CN110806300A - Measuring point arrangement method suitable for hypersonic flight test transition research - Google Patents
Measuring point arrangement method suitable for hypersonic flight test transition research Download PDFInfo
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Abstract
A measuring point arrangement method suitable for a transition study of a hypersonic flight test is realized by the following steps: s1, determining whether to measure natural transition or forced transition according to the measurement requirement, and if so, transitioning to S2; if yes, go to S3; s2, measuring a main flow transition condition and/or a transition condition of a cross flow effect according to measurement requirements, wherein when the main flow transition condition is measured, measuring points are arranged on a flow line of the hypersonic aircraft in the main flow direction; when the transition condition of the cross flow effect is measured, the measuring points are arranged at the position with the cross flow velocity in the lateral direction; the main flow direction is the central streamline direction of the aircraft and the streamline direction with an included angle of not more than 3 degrees; s3, arranging a rough element at a predetermined position on the aircraft, and arranging a measuring point at the downstream of a streamline where the rough element is located; the measuring point positions are used for measuring the surface physical quantity of the aircraft in the flight test process by installing sensors.
Description
Technical Field
The invention belongs to the field of transition research of aerospace craft.
Background
The transition is a process of transition of a flow state from laminar flow to turbulent flow, and the physical mechanism of the transition is that with the increase of a reynolds number, laminar flow starts to be unstable, so that disturbance in the flow continuously increases, and then complex nonlinear disturbance occurs, a vortex structure and vortex breakage are generated, and finally the flow is changed into a chaotic state (namely a turbulent flow state). When the boundary layer is twisted, the surface of the aircraft is transited from laminar heating to turbulent heating, and the surface heat flow is multiplied, so that the surface thermal environment is obviously changed. Fig. 1 shows the surface heat flow change of the reentry vehicle when transition occurs, and it can be seen that when transition occurs, the surface heat flow increases by about 3 times. Therefore, accurate prediction of transition is very important in engineering design, and if the transition position is earlier than the predicted value, the heat protection layer of the aircraft may be burned through, otherwise, the thickness of the heat protection layer is unnecessarily increased.
Transition is a complex flow phenomenon, and various factors may induce transition. For transition, people's understanding is incomplete, transition theory is incomplete, transition analysis and simulation methods are not mature, and particularly in the hypersonic velocity field, due to the coupling of the problem of aerodynamic heat and the difficulty of ground test simulation, relevant theories and methods are immature. The understanding of the nature of transition and the related engineering practices show that the transition problem must be solved on the engineering basis based on experimental data. Through analysis of test data, the internal mechanism of transition can be revealed, the physical rule of transition can be recognized, the transition physical model for engineering design can be refined and summarized, and an effective transition analysis prediction method can be constructed.
In the prior art, hypersonic transition measurement is mainly based on wind tunnel tests. The conventional shock wind tunnel test has strong background noise, and the noise has a significant influence on transition, so that transition information obtained by the conventional shock wind tunnel test is obviously distorted and cannot reflect a real flight condition, and test data only have reference value in comparison research; the hypersonic static wind tunnel controls strong noise in a conventional wind tunnel in principle, and is considered to be capable of approaching to a real flight environment, but the conventional hypersonic static wind tunnel only has the maximum test Mach number of 6, the maximum nozzle caliber can only reach 300mm, the total pressure of the maximum parking chamber in a silent state can only reach 1MPa, the inflow environment still has space difference, and the pneumatic heating condition is different from the flight state, so that the transition research thereof has obvious limitation. In addition, the difference between the ground wind tunnel and the flight test is obvious, and the common platinum film heat flow sensor, the flow field display technology and the like cannot be applied to the flight test.
The flight test transition measurement mainly comprises two types according to the measured physical quantity, namely a thermal parameter sensor and a pressure parameter sensor, wherein the thermal parameter sensor is divided into a heat flow sensor and a temperature sensor. The heat flow sensor and the temperature sensor can obtain a transition starting position and a transition ending position, and are mainly used for acquiring transition process and transition array surface information; the pressure measurement mainly adopts a high-frequency pulsating pressure sensor for obtaining the disturbance condition of a flow field near the real surface of the aircraft.
In recent years, the hypersonic technology is developed vigorously, the demand for transition research is more and more urgent, transition measurement data is obtained through flight tests, the internal mechanism of transition is further revealed based on the test data, the physical regularity of transition is recognized, and an effective transition analysis prediction method is constructed very urgently. The invention provides a method suitable for the arrangement of a measuring point for the transition research of a hypersonic flight test, and provides an important guarantee for the transition research of the flight test.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: aiming at the conditions that the transition theoretical research of the hypersonic aircraft is insufficient and the ground wind tunnel test cannot meet the transition research requirement, the transition data measured by the flight test is needed to provide support for the transition theoretical research and the method improvement, the invention provides the measuring point arrangement method suitable for the transition research of the hypersonic aircraft, and the method can be effectively used for measuring point arrangement during the transition research.
The technical solution of the invention is as follows: a measuring point arrangement method suitable for a transition study of a hypersonic flight test is realized by the following steps:
s1, determining whether to measure natural transition or forced transition according to the measurement requirement, and if so, transitioning to S2; if yes, go to S3;
s2, measuring a main flow transition condition and/or a transition condition of a cross flow effect according to measurement requirements, wherein when the main flow transition condition is measured, measuring points are arranged on a flow line of the hypersonic aircraft in the main flow direction; when the transition condition of the cross flow effect is measured, the measuring points are arranged at the position with the cross flow velocity in the lateral direction;
the main flow direction is the central streamline direction of the aircraft and the streamline direction with an included angle of not more than 3 degrees;
s3, arranging a rough element at a predetermined position on the aircraft, and arranging a measuring point at the downstream of a streamline where the rough element is located;
the measuring point positions are used for measuring the surface physical quantity of the aircraft in the flight test process by installing sensors.
Preferably, the sensor mounting must meet the mounting requirement of 1/8 that the amount of protrusion/recession relative to the aircraft surface is less than the thickness of the boundary layer at the corresponding location at the time of transition occurrence.
Preferably, the sensor is mounted as flush as possible with respect to the aircraft surface, preferably in the form of a recessed mounting if this cannot be achieved.
Preferably, the measuring point of S2 is arranged at a position having a cross flow velocity in the lateral direction, specifically, at a position where the cross flow velocity is greater than 5% to 10% of the total velocity.
Preferably, in S3, the corresponding measuring points are arranged at symmetrical positions of the arranged measuring points relative to the plane of symmetry of the aircraft as a comparison, or the measuring points are arranged in the upstream direction of the coarse element on the same streamline as the comparison.
Preferably, the distance between the measuring points arranged in the upstream direction and the rough element is 1-2.5 times of the requirement of the minimum distance between the adjacent measuring points.
Preferably, when the number of the measuring points arranged on the same flow line exceeds 2, the distance between the adjacent measuring points is more than 2.5 times of the requirement of the minimum distance.
Preferably, the minimum distance is determined according to the streamline of the adjacent measuring point;
when the measuring points are arranged on a mainstream direction streamline, randomly selecting one point from the streamline, calculating the flowing direction most unstable wave frequency and the flowing direction speed of the point, and obtaining the minimum distance requirement by dividing the flowing direction speed by the most unstable wave frequency;
when the measuring points are arranged at the position with the lateral cross flow velocity, one measuring point is randomly selected from the cross flow streamline, the frequency of the most unstable wave flowing in the spanwise cross flow and the cross flow velocity of the point are calculated, and the minimum distance requirement is obtained by dividing the cross flow velocity by the frequency of the most unstable wave flowing in the spanwise cross flow.
Preferably, the physical quantity on the surface of the aircraft comprises heat flow, pulsating pressure and temperature, a sensor is correspondingly arranged at the position of a measuring point, and the sensitive end of the sensor is directly contacted with incoming air.
Preferably, when the transition to be measured initially occurs on the aircraft, the measuring points are arranged at the downstream position of the surface limit streamline of the aircraft as far as possible on the premise of meeting the installation requirement of the aircraft structure; if the transition to be measured advances to the front position, the measuring point arrangement position is determined in the following way:
firstly, theoretically calculating the Reynolds number of the initial transition free incoming flow when the transition advances to the front;
then, when the Reynolds number of the starting transition free incoming flow is expanded by 1.4-2 times, the transition is recalculated to advance to the front position, and the measuring point is arranged at the position.
Preferably, the deviation between the thermal physical property parameter of the sensor and the thermal physical property parameter of the surface of the aircraft is within 30 percent, and the thermal physical property parameter is
Wherein rho is the density of the sensor material, c is the specific heat capacity of the sensor material, and k is the heat conduction coefficient of the sensor material.
Compared with the prior art, the invention has the beneficial effects that:
(1) the transition research is a complex problem, an effective measuring point arrangement method is provided, and the basic transition phenomenon can be researched.
(2) The transition measuring point installation requirement based on boundary layer thickness constraint is provided, and the method is feasible, high in measurement accuracy and strong in operability.
(3) Multiple times of practice show that the initial position of the aircraft natural transition, the transition surface morphology, the transition region and the incoming flow disturbance information are obtained according to the measuring point arrangement provided by the invention, and good flight test data are provided for transition research.
(4) The installation of the sensor needs to meet the installation requirement that the protrusion/depression amount of the surface of the aircraft is smaller than 1/8 of the thickness of the boundary layer at the corresponding position at the transition occurrence moment, the limitation of the requirement can ensure that the local disturbance caused by the installation of the sensor does not cause the change of the flow state, and the truth of the natural transition measurement information is ensured; if the required values in the claims are exceeded, the transition measurement information may be distorted by local disturbances caused by the sensor installation causing changes in the downstream flow regime.
(5) The deviation between the thermal physical property parameter of the sensor and the thermal physical property parameter of the surface of the aircraft is within 30%, so that the measured temperature of the sensor and the temperature of the surface material of the aircraft can be basically equal, and the occurrence and judgment of transition cannot be influenced due to flow change caused by local temperature inconsistency.
Drawings
FIG. 1 is a diagram illustrating a typical change in surface heat flow when transition occurs;
FIG. 2 is a schematic diagram of the arrangement of transition points in the invention;
FIG. 3 is a schematic view of the arrangement of a small number of stations;
FIG. 4 is a schematic diagram of an arrangement of a large number of stations.
Detailed Description
The invention is described in detail below with reference to the figures and examples.
And (5) determining the measurement requirement and the number of the measuring points. If the natural transition or the forced transition is measured, the transition situation of the main flow transition or the transition situation of the cross flow effect is researched; the number of the measuring points is mainly limited by the number of channels of the data acquisition equipment, the data transmission bandwidth, the aircraft installation space and the like. The surface sensor has the advantage of fast response time, and the surface sensor is preferably used for transition measurement.
If the natural transition is measured, the protrusion/depression amount of the sensor installation is controlled, the protrusion/depression amount of the relative surface of the sensor is required to be smaller than 1/8 of the thickness of the boundary layer at the corresponding position at the transition occurrence time, the sensor is installed in a flush mode as much as possible, and if the flush installation cannot be achieved, the depression installation mode is preferably selected.
Measuring the transition of the main stream direction, and arranging the measuring points in the main stream direction (on the central line) as much as possible; when the transition of the cross flow effect is measured, the measuring points are arranged at the position with obvious cross flow velocity (the cross flow velocity is more than 5-10% of the total velocity) in the lateral direction.
When the forced transition is measured, a sensor should be arranged downstream (behind) the forced transition device; if the number of the measuring points is rich, corresponding measuring points can be arranged at symmetrical positions for comparison, and measuring points can also be arranged in front of the rough element for comparison, as shown in FIG. 2.
When the number of the measuring points is not large, the measuring points can be arranged sparsely. If transition needing to be measured initially occurs on the aircraft, the measuring points are arranged at the rear position as much as possible; if the transition to be measured is promoted to the front position, the measuring points can be arranged at the front position of the aircraft according to the prediction condition and considering the deviation of the transition occurrence height of 3 km-5 km;
the disturbance is inevitably introduced to the arrangement of the measuring points, when the number of the measuring points is large, the transition caused by the disturbance caused by the arrangement of the measuring points is prevented as much as possible, and the disturbance introduced by the arrangement of the measuring points is required to be far away from the most unstable disturbance of the boundary layer. If a certain flow direction unsteady wave frequency is about 200kHz, a spanwise cross flow unsteady wave frequency is about 50kHz, and corresponding flow direction velocity and cross flow velocity are about 2000m/s and 200m/s, respectively, then the corresponding flow direction and spanwise spacing of the sensors should be 10mm and 4mm apart, respectively, and in practice should be more than 2.5 times the minimum spacing.
Cross flow velocity: the component of the local velocity of the flow field in the direction perpendicular to the direction of the outer edge of the boundary layer and the normal direction of the wall. The surface limit streamline is generally solved by a CFD (computational fluid dynamics) method, and the surface limit streamline is obtained by extracting a velocity distribution of a layer of mesh near an object plane according to a CFD calculation result.
And (3) analyzing the boundary layer of the local laminar flow by adopting a linear stability method according to the frequency of the most unstable wave at a specific position on the surface of the aircraft, wherein the obtained disturbance wave with the maximum growth rate is the most unstable wave, and the corresponding disturbance wave frequency is the most unstable wave frequency.
Surface sensor: the sensor sensitive end is positioned on the surface of the aircraft and is in direct contact with the outside air, and the sensor sensitive end is positioned in the aircraft structure and cannot be in direct contact with the outside air. Through-the-wall sensors are those that are installed by punching through the aircraft surface, which is one of the most common ways of installing surface sensors.
Example 1: arrangement of transition detection points
Transition information concerned by a small number of measuring points is limited, the conventional method mainly comprises main flow direction measuring point arrangement and cross flow direction measuring point arrangement, and the positions and directions of the measuring point arrangement generally depend on the requirements of transition research.
The wall-penetrating sensor can directly measure the physical quantity of the surface of the aircraft, the wall-penetrating sensor is selected to carry out transition measurement, meanwhile, the thermophysical parameters including density, specific heat capacity, heat conduction coefficient and the like of the sensor are consistent with or close to the surface of the aircraft as much as possible, and the evolution of the thermophysical parameters is smaller than 30% compared with the value of the surface material of the aircraft.
Of the four measurement points shown in fig. 3, measurement point ① and measurement point ② are mainstream direction measurement points, and measurement point ③ and measurement point ④ are cross-flow effect study measurement points.
When the transition condition in the main flow direction is researched, the measuring points are arranged on the central line of the windward side as much as possible (the cross flow speed is close to 0), such as the measuring point ① and the measuring point ②, when the transition starting position is concerned, the measuring points are arranged at the rear position as much as possible under the condition that constraint conditions such as structure installation are considered, when the transition advancing position in the flight time period is concerned, the measuring points are arranged at the front position as much as possible under the condition that constraint conditions such as structure installation are considered, and the position where the transition is possible can be determined according to the transition prediction result and considering a certain deviation.
When the transition of the cross flow effect is studied, the measuring points should be arranged at places where the cross flow effect is obvious (generally, a position where the cross flow velocity is 5% -10% higher than the main flow velocity is selected, and the position can be obtained through flow field numerical simulation), such as the measuring point ③ and the measuring point ④, wherein the measuring point ③ and the measuring point ④ correspond to the cross flow effect of different degrees and can be arranged appropriately according to research requirements.
The amount of relief in the sensor mounting should be less than the boundary layer thickness constraint (one proposed method is to mount the sensor with a relief less than the boundary layer thickness 1/8), and if possible, to make the sensor surface as flush with the aircraft surface, and if difficult to achieve, to preferentially recess the sensor relative to the aircraft surface.
Example 2: arrangement of a large number of transition measuring points
As shown in fig. 4, in the case of a large number of measurement points, the measurement points are arranged appropriately to obtain sufficient transition information, such as a transition occurrence start position, a transition region length, a transition front morphology and the like, and generally at least 3 to 4 measurement points are arranged in the transition region;
the through-wall sensor can directly measure the physical quantity of the surface of the aircraft, the through-wall sensor is selected to carry out transition measurement, meanwhile, the thermophysical parameters of the sensor, such as density, specific heat capacity, heat conduction coefficient and the like, are consistent with or close to the surface of the aircraft as much as possible, and the evolution of the thermophysical parameters is smaller than 30% compared with the value of the surface material of the aircraft.
The number of the measuring points is large, the sensors which are small in radius and good in shape with the surface of the aircraft are preferably installed, and the surfaces of the sensors can be polished according to the use requirements of the sensors when necessary.
For the main flow direction, the distance between the measuring points should be far away from the wavelength range corresponding to the most unstable frequency, taking the frequency of the wave flowing to the most unstable wave about 200kHz and the flowing speed about 2000m/s as an example, the distance between the sensors in the flowing direction should be far away from 10mm, and the installation distance of the sensors should be larger than 25mm as required; considering that the length of the transition region of the hypersonic transition is about 200-300 mm, at least 3-4 measuring points are arranged in the transition region, the minimum distance of sensor installation, the number of the measuring points in the transition region and the influence of the sensor installation on the structure are comprehensively considered, and the distance of the sensor flow direction is selected to be 50 mm.
For the span-wise sensors, taking the frequency of the cross-flow flowing unstable wave as about 50kHz and the flow direction velocity of the cross-flow as about 200m/s as an example, the span-wise spacing corresponding to the sensors should be far away from 4mm, the installation spacing of the sensors should be more than 10mm as required, and the arrangement spacing of the sensors is selected to be 50mm in consideration of the influence of installation on the structure.
The amount of relief in the sensor mounting should be less than the boundary layer thickness constraint (one proposed method is to mount the sensor with a relief less than the boundary layer thickness 1/8), and if possible, to make the sensor surface as flush with the aircraft surface, and if difficult to achieve, to preferentially recess the sensor relative to the aircraft surface.
The invention has not been described in detail in part of the common general knowledge of those skilled in the art.
Claims (11)
1. A measuring point arrangement method suitable for transition research of a hypersonic flight test is characterized by being realized in the following mode:
s1, determining whether to measure natural transition or forced transition according to the measurement requirement, and if so, transitioning to S2; if yes, go to S3;
s2, measuring a main flow transition condition and/or a transition condition of a cross flow effect according to measurement requirements, wherein when the main flow transition condition is measured, measuring points are arranged on a flow line of the hypersonic aircraft in the main flow direction; when the transition condition of the cross flow effect is measured, the measuring points are arranged at the position with the cross flow velocity in the lateral direction;
the main flow direction is the central streamline direction of the aircraft and the streamline direction with an included angle of not more than 3 degrees;
s3, arranging a rough element at a predetermined position on the aircraft, and arranging a measuring point at the downstream of a streamline where the rough element is located;
the measuring point positions are used for measuring the surface physical quantity of the aircraft in the flight test process by installing sensors.
2. The method of claim 1, wherein: the sensor mounting must meet the mounting requirement of 1/8 that the amount of protrusion/recession relative to the aircraft surface is less than the thickness of the boundary layer at the corresponding location at the time of the transition.
3. The method of claim 2, wherein: the sensor is mounted flush with respect to the aircraft surface as far as possible, preferably in a recessed manner if flush mounting is not possible.
4. The method of claim 1, wherein: the measuring point of S2 is arranged at the position with transverse flow velocity in the lateral direction, specifically, the position with transverse flow velocity greater than 5% -10% of the total velocity.
5. The method of claim 1, wherein: and S3, arranging corresponding measuring points at symmetrical positions of the arranged measuring points relative to the symmetrical plane of the aircraft as comparison, or arranging the measuring points in the upstream direction of the rough element on the same streamline as the comparison.
6. The method of claim 5, wherein: the distance between the measuring points arranged in the upstream direction and the rough element is 1-2.5 times of the requirement of the minimum distance between the adjacent measuring points.
7. The method of claim 1, wherein: when the number of the measuring points arranged on the same flow line exceeds 2, the distance between the adjacent measuring points is more than 2.5 times of the requirement of the minimum distance.
8. The method according to claim 6 or 7, characterized in that: the minimum distance is determined according to the streamline of the adjacent measuring point;
when the measuring points are arranged on a mainstream direction streamline, randomly selecting one point from the streamline, calculating the flowing direction most unstable wave frequency and the flowing direction speed of the point, and obtaining the minimum distance requirement by dividing the flowing direction speed by the most unstable wave frequency;
when the measuring points are arranged at the position with the lateral cross flow velocity, one measuring point is randomly selected from the cross flow streamline, the frequency of the most unstable wave flowing in the spanwise cross flow and the cross flow velocity of the point are calculated, and the minimum distance requirement is obtained by dividing the cross flow velocity by the frequency of the most unstable wave flowing in the spanwise cross flow.
9. The method of claim 1, wherein: the aircraft surface physical quantity comprises heat flow, pulsating pressure and temperature, a sensor is correspondingly arranged at a measuring point, and the sensitive end of the sensor is directly contacted with incoming air.
10. The method of claim 1, wherein: when transition needing to be measured initially occurs on the aircraft, the measuring points are arranged at the downstream position of the surface limit streamline of the aircraft as far as possible on the premise of meeting the installation requirement of the aircraft structure; if the transition to be measured advances to the front position, the measuring point arrangement position is determined in the following way:
firstly, theoretically calculating the Reynolds number of the initial transition free incoming flow when the transition advances to the front;
then, when the Reynolds number of the starting transition free incoming flow is expanded by 1.6 to 2 times, the transition is recalculated to advance to the front position, and the measuring point is arranged at the position.
11. The method of claim 1, wherein: the deviation between the thermal physical property parameter of the sensor and the thermal physical property parameter of the surface of the aircraft is within 30 percent, and the thermal physical property parameter is
Wherein rho is the density of the sensor material, c is the specific heat capacity of the sensor material, and k is the heat conduction coefficient of the sensor material.
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