CN111780948B - Method for measuring transition process characteristic of aircraft boundary layer in hypersonic flight test - Google Patents
Method for measuring transition process characteristic of aircraft boundary layer in hypersonic flight test Download PDFInfo
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Abstract
The invention provides a method for measuring the transition process characteristic of an aircraft boundary layer in a hypersonic flight test, which comprises the following steps: 1) carrying out simulation calculation on a flow field around the aircraft to obtain the flow characteristic of the surface of the aircraft; 2) determining a region of the surface of the aircraft suitable for boundary layer transition process measurement; 3) carrying out simulation calculation on the surface heat flow and the structural thermal response of the aircraft, obtaining the calculation results of the surface heat flow and the structural temperature of the aircraft along the whole flight profile, and carrying out model selection on a heat flow sensor and a temperature sensor; 4) evaluating the feasibility of sensor installation in the transition measurement region; 5) determining the positions of the sensors, the number of the sensor measuring points and the distance between the sensor measuring points according to the analysis result of the surface flow and transition characteristics of the aircraft; 6) analyzing flight test data obtained by the transition process measurement scheme determined in the steps 1) -5), and drawing a curve of surface heat flow or temperature along the change of the flow direction; 7) and analyzing the change rule of the heat flow or temperature along the flow change curve to determine the transition process characteristic of the boundary layer.
Description
Technical Field
The invention relates to a method for measuring the transition process characteristic of an aircraft boundary layer in a hypersonic flight test, and belongs to the field of aerodynamic force/heat research of hypersonic aircraft.
Background
The hypersonic aircraft spans a wider flight airspace and a wider speed domain, and the surface flow of the hypersonic aircraft can generate boundary layer transition along with the increase of the Reynolds number of the incoming flow. Because the wall friction coefficient/heat exchange coefficient of the laminar boundary layer and the turbulent boundary layer are obviously different, the boundary layer transition has obvious influence on the aerodynamic force of the surface of the aircraft, particularly the aerodynamic heat distribution. In the transition propulsion process, the aerodynamic and control characteristics of the aircraft are obviously changed, the surface heat flow is rapidly increased after the transition occurs, and the turbulent heating is 3-5 times that of laminar heating in a typical hypersonic speed state. Therefore, the boundary layer transition is an important factor that must be considered in the design of the hypersonic aircraft.
The hypersonic speed boundary layer transition is a complex flow phenomenon, the mechanism is complex, the influence factors are numerous, the transition theory is still incomplete at present, the transition prediction method is quite insufficient, and the acquisition of the information related to the transition of the boundary layer through a ground/flight test is an important means for verifying and perfecting the transition theory and the prediction method and promoting the transition recognition level to be improved. The boundary layer transition measurement technology in the ground wind tunnel test is relatively mature, generally, the surface heat flow, temperature or pulsating pressure data of the aircraft are measured, and the data are utilized to analyze so as to obtain the boundary layer transition information. The hypersonic speed real flight environment is significantly different from the ground wind tunnel inflow environment, the aircraft is subjected to severe and complex pneumatic heating, the thermal response of the surface of the aircraft is far higher than that of a ground wind tunnel test, and the arrangement and installation of the test sensors are limited by factors such as aircraft structure and space requirements, so that the difficulty of transition measurement of the development boundary layer in the hypersonic speed flight test is far higher than that of the ground wind tunnel test. At present, a small amount of scattered heat flow, temperature or pulsating pressure information can be generally obtained only through a plurality of single-point measurements in a hypersonic flight test, and analysis of the information can only obtain the starting point information of transition of a boundary layer, so that the complete transition occurrence and development process of the starting transition, the transition development (transition region) and the ending transition of the boundary layer is difficult to obtain, and the research of a transition theory and a prediction method is restricted to a great extent.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method for measuring the boundary layer transition process information in the hypersonic flight test overcomes the defects of the prior art, aims at the problem that the boundary layer transition process information is difficult to obtain in the hypersonic flight test at present, solves the problem that the current technical means can only obtain a small amount of single boundary layer transition information, can obtain complete transition process information including a boundary layer transition starting point, a transition region and a transition finishing point in the hypersonic flight test, and provides more abundant test data for the boundary layer transition theory and the prediction method research.
The technical solution of the invention is as follows: a method for measuring the transition process characteristic of an aircraft boundary layer in a hypersonic flight test comprises the following steps:
1) carrying out simulation calculation on a flow field around the aircraft, and analyzing the distribution of surface streamlines and surface flow parameters through post-processing of the flow field to obtain the surface flow characteristics of the aircraft;
2) determining a region suitable for boundary layer transition process measurement on the surface of the aircraft according to the flow characteristics of the surface of the aircraft obtained in the step 1);
3) carrying out simulation calculation on the surface heat flow and the structure thermal response of the aircraft by using a CFD numerical method and a structure thermal analysis method, obtaining the calculation result of the surface heat flow and the structure temperature of the aircraft along the whole flight section, and carrying out model selection on a heat flow sensor and a temperature sensor according to the surface heat flow and the surface temperature data of a transition measurement area, wherein the model selection comprises the measuring range, the size and the material of the sensor;
4) comprehensively analyzing the size and the installation requirements of the sensor and the structural characteristics of the aircraft near the transition measurement region and the constraint of the installation space, and evaluating the installation feasibility of the sensor in the transition measurement region, wherein if the installation of the heat flow sensor and the temperature sensor is feasible, the heat flow sensor is preferably selected for surface heat flow measurement, and if only the installation of the temperature sensor is feasible, the temperature sensor is selected for surface temperature measurement;
5) drawing a surface streamline of a transition measurement area according to an analysis result of the surface flow characteristic of the aircraft, and arranging a row of sensor measuring points along the direction of the surface streamline; evaluating the transition characteristic of the surface boundary layer of the aircraft to obtain a transition region length calculation result L of the transition measurement regioncalDetermining the number of the sensor measuring points and the distance between the sensor measuring points by combining with the transition measurement precision requirement;
6) analyzing flight test data obtained by the transition process measurement scheme determined in the steps 1) -5), and drawing a curve of surface heat flow or temperature along the change of the flow direction;
7) and analyzing the change rule of the heat flow or the temperature along the flow change curve, wherein the information of the transition starting position, the transition region and the transition ending position is the boundary layer transition process characteristic.
And in the step 1), a CFD numerical method is used for carrying out simulation calculation on the flow field around the aircraft.
In step 2), the method for determining the transition measurement region is as follows:
2a) the measuring areas are positioned in the same profile of the aircraft;
2b) the flow in the measurement area is attached flow, the flow characteristics are simple, and the complex flow characteristics of separation reattachment and shock wave-boundary layer interference do not exist;
2c) the flow characteristics of the measuring region are unchanged in the flow direction and are not disturbed by the flow of surrounding components.
When the temperature sensor is selected for surface temperature measurement, the local cabin material and the structure of the aircraft are designed, so that the cabin material and the thickness are kept consistent in the transition measurement region, and the thermal response characteristic of the structure in the transition measurement region is guaranteed to be consistent.
And 5) evaluating the transition characteristic of the surface boundary layer of the aircraft by using a transition prediction method or a ground test result.
In the step 5), when the number of the sensor measuring points and the distance between the sensor measuring points are determined, the sensor measuring points are ensured to cover the whole transition process, that is, the flow direction length covered by all the measuring points is not less than 2.0LcalThe distance between measuring points is not more than 1/5Lcal。
The specific process of the step 6) is as follows: firstly, preprocessing heat flow or temperature measurement data is carried out, invalid data including abnormal jumping points, dead points and burrs are identified and removed, and then curves of heat flow or temperature change along the flow direction at different moments are drawn by taking the flow direction length as an abscissa and the heat flow or temperature as an ordinate.
In the step 7), the determination method of the transition start position, the transition region, and the transition end position specifically includes:
when the heat flow/temperature is along the flow direction change curve, if the heat flow/temperature continues to be obviously increased after a certain measuring point position, the position is determined as a transition starting position at the moment; analyzing a heat flow/temperature curve behind the transition starting position, and if the heat flow/temperature does not rise or starts to fall after a certain measuring point position, judging the position as the transition ending position at the moment; the region between the start position and the end position of transition is the transition region of transition at the moment, and the flow direction length L is usedexpCarrying out characterization; the information of the transition starting position, the transition region and the transition ending position determined in this way is the boundary layer transition process characteristic of the windward side of the aircraft.
Compared with the prior art, the invention has the beneficial effects that:
the invention provides a method for measuring the transition process characteristic of a boundary layer in a hypersonic flight test, which overcomes the problem that the prior technical means can only obtain a small amount of single transition information of the boundary layer, and can measure the complete transition process information of the transition start, transition area and end of the boundary layer on the surface of an aircraft in the hypersonic flight test;
the invention provides a determination method of a transition process measurement region, which is based on the physical characteristic of surface flow, so that the determined measurement region is more reasonable, the actual transition physical process can be more accurately described, and the quality of a measurement result is higher;
the invention provides a solution for transition process measurement under the condition of selecting a temperature sensor, which can ensure the consistency of the thermal response characteristics of the structure in a measurement area through the design of the structure and the material of an aircraft when the heat flow measurement condition is not available, thereby using the temperature sensor to perform transition process measurement and widening the transition measurement means;
the invention provides a method for arranging the measuring points of a sensor, wherein the measuring points are arranged along the surface streamline direction, and the measuring point distance is determined by theoretical analysis or a ground test result, so that the measurement result in the transition process can be more reasonable, and the measurement precision is higher.
The measuring method provided by the invention can obtain richer boundary layer transition information, and test data verification can be provided for boundary layer transition research by utilizing the information, so that the transition phenomenon recognition level and the transition prediction method accuracy can be further improved.
Drawings
FIG. 1 is a schematic exterior view of a typical aircraft;
FIG. 2 is a schematic view of the aircraft windward surface flow line;
FIG. 3 is a schematic view of a transition measurement region and a measurement point arrangement scheme of the aircraft;
FIG. 4 is a schematic diagram illustrating the boundary layer transition process information discrimination using heat flow and temperature variation curves; fig. 4(a) is a heat flow change curve, and fig. 4(b) is a temperature change curve.
Detailed Description
The following describes embodiments of the present invention in further detail with reference to the accompanying drawings.
The specific implementation of the invention is illustrated by taking the measurement of the transition process of the boundary layer of the windward side of a typical-shape aircraft as an example. FIG. 1 is a schematic exterior view of an aircraft in a typical "fuselage + wing" configuration, the fuselage being formed by two sections, section 1 being located at the front of the fuselage and section 2 being located at the rear of the fuselage; the profiles of the two sections are not uniform and a protrusion (such as a control rudder) is located on section 2. The method for measuring the information of the transition process of the boundary layer of the hypersonic flight test aircraft comprises the following implementation steps:
1) firstly, a CFD numerical method (known in the art) is used to perform simulation calculation on a flow field around the aircraft, and surface streamlines and surface flow parameter (such as heat flow, pressure, temperature, etc.) distribution are analyzed by flow field post-processing software (known in the art), so as to obtain the surface flow characteristics of the aircraft, for the aircraft of this embodiment, as shown in fig. 2, the surface flow characteristics are: the flow of the area near the center of the section 1 and the section 2 is adhesion flow, and the flow characteristics are simple; due to the flow interference of the wings and the protrusions, the local areas of the section 2 close to the wings and the protrusions may have phenomena of separation reattachment, shock wave-boundary layer interference and the like, and the flow characteristics are relatively complex, and the areas of the section 2 affected by the wings and the protrusions are marked in fig. 2;
2) determining a region suitable for boundary layer transition process measurement on the surface of the aircraft according to the aircraft surface flow characteristics obtained in the step 1), as shown in fig. 3, the transition measurement region determining method includes:
a) the measurement areas are located in the same profile of the aircraft, namely the measurement areas are respectively located on the section 1 and the section 2 and cannot cross the section 1 and the section 2 of the aircraft;
b) the flow in the measurement area is attached flow, the flow characteristics are simple, and complex flow characteristics such as separation reattachment, shock wave-boundary layer interference and the like cannot exist, namely the measurement area cannot be positioned in the wing influence area and the bulge influence area;
c) the flow characteristics of the measuring region are unchanged along the flow direction and are not disturbed by the flow of surrounding components, i.e. the measuring region should be located completely within the central vicinity of the section 1 or the section 2 and cannot span the wing and bulge influence regions;
3) carrying out simulation calculation on the surface heat flow and the structure thermal response of the aircraft by using a CFD (computational fluid dynamics) numerical method and a structure thermal analysis method (known technology in the field), obtaining the calculation result of the surface heat flow and the structure temperature of the aircraft along the whole flight section, and carrying out model selection on a heat flow sensor and a temperature sensor according to the surface heat flow and the surface temperature data of a transition measurement area, wherein the model selection comprises the measuring range, the size, the material and the like of the sensor;
4) comprehensively analyzing the size and the installation requirements of the sensor and the structural characteristics of the aircraft near the transition measurement region and the constraint of the installation space, and evaluating the installation feasibility of the sensor in the transition measurement region, wherein if the installation of the heat flow sensor and the temperature sensor is feasible, the heat flow sensor is preferably selected for surface heat flow measurement, and if only the installation of the temperature sensor is feasible, the temperature sensor is selected for surface temperature measurement;
5) according to the sensor selection result in the step 4), if a temperature sensor is selected for surface temperature measurement, the local cabin material and structure of the aircraft need to be designed, so that the cabin material and the thickness are kept consistent in the transition measurement region range, and the thermal response characteristic of the structure in the transition measurement region is guaranteed to be consistent;
6) drawing a surface streamline of a transition measurement area (shown in FIG. 2) according to an analysis result of the surface flow characteristic of the aircraft, and arranging a row of sensor measuring points (shown in FIG. 3) along the direction of the surface streamline; evaluating the transition characteristic of the aircraft surface boundary layer by using a transition prediction method (such as transition engineering prediction method, transition mode method, etc., which are well known in the art) or a ground test result, and calculating the length of the transition region according to the transition measurement region transitioncalAnd determining the number and the interval of the sensor measuring points according to the transition measurement precision requirement, wherein the number of the measuring points needs to cover the whole transition process, and the flow direction length covered by the recommended measuring points is not less than 2.0LcalThe distance between measuring points is not more than 1/5Lcal；
7) And finishing the steps 1) to 6), and finishing the formulation of the transition process measurement scheme. Analyzing the heat flow or temperature data of the flight test acquired by the scheme, firstly preprocessing heat flow or temperature measurement data, identifying and eliminating invalid data including abnormal jumping points, dead points and burrs, and then drawing heat flow/temperature change curves along the flow direction at different moments by taking the flow direction length as an abscissa and the heat flow/temperature as an ordinate, wherein the heat flow/temperature change curves along the flow direction are shown in fig. 4;
8) analyzing the change rule of the heat flow/temperature along the flow direction change curve, taking the heat flow change curve analysis as an example, as shown in fig. 4, if the heat flow continues to rise significantly after a certain measuring point position, determining the position as a transition starting position at the moment; analyzing a heat flow curve behind the transition starting position, and if the heat flow does not rise or starts to fall after a certain measuring point position, judging the position as the transition ending position at the moment; the region between the start position and the end position of transition is the transition region at that moment, and the flow direction length L is used generallyexpCarrying out characterization; the information of the transition starting position, the transition region and the transition ending position determined in this way is the boundary layer transition process characteristic of the windward side of the aircraft. The method for judging the transition process characteristic by using the temperature change curve is consistent with the method. The above examples are provided only for illustrating the process of the present invention for a specific example, and the shape of the aircraft, the measurement area, the type of sensors used for measurement, the arrangement scheme of the measurement points, etc. may vary according to the actual application scenario. The above-mentioned embodiments are merely illustrative of the present invention, and should not be construed as limiting the present invention, so that all embodiments similar to the inventive concept are within the scope of the present invention.
Those matters not described in detail in the present specification are well known in the art.
Claims (7)
1. A method for measuring the transition process characteristic of an aircraft boundary layer in a hypersonic flight test is characterized by comprising the following steps:
1) carrying out simulation calculation on a flow field around the aircraft, and analyzing the distribution of surface streamlines and surface flow parameters through post-processing of the flow field to obtain the surface flow characteristics of the aircraft;
2) determining a region suitable for boundary layer transition process measurement on the surface of the aircraft according to the flow characteristics of the surface of the aircraft obtained in the step 1);
3) carrying out simulation calculation on the surface heat flow and the structure thermal response of the aircraft by using a CFD numerical method and a structure thermal analysis method, obtaining the calculation result of the surface heat flow and the structure temperature of the aircraft along the whole flight section, and carrying out model selection on a heat flow sensor and a temperature sensor according to the surface heat flow and the surface temperature data of a transition measurement area, wherein the model selection comprises the measuring range, the size and the material of the sensor;
4) comprehensively analyzing the size and the installation requirements of the sensor and the structural characteristics of the aircraft near the transition measurement region and the constraint of the installation space, and evaluating the installation feasibility of the sensor in the transition measurement region, wherein if the installation of the heat flow sensor and the temperature sensor is feasible, the heat flow sensor is preferably selected for surface heat flow measurement, and if only the installation of the temperature sensor is feasible, the temperature sensor is selected for surface temperature measurement;
5) drawing a surface streamline of a transition measurement area according to an analysis result of the surface flow characteristic of the aircraft, and arranging a row of sensor measuring points along the direction of the surface streamline; evaluating the transition characteristic of the surface boundary layer of the aircraft to obtain a transition region length calculation result L of the transition measurement regioncalDetermining the number of the sensor measuring points and the distance between the sensor measuring points by combining with the transition measurement precision requirement;
6) analyzing flight test data obtained by the transition process measurement scheme determined in the steps 1) -5), and drawing a curve of surface heat flow or temperature along the change of the flow direction;
7) and analyzing the change rule of the heat flow or the temperature along the flow change curve, wherein the information of the transition starting position, the transition region and the transition ending position is the boundary layer transition process characteristic.
2. The method for measuring the transition process characteristic of the boundary layer of the aircraft in the hypersonic flight test according to claim 1, is characterized in that: and in the step 1), a CFD numerical method is used for carrying out simulation calculation on the flow field around the aircraft.
3. The method for measuring the transition process characteristic of the boundary layer of the aircraft in the hypersonic flight test according to claim 1, is characterized in that: in step 2), the method for determining the transition measurement region is as follows:
2a) the measuring areas are positioned in the same profile of the aircraft;
2b) the flow in the measurement area is attached flow, the flow characteristics are simple, and the complex flow characteristics of separation reattachment and shock wave-boundary layer interference do not exist;
2c) the flow characteristics of the measuring region are unchanged in the flow direction and are not disturbed by the flow of surrounding components.
4. The method for measuring the transition process characteristic of the boundary layer of the aircraft in the hypersonic flight test according to claim 1, is characterized in that: when the temperature sensor is selected for surface temperature measurement, the local cabin material and the structure of the aircraft are designed, so that the cabin material and the thickness are kept consistent in the transition measurement region, and the thermal response characteristic of the structure in the transition measurement region is guaranteed to be consistent.
5. The method for measuring the transition process characteristic of the boundary layer of the aircraft in the hypersonic flight test according to claim 1, is characterized in that: and 5) evaluating the transition characteristic of the surface boundary layer of the aircraft by using a transition prediction method or a ground test result.
6. The method for measuring the transition process characteristic of the boundary layer of the aircraft in the hypersonic flight test according to claim 1, is characterized in that: in the step 5), when the number of the sensor measuring points and the distance between the sensor measuring points are determined, the sensor measuring points are ensured to cover the whole transition process, that is, the flow direction length covered by all the measuring points is not less than 2.0LcalThe distance between measuring points is not more than 1/5Lcal。
7. The method for measuring the transition process characteristic of the boundary layer of the aircraft in the hypersonic flight test according to claim 1, is characterized in that: the specific process of the step 6) is as follows: firstly, preprocessing heat flow or temperature measurement data is carried out, invalid data including abnormal jumping points, dead points and burrs are identified and removed, and then curves of heat flow or temperature change along the flow direction at different moments are drawn by taking the flow direction length as an abscissa and the heat flow or temperature as an ordinate.
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| CN113515804B (en) * | 2021-03-31 | 2023-04-14 | 北京临近空间飞行器***工程研究所 | Method for determining flow heat transfer in aircraft heat sealing structure |
| CN113221350B (en) * | 2021-05-10 | 2022-02-18 | 天津大学 | Hypersonic aircraft transition prediction method based on global stability analysis |
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| CN113532722B (en) * | 2021-05-25 | 2023-04-14 | 北京临近空间飞行器***工程研究所 | Flight test pulsating pressure data-based double-spectrum analysis transition identification method |
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| CN115356372B (en) * | 2022-10-24 | 2023-03-10 | 中国空气动力研究与发展中心计算空气动力研究所 | Time-varying thermal response testing method and system for novel material in flight test |
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