Ground test simulation and characterization method for reentry of aircraft into atmospheric environment

Abstract

The invention discloses a ground test simulation and characterization method for reentry of an aircraft into an atmospheric environment, and the method comprises the steps: building a thermal environment parameter and component parameter test system based on a high-enthalpy shock tunnel, and obtaining a heat flow environment parameter and a component parameter of a target model through a test means; and the thermal environment and the real gas environment in the reentry process of the aircraft are represented. According to the method, the application range of the high-enthalpy shock tunnel is expanded, and a new thought is provided for quantitative characterization of the high-temperature real gas environment.

Description

Technical Field

[0001] The present invention relates to the field of high-enthalpy flow field tests, and particularly to a ground test simulation and characterization method for an aircraft re-entering the atmospheric environment. Background Art

[0002] When an aircraft lands and re-enters the atmosphere, the flight Mach number is relatively high, and the gas temperature around the aircraft can reach thousands or even tens of thousands of degrees Celsius, resulting in a strong high-temperature real gas effect problem. Moreover, the entire re-entry process changes with the change of flight parameters. The high-temperature real gas effect has a great impact on the flight thermal environment parameters and brings a large deviation to the heat flux prediction. In order to study the influence of the high-temperature real gas encountered during the re-entry process of the aircraft on the flight parameters, it is necessary to simulate the above high-temperature real gas effect environment on the ground. For this reason, the present invention proposes a ground test and characterization method for simulating the high-temperature real gas environment of an aircraft re-entering the atmosphere, which is used to solve the problems of test simulation and characterization of the high-temperature real gas environment. Summary of the Invention

[0003] The technical problem solved by the present invention is: aiming at the simulation problem of the changing environment of the high-temperature real gas effect encountered during the re-entry process of the aircraft, a ground test simulation and characterization method for an aircraft re-entering the atmospheric environment is proposed.

[0004] The technical solution of the present invention is: a ground test simulation and characterization method for an aircraft re-entering the atmospheric environment, including:

[0005] Based on a high-enthalpy shock tunnel, a test system for thermal environment parameters and component parameters is established; the test system includes a temperature sensor, a quantum cascade laser, a mid-infrared window, a condenser lens, and a detector; the temperature sensor is located at the stagnation point position of the target model, the transmission bands of the two mid-infrared windows are 5262 nm, and they are respectively located on the side windows on both sides of the wind tunnel test section. The connection line of the centers of the two windows is perpendicular to the central axis of the target model. The quantum cascade laser is located on one side of the wind tunnel test section, and the emitted laser passes through the two mid-infrared windows. The condenser lens is arranged opposite to the quantum cascade laser on the other side, and the light receiving part of the detector is located at the focal position of the condenser lens;

[0006] Start the test system, establish the wind tunnel flow field, and after the oncoming flow pressure meets the trigger condition, obtain the temperature sensor data and detector data within a fixed-length time period;

[0007] Based on the temperature sensor data, the heat flux data at each moment is obtained to characterize the high-temperature environment during the re-entry process of the aircraft. Based on the detector data, the NO component content at each moment is obtained to characterize the gas environment during the re-entry process of the aircraft.

[0008] Further, the test system further includes a laser controller and a signal generator; the signal generator is connected to the laser controller and continuously outputs harmonic signals, and satisfies: amplitude 2 - 2.5V, repetition frequency not less than 2KHz, duty cycle 50%, high-impedance output mode; the laser controller outputs two signals to be connected to the quantum cascade laser, one is a temperature control signal, and the other is a current control signal. The current control range of the laser controller covers 0 - 500mA, and the temperature control range covers T L , T L is the control temperature when the output center wavelength of the quantum cascade laser is 5262nm.

[0009] Further, the test system further includes a detector controller. The detector controller outputs one temperature control signal to be connected to the detector, and the temperature control range covers T D , T D is the high-sensitivity temperature of the detector.

[0010] Further, the peak response wavelength of the detector is between 4 - 8μm, the response time is not higher than 50ns, and the photosensitivity is not less than 1.0A / W.

[0011] Further, the test system further includes a pressure sensor, a trigger, and a data workstation; the output signals of the temperature sensor and the detector are both connected to the data workstation. The pressure sensor is located at the stagnation chamber position of the high-enthalpy shock tunnel. The output signal is divided into two paths, one path is output to the data workstation, and the other path is output to the trigger. When the voltage signal output by the pressure sensor increases to the trigger voltage of the trigger, the trigger outputs a trigger signal to make the data workstation start storing data, including the temperature sensor data and the detector data.

[0012] Further, the condenser lens is a concave mirror with a diameter of not less than 1cm and a focal length between 1 - 3cm.

[0013] Further, the transmittance of the mid-infrared window is not less than 95%, and the diameter is not less than 20cm.

[0014] Further, the heat flux data at each moment is calculated through the temperature sensor data. The specific method is as follows:

[0015] Substitute the temperature sensor data value U(t) at each moment t within a fixed length period after triggering into the following formula to calculate the heat flux data Q(t) at each moment t,

[0016]

[0017] In the formula, k, ρ, and c are respectively the thermal conductivity, density, and specific heat capacity of the temperature sensor material, τ is the independent variable, and π is the pi;

[0018] T(t) is the temperature value at time t, which is obtained from the following formula:

[0019]

[0020] where η is the calibration coefficient of the temperature sensor, and U(0) and T(0) are the temperature sensor value and the ambient temperature before the experiment.

[0021] Furthermore, the NO component content at each moment is calculated through the detector data. The specific method is as follows:

[0022] Based on the Beer-Lambert absorption spectroscopy principle, the detector data at each moment t within a fixed-length period after triggering is used to calculate the NO component content X at each moment according to the following formula NO ,

[0023]

[0024] where A is the integrated absorption rate; P is the static pressure of the flow field; L is the diameter of the uniform region of the flow field; and S is the line intensity at the static temperature T of the flow field.

[0025] The present invention also provides a test system for thermal environment parameters and component parameters, which is coupled to a high-enthalpy shock tunnel and includes a temperature sensor, a quantum cascade laser, a mid-infrared window, a condenser lens, and a detector; the temperature sensor is located at the stagnation point of the target model. The transmission bands of the two mid-infrared windows are 5262 nm, and they are respectively located on both side windows of the test section of the wind tunnel. The line connecting the center points of the two windows is perpendicular to the central axis of the target model. The quantum cascade laser is located on one side of the test section of the wind tunnel, and the emitted laser passes through the two mid-infrared windows. The condenser lens is arranged opposite to the quantum cascade laser on the other side, and the light-receiving part of the detector is located at the focal position of the condenser lens; the heat flux data and the NO component content at each moment are obtained through the temperature sensor and the detector respectively, characterizing the high-temperature environment and the gas environment during the re-entry process of the aircraft.

[0026] The advantages of the present invention compared with the prior art are as follows:

[0027] The present invention establishes a test system for measuring thermal environment parameters and component parameters based on a high-enthalpy shock tunnel, proposes a simulation method for the re-entry atmospheric environment of an aircraft, and for the first time proposes to combine the stagnation heat flux and the NO component concentration at the nozzle exit to quantitatively characterize the simulation state of the re-entry atmospheric environment. Through the air dissociation thermochemical phenomenon existing in the re-entry atmospheric environment, the present invention analyzes the composition of the dissociation products, and considering factors such as the feasibility of current tests, it is concluded that the NO concentration is an air thermochemical product related to the total temperature, with stable content under the same conditions and high measurement accuracy. Therefore, the NO component concentration is selected to characterize the degree of air dissociation in the re-entry atmospheric environment. There are strong thermochemical reactions related to air dissociation in the re-entry atmospheric environment of an aircraft. The simulation and quantitative characterization of this environment can improve the understanding of the aerodynamic characteristics of such aircraft and the level of layout optimization, while no clear characterization method is seen in the prior art. The present invention combines the stagnation total temperature parameter and the NO concentration parameter of the air dissociation product to characterize the ground simulation state of the re-entry atmospheric environment from two aspects of temperature and components, solves the problem in the prior art of only distinguishing the ground simulation state through total temperature information without considering the simulation of different degrees of air dissociation during re-entry, expands the application scope of the high-enthalpy shock tunnel, and provides an idea for quantitatively characterizing the variable high-temperature real gas environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 It is a schematic diagram of the test system for measuring thermal environment parameters and component parameters of the present invention.

[0029] In the figure: 1 - temperature sensor, 2 - pressure sensor, 3 - target model, 4 - quantum cascade laser, 5 - laser controller, 6 - signal generator, 7 - condenser lens, 8 - detector, 9 - detector controller, 10 - mid-infrared window, 11 - trigger, 12 - data workstation. DETAILED DESCRIPTION OF THE INVENTION

[0030] In order to better understand the technical solution of the present invention, the specific implementation manner of the present invention will be described below.

[0031] The present invention provides a ground test simulation and characterization method for the re-entry atmospheric environment of an aircraft, mainly including the following steps:

[0032] The first step is to establish a test system for measuring thermal environment parameters and component parameters based on a high-enthalpy shock tunnel, which can be referred to Figure 1 . It includes temperature sensor 1, pressure sensor 2, target model 3, quantum cascade laser 4, laser controller 5, signal generator 6, condenser lens 7, detector 8, detector controller 9, mid-infrared window 10, trigger 11, and data workstation 12.

[0033] The target model is installed in the test section of the high-enthalpy shock tunnel, and the model axis coincides with the center axis of the nozzle.

[0034] The maximum temperature measurable by the temperature sensor is not less than 400 °C. It is installed at the stagnation point of the target model, and one electrical signal is output, amplified and then transmitted to the data workstation.

[0035] The maximum pressure measurable by the pressure sensor is not less than 100 MPa. It is installed at the stagnation chamber of the wind tunnel. The output signal is divided into two paths. One path is output to the data workstation, and the other path is output to the trigger.

[0036] There are two mid-infrared windows in total, both of which meet the requirements that the transmittance in the 5262 nm band is not less than 95% and the diameter is not less than 20 cm. The two windows are respectively installed on both side windows of the wind tunnel test section, and the line connecting the center points of the two windows is perpendicular to the central axis of the spherical head model.

[0037] The output band of the quantum cascade laser covers 5262 nm. The control temperature when the center wavelength of the laser output is 5262 nm is queried, which is T L The quantum cascade laser is installed on one side of the wind tunnel test section, and the laser emitted by the laser must pass through two mid-infrared windows.

[0038] The laser controller mentioned above satisfies that the current control range covers 0 - 500 mA and the temperature control range covers T L The laser controller outputs two paths of signals to be connected to the quantum cascade laser. One path is the temperature control signal, and the other path is the current control signal.

[0039] The signal generator continuously outputs harmonic signals and satisfies: amplitude 2 - 2.5 V, repetition frequency not less than 2 KHz, duty cycle 50%, high-impedance output mode. The signal generator outputs one electrical signal to be connected to the laser controller.

[0040] The condenser lens is a concave mirror with a diameter not less than 1 cm and a focal length between 1 - 3 cm. Opposite to the quantum cascade laser, the condenser lens is placed on the other side of the wind tunnel test section for collecting the transmitted laser.

[0041] The peak response wavelength of the detector is between 4 - 8 μm, the response time is not higher than 50 ns, and the photosensitivity is not less than 1.0 A / W. The high-sensitivity temperature of the detector is queried and denoted as T D The light-receiving part of the detector is located at the focal point of the condenser lens, and one electrical signal is output, amplified and then stored in the data workstation.

[0042] The temperature control range of the detector controller mentioned above covers T D The detector controller outputs one temperature control signal to be connected to the detector.

[0043] In the second step, start the test system, establish the wind tunnel flow field, and obtain the temperature sensor data and detector data, specifically as follows:

[0044] S1. Power on all components of the test system, and set the control temperature of the laser controller to T. L , with the temperature control in the on state, the current control in the external control state, and the temperature control of the detector controller turned on.

[0045] S2. Output a harmonic signal from the signal generator to the laser controller. The laser controller drives the quantum cascade laser to emit light. The emitted laser passes through the wind tunnel flow field and is received by the detector, converted into an electrical signal, and transmitted to the data workstation.

[0046] S3. Set the pressures of the high-enthalpy shock tunnel driver section, compression section, and shock section to establish the condition for the wind tunnel flow field. Among them, the pressure values of the high-enthalpy shock tunnel driver section, compression section, and shock section with a total flow field temperature above 3000K can be obtained through the flow field calibration test.

[0047] S4. Set the trigger to the waiting trigger state and then start the flow field. As the flow field starts, the pressure sensor is subjected to the oncoming flow pressure, and the output voltage signal increases to the trigger voltage of the trigger. The trigger outputs a trigger signal to make the data workstation start storing data, and obtain the temperature sensor data and detector data within 10 ms after triggering.

[0048] The third step is to calculate the characterization parameters:

[0049] Substitute the temperature sensor data value U(t) at each moment t within 10 ms after triggering into the following formula to calculate the heat flux data Q(t) at each moment t.

[0050]

[0051] In the formula, k, ρ, and c are the thermal conductivity, density, and specific heat capacity of the temperature sensor material, respectively.

[0052] T(t) is the temperature value at moment t, which is obtained from the following formula.

[0053]

[0054] In the formula, α is the calibration coefficient of the temperature sensor, and U(0) and T(0) are the temperature sensor values before the test and the ambient temperature.

[0055] Based on the Beer-Lambert absorption spectroscopy principle, calculate the NO component content X at each moment according to the following formula for the detector data at each moment t within 10 ms after triggering. NO ,

[0056]

[0057] Where A is the integral absorption rate; P is the static pressure of the flow field; L is the diameter of the uniform region of the flow field; S is the line intensity at the static temperature T of the flow field, which can be obtained by querying the HITRAN database.

[0058] The high-temperature environment during the reentry process of the simulated aircraft is characterized by the heat flux Q(t). At the same Mach number, the larger the Q value, the lower the reentry altitude of the simulated aircraft. Through the NO component content X NO characterizes the high-temperature real gas environment. At the same Mach number, the larger X NO is, the higher the degree of gas dissociation.

[0059] It can be understood that the present invention is described by way of examples. As is known to those skilled in the art, without departing from the spirit and scope of the present invention, various changes or equivalent substitutions can be made to these features and examples. Additionally, under the teachings of the present invention, these features and examples can be modified to adapt to specific situations without departing from the spirit and scope of the present invention. Therefore, the present invention is not limited by the specific examples disclosed herein, and embodiments that can fall within the scope of the claims of this application all belong to the scope protected by the present invention.

[0060] The content not detailedly described in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims

1. A ground test simulation and characterization method for an aircraft re-entry atmospheric environment, characterized in that: include: Based on the high enthalpy shock wave wind tunnel, a thermal environment parameter and component parameter testing system was established; The test system comprises a temperature sensor (1), a quantum cascade laser (4), a mid-infrared window (10), a condenser (7), and a detector (8); the temperature sensor is located at the stationary point of the target model; the transmission bands of the two mid-infrared windows include the 5262 nm band and are respectively located at the side windows on both sides of the wind tunnel test section; the line connecting the center points of the two windows is perpendicular to the central axis of the target model; the quantum cascade laser is located at one side of the wind tunnel test section; the emitted laser passes through the two mid-infrared windows; the condenser is arranged on the other side opposite to the quantum cascade laser; and the light receiving part of the detector is placed at the focus position of the condenser; Start the test system, establish the wind tunnel flow field, and obtain the temperature sensor data and detector data within a fixed time period after the incoming flow pressure meets the trigger condition; Based on the temperature sensor data, the heat flow data at each moment is obtained to characterize the high temperature environment during the spacecraft re-entry process. Based on the detector data, the NO component content at each moment is obtained to characterize the gas environment during the spacecraft re-entry process.

2. The ground test simulation and characterization method for the aircraft re-entry atmosphere environment according to claim 1, characterized in that: The test system further comprises a laser controller (5) and a signal generator (6); the signal generator (6) is connected to the laser controller (5) and continuously outputs harmonic signals and meets the following conditions: amplitude 2-2.5V, repetition frequency not less than 2KHz, duty cycle 50%, and high impedance output mode; the laser controller (5) outputs two signals connected to the quantum cascade laser (4), one of which is a temperature control signal and the other is a current control signal. The current control range of the laser controller covers 0-500mA and the temperature control range covers T L , T L It is the control temperature when the output center wavelength of the quantum cascade laser is 5262nm.

3. The ground test simulation and characterization method for the aircraft re-entry atmosphere environment according to claim 1, characterized in that: The test system further comprises a detector controller (9), which outputs a temperature control signal connected to the detector (8), and the temperature control range covers T D , T D is the high sensitivity temperature of the detector.

4. The ground test simulation and characterization method for the aircraft re-entry atmosphere environment according to claim 1, characterized in that: The peak response wavelength of the detector (8) is between 4 and 8 μm, the response time is no higher than 50 ns, and the photosensitivity is no lower than 1.0 A / W.

5. The ground test simulation and characterization method for the aircraft re-entry atmosphere environment according to claim 1, characterized in that: The test system further comprises a pressure sensor (2), a trigger (11) and a data workstation (12); the output signals of the temperature sensor and the detector are both connected to the data workstation; the pressure sensor is located in the stationary position of the high enthalpy shock wave wind tunnel; the output signal is divided into two paths, one path is output to the data workstation (12), and the other path is output to the trigger (11); when the voltage signal output by the pressure sensor increases to the trigger voltage of the trigger, the trigger outputs a trigger signal to enable the data workstation to start storing data, including temperature sensor data and detector data.

6. The ground test simulation and characterization method for the aircraft re-entry atmosphere environment according to claim 1, characterized in that: The condenser is a concave mirror with a diameter of not less than 1 cm and a focal length between 1-3 cm.

7. The ground test simulation and characterization method for the aircraft re-entry atmosphere environment according to claim 1, characterized in that: The transmittance of the mid-infrared window is not less than 95% and the diameter is not less than 20 cm.

8. The ground test simulation and characterization method for an aircraft re-entry into the atmosphere according to claim 1, characterized in that: The heat flow data at each moment is calculated using the temperature sensor data. The specific method is: Substitute the temperature sensor data value U(t) at each time t within a fixed-length period after triggering into the following formula to calculate the heat flow data Q(t) at each time t: In the formula, k, ρ, c are the thermal conductivity, density and specific heat capacity of the temperature sensor material, τ is the independent variable, and π is the pi; T(t) is the temperature value at time t, which is obtained by the following formula: Where η is the calibration coefficient of the temperature sensor, U(0) and T(0) are the temperature sensor value and ambient temperature before the test.

9. The ground test simulation and characterization method for an aircraft re-entry into the atmosphere according to claim 1, characterized in that: The NO component content at each moment is calculated using the detector data. The specific method is as follows: The detector data at each time t within a fixed length period after triggering is calculated based on the Beer-Lambert absorption spectrum principle and the NO component content X at each time according to the following formula: NO , Where A is the integrated absorption rate; P is the static pressure of the flow field; L is the diameter of the uniform area of ​​the flow field; and S is the linear intensity at the static temperature T of the flow field.

10. A thermal environment parameter and component parameter testing system, characterized in that: Coupled to a high enthalpy shock wave wind tunnel, the device comprises a temperature sensor (1), a quantum cascade laser (4), a mid-infrared window (10), a condenser (7), and a detector (8); the temperature sensor is located at the stationary point of the target model; the transmission bands of the two mid-infrared windows include the 5262nm band and are respectively located at the side windows on both sides of the wind tunnel test section; the line connecting the center points of the two windows is perpendicular to the central axis of the target model; the quantum cascade laser is located at one side of the wind tunnel test section; the emitted laser passes through the two mid-infrared windows; the condenser is arranged on the other side opposite to the quantum cascade laser; and the light receiving part of the detector is placed at the focus position of the condenser; the heat flow data and NO component content at each moment are obtained through the temperature sensor and the detector, respectively, to characterize the high temperature environment and gas environment during the reentry of the aircraft.

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