A high-speed aircraft thermal environment calculation method

Abstract

The application provides a high-speed aircraft thermal environment calculation method, which respectively calculates the radiation equilibrium temperatures of the bottom of the tail cabin in two states of the tail nozzle not being elongated and being elongated to the maximum position, takes the maximum value, takes the reference temperature, the maximum value of the radiation equilibrium temperature and a plurality of temperature values between the two as the thermal wall temperature data set for the calculation of the thermal wall heat flow of the bottom, obtains the thermal wall heat flow value in the elongation process of the tail nozzle through the interpolation of the thermal wall heat flow of the bottom of the tail cabin to time and the actual temperature of the bottom of the tail cabin, so that the influence of the elongation of the tail nozzle on the thermal environment of the bottom of the tail cabin is considered in the heat transfer calculation of the tail cabin. The aerodynamic heat, the typical time of the inner wall temperature of the tail nozzle and the thermal wall temperature value of the thermal wall heat flow of the bottom of the tail cabin are combined into a feature point data set, the pre-processing of three different boundary conditions is realized, the real-time values of the three different boundary conditions are obtained through an interpolation function in the heat transfer calculation, the complexity of the boundary condition processing is reduced, and the calculation efficiency is improved.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of aerodynamic heating and heat transfer, and particularly relates to a high-speed aircraft thermal environment calculation method. BACKGROUND

[0002] The external heating effect on the tail cabin of a high-speed aircraft is derived from three aspects: aerodynamic heating of the outer surface of the aircraft, convective heating of the bottom of the tail cabin of the aircraft, and heating of the high-temperature tail nozzle of the aircraft. The high-speed incoming flow produces severe aerodynamic heating effect on the outer surface of the aircraft, causing the temperature of the aircraft structure to rise. The bottom of the tail cabin is in the leeward area and is not directly subjected to the blowing of the high-speed incoming flow, but the pressure in the bottom area of the tail cabin is relatively low, and the high-temperature and high-speed gas flowing along the outer surface of the aircraft flows to the low-pressure area of the bottom of the tail cabin under the driving of the pressure. At the same time, the engine exhaust jet produces a certain blowing effect on the bottom of the tail cabin, and the two together cause the convective 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 takes the aerodynamic heating of the outer surface of the aircraft, the convective heating of the bottom of the tail cabin of the aircraft, and the heating of the high-temperature tail nozzle of the aircraft as three boundary conditions for the calculation of the overall thermal environment of the tail cabin structure and the cabin. The three boundary conditions are obtained through calculation or test: the cold-wall heat flux and the recovery enthalpy of the outer surface of the aircraft are calculated through numerical simulation of the outer flow field, the structural temperature of the outer surface of the tail cabin is extracted in real time along the time history of flight, the aerodynamic heating boundary condition is loaded on the outer surface of the tail cabin through cold-wall heat flux and recovery enthalpy cold-hot wall heat flux conversion; the hot-wall heat flux of the bottom of the tail cabin is calculated through the overall numerical simulation of the outer flow field and the exhaust jet, and is taken as the boundary condition of the bottom of the tail cabin; the temperature data of the inner wall of the engine tail nozzle insulation layer are obtained through numerical simulation of the flow field and heat transfer of the engine inner flow and the engine bearing structure, and are taken as the heating boundary condition of the high-temperature tail nozzle of the aircraft. The connection position between the tail cabin of the aircraft and other parts of the aircraft is set as an adiabatic boundary condition. The heat transfer numerical simulation of the air domain and equipment in the tail cabin structure and the overall thermal environment data of the tail cabin structure and the cabin are obtained.

[0004] The prior art has the following defects:

[0005] (1) In the calculation of the hot-wall heat flux of the bottom of the tail cabin, the hot-wall temperature of the bottom needs to be set. 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 calculation personnel, which may cause unreasonable hot-wall temperature value and calculation error.

[0006] (2) The aircraft tail nozzle structure has the property of thermal expansion and contraction. After the engine is ignited and started, the aircraft tail nozzle is heated to cause lengthening, the elongation of the tail nozzle changes the aircraft shape, which will cause the change of the degree of the tail jet flow blowing against the bottom of the tail cabin, affecting the heat flux of the bottom of the tail cabin. The prior art does not consider the influence of the elongation of the tail nozzle.

[0007] (3) The external aerodynamic heating boundary condition, the bottom heat wall heat flux / convection boundary condition and the tail nozzle inner wall temperature boundary condition are obtained through different numerical simulation processes, and correspond to different typical time points of the flight trajectory. The prior art does not merge the typical time points corresponding to the three boundaries. In the tail cabin heat transfer calculation, the actual boundary condition at the current time point is calculated by interpolating the boundary condition through three different functions according to the time history time point in the flight trajectory. The data processing process is relatively complex and the calculation time is relatively long. SUMMARY

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

[0009] The present application provides a high-speed aircraft thermal environment calculation method, which comprises:

[0010] Step one, at the reference temperature, through the joint numerical simulation of the tail cabin external flow field and the tail jet flow field, the radiation equilibrium temperature of the tail cabin bottom outer surface corresponding to two states of the tail nozzle not elongated and the tail nozzle reaching the maximum elongation is calculated respectively;

[0011] Step two, constructing a typical heat wall temperature value data set according to the radiation equilibrium temperature of the tail cabin bottom outer surface and the reference temperature;

[0012] Step three, for the two states of the tail nozzle not elongated and the tail nozzle reaching the maximum elongation, all the typical heat wall temperature values are set in turn, the joint simulation of the tail cabin external flow field and the tail jet flow field is carried out, and two groups of heat wall heat flux data of the tail cabin bottom corresponding to the two states are obtained;

[0013] Step four, constructing a boundary condition feature point set, the boundary condition feature point set includes all the typical time points corresponding to the aerodynamic heating, the tail nozzle heat insulation layer inner wall temperature and the typical heat wall temperature values corresponding to the heat wall heat flux of the bottom of the tail cabin in two states, respectively, linear interpolation processing of each boundary condition data at each feature point is carried out, and a new boundary condition data set is generated;

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

[0015] Step 6: Interpolate the heat flux data of the bottom hot wall of the stern compartment under two states at the time of heat transfer calculation to determine the actual heat flux boundary conditions of the bottom hot wall of the stern compartment at the time of heat transfer calculation.

[0016] Step 7: Divide the heat transfer calculation grid for the tail compartment, read the heat transfer calculation time and real-time temperature, load the actual tail compartment bottom hot wall thermal boundary conditions at the calculation time, load the aerodynamic heating boundary conditions on the surface of the tail compartment outer heat shield, load the temperature boundary conditions on the inner wall of the tail nozzle insulation layer, and carry out heat transfer numerical calculations to obtain the overall thermal environment of the tail compartment, including the tail compartment structural temperature and the thermal environment inside the tail compartment.

[0017] Furthermore, in step one, corresponding to the different shapes of the tail nozzle when it is not extended and when the tail nozzle reaches its maximum extension, the joint simulation calculation grid of the tail compartment external flow field and the tail nozzle flow is divided. The joint simulation calculation grid includes the tail compartment external heat shield surface, the tail compartment 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. Through numerical simulation, the Navier-Stokes equation and the radiation transport equation are solved simultaneously to calculate the radiation equilibrium temperature of the tail compartment bottom under the two states.

[0018] Further, in step two, the maximum radiation equilibrium temperature value among the radiation equilibrium temperatures of the outer surface of the bottom of the stern compartment corresponding to the two states is taken and denoted as Twr_max; at least one temperature value between the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c is selected, and together with the maximum radiation equilibrium temperature Twr_max and the reference temperature Twr_c, they constitute a typical hot wall temperature value dataset, denoted as Twr. Where twr is a typical hot wall temperature value, and R is a set of real numbers.

[0019] Furthermore, in step four, the aerodynamic heating boundary conditions are determined by the cold wall heat flux dataset Qw(T) q ,q w ) and recover the enthalpy dataset Hr(T) q Composed of ,hr), where T q This is a typical moment selected when performing cold wall heat flux and recovery enthalpy calculations, q w For typical time T q The corresponding cold wall heat flux, hr is the typical time T q The corresponding recovery enthalpy; the dataset for the inner wall temperature of the tail nozzle insulation layer is Tw(Tf t w ), wherein T f is a typical time corresponding to the calculation of the inner wall temperature of the nozzle heat shield, t w is the inner wall temperature of the nozzle heat shield corresponding to the typical time T f .

[0020] Further, the thermal wall temperature values twr in the aerodynamic heating boundary condition and the typical time used in the boundary condition of the inner wall temperature of the nozzle heat shield, and the thermal wall heat flux Qr s (twr, q r ) corresponding to the bottom of the tail cabin when the nozzle is not elongated, and the thermal wall heat flux Qr e (twr, q r ) corresponding to the bottom of the tail cabin when the nozzle reaches the maximum elongation, are combined to form a boundary condition feature point set T.

[0021] Further, the cold wall heat flux data set Qw(T q , q w ), the recovery enthalpy data set Hr(T q , hr), the inner wall temperature data set Tw(T f , t w ) of the nozzle heat shield, the thermal wall heat flux value Qr s (twr, q r ) or Qr e (twr, q r ) corresponding to each feature point in the boundary condition feature point set T are linearly interpolated to form a new boundary condition data set, which includes a new cold wall heat flux data set Qw'(T, q w ), a new recovery enthalpy data set Hr'(T, hr), a new inner wall temperature data set Tw'(T, t w ) of the nozzle heat shield, a new thermal wall heat flux data set Qr s '(T, q r ), and a new thermal wall heat flux data set Qr e '(T, q r ).

[0022] Further, in step five,

[0023] the new boundary condition data set is interpolated according to , wherein T1 and T i_max are the minimum feature point and the maximum feature point in the boundary condition feature point set T, T1, T2, … T i , … T i_max are each feature point in the boundary condition feature point set T; when interpolating the new cold wall heat flux data set Qw'(T, qw ), or a new recovery enthalpy dataset Hr'(T,hr) or a new tailpipe insulation inner wall temperature dataset Tw'(T,t) w When performing interpolation, x i For heat transfer calculation time t i f(x) i (t) represents the time t for heat transfer calculation. i The corresponding actual cold wall heat flux, actual recovery enthalpy, or actual tail nozzle insulation inner wall temperature; when the new tail section bottom hot wall heat flux dataset Qr is used... s '(T,q r ) or a new tail section bottom hot wall heat flow dataset Qr e '(T,q r When performing interpolation, x i The real-time temperature T at the bottom of the tail section was extracted. wi f(x) i (T) represents the real-time temperature. wi At the moment t begins to elongate of the tail nozzle rs0 The corresponding hot wall heat flux value Qr s '(T wi ), or the time t when the tail nozzle extends to its maximum position. rsmax The corresponding hot wall heat flux value Qr e '(T wi ).

[0024] Furthermore, in step six, when t i <t rs0 At this time, between the start of flight and engine ignition, the bottom of the tail section is not heated, and the heat flow Qr from the bottom hot wall is... i ′(t i ,T wi ) = 0;

[0025] When t rs0 ≤t i ≤t rsmax At time t, then at time point t i The actual bottom hotwall heat flow is Qr s '(T wi ) and Qr e '(T wi The wall temperatures T are respectively wi Below, the moment t when the tail nozzle begins to elongate rs0 and the time t when it stretches to its maximum position rsmax The corresponding hot wall heat flux value;

[0026] When t i >t rsmax At that time, in the time before the engine is shut down, the hot wall heat flow and t rsmaxthe same time, then at time point t i the actual bottom heat wall heat flow of the tail cabin at time point t i '(t i , T wi ) = Qr e '(T wi ).

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

[0028] The technical scheme of the present application provides a high-speed aircraft thermal environment calculation method, which respectively calculates the radiation equilibrium temperature of the tail cabin bottom in two states of the tail nozzle not being elongated and being elongated to the maximum position, takes the maximum value, takes the reference temperature, the maximum radiation equilibrium temperature, and a plurality of temperature values between the two as the heat wall temperature data set for the bottom heat wall heat flow calculation; obtains the heat wall heat flow value in the tail nozzle elongation process through the interpolation of the tail cabin bottom heat wall heat flow of the tail nozzle in the two states of not being elongated and being elongated to the maximum position with respect to time and the tail cabin bottom actual temperature, so that the influence of the tail nozzle elongation on the bottom thermal environment of the tail cabin is considered in the tail cabin heat transfer calculation; and the typical time of the aerodynamic heat and the inner wall temperature of the tail nozzle and the heat wall temperature value of the tail cabin bottom heat wall heat flow are combined into a feature point data set to realize the pre-processing of the three different boundary conditions. The high-speed aircraft thermal environment calculation method can improve the accuracy and calculation efficiency in the tail cabin heat transfer calculation. Compared with the prior art, the technical scheme of the present application can solve the technical problems of insufficient accuracy and efficiency in the tail cabin thermal environment calculation in the prior art. BRIEF DESCRIPTION OF DRAWINGS

[0029] The accompanying drawings included to provide a further understanding of the embodiments of the present application, constitute a part of the specification and serve to explain the principles of the present application together with the text. Obviously, the accompanying drawings in the following description are only some embodiments of the present application, and other drawings can be obtained by those skilled in the art without creative labor.

[0030] Figure 1 Fig. 1 shows a flowchart of the high-speed aircraft thermal environment calculation method according to the specific embodiments of the present application;

[0031] Figure 2 Fig. 2 shows a schematic diagram of the aircraft tail cabin structure and boundary conditions when the tail nozzle is not elongated according to the specific embodiments of the present application;

[0032] Figure 3Fig. 1 shows a schematic diagram of a vehicle tail cabin structure and boundary conditions when the tail nozzle reaches maximum elongation according to specific embodiments of the present application. DETAILED DESCRIPTION

[0033] It should be noted that the embodiments and features of the embodiments in the present application can be combined with each other without conflict. The technical solutions in the embodiments of the present application will be described clearly and completely in combination with the drawings of the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. The description of the at least one example embodiment is actually only illustrative, but not as any limitation on the present application and its application or use. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative labor fall within the scope of the present application.

[0034] It should be noted that the terms used herein are only intended to describe specific embodiments, and are not intended to limit the exemplary embodiments according to the present application. As used herein, the singular form is intended to include the plural form, unless the context clearly indicates otherwise, and it should also be understood that when the terms "comprise" and / or "include" are used in the specification, there is a reference to the presence of a feature, step, operation, device, component, and / or combinations thereof.

[0035] Unless specifically stated otherwise, the relative arrangements of the components and steps illustrated in these embodiments and the numerical expressions and values set forth herein are not limiting of the scope of the application. Also, it is to be understood that the various features, steps, and components described herein can be combined in any suitable manner without departing from the scope of the present application. Additionally, it should be noted that the drawings are not drawn to scale and that elements of the drawings can be exaggerated in size for the purposes of illustration and description. Techniques, methods, and apparatus known to those of ordinary skill in the art can not be discussed in detail herein. Any specific measures, values, and examples set forth herein are intended to be exemplary only, and it is contemplated that other examples can exist. It should be noted that similar reference numerals and letters in the various figures indicate similar items, and, as such, no further discussion relating to such similar items shall occur in the specification. It is to be understood that this application is not limited to particular examples described herein, which can be interpreted as illustrative only, but rather is intended to cover modifications and variations of those described herein. It is further noted that skilled artisans can employ other features not specifically described herein in this application without departing from the scope and spirit of the application. It is therefore apparent that there has been provided a method and apparatus in accordance with the present application, which are both economical, effective and which are susceptible of a broad scope of employments. While this application has been described as having exemplary designs, the present application can be further modified within the spirit and scope of this disclosure. This application is therefore intended to cover any and all changes which come within the scope of the

[0036] As shown in FIG. 1, according to specific embodiments of the present application, a high-speed aircraft thermal environment calculation method is provided, which comprises: Figure 1

[0037] At the reference temperature, by 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 two states, i.e., when the tail nozzle is not elongated and when the tail nozzle reaches maximum elongation, is calculated respectively; ​

[0038] A typical hot-wall temperature value dataset is constructed according to the radiation equilibrium temperature and the reference temperature of the outer surface of the tail cabin bottom;

[0039] For the two states of the tail nozzle not being elongated and the tail nozzle reaching the maximum elongation, all the typical hot-wall temperature values are set in turn, the combined simulation of the tail cabin outflow field and the tail jet flow is carried out, and two groups of hot-wall heat flux data of the tail cabin bottom corresponding to the two states are obtained;

[0040] A boundary condition feature point set is constructed, the boundary condition feature point set includes all the typical time corresponding to the aerodynamic heating and the tail nozzle heat shield inner wall temperature, and the typical hot-wall temperature values corresponding to the two groups of hot-wall heat flux data of the tail cabin bottom in the two states, and the linear interpolation processing of each boundary condition data at each feature point is carried out to generate a new boundary condition dataset;

[0041] The new boundary condition dataset of the aerodynamic heating and the tail nozzle heat shield inner wall temperature at the heat transfer calculation time is interpolated, the two groups of hot-wall heat flux boundary condition datasets of the tail cabin bottom in the two states are interpolated for the tail cabin bottom temperature, the boundary condition data of the aerodynamic heating and the tail nozzle heat shield inner wall temperature at the heat transfer calculation time corresponding to the real-time temperature are obtained, and the two groups of hot-wall heat flux data of the tail cabin bottom in the two states are obtained;

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

[0043] The tail cabin heat transfer calculation grid is divided, the heat transfer calculation time t i and the real-time temperature T wi are read, the actual hot-wall heat flux boundary condition of the tail cabin bottom at the time t i is loaded to the tail cabin bottom, the aerodynamic heating boundary condition is loaded to the tail cabin out heat shield surface, the temperature boundary condition is loaded to the tail nozzle heat shield inner wall, the heat transfer numerical calculation is carried out, and the overall thermal environment of the tail cabin including the tail cabin structure temperature and the tail cabin internal thermal environment is obtained.

[0044] With the configuration, a high-speed aircraft thermal environment calculation method is provided, which respectively calculates the radiation equilibrium temperatures of the bottom of the tail cabin in two states of the tail nozzle not being extended and being extended to the maximum position, takes the maximum value, takes the reference temperature, the maximum radiation equilibrium temperature and a plurality of temperature values between the two as the thermal wall temperature data set for the calculation of the thermal wall heat flux of the bottom, obtains the thermal wall heat flux value in the extension process of the tail nozzle through the interpolation of the thermal wall heat flux of the bottom of the tail cabin in the two states of the tail nozzle not being extended and being extended to the maximum position with respect to time and the actual temperature of the bottom of the tail cabin, so that the influence of the extension of the tail nozzle on the thermal environment of the bottom of the tail cabin is considered in the heat transfer calculation of the tail cabin, and the pre-processing of the three different boundary conditions is realized by combining the aerodynamic heat, the typical time of the inner wall temperature of the tail nozzle and the thermal wall temperature value of the thermal wall heat flux of the bottom of the tail cabin into the feature point data set. The high-speed aircraft thermal environment calculation method can improve the accuracy and calculation efficiency in the heat transfer calculation of the tail cabin.

[0045] Firstly, in the present application, under the reference temperature, the radiation equilibrium temperatures of the outer surface of the bottom of the tail cabin corresponding to the two states of the tail nozzle not being extended and being extended to the maximum length are calculated through the joint numerical simulation of the tail cabin outer flow field and the tail jet flow field.

[0046] As a specific embodiment of the present application, the present application divides the tail cabin outer flow field and the tail jet flow field joint simulation calculation grid for the different shapes corresponding to the tail nozzle not being extended and being extended to the maximum length, and the joint simulation calculation grid includes the tail cabin outer heat protection surface, the tail cabin bottom surface and the engine tail nozzle surface. The outer flow incoming flow parameters and the flow field parameters at the inlet of the engine tail nozzle are set, the N-S equation and the radiation transport equation are solved through the numerical simulation method, and the radiation equilibrium temperatures of the bottom of the tail cabin under the two states are calculated.

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

[0048] Further, in the present application, after the radiation equilibrium temperatures of the outer surface of the bottom of the tail cabin are obtained, the typical thermal wall temperature value data set is constructed according to the radiation equilibrium temperatures of the outer surface of the bottom of the tail cabin and the reference temperature. In the present application, the obtained typical thermal wall temperature value data set is used for the thermal wall heat flux calculation of the outer surface of the bottom of the tail cabin.

[0049] As a specific embodiment of the present application, the maximum radiation equilibrium temperature value in the radiation equilibrium temperatures of the outer surface of the bottom of the tail cabin corresponding to the two states is taken, which is denoted 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 the maximum radiation equilibrium temperature Twr_max, the reference temperature Twr_c and the temperature value together constitute the typical thermal wall temperature value data set, which is denoted as Twr, Wherein, twr is a typical hot wall temperature value, and R is a real set.

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

[0051] Further, in the present application, after constructing the typical hot wall temperature value data set, after the engine is ignited, the two groups of hot wall heat flux data of the tail cabin bottom corresponding to the two states are obtained by sequentially setting all the typical hot wall temperature values for the two states and carrying out joint simulation of the tail cabin outflow field and the tail jet.

[0052] As a specific embodiment of the present application, the hot wall heat flux data of the tail cabin bottom corresponding to the state when the tail nozzle is not elongated is denoted as Qr s (twr, q r ), and the hot wall heat flux data of the tail cabin bottom corresponding to the state when the tail nozzle reaches the maximum elongation is denoted as Qr e (twr, q r ), wherein twr represents different typical hot wall temperature values, and q r is the hot wall heat flux value corresponding to the hot wall temperature value twr. The moment when the engine is ignited, the tail nozzle is about to start elongation but has not yet elongated, which is denoted as t rs0 , and the moment when the tail nozzle is elongated to the maximum position is denoted as t rsmax , which together constitute the time set T r ={t rs0 ,t rsmax}.

[0053] Further, in the present application, after obtaining the two groups of hot wall heat flux data of the tail cabin bottom corresponding to the two states, the boundary condition characteristic point set including the aerodynamic heating, the hot wall heat flux of the tail cabin bottom, and the inner wall temperature of the tail nozzle heat insulation layer is constructed, and the boundary condition data set corresponding to each boundary condition characteristic point is processed by linear interpolation to generate a new boundary condition data set. In the present application, the boundary condition characteristic point set is used for tail cabin heat transfer calculation.

[0054] As a specific embodiment of the present application, the aerodynamic heating boundary condition is composed of the cold wall heat flux data set Qw(T q ,q w ) and the recovery enthalpy data set Hr(T q ,hr), wherein T q is a typical time selected for heat flux and recovery enthalpy calculation, q w is the cold wall heat flux corresponding to the typical time T q , and hr is the recovery enthalpy corresponding to the typical time T q . The tail nozzle heat insulation layer inner wall temperature data set is Tw(T f ,tw ), where T f This is the typical moment corresponding to the calculation of the inner wall temperature of the tailpipe insulation layer, t. w For typical time T f The corresponding temperature of the inner wall of the tail nozzle insulation layer. The typical moment used to combine the aerodynamic heating boundary conditions and the tail nozzle insulation layer inner wall temperature boundary conditions, while also including the heat flux Qr from the bottom of the tail section hot wall. s (twr,q r ) and Qr e (twr,q r The hot wall temperature value twr in the equation is combined with the typical moment to form the boundary condition feature point set T.

[0055] Then, the cold wall heat flux dataset Qw(T) corresponding to each feature point in the boundary condition feature point set T is analyzed. q ,q w ), Recover the enthalpy dataset Hr(T) q Data set of tailpipe insulation layer inner wall temperature Tw(T hr), f ,t w ), Heat flux value Qr of the bottom hot wall of the stern compartment s (twr,q r ) or Qr e (twr,q r Linear interpolation is performed separately to form a new boundary condition dataset, which includes a new cold wall heat flux dataset Qw'(T,q). w A new recovery enthalpy dataset Hr'(T,hr) and a new tailpipe insulation inner wall temperature dataset Tw'(T,t) w ), new tail section bottom thermal wall heat flux dataset Qr s '(T,q r ), and the new tail section bottom thermal wall heat flux dataset Qr e '(T,q r ).

[0056] Furthermore, in this invention, after generating a new boundary condition dataset, the new boundary condition dataset is interpolated to obtain the boundary condition data corresponding to the heat transfer calculation time and real-time temperature.

[0057] As a specific embodiment of the present invention

[0058] according to Interpolation is performed on the new boundary condition dataset, where T1 and T i_max These are the minimum and maximum feature points in the set of feature points T for boundary conditions, T1, T2, ..., T. i_max For each feature point in the boundary condition feature point set T; when applying the new cold wall heat flux dataset Qw'(T,q)w ), or a new recovery enthalpy dataset Hr'(T,hr) or a new tailpipe insulation inner wall temperature dataset Tw'(T,t) w When processing, x i For heat transfer calculation time t i f(x) i (t) represents the time t for heat transfer calculation. i The corresponding actual cold wall heat flux, actual recovery enthalpy, or actual tail nozzle insulation inner wall temperature; when the new tail section bottom hot wall heat flux dataset Qr is used... s '(T,q r ) or a new tail section bottom hot wall heat flow dataset Qr e '(T,q r When processing, x i The real-time temperature T at the bottom of the tail section was extracted. wi f(x) i (T) represents the real-time temperature. wi At the moment t begins to elongate of the tail nozzle rs0 The corresponding hot wall heat flux value Qr s '(T wi ), or the time t when the tail nozzle extends to its maximum position. rsmax The corresponding hot wall heat flux value Qr e '(T wi ).

[0059] Furthermore, in this invention, after obtaining the boundary condition data corresponding to the heat transfer calculation time and real-time temperature, the actual heat flux value of the bottom hot wall of the stern compartment at the heat transfer calculation time is determined based on the heat transfer calculation time and real-time temperature.

[0060] As a specific embodiment of the present invention, when heat transfer calculation time t i <Time t when the tail nozzle begins to elongate rs0 At this time, between the start of flight and engine ignition, the bottom of the tail section is not heated, and the heat flow Qr from the bottom hot wall is... i '(t i ,T wi ) = 0.

[0061] When t rs0 ≤t i ≤t rsmax At time t, then i The actual bottom hotwall heat flow is Qr s '(T wi ) and Qr e '(T wi The wall temperatures T are respectively wi Below, the time t when the tail nozzle begins to elongate rs0and the maximum elongation time t rsmax The corresponding hot-wall heat flux value.

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

[0063] Further, in the present application, after determining the actual bottom hot-wall heat flux value of the tail cabin, the tail cabin heat transfer calculation grid is divided, the heat transfer calculation time t i and the real-time temperature T wi are read, the actual bottom hot-wall heat flux value at t i is loaded, the aerodynamic heating boundary condition of the outer heat protection layer of the tail cabin is loaded, the temperature boundary condition of the inner wall of the tail nozzle heat insulation layer is loaded, and the heat transfer numerical calculation is carried out to obtain the overall thermal environment of the tail cabin including the tail cabin structure temperature and the tail cabin internal thermal environment.

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

[0065] The high-speed aircraft thermal environment calculation method of the present application has the following beneficial effects compared with the prior art:

[0066] (1) The present application proposes a method for determining the hot-wall temperature value used for calculating the bottom hot-wall heat flux of the tail cabin, which solves the problem of unreasonable hot-wall temperature value caused by relying on the experience of calculation personnel to give the hot-wall temperature value, which causes calculation errors.

[0067] (2) The present application proposes a method for processing the bottom hot-wall heat flux of the tail cabin before and after the elongation of the aircraft tail nozzle structure and during the elongation, which solves the problem that the prior art cannot consider the influence of the hot elongation of the tail nozzle on the bottom heat flux.

[0068] (3) The present application proposes a merging processing method for the typical time of three boundary conditions, i.e. the aerodynamic heating boundary condition of the outer surface of the tail cabin, the bottom hot-wall heat flux boundary condition of the tail cabin, and the temperature boundary condition of the inner wall of the tail nozzle heat insulation layer, and the typical hot-wall temperature value, which can realize the interpolation processing of the three different types of boundary conditions through one same function, reduces the complexity of the boundary condition data processing process during the tail cabin heat transfer calculation, and improves the calculation efficiency.

[0069] To provide a further understanding of the present invention, the following example uses a numerical simulation of the thermal environment of the tail section of a high-speed aircraft to illustrate the calculation method for the thermal environment of a high-speed aircraft.

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

[0071] Step 1: The reference temperature is defined as 280K, corresponding to the shape of the exhaust nozzle when it is not extended (i.e., the original shape, such as...). Figure 2 (as shown), and the shape of the tail nozzle when it reaches its maximum elongation (as shown). Figure 3 As shown, the joint simulation calculation meshes for the external flow field and the exhaust jet flow of the stern compartment are divided, including the external heat shield surface of the stern compartment, the bottom surface of the stern compartment, and the surface of the engine exhaust nozzle. High-speed inflow parameters and flow field parameters at the engine exhaust nozzle inlet are set. The Navier-Stokes equations and radiation transport equations are solved simultaneously using numerical simulation methods to calculate the radiation equilibrium temperature at the bottom of the stern compartment under two conditions.

[0072] Step 2: Extract the maximum radiation equilibrium temperature value from the radiation equilibrium temperatures at the bottom of the stern compartment under the two conditions. 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, this constitutes a typical hot wall temperature value dataset Twr={280k,500k,800k}, which is used to calculate the hot wall heat flux on the outer surface of the bottom of the stern compartment.

[0073] Step 3: After engine ignition, for both the nozzle not extended and the nozzle extended to its maximum position, all typical hot wall temperature values ​​are sequentially set, i.e., 280K, 500K, and 800K are set sequentially for the bottom of the tail section. Joint simulations of the external flow field and the exhaust flow are then performed to obtain two sets of hot wall heat flux datasets for the bottom of the tail section when the nozzle is not extended and when it is extended to its maximum position: Qr s (twr,q r ) and Qr e (twr,q r The co-simulation software used is the self-developed XCFD software; similar commercial software, such as Fluent, can also be used for calculations. In the obtained hot-wall heat flux dataset, twr represents different typical hot-wall temperature values, including 280K, 500K, and 800K. r It is the heat flux value of the hot wall corresponding to the typical hot wall temperature value.

[0074] In this embodiment, the tailpipe is about to extend but has not yet extended at the moment of engine ignition, denoted as t. rs0 It is the 60th second in the flight trajectory, i.e., t rs0 =60s. The time when the tail nozzle reaches its maximum extension position is denoted as t.rsmax is the 560s in the flight trajectory, i.e. t rsmax = 160s. Together, they constitute a set T r = {60s, 560s}.

[0075] Step four, constructing a boundary condition feature point set, which includes all the typical time corresponding to the aerodynamic heating, the inner wall temperature of the tail nozzle heat shield, and the typical hot wall temperature value corresponding to the bottom hot wall heat flux of the tail cabin in two states. Linear interpolation is performed on each boundary condition data at each feature point to generate a new aerodynamic heating, bottom hot wall heat flux, and tail nozzle heat shield inner wall temperature boundary condition data set. The specific method is as follows:

[0076] (1) The aerodynamic heating boundary condition is composed of the cold wall heat flux data set Qw(T q , q w ) and the recovery enthalpy data set Hr(T q , hr), wherein T q is the typical time selected for heat flux and recovery enthalpy calculation. In this embodiment, the typical time is 30s, 60s, 350s, 560s, and 600s. The tail nozzle heat shield inner wall temperature data set is Tw(T f , t w ), wherein T f is the typical time corresponding to the tail nozzle heat shield inner wall temperature calculation. In this embodiment, the typical time is 30s, 60s, 300s, 500s, and 650s. The typical time for aerodynamic heating and tail nozzle heat shield inner wall temperature boundary conditions is combined, and the hot wall temperature value twr in Qr s (twr, q r ) and Qr e (twr, q r ) is combined with the typical time to form the boundary condition feature point set T. That is: T = {30, 60, 280, 300, 350, 500, 560, 600, 650, 800}.

[0077] (2) Linear interpolation is performed on the cold wall heat flux data set Qw(T q , q w ), the recovery enthalpy data set Hr(T q , hr), the tail nozzle heat shield inner wall temperature data set Tw(T f , t w ), the bottom hot wall heat flux value Qr s (twr, q r ), and Qr e (twr, q r ) corresponding to the feature points in the feature point set T 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 five, interpolation calculation of the new boundary condition data set is carried out by the following function,

[0079]

[0080] where 30 and 800 are the minimum and maximum characteristic points in the boundary condition characteristic point set T.

[0081] When processing Qr w '(T, q w ), Hr'(T, hr), Tw'(T, t i ), the data set and the time t i of heat transfer calculation are brought into the above formula, and the actual cold wall heat flow, actual recovery enthalpy and actual inner wall surface temperature of the tail nozzle heat insulation layer at the time t s of heat transfer calculation can be obtained.

[0082] When processing Qr r '(T, q e ), Qr r '(T, q wi ), the real-time temperature T wi of the tail cabin bottom is extracted, and the real-time temperature T rs0 is brought into the above formula, and the hot wall heat flow value Qr rsmax '(T wi ) and Qr s '(T wi ) corresponding to the time t e when the tail nozzle starts to stretch, the time t wi when the stretching reaches the maximum position and the wall surface temperature T i can be obtained.

[0083] Step six, further processing the hot wall heat flow data of the tail cabin bottom, extracting the time point t wi of heat transfer calculation of the tail cabin and the real-time temperature T i of the tail cabin bottom, and determining the actual value of the boundary condition of heat transfer calculation at t i :

[0084] If t i < 60s, the time point is between the start time of flight and the ignition time of the engine, the bottom of the tail cabin is not affected by heating, and the bottom hot wall heat flow Qr i '(t wi , T i ) = 0.

[0085] If 60s≤t i ≤560s, the moment of the beginning of the nozzle elongation 60s, the moment of the elongation to the maximum position 560s, and the wall temperature T wi The corresponding heat flux value of the hot wall Qr s '(T wi ) and Qr e '(T wi ) are obtained through step four. Then at the moment point t i , the actual bottom heat flux value of the hot wall is:

[0086]

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

[0088] Step seven, divide the heat transfer calculation grid of the tail compartment, including the grid of the tail compartment outer heat protection layer, the tail nozzle heat insulation layer, the tail compartment bearing structure and other solid domains, and the air domain grid in the tail compartment. According to the foregoing calculation method, read the calculation moment t i and the real-time temperature T wi in the heat transfer calculation process, load the actual heat flux value at the moment t i , load the aerodynamic heating boundary condition on the surface of the tail compartment outer heat protection layer, load the temperature boundary condition on the inner wall of the engine heat insulation layer, and carry out heat transfer numerical calculation. After the calculation is completed, the overall thermal environment of the tail compartment including the tail compartment structure temperature and the tail compartment internal thermal environment can be obtained. The numerical calculation is completed by using self-developed software XCFD, and can also be completed by using commercial software of the same type, such as Fluent.

[0089] In summary, the application provides a high-speed aircraft thermal environment calculation method, which respectively calculates the radiation equilibrium temperature of the bottom of the tail cabin in two states of the tail nozzle not being elongated and being elongated to the maximum position, takes the maximum value, and takes the normal temperature as the reference temperature, the maximum value of the radiation equilibrium temperature and a plurality of temperature values between the two as the bottom hot wall heat flow calculation hot wall temperature data set. The interpolation of the bottom hot wall heat flow of the tail cabin in the two states of the tail nozzle not being elongated and being elongated to the maximum position with respect to time and the actual temperature of the bottom of the tail cabin obtains the hot wall heat flow value in the elongation process of the tail nozzle, so that the influence of the elongation of the tail nozzle on the thermal environment of the bottom of the tail cabin is considered in the heat transfer calculation of the tail cabin. By combining the aerodynamic heat, the typical time of the inner wall temperature of the tail nozzle and the hot wall temperature value of the bottom hot wall heat flow of the tail cabin into a feature point data set, the pre-processing of the 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 in the heat transfer calculation, the complexity of the boundary condition processing is reduced, and the calculation efficiency is improved.

[0090] The preferred embodiments of the application have been described above with the purpose of not being limited to the application, and for those skilled in the art, the application can have various changes and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the application shall be included in the protection scope of the application.

Claims

1. A method for calculating a thermal environment of a high-speed aircraft, characterized by, The high-speed aircraft thermal environment calculation method comprises: Step one, at a reference temperature, through 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 two states of the tail nozzle not being elongated and the tail nozzle reaching the maximum elongation is calculated respectively; Step two, a typical thermal wall temperature value data set is constructed according to the radiation equilibrium temperature of the tail cabin bottom outer surface and the reference temperature; Step three, for the two states of the tail nozzle not being elongated and the tail nozzle reaching the maximum elongation, all the typical thermal wall temperature values are set in turn, joint simulation of the tail cabin outer flow field and the tail jet flow is carried out, and two groups of thermal wall heat flow data of the tail cabin bottom corresponding to the two states are obtained; Step four, a boundary condition feature point set is constructed, the boundary condition feature point set comprises all the typical time corresponding to the aerodynamic heating and the tail nozzle heat insulation layer inner wall temperature, and the typical thermal wall temperature values corresponding to the tail cabin bottom thermal wall heat flow of the two states, linear interpolation processing of each boundary condition data on each feature point is carried out respectively, and a new boundary condition data set is generated; Step five, the new boundary condition data set of the aerodynamic heating and the tail nozzle heat insulation layer inner wall temperature at the heat transfer calculation time is interpolated, the new boundary condition data set of the tail cabin bottom thermal wall heat flow at the heat transfer calculation time is interpolated, the aerodynamic heating and the tail nozzle heat insulation layer inner wall temperature boundary condition data corresponding to the heat transfer calculation time and the two groups of tail cabin bottom thermal wall heat flow data corresponding to the real-time temperature are obtained; Step six, the two groups of tail cabin bottom thermal wall heat flow data at the heat transfer calculation time are interpolated, and the actual tail cabin bottom thermal wall heat flow boundary condition at the heat transfer calculation time is determined; Step seven, tail cabin heat transfer calculation grids are divided, the heat transfer calculation time and the real-time temperature are read, the actual tail cabin bottom thermal wall heat boundary condition at the calculation time is loaded to the tail cabin bottom, the aerodynamic heating boundary condition is loaded to the tail cabin outer heat protection layer surface, the temperature boundary condition is loaded to the tail nozzle heat insulation layer inner wall, heat transfer numerical calculation is carried out, and the whole tail cabin thermal environment including the tail cabin structure temperature and the tail cabin internal thermal environment is obtained.

2. The method of claim 1, wherein, In step one, the joint simulation calculation grids of the tail cabin outer flow field and the tail jet flow are divided respectively corresponding to different shapes of the tail nozzle not being elongated and the tail nozzle reaching the maximum elongation, the joint simulation calculation grids comprise the tail cabin outer heat protection surface, the tail cabin bottom surface and the engine tail nozzle surface; the outer flow inflow parameters and the engine tail nozzle inlet flow field parameters are set, the N-S equation and the radiation transport equation are solved simultaneously through the numerical simulation method, and the radiation equilibrium temperatures of the tail cabin bottom under the two states are calculated respectively.

3. The high-speed aircraft thermal environment calculation method according to claim 1 or 2, characterized by, In step two, the maximum radiation equilibrium temperature value in the radiation equilibrium temperature of the tail cabin bottom outer surface corresponding to the two states respectively is taken, denoted 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 the maximum radiation equilibrium temperature Twr_max, the reference temperature Twr_c together constitute a typical hot wall temperature value data set, denoted as Twr, Wherein, twr is a typical hot wall temperature value, R is a real number set.

4. The method of claim 3, wherein, In step four, the aerodynamic heating boundary condition is composed of 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 instant selected for the cold-wall heat flux and recovery enthalpy calculation, q w is the cold-wall heat flux corresponding to the typical time instant T q , and hr is the recovery enthalpy corresponding to the typical time instant T q ; the inner wall temperature data set of the nozzle insulation is Tw(T f , t w ), where T f is the typical time instant corresponding to the nozzle insulation inner wall temperature calculation, and t w is the nozzle insulation inner wall temperature corresponding to the typical time instant T f .

5. The method of claim 4, wherein, The typical time instants used for the aerodynamic heating boundary condition and the inner wall temperature boundary condition of the nozzle insulation, and the corresponding hot-wall heat flux Qr at the bottom of the aft compartment when the nozzle is not extended s (twr, q r ) and the corresponding hot-wall heat flux Qr at the bottom of the aft compartment when the nozzle is at maximum extension e (twr, q r ) are combined to form a set of boundary condition characteristic points T.

6. The method of claim 5, wherein, linearly interpolating the cold-wall heat flux data set Qw(T q , q w ), the recovery enthalpy data set Hr(T q , hr), the inner-wall temperature data set Tw(T f , t w ) of the nozzle insulation layer, the hot-wall heat flux value Qr s (twr, q r ) or Qr e (twr, q r ) of the bottom of the tail compartment respectively, to form a new boundary condition data set, which includes a new cold-wall heat flux data set Qw'(T, q w ), a new recovery enthalpy data set Hr'(T, hr), a new inner-wall temperature data set Tw'(T, t w ) of the nozzle insulation layer, a new hot-wall heat flux data set Qr s '(T, q r ) of the bottom of the tail compartment, and a new hot-wall heat flux data set Qr e '(T, q r ) of the bottom of the tail compartment.

7. The method of claim 6, wherein, In step five, according to Interpolation is performed on the new boundary condition dataset, where T1 and T i_max These are the minimum and maximum feature points in the set of feature points T for boundary conditions, T1, T2, ..., T. i ...T i_max For each feature point in the boundary condition feature point set T; when applying the new cold wall heat flux dataset Qw'(T,q) w ), or a new recovery enthalpy dataset Hr'(T,hr) or a new tailpipe insulation inner wall temperature dataset Tw'(T,t) w When performing interpolation, x i For heat transfer calculation time t i f(x) i (t) represents the time t for heat transfer calculation. i The corresponding actual cold wall heat flux, actual recovery enthalpy, or actual tail nozzle insulation inner wall temperature; when the new tail section bottom hot wall heat flux dataset Qr is used... s '(T,q r ) or a new tail section bottom hot wall heat flow dataset Qr e '(T,q r When performing interpolation, x i The real-time temperature T at the bottom of the tail section was extracted. wi f(x) i (T) represents the real-time temperature. wi At the moment t begins to elongate of the tail nozzle rs0 The corresponding hot wall heat flux value Qr s '(T wi ), or the time t when the tail nozzle extends to its maximum position. rsmax The corresponding hot wall heat flux value Qr e '(T wi ).

8. The method of claim 7, wherein, 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, and the heat flux Qr i ′(t i ,T wi ) = 0; When t rs0 ≤t i ≤t rsmax At time t, then at time point t i The actual bottom hotwall heat flow is Qr s '(T wi ) and Qr e '(T wi The wall temperatures T are respectively wi Below, the moment t when the tail nozzle begins to elongate rs0 and the time t when it stretches to its maximum position rsmax The corresponding hot wall heat flux value; When t i > t rsmax , the time before the engine is turned off, the hot wall heat flux is the same as the time t rsmax , then the actual bottom hot wall heat flux at the time point t i is Qr i '(t i , T wi ) = Qr e '(T wi ).

9. The high-speed aircraft thermal environment calculation method according to any one of claims 1 to 8, characterized by, In step seven, the tail cabin heat transfer calculation grids comprise the grids of the solid domain and the air domain grids in the tail cabin, the solid domain comprises the tail cabin outer heat protection layer, the tail nozzle heat insulation layer and the tail cabin bearing structure.

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