 5: Interpolate the new boundary condition data set of aerodynamic heating and the temperature of the inner wall of the tail nozzle insulation layer at the time of heat transfer calculation, and interpolate the new boundary condition data set of the tail cabin bottom hot wall heat flow in two states for the tail cabin bottom temperature, and obtain the boundary condition data of aerodynamic heating and the temperature of the inner wall of the tail nozzle insulation layer at the time of heat transfer calculation, and the two-state tail cabin bottom hot wall heat flow data corresponding to the real-time temperature; Step 6, interpolating the heat flux data of the hot wall at the bottom of the tail cabin in two states at the time of heat transfer calculation, and determining the actual heat flux boundary conditions of the hot wall at the bottom of the tail cabin at the time of heat transfer calculation; Step seven, divide the tail cabin heat transfer calculation grid, read the heat transfer calculation time and real-time temperature, load the actual tail cabin bottom thermal wall thermal boundary conditions at the tail cabin bottom loading calculation time, load the aerodynamic heating boundary conditions on the surface of the tail cabin outer heat protection layer, load the temperature boundary conditions on the inner wall of the tail nozzle insulation layer, carry out heat transfer numerical calculation, and obtain the overall thermal environment of the tail cabin including the tail cabin structure temperature and the thermal environment inside the tail cabin.

2. The high-speed aircraft thermal environment calculation method according to claim 1, characterized in that: In step one, corresponding to the different shapes of the tail nozzle when it is not extended and when it reaches the maximum extension, the tail cabin external flow field and the tail jet joint simulation calculation grid are divided respectively, and the joint simulation calculation grid includes the tail cabin external heat protection surface, the tail cabin bottom surface, and the engine tail nozzle surface; the external flow parameters and the flow field parameters at the engine tail nozzle inlet are set, and the NS equation and the radiation transport equation are solved simultaneously through the numerical simulation method to calculate the radiation equilibrium temperature of the tail cabin bottom under the two states respectively.

3. The high-speed aircraft thermal environment calculation method according to claim 1 or 2, characterized in that: In step 2, the maximum radiation equilibrium temperature value of the radiation equilibrium temperature of the outer surface of the bottom of the tail cabin corresponding to the two states is taken, which is recorded as Twr_max; at least one temperature value is selected between the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, and together with the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, it constitutes a typical hot wall temperature value data set, which is recorded as Twr, Among them, twr is the typical hot wall temperature value, and R is a real number set.

4. The high-speed aircraft thermal environment calculation method according to claim 3, characterized in that: In step 4, the aerodynamic heating boundary condition is determined by the cold wall heat flux data set Qw(T q ,q w ) and the recovery enthalpy data set Hr(T q ,hr),where T q is the typical time selected for the calculation of cold wall heat flux and recovery enthalpy, q w is the typical time T q The corresponding cold wall heat flux, hr is the typical time T q The corresponding recovery enthalpy; the temperature data set of the inner wall of the tail nozzle insulation layer is Tw(Tf,tw), where Tf is the typical time corresponding to the calculation of the temperature of the inner wall of the tail nozzle insulation layer, t w is the typical time T f The corresponding inner wall temperature of the tail nozzle insulation layer.

5. The high-speed aircraft thermal environment calculation method according to claim 4, characterized in that: The typical time used for the aerodynamic heating boundary condition and the temperature boundary condition of the inner wall of the tail nozzle insulation layer, as well as the hot wall heat flux Qr at the bottom of the tail cabin corresponding to the tail nozzle not extended s (twr,q r ) and the corresponding thermal wall heat flux Qr at the bottom of the tail cabin when the tail nozzle reaches the maximum extension e (twr,q r ) are combined together to form the boundary condition feature point set T.

6. The high-speed aircraft thermal environment calculation method according to claim 5, characterized in that: The cold wall heat flux data set Qw(T q ,q w ), recovery enthalpy data set Hr(T q ,hr), the temperature data set of the inner wall of the nozzle insulation layer Tw(T f ,t w ), heat flux value Qr of the hot wall at the bottom of the tail cabin s (twr,q r ) or Qr e (twr,q r ) are linearly interpolated to form a new boundary condition data set, which includes a new cold wall heat flow data set Qw'(T,q w ), new recovery enthalpy data set Hr'(T,hr), new nozzle insulation layer inner wall temperature data set Tw'(T,t w ), new heat flux dataset Qr of the bottom wall of the tail cabin s '(T,q r ), and the new tail cabin bottom thermal wall heat flux dataset Qr e '(T,q r ).

7. The high-speed aircraft thermal environment calculation method according to claim 6, characterized in that: In step five, according to Interpolate the new boundary condition data set, where T1 and T i_max are the minimum and maximum feature points in the boundary condition feature point set T, T1, T2, ...T i ,…T i_max are the characteristic points in the boundary condition characteristic point set T; when the new cold wall heat flux data set Qw'(T,q w ), a new recovery enthalpy data set Hr'(T,hr) or a new nozzle insulation layer inner wall temperature data set Tw'(T,t w ) is used for interpolation, x i is the heat transfer calculation time t i ,f(x i ) is the heat transfer calculation time t i The corresponding actual cold wall heat flux, actual recovery enthalpy, or actual tail nozzle insulation layer inner wall temperature; when the new tail cabin bottom hot wall heat flux data set Qr s '(T,q r ) or the new heat flux data set Qr of the bottom wall of the tail cabin e '(T,q r ) is used for interpolation, x i is the real-time temperature T of the bottom of the tail tank wi ,f(x i ) is the real-time temperature T wi At the moment when the tail nozzle begins to extend, rs0 The corresponding wall heat flux value Qr s '(T wi ), or when the tail nozzle extends to the maximum position t rsmax The corresponding wall heat flux value Qr e '(T wi ).

8. The high-speed aircraft thermal environment calculation method according to claim 7, characterized in that: In step 6, when t i <t rs0 When the time point is between the start of flight and the engine ignition, the bottom of the tail cabin is not heated. At this time, the heat flux Qr i ′(t i ,T wi )=0; When t rs0 ≤t i ≤t rsmax Then at time point t i The actual bottom wall heat flux is QUR s '(T wi ) and Qr e '(T wi ) are the wall temperatures T wi At the moment t when the tail nozzle starts to extend rs0 and the moment t at which it stretches to its maximum position rsmax The corresponding hot wall heat flux value; When t i >t rsmax At the time before the engine is turned off, the heat flux of the hot wall is related to t rsmax At the same time, at time t i The actual bottom wall heat flux is Qr i '(t i ,T wi )=Qr e '(T wi ).

9. The method for calculating the thermal environment of a high-speed aircraft according to any one of claims 1 to 8, characterized in that: In step seven, the heat transfer calculation grid of the tail cabin includes the grid of the solid domain and the grid of the air domain in the tail cabin. The solid domain includes the external heat protection layer of the tail cabin, the tail nozzle insulation layer and the tail cabin bearing structure.

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