CN107832530A - A kind of hypersonic boundary layer transition decision method of complex appearance - Google Patents
A kind of hypersonic boundary layer transition decision method of complex appearance Download PDFInfo
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- CN107832530A CN107832530A CN201711116280.2A CN201711116280A CN107832530A CN 107832530 A CN107832530 A CN 107832530A CN 201711116280 A CN201711116280 A CN 201711116280A CN 107832530 A CN107832530 A CN 107832530A
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- G—PHYSICS
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- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
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- G—PHYSICS
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- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
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Abstract
A kind of hypersonic boundary layer transition decision method of complex appearance：(1) wind-tunnel inlet flow conditions and model contracting ratio are based on, using numerical simulation primary Calculation complex appearance test surfaces turn twist shape and position；(2) edge flows to array arrangement thin film resistance thermometer；(3) carry out the experiment of wind-tunnel calorimetric, obtain the heat flux distribution of different buses and the hot-fluid time-serial position of each measuring point on test surfaces；(4) it is based on wind-tunnel inlet flow conditions, calculates under the conditions of two kinds of holostrome stream and full turbulent flow the heat flux distribution of different buses and the hot-fluid time-serial position of corresponding measuring point on test surfaces；(5) for every bus determine every bus turn twist position；(6) according in step (5) determine every bus on turn twist original position and/or turn twist final position, original position and/or final position line are twisted by each turn, and do smoothing processing, it is final obtain complex appearance turn twist shape line and position.
Description
Technical field
The present invention can arrange thin film resistance thermometer by array, measure and judge superb under the auxiliary of numerical computations
The boundary layer transition position of velocity of sound complex appearance.
Background technology
Hypersonic boundary layer transition is widely present in the interior outflow of aerospace craft, and its aerodynamic force to aircraft/
Thermal force, control efficiency, propulsive performance etc. all have a major impact.The boundary layer transition of aircraft surface is measured by ground experiment
Position, necessary guidance can be provided for Flight Vehicle Design.
In hypersonic field, boundary layer transition can cause hot-fluid and the frictional resistance increase of aircraft wall, relative to laminar flow
Boundary layer, the hot-fluid and Value of Friction Loss of turbulent boundary layer will increase 3-5 times.Therefore, conventional boundary layer transition decision method is profit
With the increase of hot-fluid and frictional resistance, to judge to turn to twist starting point.However, compressing surface or prominent is commonly present on hypersonic aircraft precursor
Thing is played, it can cause the change of hot-fluid and frictional resistance, so that conventional boundary layer transition decision method failure, as U.S. X-51 flies
The precursor compressing surface of row device.
Transition measurement both domestic and external is concentrated mainly on the simple profile such as circular cone and flat board.The Development of Boundary Layer of this kind of profile exists
It is synchronous to open up to (or circumferential), it is only necessary to measures any one and flows to bus and turns to twist.On complex appearance, by three-dimensional
Flow over the influence in whirlpool, large-scale vortex structure etc., open up asynchronous to Development of Boundary Layer, and then cause to turn to twist position and larger difference be present
It is different.It is external to turn to twist frequently with the planar survey such as temperature-sensitive paint or infrared chart technology development complex appearance.The country, although phosphorescence thermal map
Technology development is more ripe, but the measurement accuracy of its low-heat stream there is no method to fully meet the demand of transition measurement.
The content of the invention
The technology of the present invention solves problem：A kind of overcome the deficiencies in the prior art, there is provided the high ultrasound of complex appearance
Fast boundary layer transition decision method, boundary layer transition measurement and the decision problem of complex appearance are realized, effectively solves complex appearance
Caused boundary layer transition is asynchronous and local interference caused by the problem of routinely turning to twist decision method failure.
Technical scheme one：A kind of hypersonic complex appearance turns to twist decision method, and step is as follows：
(1) wind-tunnel inlet flow conditions and model contracting ratio are based on, utilizes turning for numerical simulation primary Calculation complex appearance test surfaces
Twist shape and position；
(2) turn to twist the profile of shape and position with reference to aircraft according to the complex appearance test surfaces calculated in step (1),
Along array arrangement thin film resistance thermometer is flowed to, wherein array arrangement includes the bus quantity and every bus of parallel flow direction
On thin film resistance thermometer position be measuring point and quantity；
(3) carry out the experiment of wind-tunnel calorimetric, obtain the heat flux distribution of different buses on test surfaces；
(4) it is based on wind-tunnel inlet flow conditions, calculates the heat of different buses on test surfaces under the conditions of two kinds of holostrome stream and full turbulent flow
Flow distribution；
(5) for every bus respectively according to following processing determine every bus turn twist position：
The heat flux distribution that a certain bar bus is obtained in step (3) is corresponding with the conditions of holostrome stream in step (4) female first
The heat flux distribution of line is compared along flow direction, when the comparison difference for occurring a certain measuring point on the bus for the first time exceeds default model
When enclosing, then the measuring point is that this bus turns to twist original position；
Since above-mentioned determination turn to twist original position, by the heat flux distribution obtained in this bus step (3) with it is complete rapid
Heat flux distribution under the conditions of stream is compared along flow direction, when the comparison difference for occurring a certain measuring point on the bus for the first time enters in advance
If scope when, then the measuring point is that this bus turns to twist final position；
(6), will be each according to turning to twist original position and/or turn to twist final position on the every bus determined in step (5)
Turn to twist original position and/or final position line, and do smoothing processing, it is final obtain complex appearance turn twist shape line and position.
Technical scheme two：A kind of hypersonic complex appearance transition measurement and decision method, step are as follows：
(1) wind-tunnel inlet flow conditions and model contracting ratio are based on, utilizes turning for numerical simulation primary Calculation complex appearance test surfaces
Twist shape and position；
(2) turn to twist the profile of shape and position with reference to aircraft according to the complex appearance test surfaces calculated in step (1),
Along array arrangement thin film resistance thermometer is flowed to, wherein array arrangement includes the bus quantity and every bus of parallel flow direction
On thin film resistance thermometer position be measuring point and quantity；
(3) carry out the experiment of wind-tunnel calorimetric, obtain the hot-fluid time series signal of each measuring point on test surfaces；
(4) it is based on wind-tunnel inlet flow conditions, calculates the heat of each measuring point on test surfaces under the conditions of two kinds of holostrome stream and full turbulent flow
Flow time series signal；
(5) for every bus respectively according to following processing determine every bus turn twist position：
The hot-fluid time series signal for each measuring point that a certain bar bus is obtained in step (3) is designated as curve A, step
(4) corresponded under the conditions of holostrome stream corresponded on bus measuring point hot-fluid time series signal be designated as curve B and for curve B set miss
Difference band；Corresponded in step (4) under full turbulent-flow conditions corresponded on bus measuring point hot-fluid time series signal be designated as curve C and for song
Line C step-up error bands；
Along flow direction curve A is compared with curve B, when occur for the first time curve A fall into curve B error bands ratio it is small
When default value, measuring point corresponding to curve A be this bus turn twist original position；
Along flow direction curve A is compared with curve C, when occur for the first time curve A fall into curve C error band ratio it is big
When default value, measuring point corresponding to curve A be this bus turn twist final position；
(6), will be each according to turning to twist original position and/or turn to twist final position on the every bus determined in step (5)
Turn to twist original position and/or final position line, and do smoothing processing, it is final obtain complex appearance turn twist shape line and position.
Technical scheme three：A kind of hypersonic complex appearance transition measurement and decision method, step are as follows：
Be utilized respectively above two scheme first and obtain two groups of results, using two groups of results to every bus turn twist position
Verified, when the result of a certain bus is inconsistent, this bus iteration scheme one or scheme two are verified again
Or the result of selection scheme one is as final result.
Further, adjacent bus, which is extended to spacing, is not less than 10mm.
Further, the point position on every bus comprises at least turning to twist starting point, turning to twist end for the middle determination of step (1)
Stop, open up to turn twist line deviation point, turn twist starting point upstream position, turn twist terminating point downstream position.
Further, the wind-tunnel inlet flow conditions in step (4) are the wind-tunnel incoming that the experiment of wind-tunnel calorimetric determines in step (3)
Condition.
Further, the wind-tunnel inlet flow conditions in step (1) are the wind-tunnel inlet flow conditions estimated before wind tunnel experiment.
The present invention has the beneficial effect that compared with prior art：
Array of the present invention arranges thin film resistance thermometer, solves every streamline and turns to twist asynchronous caused transition measurement
Problem；The contrast of numerical computations is introduced, it is a certain degree of to eliminate separation flowing and compress the factor such as used to boundary layer transition
The influence of judgement；Pulsation hot-fluid judges to judge to be combined with heat flux distribution, improves the precision for turning to twist location determination.The present invention is
Complex appearance, which turns to twist to study, provides strong guarantee.
Brief description of the drawings
Fig. 1 is point layout schematic diagram；
Fig. 2 is that the boundary layer fluidised form based on heat flux distribution judges schematic diagram；
Fig. 3 is that the boundary layer fluidised form based on hot-fluid time series signal judges schematic diagram；
Fig. 4 is the boundary layer transition original position line schematic diagram on model multi-stage compression face.
Embodiment
Below in conjunction with the accompanying drawings and example elaborates to the present invention.A kind of hypersonic border of complex appearance of the present invention
It is as follows that layer turn twists decision method step：
(1) it is based under the conditions of wind tunnel test, carries out the numerical computations of compressible turbulent flow-Transition model for complex appearance,
The turning of preliminary prediction model experiment face twists form and position, instructs the point layout of wind tunnel test.
(2) according to above-mentioned numerical result estimate turn twist form and position, targetedly arrange film resistor temperature
Degree meter, including turn twist original position, turn twist final position, open up to turn twist nonsynchronous position etc..When yaw angle is not present in model
When, face symmetric profile can only arrange side measuring point, as shown in Figure 1.
(3) carry out the wind-tunnel calorimetric test of complex appearance, obtain the hot-fluid time series of each measuring point on model measurement face
The heat flux distribution of signal and different buses.
(4) wind-tunnel inlet flow conditions and model attitude are based on, carries out laminar flow Navier-Stokes equations numerical simulation and turbulent flow
RANS is simulated, and calculates the hot-fluid of complex appearance test surfaces under holostrome stream and full turbulent-flow conditions, is obtained corresponding with wind tunnel test
The heat flux distribution of hot-fluid time series signal and different buses.
(5) in terms of heat flux distribution and hot-fluid time-serial position two, comparative test result and result of calculation, realize multiple
The boundary layer transition that every of miscellaneous aircraft profile flows to bus judges.
● a) by the heat flux distribution q (x) for every bus for contrasting wind tunnel test and numerical computations, determine every bus
Turn to twist original position and final position, as shown in Figure 2.
The heat flux distribution curve for every bus that wind tunnel test obtains is designated as A；Under the conditions of the holostrome stream that numerical computations obtain
The heat flux distribution curve of corresponding bus is designated as B；The heat flux distribution curve of bus is corresponded under the full turbulent-flow conditions that numerical computations obtain
It is designated as C；The preset range of wind tunnel test and numerical result contrast is 10%.
First, heat flux distribution curve A and B are contrasted, when the contrast difference for occurring a certain measuring point for the first time on the bus exceeds
During default scope, then the measuring point is that this bus turns to twist original position；
Then, since above-mentioned determination turn twist original position, heat flow curve A and C is contrasted.When on the bus
When the comparison difference for occurring a certain measuring point for the first time enters default scope, then the measuring point is that this bus turns to twist final position.
● b) by the hot-fluid time series signal q for each measuring point of every bus for contrasting wind tunnel test and numerical computations
(t), determine every bus turn twist original position and final position, as shown in Figure 3.
The hot-fluid time series signal that wind tunnel test obtains each measuring point of a certain bar bus is designated as curve A；Numerical value meter
Corresponded in calculation under the conditions of holostrome stream and the hot-fluid time series signal of measuring point is corresponded on bus be designated as curve B；It is complete rapid in numerical computations
Corresponded under the conditions of stream and the hot-fluid time series signal of measuring point is corresponded on bus be designated as curve C；Curve B and C error band are 10%.
Along flow direction curve A is compared with curve B, when occur for the first time curve A fall into curve B error bands ratio it is small
In it is default value (95%) when, measuring point corresponding to curve A be this bus turn twist original position；
Along flow direction curve A is compared with curve C, when occur for the first time curve A fall into curve C error band ratio it is big
In it is default value (95%) when, measuring point corresponding to curve A be this bus turn twist final position.
(6) in step (5) two methods determine every bus turn twist position (including turn twist original position and turn twist
Final position) it is compared.When result is inconsistent, for bus repeat step (1)-(5), or heat flux distribution curve is true
Fixed result is as final result.
(7) on the every bus judged according to the above method turn twist original position, twist starting point line by each turn,
And do smoothing processing, it is final obtain complex appearance turn twist shape line and position, as shown in Figure 4.
The present invention can also use following technological means, for example, be utilized respectively first it is above-mentioned a) in heat flux distribution q (x), b)
The mode of middle hot-fluid time series signal q (t) obtain complex appearance turn twist shape line and position, obtain two groups of results, utilize
Two groups of results turn to twist position and verify to every bus, and when the result of a certain bus is inconsistent, this bus is repeated
Heat flux distribution q (x), hot-fluid time series signal q (t) mode content carry out verifying or selecting a) heat flux distribution q (x) again
Result as final result.
Unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.
Claims (7)
1. the hypersonic boundary layer transition decision method of a kind of complex appearance, it is characterised in that step is as follows：
(1) wind-tunnel inlet flow conditions and model contracting ratio are based on, using numerical simulation primary Calculation complex appearance test surfaces turn twist shape
Shape and position；
(2) turn to twist the profile of shape and position with reference to aircraft according to the complex appearance test surfaces calculated in step (1), along stream
Thin film resistance thermometer is arranged to array, wherein in bus quantity of the array arrangement comprising parallel flow direction and every bus
Thin film resistance thermometer position is measuring point and quantity；
(3) carry out the experiment of wind-tunnel calorimetric, obtain the heat flux distribution of different buses on test surfaces；
(4) it is based on wind-tunnel inlet flow conditions, calculates the hot-fluid point of different buses on test surfaces under the conditions of two kinds of holostrome stream and full turbulent flow
Cloth；
(5) for every bus respectively according to following processing determine every bus turn twist position：
By the heat flux distribution that a certain bar bus is obtained in step (3) first with corresponding to bus under the conditions of holostrome stream in step (4)
Heat flux distribution is compared along flow direction, when the comparison difference for occurring a certain measuring point on the bus for the first time exceeds default scope
When, then the measuring point is that this bus turns to twist original position；
Since above-mentioned determination turn twist original position, by the heat flux distribution obtained in this bus step (3) and full turbulent flow bar
Heat flux distribution under part is compared along flow direction, when the comparison difference for occurring a certain measuring point on the bus for the first time enter it is default
During scope, then the measuring point is that this bus turns to twist final position；
(6) according in step (5) determine every bus on turn twist original position and/or turn twist final position, each turn is twisted
Original position and/or final position line, and do smoothing processing, it is final obtain complex appearance turn twist shape line and position.
2. the hypersonic boundary layer transition decision method of a kind of complex appearance, it is characterised in that step is as follows：
(1) wind-tunnel inlet flow conditions and model contracting ratio are based on, using numerical simulation primary Calculation complex appearance test surfaces turn twist shape
Shape and position；
(2) turn to twist the profile of shape and position with reference to aircraft according to the complex appearance test surfaces calculated in step (1), along stream
Thin film resistance thermometer is arranged to array, wherein in bus quantity of the array arrangement comprising parallel flow direction and every bus
Thin film resistance thermometer position is measuring point and quantity；
(3) carry out the experiment of wind-tunnel calorimetric, obtain the hot-fluid time series signal of each measuring point on test surfaces；
(4) it is based on wind-tunnel inlet flow conditions, calculates under the conditions of two kinds of holostrome stream and full turbulent flow on test surfaces during the hot-fluid of each measuring point
Between sequence signal；
(5) for every bus respectively according to following processing determine every bus turn twist position：
The hot-fluid time series signal for each measuring point that a certain bar bus is obtained in step (3) is designated as curve A, step (4)
Corresponded under the conditions of middle holostrome stream and the hot-fluid time series signal of measuring point is corresponded on bus be designated as curve B and be curve B step-up errors
Band；Corresponded in step (4) under full turbulent-flow conditions and the hot-fluid time series signal of measuring point is corresponded on bus be designated as curve C and be curve
C step-up error bands；
Curve A is compared with curve B along flow direction, the ratio of curve B error bands is fallen into less than pre- when occurring curve A for the first time
If value when, measuring point corresponding to curve A be this bus turn twist original position；
Curve A is compared with curve C along flow direction, the ratio of curve C error band is fallen into more than pre- when occurring curve A for the first time
If value when, measuring point corresponding to curve A be this bus turn twist final position；
(6) according in step (5) determine every bus on turn twist original position and/or turn twist final position, each turn is twisted
Original position and/or final position line, and do smoothing processing, it is final obtain complex appearance turn twist shape line and position.
A kind of 3. hypersonic boundary layer transition decision method of complex appearance, it is characterised in that：
Claim 1 or 2 is utilized respectively first and obtains two groups of results, and the position that turns to twist of every bus is carried out using two groups of results
Checking, when the result of a certain bus is inconsistent, to this bus repeat claim 1 or 2 content carry out again checking or
Person selects the result of claim 1 as final result.
4. according to the method described in claim 1 or 2 or 3, it is characterised in that：Adjacent bus, which is extended to spacing, is not less than 10mm.
5. according to the method described in claim 1 or 2 or 3, it is characterised in that：Point position on every bus comprises at least step
Suddenly what is determined in (1) turns to twist starting point, turns to twist terminating point, opens up to turning to twist line deviation point, turn to twist starting point upstream position, turn to twist end
Stop downstream position.
6. method according to claim 1 or 2, it is characterised in that：Wind-tunnel inlet flow conditions in step (4) are step (3)
The wind-tunnel inlet flow conditions that middle wind-tunnel calorimetric experiment determines.
7. method according to claim 1 or 2, it is characterised in that：Wind-tunnel inlet flow conditions in step (1) are wind tunnel experiment
Before the wind-tunnel inlet flow conditions estimated.
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109033525A (en) * | 2018-06-27 | 2018-12-18 | 浙江大学 | A kind of hypersonic transition prediction method based on simplified three equation transition models |
| CN110806300A (en) * | 2019-10-12 | 2020-02-18 | 北京临近空间飞行器系统工程研究所 | Measuring point arrangement method suitable for hypersonic flight test transition research |
| CN111551341A (en) * | 2020-05-29 | 2020-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Low-temperature transonic equipment TSP transition measurement test method |
| CN112304563A (en) * | 2020-10-30 | 2021-02-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Wind tunnel test method for researching influence of transition on aerodynamic characteristics of hypersonic aircraft |
| CN112837291A (en) * | 2021-02-03 | 2021-05-25 | 中国空气动力研究与发展中心高速空气动力研究所 | Laminar flow wing transition position measurement image processing method based on temperature-sensitive paint technology |
| CN116612049A (en) * | 2023-07-18 | 2023-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Transition line extraction method based on temperature distribution image |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103954425A (en) * | 2014-04-30 | 2014-07-30 | 北京大学 | Hypersonic velocity static wind tunnel nozzle design method and hypersonic velocity static wind tunnel nozzle transition position determining method |
| CN106908352A (en) * | 2017-02-22 | 2017-06-30 | 西北工业大学 | Airfoil surface boundary layer transition location measurement method based on distributed temperature sensitive optical fiber |
| US20170240271A1 (en) * | 2015-11-11 | 2017-08-24 | The Arizona Board Of Regents On Behalf Of The University Of Arizona | Control of hypersonic boundary layer transition |
-
2017
- 2017-11-13 CN CN201711116280.2A patent/CN107832530B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103954425A (en) * | 2014-04-30 | 2014-07-30 | 北京大学 | Hypersonic velocity static wind tunnel nozzle design method and hypersonic velocity static wind tunnel nozzle transition position determining method |
| US20170240271A1 (en) * | 2015-11-11 | 2017-08-24 | The Arizona Board Of Regents On Behalf Of The University Of Arizona | Control of hypersonic boundary layer transition |
| CN106908352A (en) * | 2017-02-22 | 2017-06-30 | 西北工业大学 | Airfoil surface boundary layer transition location measurement method based on distributed temperature sensitive optical fiber |
Non-Patent Citations (4)
| Title |
|---|
| S.F.XIE 等: "Research of flow field characteristics of hypersonic shock wave / transitional boundary layer interaction", 《20TH AIAA INTERNATIONAL SPACE PLANES AND HYPERSONIC SYSTEMS AND TECHNOLOGIES CONFERENCE》 * |
| TANNO HIDEYUKI 等: "Measurement of hypersonic boundary layer transition on cone models in the free-piston shock tunnel HIEST", 《47TH AIAA AEROSPACE SCIENCES MEETING INCLUDING THE NEW HORIZONS FORUM AND AEROSPACE EXPOSITION》 * |
| 宋博 等: "高超声速边界层转捩及基于层流脉动能的转捩预测", 《飞航导弹》 * |
| 解少飞 等: "边界层转捩与压缩拐角分离流动的非定常作用", 《空气动力学学报》 * |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109033525A (en) * | 2018-06-27 | 2018-12-18 | 浙江大学 | A kind of hypersonic transition prediction method based on simplified three equation transition models |
| CN109033525B (en) * | 2018-06-27 | 2022-08-30 | 浙江大学 | Hypersonic transition prediction method based on simplified three-equation transition model |
| CN110806300A (en) * | 2019-10-12 | 2020-02-18 | 北京临近空间飞行器系统工程研究所 | Measuring point arrangement method suitable for hypersonic flight test transition research |
| CN111551341A (en) * | 2020-05-29 | 2020-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Low-temperature transonic equipment TSP transition measurement test method |
| CN112304563A (en) * | 2020-10-30 | 2021-02-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Wind tunnel test method for researching influence of transition on aerodynamic characteristics of hypersonic aircraft |
| CN112837291A (en) * | 2021-02-03 | 2021-05-25 | 中国空气动力研究与发展中心高速空气动力研究所 | Laminar flow wing transition position measurement image processing method based on temperature-sensitive paint technology |
| CN112837291B (en) * | 2021-02-03 | 2022-07-29 | 中国空气动力研究与发展中心高速空气动力研究所 | Laminar flow wing transition position measurement image processing method based on temperature-sensitive paint technology |
| CN116612049A (en) * | 2023-07-18 | 2023-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Transition line extraction method based on temperature distribution image |
| CN116612049B (en) * | 2023-07-18 | 2023-10-10 | 中国空气动力研究与发展中心高速空气动力研究所 | Transition line extraction method based on temperature distribution image |
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