A method for correcting total temperature of supersonic flow field

Abstract

The application belongs to the technical field of flow field temperature test, and relates to a correction method for total temperature of supersonic flow field air flow, comprising: establishing a supersonic calibration wind tunnel experiment system to simulate a supersonic flow field; performing calibration experiments by using the supersonic calibration wind tunnel experiment system to obtain the total temperature, total pressure and static pressure of the supersonic flow field air flow, and the effective temperature behind the shock wave under different shock wave front Mach numbers; according to the total temperature, total pressure, static pressure, shock wave front Mach number and effective temperature behind the shock wave, considering the non-isentropic effect caused by the shock wave of the supersonic flow field, obtaining the complex temperature coefficient; and according to the shock wave front Mach number, the effective temperature behind the shock wave corresponding to the shock wave front Mach number and the complex temperature coefficient corresponding to the shock wave front Mach number and the effective temperature behind the shock wave, correcting the total temperature of the air flow. The application can correct the problem that the total temperature measurement value of the flow field is lower than the actual value, and reduce the measurement error of the total temperature of the supersonic flow field air flow.

Description

Technical Field

[0001] This application relates to the field of flow field temperature testing technology, and in particular to a method for correcting the total temperature of a supersonic flow field. Background Technology

[0002] Accurate measurement of the total incoming flow temperature is required for supersonic aircraft, scramjet engines, and supersonic / hypersonic wind tunnel tests. The total incoming flow temperature is used to obtain aircraft / engine performance, calculate flow field velocity, static temperature, and other flow field parameters.

[0003] Currently, the commonly used measurement method involves fabricating a total temperature probe based on contact temperature measurement methods such as thermocouples and placing it in the flow field being measured. The probe head usually has a stagnation / shielding cover to ensure that the high-speed airflow is effectively stagnated at the probe's sensing end, thus obtaining the total airflow temperature. However, in actual measurements, thermal conductivity error, radiation error, and velocity error are unavoidable. Due to the influence of these factors, the temperature measured by the probe (effective temperature) is always lower than the true total airflow temperature. Among these, thermal conductivity error and radiation error can be effectively reduced by filling the probe with insulating material, coating the shielding cover with a thermal barrier coating, and appropriately selecting the thermocouple type. However, when the total airflow temperature being measured is low (below 500K), the velocity error caused by the airflow not being completely stagnant is the main factor affecting the accuracy of supersonic airflow total temperature measurement. Related studies have shown that for temperature measurement systems with a coefficient of restitution of 0.8 to 0.9, when the flight Mach number is 2, the relative velocity error can be as high as 4.4% to 8.9%. To reduce the impact of speed error, a dimensionless total temperature recovery coefficient (rewarming coefficient) is needed to correct the speed error.

[0004] In the prior art, the total temperature recovery coefficient is defined as shown in equation (1):

[0005]

[0006] In the formula: r is the coefficient of restitution, T a T is the effective temperature of the airflow (K), T is the static temperature of the airflow (K), and T0 is the total temperature of the airflow (K).

[0007] Since the static temperature of the airflow exists in equation (1), it is difficult to obtain in actual measurement. Based on the isentropic relationship between total and static temperature, the coefficient of restitution r is identically transformed into equation (2) for calculation:

[0008]

[0009] In the formula: κ is the airflow adiabatic index, which is taken as 1.4 for air, and M is the incoming Mach number.

[0010] Currently, the total temperature recovery coefficient of supersonic flow is mostly calculated using the formula derived from the isentropic relationship under subsonic conditions. However, when the Mach number of the incoming flow is greater than 1, when the total temperature probe is placed in the flow field being measured, a detached normal shock wave / bow shock wave is usually generated upstream of the temperature sensing end of the total temperature probe due to the influence of the flow around the blunt body. The actual temperature measured by the total temperature probe is the effective temperature of the airflow after the shock wave. Since the airflow passing through the shock wave is an adiabatic process, the total temperature of the airflow before and after the shock wave is equal. Therefore, the recovery coefficient can still be calculated using the isentropic relationship of the airflow after the shock wave to correct the total temperature of the incoming flow. However, the Mach number of the airflow after the shock wave is lower than that before the shock wave (and less than 1). If the isentropic Mach number obtained from the total-static pressure before the shock wave is still used for calculation, it can be seen from equation (2) that this will make the total temperature recovery coefficient higher than the actual value (higher than the recovery coefficient of the exposed temperature sensor under subsonic conditions). Moreover, this deviation will increase with the increase of the Mach number of the flow field being measured. The higher recovery coefficient will make the corrected total temperature of the airflow lower, resulting in a large measurement error. Moreover, if the total temperature of the supersonic flow field being measured is high, the effective temperature measured by the probe will be further reduced due to the influence of radiation error and thermal conduction error. At this time, the total temperature of the incoming flow obtained by correcting the "artificially high" total temperature recovery coefficient will deviate significantly from the true total temperature. Summary of the Invention

[0011] Therefore, it is necessary to provide a method for correcting the total temperature of airflow in a supersonic flow field to address the above-mentioned technical problems. This method can correct the problem that the measured total temperature of the flow field is lower than the actual value due to the high total temperature recovery coefficient caused by the shock wave in the supersonic flow field, thereby reducing the measurement error of the total temperature of airflow in the supersonic flow field.

[0012] A method for correcting the total temperature of a supersonic flow field includes:

[0013] A supersonic calibration wind tunnel experimental system was established to simulate a supersonic flow field;

[0014] A calibration experiment was conducted using a supersonic calibration wind tunnel experimental system to obtain the inflow total temperature, inflow total pressure and inflow static pressure of the supersonic flow field, and to obtain the effective temperature after the shock wave at different shock wave front Mach numbers.

[0015] Based on the total incoming temperature, total incoming pressure, static incoming pressure, Mach number before the shock wave, and effective temperature after the shock wave, the complex temperature coefficient is obtained by considering the non-isoentropic effect caused by the supersonic flow field shock wave.

[0016] The total airflow temperature is corrected based on the pre-shock Mach number, the post-shock effective temperature corresponding to the pre-shock Mach number, and the rewarming coefficient corresponding to the pre-shock Mach number and the post-shock effective temperature.

[0017] In one embodiment, based on the total incoming flow temperature, total incoming flow pressure, incoming flow static pressure, pre-shock Mach number, and post-shock effective temperature, and considering the non-entropic effect caused by the supersonic flow field shock wave, the complex temperature coefficient is obtained, including:

[0018]

[0019] In the formula, r is the retemperature coefficient, T0 is the total incoming flow temperature, and T a κ is the effective temperature after the shock wave, M is the Mach number of the incoming flow before the shock wave, P0 is the total pressure of the incoming flow, and P is the static pressure of the incoming flow.

[0020] In one embodiment, the total airflow temperature is corrected based on the pre-shock Mach number, the effective post-shock temperature corresponding to the pre-shock Mach number, and the rewarming coefficient corresponding to the pre-shock Mach number and the effective post-shock temperature, including:

[0021]

[0022] In the formula, T t This is the corrected total airflow temperature.

[0023] In one embodiment, the supersonic calibration wind tunnel experimental system includes: a drive motor, a compressor, a heater, a regulating valve, a wind tunnel expansion section, a stabilization section, a probe, a contraction-expansion section, a test section, a second contraction-expansion section, a post-diffuser section of the test section, a gas collection tank, a probe displacement mechanism and control system, and a data acquisition and fan / valve control system.

[0024] In one embodiment, a calibration experiment is performed using a supersonic calibration wind tunnel experimental system to obtain the inflow total temperature, inflow total pressure, and inflow static pressure of the supersonic flow field, including:

[0025] In the stable section of the calibration wind tunnel, a total temperature probe and a total pressure probe with a stagnation shroud are installed in the main flow area, with the probe head parallel to the main flow, to measure the total pressure and total temperature of the incoming flow.

[0026] Four sets of static pressure holes are symmetrically opened on the wall at the exit and inlet of the contraction-expansion nozzle to measure the static pressure of the incoming flow.

[0027] In one embodiment, obtaining the effective post-shock temperature at different pre-shock Mach numbers includes:

[0028] The total temperature probe to be calibrated is installed on the displacement mechanism and placed in the mainstream area of ​​the calibration wind tunnel test section, ensuring that the airflow direction is parallel to the temperature sensing end;

[0029] By adjusting the shape of the calibrated wind tunnel contraction section to change the Mach number of the incoming flow before the shock wave, and by changing the total temperature of the incoming flow through a heating device, the effective temperature after the shock wave was measured at different Mach numbers before the shock wave and different total temperatures of the incoming flow.

[0030] The aforementioned method for correcting the total temperature of airflow in a supersonic flow field, based on existing methods for calculating the total temperature restitution coefficient of subsonic adiabatic isentropic flow, couples the influence of non-isentropic processes corresponding to the normal / bow-shaped shock waves upstream of the temperature sensor in the supersonic flow field. It employs a one-dimensional correspondence between flow parameters before and after the normal shock wave, expressing the flow parameters after the shock wave as parameters before the shock wave, and derives a restitution coefficient correction formula suitable for measuring the total temperature of airflow in a supersonic flow field. The restitution coefficient can be obtained solely through calculation of the total / static flow parameters before the shock wave. By correcting the Mach number, it overcomes the difficulties in measuring flow field parameters after the shock wave in existing technologies, as well as the problem of a high total temperature restitution coefficient and a low measured total temperature due to a high Mach number. Furthermore, it establishes a correspondence between different incoming Mach numbers in supersonic flow and the total temperature restitution coefficient, reducing the total temperature measurement error and improving the accuracy of total temperature measurement in supersonic flow fields. Attached Figure Description

[0031] Figure 1 This is a flowchart illustrating a method for correcting the total temperature of a supersonic flow field in one embodiment.

[0032] Figure 2 This is a schematic diagram of the shock wave structure in front of the total temperature measurement probe in a supersonic flow field in one embodiment.

[0033] Figure 3 This is a schematic diagram of a thermocouple with a stagnation shield in one embodiment;

[0034] Figure 4 This is a schematic diagram of a supersonic total temperature probe calibration wind tunnel test system in one embodiment. Detailed Implementation

[0035] To make the objectives, technical solutions, and advantages of this application clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the scope of this application. All other embodiments obtained by those skilled in the art based on the embodiments in this application without inventive effort are within the scope of protection of this application.

[0036] It should be noted that all directional indicators (such as up, down, left, right, front, back, etc.) in the embodiments of this application are only used to explain the relative positional relationship and movement of each component in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indicator will also change accordingly.

[0037] Furthermore, the use of terms such as "first" and "second" in this application is for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined as "first" or "second" may explicitly or implicitly include at least one of those features. In the description of this application, "multiple sets" means at least two sets, such as two sets, three sets, etc., unless otherwise explicitly specified.

[0038] In this application, unless otherwise expressly specified and limited, the terms "connection," "fixed," etc., should be interpreted broadly. For example, "fixed" can mean a fixed connection, a detachable connection, or an integral part; it can mean a mechanical connection, an electrical connection, a physical connection, or a wireless communication connection; it can mean a direct connection or an indirect connection through an intermediate medium; it can mean the internal communication of two elements or the interaction between two elements, unless otherwise expressly limited. Those skilled in the art can understand the specific meaning of the above terms in this application according to the specific circumstances.

[0039] Furthermore, the technical solutions of the various embodiments of this application can be combined with each other, but only if they are based on the ability of those skilled in the art to implement them. When the combination of technical solutions is contradictory or cannot be implemented, it should be considered that such combination of technical solutions does not exist and is not within the scope of protection claimed by this application.

[0040] This application provides a method for correcting the total temperature of airflow in a supersonic flow field, such as... Figure 1 The flowchart shown, in one embodiment, includes:

[0041] Step 102: Establish a supersonic calibration wind tunnel experimental system to simulate a supersonic flow field.

[0042] Specifically, such as Figure 2 As shown, the supersonic calibration wind tunnel experimental system includes: a drive motor 1, a compressor fan 2, a heater 3, a regulating valve 4, a wind tunnel expansion section 5, a stabilization section 6, probes 7 (including a total temperature probe and a total pressure probe), a contraction-expansion section 8, a test section 9, a second contraction-expansion section 10, a post-test section diffuser 11, a gas collection tank 12, a probe displacement mechanism and control system, and a data acquisition and fan / valve control system. The specific structure and connections are existing technology and will not be elaborated upon here.

[0043] In this step, the wind tunnel contraction-expansion section uses contraction-expansion nozzles, such as solid-walled Laval nozzles, with each nozzle corresponding to a fixed airflow Mach number.

[0044] Step 104: A calibration experiment is conducted using a supersonic calibration wind tunnel experimental system to obtain the total incoming temperature, total incoming pressure, and static incoming pressure of the supersonic flow field, as well as the effective temperature after the shock wave at different Mach numbers before the shock wave.

[0045] Specifically:

[0046] In the main flow zone of the stable section of the calibration wind tunnel, total temperature probes and total pressure probes with stagnation hoods were installed respectively. The probe heads were parallel to the main flow to measure the total pressure P0 and total temperature T0 of the incoming flow. Since the airflow in the stable section is subsonic and the total static temperature itself is not much different, the measurement is relatively accurate. Therefore, it is used as the total temperature of the incoming flow of the supersonic airflow in the test section.

[0047] Four sets of static pressure holes are symmetrically opened on the wall at the exit and inlet of the contraction-expansion nozzle to measure the incoming static pressure P and further calculate the incoming Mach number.

[0048] The total temperature probe to be calibrated (with a K-type thermocouple at the sensing end and a shielding cover at the probe) is installed on the displacement mechanism and placed in the mainstream area of ​​the test section of the calibration wind tunnel, ensuring that the airflow direction is parallel to the sensing end. The Mach number before the shock wave is changed by adjusting the shape of the contraction section of the calibration wind tunnel, and the total temperature of the incoming flow is changed by the heating device. The effective temperature T behind the shock wave is measured under different Mach numbers before the shock wave and different total temperatures of the incoming flow. a .

[0049] In this step, such as Figure 3 and Figure 4 As shown (7 is an L-shaped probe, R is a thermocouple, J is a shock wave, and the arrows indicate the direction of the incoming flow, which is a flow with a Mach number > 1), while measuring the effective temperature after the shock wave, the deflection angle of the probe is changed using a displacement mechanism to obtain the insensitive angle range of the total temperature probe. That is, the recovery characteristics of the total temperature probe are calibrated in the calibration wind tunnel to obtain the effective temperature T after the shock wave of the temperature sensor. a The correspondence between the incoming flow parameters and different incoming flow Mach numbers M and different airflow angles.

[0050] Step 106: Based on the total incoming flow temperature, total incoming flow pressure, incoming flow static pressure, Mach number before the shock wave, and effective temperature after the shock wave, the abrupt change in Mach number before and after the shock wave caused by the supersonic flow field shock wave is considered to obtain the rewarming coefficient.

[0051] Specifically:

[0052]

[0053] In the formula, r is the retemperature coefficient, T0 is the total incoming flow temperature (equivalent to the standard value), and T a κ is the effective temperature after the shock wave (the actual measured value), M is the Mach number of the incoming flow before the shock wave, P0 is the total pressure of the incoming flow, and P is the static pressure of the incoming flow.

[0054] Considering the influence of high-temperature incoming flow, the adiabatic index κ (i.e., specific heat ratio) of the airflow is taken between 1 and 2. When the total temperature of the supersonic flow field being measured is below 600K, the values ​​of κ are 1.4 (air) and 1.33 (fuel gas). When the total temperature of the supersonic flow field being measured exceeds 600K, the value of κ needs to be obtained by referring to a table.

[0055] In this step, the relationship between the rewarming coefficient r of the temperature sensor in supersonic flow and different incoming flow parameters such as Mach number M and different airflow angle is obtained. This allows for the selection of an appropriate rewarming coefficient and correction based on the Mach number of the actual measurement conditions during the correction process.

[0056] Step 108: Correct the total airflow temperature based on the pre-shock Mach number, the effective post-shock temperature corresponding to the pre-shock Mach number, and the rewarming coefficient corresponding to the pre-shock Mach number and the effective post-shock temperature.

[0057] Specifically:

[0058]

[0059] In the formula, T t This is the corrected total airflow temperature.

[0060] In this step, the total temperature recovery coefficient can be corrected simply by measuring the flow parameters in front of the shock wave.

[0061] In this embodiment, the total temperature recovery coefficient can be corrected simply by measuring the flow field parameters in front of the shock wave. This method is applicable to situations where there is a normal shock wave / bow shock wave upstream of the temperature sensor. The applicable scope includes measurement probes using temperature sensors such as thermocouples and resistance temperature detectors, including single-point temperature sensors, multi-point temperature sensors, etc., such as total temperature probes, total temperature probe combs, total temperature probe rakes, and blade-shaped total temperature sensing parts.

[0062] Existing definitions of total temperature restitution coefficient are mostly based on the definition of isentropic processes under subsonic conditions, without considering the influence of non-isentropic effects of shock wave noise in supersonic flow fields. In actual supersonic flow total temperature measurements, the influence of normal / bow shock waves upstream of the probe is present. If the same total temperature restitution coefficient calculation method as under subsonic flow conditions is used, the Mach number of the incoming flow in the denominator is significantly higher than the actual Mach number at the measurement point. Therefore, the higher Mach number of the wavefront will affect the restitution coefficient, making it higher than the actual value. Consequently, the measured total temperature after the shock wave will be lower than the actual value, increasing the error.

[0063] The aforementioned method for correcting the total temperature of airflow in a supersonic flow field, based on existing methods for calculating the total temperature restitution coefficient of subsonic adiabatic isentropic flow, couples the influence of non-isentropic processes corresponding to the normal / bow-shaped shock waves upstream of the temperature sensor in the supersonic flow field. It employs a one-dimensional correspondence between flow parameters before and after the normal shock wave, expressing the flow parameters after the shock wave as parameters before the shock wave, and derives a restitution coefficient correction formula suitable for measuring the total temperature of airflow in a supersonic flow field. The restitution coefficient can be calculated using only the total / static flow parameters before the shock wave. By correcting the Mach number, it overcomes the difficulties in measuring flow field parameters after the shock wave in existing technologies, as well as the problem of a high total temperature restitution coefficient and a low measured total temperature due to a high Mach number. Furthermore, it establishes a correspondence between different incoming Mach numbers in supersonic flow and the total temperature restitution coefficient, reducing the total temperature measurement error and improving the accuracy of total temperature measurement in supersonic flow fields.

[0064] It should be understood that, although Figure 1 The steps in the flowchart are shown sequentially as indicated by the arrows, but these steps are not necessarily executed in the order indicated by the arrows. Unless otherwise specified herein, there is no strict order in which these steps are executed, and they can be performed in other orders. Figure 1 At least some of the steps in the process may include multiple sub-steps or multiple stages. These sub-steps or stages are not necessarily completed at the same time, but can be executed at different times. The execution order of these sub-steps or stages is not necessarily sequential, but can be executed in turn or alternately with other steps or at least some of the sub-steps or stages of other steps.

[0065] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0066] The embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are specific and detailed, they should not be construed as limiting the scope of this patent application. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this patent application should be determined by the appended claims.

Claims

1. A method for correcting total temperature of a supersonic flow field, characterized in that, The method comprises the following steps: An ultrasonic calibration wind tunnel experimental system is established to simulate an ultrasonic flow field; Calibration experiments are carried out by using the ultrasonic calibration wind tunnel experimental system to obtain the total temperature, total pressure and static pressure of the incoming flow of the ultrasonic flow field, and the effective temperature behind the shock wave under different shock wave front Mach numbers; According to the total temperature, total pressure, static pressure, shock wave front Mach number and effective temperature behind the shock wave, the non-isentropic effect caused by the shock wave of the ultrasonic flow field is considered to obtain the polytropic coefficient; The total temperature of the airflow is corrected according to the shock wave front Mach number, the effective temperature behind the shock wave corresponding to the shock wave front Mach number and the polytropic coefficient corresponding to the shock wave front Mach number and the effective temperature behind the shock wave. According to the total temperature, total pressure, static pressure, shock wave front Mach number and effective temperature behind the shock wave, the non-isentropic effect caused by the shock wave of the ultrasonic flow field is considered to obtain the polytropic coefficient, including: wherein is the recovery coefficient, is the total temperature of the incoming flow, is the effective temperature behind the shock wave, is the adiabatic exponent of the flow, is the Mach number of the incoming flow before the shock wave, is the total pressure of the incoming flow, is the static pressure of the incoming flow; The total temperature of the airflow is corrected according to the shock wave front Mach number, the effective temperature behind the shock wave corresponding to the shock wave front Mach number and the polytropic coefficient corresponding to the shock wave front Mach number and the effective temperature behind the shock wave, including: In the formula, Tt is the corrected total temperature of the airflow.

2. The method of claim 1, wherein, The ultrasonic calibration wind tunnel experimental system comprises a driving motor, a compression fan, a heater, an adjusting valve, a wind tunnel expansion section, a stable section, a probe, a contraction-expansion section, a test section, a second contraction-expansion section, a test section rear diffuser section, a gas collecting tank, a probe displacement mechanism and a control system, a data acquisition and fan / valve control system.

3. The method of claim 1 or 2, wherein, Calibration experiments are carried out by using the ultrasonic calibration wind tunnel experimental system to obtain the total temperature, total pressure and static pressure of the incoming flow of the ultrasonic flow field, including: A total temperature probe and a total pressure probe with a stagnation cover are respectively installed in the main flow area of the calibration wind tunnel stable section, and the probe head is parallel to the main flow to measure the total pressure and total temperature of the incoming flow; Four groups of static pressure holes are symmetrically arranged on the wall surface at the outlet and inlet of the contraction-expansion nozzle to measure the static pressure of the incoming flow.

4. The method of claim 1 or 2, wherein, The effective temperature behind the shock wave under different shock wave front Mach numbers is obtained, including: The calibrated total temperature probe is installed on the displacement mechanism and placed in the main flow area of the calibration wind tunnel test section, and the airflow direction is parallel to the temperature sensing end; The shape of the calibration wind tunnel contraction section is adjusted to change the shock wave front Mach number of the incoming flow, the total temperature of the incoming flow is changed by the heating device, and the effective temperature behind the shock wave is measured under different shock wave front Mach numbers and different total temperatures of the incoming flow.

Images