A device for measuring total pressure loss of an ultrasonic internal flow field

Abstract

The application provides a device for measuring total pressure loss of an ultrasonic internal flow field, comprising a total pressure probe, a support rod and a wind tunnel bottom plate; the total pressure probe is arranged on the wind tunnel bottom plate through the support rod. An airflow delivery channel is arranged in the total pressure probe; a pressure measuring hole is arranged at the central position of a sharp wedge structure at the front end of the total pressure probe; an airflow outlet is arranged at the rear end of the total pressure probe; and the airflow outlet is connected with a pressure scanning valve through a hose. The device can realize simultaneous measurement of total pressure and Mach number in a single measurement without twice measurement of total pressure and Mach number respectively, and has easy operability and measurement accuracy compared with a traditional circular probe; compared with a normal shock wave in front of the circular probe, the oblique shock wave in front of the probe has smaller plugging degree to the flow field, and can reduce the influence of the invasive device on the start of the flow field; when the array arrangement probe realizes simultaneous measurement of multiple points of total pressure, the influence of the oblique shock wave-oblique shock wave interference in front of the probe on the measurement accuracy is very low, and the accuracy of the result is higher.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aerodynamic flow measurement, and particularly relates to a device for measuring total pressure loss of supersonic internal flow field. BACKGROUND

[0002] Supersonic internal flow field widely exists in the inlet, isolator and combustion chamber of scramjet, and the flow thereof contains shock wave / shock wave interference, shock wave / boundary layer interference, shock wave / jet interference, expansion wave, shear layer and other features, and has strong non-uniformity. For the entire flow passage, the total pressure recovery coefficient of the inlet and the combustion chamber can reflect the thrust loss, and is an important index for evaluating the performance of the engine. Therefore, experimental measurement of the total pressure at the inlet and outlet of the inlet and the combustion chamber can extract key pressure data from the supersonic complex internal flow field, so as to optimize the design of the inlet profile and the combustion chamber structure.

[0003] The prior art generally replaces the inlet, combustion chamber and other flow passages with a wind tunnel test section, and uses the wind tunnel experiment to realize the diagnosis and measurement of the supersonic internal flow field. For total pressure measurement, the most commonly used is a circular hollow pipe total pressure probe. When the wind tunnel flow (if the Mach number is known) passes through, a normal shock wave is generated in front of the probe, and then the subsonic flow is reduced and enters the pipe to be measured. Through the relationship between the total pressure of the wave behind and the normal shock wave, the actual total pressure of the flow at the probe can be deduced. If the total pressure of the flow is known, the Mach number of the flow can also be deduced, and therefore, this probe is also commonly used to calibrate the Mach number of the wind tunnel flow. This traditional total pressure probe needs to know the total pressure or the Mach number when measuring, so as to realize the measurement or deduction of another parameter. However, for the supersonic complex internal flow field, such as the inlet and combustion chamber of the scramjet, the Mach number and the total pressure of the flow are unknown, and therefore, even if the total pressure behind the wave of the probe is known, the actual total pressure in front of the wave at the probe cannot be deduced. In the previous operation method, two wind tunnel experiments are needed. The first experiment uses a circular total pressure probe to measure the total pressure behind the normal shock wave; the second experiment removes the probe and adds a slope with a known angle at the probe, and the shock wave angle of the slope is measured by a schlieren system, the Mach number at the probe is deduced according to the relationship of the oblique shock wave, and finally the total pressure in front of the wave is deduced according to the relationship of the normal shock wave.

[0004] The existing circular total pressure probe is used to measure the total pressure of the supersonic complex internal flow field, and two experiments are needed to deduce the total pressure in front of the wave. On the one hand, the operation is complicated, and the probe and the slope need to be replaced repeatedly, and on the other hand, for the transient flow field, the existing technology cannot realize the simultaneous measurement of the Mach number and the total pressure, and the total pressure of the flow cannot be measured. Moreover, when measuring the total pressure distribution of a certain section, a plurality of probes need to be arranged in an array, and due to the normal shock wave in front of the probe, the problem of mutual interference of the normal shock waves in front of the probes also occurs, which to some extent limits the experimental measurement accuracy of the total pressure distribution of the section. Summary of the Invention

[0005] To address the shortcomings of existing technologies, this invention provides a device for measuring the total pressure loss in a supersonic internal flow field. This invention can simultaneously measure the Mach number and the total pressure behind the wave, and also reduces shock wave interference when the probe is arranged in a multi-point array, thereby improving the measurement accuracy of the total pressure distribution across the cross section.

[0006] The technical solution of the present invention is: a measuring device for total pressure loss in a supersonic internal flow field, comprising a total pressure probe, a support rod, and a wind tunnel base plate; the total pressure probe is mounted on the wind tunnel base plate via the support rod.

[0007] Preferably, the total pressure probe has an internal airflow delivery channel, the front end of the total pressure probe has a wedge structure, and a pressure measuring hole is provided at the center of the wedge structure; the rear end of the total pressure probe has an airflow outlet, and the airflow outlet is connected to a pressure scanning valve through a flexible tube.

[0008] Preferably, the angle θ between the wedge structure and the horizontal plane is 20 degrees.

[0009] Preferably, the length and width of the pressure measuring hole are both 0.5 mm.

[0010] As a preferred method, the total pressure loss is measured using the aforementioned total pressure probe as follows:

[0011] S1) Assume that the Mach number and total pressure of the airflow at the upstream end of the total pressure probe are Ma1 and P, respectively. 01 After passing through the oblique shock wave at the leading edge of the wedge, the flow Mach number and total pressure drop are Ma2 and P, respectively. 02 ;

[0012] S2) After the airflow enters the probe, since the flow is still not supersonic, a normal shock wave is generated inside the airflow delivery channel of the probe, causing the flow to slow down to subsonic speed. The Mach number of the flow after the normal shock wave is Ma3, P 03 Subsonic flow total pressure P 03 The subsonic flow total pressure P is collected and recorded by the pressure scanning valve and then processed over time. 03 The specific value;

[0013] S3) The total pressure behind the wave measured by the total pressure probe is the total pressure behind both the oblique shock wave and the normal shock wave. The actual total airflow pressure P at the probe is calculated using the shock wave angle β of the oblique shock wave. 01 .

[0014] Preferably, in step S3), P is obtained based on the total pressure relationship before and after the normal shock wave and the oblique shock wave. 01 and P 02 ;Right now:

[0015]

[0016] In the formula, β is the included angle between the oblique shock wave and the horizontal plane, i.e. the shock wave angle; γ is a preset parameter.

[0017] As preferred, in step S3), the Mach numbers Ma1 and Ma2 are calculated by the following formulae:

[0018]

[0019]

[0020] In the formula, β is the included angle between the oblique shock wave and the horizontal plane, i.e. the shock wave angle; γ is a preset parameter; θ is the included angle between the wedge structure and the horizontal plane θ = 20 degrees.

[0021] As preferred, the total pressure probes are arranged in a vertical array to simultaneously measure the total pressure at multiple points.

[0022] The present application has the following beneficial effects:

[0023] 1. The present application does not need to measure the total pressure and the Mach number twice, but can realize the simultaneous measurement of the total pressure and the Mach number in a single measurement, and compared with the traditional circular probe, has the advantages of easy operation and measurement accuracy.

[0024] 2. Compared with the normal shock wave in front of the circular probe, the oblique shock wave in front of the probe in the present application has a smaller degree of blockage to the flow field, and can reduce the influence of the intrusive device on the start of the flow field.

[0025] 3. When the array arrangement probe realizes the simultaneous measurement of the total pressure at multiple points, the influence of the oblique shock wave-oblique shock wave interference in front of the probe in the present application on the measurement accuracy is very low, and the accuracy of the result is higher. BRIEF DESCRIPTION OF DRAWINGS

[0026] Figure 1 Fig. 1 is a structural schematic diagram of the measurement device of the present application;

[0027] Figure 2 Fig. 2 is a flow field structural schematic diagram of the total pressure probe of the present application;

[0028] Figure 3 Fig. 3 is a flow schematic diagram of the measurement of the present application;

[0029] In the figure, 1 is a total pressure probe; 2 is a support rod; 3 is a wind tunnel bottom plate;

[0030] 11 is a pressure measuring hole; 12 is an air outlet. DETAILED DESCRIPTION

[0031] The specific embodiments of the present application are further described below in combination with the drawings:

[0032] As Figure 1 shown, the present embodiment provides a device for measuring total pressure loss of supersonic internal flow field, comprising a total pressure probe 1, a support rod 2 and a wind tunnel bottom plate 3; the total pressure probe 1 is arranged on the wind tunnel bottom plate 3 through the support rod 2.

[0033] As preferred in the present embodiment, the inside of the total pressure probe 1 is provided with an airflow delivery channel, the front end of the total pressure probe 1 is provided with a wedge structure, and the wedge structure is horizontally symmetrical. The included angle θ between the wedge structure and the horizontal plane is 20 degrees. A pressure measuring hole 11 is arranged at the central position of the wedge structure; in the present embodiment, the length and width of the pressure measuring hole 11 are both 0.5 mm. An airflow outlet 12 is arranged at the rear end of the total pressure probe 1, and the airflow outlet 12 is connected with a pressure scanning valve through a hose.

[0034] As preferred in the present embodiment, as shown in Figure 2 and 3 , the method for measuring total pressure loss by using the total pressure probe 1 is as follows:

[0035] S1), assuming that the Mach number and total pressure of the airflow at the most upstream position of the total pressure probe 1 are Ma1 and P 01 , respectively; after passing through the oblique shock wave of the wedge front edge, the flow Mach number and total pressure are Ma2 and P 02 , respectively.

[0036] S2), after the airflow enters the total pressure probe 1, a normal shock wave is generated inside the airflow delivery channel of the total pressure probe 1 to reduce the flow to subsonic speed, and the flow Mach number and total pressure after the normal shock wave are Ma3 and P 03 , respectively; the subsonic flow total pressure P 03 is collected and recorded by the pressure scanning valve, and the specific value of the subsonic flow total pressure P 03 after time-averaging treatment;

[0037] S3), the post-shock total pressure measured by the total pressure probe 1 is the total pressure after the oblique shock wave and the normal shock wave, and the actual airflow total pressure P 01 before the shock wave at the probe position is calculated through the shock wave angle β of the oblique shock wave; the specific calculation is as follows:

[0038] The Mach numbers Ma1 and Ma2 are calculated by the following formula in the present embodiment, i.e.:

[0039]

[0040] In the formula, β is the included angle between the oblique shock wave and the horizontal plane, i.e. the shock wave angle; γ is a preset parameter, and γ = 1.4 in the present embodiment; θ is the included angle between the wedge structure and the horizontal plane, θ = 20 degrees.

[0041] Then, P 01 is obtained according to the total pressure relationship before and after the normal shock wave and the oblique shock wave.01 and P 02 ; that is:

[0042]

[0043] In the formula, β is the included angle between the oblique shock wave and the horizontal plane, that is, the shock wave angle; γ is a preset parameter.

[0044] As preferred in the embodiment, the total pressure probe 1 is arranged in a vertical array to simultaneously measure total pressure at multiple points.

[0045] The embodiments and the description in the specification are only illustrative of the principles and the best mode of the present application, and various changes and modifications can be made to the present application without departing from the spirit and the scope of the present application, and such changes and modifications fall within the scope of the present application.

Claims

1. A device for measuring the total pressure loss in a supersonic internal flow field, characterized in that, It includes a total pressure probe (1), a support rod (2), and a wind tunnel base plate (3); the total pressure probe (1) is mounted on the wind tunnel base plate (3) via the support rod (2); The total pressure probe (1) is provided with an airflow delivery channel inside. The front end of the total pressure probe (1) is a wedge structure, and a pressure measuring hole (11) is provided at the center of the wedge structure. The rear end of the total pressure probe (1) is provided with an airflow outlet (12), and the airflow outlet (12) is connected to the pressure scanning valve through a hose. The method for measuring total pressure loss using the total pressure probe (1) is as follows: S1) Assuming the Mach number and total pressure at the upstream end of the total pressure probe (1) are respectively , After passing through the oblique shock wave at the leading edge of the wedge, the flow Mach number and total pressure are respectively , ; S2) After the airflow enters the total pressure probe (1), since the flow is still supersonic at this time, a normal shock wave is generated inside the airflow delivery channel of the total pressure probe (1), so that the flow is reduced to subsonic speed. The Mach number and total pressure of the flow after the normal shock wave are , Subsonic flow total pressure The pressure is collected and recorded by the pressure scanning valve, and after time-averaging processing, the subsonic flow total pressure is obtained. The specific value; Mach number is calculated using the following formula. and ,Right now: In the formula, The angle between the oblique shock wave and the horizontal plane is called the shock wave angle. These are preset parameters. The angle between the wedge structure and the horizontal plane; S3), The total pressure behind the shock wave measured by the total pressure probe (1) is the total pressure behind the two shock waves, the oblique shock wave and the normal shock wave, and is measured by the shock wave angle of the oblique shock wave. Calculate the actual total pressure of the airflow at the probe. Based on the total pressure relationship before and after the normal and oblique shock waves, the following equations are obtained: and ;Right now: In the formula, The angle between the oblique shock wave and the horizontal plane is called the shock wave angle. These are preset parameters.

2. The measuring device for total pressure loss in a supersonic internal flow field according to claim 1, characterized in that: The wedge structure is horizontally symmetrical.

3. The measuring device for total pressure loss in a supersonic internal flow field according to claim 1, characterized in that: The angle between the wedge structure and the horizontal plane Spend.

4. The measuring device for total pressure loss in a supersonic internal flow field according to claim 1, characterized in that: The length and width of the pressure measuring hole (11) are both 0.5 mm.

5. The measuring device for total pressure loss in a supersonic internal flow field according to claim 1, characterized in that: The total pressure probes (1) are arranged in a vertical array to measure the total pressure at multiple points simultaneously.

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